Embed Size (px)
Citation preview
jt*1 A
CO CO ID
c o o
i^^-^Jf
Outer Planet Entry Probe System Studys^^ Final Report
Volume IV Common Saturn/Uranus Probe Studies January 1973
MJWTiiV MJmiETTA
https://ntrs.nasa.gov/search.jsp?R=19730011156 2020-02-03T18:08:13+00:00Z
JPL Contract 953311
Volume IV
Common Saturn/Uranus Probe Studies January 1973
OUTER PLANET ENTRY PROBE SYSTEM STUDY
FINAL REPORT
Approved
R. S. Wiltshire Program Manager
This work was performed for the jet Propulsion Laboratory, Califoenia Institute of Technology, sponsored by die National Aeronautics and Space Administration under Contract NAS7-100,
MARTIN MARIETTA CORPORATION P.O. Box 179 Denver, Colorado 80201
FOREWORD
This final report has been prepared in accordance with requirements of Contract JPL-953311, Modification No. 7, Task Order No. RD-10,:(follow-on to the basic contract) to present data and conclusions drawn from a three-month study by Martin Marietta Corporation, Denver Division, for the Jet Propulsion Laboratory. The report is Volume IV; the first three volumes documented the basic contract results. The four volumes are:
Volume I - Summary
Volume II - Supporting Technical Studies
Volume III - Appendixes
Volume IV - Common Saturn/Uranus Probe Studies.
iii
ACKNOWLEDGEMENTS
The following Martin Marietta sonnel participated in this st appreciated:
Raymond S. Wiltshire Allen R. Barger Eugene A. Berkery
Dennis V. Byrnes Philip C. Carney Patrick C. Carroll Revis E. Compton, Jr. Robert G. Cook W. Sidney Cook Douglas B. Cross Ralph F. Fearn Robert B. Fischer Thomas C. Hendricks John W. Hungate Carl L. Jensen Rufus 0. Moses
Kenneth W. Ledbetter Paula S. Lewis John R. Mellin Jack D. Pettus Robert J. Richardson Arlen I. Reichert E. Doyle Vogt Donald E. Wainwright Clifford M. Webb Charles E. Wilkerson
irporation, Denver Division, per-ly, and their efforts are greatly
Study Leader, Program Manager Science Integration Telecommunications, Data Handling, Power, and ACS, Lead
Navigation Mission Analysis Systems Telecommunications Mechanical Design Cloud Structures Analysis Mission Analysis Propulsion Mission Analysis Mission Analysis, Lead Systems, Lead Thermal Analysis Mechanical/Structural/Probe Integration, Lead
Science, Lead Mission Analysis Structures Data Link Analysis Receiver Systems Propulsion Mission Analysis Systems Thermal Analysis Data Handling
IV
CONTENTS
Page
INTRODUCTION 1-1 and 1-2
SUMMARY, CONCLUSIONS AND RECOMMENDATIONS II-l Summary II-l Conclusions and Recommendations 11-21
thru 11-25
SATURN/URANUS PARAMETRIC ANALYSIS III-l Mission Design Considerations III-2 Science Investigations 111-24 System Integration 111-51 Electrical and Electronics Subsystems • 111-57 Thermal Control Subsystem . 111-97 Parametric Analysis Summary III-104 References III-109
SATURN/URANUS PROBE SYSTEM DEFINITION WITH SCIENCE ADVISORY GROUP'S EXPLORATORY PAYLOAD IV-1 Mission Definition IV-1 Science Instrumentation and Performance IV-11 System Design and Integration IV-14 Telecommunications Subsystem IV-21 Data Handling Subsystem IV-29 Power and Pyrotechnic Subsystem IV-31 Attitude Control IV-33 Structural and Mechanical Subsystems IV-35 Propulsion Subsystem IV-48 Thermal Control Subsystem IV-51 Probe-to-Spacecraft Integration IV-56
thru IV-59
SATURN/URANUS PROBE SYSTEM DEFINITION WITH EXPANDED SCIENCE COMPLEMENT V-l Mission Definition V-l Science Instrumentation and Performance V-l System Design and Integration V-6 Telecommunications Subsystem V-ll
E. Data Handling Subsystem V-18 F. Power and Pyrotechnic Subsystem V-19 G. Attitude Control V-21 H. Structural and Mechanical Subsystem . V-23 I. Propulsion Subsystem V-29 J. Thermal Control Subsystem V-29 K. Probe-to-Spacecraft Integration V-33
Figure
1-1 Follow-On Study Task Definition and Flow Diagram . . . . 1-1 II-l Science Instrument Locations, Expanded
Science Payload II-5 II-2 Pre-entry Measurement Performance II-6 II-3 Science Constraints on Mission Analysis II-8 II-4 Mission Design Configurations II-8 II-5 Saturn (SU 80) Mission II-9 II-6 Uranus (SU 80) Mission Design II-9 II-7 Pictorial Sequence of Events (Task II) 11-10 II-8 Data Profile (Task II) 11-12 II-9 Power Profile (Task II) 11-12 11-10 Probe Configuration Description 11-16 11-11 General Entry Probe Configuration 11-16 11-12 Entry Probe Configuration Comparison 11-17 11-13 Descent Probe Configuration Comparison 11-17 11-14 Probe Integration with Mariner Spacecraft 11-19 11-15 Cloud Locations and End-of-Mission Limits
for Various Model Atmospheres 11-20 III-l 1979 Saturn Interplanetary Trajectory III-3 III-2 1980 Saturn/Uranus Interplanetary Trajectory III-3 III-3 Estimated Shuttle Performance Data III-5 III-4 Titan Ill/Centaur Performance Data III-5 III-5 Saturn 1979 and Saturn/Uranus 1980 Launch
Energy Contours III-8 III-6 Saturn and Uranus Entry Velocity Variations III-8 III-7 Vu_,, ZAE, and ZAP Variations with Arrival Date III-9
HP III-8 Navigation Uncertainties 111-10 III-9 Saturn and Uranus Look-Direction Dispersions 111-12 111-10 Saturn Entry Site Dispersions 111-13 III-ll Uranus Entry Site Dispersions 111-13 111-12 Relay Link Parameters 111-15 111-13 Deceleration to Subsonic Speeds at Saturn 111-21 111-14 Deceleration to Subsonic Speeds at Uranus 111-21 111-15 Maximum Deceleration at Saturn and Uranus
vs Entry Angle 111-22
vi
111-16 Maximum Dynamic Pressure at Saturn and Uranus vs Entry Angle 111-22
111-17 Saturn Deceleration Profile 111-23 111-18 Uranus Deceleration Profile 111-23 111-19 Pressure vs Temperature and Cloud Levels
for Saturn Model Atmospheres 111-25 111-20 Pressure vs Temperature for Uranus Model
Atmospheres 111-25 111-21 Pressure vs Altitude for Saturn Model
Atmospheres 111-26 111-22 Pressure vs Altitude for Saturn Model
Atmospheres 111-26 111-23 Ion-Retarding Potential Analyzer .111-31 111-24 Neutral-Retarding Potential Analyzer 111-33 111-25 Langmuir Probe and RPA Location 111-36 111-26 Descent Side-Looking Nephelometer 111-38 111-27 Pre-entry Velocity/Altitude Time History 111-45 111-28 Probe Velocities at Entry vs Saturn
Entry Angle 111-45 111-29 Pre-entry Performance vs Saturn Entry Angle 111-47 111-30 Pre-entry Instrument Measurement Performance
vs Sampling Time 111-47 111-31 Pressure Descent Profiles 111-50 111-32 Parachute Deployment Analysis,
Saturn and Uranus . 111-52 111-33 Data Profile for Task II Data Mode Tradeoff 111-54 111-34 Preliminary Data Profile for Saturn/Uranus
Probe (Task II) 111-56 111-35 Thermal Noise Temperature of the Milky Way
Galaxy 111-60 111-36 Trajectory Geometry for the Saturn/Uranus
1980 Mission 111-60 111-37 Receiving System Noise Temperature for Saturn 111-62 111-38 Receiving System Noise Temperature for Uranus 111-62 111-39 Location of Clouds That Affect Microwave
Absorption 111-64 111-40 Absorption Coefficient for the Cool
Jupiter Atmosphere 111-66 111-41 Zenith Absorption for Saturn Nominal Atmosphere . . . . 111-67 111-42 Zenith Absorption at Selected Pressures for the
Saturn Nominal Atmosphere 111-67 111-43 Zenith Absorption for the Saturn Cool Atmosphere . . . . 111-68 HI-44 Zenith Absorption at Selected Pressures for the
Saturn Cool Atmosphere 111-68 111-45 Zenith Absorption for the Uranus Nominal
Atmosphere 111-69
vii
111-46 Zenith Absorption at Selected Pressures for the Uranus Nominal Atmosphere 111-69
111-47 Circularly Polarized Turnstile/Cone Probe Antenna Patterns 111-71
111-48 Probe Turnstile/Cone Antenna 111-73 111-49 Prototype Turnstile/Cone Antenna Mounted
on Probe Model 111-74 111-50 Antenna Pattern Polarization Loss 111-76 111-51 Probe Aspect Angle for the Three Missions 111-82 111-52 Spacecraft Antenna Requirements for the
Three Missions 111-83 111-53 Definition of Spacecraft-to-Probe Look Direction . . . . 111-85 111-54 Data Handling Subsystem Functional Diagram 111-85 111-55 Power and Pyrotechnic Subsystem Block Diagram 111-90 111-56 Probe Thermal Control Concept 111-98 111-57 Worst-Case Saturn/Uranus Atmospheric
Temperature Profiles III-100 111-58 Cruise/Coast Probe Equilibrium Temperatures
Based on Thermal Coating Selection III-100 111-59 Cloud Locations and End-of-Mission Limits
for Various Model Atmospheres III-108 IV-1 1979 Saturn Mission Definition IV-2 IV-2 Saturn/SU 80 Mission Definition IV-5 IV-3 Uranus/SU 80 Mission Definition IV-9 IV-4 Sequence of Events (Task II) IV-14 IV-5 Data Profile for Saturn/Uranus Probe (Task I) IV-18 IV-6 Power Profile for Saturn/Uranus Probe (Task I) IV-19 IV-7 Probe-to-Spacecraft Communications Range for
the Saturn/SU 80 Mission IV-22 IV-8 Saturn/SU 80 Mission Probe Aspect Angle IV-23 IV-9 Probe RF Power Requirements (Task I) IV-26 IV-10 Saturn/Uranus Common Entry Probe (Task I) IV-37 IV-11 Saturn/Uranus Common Entry Probe
Structural Breakdown . IV-41 IV-12 Attitude Control Propulsion System IV-49 IV-*13 Saturn/Uranus Probe Thermal History for
Saturn Mission IV-54 IV-14 Saturn/Uranus Probe Thermal History for
Uranus Mission IV-55 IV-15 Probe and Spacecraft Interface Arrangement IV-57 V-l Data Profile for Saturn/Uranus Probe (Task II) V-10 V-2 Power Profile for Saturn/Uranus Probe (Task II) . . . . V-10 V-3 Task II Probe RF Power Requirements V-14 V-4 Saturn/Uranus Common Entry Probe, Task II V-25 V-5 Saturn/Uranus Probe Thermal History for Saturn
Mission (Expanded Science Payload) V-31 V-6 Saturn/Uranus Probe Thermal History for Uranus
Mission (Expanded Science Payload) V-32 viii
Table
II-l Study Constraints II-2 II-2 Science Instruments Related to Measurements II-3 II-3 Science Instrument Characteristics II-3 II-4 Descent Measurement Performance II-6 II-5 Mission Design Summary 11-10 II-6 Telecommunication Summary 11-14 II-7 Power and Pyrotechnic Summary 11-14 II-8 Structural/Mechanical Summary 11-18 II-9 Weight Comparisons 11-18 11-10 Probe Comparison Summary 11-19 11-11 Compromises for Commonality 11-22 11-12 Probe Subsystem Development Status 11-23 III-l Saturn Direct '79 Mission Launch Capability III-4 III-2 Saturn/Uranus '80 Mission Launch Capability III-4 III-3 Mission Launch Energy Requirements III-4 III-4 Navigation Uncertainties at Saturn (la) III-ll III-5 Entry and Communication Parameter Dispersions 111-14 III-6 Saturn Model Atmosphere Entry Parameter
Definition 111-19 III-7 Uranus Model Atmosphere Entry Parameter
Definition 111-19 III-8 Saturn/Uranus Modeled Cloud Locations 111-27 III-9 Instruments Related to Measurements 111-28 111-10 IRPA Characteristics 111-30 III-ll NRPA Characteristics 111-32 111-12 Langmuir Probe Characteristics 111-36 111-13 Nephelometer Characteristics 111-38 111-14 Variation of Parameters with Atmosphere 111-40 111-15 Start of Pre-entry Measurements 111-43 111-16 Latitude and Longitude Variations 111-46 111-17 Accelerometer Measurement Performance 111-48 111-18 Descent Data Rate Composition 111-50 111-19 Task II Data Mode Analysis Results 111-55 111-20 Ammonia Abundance in the Atmosphere Models 111-65 111-21 Telecommunication Subsystem Parameters for the
Common Probe Development 111-78 111-22 Telecommunications Subsystem Characteristics
Comparison Summary 111-80 111-23 Typical Probe Equipment Temperature Limits III-101 111-24 Comparison of Mission Designs III-105 111-25 Science Analysis Summary III-105 111-26 Parachute Deployment Results III-106 111-27 Task II Data Mode Analysis Results III-107 111-28 Communications Parameters III-108 IV-1 Saturn 1979 Mission Summary Data IV-3 IV-2 Saturn/SU 80 Mission Summary Data IV-6
ix
IV-3 Uranus/SU 80 Mission Summary Data IV-10 IV-4 Instrument Sampling Times and Data Rates for
Common Saturn/Uranus Probe with Exploratory Payload . . IV-12 IV-5 Saturn Descent Measurement Performance,
Exploratory Payload IV-13 IV-6 Uranus Descent Measurement Performance,
Exploratory Payload IV-13 IV-7 Saturn Sequence of Events (Task I) IV-15 IV-8 Uranus "Sequence of Events (Task I) IV-16 IV-9 Weight Summary for Saturn/Uranus Probe with
SAG Exploratory Payload IV-20 IV-10 Probe Telemetry Link Design for the Saturn/SU
80 Mission with Task I Science IV-27 IV-11 Telecommunications RF Subsystem for the Task I
Mission IV-28 IV-12 Baseline Common Saturn/Uranus Probe Weight
Breakdown IV-39 IV-13 Minimum and Maximum Probe Temperatures for
Worst-Case Atmospheric Encounter and Arrival Uncertainties IV-53
V-l Instrument Characteristics V-2 V-2 Instrument Sampling Times and Data Rates for
Common Saturn/Uranus V-3 V-3 Saturn Descent Measurement Performance,
Expanded Payload V-5 V-4 Uranus Descent Measurement Performance,
Expanded Payload V-5 V-5 Pre-entry Measurement Performance V-6 V-6 Saturn Sequence of Events (Task II) V-7 V-7 Uranus Sequence of Events (Task II) V-8 V-8 Weight Summary for Saturn/Uranus Probe with
Expanded Science Complement . . . . . V-ll V-9 Task II Telecommunication Subsystem Comparisons
for the Saturn/SU 80 Mission V-ll V-10 Probe Pre-entry Telemetry Link Design for
Task II Mission V-13 V-ll Probe Descent Telemetry Link Design for
Task II Mission V-16 V-12 Telecommunications RF Subsystem for Task II Mission . . V-17 V-13 Task II Saturn/Uranus Probe Weight Breakdown
(Added Science) V-28 V-14 Minimum and Maximum Probe Temperatures for
Worst-Case. Atmospheric Encounters and Arrival Uncertainties V-33
x
I. INTRODUCTION
The material in this volume summarizes results of a common scientific probes study to explore the atmospheres of Saturn and Uranus, This was a three-month follow-on effort to the Outer Planet Entry Probe System study. Two major tasks, shown in Figure 1-1, and their relationship to other tasks in the basic study were considered. The first task established a reference definition that was used to determine the impact of the additional science complement in the second task.
h-
SATURN PARAMETRIC ANALYSIS (VOL II, SECT VI-A)
SATURN PROBE DEFINITION, JST 77 (VOL II, SECT VI-B)
URANUS PARAMETRIC ANALYSIS (VOL II, SECT VII-A)
SATURN PROBE APPLICABILITY TO URANUS, JU 79 (VOL II, SECT VII-B)
•BASIC STUDY-
fce-..>.ttty?:"
(TASK I) DEFINE A COMMON SATURN/URANUS PROBE USING SCIENCE ADVISORY GROUP EXPLORATORY PAYLOAD
(TASK II) DEFINE A COMMON SATURN/URANUS PROBE USING ADDED SCIENCE
SATURN DIRECT '79 SATURN/SU 80 URANUS/SU 80 URANUS/JU 79 SPACECRAFT DEFLECTION WORST-CASE ATMOSPHERE
-FOLLOW-ON STUDY-
Fig. 1-1 Follow-on Study Task Definition and FLow Diagram
The report is arranged to present (1) a summary, conclusions and recommendations of this study, (2) parametric analysis conducted to support the two system definitions, (3) common Saturn/Uranus probe system definition using the Science Advisory Group's (SAG) exploratory payload and, (4) common Saturn/Uranus probe system definition using an expanded science complement. Each of the probe system definitions consists of detailed discussions of the
1-1
mission, science, system and subsystems including telecommunications, data handling, power, pyrotechnics, attitude control, structures, propulsion, thermal control and probe-to-spacecraft integration. References are made to the contents of the first three volumes where it is feasible to do so.
1-2
II. SUMMARY, CONCLUSIONS AND RECOMMENDATIONS
A. SUMMARY
1. General
The study objectives consist of using configuration concepts and subsystem definition data generated in the basic study, and extending this data to define a common:
1) Saturn/Uranus probe using the Science Advisory Group (SAG) exploratory payload (Task I);
2) Saturn/Uranus probe using an expanded science payload (Task ID.
The constraints for the study are shown in Table II-l. The Saturn Direct 79 and Uranus (SU 80) missions are terminal missions at the respective planets, and Saturn/SU 80 is a flyby mission. The science instruments, used for Task I, are the same ones used for the basic study and are discussed in detail in Volume II, Chapter III. The pre-entry Task II science instruments provide data to define the structure and composition of the upper atmosphere and the ionosphere. The added Task II descent,instrument provides data to more precisely identify location of major cloud formations.
The spacecraft deflection mode, used for this study, provides a simplier probe definition than the probe deflection mode that was used in the basic study.
The effects of warm, nominal, and cool atmospheric models were evaluated for Saturn and Uranus and the "worst case" was used for the common probe definition. In the basic study, only the nominal atmospheric model was used.
2. Science
The basic science objectives are to:
1) determine the structure and composition of the bulk atmosphere;
2) define the density, pressure, and temperature profiles;
3) determine the composition of the major clouds.
II-l
Table II-l Study Constraints
Missions: Saturn Direct 79, Saturn (SU'80) and Uranus (SU'80)
Science Instruments:
Task I (SAG Exploratory Payload)
Neutral Mass Spectrometer \
Temperature Gage / Descent
Pressure Gages J
Accelerometers - Entry and Descent
Task I I (Preentry Instruments)
Langmuir Probes
Ion Retarding Potential Analyzer (IRPA)
Neutral Particle Retarding Potential Analyzer (NRPA)
Task II (Descent Additional Instrument)
Nephelometer
Deflection Mode: Spacecraft
Atmospheric Models: Worst Case at Saturn and Uranus
Entry Angles, Latitude and Descent Depth: Compatible with Mission Set and Science Objectives
Secondary objectives, requiring the addition of extra instrumentation are to:
1) identify the location of the major cloud formations;
2) define the structure and composition of the upper atmosphere;
3) define the structure and composition of the dayside ionosphere.
The science instruments and their relationship to the measurements for the three science mission phases of pre-entry, entry, and descent are shown in Table II-2. The entry and descent instruments (except for the nephelometer) comprise the Task I science payload. At least one instrument contributes directly to all descent measurements except the atmospheric turbulence measurement. In this application, four instruments provide related data from which information is derived.
The science instrument characteristics for both study tasks are shown in Table II-3 in terms of weight, volume, power requirement, sampling times, and science data rate. Task II totals for weight, volume, and power are approximately double that for Task I. Although the science weight differs by 6.4 kg, the total entry probe weight difference is only 14 kg, and the probe diameter difference is only- 5 cm.
II-2
Table II-2 Soienoe Instruments Related to Measurements
****•—^^^ Instruments
Measurements ' -—.^
P reentry Ion Concentration Profiles Neutral Concentration
Profiles Electron Density & Temp
Entry Deceleration Loads
Descent H/He Ratio Isotopic Ratios Minor Constituents Atmospheric Temperature Atmospheric Pressure Cloud Location/Structure Cloud Composition Atmospheric Turbulence
Temp Gage
R N R D R R R R
Press Gage
R N R R D R R R
Acceler-ometers
D
R N N R R R R R
Neutral Mass Spec
D D D R R R D N
Nephe-lometer
N N R N N D R R
Langmuir Probes
D
N D
IRPA
D
R R
NRPA
N
D N
D = Direct Measurement, R = Related Measurement, N = Little or No Relation
Table II-S Science Instrument Characteristics
Task I, Exploratory Pay load Neutral Mass Spectrometer Temperature Gage Pressure Gages Accelerometers: Entry
Turbulence
Totals - Task I
Weight, kg
5.45 0.45 0.68 1.50 Same
8.08
Task 11, Expanded Payload Additions Nephelometer (Descent) 1.14 P reentry
Langmuir Probes 1.36 Ion Retarding Potential Analyzer 1.59 Neutral Particle RPA 2.27
Volume, Power, Sampling Science Data cm3 Watts Time, sec Rate, bps
5,314 426 246 918
Same
6,904
1,312
1,200 2,200 2,500
14.0 1.4 1.3 2.8
Same
19.5
3.0
3.0 3.0 5.0
50 4 4 0.2 8
0.5 2.0 3.0
8.7 2.6 2.6
100* 7.8
3.4
60 60 93.3
Totals - Task II 14.4 14,116 33.5
•Stored for later transmission at 3.2 bps.
n - 3
The location of each instrument of the expanded science payload complement is shown in Figure II-l. The neutral retarding potential analyzer (NRPA) and ion retarding potential analyzer (IRPA) are located so that the apertures of the instruments are slightly forward of the probe apex to minimize the possibility of heat shield particles reaching the aperture. Of the two Langmuir probes, one is mounted perpendicular to the flight velocity vector and the other is parallel. The sampling port for the neutral mass spectrometer (NMS) is located at the stagnation point of the descent probe, and the analyzer, electronics, and ballast volume tank are located nearby. The accelerometers are located on the longitudinal axis aft of the NMS. The temperature gage, pressure gages, and nephelometer are all mounted on the descent probe outer wall to have access to sensing ports.
The science measurement performance for the descent and preentry instruments are shown in Table II-4 and Figure II-2, respectively. The table shows the NMS sampling duration to be 50 seconds. The high methane cloud in the Uranus warm atmosphere necessitated this NMS interval as compared with 60-second sampling time in the basic study. The warm atmospheres generally present the worst case for performance because of the large pressure and density scale heights. This means that a smaller pressure differential is detected with a given change in altitude, thus with the descent based upon pressure, there are fewer measurements per kilometer. The Saturn warm atmosphere has slightly larger scale heights than for Uranus.
The criteria denoted in the table shows that three general types of measurements are:
1) a specified number of measurements per scale height or per °K or per km, measured throughout descent;
2) a specified number of measurements within each modeled cloud;
3) total number of measurements.
The NMS performance, when defined to provide two measurements in the CH^ cloud, greatly surpasses the other requirements. The ac-celerometer has a criteria of one measurement per kilometer. It makes 0.7 measurements in the worse case atmosphere and 2.4 in the best case atmosphere at the highest altitude. This performance increases up to a maximum of 7.2 measurements.
1,1-4
Tem
pera
ture
Gag
e
Pre
ssur
e G
age
(2)
Ion
Ret
ardi
ng
Pot
entia
l A
naly
zer
Acc
eler
omet
ers
Nep
helo
met
er
Mas
s S
pect
rom
eter
Sy
stem
Neu
tral
R
etar
ding
P
oten
tial
Ana
lyze
r
Lang
mui
r Pr
obes
-
i F
ig.
II-l
Sc
ienc
e In
stru
men
t L
ocat
ions
-
Exp
ande
d Sc
ienc
e P
aylo
ad
Table II-4 Descent Measurement Performance
Instrument and Sampling Time Measurement Criteria
Saturn/Uranus Range of Performance*
Neutral Mass Spectrometer (50 sec)
Minor Constituents Cloud Composition H/He Ratio Isotopic Ratios Molecular Weight
2 per scale height 2 inside each cloud 1.
1_ 4 measurements
4.6/7.7 to 38 2 to 18 in CH4 5 to 12 in NH3
(10 to 16 in H2O 52 measurements
Temperature Gage (4 sec)
Temperature Profile Cloud Structure
1 per °K 2 inside each cloud
1.2/3.1 to 10.9 26 or greater in all clouds
Pressure Gage (4 sec)
Pressure Profile Cloud Structure
2 per kilometer 2 inside each cloud
1.5/4.8 to 14.4 26 or greater in all clouds
Accelerometers (8 sec)
Turbulence 1 per kilometer 0.7/2.4 to 7.2
Nephelometer (3 sec)
Cloud Location Cloud Structure
1 per kilometer 2 inside each cloud
1.3/3.2 to 19 34 or greater in all clouds
'Warm atmospheres give worst-case performance.
Vertical Distance to Make One Measurement, 50
km
uu
90
80
70
60
50
40
30
20
10
0
Langmuir
1 '
'robe
1
r 1
Jranu V = -
S = a ,
60°
LIRPA i
^ S a 7
turn • -3C
kNRPA
)°
0 1.0 2.0 3.0 Sampling Times, sec
Fig. II-2 Preentry Measurement Performance
4.0 5.0
II-6
The pre-entry science instrument sampling times versus vertical distance to make one measurement for Saturn and Uranus are shown in Figure II-2.
The science constraints on the mission analysis is shown in Figure II-3 in terms of entry site selection, approach conditions, and entry attitude.
3. Mission Analysis and Design
The mission analysis and design considerations from launch through the end of mission are shown in Figure II-4. The interplanetary trajectory is selected primarily on the basis of launch payload analysis when considering such constraints as daily launch window, range safety, and parking orbit coast time. The primary goals for encounter are to send the probe to the desired entry site at the proper attitude and to establish an effective communication geometry.
The Saturn/SU 80 mission is described in Figure II-5 in terms of relative probe and spacecraft trajectories, mission summary definition and deflection summary. Just before entry, the spacecraft leads the probe and subsequently the probe catches up. Entry is at -30° and well away from the terminator on the light side.
Figure II-6 summarizes the Uranus/SU 80 mission. At entry, the probe leads the spacecraft and subsequently the spacecraft catches up. The probe is deflected below the spacecraft trace and is carried vertically by the planet rotation across the spacecraft trace. Entry is at -60° and well away from the terminator on the light side.
The mission summary for the three design missions: Saturn direct 79, Saturn/SU 80, and Uranus/SU 80 is shown in Table II-5. The first mission is a Saturn terminal mission, the next a Saturn fly-by, and the last a Uranus terminal mission. The Saturn/SU 80 mission, with a high periapsis radius of 3.8 R , causes a large com-
s munication range of 1.97 x 105 km at entry, which is the worst case of the three design missions. The missions were slightly altered to provide a single spacecraft antenna pointing for both planets. The table shows the cone angle at entry to be similar for all three missions. In addition, the probe aspect angle for Uranus at the end of mission is shown to be reaching an extreme of 43°, which is an important factor of common missions. The entry time dispersion of 4.4 minutes for Saturn and 28.9 minutes for Uranus has a significant impact on power requirements and acquisition time.
II-7
Probe
Trajectory
Entry Site Selection - Saturn - Prefer Posigrade Approach
near Equator to Reduce Relative Velocity
Entry Site Selection - Uranus - High North Latitude Entry
to Avoid Terminator
7
y =
= V
- -90°
-60°
min Mask Angle
Approach Conditions - Satisfy 20° Mask Angle Constraint at Entry
(Lightside Passage through Ionosphere)
Entry Attitude - Obtain 0° Relative Angle of Attack
Fig. II-3 Science Constraints on Mission Analysis
Launch Analysis Launch Period Launch Window Range Safety
Second M/C Correction 13 days before Deflection
Initiate Tracking for Approach
End of Mission
-Entry Entry Conditions Dispersions
-Acquisition Relative Geometry Dispersions
Deflection Maneuver 10-50 days before Encounter
4V Requirements Implementation Radius Mode
Fig. II-4 Mission Design Considerations
II-8
)
Probe Trajectory
E-5 hr
Mission Summary Launch Date Saturn Arrival Flight Time
12/03/80 5/9/84 1258 days (3.4 years)
Spacecraft Periapsis 3.8 R$ Entry Angle -30° Entry Latitude 20°N Deflection Summary Ejection Radius Coast Time 4V Application Angle
30 x 106 km 35.9 days 96 m/sec 73°
-Spacecraft Trajectory
142.7° Spacecraft on Earth Lock
10.6°
Probe Release
E (Entry) End of Mission E+0.75
Periapsis E+2.27
7\ ' \
Spacecraft Deflection (4V =96 m/s)
r ,_ , , End of E (En,r?> Mission
Fig. IIS Saturn (SU'80) Mission
Mission Summary Launch Date Saturn Flyby Uranus Arrival Flight Time
Spacecraft Periapsis Entry Angle Latitude
12/03/80 5/9/84 8/24/88 2826 days (7.7 years) 2.0 Ru
-60° 30° N
Probe Trajectory
Spacecraft Trajectory
End
W Mission
Periapsis E+1.85
^Z7 172.5° Spacecraft on Earth Lock
Probe Release
Spacecraft Deflection dV = 85 m/sec
Return to Earth Lock
Deflection Summary Ejection Radius 10 x 10° m Coast Time 9.5 days 4V 85 m/sec
Spacecraft Trajectory
E (Entry)
End of Mission (E+46 min)
Spacecraft Trace
Fig. US' Uranus (SU'80) Mission Design
II-9
Mission
VHP, km/sec Entry Angle, deg Periapsis, Rs or Ru
Deflection Radius, 106 km Coast Time, days Probe Release Angle, deg Spacecraft AV Angle, deg A V Magnitude, m/sec Lead Angle, deg Lead Time, hr Range at Entry, 105 km PAA* at Entry PAA at End of Mission CA+ at Entry CA at End of Mission Entry Time Dispersion, min Entry Angle Dispersion, deg Angle of Attack, Dispersion, deg
Design Missions Saturn 79
8.2 -30
2.3 30 39.4 14.7 57.3 52 5.3 1.48 1.08 8.2 (42.4)
13.8 122 102
4.4 1.23 2.1
Saturn (SIT 80)
9.11 -30
M 30 35.9 10.6 59.1 96 3.7 2.1 1.97 4.9 (46.7)
20.0 118 112
4.4 1.23 2.1
Uranus (SU'80)
11.87 -60
2.0 10 9.5
21.4 46 85 1.4 1.85 0.87 1.9 61.0)
43 131 105 28.9 6.24 3.0
*PAA = Probe Aspect Angle +CA - Cone Angle
Table II-5 Mission Design Summary
Reorient to Earth Lock, Acquire
Receive
Activate Subsystems Measure & Transmit Preentry Data
Orient for Separation Separate Probe Probe Spin-Up (5 rpm)
M - 0.7 (E+59 secl(5g+15 sec) Deploy Chute Separate Aeroshell/Heat Shield Measure and Store Data; Start Acquisition
/ 'E+l min 14 sec (5g+30 sec) Eject Chute and Uncover Instruments
7E+44 min 59 sec End of Mission (Approx 8 bars)
E+2 min 39 sec (5g+115 sec) Communication Link Established Start Data Transmission
Fig. II-7 Pictorial Sequence of Events (Task I)
II-IO
3 O
3 rt
rt
3 OQ
O
hf
-3
O
CD
I-!
hi
CD
.n
3 H-
hi
CD
3 CD
3
3 3 O CD
i-
l rt
03
M
H
- 03
3
rt
rt
CD
V!
03
Hi
hj
CD
H-
CD
3 C
u
CD
CD
CD
A
. •
O
3 H
3
* 0
3 H
- 3
CO
CL
3
"
O
H <
C/l
03
03
h-1
hi
CD
3 .Q
3
H-
H-
CO
l-i
CD
O
CD
3 r-
» rt
^
3*
CD
• 3
•P-
03
X
3 H
-l|
C
3
rt
CD
CD
CD
3 -
CD
i-l
rt
OQ
3
**
<
CD
• H
CD
Hi
H
O
3*
hi
CD
CD
03
rt
O
3"
rt
CD
H
-<
13
P h
j rt
O
CD
W
CO
C
D
rt
p
3"
i-t
CD
H
CD
< 3
CD
0
4 C
D
CD
^ P
CD
rt
rt
CD
C
D
3 3
CD
rt
i-i
a4 ^
J CD
H
i C
O
O
O
H CD
h-1
P rt
rt
3
* C
D
CD
•<
rt
3*
P
P
H
rt
Hi
h-1
3 O
P
CD
hf
rt
CD
CD
C
D
p
3 3
Q
c 3
CD
H'
3 O
P
P
H
i (-
• rt
>»
rt
to
C
D
vo
p
•
3 3
CL
H
- H
3
CD
3"
C
P
CD
rt
(-(
CD
M
H
i CD
^
H-
0Q
P
3 H
H
tr1 H
P
H
' rt
<
CD
P
>d
TJ
hf
Q
H
3"
CD
CD
p
O
O
P
CD
H
rt
CD
» *3
CD
H
13
P
3 hi
..
C
o
hi
K
CD
CD
H-
"3
H
hj
CD
O
3-
Hi
CD
Hi
O
hi
Hi o
CD
3 CD
T3
H-
hj
3 CD
H
0Q
CL
CD
rt
hi
3*
CD
CD
P M
rt
M
H-
^ 3 CD
h
i
I CD
rt
3 3
" rt
CD
i-i
•<
! CD
P
H
i hf
H
H
. h
i CD
3
H-4
3 CD
<
C
CL
p
H-
M
h{
P
CD
CO
3 3
3 CD
rt
O
3
3"
CD
rt
CD
hi
CD
rt
n
p
Hi
~ H
- "
3
hi
CD
43
[B
3 3
H-
a h
i CD
h
i 3
- C
D
43
H-
3 03
H
-h
i H
i C
D
O
CD
CL
3
" h
j CD
C
u
CD
*d
CD
h
j o
o CD
a*
CD
H
p
CD
rt
M
< -
13
CD
3
" h
j p
01
C
u
CD
CO
3 "
3
I
3*
ON
CD
^
J
> S3
o
i n
3"
3 h
i rt
hi
O
O
H
i
3 •3
rt
O
3
CD
CD
hi
hi
•
3^
O
H-
3 O
h
i 3
* I
CD
CO
CO
•
rt
H> 3
H
CD
3"
CD
Hi
O
hi
CO
CD
13
CD
OQ
" ~
C
CD
O
rt
CD a 3 *i
P
CD
hi
-3
P
P
rt
hj
H-
p
O
rt
3 H
O
1
3 3
3*»
P
hi CD
I VO
03
O
CD
O
H
-CD
r
t
CO
cr
o P
s;
01
03
3*
CD
•3 P
O
CD
o co
h
i H
-p
rt
H
i H
-rt
O
3 [D
w
rt
h
i U
l CD
O
P
•
M
Ul
I rt
a*
H-
H-
3
p
CL
n
CD
4
3 "3
3
M
H-
O
03
V-
H-
3 rt
CD
, H
- 3
O
rt
3 »
h-1
P
CD O
Q
O
CD
rt
CD
p
3
, rt
p
S3
rt
U3
p
rt
3 H
-C
u
Hi
•o n
h
i
U)
O
H
Hi
3*
a*
rt
CD
p p
p 3 3-
3 3*
P
CD
S
^
3 CD
0Q
03
O
p
CD
3
rt
rt
O
rt
CL
P
P
W
H-
rt
03
03
CD
. o
CD
3 p
CD
rt
O
hi
CD a.
CL
P 3*
CD
rt
H-
O
03
H
P
p
OQ
M
CD
03
O
O
H
i O
O
O
N
M
O
h-1
• C
D
K>
O
ft
X
CD
3-
H1
O
P
oo
3 Cb
cr
H-
03
rt
rt
CO
o
• H
CD
> hh
3
rt
3 CD
rt
i-i
H
-
CO
rt
i-i
O
v-
hi CD
Cu
(t
h-1
3"
CD
P n
o
CD
h-1
CD
hi
O 3 CD
rt
CD
hf
Cu
CO
CD
3 hi
03
CD
H-
3 0
Q
CD
3
3 P
rt
M
03
03
P
Hi
It
O
hi
p
13
C1
3 O
l-i
ft
O
3*
X
H-
hi
3 CD
p
P
it
CD
< CD
13
O4
hj
hj
CD
CO
O
H
i C
H
i O
03
H
-CD
> a.
H
i p
rt
rt
CD
P
h
i
CD
3 H
i 3
CD
hi
Hi
^ O
-
H
rt
13
3-
h(
CD
CD
I f
CD
P
3 3
rt
0Q
h
(
O
H
H
P
CD
7?
3 ^
0Q
rt
^
3 -e
-CD
O
J O
rt
O
h
i §r
en
3 P
H
- cr
co
o
co
<
h1-
CD
O
3
3 hi
O
3 P
CD
3 C
L
CU
rt
CL
3 P
O
It
03
p
13
3
* CO
CD
It
h
i O
CD
h
i I
3 -
H hi
CD
C
D
3 fl
rt
H
h
i h
j H
O
^
I &
«•
0
0 CD
»
03
p
3 CD
P
C
L
3*
0 9
CL
CL
<
CD
CD
M
03
?
d
O
CL
hc
j CD
p
>
3 a
-"
* CD
1
3 H
-O
Q
3"
3 H
- p
3
CO
CD
03
03
3"
O
3 03
P
3 CL
CO
3 P
3
3 r
t C
L
H-
M
hf
CD
rt
p
3-
H
CD
1 rt
CD
H-
3 3
Cu
CD
O
p
hh
3
o
3 O
3
* C
u
3 03
p
Hi
rt
CD
hi 3 p
H-
3
CL
13
rt
V
CO
CD
rt
O
P-
CD
CO
CL
CO
H
- C
L
O
p
3 r
t •
P
P
hi
CD
It
h! P 3 03 3 H-
rt
it
CD
CL
rt
O
rt \X
CD
CD
13 P
O
CD
n
hi
p
Hi
It
p hi p n
It CD
H-
CO
CD
(_i.
CD
O
It
CD
C
u
P
3 Cu
3^
rt
3 rt
CD
CL
CD
1
3 h-1
O 1 CD 3 rt >#
hi
CD
P
O
XI 3 H-
CD
H-
rt
H-
O 3 H-
CO
O
O 3 1
3 h-1
CD
rt
CD
3*
CD
Cu
CD
03
o
CD
3 rt
O
3*
C
rt
CD
H-
CD
Cu
CD
1
3 h-1
O
^ CD
Cu
• > rt
M
O
O
CO
CD
n i
03
3 3
* P
CD
H
-M
3
h-1
Hi
hi
O 3 rt
3*
CD
Cu
CD
CO
O
CD
13 P
hi P n
3 3 rt H-
h-1
hi
CD
P
O
3* .
a c r
t CD
H-
CO
CL
CD
3 1
3 I
t
13 hi
O
U"
CD
• h-J
H-
Hi
rt
CD
CD
3 CD
CD
o
o
3 Cu
03
M
P
rt
CD
hi
V r
t
M
O
^ CD
CL
rt
O
CO
CD
13 P
hi P
It
CD
It 3*
CD
3*
CD
P
It
CO
3*
H-
CD
H
CL
3*
P
CD
M
P
hi
OQ
CD
3 CL P
CD
hi
O 1
C
H-
CD
H-
rt
H-
O 3 • > rt
Ul
0Q
•->
*
Cu
CD
CO
o
CD
3 Cu
H
-3 OQ
>*
-' •
3 h-1
3 CD
h-1
Ul
03
CD
O
O 3 CL
03
<«
It
rt
3^
CD
It
hi P 3 CD
3 p-
rt
rt
CD
H
H-
CD
rt 3 hi 3 CD
C
u
O
hh
h
h
P
3 Cu
P
O n
CD
h->
CD
hi o
3 CD
It
CD
hi
Cu
p
rt P
H-
03
03
rt
O
hi
3"
CD
CD
Cu
rt
hi
P
3 03
3 H-
ft It
CD
P 3 CL
It
3*
CD
O
rt
CL
3
"
P
Hi
rt CD
h
i
13 hi
O cr
CD
P o
.a
c H-
CO
H-
rt
H-
O
3
CD
hi
CO
3 u-
CD
^ CO
rt
CD
3 CO
• M 3 0
Q
H-
3 CD
CD
hi
H-
cr 3
v
< O
Q
rt
S4
CD
03
•3
P n
CD
n
hi p
Hi
rt • > rt CD
3 r
t h
i
Cu
P
It
P
H-
CO
O
O
H
M
CD
n
rt
CD
CL
p 3 CL
It
H-
3 CD
hi
H-
3 rt
3^
CD
13 hi
O cr
CD
P n
rt
H« < P
rt
CD
CD
rt
H-
CO
Hi c hi I
t
< CD
M
O
O
H-
rt
3*
V)
CD
H
hf
O
It
P
rt
CD
Cu
rt
O W
P
hi
It
3"
h-1
O
3*
O
CD
13 hi
CD
1 CD
3 it
hi
*<: 92
3 CL
Cu
CD
CO
n
CD
3 rt
O1
P
rt
rt
CD
hi
v*.
!^
• > rt
13 H
CD
1 CD
3 rt
hi
^ \«
It
3 o
rt
O
hi
H-
CO
Hi
H-
hi
CD
CL
Hi
O
hf
ft
3 3 rt
H-
H
3*
CD
H-i
"
3
•d
hi
CD
1 CD
3 It
hi
^ • H
3"
CD
CO
•d
p o
CD
n
hi
jr
p
CD
P
•d
•d
h
i o
13 hi
H-
p rt
CD
Hi
M
V! V
VJ
3f
0Q
CD
> n n
3 rt
hi
O 3
CD
O 3 CD
rt
hf ^ P
3 CL
Hi
rt
H-
CO
rt
3"
CD
3 h{
O
It P
It
CD
Cu
V
rt 3*
CD
Cu
CD
M
r
t P 1
hi
O
CT
CD
CO
13 H
-3 CD
3 -d
it o
Ul
hi
•d 3 P
3 CL
H-
CD
hi
CD
It
C
hi 3 CD
CL
rt
O P
43 3 H
-CD
03
O
CD
3 rt
CO
It
p
It
CD
*d
r
t h
i p
O
h
) O
*0Q
CD
CD
•
It
CD
CL
HI
O
hi P 3 H-
3 *d
P n
rt
H-
3 OQ
rt
hi P
l_l.
CD
D
It
O
H
V!
1* h
i O
I
t P
It
CD
03
82
3 CL
hf CD
h
-1
CD
P
CO
CD
03
rt
3*
CD
rt
H-
O 3 rt o
rt
3"
CD
CD
3 CL
O
Hi
It
H
P
CO
K
M
H-
CO
03
3*
O sj
3 H-
3 ^ 3
* H
-CD
3 H-
CO
CD
H-
O 3 • > rt
CD
CD
13 P
hi p
It H-
O 3 *»
rt
3"
CD
CD
•d
P n
CD
n
hi
P
Hi
rt
0Q
C
h
i CD
M
M 1 ^J
Hi
H
O 3 1
3 H
O^
C
CD
1 it
O 1 03
•d
p n
CD
n
hi
P
Hi
rt
CO
CD
13 P
H
P 1
> •d
H-
O
It o
hi
H-
P
H
03
CD
43 3 CD
3 O
CD
O
Hi
CD
< CD
3 rt
CD
Hi o
H
It a4
CD
CO
13 P
O
CD
O
H
P
hh
r
t
P 3 Cu
•d
h(
O a4
CD
Hi o
hi
w
^ 03
It CD
3 a CD
03
H-
JQ
3 S
3 CL
h-l
3 f
t (D
JQ
h
i P
f
t H-
O 3
P reentry Entry Descent
Data, Bits x 1000
0 10 20
Time from Entry, min
Fig. II-8 Data Profile (Task II)
160
140
120
100 Power, Watts go
60
40
20
0
Separation
10 W 0.67 W-hr
Coast
35.8 days (Saturn)
9.5 days (Uranus)
A -v
P reentry
Late Arrival (LA)
Entry
T 161.3 Transmitter On
(RF Power 41 W H
Nominal Arrival (NA)
Early Arrival (EA)
68.9 W
T Science I On I 40.4 WJ
(Transmitter On . „ „ . (RF Power 7 Wl Amp Off
i J mm r _ j Power O n - '
J I L
P reentry Science &~j~]\
p RF Power r !
—-Transmitter On (RF Power 41 W)
Descent
150.3 W
End of Mission
\
(EA) = 124.16 W-hr (NA) ' 157.47 W-hr (LA) = 190.77 W-hr 36.3 W
Time from Separation, min
•70 -60 -50 -40 -30 -20 -10 0 10 20 30 40
Time from Entry, min
Fig. II-9 Power Profile (Task II)
11-12
5. Electrical and Electronic Subsystems
The telecommunications subsystem definitions for both study tasks are summarized in Table II-6 in terms of frequency, modulation, spacecraft and probe antennas, probe RF power, probe data rate, and probe data storage. The frequency of 0.86 GHz was selected during the basic study and used for the follow-on effort. The optimum frequency could be less than 0.86 GHz, but accommodation of the larger antennas becomes a controlling factor. Also, it was found that hardware is being developed for the frequency selected. The descent modulation was selected to be FSK because of the expected turbulence. Using the Task II pre-entry data rate of 235 bits/sec with FSK modulation required 120 watts of RF power compared with 41 watts for PSK modulation. Therefore, for Task II, PSK modulation was used for pre-entry and FSK for descent.
Two levels of RF transmitter power was required for each task: Task I descent used 25 watts and Task II pre-entry used 41 watts. These same power levels for pre-entry acquisition caused transmitter overheating due to the long RF power-on time required by the 29-minute entry arrival uncertainty for Uranus. Therefore, a pre-entry RF power of 7 watts was established as the minimum requirement. The high data rate for Task II resulted in a data storage of approximately 60 x 103 bits.
The power and pyrotechnic subsystems are essentially unchanged from the basic study. However, the definitions are summarized in Table II-7. As was seen in the power profile in Figure II-9, the separation power requirements are very small and the entry and descent batteries are determined by the RF power requirements and the entry arrival uncertainty for Uranus. The pre-entry and descent watt-hour requirements of Figure II-9 are lower than those shown on the table because the profile represents the power required at the equipment and does not reflect losses in the bus and battery margins. The separation battery was selected on the basis of availability (EP GAP 4384) rather than system optimization. The pyrotechnic subsystem uses a capacitor discharge method, dual bridgewires, and latching relays for arming and safing.
11-13
Table II-6 Telecommunication Summary
Parameter
Frequency
Modulation
Task 1 Task I I , Preentry Task I I , Descent
Spacecraft Antenna
Probe Antenna
RF Transmitter Power, Watts Task 1 Task II
Data Rate, bps Task 1 Task II
Data Storage, bits Task 1 Task II
Value
0.86 GHz
PCM/FSK + Tone PCM/PSK/PM PCM/FSK + Tone
Parabolic Dish, 1.3 18.3 dB Gain, 20°
m Diameter Beamwidth
Axial Pattern, Turnstile/Cone 6.5 dB Gain, 100°
7/25 7/41
1/28 235/50.5
11,600 60, 200
"able II-7 Power and Pyrotechnics
Beamwidth
Summary
Note
Minimizes Link Losses
Low Bit Rate High Bit Rate Low Bit Rate
^
Saturn (SU'80), End of Mission Critical
P reentry/Descent P reentry/Descent
Entry Blackout Preentry and Entry Blackout
Power Subsystem Configuration
Post Separation Battery Task I
Task I I
Entry/Descent Battery Task I
Task I I
- Remote Activated Ag-Zn Prime Power Source'
- Power Conditioning and Regulation at User Subsystem
- 0.25 kg, 94 cm3
- 0.25 kg, 94 cm3
- 160 Watt-hours at 28 Volts, 2.4 kg, 1196 cm3
- 227 Watt-hours at 28 Volts, 3.3 kg, 1639 cm3
Pyrotechnics Subsystem Configuration - Capacitor Bank Discharge Pyrotechnic Initiation
- Latching Relay Safe/Arm
- Dual Bridgewire Initiation from Separate Isolated Power Conditioning
Task I - 15 Events, 2.5 kg, 3146 cm3
Task I I - 16 Events, 2.54 kg, 3195 cm3
11-14
Mechanical and Structural Subsystems
The probe, as separated from the spacecraft, along with the three major elements, entry aeroshell and heat shield assembly, descent probe and parachutes, is shown in Figure 11-10.
A general entry probe configuration, Figure 11-11 depicts a cutaway showing the internal parts of the probe and the afterbody segment location. The entry probe configuration comparisons of the SAG exploratory payload (Task I) and the expanded science pay-load (Task II) are shown in Figure 11-12. Task I configuration is 86.8 cm in diameter and weighs 89 kg compared with 92.1 cm and 103 kg for the Task II configuration. The basic difference between the two configurations is the pre-entry science and descent nephelometer which were added to the larger probe.
The descent probe configuration comparisons of the SAG exploratory payload (Task I) and the expanded science payload (Task II) are shown in Figure 11-13. Task I configuration is 47.0 cm in diameter and weighs 40 kg compared with 48.9 cm and 50 kg for Task II configuration. The basic difference between the two configurations is the added nephelometer for Task II.
The structural and mechanical definitions for the two study task configurations are summarized in Table II-8. The design deceleration loads are based upon -65° entry angle in the Uranus cool atmosphere, which results in a deceleration of 865 g compared with 837 g and -60° for the design mission. Therefore, the definition considers a 5° entry angle uncertainty.
The subsystem weight buildup and dropoff for the two study task configurations is shown in Table II-9. These weights include a 15% margin for uncertainties.
Probe to Spacecraft Integration
The common Saturn/Uranus probe is shown mounted on a Mariner-type spacecraft in Figure 11-14. The interface is briefly described in terms of mission cruise, pre-separation checkout, separation of probe and spacecraft, and post-separation.
Probe Comparison
An overall comparison of the two probe configurations is summarized in Table 11-10.
11-15
Separation and Descent Parachute
Descent Probe
Entry Aeroshell/ Heat Shield Assembly
Entry Probe
Fig. 11-10 Probe Configuration Description
•P reentry Antenna Descent Probe
Afterbody Ablator (ESA 3560)
Triadic Afterbody Segment (Opens After Entry)
Pyrotechnic Actuator for Afterbody Cover
Preentry Equipment
Aeroshell Structure
ATJ Graphite Heat Shield
Aeroshell Payload Support Ring
Carbonaceous Felt Insulator
Fig. 11-11 General Entry Probe Configuration
11-16
63.7 cm (25.1 in
-86.8 cm Diameter-(34.2 in.1
-92.1 cm Diameter-(36.2 in.
re-entry Antenna
NRPA Sensor
73.7 cm --? (29.0 in.)
MRPA Sensor Langmuir Probes
Weight = 103.2 kg (227.4 Ibm) Weight = 89 kg (196.1 Ibm)
a) Task I Entry Probe b) Task 11 Entry Probe
Fig. 11-12 Entry Probe Configuration Comparison
Temperature Gage
47.0 cm Diameter (18.5 in.)
Accelerometer
46.2 cm (18.2 in.)
-Pressure Gage I
Weight = 40.5 kg (89.1 Ibm)
Mass Spectrometer
Nephelometer on Far Side (Not Shown)-
48.9 cm Diameter (19.25 in.)
47.5 cm (18.70 in.)
Weight = 50.3 kg (110.8 Ibm)
a) Task I Descent Probe b) Task I I Descent Probe
Fig. 11-13 Descent Probe Configuration Comparison
11-17
Entry Probe Weight, kg Entry Probe Diameter, cm Descent Probe Weight, kg Descent Probe Diameter, cm Separation Parachute Diameter, cm Descent Parachute Diameter, cm Probe Entry Deceleration Load,
65° Uranus Entry Angle Forebody Heat Shield Mass Fraction,
35° Saturn Entry Forebody Heat Shield Material Afterbody Heat Shield Material Spin Stabilized, rpm
S , y- r\ * s .
SAG Exploratory Expanded Science Payload (Task 1)
89.0 86.8 40.5 47.0
232 73
865 g
0.21
Graphite (ATJ)
Payload (Task II)
103.2 92.1 50.3 50.1
250 88
865 g
0.21
Graphite (ATJ) ESA 3560 (13.7 kg/mz) ESA 3560 (13.7 k 5.0 5.0
Descent Probe
kg/m2)
Entry Probe
Table II-8 Structural/Mechanical Summary
Science Power and Power Conditioning Cabling Data Handling Attitude Control System Communications Pyrotechnic Subsystem Structures and Heat Shield Mechanisms Thermal Propulsion Margin (15% of Above Total)
Probe Total Weight Probe Entry Weight Post Entry Weight Descent Weight
Table II-9 Weight Comparisons
SAG Exploratory Payload (Task 1)
8.08 kg 4.70 3.86 2.43 2.84 3.86 6.38
34.00 6.03 5.22 0.01
11.61
89.02 89.01 78.48 40.50
Expanded Science Payload (Task II)
14.44 kg 5.79 4.50 2.78 2.84 3.95 6.44
37.12 6.67 5.22 0.01
13.46
103.22 103.21 91.32 50.32
11-18
Spacecraft/Probe Interface
Cruise Environmental Cover Monitor (On Demand) ^ 6 ^ High Conduction Mechanical/Thermal Attachment
Preseparation Checkout Power - 8 W Avg; 30 W Peak Checkout Signals Data Monitor - 1400 bits
Separation Probe Pointing - 2° Battery Activation Signal Separation Signal Spring Separation System ( A V = 1 m/s) Probe Self Spin (0.52 rad/sec)
Post Separation Track Probe Receive Data 1 to 50.5 bps Up to 128 Kbits
Fig. 11-14 Probe Integration with Mariner Spacecraft
Mariner -Jupiter/Saturn 77 Spacecraft
Sat urn/Uranus Probe
Table II-IO Probe Comparison Summary
Parameters Task
Mission
Entry Angle, deg
Entry Latitude, deg
Spacecraft AV, m/sec
Max Deceleration, g
Science Weight, kg
RF Power at 0.86 GHz
Max Realtime Data Rate, bps 28 Descent
Probe Data Storage, Kbits 11.6
Ejected Weight, kg 89.0
Heat Shield Diameter, cm 86.8
SU'80
-30(S), -60(U)
20(S), 30(U)
96(S), 85(U)
837 (U Cool Atmosphere)
8.1
25(S)
Task II
SU'80
-30(S), -60(U)
20(S), 30(U)
96(S), 85(U)
837 (U Cool Atmosphere)
14.4
4KS)
235 Preentry, 50.5 Descent
60.2
103.2
92.1
11-19
9. Influence of the Atmospheric Models
Each of the warm, nominal and cool atmosphere models for Saturn and Uranus has a different influence upon the probe design parameters. The controlling models for seven of the most significant parameters are as follows.
a. Science Performance - The Saturn warm atmosphere model is generally the worst case for science performance because of the large pressure and density scale heights.
b. Pressure Height of the First Measurement - The Saturn cool atmosphere places the first descent measurement at the highest pressure height of 122 mb due to the highest pressure at parachute deployment. The lowest pressure height for the first measurement is 51 mb in the Saturn warm atmosphere.
c. Depth of Descent - Using a descent ballistic coefficient of 110 kg/m2 and a descent time (time from parachute deployment to the end of the mission) of 44 minutes, the deepest pressure descent of 19.2 bars is reached in the Uranus cool atmosphere compared with 3.1 bars in the Uranus warm atmosphere, as shown in Figure 11-15.
0.01
0.1
1.0
Pressure, bars
10
100 Warm Nominal Cool Warm Nominal Cool
Legend:
Pressure Depth at End of 44.0 min Descent Communication Link Geometry Limit, 0 dB Margin Communication Atmospheric Attenuation Limit, 0 dB Margin Limit for Passive Thermal Control System Thermal Control Limit for Design Mission, Nonpassive
Fig. 11-15 Cloud Locations and End-of-Mission Limits for Various Model Atmospheres
11-20
Saturn
[ NH3
° ///
I///
= 1 1 \
U. U l
d. deceleration - The Uranus cool atmospheric model, which is controlling, imposes 837 g on the probe at entry compared with 568 g for Saturn cool, and decreasing to 225 g for Uranus warm.
e. RF Link Analysis - The Saturn cool atmosphere model is controlling due to the higher ammonia abundance at the design depth of descent which attenuates the transmitted signal. Figure 11-15 shows that for all the Uranus atmospheres and for the Saturn warm and nominal atmospheres the communication geometry is controlling.
f. Entry Data Storage - The Saturn warm atmosphere is the worst case. Data is stored from 0.1 g (increasing) until 5 g (decreasing) plus 15 seconds when acquisition is complete and descent transmission is started. For Saturn warm, this time is 172.5 sec compared with 166 sec for Uranus warm and decreasing to 133 sec for Uranus cool.
g. Thermal Control - The Uranus cool atmosphere is the worst case for the thermal control subsystem because of the severity of the cold temperature at Uranus. Figure 11-15 shows that a passive subsystem is adequate for Saturn, but a nonpassive system is required for the Uranus nominal and cool atmospheres.
Compromises Necessary to Achieve Commonality
In designing a common probe for use at Saturn and Uranus and for all three atmosphereic models at each planet, certain compromises were made. Compared with a probe optimized for a single set of constraints, such as the Saturn warm atmosphere, the compromises for this common probe are as shown in Table 11-11.
Probe Subsystem Development Status
Some probe components are presently available while others are not. The development status of the major probe components are shown in Table 11-12.
CONCLUSIONS AND RECOMMENDATIONS
The study showed that a common Saturn/Uranus probe is technically feasible for the 1980s using the Science Advisory Group (SAG) exploratory payload and the expanded science payload along with three atmospheric models, defined in the study monographs, for
11-21
Tab
le
11-1
1 C
ompr
omis
es f
or
Com
mon
alit
y
DESIGN
PARAMETER
1. RF Power
2. Temperature Control
3. Heat Shield
4. Structural Design
5. Descent Start
6. NMS Performance
in Cloud
7. Temperature Pressure Profile
Performance
8. Entry Data
9. Battery Sizing
CONTROLLING
FACTORS
Lead lag commonality
at Saturn and Uranus
Passive: S & UW
Active: UN & UC
Mass Fraction:
S = 0.21; U = 0.13
Best Case: UW = 225 g
Worst Case: UC = 837 g
Best Case: SC = 102 mb
Worst Case: UC = 41 mb
Best.Case: CHi+ cloud in UC
Worst Casei' CHi+ cloud in UW
Best Case: UC
Worst Case: SW
Best Case: SW = 8450 bits
Worst Case: SC = 4400 bits
Shortest Pre-entry-S
Longest Pre-entry-U
S = Saturn
C = Cool atmosphere
N = Nominal atmosphere
U = Uranus
W = Warm atmosphere
COMPROMISE
8-watt increase at
Active design
Design for Saturn
Design for 837 g
Minimum Data in UC
Minimum Data in UW
Minimum Data in SW
Minimum Data in SC
Design for Uranus
Saturn
Table 11-12 Probe Subsystem Development Status
SUBSYSTEM ELEMENTS
Science
IRPA NRPA Langmuir
Probe NMS NMS Inlet Accelero-meters
Pressure Gages
Temperature Gage
Nephelo-meter
STATE OF THE ART*
TECHNOLOGY
X
Electronics Subsystems
RF Transmitter Antennas RF Switch Modulator Data Handling Battery Pyrotechnics
Structural/Mecha
Heat Shield Parachutes Cold Gas System
Devices Thermal
Control Radio Isotope Heaters
DESIGN
X
nical Subsystems
X
HARDWARE
X
X X
X
X
X
X
X X X X X X X
X
X X
X
X
DEVELOPMENT REQUIRED
NONE
X
X
X
X
X
MINOR MAJOR
X X
Venus-Pioneer X
X
X
X
Viking
ISIS-B
X X
Titan Missile X X X X
X X
X X
X
Pioneer Spacecraft
*Technology--limited to concepts and ideas; no usable design. Design—hardware definition in process; no similar hardware available. Hardware—actual or similar hardware available.
II-23
Saturn and Uranus. The study also clearly showed that further research, design, and development activities are required in such areas as new science payloads, better atmospheric model definition, heat shield, development, cloud condensation in the instrument sensing ports, NMS inlet definition, dynamic instability, communications, data handling, and parachute materials.
1. Science
Although the SAG exploratory payload definition was provided at the beginning of the basic study and the expanded science payload was provided at the start of the follow-on study, the scientists are not satisfied with either payload and continued study should ensue to obtain a better science complement.
2. Model Atmospheres
The warm, nominal, and cool atmosphere models for Saturn and Uranus were provided at the start of the basic study. The scientists believe that additional data should be generated to more precisely define these atmospheric models.
3. Heat Shield Development
NASA-ARC has done extensive development in the heat shield area in the past and has established a good reference base for further development and testing.
4. Cloud Condensation at Instrument Ports
As the probe enters the planetary atmosphere, the cloud condensation at the inlet ports for the NMS, temperature and pressure gages and on the nephelometer lense will cause erroneous readings. Further development is required to eliminate or reduce this effect.
5. NMS Inlet Definition
The inlet for the neutral mass spectrometer requires further evaluation to ensure compatibility with the masses of the primary constituents that exist in two different groups: 1 to 4 AMU and 15 to 18 AMU. The leak rates through the sintered plug might be appreciably different for each group and cause measurement distortion.
11-24
Dynamic Instability Due to Preentry Science "Burn-Off"
For Task II, the pre-entry science instruments are mounted external to the heat shield and are allowed to burn off during entry. Dynamic instability may result; this requires further analysis.
Communications
Each probe definition shows two probe antennas of the same design: one for pre-entry the other for descent. An analysis and test might show that the descent antenna with an added radome could be used for pre-entry transmission.
The thermal control analysis showed that the minimum RF transmitted power should be used during pre-entry because the entry arrival uncertainty causes the transmitter to overheat. For this reason, transmitter designs that have a variable RF power output and related dissipation control should be investigated. Reduced battery weight would also be realized.
The RF power of the binary FSK link, is designed to meet E /N
= 8.9 dB with convolutional coding and Viterbi decoding. This constraint is based on comparable improvement of PSK (BER = 5 x 10"5). Analys is or simulation should be performed to verify the validity of this constraint.
Data Handling
A nonvolatile programmable memory would provide an effective approach to fault control. Memory devices and fail-safe techniques should be evaluated to provide effective failure control.
Parachute Materials
Data shows that the parachute material, Dacron, will withstand the Uranus cool temperatures. However, a test of the material is recommended to verify that it will withstand these temperatures under deployment conditions.
11-25
III. SATURN/URANUS PARAMETRIC ANALYSIS
This section includes the system and subsystem analyses and tradeoffs that were conducted in support of the probe system definition of Sections IV (Task I) and V (Task II). The general constraints for this section follow:
Missions:
*Saturn Direct 79.,
*Saturn (SU 80),
*Uranus (SU 80),
Uranus (SU 79).
Science Instruments:
Task I (Science Advisory Group's Exploratory Payload) - Neutral Mass Spectrometer (NMS), Temperature Gage, Pressure Gages and Accelerometers;
Task II - Langmuir Probes, Ion Retarding Potential Analyzer (IRPA), Neutral Particle Retarding Potential Analyzer (NRPA) and Nephe-lometer.
Deflection Mode: Spacecraft
Atmospheric Models: Worst Case
Entry Angles, Latitude and Descent Depth: Compatible with Mission Set and Science Objectives
*Design Missions
III-l
MISSION DESIGN CONSIDERATIONS
Launch and Interplanetary Trajectories
Opportunities for Earth-Saturn transfer occur approximately every 12.4 months when the Earth is in its appropriate orbit position for the launch of a spacecraft on a direct flight to Saturn. These launch opportunities have characteristics that vary through a cycle of 29.5 Earth-years (one revolution of Saturn about the sun). Interplanetary trajectories for the Saturn Direct 79 and SU 80 missions are shown in Figures III-l and III-2. These missions to Saturn are characterized by trip times in excess of 3.5 years and type I transfers.
a. Launoh Energy - Minimum energy Earth-to-Saturn transfers occur in 19 70 and 1984. These missions have favorable characteristics in that they represent node-to-node trajectories approximating Hohmann transfers. The C3 requirement for the 1984 Saturn mission is 110 km2/sec2; for the 1979 and 1980 Saturn missions, the launch energy is in excess of 130 km2/sec2. Even though the 1979 and 1980 missions are not optimal in terms of launch energy and trip time, they are still representative and feasible missions. If for any reason later launches to Saturn are considered, launch energy requirements would be improved.
Figure III-3 presents estimated Shuttle performance data; Figure III-4 Titan/Centaur performance data. The launch vehicle combinations considered for this study are (1) Titan IIIE/Centaur/BII, (2) Titan IIIE/Centaur/TE364-4, (3) Shuttle/Std. Centaur/BII, and (4) Shuttle/45,000 lb/Agena/BII. The seven-segment Titan III is no longer considered a viable launch vehicle option and accordingly it was not considered. Performance data for all vehicles assumes a 185-km parking orbit. The payload in all cases refers to probe, spacecraft, spacecraft modifications, and adapters.
A lightweight Mariner spacecraft design, not including probe, has a launch weight of about 500 kg (1100 lb) whereas an unmodified Pioneer-class spacecraft has a launch weight of about 321 kg (710 lb). Comparisons in launch vehicle performance in terms of injected payload as a function of launch period for the Saturn Direct 79 and S/U 80 mission are shown in Tables III-l and III-2. Table III-3 summarizes the launch energy requirements for the two missions under consideration as well as the Jupiter Uranus 79 mission.
III-2
81 i 80
ORBIT 8 4 OF SATURN
85
82
86 >
ARRIVAL AT SATURN, JUL 19, 1983
87
-APHELION
81
89!
EARTH AT SATURN' ARRIVAL
.80
• #
78
77
76
PERIHELION-
LAUNCH 'FROM EARTH, NOV 23, 1979
0 5 DISTANCE, AU
Fig. III-l 1979 Saturn Interplanetary Trajectory
10
ORBIT OF URANUS
ENCOUNTER SATURN MAY 9, 1984
81
ARRIVAL AT URANUS, OCT 23, 1988
EARTH AT URANUS' ENCOUNTER
LAUNCH FROM EARTH, DEC 3, 1980
I I I I I I J 0 5 10
DISTANCE, AU
Fig. III-2 1980 Saturn/Uranus Interplanetary Trajectory
Table III-l Saturn Direct '79 Mission Launch Capability
LAUNCH PERIOD
NOMINAL
5-DAY
15-DAY
INJECTED WEIGHT, kg (lb)
TITAN HIE/ CENTAUR/ BURNER II
480 (1060)
458 (1010)
417 (920)
TITAN HIE/ CENTAUR/ TE 364-4
580 (1280)
562 (1240)
510 (1125)
SHUTTLE/ CENTAUR/ BURNER II
847 (1870)
806 (1780)
691 (1525)
SHUTTLE/ AGENA/ BURNER II
974 (2150)
956 (2110)
861 (1900)
Table III-2 Saturn Uranus '80 Mission Launch Capability
LAUNCH PERIOD
NOMINAL
5-DAY
15-DAY
INJECTED WEIGHT, kg (lb)
TITAN HIE/ CENTAUR/ BURNER II
476 (1050)
448 (990)
417 (920)
TITAN HIE/ CENTAUR/ TE 364-4
571 (1260)
552 (1220)
510 (1125)
SHUTTLE/ CENTAUR/ BURNER II
827 (1825)
788 (1740)
691 (1525)
SHUTTLE/ AGENA/ BURNER II
956 (2110)
938 (2070)
861 (1900)
Table III-2 Mission Launch Energy Requirements
LAUNCH ENERGY, C
NOMINAL
5-DAY
10-DAY
15-DAY
3, km2/sec2 S 79
132
133.9
135.8
140
SU 80
133
134.5
136
140
JU 79
102
105
107
110
III-4
LAUNCH ENERGY, C3, km
2/sec2
160
140
120k
100
80
SHUTTLE/ 45,000-lb AGENA/BURNER II
SHUTTLE/ STANDARD CENTAUR/ BURNER II
J i 1000 2000 3000 INJECTED WEIGHT, lb
J I I I I I 200 600 1000
INJECTED WEIGHT, kg
Fig. Ill-3 Estimated Shuttle Performanae Data
1400
LAUNCH ENERGY, C.,, km/sec2
160
140
120
100
-TITAN IIIE/CENTAUR/TE 364-4
TITAN IIIE/CENTAUR/ BURNER II I
750 1000 1250 INJECTED WEIGHT, lb
1500
300 400 500 600
INJECTED WEIGHT, kg
Fig. Ill-4 Titan III/Centaur Performanae Data
Figure I I I - 5 provides the C3 contours for launch years 19 79 and 1980. The two re fe rence missions analyzed i n t h i s volume are i n d i c a t e d by the dots 0 on the C3 con tou r s . Since the 1979 mission has no pos t Saturn o b j e c t i v e s , the end condi t ions in terms of p e r i a p s i s r ad ius are open and the re fe rence mission was t a r g e t e d to R = 2.3 R . The S/U 80 Co contour has the Saturn
p s 3
f ly-by r ad ius i n d i c a t e d . For the S/U miss ion , a t r a d e e x i s t s between launch energy, which a f f e c t s launch v e h i c l e i n j ec t ed weight c a p a b i l i t y , and p e r i a p s i s r a d i u s , which a f f e c t s space c r a f t - t o - p r o b e communication space l o s s ; a l a rge p e r i a p s i s r ad ius impl ies a l a rge s p a c e c r a f t - t o - p r o b e range a t e n t r y . The r e f e r ence Saturn/Uranus mission /R = 3.8 R \ was s e l e c t e d as a com-
l P s ) promise between these two f a c t o r s . By r e f e r r i n g to Figure I I I - 5 i t i s seen t h a t launch energ ies for 1979 and 1980 missions to Saturn are in the range of 130 to 140 k m 2 / s e c 2 . For comparison purposes , i t i s r e c a l l e d t h a t the r equ i r ed launch energy for the JU 79 mission was of the order of 105 k m 2 / s e c 2 . Hence, in terms of r equ i r ed launch energy, t r a j e c t o r i e s t h a t f ly by J u p i t e r before encounter ing Uranus are p r e f e r r ed to t r a j e c t o r i e s t h a t f ly by Saturn f i r s t . The primary reason for inc reased launch energy for Sa turn launches as opposed to J u p i t e r launches i s simply the g r e a t e r d i s t a n c e , 10 au versus 5 au.
b. Launch Constraints - The launch d a t e / a r r i v a l d a t e (LD/AD) s e l e c t i o n must cons ider o the r requirements i n add i t i on to launch energy. Cer ta in launch parameters cons t r a in the choice of mis s ion d a t e s . The d e c l i n a t i o n of the launch asymptote, DLA, i s l i m i t e d in magnitude to l e s s than 36° as a r e s u l t of range d e f i n i t i o n and s a f e t y c o n s i d e r a t i o n s . This DLA c o n s t r a i n t i s most r e s t r i c t i v e for the 1979 Saturn launch, e l im ina t ing app rox i mately one q u a r t e r of the a v a i l a b l e p e r i o d ; by 1980 the DLA cons t r a i n t i s of no s i g n i f i c a n c e .
The nominal range of launch azimuths i s l im i t ed between 90° and 115°, and t h i s range determines the da i l y launch window. For both miss ions cons idered , the launch window was i n excess of an hour .
Another c o n s t r a i n t t h a t i s f requent ly imposed involves a f u r t h e r r e s t r i c t i o n on DLA. The accuracy of the t r a c k i n g process i s s e r i o u s l y degraded for DLAs in the v i c i n i t y of z e r o . Accordi n g l y , the fol lowing c o n s t r a i n t i s considered in the launch a n a l y s i s .
I I I - 6
|DLA| > 2°
This constraint is checked to ensure that it is not violated.
c. Arrival Constraints - The most critical arrival constraint is the observability of the spacecraft at encounter by Earth. If the Sun is between Earth and Saturn at encounter, neither tracking nor critical communication tasks can be performed. Therefore, the SEV angle, defined as the angle between the Sun, Earth and spacecraft, at encounter must be bounded away from zero. The constraint plotted on Figure III-5 is |SEV| < 15° and occurs during a noncritical spread of Saturn arrival times. A second constraint, imposed to avoid degradation of the DSN tracking process is 6 = 0 , where 6 is the geocentric declination of
s s the spacecraft at Saturn encounter. This constraint for the missions under consideration is never violated.
Another key parameter defining the arrival geometry is the hyperbolic excess velocity V at the planet. V is no more than
rir Hr
the velocity of the spacecraft relative to the planet; it varies between 7 and 11 km/sec at Saturn and 10 and 14 km/sec at Uranus. The entry velocity at either planet is not greatly affected by the magnitude of V as is shown in Figure I1I-6. The magnitude of V influences probe coast time most significantly. Of more
Hi: importance than magnitude is the direction of the V vector.
HP Two angles, ZAP and ZAE, defined and presented in Figure III-7, are used to determine planetary lighting conditions of the approach trajectory and efficacy of approach orbit determination, respectively. Values of ZAE near 90° indicates that the velocity vector is contained in a plane perpendicular to the radius vector and hence is not conveniently determined by Doppler tracking. Approach Orbit Determination and Dispersions
No new navigation analysis was performed for this phase of the study, this section includes work previously performed but not previously reported. Using knowledge and control covariances previously obtained, dispersions in entry and communication parameters were generated for trajectory parameters corresponding to a Saturn and Uranus approach from a 1980 mission. In addition, a parametric study was performed in which approach orbit determination for a range of deflection radii was investigated. The results of this study are summarized in Figure III-8; error assumptions are presented in Table III-4.
III-7
Fig
. II
I-5
Satu
rn
1979
and
Sat
urn/
Ura
nus
1980
Lau
nch
Ene
rgy
Con
tour
s
37 2 r-
y = 3.79 x 107 km3/sec2
25 4
R = 60,291 km
ENTRY
VELO
CITY
, Vr
S km/sec
37.0 -
36.8-
36.6 —
36.4-
36.2
36.0-
25.0 —
24.6 —
ENTRY
VELO
CITY
, M
r, km/sec
24.2-
23.8 —
23.4 —
23.0 10
VHp,
km/sec
a) SATURN ENTRY
VELOCITY
VS V HP
F
ig.
III-
6 Sa
turn
an
d U
ranu
s E
ntry
V
eloc
ity
Var
iati
ons
wit
h V
u = 5.788 x 106 km3/sec2
R = 27,000 km
12
14
VHp,
km/sec
16
b)
URANUS ENTRY
VELOCITY
VS V HP
HP
III-8
b/lU
5/10
4/9 1984
3/8
2/8
1/9
12/10
11/10
SATURN 1 0 / U
ARRIVAL DATE 9 / H
8/12
7/13 1983
6/13
5/14
4/14
3/13
2/13
1/14
12/15 1982
11/15
—
—
—
~ I -
M
1
\
\ 1980 \ LAUNCH
\
\
\
\
\
\
I \
\
\ 1979 \LAUNCH
\
\
1 1 1
1979 LAUNCH
EARTH
J L 6 10 14
VHp, km/sec
a) SATURN ARRIVAL DATE v s v H P
140 150 ZAE, deg
160
1980 LAUNCH
b) SATURN ARRIVAL DATE VS ZAE
150 160 ZAP, deg
c) SATURN ARRIVAL DATE VS ZAP
Fig. III-7 Vm, ZAE, and ZAP Variations with Arrival Date
III-9
DEFLECTION RADIUS/TRACKING, 106 km
100/S 30/S 20/S 10/S 20/0
T, 103 km T, 103 km
CONTROL UNCERTAINTY, R, 103 km
KNOWLEDGE UNCERTAINTY, R, 103 km
a) NAVIGATION UNCERTAINTIES AT SATURN
T, 103 km
-CONTROL COVARIANCE
-KNOWLEDGE COVARIANCE
STANDARD DSN TRACKING, 15 x 10G km DEFLECTION
b) NAVIGATION UNCERTAINTIES AT URANUS
DEFLECTION RADIUS, 106 km
-15-
1 CONTROL UNCERTAINTY
h2 R, 1 0 3 km
R, 10 3 km
c ) STANDARD PLUS OPTICAL TRACKING
Fig. Ill-8 Navigation Uncertainties
111-10
Table III-4 Navigation Uncertainties at Saturn (la)
EPHEMERIS UNCERTAINTIES, la
SATURN URANUS
IN-ORBIT, km 750 10,000
RADIAL, km 750 10,000
OUT-OF-PLANE, km 250 2,000
EQUIVALENT STATION LOCATION ERRORS, la
DISTANCE OFF SPIN AXIS, m 1.5 LONGITUDINAL DISTANCE, m 3.0
Z-HEIGHT, m 2.0
LONGITUDINAL CORRELATION 0.97
MEASUREMENT NOISE, la
DOPPLER:
RANGE: OPTICAL:
0.3 mm/sec (l-minute COUNT TIME)
150 m 10 arc-sec
A final midcourse maneuver was scheduled 13 days before the deflection maneuver. A 40-day tracking arc was used before initiating the midcourse maneuver. Standard tracking of 10 Doppler measurements/day plus one ranging measurement, and standard plus optical tracking that consisted of three star planet angles and one apparent planet diameter measurement per day, were evaluated for a range of different deflection radii. The S and 0 located between the knowledge and control ellipses in Figure III-8 refer to standard and standard plus optical tracking, respectively. Generally, standard tracking appears to be adequate at Saturn. The large ephemeris uncertainties of Uranus coupled with the large range (approximately 20 au from Earth) result in a requirement to supplement standard DSN tracking with optical measurements. The relatively minor degradation of the effectiveness of the orbit determination process at Saturn with standard tracking and Uranus with optical plus standard tracking as the deflection radius is increased, is demonstrated in Figure III-8.
The spacecraft-to-probe communications link is designed to accommodate realistic dispersion in the communication parameters. The spacecraft antenna beamwidth selection process is based on spacecraft look-direction dispersions. Figure III-9 displays the design mission look dispersions. Even though the knowledge and control covariances at Saturn and Uranus are approximately equal, the look-direction dispersions at Saturn are considerably (factor of two) larger than the dispersions at Uranus. This is explained by the respective geometries where the tail geometry yields smaller look dispersion than a side geometry. The Uranus approach approximates a side geometry, whereas the Saturn approach approximates a tail geometry.
III-ll
a)
CROSS CONE
CROSS CONE
1.0
CONE ANGLE, deg
URANUS LOOK DIRECTION DISPERSIONS
b) SATURN LOOK DIRECTION DISPERSIONS
Fig. II1-9 Saturn and Uranus Look Direction Dispersions
Dispers ions i n the probe en t ry parameters a f f ec t the design and opera t ion of the probe i t s e l f . The sc ience performance of the probe mission i s a f fec ted by d i spe r s ion i n en t ry s i t e , en t ry ang le , and en t ry angle of a t t a c k . The u n c e r t a i n t y i n the time of en t ry i s a c r i t i c a l f ac to r i n sequencing and probe power r e q u i r e ments. Dispers ions :n en t ry angle a f f e c t the s t r u c t u r a l r e q u i r e ments imposed on the probe s ince they impact probe thermal and aerodynamic c o n s i d e r a t i o n s . Because the d i s p e r s i o n s in most of these parameters a re due s o l e l y to nav iga t ion u n c e r t a i n t i e s , Figures 111-10 , I I I - l l and Table I I I - 5 p resen t the en t ry d i s pe r s ions used i n the design of the t h r e e missions under cons ide r a t i o n .
111-12
24
23
22
21
LATITUDE, deg
20
19
18
17
16
S 79
62 63 64 65 66 67 68
LONGITUDE, deg
69 70
Fig. 111-10 Saturn Entry Site Dispersions
40,—
38
36
34
32
30
LATITUDE, deg
U/SU 80
28
26
24
22
20
18 —
16 317 319 321 323 325 327 329 331 333 335
LONGITUDE, deg
Fig. III-ll Uranus Entry Site Dispersions
111-13
Table III-5 Entry and Communication Parameter Dispersions
3d DISPERSIONS
ENTRY ANGLE, deg
ANGLE OF ATTACK, deg
LEAD TIME, minutes
ENTRY TIME, minutes
PROBE ASPECT ANGLE, deg
RANGE, km
DOPPLER VELOCITY, km/ sec
DOPPLER RATE, m/sec2
* S 79
1.1 1.8 5.7 4.4 3.3 4200
1.01
0.18
S/SU 80*
1.2 1.86
8.7
4.4 2.59
5080
0.935
0.11
*DEFLECTION RADIUS = 30 x 106 km.
'DEFLECTION RADIUS = 10 x 106 km.
U/SU 80+
6.5 3.15
1.22'.
28.9
1.17
1240
0.43
0.11
3. Planetary Encounter
The planetary encounter phase of the mission encompasses the deflection maneuver, acquisition, entry, and descent. The primary problems associated with planetary encounter include the design of the communication relay link, the selection of the approach trajectory, and the analysis of the deflection maneuver. The primary emphasis of the section will be to determine the selection of a common approach geometry for all three missions: Saturn 79, Saturn/SU 80, and Uranus/SU 80. Lesser emphasis will be placed on approach trajectory selection and the deflection maneuver. It was the intent of this study to determine a common range of spacecraft-to-probe look directions. Commonality of look directions results in a simplified spacecraft antenna design, eliminating the need for antenna gimbals on the spacecraft.
a. Design of the Relay Link - The key parameters associated with the relay link analysis are illustrated in Figure 111-12. During the pre-entry phase, the probe is assumed to move on a conic trajectory in the attitude required at entry for zero relative angle of attack. During the entry phase, the probe rotates so its axis is radial relative to the center of the planet. During the descent phase, the probe is on a parachute descending along a radius vector as that vector rotates about the center of the planet at the angular rotation rate of the planet. Definitions of the relevant parameters follow.
111-14
NOTE: 4>E IS THE ANGLE THE PROBE ROTATES
THROUGH IN GOING FROM THE PRE-ENTRY ATTITUDE TO THE DESCENT A T TITUDE.
a) PRE-ENTRY PHASE b ) ENTRY PHASE , c ) DESCENT PHASE
Fig. Ill-12 Relay Link Parameters
Probe Aspect Angle (PAA)3 \\i - The angle between the axis of the probe and the p robe - to - spacec ra f t range . The PAA time h i s t o r y has a d i s c o n t i n u i t y a t en t ry corresponding to the ins tan taneous r o t a t i o n of the probe from p r e - e n t r y to descent a t t i t u d e . The PAA would opt imal ly be zero ; p r a c t i c a l l y , t h i s i s not p o s s i b l e . The l a r g e r the va lues of the PAA, the more power i s r equ i r ed for the probe antenna.
Relay or Communication Range (p) - The d i s t ance between the probe and s p a c e c r a f t . Opt imal ly , t h i s i s he ld as smal l as p o s s i b l e .
Lead Angle (X) - The angle between the spacec ra f t r ad ius and the probe rad ius (p ro jec t ed i n t o the spacec ra f t p l a n e , i f necessary) a t e n t r y . I f n e g a t i v e , the probe leads the spacec ra f t a t e n t r y ; i f p o s i t i v e , the spacec ra f t l e a d s .
Lead Time [t \ - The time from en t ry to spacec ra f t p e r i a p s i s p a s
sage . I f t i s p o s i t i v e ( the usua l c a s e ) , en t ry occurs before
the spacec ra f t has passed p e r i a p s i s .
111-15
Cone Angle (CA) Cloak Angle (CLA) - The CA and CLA are here referenced to Earth and Canopus. The CA is the angle included by the Earth-spacecraft-probe alignment; the CLA locates that direction relative to Canopus.
The selection of an effective communication geometry requires an iterative search. It was decided to first minimize the probe-to-spacecraft communication power for each of the missions under consideration. This was accomplished by targeting the spacecraft at the deflection maneuver for zero probe aspect angle at different times after descent. Four targeted trajectories corresponding to zero PAA at 5, 15, 30, and 45 minutes after entry were obtained. The nominal descent time for all three missions is approximately 44 minutes. A preliminary communication link analysis was then performed for all targeted trajectories and the minimum power case was selected as optimum. For the task one science complement, the mission determining the size of the probe transmitter was the Saturn/SU 80 mission and the resultant power was 17 watts. The increased power requirement for the Saturn/SU 80 mission relative to the Saturn/JS 77 mission was primarily a result of spacecraft-tor-probe range at entry, which was 197,000 km. The range for the Saturn 1979 and Uranus/SU 80 missions was 108,000 km and 87,000 km, respectively.
The next phase in design of the relay link was to compromise each mission slightly to yield a common spacecraft-to-probe look geometry. Since the Saturn 1979 mission is a direct flight and there are no constraints on periapsis radius other than to avoid the rings, it is always possible to duplicate the Saturn/SU 80 geometry while simultaneously reducing the range. For this reason, in the subsequent discussions we shall only discuss commonality between the Saturn/SU 80 and Uranus/SU 80 missions.
The optimum Saturn/SU 80 mission had a cone and clock angle at entry of 109° and 256°, respectively, whereas the optimum Uranus/ SU 80 mission resulted in CA =135°.and CLA = 275°. By increasing the lead time at Saturn (and hence increasing range and PAA) and decreasing the lead time at Uranus, the differences between the respective cone angles at entry were reduced to less than 15° with CAOArT1 = 120° and CA.— ..., = 131°. At Uranus, the approach tra-SAT UKAN jectory was targeted to an inclination of 10° between the probe and spacecraft trajectory planes to minimize variations in the
111-16
PAA during the descent phase of the mission. A nonplanar deflection is desirable because the pole vector of Uranus is parallel to the ecliptic plane, resulting in probe motion during descent that is approximately perpendicular to the spacecraft trace.
To reduce the differences in clock angle at Saturn and Uranus, the deflection maneuver at Uranus was retargeted so that the inclination between the spacecraft and probe orbit planes at entry was reduced to zero. The planar deflection maneuver at Uranus had the effect of reducing the clock angle by 13° so that for the final design the difference in clock angle at Saturn and Uranus is less than 10°. This change had little effect on the cone angle.
At this point, it is interesting to point out the price that was paid to achieve Saturn and Uranus commonality. The primary criterion in determining link design effectiveness is transmitter power, and the driving mission which sizes transmitter power is Saturn/SU 80. As stated previously, the optimized Saturn/SU 80 link design resulted in a transmitter power of 17 watts. For the Saturn/SU 80 mission design to achieve commonality, this figure increased to 25 watts. Therefore, an increase of 8 watts was required to accommodate the common Saturn/Uranus link design.
Selection of Approach Trajectory - The selection of Saturn/SU 80 approach trajectory was determined primarily from launch energy considerations. Ideally, the optimum periapsis radius at Saturn to achieve effective rate matching is about 2.5 R . However, the
s increase in launch energy to achieve the appropriate periapsis radius for rate matching was significant enough to preclude use of this trajectory and a compromise was reached. By increasing the periapsis radius to R = 3.8 R , the launch energy was reduced to a realistic value (C3 = 140 km
2/sec2 for a 15-day launch period). The size of the probe transmitter design is determined by the Saturn/SU 80 mission. The approach trajectory selection for the 1979 Saturn direct and the Uranus/SU 80, while critical in the design of the communication link, is not constrained by the relationship between launch energy and periapsis radius. The Saturn 1979 approach was selected to have a periapsis radius of R =2.3
R which is the minimum value that will avoid the Saturn rings, s
The periapsis radius at Uranus was selected to be R = 2.0 R . r p u
111-17
a. Deflection Maneuver - The spacecraft deflection mode was selected for each of the three missions analyzed primarily to reduce the complexity of the probe, since for this deflection scheme, the probe needs neither an attitude control system (ACS) nor a propulsion system. For a spacecraft deflection, the spacecraft trajectory is initially targeted to impact the entry site. The probe is spun up and the spacecraft is precessed so that the.released probe has a zero angle of attack. The spacecraft is then deflected away from the planet to establish communication geometry and the required periapsis radius and Earth lock is reestablished.
Comparison of Deflection t\V Requirements - Deflection AV requirements for a wide range of parametrics have been summarized in Volume II pages IV-31 through IV-34. The primary difference between the missions analyzed and reported in Volumes I, II, and III and those discussed in Volume IV is the deflection mode. With spacecraft deflection, greater emphasis should be placed on reducing deflection AV requirements. Since a ground rule of this study was to perform no new navigation work, previous navigation results were used.
The deflection radius used at Saturn and Uranus was 30 x 106 km and 10 x 105 km, respectively. These values were primarily determined as a result of availability of previous navigation results. A topic for future study would be the determination of the maximum possible deflection radius while still maintaining reasonable dispersions. It is probable that a reduction in spacecraft deflection AV from the present value of approximately 100 m/sec to 15 m/sec can be realized.
Planetary Entry
The critical phases of the probe mission are the pre-entry, entry, and descent phases. The entry phase is initiated when the probe first experiences effects of the sensible atmosphere and terminates with staging of the aeroshell. Typically, the aeroshell is staged at a subsonic velocity above a pressure level of 100 mb. Since entry is initiated at a, particular pressure level and the pressure altitude profile is dependent upon atmosphere model, the altitude of entry is also dependent upon the atmosphere model. In all instances, altitudes are given above the 1-bar pressure level. For reference, the entry radius, pressure, and altitude of entry are displayed in Tables III-6 and III-7 for both Saturn and Uranus; all three (warm, nominal, and cool) atmosphere models are treated.
111-18
Table III-6 Saturn Model Atmosphere Entry Parameter Definition
REFERENCE PARAMETER
RADIUS AT 1.0 atm, km
ALTITUDE AT 1.0 atm, km
PRESSURE AT ENTRY, atm
ENTRY ALTITUDE, km
ENTRY RADIUS, km
SATURN A" WARM
59,735.6
0.0
1 x 10-7
568
6030.6
rMOSPHERE NOMINAL
59,805.9
0.0
1 x 10-7
366
60,171.9
COOL
59,863.4
0.0
1 x 10-7
227
60,090.4
Table III-7 Uranus Model Atmosphere Entry Parameter Definition
REFERENCE PARAMETER
RADIUS AT 1.0 atm, km
ALTITUDE AT 1.0 atm, km
PRESSURE AT ENTRY,.atm
ENTRY ALTITUDE, km
ENTRY RADIUS, km
URANUS ATMOSPHERE •WARM
25,926.2
0.0
1 x 10"7
1073.8
27,000
NOMINAL
26,468
0.0
1 x 10-7
"t
532
27,000
COOL
26,800
0.0
1 x 10-7
200
27,000
Typically, the entry phase of the mission lasts less than two minutes; however, it is during this phase of the mission that the probe experiences severe aerothermodynamic effects. These effects contribute in a significant manner to the sizing of the heat shield and structure. Sufficient ablative material must be available to withstand the heating encountered during entry; the structure must be sized to withstand peak decelerations.
a. Entry Ballistic Coefficient Selection - The primary considerations influencing the selection of the entry ballistic coefficient are staging conditions compatible with technology requirements and science objectives. It is desirable to commence science measurements above the 100 mb pressure level; aerodynamic considerations prefer a subsonic staging velocity. A parametric study was performed to investigate variations in staging conditions with entry angle, ballistic coefficient, and model atmosphere. The results of the study are displayed in Figures 111-13 and 111-14. On each figure, contours of altitude at which the prob's velocity is M = 0.7 for the particular ballistic coefficient indicated. Contours are displayed for the three atmospheric models and ballistic coefficients of 78.5, 157, and 235.5 kg/m2. The altitude for each respective atmosphere that corresponds to the 100 mb pressure level is also indicated. Ideally, the ballistic coefficient should be as large as possible consistent with staging considerations. The worst case atmosphere with regard to entry ballistic coefficient is the cool atmosphere, and Saturn entry with entry angle of y = -30° presents the most severe situation. To achieve desired staging conditions, the entry ballistic coefficient selected must be less than 78.5 kg/m2. For consistency with previous work and also to ensure some degree of safety margin, an entry ballistic coefficient of 110 kg/m2 was selected.
b. Entry Environment - The relationship between entry angle, model atmosphere, peak decelerations, maximum dynamic pressure, and their respective time profiles is shown in Figures 111-15 through 111-18.
111-20
SATURN ENTRY, WARM ATMOSPHERE
ALTITUDE OF MACH = 0.7, km
200
190
180
170
160
150
140
-
B = 78.5 kg/m2
B = 157 **s
• B = 235.5
r ."^
\ \
O B » 1
> *
100 mb
\ -10° -20° -30° ENTRY ANGLE, y, deg
-10° -20° -30° ENTRY ANGLE, y, deg
Fig. Ill-12 Deceleration to Subsonic Speeds at Saturn
90
80
70
60-ALTITUDE OF MACH = 0.7, km
30
URANUS ENTRY, NOMINAL ATMOSPHERE
170
160-
150
-100 mb
140 ALTITUDE OF MACH » 0.7, km
URANUS ENTRY, COOL DENSE ATMOSPHERE
B = 78.5
-50 -60 -70 ENTRY ANGLE, y, deg
100 mb
_1 l_
130
120
110
100
URANUS ENTRY, WARM ATMOSPHERE
B = 78.5 kg/m2
B = 157
-100 mb
-50 -60 -70 ENTRY ANGLE, y, deg
Fig. 111-14 Deceleration to Subsonic Speeds at Uranus
III
B • 7 8 . 5 - 2 3 5 . 5 kg/m2
MAXIMUM DECELERATION, EARTH g
MAXIMUM DECELERATION EARTH g 50o
B • 78.5^235.5 kg/m2
'COOL. DENSE ATMOSPHERE
-10 -20 -30 ENTRY ANGLE, Y. deg
a) MAXIMUM DECELERATION AT SATURN
ENTRY ANGLE, y. deg
b) MAXIMUM DECELERATION AT URANUS
Fig. Ill-15 Maximum Deceleration at Saturn and Uranus vs Entry Angle
MAXIMUM DYNAMIC
PRESSURE,
103 lb/ft2 '
4 4 r B • 235.5 kg/m2
- COOL ATMOSPHERE
NOMINAL ATMOSPHERE
» 235.5 kg/m
MAXIMUM DYNAMIC
PRESSURE,
103 lb/ft2
157
235.5
157
78.5
- 1 0 - 2 0 - 3 0
ENTRY ANGLE, Y , deg
- 5 0 - 6 0 - 7 0
ENTRY ANGLE, y, deg
a) SATURN ENTRY b) URANUS ENTRY
Fig. 111-16 Maximum Dynamic Pressure at Saturn and Uranus vs Entry Angle
111-22
DECELERATION, EARTH g
WARM ATMOSPHERE
20 30 40 50 TIME FROM ENTRY, sec
Fia. Ill-17 Saturn Deceleration Profile
DECELERATION, EARTH g
COOL, DENSE ATMOSPHERE
NOTE: T = -60°
WARM ATMOSPHERE
TIME FROM ENTRY, sec
Fig. Ill-18 Uranus Deceleration Profile
111-23
Peak decelerations increase with increasing entry angle and are most severe for entries into the cool atmosphere. Results indicate that peak decelerations are relatively insensitive to changes in ballistic coefficient. Maximum dynamic pressure, however, is a direct function of ballistic coefficient. For a fixed ballistic coefficient, the time profiles of deceleration and dynamic pressure have a one-to-one correspondence. For this reason, time profiles of dynamic pressure were not included. Maximum decelerations are sensitive to model atmosphere. Analysis indicates (Figure 111-15) that at Uranus for an entry angle of -60°, the nominal peak deceleration is 350 g; for entry into the cool atmosphere, this figure increases to 830 g. Maximum dynamic pressure follows a similar trend.
SCIENCE INVESTIGATIONS
Saturn/Uranus Atmospheric Models
In Chapter II of Volume II.of this report, the model atmospheres used during the first part of the study were presented. They included the nominal models for both Saturn and Uranus as well as the models for Jupiter and Neptune. In this follow-on study, the cool and warm models for both Saturn and Uranus were used and are thus presented here along with the nominals for reference.
The models used were taken from NASA monographs given in References III-l and III-2. Both of these models have been updated from those used in the first part of the study. Figures 111-19 and 111-20 present the pressure versus temperature profiles for each model atmosphere for Saturn and Uranus. At 10 bars, the temperature range is from 191°K to 424°K at Saturn and from 114°K to 300°K at Uranus.
Figures 111-21 and 111-22 connect pressure to altitude for each of the six models to aid in model description. The modeled clouds and their locations are shown in Figures 111-19 and 111-20 and tabulated in Table III-8. These are methane, ammonia, and water clouds, the last two existing in both solid and liquid phases in different models.
111-24
-NOMINAL
10
10
•6 _
• •+ _
PRESSURE, P, 10"2
bars
10
10
0 _
2 _
50 70 100 200 300 500 700 1000
TEMPERATURE, °K
Fig. Ill-19 Pressure vs Temperature and Cloud Levels for Saturn Model Atmospheres
LOG PRESSURE, P, bars
LEGEND:
1 2 3 4
CH„ ICE CLOUDS
NH3 ICE CLOUDS
H20-NH3 SOLUTION CLOUDS
H20 ICE CLOUDS
50 70 100 200 300 500 700 1000 1500
TEMPERATURE, T, °K
Fig. Ill-20 Pressure vs Temperature for Uranus Model Atmospheres
111-25
ALTITUDE, 0 km
10 ' 10
PRESSURE, P, bars
Fig. Ill-21 Pressure vs Altitude for Saturn Model Atmospheres
ALTITUDE, km 0
-800-
10 10"5 10-3 10-i PRESSURE, bars
10
Fig. 111-22 Pressure vs Altitude for Uranus Model Atmospheres
111-26
CO K O « O 3 O
<0
:§ CO
S 3
CO
CO I
H
s_ -Q „ L
U DC rr> oo oo L
U DC Q
_
BASE
Q ZZ) O _
l O o_ o h-Q rr> o _
l
o o rr> C
D _
J O L
U DC L
U m
o. ATMOS
<-CO
UD
o o oo
r.
to U
D
o o oo
oo o
m
•z.
_l
o ^:
o DC
O rD 1—
<c oo
CM
i—
l
i—I
r~-CM
r~-o
CM
CT>
UD
*d-
cr>
oo SOLUTION
oo o
Cvj m
n:
_u o o c_>
^ NAL
i—i
2: o z
LO C
M
*3-
O
CM
U
D C
M
o
oo <—
i
CM
CX> C
M
T—I
oo o
CM
H
Z
m NAL
i—i
2: o z
^ 2: DC
<c 3
CM
rn
2: Q
C
<c 2
LO C
M
UD
o i—1 «*
m
o _i
o oo
o rZ>
O
< or Z
D
r~-o 00
o o oo 00
rn 2
:
_j
0 0 0
00
<zn
0 CT>
>* O •* rn 0 NAL
1—1
2: 0 z
en U
D
UD
0 CO
«3-
«* C
O •stf-
CM
CM
0O SOLUTION
00 0
rn :z NAL
1—1
2: 0 2
:
CM
rn NAL
1—1
2: 0 •zc
CO 1—
1
0 CM
1—1
O
•*
m
<_>
2: DC
«=c 3
CT> C
M
1—1
O
I-H
0 CO
UD
t-H
r. «d-
0O O
m
^ 2: 0
:
<c 3
CM
m
2: DC
<c 3
a)
CO •U
fi 0)
e 3 M
•u CO fi
M
•H cd C O
•rl 4-1 •rl
-a *3 <J L4
O 4-1
CO O
•!-( 4J CO
•H M
QJ 4-i
a cd L<
cd 43 u 4J
c a) 6 3 k 4-1 CO fi
M OJ
O fi QJ •rl O W
CO 4-1 fi QJ
e 3 M
4-t CO fi
•rl
T3 cd O
rH >•. cd 0
.
>> ^ 0 4-> cd
u 0 rH o< X
W
O <J C/l QJ 43 4-1
MH O CO CJ
•rl 4-> CO
•rl ^1 OJ
4-1 O Cd H cd
43 a OJ .fi H
1 M
cd
^3 O Q) 43 H • M
M
0)
6 3 rH O
> m 0
M
rH M
u cu 4-1
1 CU
•a T3 fi cd
>> u 4-1 fi cu 1 CU
u (X •* CO
4-1 fi QJ
a 3 1-1 4-1 CO fi
•rl
O.
^ cd
rG
a M-l O
O c 0 •rl 4J
a CU CO
fi •rl C CU
> •rl 60
0) H cu &
cd fi O
•H 4J •rl •3 •3 cd
CU CO 0
43 4->
<4H O CO CJ
•rl 4-1 CO
•H L<
cu 4-> O cd
I c 0 0 CU ^3 4-1
4H O c 0
•rl 4J U O O
,
C O 1
& O r^
rH O >4-l
CU J2 4-1
c •rl
T3 0) 13 3 rH O C •rl CU U CU
& 4-1 cd
J2 4-1 n
4-1 C 0) CJ CO
•\ >-.
rH rH cd
O •rl m
•rl a CU
a. C/l • CO
X! CL
cd M
60 cd u cd 0
.
60
c •rl & 0 rH rH 0 M
-l 0J ,fi 4J
C •H
13 CU T3 3 rH •cj
C •H 0)
n cd
4J 0 cd L(
4-1
* CO
u CU N >>
rH cd fi cd
i-H cd •rl 4J fi a) 4-1 0 a.
60 fi
•rl *3 M
cd
4J <]) U
T~i cd !-i 4-J
3 cu c
*a fi cd
fi 0 •rl 0)
U cd
CD CO CU
X!
4J
• u aj 4-1 cu e 0
<-t CU ,fi a. cu fi cd
T3 fi cd •* CO 0)
rO 0 u 0,
L4 •rl 3 e 60 fi cd
HJ
cd I 0)
>.
J3 0)
^3 4-1
CO 4J
c 0)
& cu n 3 CO cd
• aj
S cu .fi 4J
O 4J
CO 4->
c cu a 3 u 4J CO
c •H CU CO CU
.fi 4-1
CO cu
4-1 cd
T-i cu u
CD> | M
M
M CU
•' ^
,n cd H
1 cu
a. T3
CU fi 0)
J3 4-1
V4 O
UH
4J
fi •rl
•d QJ
& •H M 0 CO
0) Q
- "3
CU O X CU •\ c 0
•H 4-1 L4 O
« 1-H M
0J
& 3 rH O
> Q
- MH
4J
c cu a CO 0)
T3 CU ^3 H • 6 L4 0
l+H
u 0) Q
.
O 4-1
-0 CU 4J a cu 0
-X QJ
O
CM 1
M
M
M cu
rH .Q cd H c •H c cu >
•H 60
CO cd & «t
K QJ 4J
0 0 r-i
•a cd
cu J3 H • CO CD QJ
c QJ 4-1 QJ
r-{ Cu 6 0 0 !-i
0 14H
>> ^ c 0 aj
u QJ ^3
TJ QJ
-a 3 rH 0 c •H CO •H
T3 fi cd #1
QJ !-i QJ
.fi 4J
rH •H cd 4J
QJ
fi •H
m 0)
•3
T3 fi cd
CO •3 3 O rH O
60 C
•H CO
c cu •3 fi O O
H O m
43 CJ M
cd QJ CO
rH rH •H & L4
QJ 4-1 QJ g 0
rH cu & p
. aj
c rH cd c 0
•rl 4-1
cd
CO cd ** QJ U 3 CO CO QJ
u Cu
•3 c cd
0) !-i
3 4J
cd u QJ 0
.
& 0) 4-1
0 4J
4-1
a 0) p< CO
cu u
43 4J
•H & CO fi 0
•H 4-1 cd 0 0
rH
!-i •rl a) 43 4J
• CO Q) CO cd 60
a) rH 43 •rl CO
c 0) -3 fi 0 0 OJ 43 4-1
<4-l
O ^ 4-1 •rl CO fi QJ
•a CU 43 4J
4H O fi O •H 4-1 O fi 3
MH
CS|
Table III-9 Instruments Related to Measurements
MEASUREMENTS
PRE-ENTRY
ION CONCENTRATION PROFILES
NEUTRAL CONCENTRATION PROFILES
ELECTRON DENSITY h TEMPERATURE
ENTRY
DECELERATION LOADS
DESCENT
H/He RATIO
ISOTOPIC RATIOS
MINOR CONSTITUENTS
ATMOSPHERIC TEMPERATURE
ATMOSPHERIC PRESSURE
CLOUD LOCATION/STRUCTURE
CLOUD COMPOSITION
ATMOSPHERIC TURBULENCE
INSTRUMENT*
LANGMUIR PROBES
D
N
D
IRPA
D
R
R
NRPA
N
D
N
TEMPERATURE GAGE
R
N
R
D
R
R
R
R
PRESSURE GAGE
R
N
R
R
D
R
R
R
ACCELER-OMETERS
D
R
N
N
R
R
R
R
R
NEUTRAL MASS SPECTROMETER
D
D
D
R
R
R
D
N
NEPHELO-METER
N
N
R
N
N
D
R
R
*D = DIRECT MEASUREMENT;
R = RELATED MEASUREMENT;
N = LITTLE OR NO RELATION.
The function of the three pre-entry instruments is to establish profiles of ion, electron, and neutral particle number densities, and electron temperatures as a function of altitude, through the ionosphere and upper atmosphere. This is accomplished before the probe's trajectory is significantly deviated from a free-space conic, and thus is prior to aerodynamic entry. This allows the pre-entry instruments to be mounted on the outside of the heat shield. They will cease to functio" and burn off shortly after entry.
The accelerometer triad measurements of deceleration loads during entry are indicated here to point out the distinction between the pre entry measurements and the descent turbulence measurements.
a. Positive Ion Retarding Potential Analyzer (IRPA) - The function of the IRPA is to establish the positive ion number density concentration profiles through the ionosphere as the probe descends in free molecular flow. The instrument has a range of 1 to 5 amu to include Kx+, H2
+, H3+, He+, and HeH+. It will begin
its operation at about 5000 km and take data for several minutes before passing through the turbopause where data will probably
111-28
cease due to grid wire burn-up. The sampling time for one complete measurement has been set at 2.0 seconds. Although this is not the minimum possible sampling time, it is a reasonable compromise between science data return and maximum allowable bit rate for entry velocities in the range being considered.
The characteristics of the instrument are shown in Table 111-10. The total weight is 3.5 lb, the power requirement is 3 watts, and the nominal bit rate is 60 bps. The configuration of the instrument is shown in Figure 111-23. It is circular with a conical entrance that has a vertex cone angle of 120° and has side vents to allow particles to flow through. The conical entrance reflects particles to the side to limit interference with incoming ions that would normally enter the aperture and be measured. The particle interference problem has been analyzed in Reference III-3 and the results used here.
The instrument is on the nose of the vehicle so that the aperture is forward of the stagnation point. This location minimizes particle reflection from the probe body and other instruments, which could interfere with incoming particles. This design resulted from References III-3 and III-4.
The aperture is circular and approximately 5 cm2 in area; below it is a grid of 1-mil wire grounded to probe surface potential. Ions enter and are retarded by a second set of grids successively varied from -3 to 63 volts in 5.5-volt steps. The ion then passes through a third grid biased at about -20 volts and collected on a plate at about -5 volts. The purpose of the last grid is to suppress emission of secondary electrons from the collector. As voltages are varied, corresponding current values resulting from collection of various positive ions are measured and telemetered back to be coupled with the present voltage values, to establish a current-voltage (I-V) curve from which density, temperature, composition, and potential can be derived.
To obtain ion density and composition, only one current value per mass number is necessary. However, there is danger of missing a step if voltage step size equals voltage per mass number, which is about 12 v/amu; two measurements per step gives better determination of step level and length. Thus, for the 1 to 5-amu range, voltage step size has been set at the previously mentioned 5.5-volt value for the 66-volt sweep. This gives twelve current values to be sent back, and, assuming a 10-bit word at the 2.0-second sampling time, gives 60 bps.
111-29
Table 111-10 IRPA Characteristics
1 WEIGHT - SENSOR = 0.23 kg (0.5 lb), ELECTRONICS = 1.36 kg (3 lb), TOTAL = 1.59 kg (3.5 lb)
2. SENSOR SIZE - 10 x 10 x 4 cm
3. SENSOR VOLUME = 400 cm3, TOTAL PACKAGE = 2200 cm3
4. MAXIMUM POWER REQUIRED - 3 w
5. WARMUP TIMES - 15 TO 30 sec
6. SAMPLING INTERVAL* - 0.5 TO 5 sec
7. DATA BITS PER SAMPLE - 120
8. DATA BIT RATE - 24 TO 240 bps
9. TEMPERATURE LIMITS - -30 TO 1000°C
10. HEAT DISSIPATED - 3 w
11. ONBOARD PROCESSING REQUIRED1 - NOMINALLY, NO; YES FOR ION TEMPERATURES
12. OPERATIONAL ALTITUDES - 5000 km TO TURBOPAUSE ( 530 km)
13. SENSITIVITY - 10 ion/cm3 — 5 ion/cm3 BY 1975
14. EXTERNAL LOCATION - ON SIDE OF PROBE FORWARD SECTION, TO BE IN LINE WITH STAGNATION POINT, OR ON NOSE AHEAD OF STAGNATION POINT
15. ORIENTATION & POINTING - APERTURE NORMAL TO INCOMING FLUX, AND PERPENDICULAR TO FLIGHT VELOCITY VECTOR
16. OTHER REQUIREMENTS - CONICAL ENTRANCE CONE AND VENTED SIDE PLATES
TWELVE SAMPLE POINTS REQUIRED, BASED ON A 5.5-v STEP SIZE (1- TO 5-amu RANGE).
FOR A MISSION THAT INCLUDES AN ONBOARD ION TEMPERATURE PROCESSOR, ADD 0.5 lb TO ELECTRONICS WEIGHT. THE DATA BIT RATE WILL NOT BE LARGER THAN THE NOMINAL VALUE GIVEN'HERE.
If ion temperature is to be determined, the I-V curve must be detailed enough to accurately define the sharp drop corresponding to a given ion mass number. For ion temperatures expected, the number of data points needed to define the curve will be excessive for the telemetry system. Therefore, onboard data processing will be necessary. The electronics will be more complicated than those necessary to determine ion composition and density alone, and the corresponding weight increase has been estimated to be a half pound. However, overall bit rate could be slightly less because the output to be telemetered is now processed, and therefore condensed.
This instrument is state-of-the-art flight hardware, and except perhaps for some development for the onboard temperature processor, will cause no major development problems.
111-30
INLET APERTURE (5 cm2
VENT
RPA ELEMENT
GRID 1
GRID 2
GRID 3
COLLECTOR'
GRID 1
GRID 2 (3 GRIDS)
GRID 3
COLLECTOR
MOUNTING BRACKET ONTO PROBE
POTENTIAL RELATIVE TO PROBE GROUND
0 v
VARIABLE RETARDING VOLTAGE (-3 TO 63 v IN 5.5-v STEPS)
-20 v (TO EXCLUDE ELECTRON COLLECTION & SUPPRESS EMISSION OF SECONDARY ELECTRONS FROM COLLECTOR)
-5 v
Fig. Ill-23 Ion Retarding Potential Analyzer
b. Neutral-Particle Retarding Potential Analyzer (NRPA) - The NRPA establishes neutral-particle number density concentration profiles through the upper atmosphere as the probe flow field goes from free molecular into the transitional region. The instrument range is 1 to 20 amu, looking primarily for H, H2, and He, but wide enough to detect compounds like CH^, NH3, and H20. The NRPA will begin operation nominally at about 5000 km altitude and take data until communications blackout or until the grid wires burn up. The sampling time for one complete measurement has been set at 3.0 seconds. Although this is not the minimum possible sampling time, it is a reasonable compromise between data return and maximum bit rate allowable.
Instrument characteristics are shown in Table III-ll. Power required for the NRPA is greater than for the IRPA, primarily because an ionizing beam is required. The bit rate is larger because of the extended sweep range.
111-31
Table III-ll NEPA Characteristics
1. WEIGHT - SENSOR = 0.45 kq (1 lb), ELECTRONICS = 1.82 kg (4 lb), TOTAL = 2.27 kg (5 lb)
2. SIZE - SENSOR = 10 x 10 x 4 cm, ELECTRONICS + SENSOR = 25 x 10 x 10 cm
3. VOLUME - SENSOR = 400 cm3, TOTAL PACKAGE = 2500 cm3
4. POWER REQUIRED - 5 w
5. WARMUP TIME - 30 sec
6. SAMPLING INTERVAL* - 1 TO 5 sec
7. DATA BITS PER SAMPLE - 280
8. DATA BIT RATE - 56 TO 280 bps
9. TEMPERATURE LIMITS - -30 TO 1000°C
10. HEAT DISSIPATED - 5 w
11. ONBOARD PROCESSING REQUIRED1" - NOMINALLY, NO; YES FOR NEUTRAL-PARTICLE TEMPERATURES
12. OPERATIONAL ALTITUDES - 5000 km TO TURBOPAUSE ( 530 km)
13. SENSITIVITY - 105 particles/cm3 — 1 0 4 particles/cm3 BY 1975
14. EXTERNAL LOCATION - SAME AS IRPA
15. PROTECTION REQUIRED - FOR PROTECTION AGAINST CONTAMINANTS, ENTRANCE & EXIT VENTS ARE COVERED AND VACUUM-SEALED UNTIL AFTER PROBE SEPARATION
16. ORIENTATION & POINTING - APERTURE NORMAL TO INCOMING FLUX, AND PERPENDICULAR TO FLIGHT VELOCITY VECTOR
17. OTHER REQUIREMENTS - CONICAL ENTRANCE CONE AND VENT IN BOTTOM
*TWENTY-EIGHT SAMPLE POINTS, BASED ON A BIAS VOLTAGE STEP SIZE OF 10 v (l- TO 20-amu RANGE).
+FOR A MISSION THAT INCLUDES AN ONBOARD NEUTRAL-PARTICLE TEMPERATURE PROCESSOR. ADD 0.5 lb TO ELECTRONICS WEIGHT. THE DATA BIT RATE WILL NOT BE GREATER THAN THE NOMINAL VALUE GIVEN HERE.
111-32
The configuration of the instrument is shown in Figure 111-24. It is basically the same in shape and location as the IRPA. The conical entrance is designed with a vertex cone angle of only 90° to lessen interference with heavier neutral particles and because the distribution inside the instrument is not as critical. This instrument is vented at the back to allow greater flow because it operates generally in a denser portion of the atmosphere. The position of the NRPA is symmetrical to that of the IRPA on the .. opposite side of the probe centerline with its aperture ahead of the stagnation point.
INLET SLIT
GROUND GRID
ION REPELLER
ELECTRON BEAM GUN
ION COLLECTOR
VENTS
ELECTRON BEAM RECEIVER
SWEEP GRID
SECONDARY ELECTRON SUPPRESSOR GRIDS
ION REPELLER
PROBE MOUNTING BRACKET
RPA ELEMENT
ION REPELLER
BEAM RECEIVER ANODE
SWEEP GRID
SECONDARY ELECTRON SUPPRESSOR GRID
POTENTIAL RELATIVE TO GROUND
300 v
0 v
0 TO 280 v
-300 v
Fig. Ill-24 Neutral Retarding Potential Analyser
111-33
The rectangular aperture is about 5 cm2 with the long axis parallel to the direction of an ionizing beam inside the sensor. Below the inlet is a ground grid of 1-mil wire at zero relative potential to the probe surface. Inside are two grids charged at +300 and -300 volts to prevent all charged particles from getting inside the instrument, allowing only neutrals to enter. The electron beam gun ionizes the neutrals, which are subsequently retarded and collected as in the IRPA.
To correspond to the range of from 1 to 20 amu, the retarding sweep voltage varies from 0 to 280 volts. The steps on the resulting I-V curve will be a little greater than 13 v/amu. Only one current value per mass number is allowed because of the greater amount of data and the limitations on bit rate. However, again there is danger of missing a step (mass number) if the voltage step size equals 13 v/amu. Thus, for this..study, 10 volts was chosen for the step size. One sample or sweep then consists of 28 words. With a 10-bit word and a 3.0-second sampling time, the nominal bit rate for the NRPA is 93.3 bps.
If the neutral-particle temperature is to be determined, the I-V curve must be detailed enough to accurately define the sharp drop corresponding to a given mass number. For particle temperatures expected, the number of data points needed to define the curve will be excessive for the telemetry system. Therefore, onboard data processing will be necessary to determine composition and density alone, and the corresponding weight increase has been estimated to.be 0.5 lb. However, overall bit rate could be slightly less because the output is now processed, and therefore condensed.
This instrument is not state-of-the-art flight hardware, and will require considerable development, although the technology gained from the operational IRPA: is applicable.
111-34
I Lo
Ui
cr
o ft)
H.
CL
13 l-(
O
O
ft cn
cn
H-
0 OQ
%»
rt
0"
ft
ft
rt
ft)
?T
ft 0 rt
O
O cr
rt ft)
H-
0 rt
0"
ft
CO
M
O
|-i
13
ft n rt
H
O
0 rt
ft 3 1
3 fD
1-1 ft)
rt
0 l-i
ft
H-
CD
O cr
rt
ft)
H- 0 ft
CL •
ft o Hi
rt
0"
0)
M 1 < o c n < fD . > Hi
rt
ft
i-i
Hi
0 H
rt
CT
ft
H
O 0 1
< o M
rt
ft)
OQ
ft
H-
cn
cn
s:
m
13 rt Hi
H
O 0 0 ft
OQ
ft
) rt
H
- < ft
rt
O
13 O
CO
H-
rt
H- < ft >«
o 0 H
H
ft 0 rt
H
ft
ft)
CL
H- 0 O
Q
CD
ft)
H
ft
3 0"
H-
O
CT
H-
CD
•d
i-i o o ft
CO
CD
ft
CL o 0 cr
o ft)
i-i
CL rt
O
OQ
H
- < ft
H- o 0 0 C
3 cr
ft
i-i
CL
ft 0 CD
H-
rt
v: • > CD
rt cr
ft
0T
H-
OQ
0*
ft)
0 CL 0 ft
OQ
ft
) rt
H
- < ft
v rt
0J
ft r1 ft)
0 OQ
3 0 H
-H
13 H
O cr
rt
3 ft ft)
cn
C
H
ft
CD
rt
0-
ft
H-
O 0 n c H
i-i rt
0 rt
># ft
) 0 C
L 0"
ft)
cn
< ft)
M
H-
ft)
cr
M
ft < o l->
rt
ft
) O
Q
ft ft)
13
X)
Hi
H- K
ft
CL
CD
O
rt
0J ft)
rt
rt
0"
ft
CO
ft 0 CD
O
H
H-
cn
H"
13
H-
ft
CL • si j?
ft 0 rt cr
H-
CD
< ft)
H
H-
ft)
cr
h-1
ft < o M
rt
ft)
OQ
ft
H-
CD
ft)
i-i ft)
M
M
ft
M
rt
O
rt cr
ft
Hi
l->
H
-O
Q cr
rt < ft
M
O n H-
rt
V! < ft n tor
o 0 cr
o ft)
i-i
CL
rt
O
VI H-
ft
M
CL
rt cr
ft
ft
i-1
rt n rt
i-i o 0 0 0 3 cr
ft
i-i
CL
ft 0 CO
H-
rt
^ • H cr
ft o rt cr
ft
H
TJ I-i
O cr
ft
CD
< ft)
H
H-
ft
CD
s:
H-
rt
0*
rt cr
ft
cn
rt
ft)
0 rt < O
M
rt
ft)
OQ
ft
ft)
13
PL
X)
m
CD
O
ft 0 rt
ft)
M
rt
H-
rt
C
CL
ft . H cr
ft cn
ft 3 ft ft)
CD
0 i-i
ft B
ft 0 rt
cn
ft)
H
ft
13 i-i
O
D
ft pass M
H
-ft
C
L
\* ft)
0 CL 3 rt
0)
CO
c H
ft
CD
rt cr
ft
ft
M rt n rt
i-i o 0 o c I-i
I-i rt
0 rt ft)
cn
H-
rt
X) H
O cr
rt
H-
cn
13 rt « T
3 rt
0 CL
H-
O c M ft)
I-l
rt
O
rt cr
rt
Hi
M
H-
OQ
cr
rr < rt
h-1 o o H-
rt
VI < rt
o rt o H >« cr
o>
cn
ft)
o o 0 1
Hj
I-i o B
CL
H-
CD
n c CO
CD
H-
O 0 CO
C
H-
rt cr
a i-s
. t-1 • w
• CO
l-
( ft) n rt
o Hi
2! > in
> 1 a CO
IT
| n V
o 0 rt
t-1
ft)
0 OQ
B
C
H
-i-
i
ft)
N) 1 0 H- 0 C
rt
rt s ft)
l-( B c X
) rt
H ft)
0 CO
H-
(t 0 rt • . cr
rt
ft
) r
t ft
« Hi
H-
(-•
ft)
B
rt
0 rt
H-
cn
5L
cr
o c rt
NJ .
cr
B) cn
CL
i-i
H-
Hi
rt
rt
Cu S3
« ft)
•<:
ft)
0 -<! cn
13 O n rt r>
i-i ft)
Hi
rt
ft 0
Hi
OQ
i-i o 3 rt cr
rt
CO
Ui
Tl
< ft)
rt
rt
CD
Hi
O
H
rt cr
rt
M o 1 3 H-
0 0 rt
ft
X> m
H
H
-O
C
L
%»
plus
ft)
o rt n H ft)
Hi
rt . hd
o C
rt
i-
i
i-i
ft
ja C
H-
H rt
CL
Hi o I-i rt &)
D cr H
-0 rt
Hi
H-
H
H-
0 OQ
CO
CO
cr
o C
M
CL cr
rt
X) rt
i-i
Hi
O
H £3
rt
H-
cn
O
M
ft ft)
0 • IS
< ft 0 ft) n o M
CL
OQ
ft
) cn
CO
^i CO
rt
rt
3 o o C
H
* C
L n ft)
0 CO
rt
ft)
CL
X)
cu
Hi
rt
rt
H
rt cr
rt
x>
robe
i-s
o cr
M rt
3 V* rt
cr
0 CD
CL
O 0 ft cr
ft
Hi o H
ft
CO
ft
XI ft)
H ft)
rt
H>
O 0 Hi
i-i
O B
CO
X) ft) n rt
o M ft)
Hi
rt
X) O
£,
ft
H
>«
H-
Hi
rt cr
rt
CO
rt
*0 ft
) H
0)
rt
H- o 0
ft>
H
i r
t ft
H
CO
ft
•o
ft)
I-i ft)
rt
H-
O 0 ft)
0 CL ft)
Hi
rt
C cr
o 0 H
CO
cr
rt
Hi o I-i
ft
ft 0 rt
H "<|
\* cr
0 rt
H-
rt
O
O 0 I-1
CL cr
rt
0 CO
ft • H cr
ft
CO
0 OQ
O
Q
ft
CO
rt
rt
CL cr
rt
CD
rt
rt
H- 3 rt
Hi o i-i
X) rt
i-i
Hi
o i-i B
H- 0 O
Q
rt
p- H-
CD
O
X) ft
i-
i ft)
rt H- o 0 H-
CD
0 H
ft 3 ft 0 rt
cn
M« H-
rt
CL
O
ft cn
0 O
rt
0T ft) < rt
rt o K
ft
ft) o a'
ft>
3 cr
H
- rt
0 rt
rt
ft B
X> ft
M
ft)
rt
0 H
ft cr
rt
Hi o i-s
rt
rt
H- 3 ft ft)
0 CL 0 o h-1 o 0 O
Q
rt
i-i
rt
0*
ft)
0
XI O
CO
H-
rt
CO
Hi
i-i
O 3 rt
0 OQ
H
- 0 rt
Hi
H-
i-i
H- 0
ro O
Q
1 3 H- 0 0 rt
ft cn
n o o M
H- 0 O
Q
rt
H- 3 ft
cr
rt
Hi
O
H rt
rt
ft)
??
H-
0 OQ
-SB3UI
CO
. H cr
rt
XI i-i o cr
rt
0 rt
rt
CL
CO
ft)
NJ 1 3 H- 0 0 rt
ft 3 ft)
i-i 3 1 0 •0
Hi
i-i
O 3 M ft)
C
0 O cr
o XI rt
i-i
0)
rt
H-
O 0 CO
V o 0 rt
O
Q
ft)
CD
CD
H- 0 O
Q
CL 0 i-i
H-
0 OQ
D
H
0 H-
cn
ft
V o i-
i
Hi
0 ft
M
Cu rt i
ft)
M
I-1
O o 0 rt ft)
3 H- 0 ft)
0 rt
-d
ft)
H
rt
H-
O
0 rt
ft
CD
cr
rt
Hi
o i-l rt
0 CD
rt » H cr
rt
h-i
13
rt
cn
rt cr
ft)
rt cr
^ <:
rt
cr
rt
rt
0 n o M
M
ft n rt
rt
CL o 0 rt ^r
rt
CO
rt
0 sor
0 I-l
XI o cn
rt
o Hi cr
rt
ft)
rt
H- 0 O
Q
rt cr
rt
XI I-I o cr
rt
H- cn
rt o I-l rt 3 o < rt
ft)
I-I rt
H-
0 o M 0 CL rt
CL
H-
0 rt cr
rt cr
o i-1
M
Q C
rt
C
cr
rt
ft)
0 CL cr
rt
ft)
rt
rt
CL
rt
O
Ln
O
O o
O
Hi
O
H
!-•
O 3 H-
0 1
M
O 0 O
Q cr
o M
M o s:
rt
0 a4
ft >. o • M
Ul
CO
n 3 —\
M
"~—.
M
C^
H- 0 • "-
o D • M
h-'
rt
o rt
I-i
H- o ft)
M cr
rt
ft)
rt
rt
i-i cn
H o 0 O
Q cr
h-1
V! VD
O
o
Hi
H
O B
rt cr
rt S > M
O o ft)
rt
H-
O 0 CO
• H cr
rt
cn
rt
0 cn
O
i-i
H-
CD
ft)
^J • ON
1 n 3 s~\ 3-in.)
»r|
H-
OQ
0 I-
i ft
M
M
I-l 1 N)
Ln . H cr
ft
OQ
0 ft
) H
C
L CD
13 n o rt
H 0 CL rt
Hi
H
O 3 rt cr
rt
0 o CD
rt o Hi
rt cr
ft <J
rt cr
H-cle,
H ^r
rt o o 0 Hi
H-
OQ
0 I-
i ft)
rt
H-
O 0 ft)
0 CL
M
O o ft)
rt
H- o 0 o Hi
rt cr
ro
H- 0 CO
rt
i-i d B
rt
0 rt
CO
ft)
I-l ro
CD
cr
g £ p H-
0
M • LO
O
N
?T
OQ
^^
(jO
h-1
cr
N-
^
ft)
0 CL
H rt
^3 0 H
-H
rt
CD
o 0 rt
CD
rt
rt o Hi
rt
h-1
rt
n rt »-(
o 0 H- n CO
H-
cn
H ro
C
o >Q
s:
B)
rt
rt
CD
rt
O o 1
3 ro
H ft)
rt
rt •
0 H-
H rt
CL
Hi o I-l cr
o rt cr
13 H
O cr
ft cn
ft)
0 CL
H-
rt
Z rt
H-ghs
o cr
0)
n ft>
o rt
rt
H
H
-cn
rt
H
-O
cn
O
Hi
rt cr
ro
H- 0 CO
rt H c 3 rt
0 rt ft)
i-i rt
CO
0"
Q
f±
0 H- 0 H
ft)
cr
t-1
rt
H
M
M 1 M
N) • Only
n o B
13 i-l o 3 H-
CO
ft cr
rt
rt t,
ro
ro
0 CL ft)
rt
ft)
t-i ro
rt
0 l-i 0 ft)
0 CL
M
H- 3 H-
rt
H- 0 O
Q
cr
H-
rt
l-i ft)
rt
ft-
CO
•
%
cr
H- o cr
H-
CD
0 o rt
rt
3 ro
0 rt
v H-
0 n h-1
0 CL
H- 0
cr O
Q ro
H
i ft
) CO
rt
rt
cn
rt
13 o cn
cn
H
- cr
M rt
rt
H- 3 rt
M cr
c:
r
t
H-
cn
ft)
i-l rt
ft)
CO
o 0 ft)
cr
h-1 ro
O 0 cr
o ft)
I-I
CL
13 H
O n ft
cn
cn
H- 0 O
Q
\* cr
&)
cn
cr
rt ro
0 cn
ro
rt
ft)
rt
O • Ln
CO
ft
O
O 0 CL
0 0 rt H-
M 0 ft ft)
H
rt cr
rt
rt 0 H cr
o 13 ft)
0 CD
rt . CO
ft
) 3 13
h-
1
H-
0 OQ
rt
H- 3 ft
Hi
O
i-i o 0 ft n o B
13 h-1
rt
rt
ft B
ft asure-
Hl \-i o z . H cr
rt
XI H o cr
rt s:
H-
H-1
h-1 cr
ro
OQ
H
- 0 o 13 rt
H ft)
rt
H-
O 0 ft)
rt
ft)
CD
rt cr
rt < ro
cr
H-
o M ro
CL rt
CO
n rt
0 CL
CO
rt cr
I-I o 0 O
Q cr
rt cr
ro
H- o 0 o cn
Ui
13
O
O
O S"
p 0 C
L
rt
ft)
?r rt
CL ft)
rt
ft)
cr
ft
I-I rt
H- 0 Hi
H ro
ro
B
o h
-1
rt o 0 lar
CL rt
0 CO
H-
rt
V! O o 0 o rt
0 rt
H ft)
rt
H- o 0 1
3 l-i
O
Hi
H-
M rt
CD
ft)
0 CL ro
M
ro
o rt
H
O 0 rt
rt
B
13 rt n ft)
rt
0 I-l ro
1
3 H o Hi
H-les
tr1
O
ft)
• 0 OQ
3
ti
0 »
H-
S
i-i
«5 2
13
g
l-i
^.
O
iS
cr
ro
•xi
CD
hS
o p)
c
r i-i
K
i ro
—
H
tq
ft
t-J
U2
K)
0 Q
i H
- <r
t-i-i
^
ro
o
CL
S
rt
i~a
O
TO
2
ro
^ cn
»
rt
^
ft)
ft
cr
u-
M
P
H-
4 CO
T
O
cr
\ t)
rt T
O
0*
S
rt
cu
^.
rt
rt-
(-.c
s;
ft o
t)
rt
^
H
O
o !y
0
TO
s—
0 c
t 3 (0
?
M
O
Table 111-12 Langmuir Probe Characteristics
1. WEIGHT - SENSOR = NEGLIGIBLE, ELECTRONICS = 1.36 kg (3 lb)
2. SIZE - SENSOR = 7.6 cm (3 in.) LONG BY 0.16 cm (0.0625 in.) IN DIAMETER ELECTRONICS = 5 x 15 x 15 cm (2 x 6 x 6 in.)
3. VOLUME - 1200 cm3 (SENSOR IS HOLLOW TUBE)
4. POWER REQUIRED - 3 W (2 w BY 1975); HEATER POWER = 5 w FOR 12 minutes BEFORE ENTRY
5. WARMUP TIME - 12 minutes (INCLUDING DECONTAMINATION)
6. NOMINAL SAMPLING INTERVAL* - 0.5 sec
7. DATA BITS PER SAMPLE* - 30
8. NOMINAL DATA BIT RATE - 60 bps
9. SENSOR TEMPERATURE LIMITS - -100 TO 700°C
10. HEAT DISSIPATED - 3 w
11. ONBOARD PROCESSING REQUIRED - YES*
12. OPERATIONAL ALTITUDE - 5000 km TO TURBOPAUSE (- 530 km)
13. SENSITIVITY - 101 TO 107 cm"3, 200 to 20,000°K (e~ temp)
14. ORIENTATION & POINTING ONE SENSOR NORMAL TO AND ONE PARALLEL TO FLIGHT VELOCITY VECTOR
ABOUT 30 CURRENT MEASUREMENTS ARE OBTAINED FROM ONE VOLTAGE SWEEP. AFTER ONBOARD PROCESSING, THE NOMINAL THREE. WORDS OF INFORMATION TELEMETERED BACK ARE THE ELECTRON NUMBER DENSITY, ION NUMBER DENSITY, AND ELECTRON TEMPERATURE. HOWEVER, AS AN OPTION, THE ENTIRE 30 CURRENT MEASUREMENTS COULD BE SENT BACK FROM EVERY 30th VOLTAGE SWEEP BY ADDING A FOURTH WORD OF PRESTORED CURRENT DATA TO BE TELEMETERED EVERY SAMPLE.
GUARD
LANGMUIR PROBE
HEATER - -SENSOR
IRPA
Fig. Ill-25 Langmuir Probe and RPA Location
111-36
The primary reasons for aligning the variable-voltage sensor parallel to the velocity vector are to (1) substantially reduce extraneous voltage induced by the planetary magnetic field as the probe moves through it; (2) keep the distribution of potential along the sensor length even; and (3) prevent translational energy of the probe from interfering with particle thermal energy measurements. This orientation is sensitive to angle of attack, the limit being about ±5°.
The nominal data telemetered consists of three 10-bit words, all processed onboard—electron number density, ion number density, and electron temperature. With the half-second sampling time, the nominal bit rate is 60 bps; however, as an optional verification of the accuracy of the data, one complete set of current readings (30) along a voltage sweep could be sent back every 30th sample by adding a fourth word (current) to the data from a storage unit. This would increase the data rate to 80 bps.
The Langmuir probe and all of its electronic equipment have been flown on Earth-orbital missions, and as such, are all state of the art.
d. Nephelometer - This instrument is the sole addition to the descent payload given in Section C of Chapter III of Volume II. It is added to the payload for the purpose of better defining the location and density of the aerosol cloud formations. It will operate from the initial sequencing of descent instruments, after the parachute is deployed, until the end of the mission. The characteristics used for probe integration design are given in Table 111-13 and the configuration and location of the sensor is shown in Figure 111-26. It is composed of a He/Ne laser light source that emits a fine beam through a window in the descent capsule. The amount of the beam reflected is a function of the density, albedo, and size of the particles. The fraction that is reflected back through the window is diverted by optics into a photomultiplier that yields an intensity reading.
A sampling time of 3.0 seconds has been selected to ensure nephelometer readings of better than one per kilometer below 100 millibars; the instrument is capable of sampling at a faster rate but at the expense of a higher bit rate. The nominal bit rate based upon one 10-bit intensity word every 3.0 seconds is 3.33 bps. The components of this system are all currently available and are being proposed for use on Venus probes. No development will be necessary.
111-37
Table III-13 Nephelometer Characteristics
1 . WEIGHT - SENSOR = 0.45 kg (1 l b ) , ELECTRONICS = 0.68 kg (1 .5 l b )
2. SENSOR SIZE = 5.1 x 6.3 x ^12.7 cm
3. VOLUME - SENSOR = 410 cm3 (^25 i n . 3 ) , ELECTRONICS = 902 cm3
4. ROWER REQUIRED - 3 w
5. WARMUP TIMES - ^5 sec
6. SAMPLING INTERVAL - 1 TO 5 sec (NOMINAL = 3 sec)
7. DATA BITS .PER SAMPLE - 10
8. DATA BIT RATE - 2 TO 10 bps
9. OPERATIONAL TEMPERATURE LIMITS - 240 TO 325°K
10 . HEAT DISSIPATED - 3 W
11 . ONBOARD PROCESSING REQUIRED - NO
12. OPERATIONAL PRESSURES - 100 mb TO DESIGN LIMIT
13. SENSITIVITY - 101* photons/measurement
14. LOCATION & ORIENTATION - ORIENT SENSOR SUCH THAT LOOK DIRECTION IS TO THE SIDE OF THE PROBE AND APPROXIMATELY HORIZONTAL
WINDOWS r " REFLECTORS LENS DETECTOR
Fig. Ill-26 Descent Side-Looking Nephelometer
111-38
Science Mission Analysis and Data Collection
a. Soienoe Event Sequence Profile (Expanded Pay load) - The_se-quence of science events is approximately the same for all Saturn and Uranus model atmospheres, with the time of occurrence for the same events varying slightly. Heater coils inside the Langmuir probes are turned on before the other instruments to decontaminate the sensor. This occurs from 20 to 45 minutes before entry depending upon trajectory uncertainties.
The Langmuir probes begin operation no later than 5000 km altitude and the RPAs 20 seconds (a-600 km) later. The probes begin transmission of real-time data to the spacecraft upon detection of nonzero (or threshold) data. If it is desired to store the data for delayed transmission during descent, both the Langmuir probes and IRPA are also monitored for the beginning of nonzero data and a signal is sent to the data handling system to begin storage, avoiding storage of the large amount of zero data due to trajectory and atmospheric uncertainties.
A g sensor will detect 0.1 g at about 5 seconds after entry, and will signal the pre-entry instruments to be turned off and the entry accelerometers to begin storing entry deceleration data. The pre-entry instruments, mounted on the heat shield, will burn off a few seconds after entry.
The remainder of entry and descent has been described in Section D.l of Chapter III of Volume II. One important difference is that a nephelometer will begin operation simultaneously with the temperature and pressure gages, and take data on cloud densities until the end of mission. A second difference is that consideration of the cool and warm atmospheres, in addition to the nominal, has caused a wider variation in some of the times and pressure of event occurrences. Values for some of these parameters are given in the following subsection.
b. Parameter Variation With Atmosphere - In order to design a common probe for both Saturn and Uranus, the warm, nominal, and cool atmospheres of each planet were evaluated to determine the worst case environment for each design parameter. Table 111-14 lists 18 parameters relating to the entry and descent in each of the six model atmospheres showing the variation and identifying the worst case for those that are important.
111-39
Tab
le
111-
14
Var
tat%
on
of
Par
amet
ers
wtt
h A
tmos
pher
e
PARAMETER
REFERENCE RADIUS (O-km ALTITUDE), km
ENTRY ALTITUDE, km
MAXIMUM DECELERATION, g
MAXIMUM DYNAMIC PRESSURE, 105 N/m2
TIME TO 0.1 g, sec
TIME TO 100 g, sec
TIME TO MAXIMUM g, sec .
ALTITUDE AT MAXIMUM g, km
TIME TO M = 0.7, sec
CHUTE DEPLOYMENT MACH NO.
CHUTE DEPLOYMENT ALTITUDE, km
CHUTE DEPLOYMENT TIME (5 g + 15
sec), sec
CHUTE DEPLOYMENT PRESSURE, mb
ACCELEROMETER MEASUREMENT PERIOD, sec
PRESSURE AT FIRST MEASUREMENT, mb
ALTITUDE AT FIRST MEASUREMENT, km
DESCENT TIME TO 7.3 bars, minute
PRESSURE REACHED IN 44 minutes, bars
SATURN ATMOSPHERE
(y = -30°)
COOL
59,863.4
297
568
5.6
2.0
8.5
11.5
92.7
36.5
0.56
48
44
102
57
122
44.3
24.4
18.2
NOMINAL
59,805.9
536
366
3.6
5.0
17.5
22
151.2
59.5
0.70
92.5
59.5
65
69.5
77
87.8
41.1
8.0
WARM
59,735.6
968.5
227
2.2
12
35
40
259
98.5
0.88
168
84.5
44
87.5
51
162
67.5
4.0
URANUS ATMOSPHERE (Y = -60°)
COOL
26,800
200
837
8.3
0.5
4.5
6.5
65.1
27.5
0.59
35.3
33.5
41
48
56
32.2
23.8
19.2
NOMINAL
26,468
532
358
3.5
4.0
15
19
138
54.5
0.72
78.9
53.5
47
64.5
55
74
44.0
7.3
WARM
25,926.2
1073.8
225
2.2
14
36
40
235
90.0
0.84
145
80
49
81
54
141.3
82.0
3.1
WORST
CASE*
UW
UC
UC
SW
SW
SW
SC
SW
SC
SW
UW
UC
*S = SATURN;
U = URANUS;
C = COOL ATMOSPHERE;
W = HARM ATMOSPHERE.
The first two lines in the table give the radius at zero altitude (1 atmosphere pressure), computed from equations in the NASA monographs referred to previously and the altitude of entry. The probe must traverse more vertical distance in the Uranus warm atmosphere, while the next two lines show that the Uranus cool atmosphere has the highest peak g-load and dynamic pressure. Times and altitudes to various event points are also given, the warm atmospheres requiring significantly longer times.
Parachute deployment occurs at 5 g sensing plus 15 seconds in all six atmospheres, which causes variation in the Mach number at deployment but allows for commonality of design. The worst case Mach number is 0.88 in the Saturn warm atmosphere. This atmosphere also requires the longest time for entry and thus sizes the entry accelerometer data storage capacity requirement. The first descent measurement is desired as high in the atmosphere as possible. The Saturn cool atmosphere is the worst case, the first measurement being taken at 122 mb. In all the other atmospheres, this measurement is made before the probe reaches 80 mb.
The descent depth is 7 bars with additional time to digitize and telemeter the final mass spectrometer sample, assuming that it was taken exactly at 7 bars, which is approximately the cloud base in both nominal models. This adds 50 seconds or about 0.3 bar to the depth, giving 7.3 bars. The times to descent to 7.3 bars, given in the table for each atmosphere, use a ballistic coefficient of 110 kg/m2, and become prohibitive for communication geometry in both warm models, exceeding one hour. However, the clouds are higher in these atmospheres and it is not necessary to descend to the same depth as for the cool and nominal atmospheres. Therefore, the worst case time is that for the Uranus nominal (44 min). Selection of a constant descent time, allows the time-dependent design values, such as bit rate and battery power requirement, to be the same for all atmospheres, as well as more closely aligning the final pressure design point with the modeled clouds. Thus, the descent is designed for 44 minutes and the final pressure reached is given in the last line of the table. The maximum pressure to be designed for is 19.2 bars in the Uranus cool.
111-41
M
H I
CO
11
rt
*-
3*
Oi
ro
o M
?r
B
P
rt
i-h
3
O
i-i
H
ro
to
l-h
p
ro
rt
i-i
C
ro
H
3 3
n
ro
(D
^ 3 cu
u>
o
o r l-h
O
H
C
i-i
P
3 C
CO
II o l-h
O
i-i
IS!
A
IS!
i—•
II ro N
> L
n
O
C
t-,
X
M o en
ro
M
ro
n
rt
H o 3 CO
n
3 CO
l-h
O
i-i
tN!
|v
CXI
e cu
ro
CO cu
a4 o < ro
H-
CO
rt
3- ro
CO
fS
3 ro
Hi o i-i
cr
o
rt
ET
TJ M
P
3 ro
rt
to
V cr
c rt N
t—
» H-
CO
Cu
CD
l-h
H
-3 CD
C
U
a.
H-
i-h
l-h
ro
I-i ro
3 rt
M
v: •
(B
t>
•a
n o X
H- 3 CD
rt
H-
O
3 Hi
O
H
T3 I-i
O
rt
O
3 cu
ro
3 CO
H-
rt
^ <!
ro
I-I CO
c CO
CD
h-1
rt
H-
rt
C a.
ro
• H
3" ro
ro
.a
c P
rt
a- g
ro
3 CO
H
-rt
^ V
C
H-
rt
3" C
3 n ro
H
rt
Co
H-
3 rt
H- ro
CO
V rt
3
" fu
rt
3 CU
^ a4 ro
c CO
ro
Pi co
CO
Co
Hi
H-
H
CO
rt o I-i
Cu
ro
3 rt
H-
O
3 ro
cu
v—'
a- o 0Q
H- <
ro
Cu
3 ro
.a
C
Co
rt
H-
O
3 Hi
O
H
rt
3" ro
13 I-i
o l_l.
ro
o rt
ro
Cu
ro
M ro
o rt
H
O
3 3 C
Hi o H
CO
Co
rt
3 H 3 O
i-i
C
H cu
3 3 CD
• EC
O s ro
<
ro
i-i
V rt
E
T
ro
2; E>
co
>
3 0 3 O
0Q
H Co
X
I ET
CO
/—s
>0 « ro
<
H
-O
C
H-
rt
CD
n —\
pa
ro
Hi
HI
M
M 1 Ui
Co
3 Cu
O
rt
ET
ro
i-i
CO
Nw
*
a*
3 rt
ET
Co
CD
3 O
rt
a4 ro
ro
3 Co
rt
rt ro
3 X) rt
ro
Cu
>#
rt
O
Cu
rt
O ro
X
H-
CO
rt
Co
CO
Co
Hi
C
3 O
rt
H-
O
3 O
Hi
CO
M
rt
H-
rt
3 Cu
ro
• H
E
T
H-
CD
3"
Co
CD
a4 ro
ro
3 Cu
0 3 ro
H
i 0 i-i
rt
3- ro
3 c 3 cr1
ro
n Cu
ro
3 CO
H
-rt
H
- ro
CO
0 Hi <
CO
H
H-
O c CD
CD
xs
CD
O
H- ro
CD
0 Hi
13 Co
i-i
rt
H
-O
M
ro
CD
£ *a
ro
0 rt ro
Cu
0 Co
3 CT
ro
C
u
0 3 ro
i-i
ro
,0 c H- i-i
ro
CD
Cu
rt
ET
ro
0 i-i
ro
rt
H-
0 Cu
I-1 3 0 Cu
ro
M
V
S!
H-
rt
ET
C
3 O ro
i-i
rt
Cu
H-
3 rt
H- ro
CD
CJ
*TJ
ro
Ml
O
I-i ro
Co
0 rt e CO
h-1
cu
ro
i-i
0 Cu
^ 3 £ 3 H
- n ro
3 rt
f-i
V! • P>
CO
H
rt
H-
O
h-1
ro
n
0 3 0 ro
3 rt
H
Co
rt
H-
O
3 CO
H- 3 rt
ET
ro
H
-O
3 O
CD
C
u1
3
ro
rt
ro
H g
H- 3 Co
rt
H-
O
3 O
Hi
3* Q
•3
s;
ro
M
Ef ro
H ro
Co
3 D- C
x>
xs
ro
I-I Co
rt
3 O
CO
x>
3
* ro to
to
rt
C
H
3 P
3 C
u
a H Co
3 C
CO
Co
i-i ro
H-
3 rt ro
3 Cu
ro
C
u
rt
O
Cu
ro
O h3
hS
CO
1 P- s <rf
•
cs; t)
» rt-
R
C5
O
^J
^J
CO
Ci
<rt- ^.
O
3 1 H
ET
ro
H
l>3
H- 3 ro
rt
3- ro
H-
0 3 Co
3 Cu
3 ro
3 rt
i-i
Cu
i->
H ro
1 ro
3 rt
i-i
"<! 3 ro
Co
CO
c H 2 3 ro
3 rt
CO
Cu
H-
ro
SC
fl
Co
C-
i —
I—
' »-
t rt
O
i-i
a
* h
- rt
C
ro
3
ro^'r
o'o
H
irt
xj1
H
>.
I O
E
T
O
i-i
H-
rt
CO
3*
c ro
3 ro
cu
CD
c H ro
Cu
K
H-
rt
Ef
O
C
rt
CD
ro
H
H-
O c CD
X) ro
3 Co
M
rt
•^
rt
O
rt
ET
ro
O
<
ro
H
CO
M
h-1
X> It
O
cr
ro
Cu
ro
CD
H
-CM
3
M
Hi
rt
ET
ro
Cu
rt
3 0 CD
XI ET
ro
H
ro
CD
Cu
H ro
M
H-
??
ro
rt
E
T
ro
0 0 0 H1
3 O
Cu
ro
H
-1
CD
V rt
E
T
ro
CO
ro
0 M
O
C
Cu
CD
n Co
3 3 O
rt
s:
H-
rt
EJ*
H-
3 rt
ET
ro
0 M
0 e Cu
• H
E
T
H-
CO
H-
CO
Co
M
CO
0 rt
i-i
3 ro
0 Hi
rt
ET
ro
CO
Cu
rt
C
H
3 c Co
rt
ro
H n
M
0 c Cu
•
rt
ET
ro
i-f
ro
H-
CD
Co
0Q
O
O
Cu
T3 H
O
cr
Co
cr
H- (->
H-
rt
v;
rt
ET
Co
rt
H-
rt
•3
H-
M
rt
ET
ro
X
) i-i
0 cr
ro
s:
0 c M Cu
3 O
rt
X> ro
3 ro
rt
H
Co
rt
ro
n
0 3 X) M ro
rt ro
M
!_•
VJ
O
cr
rt
Co
H-
3 CO
0 3 ro
3 ro
Co
CO
C
i-i ro
3 ro
3 rt
CO
rt
ET
i-i
O
3 era
E
T
rt
ET
ro
0 M
O
c Cu
t# cr
c rt
3 Co
M
rt
3" ro
H1 X
> O
n Co
rt
H-
O
3 • M
Hi
rt
ET
ro
Co
rt
3 O
CO
X
) ET
ro
H
ro
H-
CD
Co
0 rt
3 Co
M
M
"<!
O
O
M
Cu
ro
i-i
rt
ET
CO
3 3 O
3 H-
3 Cu
M
H
0 a4 ro
rt
0 x>
Co
CD
CO
rt
ET
H
O
3 0Q
3
d
rt
ET
ro
c H CU
3 3 CO
CO
5 3 O 3 H-
P
O
H-1
O e Cu
H- 3 H-
rt
CO
3 O 3 H- 1
cr
ro
CD
ro
ro
3 v# rt
ET
ro
cu
ro
CO
0 ro
3 rt
rt
H- 3 ro
O
Hi
-O
•P~ 3 H-
3 3 rt ro
CD
s:
Co
CO
CD
ro
M ro
0 rt ro
Cu
rt
O
P
h-1
h-1
O S
i-i ro
CD
X) ro
0 rt
rt
O
X) i-i ro
CD
CO
c I-i ro
P
rt
IT
S' ro
ro
3 Cu
0 Hi
p
JN
JN
1 3 H>
3 C
rt
ro
Cu
ro
CD
n
ro
3 rt
• >
CD
n
P
3
0 Hi
X) I-i
ro
CO
CO
3 i-i ro
Cu
ro
X
I rt
ET
P 3 Cu
rt
ET
ro
H
ro
i-1
P
rt
H- <
ro
h-1
0 0 P
rt
H-
O
3 O
Hi
rt
ET
ro
X
I i-i
0 cr
ro
s:
H-
rt
ET
»r|
H-
era
c I-i ro
M
M
H 1 Ol
VO
H-
3 CO
ro
0 rt
H-
O
3 •r|
CO
ET
O
£ CD
rt
ET
ro
0 M
O
3 C
u
M
O
O
P
rt
H-
O
3 CO
p
CO
p
Hi
3 3 O
rt
ion
The uncertainties given are 10 on the exponent and ± 200 km on Zj. Despite these large uncertainties, this nominal equation was used for this preliminary look at Saturn and Uranus ionospheric investigations.
From Reference III-4, the projected state of the art sensitivity for the Langmuir probe is 10 electrons/cm3, and for the IRPA it is 5 ions/cm3. Using the equation previously given, Table 111-15 shows the approximate altitudes and times before entry when each instrument should begin to detect its respective particle.
Table 111-15 State of Pre-Entry Measurements
INSTRUMENT & PARTICLE
LP (e~)
LP (e")
IRPA (Hi"1")
IRPA (Hx+)
IRPA (H!+)
SENSITIVITY, particles/cm3
10
10
5
5
5
PLANET
SATURN
URANUS
SATURN
URANUS
JUPITER
ALTITUDE, km
3350
3200
3500
3350
1375 (REFERENCE)
TIME BEFORE ENTRY, sec
151
130
159
138
It can be calculated from the equation that with each 576 km of altitude, the number density changes one order of magnitude. Assuming a liberal two-order-of-magnitude error in the exponent ( 1200 km) and the 200 km error in reference altitude, the IRPA could possibly encounter protons at Saturn as high as 4900 km. Therefore, these pre-entry instruments should begin monitoring for ions at about 5000 km altitude, which is 237 seconds before entry. This time (and altitude) also does not include the trajectory errors. Twelve minutes prior to use, the Langmuir probes should be warmed up and decontaminated. The IRPA requires
111-43
only 5 to 10 seconds for warmup. However, the NRPA may require as long as 30 seconds, so that if turned on exactly at 5000 km, it may not start measurements until about 4420 km, which is still well above the required measurement altitude (probably less than 2000 km altitude for the neutral particles).
Figure 111-27 shows the altitude and velocity time histories from 5000 km to nominal entry at 536 km (532 km Uranus). The probe inertial velocity, the velocity relative to an ionosphere rotating with the planet, and the vertical descent velocity component are shown for both planets. The great difference in inertial velocity between Saturn and Uranus is directly attributable to their difference in size and gravitational constant. At Saturn, the probe enters with the planet's rotation; thus the relative velocity is significantly reduced from the inertial. However, at Uranus, the inertial entry velocity vector is approximately perpendicular to the direction of motion of the planetary atmosphere; thus, the relative velocity is about 0.5 km/sec higher than the inertial. The relative velocity is that velocity at which the pre-entry instruments will intercept particles, and it differs between Saturn and Uranus by about 4 km/sec.
The difference between the radial velocities for the two planets is always less than about 2 km/sec. Since the rate of vertical descent through the upper atmosphere is used to judge measurement performance of the pre-entry instruments, the fact that these velocities are about the same allows for commonality of sampling times and bit rates. This occurs only because of the selected entry flight path angles of 30° for Saturn and 60° for Uranus. The large difference in entry angle also explains the fact that the radial velocity lines in Figure 111-27 diverge. The small decrease in the flight path angle along the Saturn conic trajectory offsets the increased inertial velocity for calculation of the radial component. This does not occur for Uranus; thus the former decreases and the latter increases.
If the entry angle is varied, the performance will change with the radial velocity. Figure 111-28 shows the change in the various velocities as a function of entry flight path angle for Saturn. While the inertial velocity is almost independent, the relative velocity increases significantly with increasing flight path angle, and the radial velocity increases drastically. The effect of entry angle on radial velocity dominates the effect of mission selection and VH . (See Figure III-6.)
111-44
VELOCITY, km/sec
36,
34
32
30
28
S A T U R N » £ S ^
SATURN REi£2££
26h
24
22
20
18
URANUSJADIAL
200 100 0 TIME BEFORE ENTRY, sec
A , . :
-SATURN
-URANUS
200 100 0
TIME BEFORE ENTRY, sec
Fig. 111-27 Pre-Entry Velocity/Altitude Time History
INERTIAL
PROBE VELOCITY, km/sec
36
32
28
24
20
16
12
RELATIVE
hO 20 30 40 50 60 70
SATURN ENTRY ANGLE, deg
Fig. Ill-28 Probe Velocities at Entry vs Saturn Entry Angle
The resulting effect on pre-entry performance is gxven by Figure 111-29 for Saturn. For a given instrument sampling time, the number of kilometers of altitude required for a single measurement increases with the entry angle, and fewer measurements are ob- -tained tor a given mission. For the specific entry angles chosen, Figure 111-30 shows the variation in measurement performance with sampling time for both Uranus and Saturn. The lines are relatively close together so that the performance for a given sampling time (£5 sec) is nearly the same for the two planets.
The sampling times selected for the pre-entry instruments are shown on the figure. The Langmuir probes, sampling every half second, make one measurement every 10 km, which would be on the order of 280 separate measurements of electrons and protons, based upon the previously given equation. The IRPA, sampling at 2 seconds makes one measurement in approximately 40 km, which is about 70 proton density determinations and fewer for the heavier positive ions. The NRPA sampling time has been selected at 3.0 seconds, which requires the probe to travel about 60 km (more for Uranus, less for Saturn) between measurements. Since no model for the upper atmospheric neutral particles exists, the number of measurements cannot be estimated. If the neutrals are measured at and below a conservative 1500 km altitude, a total of 16 separate measurements will be made.
Ideally, the ionosphere should be sampled along a radial line to obtain a vertical ionospheric sample. However, because of entry at angles other than 90°, the probe will traverse a small amount of latitude and longitude. Table 111-16 shows these amounts for the three missions investigated. Both variations are small and are satisfactory for ionospheric measurements.
Table 111-16 Latitude and Longitude Variations'
PLANET
MISSION
A LATITUDE
A LONGITUDE
SATURN
S 79
0.5 deg
6.6 deg
SATURN
SU 80
1.3 deg
6.6 deg
URANUS
SU 80
4.7 deg
4.8 deg
*ALTITUDES FROM 5000 km TO 530 km.
111-46
20 40 60
SATURN ENTRY ANGLE, deg 80
Fig. Ill-29 Pre-Entry Performance vs Saturn Entry Angle
no
100
90
80
70 VERTICAL DISTANCE TO MAKE ONE MEASUREMENT, 6 0
km
50
40
30
20
10
0
•
•
-
LANGMUIR PROBES y
r i
URANUS, -Y = -50°
' • IRPA
i i i
/ * — SATURN, ' Y = -30°
• NRPA
i i I I
1.0 2.0 3.0
SAMPLING TIMES, sec
4.0 5.0
Fig. III-ZO Pre-Entry Instrument Measurement Performance vs Sampling Time
d. Entry Accelerometer Data Collection - The entry accelerometers will begin to store data upon sensing of a steady 0.1 g at about 5 seconds after entry. As established in the previous part of the study, the axial sampling will be at a rate of 5 samples/sec and lateral sampling at a rate of 2.5 samples/sec. This gives a total collection data bit rate of 100 bps, assuming 10-rbit words and two lateral sensors.
The range of accelerometer operational times given in Table' 111-14, show a variation from 48 seconds in the Uranus cool to 87.5 seconds in the Saturn warm atmosphere. This sizes the accelerometer data storage memory at about 8800 bits for entry into any of the six atmospheres. Additional storage is required for engineering, formatting, and some descent data.
The performance of the instrument for a given sampling time can be gaged from how detailed it defines the peak-g point on a g-time history. Figure 111-32 shows these deceleration curves for all six atmospheres. Table 111-17 lists the peak g values and then gives the number of separate measurements above three high-g levels, based upon the sampling times given above. Comparison of the number of points to the curve shows that the chosen sampling should be more than adequate.
Table 111-17 Accelerometer Measurement Performance
ATMOSPHERE
SATURN
WARM
NOMINAL
COOL
URANUS
WARM
NOMINAL
COOL
PEAK ACCELERATION,
9
227
336
568
225
358
837
NUMBER OF MEASUREMENTS AT
> 500 g
0
0
8
0
0
11
> 300 g
0
13
19
0
12
18
> 100 g
60
50
39
51
45
27
111-48
MO
O
rt
H
i fB
1
3 rt
h
+it
) 4
H
< H
- fB
F
- O
3
H
B4
F-
M
H-
P
<•
p1
33
i-
ir
tO
fB
3P
0Q
Cr
' rt
it
fD
i-(
H
i p
3
3 H
CO
(T>
C
O&
)"
<!
H-
l-
iM
fB
i-
i(
B
fD
fi
rt
M
F-
rt
fD
o fft
ifnitr
omtto
i M
r
tH
-f
B3
H-
OC
03
*i
-i
» H
M
p.
3 a.
<
rt
fB
H
M
o
co
o
ro
>d
n
• in
c
H
I 3
rt
O
13
3 P
X
en
en
I
H
i-
to
i-
i i-
io
CH
-W
OJ
W3
h-
'f
BP
Pf
BH
>-
(3
H
3 CO
rt
O
IB
B
4 fB
0Q
13
H
im
Su
BJ
ad
co
i-i
3 3
O
C
O
3 m
(i)
C
(t
f
lr
tC
lf
ll
ll
fl
rt
ar
t O
CO
(t
(t
•«
ft
ff
(6
C
u
?!
fD
H-
3 H
» 3
* F
- CO
rt
fD
03
3
3 fH
B
•
fD
Cu
CO
O
S
4 CO
rt
P
•
1-i
fB
(t
(C
n
D-
fD
n
in
fD
fl
rt
9 fB
fB
C
u
B4
3 U
»
M
F-
It
CB
3 03
(1
h
ii
j (D
O
B
3
rt
O<
! ti
flC
Brt
SlIi
rtfl
(D
&>
rt
S
i-i
O
3*
B
F-
CT
3M
(
Df
Br
tf
Bf
Bf
Dc
rS
OP
rt
C
i-i
3 B
4 3
3^
It
Hi
H
fB
rt
ft
rt
13
rt
Hf
l CO
<
i-i
CD
M
» 0)
F
- rt
p
. i-i
-
•O
C
U
3 CO
S
4 CD
O
H
i fB
C
u
F>
3 S
4 n
ft
H
i O
C
Oi
-S
F-
3f
DS
4P
P
H-
H-
i-i
fD
Hi
CO
rt
rt
F-
3 B
. fl
P
O
OO
O
OO
Of
B
fDC
O
rt
ii
C
34
34
CO
C0
OC
DO
F
- B
CD
CO
p
3
3F
-0
B
OP
CD
PC
OF
-r
tC
DO
Qf
BH
lf
B
3 O'
(B
F-
CD
00
F- < fB
3 F- 3 n B4 P
13 rt
fB
i-
i
F- O
3 O
3 Cu
(B
CD
O
fB
3 rt &)
3 a.
rt
3- fB
rt
C
i-i
3 cu
H
fB
Cu
fB
fB
13 fB
I-
i
Hi
O
I-i
CU
M
M
fu
CD
3 (U
M
M
fB
Hi
Hi
fB
O
rt
C
X) o 3 Cu
fB
CD
cu
< (U
I-i
F- cu
cr
M
fB • a o rt
fB
rt
3*
CU
fB
&
3 cu
M
rt
O
IT S'
(B
rt
F- 3 (B
Hi
H
O B
rt
cu
3 rt
J^
-P- B
F- 3 C
rt
fB
CD
cr
c rt
rt
3*
fB
3 B
F-
CD
CD
F>
O
3 F-
CD
F- 3 Cu
F-
O
CU
rt
fB
Cu
Hi
Hi
F-
O
F-
fB
3 rt
O
Hi
o n fu
M
M
P
rt
B
0 CD
M
13
M
O ?r
OQ
*•—.
B
B4
fB
H
fB-
CD
O
N> H
i •
cr
O
Hi
rt
B4
fB
Cu
fB
CD
O
(B
3 rt
X> P n CU
3 fB
rt
CU
CD
rt
C
U
O
fB
CD
fB
O
rt
H
rt
H
CD
X
fB
3"
3* C
D
O I
3 fB
I
rt
rt
fB
i-t < \»
C/3
(B
O
rt
F-
O
3 W
•
O
•Hi
l-( fB
X) I—
1
CU
^ F- 3 O
P
13 i-i
fB 1 fB
3 rt
H
^ Cu
fu
rt
Cu
• H
3"
H-
CD
F-
CO
Cu
F
-CD
O
3 CO
CD
fB
C
u
Hi
3 i-i
rt
3*
fB
H
F- 3 o B4
CU
13 1
cu
<J
cu
F-
M
cu
w
M
fB
F-
CO
rt
O
F- 3 O
i-i
fB
Cu
CD
fB
rt
S4
fB
Cu
fD
CD
O
fB
3 rt
CO
o F-
fB
3 O
fB
B.
fB
CU
CD
3 I-i
fB
3 fB
3 rt
CD
F- 3 CD
rt
fB
CU
Cu
CD
rt
F-
M
M
i-i
fB
.n
3 F-
H
fB
CO
M
fB
CD
CD
13 O
-3
fB
I-i
rt
B4
cu
3 13
i-i
(B 1 fB
3 rt
<-i
<<
• > 3 O
rt
3"
fB
i-i
O
13 rt
F-
O
3 Cu
M
CO
O
Hi
O
i-i
cu
n o CD
rt o Hi
o 3 M
*<:
K>
ON
"3
1 B4
i-i > H
B4
(B
O4
F-
rt
i-i
Cu
rt
fB
CT
fB
O
O
B
fB
CD
C/l
rt
3*
(B
Cu
CU
rt
CU
O
O
C
M
Cu
C74
(B
CD
rt
O
H
fB
Cu
CU
3 Cu
M
fB
13 M
CU
^ fB
Cu
Hi
O
I-i
H
fB
Cu
c 3 Cu
cu
3 O
O
"-
t4
> Ln
C74
13 CO
€ B4
F-
O
B4
Cu
3 i-i
H
-3 cro
C
u
fB
CO
o ent
13 H
fB i n>
3 rt
i-i
^ i-i
CU
rt
fB
O
Hi
NO
CO
l_
n
o4
13 CD
fB
X
o fB
fB
Cu
CO
rt
B4
CU
rt
i-i
(B
^3 c H- i-i
fB
C
u
ID
X
13 cu
3 C
u
fB
Cu
13 cu
^ M
O
cu
C
u
Cu
fB
CO
n fB
3 rt
• EC
o -3
fB <!
fD
I-i
V rt
B
4
fD
13 O
-3
(B
H
i-i
fB
Hi
JZ
O
I-i
Cu
fB
CD
O
fB
3 rt
*••
rt
B4
C
CO
e H- I-i
fB
B
fB
3 rt
Hi
O
i-i
rt a- fB
H-
rt
H-
3 o I-i
fB
CU
CO
fB
CO
rt
O
u>
M • IJl
cr
13 CO
• H
B4
H-
CD
H-
CO
rt
3"
fB B
H-
3 H- 3 e 3 i-i
fB
>Q
3 H-
i-i
fB
3 fB
3 rt
Hi
o t-i
rt
B4
fB
rt
O
i-i
•<!
13 CU
<
CO
c»
g
§ 3 cu
H
M
Vi
O
CU
Cu
H-
CD
N3
^J • ^J
CT
1
13 CD
>#
CO
3 Cu
K
H-
rt
B4
rt
B4
fB
Co
Cu
C
u
H-
rt
H-
O
3 O
Hi
rt
B4
fD
3 fB
13 B
4
fB
M
O 3 fB
ter
Hi
O
i-i
Cu
H
-CO
O
r- CO
CO
H- o n
B4
fD
H
fB • H
B4
fD
rt
O
rt
P
h-1
Cu
CU
rt
P
i-i
CO
rt
fD
Hi
O
t-i
rt
O
B4
Co
13 rt
(B
i-i
CO
o o 3 O
CD
I-i
3 H-
3 00
CO < CO
rt
fB
B
Cu
fD
H
i H
-3 H
-rt
H
-O
3 CD
*#
a4
3 rt
H
P
C34
M
fD
M
M
M 1 M
00
p'
X)
fD
W
X
X) M
ora-
i-i
O < H- Cu
fB
CD
H
3"
fB
Cu
fD
rt
P
H
-M
fD
C
u
Cu
P
rt
P
H
P
rt
fB
a4
i-i
fB
P
?r
Cu
o si
3 H
i O
H
.
fB
P
O
B4 3 H-
CO
CO
H-
O
3 H-
CD
OQ
H- < fB 3 H-
3 rt
B4
fD
3 P
*<:
a4
(B
H-
3 rt
fB
i-i
M
fD
P < fB Cu
C
H-
rt
B4
rt
B4
fB
H
fB
P
M
rt
H- 3 fD •
=3
H-
rt
B4
rt
B4
fD
IB
X
13 P
3 Cu
fB
C
u
13 P
VJ
CD
O
(B
3 rt
*»
rt
B4
fB
CO
rt
O
H
fB
fB
3 rt
i-i
M
^ O
P
Cu
^ rt
B4
fB
fD
3 rt
H-
H
fD
13 i-l
fD
1 fD
3 rt
i-l
"<!
O o
P
O
O
fB
M
fB
H
O 3 fD
rt
fD
H
Cu
p
rt
P B
3 CO
rt cr
fD
M
13
M
fD
O
rt
H-
O
3 O
Hi
Cu
P
rt
P
M
p
^ fD
Cu
D4
P
O
!V
*• P 3 Cu
3 fB
13 B
4
fB
M
O 3 fB
rt
fB
I-l
•3
H-
(-•
M
CD
ID
P
H
O
B4
H-
3 Cu
(B
rt
P
H
-M
Hi
o I-l
O
M
O
3 Cu
CO
• >
•3
H-
M
M
a4
fB
o 13 fB
i-l
P
rt
H
-3 OQ
• M
3 p
Cu
C
u
H-
rt
H-
O
3 *• O
3 rt
V
ID
(B
X
13 P
3 Cu
fB
C
u
|-i
x)
CD
O
V C
u
3 i-i
H-
3 00
Cu
fB
1
P
V! M
O
P
Cu
>#
p
p 3 Cu
rt
B4
fD
P
O
O
fD
M
fD
I-i
O 3 fD
rt
fB
i-i
rt
i-i
H-
P
Cu
•-N
CO
-3
H
-rt
O
rt
B4
fB
3 fB
3 rt
H
P
M
3 P
CD
CD
CO
XI (B
O
rt
I-i
O
B
fB
rt
fB
I-i
>•
rt
fB
3 B
4 1
3 fB
C
u
rt
O
rt
B4
fB
rt C
H
O4
C
M
fB
3 O
(B
3 o Cu
fB
^s
0)
H P
rt C
H
fD
P 3 Cu
X) H
fD
CD
CD
C
H
fB
OQ
P
OQ
(B
CO
v
TO
• t)
C6
Co
Ci
CG
S
<D-
t)
» <rt-
ft
Qi
O
w
^~
i C<
b
CJ
<i- ^.
O s » s Ro
»-a
hS
R s CO
3 ^.
CO
CO
^.
o S 1 a c H H-
3 OQ
Cu
ID
CD
O
ID
3 rt
*•
I
Table III-18 Descent Data Rate Composition
COMPONENT
EXPLORATORY PAYLOAD SCIENCE
FORMATTING
ENGINEERING
EXPLORATORY PAYLOAD TOTAL
NEPHELOMETER ADDITION
FORMATTING
EXPANDED PAYLOAD TOTAL (LESS
TOTAL STORED PRE-ENTRY
STORED FORMATTING
STORED ENGINEERING
TOTAL DESCENT
PRE-ENTRY)
BIT RATE, bps
24.7 (Table IV-4)
2.5
0.5
27.7
3.4
0.4
31.5
16.8
1.7
0.5
50.5
PRESSURE, 2 bars
LEGEND:
SATURN
URANUS
10 20 30 TIME FROM ENTRY, minutes
40
FINAL EOM PRESSURE, bars
50
Fig. Ill-31 Pressure Descent Profiles
111-50
SYSTEM INTEGRATION
Parachute Deployment Analysis
The expected deceleration of the probe varies significantly for the cool, nominal, and warm atmospheric models for Saturn and Uranus as shown in Figure 111-32. Accordingly, it was necessary to investigate the parachute deployment to meet the following criteria:
1) The parachute shall be subjected to Mach numbers no greater than 0.9.
2) The parachute shall be deployed at pressure altitudes higher than 100 mb (a soft constraint).
From the initial portion of the contract, the 100 mb pressure level for an entry ballistic coefficient of 102 kg/m2 (0.65 slug/ ft2) occurs at or near Mach 0.7. This Mach number varies from 1.38 g for the Uranus warm atmosphere to 2.4 g for the Saturn cool atmosphere. Using 2.4-g sensor to deploy the parachute would expose the chute to Mach 0.95 in the Uranus warm atmosphere and Mach 1.0 in the Saturn warm atmosphere. Using 10-g (decreasing) sensing for the Saturn nominal atmosphere plus 20 seconds, the highest velocity encountered is Mach 0.9 in the Saturn warm atmosphere. Due to errors expected in sensing devices, timers, and determining ballistic coefficients, 5 g (decreasing) plus 15 seconds was selected for parachute deployment. As a result, the following Mach numbers and pressure altitudes are expected for the various atmospheric models:
Mach No. Presssure Altitude, mb
Saturn Atmospheres
Cool 0.56 102
Nominal 0.70 65
Warm 0.88 44
Uranus Atmospheres
Cool 0.59 41
Nominal 0.72 47
Warm 0.84 49
111-51
1000
50
0-
-C00
L
ATM
OSP
HER
E
100
50
DECELERATION, EARTH g
2.0r
1.5
MACH N
O. 1.0
0.5
10
OL
1
-NOMINAL A
TMOSPHERE
WARM A
TMOSPHERE
ENABLE
5-g
SENSOR
NOTE
: 1. HEAVY
LINES DENOTE URA
NUS;
LIGHT
LINES DENOTE S
ATURN.
2. B = 102 kg/m2?(0.65
sl
ug/f
t2).
LEGEND:
DECELERATION
MACH NO.
M = 0.7
P = 65 mb
"V
rM = 0.88
: 44 mb
P = 49 mb
M = 0.84
M = 0.59
40
60
TIME FROM
ENTRY, sec
100
Fig
. 11
1-32
P
arac
hute
D
eplo
ymen
t A
naly
sis
for
Satu
rn
and
Ura
nus
The 5-g sensor would be enabled by 100 g (sensing) as shown in
the figure.
Task II Data Mode Analysis
The data mode for Task I is very similar to that for the Probe Dedicated Jupiter Mission described in Volume II, Chapter V, Section C. Since the probe is deflected by the spacecraft, the probe separation time is very short, i.e., a few minutes, during which time data is not collected. During the pre-entry phase, engineering data is collected and transmitted in real time until entry when the transmitter is turned off due to the blackout encountered. Science and engineering data is collected and stored until probe acquisition and interleaved with real-time data during descent at approximately 28 bps.
Compared with Task I, the addition of the pre-entry science instruments and the nephelometer for Task II caused the pre-entry data rate to increase from 1 bps to 235 bps and the descent data rate to increase from 28 bps to 31.5 bps, assuming that pre-entry science data is not stored. The high pre-entry data rate precipitated a data mode tradeoff as follows:
1) Keeping the pre-entry transmission real time and using FSK modulation will require a high RF power; changing the modulation to PSK should reduce the RF power required.
2) Storing the pre-entry data (i.e., no pre-entry transmission) and interleaving it with real-time data during descent would increase data storage requirement, but decrease the RF power requirements.
The data profiles for the primary and optional data modes are shown in Figure 111-33 with the optional mode corresponding to 2) above. In the primary mode, the pre-entry science starts collecting data at the time corresponding to 5000 km above 1 atmosphere, but due to the entry arrival uncertainty of 29 minutes for Uranus and 4.4 minutes for Saturn, the pre-entry science starts collecting data earlier than required. For the optional mode, the Langmuir probes and IRPAs are used to sense data at approximately 4300 km above 1 atmosphere and instruct the Data Handling Subsystem to start storing data. This adaptive science method eliminates the entry arrival uncertainty and accounts for less data collected in the optional mode.
111-53
0 12
24
36
48
TIME FROM ENTRY, m
inutes
a]
OPTIONAL MODE
36
-24 -12
0
12
TIME FROM ENTRY, m
inutes
b) PRIMARY
MODE
111-
33
Dat
a P
rofi
le
for
Tas
k II
, D
ata
Mod
e T
rade
off
Results of these trades, shown in Table 111-19 were presented at the Mid-Term Review held at Martin Marietta Denver Division on 13 October 1972. Since data modes 2 and 3 require the same RF power without increasing the probe weight or size, it was recommended at the Mid-Term review that the pre-entry data be transmitted real time as in mode 2 and also stored for descent transmittal as in mode 3, as shown in Figure 111-34. These trades are also discussed in Sections B and D.
After conducting the thermal control analysis of the combined data modes 2 and 3 shown in Figure 111-34, the long real-time transmission times for Uranus caused the transmitter to exceed the upper temperature limit. As a result, a 7-watt transmitter amplifier was added for pre-entry tracking and acquisition. At 3.3 minutes before entry, the adaptive science method starts the real-time transmission at 41 watts and begins storing data. This condition is further discussed in Sections D and E.
The pre-entry RF power required for mode 2 is 41 watts and the descent RF power required is only 29 watts. As discussed in Section B, it might be more desirable to use the increased descent RF power by improving the descent science performance, i.e., increase the NMS sweeps instead of the combined data mode 2 and 3.
Table 111-19 Task II Data Mode Analysis Results
DATA MODE
PRE-ENTRY DATA,
TRANSMITTED
REAL-TIME (NO STORAGE)
DESCENT
PRE-ENTRY DATA
TRANSMITTED
REAL-TIME (NO STORAGE)
DESCENT
PRE-ENTRY DATA STORED & TRANSMITTED DURING DESCENT
DATA RATE, bps
235
31.5
235
31.5
50.5
MODULATION TYPE
FSK
FSK
PSK
FSK
FSK
RF POWER, w
120
29
41
29
40
PROBE WEIGHT, kg
116
103
103
PR0GE DIAMETER, cm
96.5
91.4
91.4
PROBE DATA STORAGE 103 bits
11.9
11.9
60.2
111-55
I
1/1
as
PRE-ENTRY
ENTRY
DESCENT
534,614-
DATA, 103 bits
END OF
MISSION
-60
-48 -36
-24
-12
0
12
24
36
TIME FROM ENTRY, minutes
Fig
. 11
1-34
P
reli
min
ary
Dat
a P
rofi
le
for
Satu
rn/U
ranu
s P
robe
(T
ask
II)
ELECTRICAL AND ELECTRONIC SUBSYSTEMS
Additional analyses have been performed in several areas of the electrical and electronic subsystems to further refine hardware design criteria. Only the changes and improvements made to the original subsystem, which are described in Volumes I through III, are discussed in the following sections.
Telecommunication Subsystem
Parametric variations were made in arriving at an optimum geometry that results in a common Saturn/Uranus probe from the standpoint of the telecommunication subsystem. Results of parametric variations are discussed together with changes and improvements to some preliminary data used previously. New hardware information is also included.
a. System Noise Temperature - Previous determination of the system noise temperature did not include effects of antenna feedline noise or the location of the galactic center. Disk noise for Uranus was also preliminary data obtained before release of the Uranus design criteria monograph (Ref III-2). Improvements in system noise temperature calculations were made for Saturn and Uranus and are described in the following paragraphs.
Various noise sources reduce the useful operating range of the receiver system. There are several external sources of radio noise such as galactic, star, magnetosphere, planet, and atmospheric noise. Internally, the antenna and receiver also have a noise level which is expressed in terms of noise temperature. Noise temperature can also conveniently be expressed in terms of a noise figure as defined in Volume III, Appendix B, p B-10. The thermal noise sources affecting the probe receiver subsystem are shown in following sketch.
— • Antenna Feed
TA
Feedline
F' F
RF Amplifier
TR
Mi
i xer
r IF Amplifier
Local Oscillator
Detector
111-57
The system noise temperature is used in all considerations of sensitivity and in calculating the total effective noise power density per modulation bandwidth (kT \. System noise temperature is defined as
Ts - TA + TF + VR t1]
where T = receiving system noise temperature, °K
T = antenna noise temperature, °K
= TG + T B g + T B D, as applicable
T = galactic noise temperature, °K
T = synchrotron noise temperature, °K Bo
T = planet disk noise temperature, °K ou
T = antenna feedline noise temperature, °K r
T = receiver noise temperature, °K K
L = feedline power loss ratio, r
Feedline loss affects the noise temperature of the receiver and the feedline temperature in °K is expressed by
T p = 290° (LF - 1). [2]
The loss in 3 ft of coaxial cable (from antenna feed to receiver) is 0.2 dB/ft at 0.86 GHz or 0.6 dB. The two connectors add 0.4 dB for a total loss of 1 dB (power ratio = 1.26). From Equation [2], we have
T_ - 290 (1.26 - 1) = 75.5°K. [3] r
It is important to have the receiving anfenna close to the receiver with a minimum feedline cable length. Low loss cable and connectors should be used on the spacecraft to keep reflected noise temperature to a minimum.
111-58
The receiver is an amplifier and mixer chain (see sketch) to the detector, which determines the receiver noise temperature. The receiver noise temperature, T , of a typical superheterodyne re-
K ceiver, with RF amplifier and a mixer down converter to IF frequencies, is given by the relationship
T2 TR = Tl + GT '
where
T = receiver noise temperature, °K
Ti = RF amplifier noise temperature, °K
T2 = mixer noise temperature, °K
Gi = power gain ratio of RF amplifier.
For an amplifier and mixer package where T1 = 289°K, T2 = 3000°K, and Gi = 30 dB (ratio = 1000), the receiver noise temperature from Equation [4] is
1000° TR " 2 8 9°. + I56F " 2 9 2 ° K
As can be seen, the noise temperature if this receiver is controlled by the first amplifier stage and the other, high gain stages in the receiver only add a small amount of noise.
Noise contributions from the magnetosphere and planet disk vary during the probe mission. Background cosmic noise from the galaxy can be significant, as seen in Figure 111-35 (Ref III-6), depending on whether or not the center of the galaxy is within the beam-width of the probe receiving antenna on the spacecraft. The position of the galactic center relative to the antenna look direction is dependent upon the arrival date, the planet of interest, and the spacecraft-to-probe geometry. As seen in Figure III-36a, the galaxy is occulted by Saturn during probe acquisition and during the active portion of the probe mission. The planet disk subtends an angle at the spacecraft of 28° and galactic background noise is overshadowed by the disk noise. Cosmic noise present between the outer limit of the magnetosphere and the spacecraft is negligible (< 10°K).. The direction to the center of the galaxy is also shown in the figure; it is 19° off the spacecraft antenna boresight at acquisition.
111-59
GALACTIC NOISE TEMPERATURE, Tc, -K
To3
RF FREQUENCY, MHz
GALACTIC . CENTER
Fig. 111-35 Thermal Noise Temperature of the Milky Way Galaxy
E - 0.1 (ACQUISITION) l-ENTRY
a) SATURN GEOMETRY
PROBE RECEIVING ANTENNA LINE-OF-SIGHT
ENTRY GALACTIC CENTER
b) URANUS GEOMETRY
Fig. 111-36 Trajectory Geometry for the Saturn/Uranus 1980 Mission
111-60
Therefore, only the variable synchrotron and planet disk noise is present at Saturn encounter as a function of the relative geometry of the probe, planet, and spacecraft. Within the operating frequency range of the probe receiving antenna on the spacecraft, noise power produced by black body radiation from the planet disk and magnetosphere is received through the antenna radiation envelope and is coupled to the receiver through the radiation resistance of the antenna. The fraction of noise power received by the spacecraft antenna is a function of the directivity of the antenna with most of the power received via the main beam. The noise temperature of the antenna from each direction is a function of the power (flux) emitted into the antenna from that particular direction.
The revised receiving system noise temperature for Saturn is shown in Figure 111-37. The original curve is shown in Figure VI-7 of Volume II. The revised curve represents an increase of 149°K at 0.86 GHz, resulting from feedline noise and its impact on receiver front-end noise temperature. A feedline loss of 1 dB is used for 0.88 m (3 ft) of coaxial cable. The antenna temperature remains unchanged, based on the latest values for synchrotron and disk brightness temperatures for Saturn (Ref III-l).
The trajectory geometry for Uranus is shown in Figure III-36b and indicates that the line-of-sight (LOS) of the receiving antenna grazes the planet disk at acquisition. For a 20° antenna beam-width, half of the beam encounters space noise and half disk thermal noise with the planet disk emanating the higher noise level. Also depicted in the figure is the direction to the galactic center. The LOS vector is 53° away from the galactic center at acquisition. Therefore, the antenna noise temperature is attributed only to the influence of the planet disk. The planet disk subtends an included angle of 24° at the spacecraft position during acquisition.
The revised receiving system noise temperature for Uranus is shown in Figure 111-38. The original curve is shown in Figure VII-2 of Volume II. The revised figure depicts a new upper-limit curve for the planet disk brightness temperature based on measured data reported in the planet monograph (Ref III-2). The basic study assumed a constant disk brightness temperature of 300°K. The monograph also discusses the possibility of trapped radiation belts existing at Uranus. The conclusion is that radio emission measurements do not establish the planet as a synchrotron source since the existing measurements are in rough agreement with what would be expected of thermal emission from the atmosphere (Ref III-2, p 28). Therefore, the antenna noise temperature is equal to the planet disk brightness temperature, as shown in the figure, and is constant during the active portion of the probe mission.
111-61
.TEMPERATURE, "K
150 50 30 20 15 UAVELENGTH, A, cm
Fig. 111-37 Receiving System Noise Temperature for Saturn
100O
500
EFFECTIVE NOISE TEMPERATURE, 'K
300
200
100 0
1 2
1
NOTE
I
.^•^"""sYSTEM NOISE TEMPERATURE, T $
ANTENN
DISK 8RIGHTNE TEMPERATURE,
TA = TBD•1
1 1 1 1.0
LEGEND: O = PLANET DISK
TEMPERATURE OBSERVATIONS (1966 1 1969)
A TEMPERATURE, Tft
S
BD
1 1 1 1 1 2.0
o
o
1 1 1 3.0
RF FREQUENCY, GHz
L-J I I I 1 I I I I I I I I 150 50 30 20 15 12 10
UAVELENGTH, J , cm
Fig. 111-38 Receiving System Noise Temperature for Uranus
111-62
The average receiver noise temperature curve of Figure V-ll, p V-21, Volume II, is used and a feedline loss of 1 dB is used for 0.88 m (3 ft) of coaxial cable. The revised system noise temperature curve represents an increase of 102°K at 0.86 GHz resulting from feedline noise and its impact on receiver front-end noise temperature and a revised antenna noise curve. The revised system noise temperature curves are used to determine the receiver noise spectral density in the RF link analysis.
b. Atmosphere Microwave Losses - The basic method of computing absorption and defocusing losses in the planetary atmosphere is described in Volume III, Appendix A. Results were given only for the nominal atmosphere models of Saturn and Uranus. For the common probe design, the worst-case atmosphere was used to define the communication subsystem.
For an atmosphere such as that found on Saturn and Uranus, the principal source of microwave absorption is gaseous ammonia. Water vapor and liquid water/ammonia solution clouds also contribute to absorption losses. Absorption in solidified ammonia or water clouds is negligible. During descent, the solid ammonia clouds are encountered before the water/ammonia solution clouds, as seen in Figure 111-39. In general, the atmosphere model having the highest ammonia concentration will result in the greatest microwave absorption if the probe penetrates the ammonia saturation level; descent into the cool Uranus atmosphere is an exception because the probe does not reach the saturation level (80.7 bars) as seen in the figure. The saturation level is approximately 6.7 bars for the nominal Uranus atmosphere and the probe descends to approximately 7.3 bars at mission completion. Ammonia abundances are shown in Table T.II-20 for Saturn (Ref III-j., p 43) and Uranus (Ref III-2, p 33). As seen in the table, cne warm models for both planets are very low in ammonia and were therefore not investigated because the resulting micro-Wave absorption will be smaller than for the nominal or cool models. Methane clouds were not considered in the analysis since their microwave absorption is negligible.
There have been two changes made in the computation method described in Volume III, Appendix A. The improvements, which are based on more recent analysis, did not have a large effect on the numerical results. The nominal planet atmosphere models for Saturn and Uranus were also recalculated to update the data on microwave absorption.
111-63
ioor
5 0 -
0 -
-50
ALTITUDE, z , km
-100
•150
-200
-250
-300
SATURN W N C
NH<
H 2 0 § I C E ^
mm
NH3:
H20 :ICE S§
H20/ NH3
SOLUTION^/
JH 20§
URANUS W N C
J I I LEGEND:
W
N
C
<
END OF DESIGN MISSION
WARM MODEL
NOMINAL MODEL
COOL MODEL
AMMONIA SATURATION LEVEL
!NH
H20/: 5ICE :
HI
|NH-
!!&
H20/ ', NH3 ll SOLUTION/
vim i
NH:
Fig. 111-39 Location of Clouda that Affect Microwave Absorption
111-64
Table III-20 Ammonia Abundance in the Atmosphere Models
Ammonia Mass Fraction,^
SATURN ATMOSPHERE
WARM NOMINAL COOL
0.037 0.113 0.34
URANUS ATMOSPHERE
WARM NOMINAL COOL
0.038 0.095 0.168
The first change consisted of an alteration of the method used to compute ammonia (NH3) absorption coefficients. The method outlined in Volume III, Appendix A was based on work done by A. A. Maryott at National Bureau of Standards, Gaithersburg, Maryland and used in the original analysis (Ref III-7). The new method is based on work recently completed by D. A. DeWolf of RCA, Princeton, New Jersey (Ref III-8).
Figure 111-40 shows a comparison of the two methods of computing absorption coefficients, a(z), using the Jupiter cool atmosphere at 1 GHz. Jupiter is shown because an atmosphere profile that extended to very high pressures was needed to show the turnover that occurs at -100 km (approximately 65 bars). Above this altitude, the two coefficients are very close. Below this point, the high pressure asymptotic forms prevail: T-1 for Maryott's coefficients and T-1,1+for the DeWolf coefficients. Therefore, the curves will continue diverging as depth and temperature are increased. Similar differences in the two methods of computing the coefficient exist for Saturn and Uranus, which result in differences in absorption as seen in Figure 111-41. The zenith microwave absorptions for the nominal and cool Saturn atmosphere models and the Uranus nominal atmosphere model are shown in Figure 111-41, through 111-46. Each model has two figures associated with it to depict absorption as a function of altitude and pressure at two selected frequencies, or as a function of frequency at several selected altitudes as represented by pressure. Use of the coefficients developed by Maryott and DeWolf is shown for the nominal atmospheres for purposes of comparison. The curves were calculated only to a pressure of 30 bars and extrapolated to higher pressures. The zenith absorption is equal to the total microwave attenuation since the defocusing losses are zero at zenith. The zenith absorption is the integral of the coefficients given in a plot such as shown in Figure 111-39. The difference between the two methods of computation is only about 23% over the altitude range of interest for this program. However, the change to the DeWolf method was made because it is based on more recent and extensive experimental work. Defocusing losses for the two planets are similar to Figures A-8 and A-9 in Volume III, Appendix A, and are less than 0.1 dB for the altitudes and aspect angles encountered during the missions.
111-65
ABSO
RPTI
ON
COEF
FICI
ENT, 10
ct(z), d
B/km
LEGEND:
——
NBS,
MA
RYOT
T
RCA,
DEWOLF
NOTE
: FR
EQUE
NCY = 1
GHz
.
J I
AL
TIT
UD
E,
z,
km
Fig
. 11
1-40
A
bsor
ptio
n C
oeff
ioie
nts
for
the
Coo
l Ju
pite
r A
tmos
pher
e
-28
0
20 RF ABSORPTION,
Fig. 111-41 Zenith Absorption for the Saturn Nominal Atmosphere
RF ABSORPTION
' 10|
LEGEND:
RCA
NBS
SATURN PRESSURE,
7 / / /
bar
/
1
0 . 2 0 / 6 1.0 1.4 1.1
RF FREQUENCY, GHz
2.2 2.6 3.0
Fig. 111-42 Zenith Absorption at Selected Pressures the Saturn Nominal Atmosphere
LEGEND:
RCA
EXTRAPOLATION
l1
IV AMMONIA ICE CLOUD
20 RF ABSORPTION,
WATER/AMMONIA SOLUTION CLOUD
ATMOSPHERIC PRESSURE, bar
Fig. III-4S Zenith Absorption for the Saturn Cool Atmosphere
24
RF ABSORPTION,
SATURN PRESSURE, bars
20
0.2 0.6 1.0 1.4 1.8 2.2 2.6 3.0
RF FREQUENCY. GHz
Fig. 111-44 Zenith Absorption at Selected Pressures for the Saturn Cool Atmosphere
111-68
ALTITUDE, z , km - 2 0 0
\ \
WATER/AMMONIA SOLUTION CLOUD,.
w N
V; i
20
ATMOSPHERIC
PRESSURE, bars
10 20 30 40
RF ABSORPTION, <)B
Fig, III-4.5 Zenith Absorption for the Uranus Nominal Atmosphere
12
RF ABSORPTION,
Fig. 111-46 Zenith Absorption at Seleoted Pressures for the Uranus Nominal Atmosphere
III
The second change in the computation consists of implementing a calculation of the variable ammonia and water abundances in the saturated region in and above the clouds, rather than setting the abundances equal to zero above the saturation elevation. Absorption in this region is very low compared to that below the saturation elevation. Therefore, it can be neglected for probe missions that penetrate any distance below the ammonia saturation elevations, shown in Figure 111-39 by the triangles. The necessity of adding this computation arose from examination of the cool Uranus model. The ammonia saturation level occurs at a pressure of 80.7 bars, so a 20-bar mission would be flown completely in the saturated region of the atmosphere. The equation used to compute ammonia abundance is shown as Equation [IVK-25], page IV-99 of Reference III-7, and was taken from Reference 111-97 A similar expression was also used for the water abundance.
Computation of the absorption loss in the cool Uranus atmosphere model indicates that it is completely negligible due to the very low ammonia abundance in the saturation region and the fact that the mission is over before the ammonia cloud level is reached. Zenith absorption at 1 GHz and a depth of 28 bars was only 0.007 dB. At 3 GHz and 28 bars, absorption is only 0.06 dB. Absorption curves for the cool Uranus atmosphere similar to Figures 111-45 and 111-46 were therefore not plotted because the absorption is essentially zero to 28 bars. The nominal Uranus atmosphere and cool Saturn atmosphere are the two worst case atmospheres, with the cool Saturn model providing the greatest microwave absorption as seen in Figures 111-43 and 111-44.
o. Probe Antenna - One result of the parametric study to achieve a common probe design for the Saturn and Uranus missions was a decrease in probe aspect angles before entry. An axial pattern with 100° beamwidth will satisfy the requirements for pre-entry and descent antennas. A pair of identical antennas will satisfy the design requirements; the turnstile/cone antenna provides the most compact design among, several considered.
An internally funded study (Ref 111-10) has developed prototype turnstile/cone antenna design with circular polarization. The measured radiation and polarization patterns are shown in Figure 111-47. The main beam radiation pattern is uniform and approximately symmetrical in both the E- and H-planes. The on-axis ratio is 1.8 dB and only 2.8 dB at the 3-dB points. High circularity is a result of improved feed and ground-plane design.
111-70
«•- CONDITIONS
E-PLANE VOLTAGE 100° BEAMWIDTH 0.86 GHz LINEAR SOURCE
AXIAL SCALE = 5 dB/RING
b) CIRCULAR POLARIZATION PATTERN
Fig. 111-47 Circularly Polarized Turnstile/Cone Probe Antenna Patterns
111-71
A revised probe antenna design is shown in Figure 111-48 for the turnstile/cone antenna. Comparison with the original design (Vol III, Figure D-7, p D-12) indicates an envelope diameter reduction from 20.3 to 15.2 cm (8 to 6 in.) and slightly increased (0.6 cm, 0.25 in.) height. The dipoles are equal in length, and wings have been added to the truncated cone apex to control pattern circularity. The weight is unchanged. As seen in the perspective view of the figure, the dipoles and wings must be supported by dielectric material (Teflon, etc) in order to withstand the high g-forces during planet entry. The antenna patterns were measured using a prototype antenna design mounted on an aluminum hemisphere 61 cm (24 in.) in diameter. The test model is shown in Figure 111-49. Patterns were measured on domes of several sizes and the ground plane area affects the backlobe distribution. Therefore, full-scale radiation and circularity patterns should be recorded during hardware development to verify the predicted pattern coverage.
Thermal analysis of the descent probe indicates that the antenna will be exposed to a temperature of approximately 50°K (-370°F). The turnstile/cone antenna was qualified for operation at 172°K (-150°F) for the Viking mission. Cryogenic operation at the low temperature could result in problems with the printed circuit balun feed and the dipole elements. Preliminary analysis indicates that a radome may be required over the probe descent antenna with a heating element or thermal insulation such as spun fiberglass. Fiberglass or Teflon would be a suitable radome material and would provide thermal insulation.
d. Polarization Loss - In the previous study an arbitrary value of 0.2 dB was used throughout all the RF link analyses as an adverse tolerance resulting from polarization loss between the probe and spacecraft antennas. During the follow-on study, an equation was added to the computer program (LINK) to calculate the cross-polarization loss as a function of the axial ratios of the two antennas and the ellipse axis aspect angle. The equation that relates this loss is expressed by:
!
1 2 Ri R2 (l - Ri )(l - Ro, L + ' + _1 LA |i_ c o s C2g) 2 (l + Rf)(l + Rl) 2(l + Ri)(l + Ri)
where
P = polarization loss, dB i-i
R = probe antenna axial ratio (voltage)
111-72
PERSPECTIVE VIEW
14.7 cm (5.8 in.)
NOTE: 1. f = 0.86 GHz. 2. • X = 34.88 cm (13.73 in.). 3. WEIGHT = 0.45 kg (1 lb). 4. CIRCULAR POLARIZATION. 5. BEAMWIDTH = 100°. 6. MAXIMUM GAIN = 6.5 dB
4.4 cm (1.75 in.)
X/2 = 17.2 cm (6.8 in.)
Fig. Ill-48 Probe Turnstile/Cone Antenna
R2 = spacecraft antenna axial ratio (voltage)
3 = ellipse axis aspect angle as defined in Figure 111-50, deg (0 < S 4 90°).
The equation is plotted in Figure 111-50 along with a definition of the ellipse axis aspect angle (Ref III-ll). As seen from the insert in the figure, the best case is when the ellipses are coincident (& = 0°); the worst case is when the two patterns are orthogonal (£ = 90°).
A circularly polarized antenna pattern may be considered as an elliptical pattern with a certain circularity expressed in terms of the axial ratio. As the probe rotates about its spin axis, the probe antenna pattern will rotate in space with respect to the spacecraft antenna pattern. As seen in Figure 111-47, the axial ratio for the probe antenna is 2.9 dB at the half-power points. An axial ratio of 2 dB was used for the parabolic dish antenna on the spacecraft and £ = 90° for worst-case relative position. As seen from Figure 111-50, the polarization loss is 0.35 dB and this value was used in all RF link calculations.
e. Probe Transmitter - Another attempt was made to survey the industry and determine availability of transmitters in the 0.8 to 1.2 GHz frequency range. Technical information was requested of 19 suppliers and 13 did not respond with data. The six suppliers who did respond provided various technical brochures on hardware that could be modified for the frequency and/or type of modulation required. Philco-Ford Western Development Labs Division, supplied the most technical information. Transmitter efficiency and package density remain the same as determined previously (Vol III, p V-23) since additional technical data indicated the same general values.
/. Spacecraft Receiver - Seventeen suppliers were requested to furnish technical information on hardware that could be modified to the probe mission requirements. Eleven vendors did not have receiver hardware that could be modified to the probe mission, and six suppliers sent general technical brochures that were not particularly applicable to the specific requirements of the receiver with the unique modulation and tone. Therefore, noise figures, weights, and package densities remain unchanged from the values used on the original portion of the study (Vol III p V-23).
111-75
I ON
PO
LAR
IZA
TIO
N
LOS
S,
dB
23
45
67
89
AX
IAL
RA
TIO
O
F T
RA
NS
MIT
TIN
G
AN
TEN
NA
, dB
AX
IAL
RA
TIO
OF
RE
CE
IVIN
G
AN
TEN
NA
,
dB
Fig
. II
I-SO
A
nten
na
Pat
tern
P
olar
izat
ion
Los
s
g. Parametria Designs - RF link analyses were performed first to optimize several critical parameters for the individual missions to Saturn and Uranus using the Task I science payload. This initially sized the beamwidths of the pre-entry and descent antennas and the spacecraft antenna for the optimum trajectory to each planet. Key mission communication parameters are shown in Table 111-21 under Configuration 1. Parameters that were not varied are listed at the bottom of the table. The second configuration considered using the optimum probe descent antenna beamwidth for Saturn for the Uranus mission, and vice versa. As seen in the table, the RF power requirement for Saturn increased 1.2 w.
Early in the analysis, it was evident that the Saturn encounter from the common Saturn/Uranus-1980 trajectory would create the worst-case conditions. (See Table 111-21.) The direct Saturn mission (Saturn 1979) did not create worst-case conditions since the trajectory could be tailored to accommodate the communications geometry required for the Saturn/SU 80 mission. The communication range is also smaller because the direct mission has a periapsis radius of 2.3 Rc as compared to 3.7 R for the Saturn/
b S SU 80 mission. Saturn also has the higher atmosphere microwave absorption of the two planets, as discussed in subparagraph b. Because of these reasons, the direct Saturn mission was not investigated fully until a common probe trajectory had been fin finalized.
Configuration 3 considered two identical axial-beam antennas on the probe that would satisfy the probe aspect angle requirements for both planets. This was achieved by varying trajectory lead times for the two trajectories until a similar probe aspect angle profile versus mission time was established. At the same time, attempts were made to achieve a common spacecraft look direction for probe tracking at both planets. This was achieved by trajectory variations that resulted in a common spacecraft cone angle to cover both planet missions. The clock angles were also close at this time but not as close as the cone angles.
The position of the spacecraft antenna pointing for the common missions was further refined for Configuration 4, as seen in Table 111-21. Minor modifications were made at this time to reduce the cone angle and optimize the spacecraft antenna gain, which further reduces the RF power required. The modifications had more impact on the RF power for Saturn than for Uranus. Configuration 4 results in missions that are common from the communications hardware standpoint, except for the disparity in spacecraft clock angle.
111-77
Tab
le
111-
21
Tel
eaom
mun
iaat
ion
Subs
yste
m
Par
amet
ers
for
Com
mon
Pro
be D
evel
opm
ent
PARAMETER
FIXED LOSSES, dB
PROBE ANTENNA PATTERN TYPE*
PROBE ANTENNA BEAMWIDTH, degt
PROBE ANTENNA BEAM DIRECTION,
deg
SPACECRAFT ENTRY CONE ANGLE,
deg
SPACECRAFT ENTRY CLOCK ANGLE,
deg
MAXIMUM RF TRANSMITTER OUTPUT
POWER, w
1-S**
5.9
T/A
30/90
60/0
105
256
17.4
INVARIENT PARAMETERS
SPACECRAFT ANTENNA BEAMWIDTH = 20°
FREQUENCY =0.86 GHz
BIT RATE = 28 bps (TASK I DESCENT)
E./N„ = 8.9 dB
b
°
-5
BER = 5 x 10
BINARY FSK MODULATION, K = 8
CONVOLUTIONAL ENCODER, M = 2, V = 2
SIGNAL-TO-NOISE RATIO = 10 dB
*PROBE ANTENNA PATTERN AND TYPE
T = TORUS (BUTTERFLY), ANNULAR SLOT
A = AXIAL, TURNSTILE/CONE
1-U
5.9
A/A
90/110
0/0
125
275
2.2
, Q = 8
CONFIGURATION
2-S
5.4
T/A
30/110
60/0
105
256
18.6
2-U
7.1
A/A
90/90
125
275
2.6
3-S
6.3
A/A
100/100
119
257
28
3-U
7.4
A/A
100/100
119
275
5.5
4-S
6.2
116
256
25
4-U
7.3
116
275
5.3
5-U
9.0
116
263
9.2
5-SD
6.2
116
254
18.4
DATA CHANNEL LOSSES = 2 dB
POLARIZATION LOSS = 0.35 dB
SATURN COOL AND URANUS NOMINAL ATMOSPHERE MODELS
USED
FIXED LOSSES INCLUDE CIRCUIT LOSSES, ADVERSE
TOLERANCES, AND 3a VALUES OF SPACE
LOSS,
PROBE POINTING AND SPACECRAFT POINTING LOSSES
SYSTEM NOISE TEMPERATURE
1750°K FOR SATURN
690°K FOR URANUS
tDUAL VALUES ARE FOR PRE-ENTRY/DESCENT TRAJECTORY CONFIGURA
TION
**MISSIONS
S = SATURN/SU 80 MISSION
U = URANUS/SU 80 MISSION
SD = SATURN 79 DIRECT MISSION
Adjustments were made in probe deflection for Uranus to reduce the clock angle and provide for a common spacecraft receiving antenna LOS direction. Configuration 5 in the table shows the results after adjustment was made for Uranus. The spacecraft clock angle was reduced 12° with only a slight change in cone angle. The 116° nominal cone angle position covers the mission satisfactorily. The probe aspect angle increased from 35° to 43° at mission completion. The increased fixed losses result in the RF power increase to 9.2 w required at probe entry, which is the worst-case point for the Uranus mission.
With the communication parameters determined for the common Saturn/ Uranus 1980 mission, the direct Saturn 1979 mission was investi-gated to determine the RF power required. The results are shown in Table 111-21 as Configuration 5-SD. To satisfy the common mission requirements, 18.4 w of RF power is required.
Parametric variations in certain key communication parameters (probe aspect angle, clock angle, cone angle, etc) have resulted in a common probe design with a communication subsystem that will operate for either the S/U 80 or S 79 missions. The Saturn/SU 80 mission (Configuration 4-S) contains worst-case conditions from the standpoint of communication subsystem design. The direct Saturn 1979 mission (Configuration 5-SD) was second and Uranus/ SU 80 mission (Configuration 5-U) was third in order of RF transmitter output power required for the Task I science payload, as seen in Table 111-21. Design details of the Saturn/SU 80 mission are described in Chapter IV, and used in Chapter V, Section D since it is the worst-case mission.
Table 111-22 depicts a comparative summary of the three missions for the follow-on study and the two missions for Saturn (JST 77) and Uranus (JU 79) of the basic study described in Volume II, Section B of Chapters VI and VII. Communication parameters that were sensitive to the communication geometry are shown in the table for the five missions. The new Saturn mission (S/SU 80) has increased range, bit rate, fixed losses, and atmosphere at- . tenuation. The new Uranus mission (U/SU 80) uses a spacecraft antenna with higher gain to offset a smaller probe antenna gain, higher bit rate, and increased fixed losses in the link. The net result is 2.7 w more RF power required for the Uranus encounter from the S/U 80 trajectory.
111-79
Table 111-2,2 Telecommunications Subsystem Charaoteristias Comparison Summary
PARAMETER
ENTRY ANGLE, Y £» cleg
PERIAPSIS RADIUS, Rp
EJECTION RADIUS, 106 km
DEFLECTION MODE
ENTRY COMMUNICATION RANGE, km
PROBE ASPECT ANGLE, *, deg
ACQUISITION
ENTRY
END OF MISSION
SPACECRAFT CONE ANGLE, <j>, deg
ENTRY
END OF MISSION
SPACECRAFT CLOCK ANGLE, e, deg
ENTRY
END OF MISSION
PROBE PRE-ENTRY ANTENNA
ANTENNA TYPE
PATTERN TYPE
BEAMWIDTH, deg
MAXIMUM GAIN, dB
PROBE DESCENT ANTENNA
ANTENNA TYPE
PATTERN TYPE
BEAMWIDTH, deg
MAXIMUM GAIN, dB
SPACECRAFT RECEIVING ANTENNA
ANTENNA TYPE
BEAMWIDTH, deg
MAXIMUM GAIN, dB
FREQUENCY ACQUISITION TIME, sec
RF POWER OUTPUT AT 0.86 GHz, w
BIT RATE, bps
WORST-CASE POINT
WORST-CASE ATMOSPHERE
JST 77
-25
2.3 Rs
10
PROBE
9.6 x 104
55
48
8
91
59
177
173
SLOT
TORUS
30
5.2
TURNSTILE
AXIAL
90
7
HELIX
35
13.5
40
6.5
26
END OF MISSION
NOMINAL
SATURN
S/SU 80
-30
3.8 Rs
30
SPACECRAFT
2 x 105
46
44
24
120
113
257
254
TURNSTILE
AXIAL
100
• 6.5
TURNSTILE
AXIAL
100
6.5
DISH
20
18.3
31
25
28
END OF MISSION
COOL
S 79
-30
2.3 Rs
30
SPACECRAFT
105
45
42
14
122
103
254
250
TURNSTILE
AXIAL
100
6.5
TURNSTILE
AXIAL
100
6.5
DISH
20
18.3
29
18.4
28
END OF MISSION
COOL
URANUS
JU 79
-60
2.4 Ru
10
PROBE
1.5 x 105
19
18
10
159
150
260
257
TURNSTILE
AXIAL
90
7
TURNSTILE
AXIAL
90
7
HELIX
35
13.5
25
6.5
26
ENTRY
NOMINAL
U/SU 80
-60
2.0 Ru
10
SPACECRAFT
8.8 x 104
34
31
43
132
106
263
251
TURNSTILE
AXIAL
100
6.5
TURNSTILE
AXIAL
100
6.5
DISH
20
18.3
18
9.2
28
ENTRY
NOMINAL
111-80
The three major communications geometry parameters that determine a common telecommunications subsystem design are variations in probe aspect angle, spacecraft cone angle, and spacecraft clock angle. Differences in communications range relate directly to the RF power required to overcome greater space loss. These three angles were varied to achieve commonality that results in a common probe pre-entry and descent antenna beamwidth and a probe receiving antenna on the spacecraft that is fixed in look direction as defined by the cone and clock angles.
Probe aspect angles for the three missions are shown in Figure 111-51. End-of-mission conditions for the Saturn/SU 80 mission was used to size the descent antenna beamwidth because maximum gain and minimum beamwidth is required for that mission. The angle is 24° and 3° for 3a variations and 20° to account for parachute swing is added to the nominal value. The antenna should have a half-power (3 dB) point at approximately 50° (24° + 20° + 3°) or a beamwidth of 100". As seen in the figure, the end-of-mission angle for Uranus is 43° and 3a variations are 6°. An optimum beamwidth for Uranus is 140°. The 100° beamwidth for Uranus is down by 6 dB from a maximum of 6.5 dB. The 140° beam-width requirement for Uranus is less severe on a 100° beamwidth than the performance requirements at Saturn. Adverse tolerances have been added to the Uranus RF link design to compensate for the low probe antenna gain. Therefore, two identical probe antennas are used for the common probe design. A circularly polarized axial beam antenna with a maximum gain of 6.5 dB and 100° beamwidth will cover probe aspect angle variations for the three missions and is optimized for the Saturn/SU 80 encounter, which requires maximum antenna gain.
Spacecraft cone and clock angles for the three missions are shown in Figure 111-52. Relative probe movement in the elevation plane is accurately depicted by the cross-cone angle as defined in Volume II, Chapter IV.E.l, pp IV-37 through IV-39. The cross-cone angle has a zero reference at a particular clock angle (Vol II, Figure IV-16, p IV-39), and plots of several missions on the same cross-cone angle versus cone angle would not be realistic because each zero-reference cross-cone point has a different clock angle value. For the three missions of interest, the cone angle is near 115° and the clock and cross-cone plots are nearly identical. Therefore, plots were made of the relative probe movement including 3a dispersions on a cone angle versus clock angle plot as shown in the figure. A common plot was required in order to optimize the position of the probe receiving antenna on the spacecraft, and modify the trajectories so that clock and cone angles were thinin 10° of each other. The design goal was to eliminate spacecraft antenna pointing complexity by having the relative probe positions for the three missions occurring in the same general point in space. Thus, a fixed antenna could be employed on the spacecraft.
111-81
LOOK VECTOR
DESCENT ANTENNA
MAIN BEAM AXIS
SATURN ACQUISITION (E - 7.5)
-PRE-ENTRY
50
40
30
URANUS ACQUISITION (E - 32 minutes)
PROBE ASPECT ANGLE, *, deg
20
10
LEGEND: .t/pii on
S 79
U/SU 80
!
*±3* 3a VARIATIONS
±20* PARACHUTE SWING
-0.6 -0.4 -0.2 0 0.2 0.4 0.6 0.8 TIME FROM ENTRY, hr
Fig. Ill-SI Probe Aspect Angle for the Three Missions
111-82
270,—
M I 00
u>
265
260
CLOCK
ANGL
E,
o, deg
255
250
245 95
LEGE
ND:
• ^
i^
v
A
EOM
• SATURN/SU 80
• SATURN 79
• URANUS/SU 80
NOMINAL
POSITION
END OF MISSION
EOM
j«—
-Q^
k '*
. -
-
ENTRY
100
105
110
115
120
CONE A
NGLE
, «,
deg
125
Fig
. Il
l-52
Sp
aaeo
raft
A
nten
na
Req
uire
men
ts
for
the
Thr
ee
Mis
sion
s
i&p&
130
135
The cone and clock angles are defined in Figure 111-53. These two angles determine a vector direction that represents relative probe movement from a fixed reference system on the spacecraft. The probe vector direction also defines the probe receiving antenna boresight position. The planes shown in the figure contain the Canopus meridian and cone meridian. The cone angle is the angle included by the Earth-to-spacecraft-to-probe alignment; the clock angle is the angle between the Canopus meridan and the cone meridian measured as shown in the figure.
The parametric studies indicated the Saturn encounter from the SU 80 trajectory as having maximum communications range; therefore, the spacecraft antenna gain was maximized in order to minimize the RF power required in a manner similar to the probe aspect angle adjustment. As seen in Figure 111-52, the antenna bore-sight (look direction) was chosen for optimum coverage with minimum beamwidth for the Saturn/US 80 mission. A beamwidth of 20° with a peak gain of 18.3 dB and circular polarization was chosen, which can be met with a parabolic dish antenna. The chosen beam-width will also cover the other two missions, but less efficiently. The end of mission is outside the 3-dB beamwidth for the direct Saturn mission and conditions are worst for the Uranus mission. Entry is 17° off boresight with a gain of 8.6 dB at that point. The gain improves during descent and then is again 14° off boresight at mission completion and moving away rapidly. The gain at the end of mission is 12.1 dB. The Uranus mission certainly is not optimized from the standpoint of antenna gains, but less gain is required since the communications range and atmosphere losses are less.
2. Data Handling and Sequencing (DHS) Subsystem
The basic concepts of the DHS are unchanged from those of the previous volumes of this report. The functions provided are data acquisition from digital and analog science and engineering interfaces, storage, sequencing, formatting, and coding. There are no provisions for data compression or adaptive control with the minor exception of the use of discrete signals from the science instruments and g-switches to provide effective control of data storage and descent format. A functional block diagram of the DHS is shown in Figure 111-54.
111-84
PLANE CONTAINING SPACECRAFT-TO-EARTH & SPACECRAFT-TO-PROBE VECTORS
-ORTHOGONAL PLANE
PLANE CONTAINING SPACECRAFT-TO-EARTH & SPACECRAFT-TO-CANOPUS VECTORS
NOTE: 1. 4 = CONE ANGLE MEASURED FROM EARTH VECTOR TO PROBE VECTOR, 0 < • < 180° 2. e = CLOCK ANGLE MEASURED FROM CANOPUS VECTOR TO PROBE VECTOR,"COUNTERCLOCKWISE LOOKING
AT ANTI-EARTH, 0 < 9 < 360°
Fig. 111-53 Definition of Spaaearaft-to-Probe Look Direction
TIMING
1 USLILLAIUK
CLOCK GATE *
.DATA BUFFER
„ DATA _
SPACECRAFT INTERFACE
CLOCK ' GATE .
BUFFER
COMMAND t
COMMAND ^
" STORED COMMAND PROGRAMMER A
SATURN SEQUENCE (ROM)
•
DATA STORAGE i
* '
COMBINER
1 ENCODER
8-BIT REGISTER
«-
T • URANUS SEQUENCE (ROM)
TIMING
FORMAT GENERATOR
>
1 1
FORMAT (ROM)
TIM
1 - '• . GATES MULTIPLEXER A/D CONV CHANNELS
TRANSMITTER
ING
•
DISCRETE COMMANDS
» DECODE LOGIC SERIAL COMMANDS
PflUFR TNTTTflTF
COAST TIMER SCIENCE
0.1 g
100 g/5 g
— •
SCIENCE & PROBE SUBSYSTEM
Fig. 111-54 Data Handling Subsystem
111-85
a. Probe/Spacecraft Interface Functions - All probe checkout functions are implemented by commands from the spacecraft through the in-flight disconnect to the probe DHS. Although launch and cruise operations have not been defined, it is expected that the DHS and accelerometers will be operating during launch and some minor maintenance functions such as removal of outgassing products from the science instruments, will probably be required during cruise. The capability to operate the probe subsystems individually and in reduced modes is necessary in order to maintain the peak power demand from the spacecraft during cruise and pre-sep-aration checkout within acceptable limits. This capability is provided by the power distribution unit (power and pyrotechnic subsystem) which is controlled by the DHS. Pre-separation checkout must be initiated early enough to allow at least one update
by the ground station. Capability is provided to switch out failed subsystems that are not essential to the operation of the probe in a reduced data mode mission, and it is recommended that pre-separation checkout be initiated early enough to allow several revisions of the probe programming. The present concept of the coast sequencing is implemented by an Accutron timer. Since this component is factory set and not continuously programmable, operation must be initiated at a precise interval prior to probe entry. In order to meet requirements for the two-planet mission, an extra set of output contacts must be provided and will be selected by latching relay before separation. The coast timer is the only active component in the probe subsystem during the coast period, and it contains its own power source. Although basically a sequencing function, the timer is shown as part of the power and pyrotechnic subsystem since its only purpose is to initiate the first pre-entry pyrotechnic event. The only other active component during separation is the spin-up electronics. Spinup is implemented by a simple timer and two pyrotechnic valves. These pyrotechnics will be operated directly off the post-separation battery through relay contacts. An RC timer will provide adequate accuracy and is naturally resistant to EMI produced by the pyro events. Operation will be initiated by the separation event that removes a ground from the timer. The first pyro event occurs 0.5 seconds after separation; the second at 1.26 seconds after separation (0.76 seconds spin-up time).
b. Data Handling Subsystem Design - As stated, the basic concept of the DHS is unchanged fromthe design described in previous volumes of this report, but some modifications have been made to optimize the design for the two-planet mission. Program sequencing is provided by a timer and two ROMs; the correct ROM is selected during the pre-separation period by a latching relay.
111-86
An additional discrete is included in this design to improve descent sequencing. Sensing of the deceleration decreasing to 5.0 g subsequent to 100-g deceleration provides more accurate data for descent sequence control than the 100-g discrete alone.
The memory electronics have been modified from previous bipolar IC design to COS/MOS electronics. Considering the present state of the art, this approach provides a more realistic design. It is anticipated that in the expected time frame of hardware development, all the DHS electronics will be COS/MOS. It is also estimated that a completely redundant DHS would have weight volume, and power characteristics similar to that of the present nonredundant bipolar data processing electronics alone (2.13 kg, 6.9 w, 2330 cm 2). The basis used for sizing the memory consisted of the following assumptions:
1) use of 1024-bit chip (Intel-1103 or TI-TMS4062JC)
2) 30 chips plus 10% of memory for peripheral circuits
3) 5 pw/bit, standby; 25 pw/bit, operating
4) 50% power conditioning efficiency
5) 9.7 cm2/chip, 0.76 cm board thickness, 1.5 cm board spacing
6) 0.91 grm/cm3 (board only)
On this basis the following memory parametrics were generated:
1) volume =443+16.0 M (10-3) cm3
2) weight = 0.2 + 0.715 M (10~5) kg
3) power (standby) = 0.2 + 11.0 M (10~6) w
4) power (operate) = 1.0 + 55.0 M (10~6) w
5) memory = M (number of bits stored)
The DHS begins initial descent sequence shortly after the descent battery is energized. The initiation of the sequence and the initial state of the DHS will be established by a timing circuit working off battery voltage. This will ensure that the turn-on battery transient has subsided, and will eliminate the need of an additional event from the coast times. The required sequence
111-87
of events is illustrated in Table IV-1 (Saturn Sequence of Events, Task I). The pre-entry sequence occurs from item 19 to item 22 inclusive. The actual time with respect to entry of these events will vary (±4.4 min, Saturn; ±29 min, Uranus) due to trajectory uncertainties. The control state established at item 22 will remain fixed regardless of the actual time before entry until 0.1-g sensing (item 24). The functions required of the DHS formats with respect to the Task I Saturn sequence are as follows.
FROM ITEM
19
22
24
30
31
TO ITEM
22
24
30
31
33
FORMAT
STORE ENGINEERING DATA
INTERLEAVE STORED AND REAL-TIME ENGINEERING DATA
STORE ENTRY ACCELEROMETER SCIENCE AND ENGINEERING DATA
STORE DESCENT SCIENCE AND ENGINEERING DATA
INTERLEAVE STORED AND REAL-TIME SCIENCE AND ENGINEERING DATA
During periods 22 to 24 and 31 to 33, the interleaved data is sequenced to the encoder and RF transmitter. The sequence and formatting for Task II follows a similar pattern (Table V-l).
FROM ITEM
19
25
25
27
33
34
TO ITEM
25
27
27
33
34
35
FORMAT
STORE ENGINEERING DATA
INTERLEAVE REAL-TIME PRE-ENTRY SCIENCE AND ENGINEERING DATA WITH STORED ENGINEERING DATA
STORE PRE-ENTRY SCIENCE AND ENGINEERING DATA
STORE ENTRY ACCELEROMETER AND ENGINEERING DATA
STORE DESCENT SCIENCE AND ENGINEERING DATA
INTERLEAVE REAL-TIME DESCENT SCIENCE AND ENGINEERING DATA WITH STORED SCIENCE AND ENGINEERING DATA IN THIS ORDER: (a) ITEM 27 TO 34; (b) ITEM 25 TO 27
111-88
The interleaved data (item 22 to 27 and item 34 and 35) is sequenced through the encoder to the RF transmitter. The retransmission of the pre-entry data (items 34, 35) is an option that provides some data redundancy, however, other options such as increased descent science may eventually prove more favorable.
a. Coding - On these missions, the purpose of coding is to provide a significant reduction in RF power. The selection of con-volutional encoding is recommended because of the simplicity of the circuitry. The constraint length selection will ultimately be influenced by the capabilities of the spacecraft relay link and ground station decoding cost. Long constraint lengths will decrease required probe power but increase the complexity of the decoding equipment. Although the Viterbi decoding algorithm is recommended for this design, longer constraint lengths would favor sequential decoding. The trades between decoding methods would depend on a critical evaluation of the required equipment for each approach with variations in constraint length and code rate. The result would be weighted by the available onboard computational capability of the spacecraft if decoding on the spacecraft is a consideration. Since adequate models for such an evaluation are not defined, this trade is considered beyond the scope of this study. The classic comparisons between con-volutional and block codes are also applicable. The use of block codes would suggest decoding at the ground station due to the cost of possible loss of block synchronization. The increased number of symbols of coded data place an increased burden on the spacecraft relay link and/or data storage. If sufficient spacecraft storage is available or it is considered feasible to relay the coded data in real time, it is recommended that decoding be performed at the ground station. With the code presently recommended, the total number of symbols of the coded data is increased by a factor of eight over the basic data due to the rate 1/2 code and the required quantization. The Viterbi algorithm is recommended for the decoder (constraint length = 8) if decoding is required on the spacecraft. If additional storage capability is required on the spacecraft for probe data, the state of the art of tape recorders must be traded off against COS/MOS approaches. A tape recorder in general represents an overdesign of spacecraft memory with respect to total probe data link storage requirements. It is anticipated that state of the art design could reduce the estimated COS/MOS weight, size, and power by a factor of four. Reliability of either approach represents the greatest uncertainty at this time. Present development of COS/MOS for space applications is advancing rapidly. It is anticipated that a COS/MOS memory could be tested in flight and failed blocks programmed out by spacecraft control. Catastrophic failure of data storage would then be an extremely remote possibility. This is the recommended approach.
111-89
3. Power and Pyrotechnic Subsystem
The configuration of the power and pyrotechnic subsystem is illustrated in Figure 111-55 and is unchanged from the basic con-concepts of previous volumes of this report.
Power is supplied, as required, to the probe from the spacecraft power system during all periods before separation. Remote activated Ag-Zn batteries provide probe power during post-separation activity (4 min) and the entry and descent period (<2 hr). Two Hg-Zn batteries provide power for the coast timer (1.3 v) and the first pre-entry pyrotechnic event (36 v). The power distribution unit responds to commands from the DHS to provide all required power sequencing of the probe subsystems; latching relays provide the switching function. The pyrotechnic subsystem uses capacitor bank discharge to fire entry and descent ordnance devices. SCR switches discharge the banks and latching relays perform arming and safing. The pyrotechnic electronics provide requires power conditioning, ordnance circuitry isolation, and logic required to decode the commands from the DHS. The two pyrotechnic events that occur during spinup derive initiation energy directly from the battery. This avoids the necessity of activating the descent pyrotechnic subsystems and the DHS. A simple RC timer that is highly resistant to pyrotechnic EMI is sufficiently accurate to provide the required discrete signals.
POWER DISTRIBUTION CONTROL
Ag-Zn BATTERY (POST-SEPARATION)
Hg-Zn BATTERY (COAST)
Hg-Zn BATTERY (COAST)
Ag-Zn BATTERY (DESCENT)
SPACECRAFT POV
DHS COMMAND
28 v
1.3 v COAST TIMER
36 v
(INITIAL PYRO EVENT
28 v
JER
)
ii i k
SPINUP ELECTRONICS
PYROTECHNICS
CAPACITOR BANKS
SCR CONTROL
RELAY SAFE/ARM
POWER ELECTRONICS
15 EVENTS 16 EVENTS
TASK I TASK II
Fig. 111-55 Power and Pyrotechnic Subsystem
111-90
a. Power Sources - The post-separation battery energy requirement is composed of power required for the spin-up electronics and two pyrotechnic events. The battery selection was based on available rather than optimum design. Energy requirements for the simple circuitry indicated would not exceed 0.1 w-hr as compared to the indicated capability of approximately 12.0 w-hr of the selected battery. The excess capacity provides capability for limited pyrotechnic firing. Some test verification or cell redesign may be required. The battery characteristic and the number of cells (20) indicate a voltage of 1.4 volts per cell at 1.0 amperes. Assuming a typical linear characteristic and a minimum battery voltage of about 10.0 volts, a maximum current of approximately 10.0 amperes could be developed. The basis of these estimates is as follows:
I . =5.0 amp (minimum bridgewire current) mxn ° R = 2.0 ohms (assumed total bridgewire and circuit resistance)
V . =10.0 volts = 5.0 amp x 2.0 ohms m m
I = 10 amp (minimum battery current for two bridgewires)
V (1.0) = 1.4 volts (28 volts per 20 cells at 1.0 amp) c
V (0.0) = 1.5 volts AV (1.0) = 0.1 volts c c
AV (10.0) = 1.0 volt V (10.0) = 0.5 volt c c
These calculations are based on the preliminary test data for this battery. The cell voltage as a function current density is higher than that indicated from standard sources (NASA SP-172 Batteries for Space Power Systems) but is conservative with respect to more recent data provided by the MMA-D battery area for new Ag-Zn primary cells. Since this battery has very short required life and the circuitry may be designed around the voltage fluctuations, a significant decrease in post-separation system complexity may be achieved by using direct battery operated pyrotechnics.
The weight and volume of Ag-Zn remote activated battery for the entry and descent phase are estimated on the same basis as in previous volumes of this report (Vol III, Appendix G). The required commonality necessitates the use of the higher power at Saturn and the longer trajectory (due to uncertainties) as Uranus.
111-91
H
P
H
J-i
S3
1
(B
<^5
P-
N3
OT P'
rt
(t>
i-h
i-h
fD
o rt
P- < CD
ft
ft) co
P-
00 P •
X) i-i
O V
M
ft) s en
V P
P
ft
w» l-h
O
i-i rt
P4
fD
M
O a rt o rt
P
M
P g S
cr
CD
H
O
Hi <S
< fD
P
rt
CD
•^
s
M
Ul
(D
< fD
P
rt cn
s-'
* p- tn
rt
O
i-i cr
P
p ?f
P
na
-a
H 0 P o cr
pr
P
CO
00 M
fD
(13
rt fD
H
P-
P
P4
fD
H
fD
P
rt
i-t
fD
ft e 3 ft P
3 n v:
s# ft
fD
o i-i
fD
P en
fD
rt O
i-i a4
P P
FT
CD
i-h
i-i O 9 rt
P4
P- cn
fD
< fD
P
rt
P
h-1
O P
fD
• P3
P4
fD
CO
13
X) h->
P-
O
P
rt P-
O P
o I-h
rt cr
fD
o P
ft
XI
M s H
P
O
P- 1
cr
fu
rt
rt ft)
i-i
P-
fD
cn
/-^
Oi a i cr
K
. M
cr
—'
o it cr
•<! cr
i-t
P-
ft
CD
•<;
CD
rt £8
s rt cr
Co
rt
P- 3 o o i-i
XI o I-i
Co
rt fD
ft
O
Co
XJ Co
o P- 1
H
^-v
cr co
fD
X
) X
I P
- M
P
O
P
- X
n
a-P
- S
P
Co
M
rt
fD
X
) M
•<
: ^
i-t
O
4>
rt
O
fD
1 O
Co
P4
3 P
X
) P
- n xi
c
ro
M
< CD
fD
fD
P
^ r
t i-
h S
i-
i O
O
e 3
M
ft
rt cr
l-S
fD
fD
ja
B
C
Co
P-
P-
i-i
P
fD
X)
rt
o
P4
< fD
(T
> i-i
C cn
cr
fD
C cn
o l-h
CO
P
a ft
P- 1
OT
o i-
i ft
O
c p ft •
cr
Co
rt
rt
fD
i-i ^ 1 P-
P
P-
rt P-
Co
rt fD
ft
n cr
Co
I-i
00 fD
CO
P
P- P
H-
rt
P-
P
rt
P-
P
ft oo
ft
fD
i-
i H- < P-
P
ft
00
fD
CD
o fD
p rt
X)
•<:
H o rt
fD
n cr
3 P- o cn
K
o c M
ft
i-i fD
•Q e P-
i-i
fD
ft
P- cn
o M
Co
rt P-
O 3
n e i-i
i-i
fD
P
rt
ft
P-
i-i
fD
n rt
M
*<
l-h
i-i
O S
Co
cr
Co
rt rt
fD
i-i
•<! • H cr
fD
C
cn
(D
o l-h
fD
P
rt
i-i
^ —•
-.
ft
fD
cn
o (D
P
rt
O
H
ft P
Co
3 n CD
ft
fD
< H-
O
fD
cn
Co
H
fD
O
Co
XI CO
O
H
> r
t O
i-
i cr
Co
3 ?r
ft
H- cn 1
tr
• ^ tc ^ o rt- ro
Ci
S" s ^.
Ci
CO
K cr
CO
ce
CO
C+- ro
3 1 H cr
fD
rt K
o o o 3 rt fD
3 ft
H- 3 0
0 Co
XI
XI i-i o CO
o pr
fD
CO
l-h
O
i-i
rt
CO
XI
^<
3 fD
X
) ft
3
fD
ft
< H
-O
OP
CD
M
0
0 M
C
M
B
l-
h •
fD
C
i-i
cn
rt
xi
cr
fD
fD
O
l-{
H-
i-h
cn
H-
rt
fD
C
ft ft
"<!
l-h
O
O
i-i
l-h
rt
M
cr
o H
- P
cn
oo
Co
cn
X4
cr
X)
fD
t-1
\-~
H-
l-h
n I
Co
I-1
rt
H-
H-
l-h
O
(D
P
X)
S!
H
H-
O
H
cr
I-1
I-1
fD
tu
a n
cn
n •
o B
3 H
o
cr
ft
fD
Co
rt
SI
CD
fD
H-
CD
00
H-
cr
rt
rt
cr
fD
i-i
x)
n o
o cn
c
fD
M
cn
ft
o cr
l-
h fD
rt
Co
cr x
) H
- X
I cn
H
-H
-cn
fD
r
t ft
c ft ft
•xi
e »
i-i
H-
rt
p
cr o
o fD
l-
h cn
i-
1
fD
H-
MO
O
fD
cr
O
rt
rt
H-
Co
o o
P
n fD
O
XI
l-h
rt
Co
rt
p
cr
o fD
fD
ce
rt
00
fD
1 cn
C
-j r
t P
H
-P
O
0
0 fD
•
M
M
rt
fD
cn
rt
rt
O
ft
fD
rt
fD
O
rt
M
fD
CO
?r
Co
00 fD
H- 3 ft
C o fD
ft
O c rt 0
0 P cn
CD
H- p O
T CO
i-i
fD
rt
fD n cr
P
H-
hr|
X
I H
" O
CD
i-f
Co
rt
o cr
Co
fD
ft
fD
X
I 3
C
H-
i-i
n i
**
C
fD
cn
rt cr
Co
rt
rt
H-
O 3 P
rt
rt cr
H- cn
rt
H- 3 fD • U
fD
rt P
H-
M
ft)
ft X
1 I-i
p *<
fD
X
p 3 H- P
P
rt
H-
O 3 P 3
o fD
h-1
O
P
!-•
XI
CO
l-h
i-i O 3 rt cr
fD
CD
p 3 fD 3 P
P
e l-h
P n rt c l-(
fD
ft M
O
rt
H-
CD
rt cr
fD
3 o CD
rt
ft
XI
cr
P n ft
< P o e c B
i-i
O B
H- en
H-
P
00 cn
o M
c 1
H B"
P
i-i
^1
n fD
i-1
M
cn
>«•
P n n fD
M
fD
I-i
P
rt fD
ft
rt fD
CD
rt
H-
P
00 s~
\ p r
t cr
H-
00 cr
r
t £S
3 X) CD
H
P
r
t C
f-
i fD
s—
'
O
l-h
£ H-
rt cr
(D
H-
rt cr
(D
i-i n fD
M (-1
ft
fD
cn
H-
OT 3 H- cn
X) i-i
fD
ft
H-
O
rt
H-
3 00 H
* O
3 OT 1 rt
(D
K
B
l-h
P
H-
M
C
i-i fD
l-(
P
rt
fD
cn
n P
rt
fD
rt cr
H- cn
H- cn
P cn
o h-1
< P cr
M
fD
XI n Q cr
M SB
3
X) fD
P
i-
i cn
rt o cr
(D
H- P
rt
fD
l-i 3 P
M n pr
fD 3 H- n P
O
O P cn
H-
ft fD
n p rt H-
O P
l-h
O
i-i
M
O P
00 cn
r
t O
H
P
I-
" O
T
n pr
P
3 H
oo
pr
fD
O cr
< H-
O e CD
3 P
i_i.
o i-i
ft
H-
Hi
l-h
H-
O e M
rt ^
fD
CD
V p H
* rt
cr
o e oo
cr
rt
fD
cn
rt
ft
P
rt
P
H-
P
ft
H- 1
fD
M
H-
l-h
fD
l-h
O
i-i (D
H-
rt pr
fD
i-i
rt ^ XI fD
O
l-h n (D
(-•
M
p XI 1
c 3 ft
fD
l-( cn
rt
C
ft •<:
pr
H- cn
rt o l-f
^ o l-h
P
l-h
XI
o I-i
X) p n fD
B
P pr
fD
i-i
P
XI
XI M
H- n P
rt
H-
O P
CD
o H cr
fD
3 p <
_l
.
o I-i
l-(
fD
h-1
H-
P cr
H-
K1
H-
rt ^
X3 M
H- n p rt
H-
O 3 P
3 ft
M o 3 oo
1 M
H-
Hi
fD
l-i fD
t-1
H-
P cr
H-
H1
H-
rt ><;
X) H
O cr
M ffi
3 cn
P
i-i
fD
(_•
p fD
X
I
3 x)
P
rt
MX
)
s:
H-
cr o
H
- P
O
r
t
cr
H-
o 3
3 fD
»
• fD
ft
M
cn
fD
P
HI
pr
C
P
l-i
OT
rt
fD
pr
fD
O
H
Hi
CD
CD
rt
3
e P
ft
M
i-i (D
cn
fD
P
rt 3 P
3 C
Hi
P n r
t e H
fD
ft
H-
P
P
•<
Ht
)
• O
fD
XI H o
CC
M
XJ
OT
M
1 cn
C
N
3 H
-CD
O
fD
H
-M
P
M
ft
cn
H-
o pr
p
P
rt
< fD
CD
ft
P
P
p cn
fD
P
X
rt
ft
fD
fD
p cn
cn
p
-H
- cro
<
3 fD
X
I H o cr
I
H
H>
P
ft
fD
cn
H-
N
(D
cn
Hi
O
i-i
rt P4
H- cn
M
O C 1
X) 0 5S
fD
l-
i
3 o i-i
fD
i-i
fD
cn
H- cn
rt P
P n fD
rt
O
fD
X
rt
H ffi
3 fD
rt ffi
3 X) (D
i-<
P
r
t C
i-
i fD
fD
P
< F-
H
O 3 3 fD
3 rt
CD
cr
e rt
H- cn
3 O
rt
Hi
fD
O
rt
O
Hi
O cr
fD 3 H- n P
h-1
o cr
P 3 O
T (D
cn
C
H-
rt P*
H
-P
rt cr
fD
cr
P
rt
rt fD
H
•<
H cr
(D
cc
oo 1 n ft
o fD
M
M
H-
CD
H-
ft
P
fD
P
cn
n rt
fD
O
i-i
H
-P
3
00 fD
fD
X
P
XJ
H
fD
fD
n rt
rt
fD
cr
ft fD
cn
P
rt
o o
pr
i-i
H-
P
fD
00
< fD
B
M
fD
H-
3 H
i ft
fD
• O
H
i H
P
cr
fD
t-1
fD
B
p p
CK
(_..
1
O
X)
H
i-i O
l-i
O
fD
Hi
M
H-
cn
P
ID
cr
P
H-
H1
M
H-
p r
t 3
•<;
ft
X3
rt
i-i
cr o
fD
cr
M
fD
fD
H
i 3
1 CD
< cn
cn
o
H-
cr
M
N
O
< fD
s;
fD
ft
P
r
t P
cr
H
- ft
P
Cn
rt
£ cn
fD
r
t
B
H-
cr
P
00
fD
i-
cr
H
rt
K
» •
OT 1
rt
M
pr
CC
3 fD
O
t, cr
C
fD
P
M
<
rt
rt
fD
rt
H-
i-i
fD
B
- i-
i P
<5
rt
cn
fD
F
- cr
3
P
ft o
cn
fD
fD
o p
H-
rt
X)
CD c
r x)
H
- fD
i-
i O
O
3
rt
X
cn
rt
3 cr
p
P
O
M
rt
e fD
(-
'<»
-'
ft
O
^ M
cr
e p
fD
3 fD
U>
cr
o P
P
S
9 cn
3
fD
ft
P
ft ft
s:
< O
(D
P
3
H-
3 0
0 r
t
n pr
P
O
r
t 0
0 3
(D
Hi
H-
p.
3 p
. 1
1 3
CD
C
/-v
cr
g to
S
O
» r
t H
i 0
0 r
t rt
O
H
i P
-O
P
-fD
3 rt CD
cr
fD
M
Hi
M
P-
Hi
fD
1 r
t fD
i-
i
n it
H
i»
ft
H
^ O
T ^
"<
! fD
C
O
P
P
M
3 3
V!
p.
ft
ft
t-j
Hi
^ fD
S
P
• p
l-i
M
CC
fD
p
« 0
0 H
&
o i
pr
c I-
I rN
fD
p
-^
3 i-i
r
t fD
^ cr
s;
3
CC
P
O
fD
H
i 0
0 r
t P
O
H
rt pr
p- cn
P
X3
XI M
P- o p rt h1-
O P
!-d
i-i
fD
M £ p- 3 P
i-i
^ CD
rt e ft
p-
fD
CD
1 r
t x
i rt
N
fD
t-
j cn
3
i-i
p.
w^
3 M
> •
O
O
P-
P-
H
3 xi
ft
M
P
r
t
p-
P
M
cr
O
Hi
(D
PO
O
rt
3 O
B
fD
3
P
P
Cn P
rt
X
r
t fD
p
-rt
p
. p
3
cr
o ft
c p
3 p.
g
rt
3 H
i O
T O
c
r i-i
i-
i O
O
r
t C
r
t 3
V!
p.
cr
x)
cn
M
fD
fD
C
P
CO
P
fD
XI
rt
O
P
(D
rt
rt
i-i
i-i
fD
l-i
fD
P-
K
O
O
p.
O
P
ft
fD
pr
9)
ro
K/-
s 3
OT
P
pr
p-
1 xi
p
o o
xi
< ft
H
fD
O
X
p- 1
< O
P
M
r
t O
(D
cn
CD
(D
CD
r
t cr o
fD
X
) fD
fD
l-
i 3
cu
rt
rt
H
P-
^ 3
--
OT
ft
fD
P
CD
O
H
fD
fD
3 M
r
t P
v:
cr
p r
t r
t O
r
t (D
Hi
i-i
P-
VS
H
•
fD
rt
w
cr
P
fD
O cr o
i-
i
cr
ft P
3
rt
p r
t 3
fD
O
l-i
fD
*<
ft
P-
fD
CD
< P-
l-i
O
fD
fD
•Q C
£ P-
pr
H
p-
fD
O
ft cr
r
t p
o o rt
3 P-
fD
1 fD
r
t
M
fD
P pr
P
00 fD
o e H
i-i
fD
3 rt
Hi o i-i o 3 fD cr
o c K •
o O
,n
pr
H
i c
P
p-H
W
H
O
T U
i fD
fD
ft
ft
p
P
rt
rt
«<!
O
cn
M
ft
fD
^^
fD
rt
pel
cr O
T
fD
1 N
U>
P
ON
1
cr
< P
O
r
t M
r
t r
t fD
i-
i p
W
M
XI
P-
CD
P
P-
rt
rt
< C
fD
r
t i-i
i-
i
S
P
O
-^
M
• fD
a
4 CD
P
CD
3 H
pr
cr
rt
CD
fD
cr
P
O
XI
P
Hi
•<!
i-i
rt
O
O
fD
P
rt
3 ;>
x>
fD
rt
rt P4
P
rt
rt 3- fD
P
O
O 3
P
O
B
o cr
p.
p-
3
o r
t p
. i-
i O
O
O
i-
i 1
cn
cr
P
P
3 p
rt
xi
P
rt
ro
ft
fD
H
i-i
fD
cn "
~<
cn
e XI
XI
Hi
XI
H
O
M
O
l-i
rt
^ <
P o rt
O 3 rt cr
fD
It g"
fD
i-i
p.
co
P
ft
cn
fD
B
CO
CO
P
O
X
O
XI
P"
p.
O
B
P
£•
C
rt
fD
B
fD
l-i
ft
X)
rt
fD
O
H
P-
O
ft
^ fD
l-
i CO
O
r
t p
fD
P
O
fD
pr
p co
P
- (D
O
M
fD
cr
o P
r
t r
t fD
r
t ft
(D
l-
i H
i ^
O
• l-
i
rt
d ^
pr
fD
fD
M
O
• O
<
JJ
P
1 CD
<
rt
O
P->
r
t r
t
g'o
fD
O
l-i
P
CD
cr
rt
P
rt
rt
fD
3 i-i
fD
^
H
p.
fu
cn
3 ft
i-i
fD 1
The subsystem defined for the Saturn/Uranus entry probe uses SCR switches to initiate firing. The requirements and constraints for for this design are:
1) maximum current less than 34.5 amp
2) minimum current greater than 8.0 amp 0.5 msec after current initiation;
3) all fire energy 100 m joules in less than 5.0 msec;
4) bridge wire resistance 0.95 to 1.15 ohms;
5) SCR voltage drop defined by E = V + IR, V = 1.25 volts, R = 0.043 ohms;
6) maximum rate of SCR current 150 amp/ysec;
7) capacitor, tantalum porous plug, C = n82 yf ± 5%;
8) equivalent series resistance (ESR) less than 4.2 ohms/capacitor over temperature range.
For total capacitance less than 1000 yf, 5.0 msec is long compared to the circuit time constant and the energy dissipated is:
E 2
H " 2iT V
R = bridgewire resistance = 1.05 ± 10% ohms
C = total capacitance = n82 ± 5% yf
n = number of 82 yf capacitors
R = series resistance at bridgewire, SCR, relay contacts, ESR (4.2/n) and circuit wiring
E = charge voltage minus SCR drop = 36 ± 10% - 1.25 volts
0.034 n _ . H = 7—^ Joules
1 .35+^1 n
H = 0.146 Joules n = 8
Nominal capacitance/band = 8(82) = 656 yf.
111-93
The current rate of rise constraint is met with the use of a 2.0 yh choke in series with the SCR. This does not have a significant influence on the delivered energy at 5.0 msec. For actual hardware design, the total resistance and inductance of the initiator circuit must be carefully scrutinized; however, this analysis closely approximates values determined in Viking tests.
4. Attitude Control
The attitude control dynamics are minimal for the design configuration considered for the Saturn/Uranus probe. The errors involved are:
1) initial error due to spacecraft pointing;
2) drift error accumulation during period between separation and spinup due to tipoff rate;
3) error due to tipoff rate combined with spinup;
4) error due to spinup jet misalighnment.
Only momentum vector errors are of significance in this analysis since nutation errors are damped out and the spin axis (principal moment of inertia) will align with the momentum vector. The momentum vector errors involved are discussed in the following:
1) Initial error is a function of the spacecraft and will not be discussed.
2) Drift error - Assuming an initial tipoff rate of 0.5 deg/sec about a nominal X-axis, the spin axis is displaced in the minus y direction.
8 = drift error
W = tipoff rate
t = drift period
3) Tipoff rate combined with spinup - This error was analyzed in Vol III, Appendix F of this report. The analysis results in a series solution. The initial (significant) terms are summarized:
111-94
6 (p) = P (o)/P [1 + q2/24 - q^/2688 + q&/506880 . .]
0 (p) = P (o)/P [- q/2 + q3/240 - q5/54569 ]
the position of the spin axis is given by
6 (K) = P (o)/P [q2/6 - qV366 + q6/42260 ] X X
6 (K) = P (o)/P [-q + q3/40 - q6/5456 ]
P = angular momentum
q = P2/mIt
m = torque
1 = transverse moment of inertia
R = spin-axis unit vector
4) Spin jet misalignment - Lower moments of inertia and higher torques enable a simpler analysis of the effect of this misalignment. Since the probe rotates less than 15° for Task I and Task II, the error clue to the spin jets may be based on impulse analysis. From Appendix F, Vol III, the 3o" tolerances involved are as follows. The resulting RSS error applies to Task I and Task II.
Nozzle/flange, cm 0.0254
Flange, cm 0.0762
Cg location, cm 0.038
Mounting surface, deg 0.1
Thrust moment arm, cm 38
Thrust vector, deg 0.1
Number of thrusters 2
RSS torque vector error, deg 0.136
111-95
The time constant of the damper for the previous study was calculated to be twenty hours at 5 rpm. The spin moment of inertia for that calculation was 12.1 kg-m2.* Since the time constant is proportional to this parameter, the time constant for Task I and II is
20 T = -r-fT- I = 2.22 I hr
12.1 s s
T = time constant
I = spin moment of inertia s
It appears that some reduction in size and weight of the nutation damper could be made, but since
T a 7-5- (effect of radius and mass)
m a — (annular ring), K
the decrease in size and weight would be small and was not evaluated. The design constraints for damper with spacecraft deflection mode have not been established. The prime concerns are the effect of nutation on the entry conditions and pre-entry instrumentation. In the absence of more definitive constraints, it has been assumed that the damping time constant should be less than one-tenth of the probe coast time. Analysis of the final battery configuration may show that the stored electrolyte for the remote activated entry/descent battery provides sufficient damping and eliminates all need for a nutation damper.
111-96
E. THERMAL CONTROL SUBSYSTEM
Thermal control is provided as an integral part of the mechanical subsystem design to ensure that all probe systems and individual subsystem components will be maintained within acceptable temperature limits during all phases of the mission. During recent studies (GSFC NAS5-1145 and JPL 953311), various conceptual thermal control approaches for outer planet probe missions were investigated. The results of these studies identified important tradeoffs as well as logical thermal design; these results have been used to arrive at the baseline probe thermal control design presented here. A detailed discussion of the important parameters and thermal design approach and analysis have been previously presented in Volume II, Chapter V, Section A.10, (pp V-79 through V-103).
The most critical thermal control problem from a standpoint of thermal design is the high rate of thermal energy heat loss to the the cryogenic atmospheric environment during descent. To compensate for the cold atmospheres encountered, elevation of the probe internal temperature to maximum before entry allows optimum leeway for internal temperature decrease during descent. Since the probe must be thermally self-sufficient during coast, the probe internal temperature desired for planetary encounter becomes an important consideration.
To satisfy the mission constraints imposed by the anticipated environments, the thermal control concept shown in Figure 111-56 was parametrically evaluated for common Saturn/Uranus probe design. The concept uses:
1) radioisotope heaters located in the service area within the aeroshell to provide heating during spacecraft cruise and coast;
2) multilayer insulation to encapsulate and isolate the probe from the space environment;
3) an environmental spacecraft cover for meteoroid protection during spacecraft cruise;
4) thermal coatings to conserve the internal heating required during probe coast;
5) graphite ablator and fibrous aeroshell insulation for probe entry and to isolate the descent probe capsule from the entry heat pulse;
111-97
LOW-DENSITY INTERNAL FOAM INSULATION
DESCENT PROBE EQUIPMENT
ENVIRONMENTAL GAS BOTTLES
RADIOISOTOPE HEATERS
EXTERIOR MULTILAYER INSULATION
SPACECRAFT ENVIRONMENTAL COVER
INSULATED PRIMARY BATTERIES
CARBON FELT INSULATION
GRAPHITE HEAT SHIELD
Fig. Ill-56 Probe Thermal Control Concept
111-98
6) internal probe low-density foam insulation, located adjacent to the structural probe shell, for thermal isolation of the internal components from the descent atmospheric environment.
To evaluate the thermal control subsystem, worst-case atmospheric encounters and navigation uncertainties were analyzed for the SAG exploratory payload. Figure 111-57 presents the Saturn/Uranus worst-case model atmosphere temperature profiles for descent, using the baseline descent ballistic coefficient of 110 kg/m2
(0.70 slug/ft2). To increase the passive ability of previous probe designs, the probe arrival temperature was optimized to maximum and the more critical primary descent batteries were insulated with 1-cm foam insulation.
Table 111-23 presents typical probe equipment temperature limits. As previously discussed, the probe batteries bound the allowable payload storage limit during spacecraft cruise (300°K upper limit to maintain the dry stored battery charge characteristics). Figure 111-58 presents the cruise/coast probe thermal control radioisotope heating requirement. Since cool-biased probe cruise temperatures are desired for long-term equipment storage and warm-biased probe desired for long-term equipment storage and warm-biased probe temperatures are desired for descent, the probe/spacecraft structural attachments were carefully considered for heat rejection. Results indicated that approximately 8 watts could be dissipated to establish the desired cruise equilibrium temperatures. The hard-mount attachments disconnect at separation and thus provide for probe tempeature increase during coast. On the basis of preliminary probe analysis, an internal probe coast temperature of 302°K with a ±10°K uncertainty would provide optimum leeway for probe temperature decrease and an acceptable margin for possible over-temperature problems at the beginning of descent. Although not included in the probe design for this study, the use of a probe louver system warrants consideration to establish the probe coast temperature and reduce the thermal uncertainty at planetary encounter.
Using the optimum probe entry temperature of 302°K, the SAG exploratory probe was evaluated for a 100% passive thermal control protection system. Late arrival uncertainties were assumed for warm atmospheric encounters where over-temperature conditions would terminate the mission, and early arrival uncertainties were assumed for cool atmospheric encounters where under-temperature conditions would terminate the mission.
111-99
300r
ATMOSPHERIC TEMPERATURE, °K
BE = 110 kg/m2
20 30
DESCENT TIME, minutes 40 50
Fig. Ill-57 Worst-Case Saturn/Uranus Atmospheric Temperature Profi les
25r-
20f
HEATER
REQUIREMENT
w 151—
10 250
oi/eAFT = 0.62/0.13
(IRIDITE 14-2)
FORWARD = ° - 0 4
(ALUMINIZED MYLAR)
,c£o
LEGEND: — —
. . . .
_ _ —
SPACECRAFT CRUISE URANUS COAST SATURN COAST
''* DESIGN (18 w)
325 275 300 PROBE EQUILIBRIUM TEMPERATURE, °K
Fig. 111-58 Cruise/Coast Probe Equilibrium Temperatures Based on Thermal Coating Selection
III-lOO
Table III-23 Typical Pvobe Equipment Tempevatuve Limits
PROBE SUBSYSTEM & SCIENCE COMPONENTS
SCIENCE INSTRUMENTS & ELECTRONICS
NEUTRAL MASS SPECTROMETER TEMPERATURE PROBE TEMPERATURE PROBE ELECTRONICS PRESSURE GAGES ACCELEROMETERS ION RETARDING POTENTIAL ANALYZER NEUTRAL-PARTICLE RETARDING POTENTIAL ANALYZER
LANGMUIR PROBE LANGMUIR PROBE ELECTRONICS NEPHELOMETER
COMMUNICATIONS
RF TRANSMITTER POSTENTRY ANTENNA
DATA HANDLING SYSTEM
ELECTRONICS MEMORY UNIT
POWER & PYROTECHNICS
POWER CONDITIONER POWER DISTRIBUTION BOX PYROTECHNICS (Hg-Zn BATTERY) PRIMARY ENTRY BATTERY PYROTECHNIC ELECTRONICS COAST TIMER
TYPICAL TEMPERATURE LIMITS, °K
STORAGE (MIN/MAX)
235/325 175/700 210/350 220/325 225/375 240/500
240/500 175/700 230/350 240/325
225/350 125/365
220/400 220/400
225/350 225/340 225/325 225/300 225/350 240/350
OPERATING (MIN/MAX)
250/320 175/500 230/325 240/325 230/370 240/500
240/500 175/500 235/350 240/325
250/345 125/365
250/350 250/350
250/340 250/340 255/325 275/325 250/350 280/350
• 1
III-101
S3
o
3 (->
Cu
i-i
fB
,Q e H- I-i
(B
3 tB
O
3 (B
3 <
H-
H
O P 3 (B
P
rt
o o 3 rt
i-i
O
M
Hi
O
i-i
rt
3"
(B
fB
P
It
H-
i-t
rt>
Cu
fB
en
o
ro
3 rt •
H
fD
W
O
nr
fD
CO
P
•3
O
H-
3 rt
O
Hi
Cu
H
-S
H
-3 H
-en
3
* H
-P
OQ
H
tB
rt
C
H
P
** P
P
Cu
P
Cu
(B
en
n fD
P
P
o rt w
(B
P
Cu
(B
»
3 e p
rt
fB
• H
P4
fD
f- en
tB
o Ml
P
cm
P
en
"3
i-i
tB
en
en
3 rt
H-
N
P rt
M-
O
3 en
rt
^
rt
O
OJ
Ui
en
rt
fB
3 rt
P
cr -
3 0)
i-i
en
H-
Cu
M
^!
er
P
H
en
H-
P
rt
P4
fD
a H P 3 C
en
O
O
o M
P
rt S
o en
•3
3*
tB
i-i
fB
W rt
P
4
fB
P
fB
O
P
OQ
P
en
P
•3
*3
i-i
O
P
O
P4 -s
o 3 M
Cu
O
H-
P
rt
fB
Cu
3*
tB
P
rt
H-
3 OQ
n H-
>-i o c H-
o • Ul
->
f^
OQ
• H
P4
H-
en
"3
IB
P
P
M
rt
rt
V4
en
P 3
n o 3 C
u*3
en
s p- rt n 3^
fD
en
• n
j o rt
fD
X
rt
fD
P
Cu
fB
C
u
Cu
fD
en
O
fD
P
rt
rt
O
OJ
Ul
p
i-i
(B
en
s:
H-
rt
P4
N> • O ?T
OQ
O
i-h
cr
P
rt
rt
IB
i-i
H-
fB
en
P P
o o
•
Hi
-P~
O
rt
P4
M
P-
H-
en
rt
fD
en
•-!
rt
en
e Cu
o
^ H
i •
<
O
H
M
P4
C
(B
s fB O
• H
- N
> 3
Ul
C!
* K
OQ
o
o
rt
(D
g 3"
P en
»•
rt
3"
fB
rt
fB
Hi
O
i-i
fD
w er
fD
fD
p
H-
P
rt
fD
a4
HI
OQ
o rt
OQ
rt
P
i-
1 en
fD
en
i-i
-.
<"
"3
^>
H-
C
rt
H-
3*
fi
fB
P
Cu
rt
O
O
O
rt
r>
P
C
H
"3
H-
Cu
fD
ID
en
M
rt
p
P
"3
„ "
=3
»-
i fB
O
H
- X
C
uO
Q
H-
P
en
en
O
P4
3 rt
p
rt
O
fD
l-i
P
rt
fB
Cu
H-
P
rt
O
rt cr
fD
<T
P en
(B
M
H-
P
fD
*3
H
O
H-
3 en
3 h-1
P
rt
H-
O
P ^^
•3
i-i
O
W
(B
en
3*
(B
I-1
M
cr
o 3 p
Cu
p
I-i
•<
• >
M
O
O ^ H
(D
Cu
3 P
C
u
P P
rt
OQ
P
en
en
3 •3
04
T3
fD
Cu
(B
en
H
-H
i (-
.fJ
Q
M
V! en
V
! en
i
T) H o cr
P cr
H-
t-1
H-
rt
^ o Hi
n o P
Cu
fD
P
en
P
rt
H
-O
P
P
3 Cu
Hi
»-{
(B
fD
N
H-
P
OQ
O
rt
C
i-i
tB
en
P
P
rt
H-
O
H-
•3
P
rt
fB
Cu
H-
P
rt
P4
(B
a H P P c en
o o o M
P
rt B
O
en
Hi-
3
3 H-
rt
H
O
OQ
fD
P
OQ
P
en
P
rt
rt P4
fD
O
P4
o en
(B
P
o <
(B
H P
H-
rt
i-i
O
OQ
(B
3 OQ
P
en
cr
fD
O
P
3 en
fB
o Hi o
o
o
M
fD
rt
3"
3 fB
H
fD
P
3 Cu
rt
P4
fD
H-
P
O
H
fD
P
en
(B
Cu
rt
O
fD
M
H-
S
H-
3 P
rt
(B
rt
P4
fD
H
fD
•a
C
H-
H
fB
3 fD
3 rt O
i-h
fD
M
fD
O
rt
K
H-
O
P
M s4
fD
P
rt
H-
3 a
4 O
Q
fD
en
P4
fD
i-1
t-1
rt
(B
0 X) fD
I-i
P 1
• S
fD
O
3 OQ
P cn
s!
P en T
3 11
O
cr
fD
Hi
o i-i
rt
P4
fD
Hi
H-
H en
rt
N3 • Ul
w
p l-(
en
o Mi
Cu
fD
en
n fD
3 rt
£ O c M
Cu
cr
(B
en
3 Hi
Hi
H-
n H-
(B
3 rt
rt
O
i"(
tB
en
3 !-•
rt en
H-
3 Cu
H
-O
P
rt
fD
rt P4
P
rt
P
3 o 3 cr
o P
K
Cu
en
3 X) x>
M
^ o Hi
3 fD
O
3 OQ
P
en
rt o X
I c H
OQ
(D
rt
P4
fB
H
fD
&
C
H-
H
fD
Cu
n o o M
C
i-i
P
3 3 en
P
rt 3 o en
"3
P4
fD
K
fD
O
P
I-1
I-1
cn
Ml
o H
P
Cu
fD
CO
n
(B
3 rt
rt
O
H"
VD
• IS3
cr
P
i-i
cn
>•
3 P
rt
H- <
fD
CO
O P4
fD
3 fD
O
Hi
rt P4
fD
I-i
3 P
H1
>3 H
O
rt
(D
n rt
H-
O
3 C
P
Cn
n o 3 cn
H-
n
o
3 o fD
XI rt en
P en
S
fD
l-J
M
P cn
cn
fD
<
(D
H
fD <
O
M
3 3 tB
P
3 Cu
-3
(D
H-
OQ
P4
rt
Cu
-3
fD
i-i
(B
Cu
• P>
H4
rt
p4 5 3 OQ
P4
rt p4
fD
fD
p
p
H-1
rt
H-
fD
cn
•«
p
p
p
I-1
rt
CD
i-i 1
CO
H-
P
O
fD
fD
M
fB
n rt
rt
H-
O
P !-•
P4
fD
P
rt
H-
P
OQ
O
Hi
rt P4
fD
XI H
O
a4
fD
-3
p •<!
M
O
P
Cu
i-i
(B
•O
C
H-
H
fD
cn
n o 3 -3
(-•
(B
X
P4
fD
P
rt
H-
P
OQ
w
O
i-i
-P- • OJ
c^
?r
OQ
o Hi
a4
P
rt
rt
(B
i-i
H-
(B
CO
•
3 W
CO
O
h-1
O
3 Cu
i--^
Ul
Ul 1 cr
P
i-i
Cu
(B
CD
n fD
p
rt
v
-'
s:
o 3 M
Cu
H
CD
•a c H-
i-i
fD
P
Xi
XI K
O
X
M*
3 p
rt
(B
M <
OJ o o s3 1 P
4
i-i o Ml
P
3 fD
X
rt
fD
3 Cu
fD
a 3 H
-cn
en
H
-o P
rt
O
O
D4
rt
P
H-
P
3 fD
P en
c I-i ft)
y fD
P
rt en
H-
P
rt P4
fD
n o o M
P
rt 3 o cn
Xi P
4 ere
fD
3 o 0 3 3 rt
fD
i-i
o M • O o for
OQ
o Ml
cr
P
rt
rt
fD
i-i
H-
fD
CD
•'
• >
a4
P
rt
rt
(B
i-i
H-
fD
CD
v-^
s* cr
3 rt
h-1
fD
i-i
OQ
^ K o
3 I-1
Cu
CO
P
rt
H-
en
Ml
OJ
1-^
U
l -c
1 P4
H s o 3 M
Cu
cr
fD
»-!
(B
Cu
,0
Cu
H
-rt
H
-0 3 P
H
-1
P
3 P
h-1
^ cn
n>
en
cn
P4
0 K
rt
3 H-
i-i
fD
Cu
Ml
O
M
rt
P4
fD
O o 0 M
P
rt 3 O
CO
pr
-3
P
rt
here
rt
P4
fD
3 o 3 H-
3 P
M
P
rt 3 o CD
Xi P
4
fD
l-(
fD
(B
P
n o 3 P
rt
(B
i-i
<»-\
O >
cn
•3
p4
fD
I-I
fD
10 • P
d
(B
cn
3
P4
(B
P
rt
H» P
0Q
rt
rt
O cn
P
rt
H- en
Mi
P4
^ (D
•3
P
(_>
v<;
rt cn
CD
p4 o S3
rt
P4
P
rt
U>
00 "3
1 P4
rt
O
Hi
fD
h-1
ID
O
rt
M
H-
O
P
h-1
O
P Cu
s p CO
Hi
H-
H en
rt
fD
<
P
i-1
3 P
rt
IB
Cu
Hi
O
H
rt
P4
fD
h-"
3 h-
1
^ cn
3 •3
Ui-
3 U
l
?r
OQ
o Ml
r-1
H-
(B
Cu
IB
3 1
fD
P
P
H4
H-
N
H-
3 0Q
P
rt
P4
fD
3 H-
co
CD
H-
O
3 H
(B
^2 3 H-
i-i ffi
3 IB
3 rt • IT|
o t-t
C3
H
P 3 3 CO
v fD
H1
fD
n rt
i-i
H-
O
P
h-1
M
rt
^ 3 O
1
*i
o
i-i
C/l P
rt
3 i-i
3 fD
3 O
O
3 3 rt
fD
"I
en
** rt P4
fD
XI P cn
CD
H> <
ID
rt P4
fD
I-i
3 P
M
O
O
P
rt
i-i
O
H1
H-
CD
CO
3 Hi
Mi
H-
n p-
IB
P
rt
* O
O m
tn ^ o —1
s m
m
—I
^ l-H
S cr
2 S
i—i
OO
O
O
r-1
O ^ x>
m
/D
cr
i—
i ?
3 m
3:
m
z —I
4=>
-P»
3 3
m
—i
x m
o
S
m
-o
m
m
O
7
3 m
3=
o
—
• c
r 7
3 m
r-
i—i
IS
J—i
—i
cr
-a
-3 m
73 cr
"O
-a
m
73
I-
O s:
m
73
1—
O s:
m
73
f—
o E: rn
73
i-
o s:
m
73
"3 73 m
oo
OO
c
r 7
3 m
<^ cr
Oi
-i cn
• CO
I—1
^J • -p>
ro
cn
• VD
OO
• I—1
cn
ro
CO
lO
—I
i—i
3^ n~;
—
3 _J
. 3 •k
^^
cn
•f*
•vj
-P>
cn
c
n
4i>
cn
co
cn * ro
^j sf-
<^-~
en
i—1
m
CD
-O
O
—
1 s«
az
"3
O
> -n
O
O
oo
o I—
i m
«=
: oo
m
o
-H
s:
3=
73 s:
2:
o s:
l-H
^ J=
I-
o o o r~
s: >
73
3: ^ o IS
INAL o o o r—
oo
3>
—1
cr
73
S
cr
73
3=
Z cr
00
S
en
ft M
W
-d
rf
3
«i
<
» rt
V
(-3
H
'Cf
l>
(D
03
'CD
n>
0 H
I p
* (B
?
P-
>O
C0
Oi-
((
Ui-
t3
i-{
t>
rr
|D
C0
rD
M
3'
nr
fi
tc
n3
rt
i-iH
- en
(D
H
(1
ID
(J
rt
(D
[B
TJ
H-
t)
O
(t
3"
<
3 H
fl
D*
II)
i-h
H
<
H
OX
rt
rt
^ i-i
ID
ID
*<!
O
ID
i-(
(i
r
t»
it
H-
ru
o 3
a*
co
CH
(
(H
i-
if
lc
fw
rt
rt
ur
tg
nn
>r
f,tf
or
De
a>
>
3*
rr i
t g
3*
O
H
• H
- O
n ID
H
i (D
rt
3
ID i
t en
-
'da
- o
oi-
jhj
3"
^ (i
rt
g
tD
fo
gg
HS
H-
It
rt
Cu
p'r
tf
DC
OX
'OH
- C
o O
3
ID
g It
It
3
H-f
l P
U
ft
H
gC
rt
co
gr
tr
jO
P-
'I
Dc
np
* co
rt
tl
IB
O
ID
1)
3 O
rt
H
- ID
O
rt
3
'ftr
t fD
oc
oi-
trto
o
p C
Di-
i O
g
H
H
• Co
3
Ph
3 Q
-co
fo
a
* »-(
H
>
0)
rt
"O
h rt
h
(t
rt
ID ^
01
rt
n
Ot
-t
pih
ii-
tf
D
C
OC
nC
Oi
-3
i-
fO
3C
0O
C0
O
H
Co O
Q
H>
H
g 3
"*S
C
Cr
t p
. O
lD
CT
ft
Ol
tf
ll
D
It
gc
op
->
33
O
*d
CD
ro
fur
t *
0«
It
Pf
fl
tt
S'
rt
»r
t^
''
rj
p'
O
D'i
-i
n>
(t
v^
3*
CD
Ln
T3
fD
3
tfrt
O
3 g
M
ft
rt
i-f
H
fD
CO
rt
OT
fO
OH
It
r/
Od
3 ft
r
Dt
o(
tc
or
D(
ug
p[
uf
ui
-h
rt
sig
n-
- h
<a
,co
i-jn
o
CO
CB
H'1
33
CL
fB
p-
'O-
CO
p'H
H
rt
it
nt
aC
rt
pj
ft
ID
Cu
co
3'
'-
in
>3
i-
iC
33
'C
OO
£;
» hi
O
CD
(B
r
tP
Hf
rP
PH
iP
CD
rt
rt
rt
It
0 It
CO
OQ
rt
p
rt
(i
C
3"3
(IO
rt
rt
3
rt
p*
3 C
L 3
*3
1 It
3
H
3*
O
p*
O-
(l
(t
CL
ID
(l
l|
li
rt
p(
tH
ni
t3
P
•a
cn
ct»
3*
g H
- n
it n
P9
1 3
(t
Pl
td
Pf
lO
II
) O
CO
^v
OX
I rt
)(
«
i ft
p
3 rt
^
ft
rr
cn
OC
oC
oc
nT
d It
fu
ID
ItC
u o
gi
-t
rt
Co
Ph
t-S
3 OQ
c
uro
o o
'd
jj
go
SN
Pi
tf
ft
• 3
3^
0f
)O
Hi
(L
O'
Jr
t»
P'
P-"
3 D
*(B
ID
3
• D
l C
U
rt
^O
lt
CS
lB
IL
i-S
H-r
t 3
* N
3
ft
p'
s h-
1 "x
i 3
ft
rt
rt
It
It
vj
in
(B
n H
-fu
M
3'
PB
'-
'H
0C
3'
Cb
rt
g ID
ID
p
it
p 3
in
it
p 3
Co
C0
S;
SB
nX
(D
Ot
3C
n{
DH
-H
IB
»
P
H
"0
3 3
* rt
O
p-
1 3
it ^
PI
I a
o n
iDfo
p-'
H.
co
<
po
it
oi
it
rt
ti
p'
.j
g 3
*(t
C
C
Dlt
O
(II
rt
C
O
i-
iC
Cf
fn
oC
LO
H
-g
n p
u it
n
it
OG
.M
-S
3 •r
rt
p it
3(
tg
it<
H-
co
s;
^ 3
C~u
rt
«
t)
01
H
P-
ft
rt
HO
PP
I
P->
O
(to
M
O-
H-
3*
0 i
g I
DI
D3
ItO
Q
co
i-h
I I
3 O
- 3
*
PARAMETRIC ANALYSIS SUMMARY
This section summarizes results of analyses conducted during the study and presented in detail in previous sections of this chapter.
Mission Analysis and Design
A comparison of the three design missions is shown in Table 111-24. Saturn direct (S-79) is a terminal mission and is more flexible to compromises than the Saturn/SU 80 mission which flies by Saturn for a subsequent Uranus encounter. The Uranus/SU 80 mission is also a terminal mission with some flexibility to constraints. The periapsis radius of 3.8 R for the Saturn/SU 80 mission causes
a large range of 197 x 103 km at entry, which is significant for the RF link analysis and is the worst case of the three design missions.
An analysis was conducted to obtain the same spacecraft antenna look direction and the same probe pre-entry antenna for both planets. Some modifications to the missions were made so that the cone angles at entry are nearly the same. Further, the Table 111-24 shows that the probe aspect angle at Uranus at the end of the mission is reaching extreme limits and is an important factor in making the missions similar for both planets.
Entry deceleration varies from 225 g in the Uranus warm atmosphere to 837 g in the Uranus cool atmosphere.
Science Analysis
The science analysis is summarized in Table 111-25. The table denotes that the addition of the pre-entry science and nephelometer instruments, used in Task II, increase the weight, power, and volume compared to that for Task I; however, these increases are small compared to the same factors for the overall probe. The science objectives' are achieved with a single descent time of 44 minutes with the Saturn, warm atmospheric model controlling the instrument performance.
III-104
Table 111-24 Comparison of Mission Designs
MISSION
V H p, km/sec
ENTRY ANGLE, deg
PERIAPSIS, R$ OR Ry
DEFLECTION RADIUS, 106 km
COAST TIME, days
PROBE RELEASE ANGLE, deg
SPACECRAFT AV ANGLE, deg
AV MAGNITUDE, m/sec
LEAD ANGLE, deg
LEAD TIME, hr
RANGE AT ENTRY, 105 km
PROBE ASPECT ANGLE AT ENTRY, deg
PROBE ASPECT ANGLE AT EOM, deg
CONE ANGLE AT ENTRY, deg
CONE ANGLE AT EOM, deg
ENTRY TIME DISPERSION, minutes
ENTRY ANGLE DISPERSION, deg
ANGLE-OF-ATTACK DISPERSION, deg
DESIGN MISSIONS
S 79
8.2
-30 2.3
30
39.4
14.7
57.3
52
5.3 1.48
1.08
8.2 (42.4)
13.8
122
102
4.4
1.23
2.1
S/SU 80
9.11
-30 3.8
30 35.8
10.6
59.1
96 3.7
i 2.1 1.97
4.9 (46.7)
20.0
118
112
4.4
1.23
2.1
U/SU 80
11.87
-60 2.0
10 9.5 21.4
46 85 1.4
1.85
0.87
1.9 (31.0)
43 131
105 28.9
5.24
3.0
Table III-25 Saienae Analysis Summary
• PLANETARY ATMOSPHERIC VARIATIONS ALLOW FOR INSTRUMENT COMMONALITY
• EXPANDED PAYLOAD ALLOWS DETERMINATION OF UPPER ATMOSPHERE AND IONSPHERE AND BETTER DEFINES DESCENT CLOUD LOCATIONS FOR THE FOLLOWING INSTRUMENT INCREASES: 79% GREATER WEIGHT
104% GREATER VOLUME (93% INSIDE PROBE) 72% MORE POWER REQUIRED
• A CONSTANT DESCENT TIME OF 44 minutes SATISFIES THE OBJECTIVES
FINAL DESCENT PRESSURE: 3.1 TO 19.2 bars
VERTICAL DISTANCE COVERED ON CHUTE: 92 TO 320 km
DESCENT BALLISTIC COEFFICIENT: 110 kg/m2
• MEASUREMENT PERFORMANCE IS SATISFACTORY FOR WORST-CASE SATURN WARM ATMOSPHERE AND IMPROVES WITH COOLER MODELS AND WITH URANUS
,• THREE MODES OF DATA TRANSMISSION ARE POSSIBLE:
PRE-ENTRY DESCENT
1. EXPLORATORY PAYLOAD ONLY 0 bps 27.7 bps
2. REAL-TIME PRE-ENTRY TRANSMISSION (EXPANDED PAYLOAD) 235 bps 31.5 bps
3. PRE-ENTRY DATA STORAGE AND RETRANSMISSION 235 bps 50.5 bps
OJ
1-H
H
1 I-1 o
D.
O'
fl
ny
Sl
DH
'^
gt
f C
DH
ll
-i
ft
C/
lH
-C
DS
ft
O*
rt-
(B
co
^r
t rjp
,n
p.
n
^ i
c
3"
co
c
n>
n]
(C
H
rt
p4
1-3
H-
<3
0 P
rt
Of
BO
pi
MC
Bf
t O
t
) r
t H
>
S
0)
H
3
rt
p,
o
K
3
•tfc
o
3
^ p
. p
i fi
»
s;
^ on
to
n
M
co
ft
^N
H-
MP
^S
: l
-i
CL
H-
»(
D3
CO
O
ro »
PI
3
SjO
-H
-R
u
rt
M
O
a
(1
pi
»
«)
Pi
pi
pi
i-
if
Di
-h
Pi
i-
fr
t 3
3 (
DP
* rt
it
H
»
Ji
a^
p
ac
«t
) i-
ts
3 C
O
e
n
i-f
i-i
ft
ft
OJ
F-
P*
O
H-
• 0
ft
pi
^-i
•
3
iD
pU
iO
rt
rt
ff
lt
s;
O
rt
p.
0Q
CL
p
ID
p*
to
i-i
i-i
n
n>
H
- o
^
. ft
CD
CD
rt
- 3
3
O
Si
Co
p
i P
F
- P
4
• H
»
Mi
|j
CO
CO N
ft
P
. H
j F
- 0
(D
»
NJ H
- O
3
F-
pd
O
w
3
hi
CW
rt
-f
BU
JC
uH
rf
Ui
CO
3
rt
ff
ff
H
>-3
rt
- a
T
J p
i It
It
tf
PJ
fi
P4
F-
O
>3
O4
Co
(
Or
tS
lS
ig
nf
ri
Cl
lH
X'
ID
It
O
III
^
• fD
co
C
L H
- H
a
. co
H
rt
p
0
9
ID
C
fl
M
H
O
l-i
P4
i-(
MO
MM
i
-i
H-
rt
fD
r
or
ts
; n
H
1
pi
3
• H
Q
F-
fD
1
00
OP
3
o
3
H
rt
ro
ft
f CD
F
- P
4 0
9
p4
-J
ft
• C
L d
H
PJ
I-I
fD
• CO
ID
P
i It
3
TJ
It
C
(0
rt
gO
Q
H^
3
4
H'
Hn
pi
fD
ID
OC
C/
iH
rt
(BID
3
CO
a4
F-
!*!
P4
CO
CO
3
S
rt
ID
i-i
fD
7?
rt
O
rt
ID
S
O
»
ar
tD
'S
gO
fl
Hi
M
ft
O
ft
n>
ft
&>
H
M
e H
- 3
C
fD
rt
(0
u
J>
3
W
rt
H
1
p"
p.
F-
Ma
s'
PI
ft
ft
It
3
(0
a
rt
s;
rt
3
>-h
OP
rt
Sj
C
pi
F-
rt
H
p.
OP
iH
Sj
IO
Of
lU
0
4^
H
rt
Pi
Pi
3
^
CO
p.
O
It
rt
rt
01
H
FT
4
rt
co
co
F-
H
C
»
p.
s;
»
OM
a>
co
oM
O
p
i P
I S
t-
1
It
H'
H
3 rt
H
(1
CT\
£
P*
ft
rt
H
3
Mi
Pi
3
P*
3 01
p.
i-i
?r
rt
p.
O
CD
CO
O
W
rt
C
It
rt
ft
S
CO
t-j
pi
CO
1
cr § rs
oo
> —'
cz
sr
^o
cr
sr
rs
oT
S
3=
oo
oo
3=
>o
or
^ 7
3
2
O
73
3
O
2
>-i
r-
:c=
2
•-<
r—
:>
^ 3>
r-
—4
2 O
oo
-a
re
m
73
m
00
z
3>
r—
—1
2 o
oo
-a
re
m
73
m
oo
o
o
o
o
o
o
• •
• •
• 0
0 *v
l e
n
CO
"-J
o
n
-P*
r\>
1
0
co
o
en
»—>
-F^
-P>
-P
»
-f*
en
o
to
^i
i—•
-P>
o
n
ro
IS
3=>
O
re
zz.
o
• -u
^-
73
3
m
cr
oo
•—
oo
cr
73
m
^ a cy
*-»
TO
h
h
H
CO
O
i
ti
ft
hS
ft
Ci
<rf
TO
to
TO
T
3
w
O
^ 3 TO
3 <rf
>3
TO
03
£ ^ rt
co
S
&i
o
3"
3 C 0 &
ft
H
(B
3 P.
•3
l-i
ft
CO
CO e H ft
fu
M
rt
F-
rt e P- (D
l-h
O
i-i
ft
(B
O
p4
(B
rt 0 O
CD
"3
P4
rt
i-i
ft •
/~s
S
CD
P-
cr"
ri
ft
i-i
O
pi
Pi
i-i
3
H-
it
a
3
pi
rt
CO
p*
H-
H-
s;
3 K
P
i 0
9 p
4 C
O
>-'
ft
H
P
-"3
•
ft
(_•
>3
C
K.
I-'
CO
>
O
H"
"S
U
l CB
ft
3
CO
i-i
rt
ft
ft
n
co
o
O
C
l-h
3
M
p.
rt
rt
CO »
p4
ft
rt
P4
O
Hi
t f
f P
i P
c
rfl
rt
M
P
i ft
ft
i-
i P
i (B
M
o
r
t M
p
4
I-I
e
"3
1 rt
i-i
Kl
ft
ft
ON
CO
F
- CO
CD
CO
C
p
4 I"
! o
a
it
s4 rt
01
ti
p
M
M
rt
O
rt
S1
^ H
' ft
ft
r
t P
- C
H
p
. ft
p
i ft
I-
1 rt
CO
CB
rt
t_n
pi
ft
rt
P.
OP
h
-1
O
O
CD
p.
P.
O
ft
e
e
ft
o
• c
r n
o
o
U
rt
It
H
fD
fD
M
^Cl
ri
a
it
pi
R
rt
i-i
rt
hs
fD
rt
pi
B
ft
P.
O
rt
O
Ci
H
CO
S1
rt
P.
O
T3
K
O
fD
3
3*
rt
rt
fD T
O
S
ft
rt
i-i
CB
l-i
O
H-
t3
O
0
O T
O
P4
H-
<
tS
3
CB
0
w
3 ft
H
O
O
3
><!
CL
CC
3
-3
ft
3
CT
CD
t-h
M
TO
ft
i-
i i-
i CO
3
H
pi
o
d-
co
n
3
i-h
P4
O
^ 3
C
M
i-i
S
O
rt
M
ft
It
Ol
CO
M
0Q
p
i ts
; i-
i p
, O
P
rt
co
fD
ft
C
^.
pi
T3
rt
M
oj
rt
M
O
3
ft
O
1
ti
^
CO
pi
g
oo
3
en
rt
ft
^J
P.
H-
P4
P
3
pi
rt
OQ
d
n
3
»
i-i
ft
CO
CB
O
3
pi
3
rt
• O
3
C
P4
>i)
p4
CO
ft
Pi
rt
3
O
3
p*
pi
pi
pi
>
pi
I-"
3
i-i
rt
v<j
co
g
CD
CD
ft
»
It
rt
|*
O
p4
CO
rt
3
O
ft
p4
O
3
*3
ft
B
p,
o
p
p.
CB
p'
CO
fD
3
H
3
3
Pi
^ rt
O
rt
h-1
ft
O
i-i
»
O
3
><J
O
p.
I CD
3
CD
p
1 P
.
cn
v; CO
rt
fD
0 a ft CO
H-
OP
3 > 3 P
i M
^ CO
H-
CO
Table I11-27 Task II Data Mode Analysis Results
DATA MODE
1. PRE-ENTRY DATA
TRANSMITTED REAL-TIME (NO STORAGE)
DESCENT
2. PRE-ENTRY
TRANSMITTED REAL-TIME (NO STORAGE)
DESCENT
3. PRE-ENTRY DATA
STORED & TRANSMITTED DURING DESCENT
DATA RATE, bps
235
31.5
235
31.5
50.5
MODULATION TYPE
FSK
FSK
PSK
FSK
FSK
RF POWER, w 116 \
29 )
"I 29 ) 40
PROBE WEIGHT, kg
116
103
103
PROBE DIAMETER, cm
96.5
91.4
91.4
PROBE DATA STORAGE, 103 bits
11.9
11.9
60.2
c, descent Depth Sensitivity - Figure 111-59 shows the probe system definition depth at the end of the mission for all three atmospheric models at Saturn and Uranus . For Saturn, the communication subsystem is sized for the cool atmosphere primarily because of the attenuation; the nominal and warm atmospheres, however, are more critical to the link geometry. For Uranus, the communication geometry is critical because the axis of rotation for Uranus is approximately in the ecliptic plane.
Thermal control, sized by the Uranus cool atmosphere, shows that for the design mission a depth capability of 20 bars in the Uranus warm atmosphere is feasible using a nonpassive system. The atmospheres at Saturn are much less severe so that a passive system could be used for that planet.
4. Telecommunication Analysis
Table 111-28 presents the communications analysis summary to achieve common spacecraft antenna pointing and a common probe pre-entry antenna for use at both planets, as mentioned in paragraph III.F.l for the mission impact. The table shows that when tailored for each planet, the RF power required was 17 watts for
III-107
Table 111-28 Communications Parameters
PARAMETER
PROBE ANTENNA PATTERN TYPE
PROBE ANTENNA BEAMWIDTH, deg
SPACECRAFT ANTENNA BEAMWIDTH, deg
SPACECRAFT-TO-PROBE NOMINAL CONE ANGLE, deg
SPACECRAFT-TO-PROBE ENTRY CLOCK ANGLE, deg
MAXIMUM TRANSMITTER POWER, w
S/SU 80
TORUS/AXIAL
30/90
20
105
256
17.4
U/SU 80
AXIAL/AXIAL
90/110
20
125
275
2.2
S/SU 80
AXIAL/AXIAL
100/100
20
116
256
25
U/SU 80
AXIAL/AXIAL
100/100
20
116
263
9.2
S 79
AXIAL/AXIAL
100/100
20
116
256
18.4
0.01
Pressure, bars
0.01
Nominal Coot Warm Nominal Cool
Legend:
— —= — Pressure Depth at End of 44.0 min Descent —••—••= Communication Link Geometry Limit, 0 dB Margin
Communication Atmospheric Attenuation Limit, 0 dB Margin = — - = - - Limit for Passive Thermal Control System -=—•—Thermal Control Limit for Design Mission, Nonpassive
Fig. III-S9 Cloud Locations and End-of-Mission Limits for Various Model Atmospheres
III-108
Saturn and 2 watts for Uranus. When the mission definition was revised for the desired commonality, the RF power requirement was increased to 25 watts at Saturn. This definition is reflected in Chapter IV (Task I) data.
REFERENCES
III-l The Planet Saturn (1970), NASA Space Vehicle design Criteria (Environment). NASA SP-8091, June 1972.
III-2 The Planets Uranus, Neptune and Pluto (1970), NASA Space Vehicle Design Criteria (Environment). NASA SP-8103, Nov.. 1972.
III-3 S. G. Chapin, Program Manager: Jupiter Turbopause Probe Gas Physics -Environment and Instrument Response Study, Final Report. Martin Marietta Corporation Report No. MCR-71-142, August 1971.
III-4 R. S. Wiltshire, Program Manager: Systems Level Study of a Non-survivable Jupiter Turbopause Probe, Final Report. Martin Marietta Report No. MCR-72-91, June, 1972.
III-5 D. M. Hunten: "The Upper Atmosphere of Jupiter", Journal of Atmospheric Sciences, Vol 26, No. 5, Part 2, September 1969.
III-6 J. D. Krans: Radio Astronomy. McGraw-Hill, New York, New York, 1966, p 237.
III-7 S. J. Ducsai: Jupiter Atmospheric Entry Mission Study, Final Report. Martin Marietta Corporation, Denver, Colorado, MCR-71-1 (Vol III) Contract JPL 952811, April 1971, pp IV-82 through IV-158.
III-8 D. A. DeWolf and G. S. Kaplan: Investigation of Line-of-Sight Propagation in Dense Atmosphere: Phase III, Part I, Final Report. RCS Labs, Princeton, New Jersey, NASA CR-114416, NASA/ARC Contract NAS2-5310, November 1971, pp 5 through 11.
III-9 J. S. Hogan, S. I. Rasool, and T. Encrenaz: "The Thermal Structure of the Jovian Atmosphere." Journal of the Atmosphere Sciences, Vol 26, No. 5, Pt 1, September 1969, pp 898 through 905.
111-10 J. D. Osborn: An Investigation of Low Gain Antennas for Spacecraft and Planetary Probes. Martin Marietta Corporation, Denver, Colorado, Research Report, D72-48728-001, December 1972.
D. G. Fordham, Jr. and R. S. Brazil: "Polarization Loss for El-XH-11 liptically Polarized Antennas." Microwave Journal, Vol 8, No. 12,
December 1969, pp 50 through 52.
III-109
SATURN/URANUS PROBE SYSTEM DEFINITION WITH SCIENCE ADVISORY GROUP'S EXPLORATORY PAYLOAD
This section summarizes the common Saturn/Uranus probe system definition (Task 1) which uses the Science Advisory Group's (SAGs) exploratory payload and was defined in Volume II, Chapter III. The definition uses concepts that were generated during the basic study and extended to be compatible with the three new missions, the "worst case" atmospheric models, and spacecraft deflection mode. Results of the system and subsystem tradeoffs and analysis from Chapter III are included in the definition.
MISSION DEFINITION
1979 Saturn Mission Definition
The Saturn Direct 1979 mission is shown in Figure IV-1 and detailed in Table IV-1. Important mission design results are sum-
i marized in this Chapter.
a. Interplanetary Trajectory Selection - The type I Interplanetary trajectory from Earth to Saturn is shown in Figure IV-la with 1-year intervals noted. The launch date of November 23, 1979 and a Saturn arrival date of July 19, 1983 results in a total trip time of 3.7 years.
b. Launch Analysis - The launch analysis is provided in Figure IV-lb. Available payload is plotted against the launch period for four sets of launch vehicle performance data. The nominal launch window and parking orbit coast times were checked and found satisfactory.
c. Approach Trajectories - The probe trajectory was constrained to enter and remain in the sunlit side of the planet until the end of the mission (E + 44 min). This requirement set the entry angle at y = -30°. The resulting approach trajectory is pictured in Figure IV-Id and summarized in Table IV-lb. Since Saturn is the terminal planet for this mission, a spacecraft periapsis radius was selected to be 2.3 R to miss encountering the rings.
IV-1
ARRIVAL
AT S
ATURN,
JUL
19, 1983
ORBIT
OF SATURN
PAYLOAD
WEIG
HT 700-
81 L 80
14.7°
57.3°
72°
146°
''
'
\ 146°
SPACECRAFT ON
PROBE
EARTH
LOCK
RELEASE
HP
EARTH
AT SATURN
ARRIVAL
LAUNCH
FROM EARTH,
NOV
23, 1979
a)
INTERPLANETARY
TRAJECTORY
SHUTTLE/
AGENA
SHUT
TLE/
CENTAUR
TITAN
IIIE/CENTAUR/TE364-4
TITAN
IIIE/CENTAUR/BURNER II
0 5
10
15
LAUNCH PE
RIOD
, days
c) LAUNCH ANALYSIS
AV =
52.5 m/sec
SPACECRAFT
AV
RETURN
TO
EARTH
LOCK
b) DEFLECTION MANEUVER
ROTATION
POLE
TERMINATOR
PERIAPSIS
= 2.3 Rs
ENTRY
ANGLE = -30°
LATITUDE
= 21.7°N
d) PLANETARY
ENCOUNTER
Fig
. IV
-l
1979
Sat
urn
Mie
sion
D
efin
itio
n
Table IV-1 Saturn 1979 Mission Summary Data a) CONIC TRAJECTORY DATA
INTERPLANETARY TRAJECTORY
LAUNCH DATE: 11/23/79
ARRIVAL DATE: 7/19/83
FLIGHT TIME: 3.7 yr
LAUNCH TRAJECTORY
NOMINAL C3:
NOMINAL DLA:
C3 (5-day):
C3 (10-day):
C3 (15-day):
132 km2/sec2
34.5°
133.9 km2/sec2
135.8 km2/sec2
140 km2/sec2
ARRIVAL TRAJECTORY
VHP.
RA:
DEC
ZAE
ZAP
INC
V
8.2 km/sec
173.88°
-0.163°
146.0°
140.1°
3.97°
2.3 Rs
b) DEFLECTION MANEUVER AND PROBE CONIC
DEFLECTION MANEUVER
DEFLECTION MODE:
DEFLECTION RADIUS:
COAST TIME:
AV:
APPLICATION ANGLE:
OUT-OF-PLANE ANGLE:
PROBE RELEASE ANGLE:
SPACECRAFT DEFLECTION ANGLE:
SPACECRAFT REORIENTATION ANGLE:
*SUBSOLAR ORBITAL PLANE.
SPACECRAFT
30 x 106 km
39.4 days
52.5 m/sec
74.4°
10.9°
14.7°
57.3°
72°
PROBE CONIC DEFINITION
ENTRY ANGLE:
ENTRY LATITUDE*:
ENTRY LONGITUDE*:
LEAD TIME:
LEAD ANGLE:
PROBE/SPACECRAFT RANGE (ENTRY):
PROBE ASPECT ANGLE (ENTRY - ) :
PROBE ASPECT ANGLE (ENTRY +):
PROBE ASPECT ANGLE (EOM):
-30°
-4.4°
59.9°
1.48 hr
5.3°
1.08 x 105
42.4°
8.2°
13.8°
km
c) DISPERSION ANALYSIS SUMMARY
NAVIGATION
TRACKING:
SMAA:
SMIA:
6: TOF:
UNCERTAINTIES
DOPPLER & RANGE (40-DAY ARC)
2009 km
833 km
93° 87.7 sec
EXECUTION ERRORS (3a)
AV PROPORTIONALITY:
AV POINTING:
PROBE ORIENTATION POINTING:
1%
2°
2°
DISPERSIONS (3a)
ENTRY ANGLE:
ANGLE OF ATTACK:
DOWNRANGE:
CROSSRANGE:
LEAD ANGLE:
LEAD TIME:
ENTRY TIME:
1.1°
1.8°
2.02°
1.47°
2.4°
5.7 minutes
4.9 minutes
d) ENTRY AND DESCENT SUMMARY
ENTRY PARAMETERS
ENTRY VELOCITY, km/sec
ENTRY ALTITUDE, km
ENTRY B, kg/m2
MAXIMUM DECELERATION, g
MAXIMUM DYNAMIC PRESSURE, N/m2
EOM PRESSURE, bar
DESCENT B, kg/m2
CRITICAL ENTRY EVENTS
g = 0.1
g = 100
MAXIMUM DECELERATION
MACH = 0.8
MACH = 0.7
ATMOSPHERE MODEL
WARM
36.3
968.5
102
227
2.2 x 105
4.0
no ATMOSPHERE MODEL
WARM
TIME, sec
12.2
35.0
40
89.5
98.5
ALTITUDE, km
734
329
259
166
162
NOMINAL
36.4
536
102
366
3.6 x 105
8.0
110
COOL
36.5
297
102
568
5.6 x
18.2
110
NOMINAL
TIME, sec
5.0
17.5
22
54.5
59.5
ALTITUDE, km
445
218
151.2
94.5
92.5
COOL
TIME, sec
1.5
8.5
11.5
33.5
36.5
LO5
ALTITUDE, km
268
139.5
92.7
51.5
50.5
IV
The probe aspect angle at start of descent Is 8.2° and increases to 14° at the end of the mission. The spacecraft periapsis radius is such that approximate angular rate matching is achieved, which accounts for the small change in probe aspect angle for the duration of descent.
d. Deflection Maneuver - A spacecraft deflection maneuver is used in this mission at a deflection radius of 30 million km, or 39.4 days from entry. A deflection AV of 52.5 m/sec is applied to the spacecraft to establish the desired communication geometry. The implementation sequence is shown in Figure IV-lc.
e. Navigation and Dispersions - The navigation and dispersion results are given in Table IV-lc. A forty-day standard Doppler range tracking arc was used. The entry time dispersion was 4.4 min (3a). The communication parameter dispersions are discussed in Chapter IIIS Section A.
/. Entry and Descent Trajectories - Table IV-Id summarizes the entry and descent phases of the mission. For the nominal atmosphere, the entry phase starts 536 km above the 1 atmosphere pressure level and ends 59 seconds later. The descent phase starts after the aeroshell is staged and lasts until the end of the mission, 44 minutes later. Parametric data showing critical events times and altitudes for the warm, nominal, and cool atmospheres is also shown in Table IV-Id. Maximum deceleration of 568 g is experienced in the cool atmosphere.
2. Saturn/SU 80 Mission Definition
The Saturn/SU 80 mission is shown in Figure IV-2 and detailed in Table IV-2. The important mission design results are summarized in this section.
a. Interplanetary Trajectory Selection - The type I interplanetary trajectory from Earth to Saturn is pictured in Figure IV-2a. The launch date is May 9, 1984. The total trip time is 3.4 years.
b. Launch Analysis - Available payload weight is plotted against launch period for four sets of launch vehicle performance data in Figure IV-2b. The nominal launch window and parking orbit coast times were checked and found satisfactory.
c. Approach Trajectories - The probe trajectory was constrained to enter and remain in the sunlit side of the planet until the end of the mission (E + 44 min). An entry angle of y = -30° was found satisfactory. The resulting approach trajectory is pictured
IV-4
ENCOUNTER
SATURN,
82
81^,. ORBIT OF
MAY
9, 1
984
83^^
""
SATURN
84.
#1
LAUNCH
FROM E
ARTH,
DEC
3, 1980
a)
INTERPLANETARY TRAJECTORY
kg lb
PAYLOAD
WEIGHT 90
0
800
I 700-
600-
500
400 J
2100
1900
1700
1500
1300
1100
900
SATURN 1980
SHUTTLE/
AGENA
SHUTTLE/
CENTAUR
TITAN
IIIE/CENTAUR/TE
364-4
TITAN
IIIE/CENTAUR/BURNER II
0 5
10 15
LAUNCH P
ERIOD, days
c) LAUNCH A
NALYSIS
Fig
. IV
-2
Satu
rn/S
U
80 M
issi
on
Def
init
ion
SPACECRAFT
ON E
ARTH
LOCK
10.6
59.1
" 69.7°
•*?
---f
--"%
"-X
-142.7°
-7
«.„„•:.•„}„„ „
»
"HP
PROBE
RELEASE
b) DEFLECTION M
ANEUVER
y = -45
SPACECRAFT AV
RETURN TO
\ EARTH
LOCK
AV =
96 m/sec
PROBE
ACQUISITION
(E -
7.5 m
inut
es)
PERIAPSIS
END
OF
E MISSION
(ENTRY)
(E + 44 mi
nute
s)
PROBE
TRAJECTORY
SPACECRAFT
TRAJECTORY
PERIAPSIS
E + 2.27 hr
ENTRY, E
d) PLANETARY
ENCOUNTER
E + 44 minutes
Table IV-2 Saturn/SU 1980 Mission Summary Data
a) CONIC TRAJECTORY DATA
INTERPLANETARY TRAJECTORY
LAUNCH DATE: 12/3/80
ARRIVAL DATE: 5/9/84
FLIGHT TIME: 3.4 yr
LAUNCH TRAJECTORY
NOMINAL C3:
NOMINAL DLA:
C3 (5-day):
C3 (10-day):
C3 (15-day):
132.7 km2/sec2
27.5°
134.5 km2/sec2
135 km2/sec2
140 km2/sec2
ARRIVAL TRAJECTORY
VHP RA:
DEC ZAE
ZAP
INC
9.11 km/sec
188.27°
0.192°
142.7°
145.9°
2.8°
b) DEFLECTION MANEUVER AND PROBE CONIC
DEFLECTION MANEUVER
DEFLECTION MODE:
DEFLECTION RADIUS:
COAST TIME:
4V: APPLICATION ANGLE:
OUT-OF-PLANE ANGLE:
PROBE RELEASE ANGLE:
SPACECRAFT DEFLECTION ANGLE:
SPACECRAFT REORIENTATION ANGLE:
*SUBS0LAR ORBITAL PLANE.
SPACECRAFT
30 x 106 km
35.8 days
95.9 m/sec
73.0°
13.9°
10.6°
59.1°
69.7°
PROBE CONIC DEFINITION
ENTRY ANGLE:
ENTRY LATITUDE*:
ENTRY LONGITUDE*:
LEAD TIME:
LEAD ANGLE:
PROBE/SPACECRAFT RANGE (ENTRY)
PROBE ASPECT ANGLE (ENTRY - ) :
PROBE ASPECT ANGLE (ENTRY +):
PROBE ASPECT ANGLE (EOM):
-30°
-4.969°
63.3°
2.1 hr
3.8°
1.97 x 105
46.8°
4.9°
20.7°
km
c ) DISPERSION ANALYSIS SUKMARY
NAVIGATION
TRACKING:
SMAA:
SMIA:
6: TOF:
UNCERTAINTIES
DOPPLER & RANGE (40-DAY ARC)
1950 km
833 km
93° 85.5 sec
EXECUTION ERRORS (3o)
AV PROPORTIONALITY:
4V POINTING:
PROBE ORIENTATION POINTING:
U
2° 2°
DISPERSIONS (3o)
ENTRY ANGLE:
ANGLE OF ATTACK:
DOWNRANGE:
CROSSRANGE:
LEAD ANGLE:
LEAD TIME:
ENTRY TIME:
1.2°
1.86°
2.2°
1.58°
1.9°
8.7 minutes
4.6 minutes
d) ENTRY AND DESCENT SUMMARY
ENTRY PARAMETERS
ENTRY VELOCITY, km/sec
ENTRY ALTITUDE, km
ENTRY B, kg/m2
MAXIMUM DECELERATION, g
MAXIMUM DYNAMIC PRESSURE, N/m2
EOM PRESSURE, bar
DESCENT B, kg/m2
CRITICAL ENTRY EVENTS
g = 0.1
g = 100
MAXIMUM DECELERATION
MACH = 0.8
MACH = 0.7
ATMOSPHERE MODEL
WARM
36.8
968.5
102 227
2.2 x 105
4.0 110
ATMOSPHERE MODEL
WARM
TIME, sec
12.2
35.0
40.0
89.5
98.5
ALTITUDE, km
734 329 259 166 162
NOMINAL
36.8
536 102 366 3.6 x 105
8.0 110
COOL
36.8
297 102 568 5.6 X 105
18.2
110
NOMINAL
TIME, sec
5.0 17.5
22.0
545 59.5
ALTITUDE, km
445 218 151.2
94.5
92.5
COOL
TIME, sec
1.5 8.5 11.5
33.5
36.5
ALTITUDE, km
268 139 92.7
51.5
50.5
IV-6
in Figure IV-2d and summarized in Table IV-2b. The spacecraft periapsis radius of R = 3.8 R was selected as a compromise between launch energy and communication link power requirements, as discussed in Chapter III, Section A.3. The probe aspect angle at the beginning of descent is 5° and increases to 21° at the end of the mission. Approach geometry for all three missions was selected to achieve a commonality in spacecraft-to-probe directions. As a result, the probe aspect angle is not minimized, but rather is simply kept within acceptable bounds. The probe leads the spacecraft for the duration of descent.
d. Deflection Maneuver - A spacecraft deflection maneuver requiring 96 m/sec is used in this mission at a deflection radius of 30 million km. The corresponding coast time is 36 days. The implementation sequence is shown in Figure IV-2.
e. Navigation and Dispersions - The navigation and dispersion results are given in Table IV-2c. A standard Doppler tracking arc of forty days with supplementary ranging measurements was used. The three-sigma entry time dispersion was 4.4 minutes. The communication parameter dispersions are discussed in Chapter III, Section B.
f. Entry and Descent Trajectories - Table IV-2d summarizes the entry and descent phases of the mission. For the nominal atmos^ phere, the entry phase starts 536 km above the 1 atmosphere pressure level and ends 59 seconds later. The descent phase starts after the aeroshell is staged and lasts to the end of the mission 44 minutes later. Parametric data showing the critical event times and altitudes for the warm, nominal, and cool atmospheres is also shown in Table IV-2d. Maximum deceleration of 568 g is experienced in the cool atmosphere. The entry and descent trajectory for the 1979 Saturn and the Saturn/SU 80 missions are identical.
Uranus/SU 80 Mission Definition
The 1980 Saturn/Uranus mission is shown in Figure IV-3 and detailed in Table IV-3. The important mission design results are summarized in this section.
a. Interplanetary Trajectory Selection - The interplanetary trajectory from Earth to Uranus with an intermediate Saturn swing-by is pictured in Figure IV-3a. The launch date is December 3, 1980 and the Uranus arrival date is October 23, 1988. The total trip time is 7.7 years.
IV-7
b. Launch Analysis - Available payload weight is plotted against launch period for four sets of launch vehicle performance data in Figure IV-3b.
a. Approach Trajectories - The probe trajectory entry angle, y, of -60° was selected. The resulting approach trajectory is pictured in Figure IV-3d and summarized in Table IV-3b. The spacecraft periapsis radius of 2.0 IL results in an acceptable spacecraft-to-probe communication range at entry. The probe aspect angle at beginning of descent is 2° and increases to 43° at the end of the mission. The relatively large EOM probe aspect angle is the result of targeting for common spacecraft to probe look direction for the three missions being considered.
d. Deflection Maneuver - A spacecraft deflection maneuver was used to establish the above defined link geometry and acquire the entry site. A deflection radius of 10 x 105 km was used with a resulting spacecraft AV requirement of 85 m/sec. The implementation sequence is pictured in Figure IV-3c.
e. Navigation and Dispersions - This design mission requires optical tracking to supplement the standard Earth-based tracking, because Uranus' ephemeris uncertainties are about ten times greater than those of Saturn. The navigation results provided in Table IV-3c are approximately equivalent to those obtained at Saturn with standard Earth-based tracking.
f. Entry and Descent Trajectories - Table IV-3d summarizes the entry and descent phases of the mission. Both entry and descent were simulated using all three atmosphere models. The variation in entry duration is from 27 seconds for the cool atmosphere to 90 seconds for the warm. During this phase of the mission, a peak deceleration of 837 g is attained 6.5 seconds after entry into the cool atmosphere.
IV-8
< I
ARRIVAL
AT URA
NUS,
OCT
23, 1988
ENCOUNTER
SATURN,
82
MAY
9, 1
984 83,
84j
4.3 yr
0 12 3 4 5
DISTANCE, AU
PAYLOAD
WEIG
HT,
kg
900
800 -
700
600-
500
400 "-
81 ORBIT
OF
SATURN
172.5°
SPACECRAFT ON
EARTH
LOCK
S\
21-4°
/
PROBE
" /
\ 46°
EARTH
.
AT SATURN
3.4 yr
ENCOUNTER
- \
EARTH
>K"^
LAUNCH
AT URANUS /
FROM EARTH,
ENCO
UNTE
RJ DEC
3, 1980
PROBE
RELEASE
/
AV = 85.1 m/sec
SPACECRAFT AV
b) DEFLECTION M
ANEUVER
67.4°
HP
RETURN TO
EARTH
LOCK
10 a)
INTERPLANETARY
TRAJECTORY
2100
1900
1700
PAYLOAD 1500
WEIG
HT,
lb
1300
1100
900
TITAN
•IIIE/CENTAUR/TE 364-4
IIIE/CENTAUR/BURNER II
0 5
10
15
LAUNCH P
ERIO
D, days
c) LAUNCH A
NALYSIS
Fig
. IV
-3
Ura
nus/
SU
8bi M
issi
on
Def
init
ion
TERMINATOR
PROBE
SPACECRAFT
ROTATION
r\
J POLE
'~
~3^
ENTRY
EOM
(E +
44 mi
nute
s)
PERIAPSIS
ENTRY
ANGLE
ENTRY
LATITUDE =
2.0 R,
-60°
30°N U
d) PLANETARY
ENCOUNTER
PERIAPSIS
Table IV-S Uranus/SU 1980 Mission Summary Data
a) CONIC TRAJECTORY DATA
INTERPLANETARY TRAJECTORY
LAUNCH DATE: 12/3/80
ARRIVAL DATE: 10/23/88
FLIGHT TIME: 7.7 yr
LAUNCH TRAJECTORY
NOMINAL C3:
NOMINAL DLA:
C3 (5-day):
C3 (10-day):
C3 (15-day):
132.7 km2/sec2
27. 6°
134.5 km2/sec2
135 km2/sec2
140 km2/sec2
ARRIVAL TRAJECTORY
VHP
•RA:
DEC
ZAE
ZAP
INC
11.87 km/sec
276.5°
-2.17°
169.7°
172.47°
5.9°
b) DEFLECTION MANEUVER AND PROBE CONIC
DEFLECTION MANEUVER
DEFLECTION MODE:
DEFLECTION RADIUS:
COAST TIME:
AV:
APPLICATION ANGLE:
OUT-OF-PLANE ANGLE:
PROBE RELEASE ANGLE:
SPACECRAFT DEFLECTION ANGLE:
SPACECRAFT REORIENTATION ANGLE:
*SUBSOLAR ORBITAL PLANE.
SPACECRAFT
10 x 106 km
9.5 days
85.1 m/sec
127°
180°
21.3°
22.4°
43.7°
PROBE CONIC DEFINITION
ENTRY ANGLE:
ENTRY LATITUDE*:
ENTRY LONGITUDE*:
LEAD TIME:
LEAD ANGLE:
PROBE/SPACECRAFT RANGE (ENTRY):
PROBE ASPECT ANGLE (ENTRY - ) :
PROBE ASPECT ANGLE (ENTRY +):
PROBE ASPECT ANGLE (EOM):
-60°
-2.5°
47.3°
1.8 hr
1.5°
8.7 x lO1*
31°
2°
43°
km
c) DISPERSION ANALYSIS SUMMARY
NAVIGATION
TRACKING:
SMAA:
SMIA:
6:
TOF:
UNCERTAINTIES
OPTICAL S DOPPLER
1682 km
702 km
20°
635.8 sec
EXECUTION ERRORS (3o)
AV PROPORTIONALITY:
AV POINTING:
PROBE ORIENTATION POINTING:
1%
2°
2°
DISPERSIONS (3a)
ENTRY ANGLE:
ANGLE OF ATTACK:
DOWNRANGE:
CROSSRANGE:
LEAD ANGLE:
LEAD TIME:
ENTRY TIME:
6.5°
3.15°
9.3°
3.4°
13.7°
1.22 minutes
31 minutes
d) ENTRY AND DESCENT SUMMARY
ENTRY PARAMETERS
ENTRY VELOCITY, km/sec
ENTRY ALTITUDE, km
ENTRY B, kg/m2
MAXIMUM DECELERATION, g
MAXIMUM DYNAMIC PRESSURE, N/m2
EOM PRESSURE, bar
DESCENT B, kg/m2
CRITICAL ENTRY EVENTS
g = 0.1
g = 100
MAXIMUM DECELERATION
MACH = 0.8
MACH = 0.7
ATMOSPHERE MODEL
WARM
22.1
1073.8
102
225
2.2 x 105
3.1
110
ATMOSPHERE MODEL
WARM
TIME, sec
14.0
36
40
82.5
90
ALTITUDE, km
770
300
235
143
140
NOMINAL
22.2
532
102
358
3.5 x 105
7.3
110
NOMINAL
TIME, sec
4.0
15
19
49.5
54.5
COOL
22.3
200
102
837
8.3 x 105
19.2
110
ALTITUDE, km
444
208
138
80.3
78.5
COOL
TIME, sec
0.5
4.5
6.5
24.5
27.5
ALTITUDE, km
188
102
65
37.3
36.5
IV-10
B. SCIENCE INSTRUMENTATION AND PERFORMANCE
The instruments for the basic common Saturn/Uranus probe are the SAG Exploratory payload instruments. Their characteristics were given in Table V-29 of Volume II and are also given in the first part of Table V-l of this volume. They are discussed in detail in Chapter III, Section C.l of Volume II. These instruments are unchanged from the first part of the study.
The descent profiles shown in Figure 111-31 are different from those in the previous part of the study only because three atmospheres are considered. The ballistic coefficient is 110 kg/m2
for all descents. Values for descent parameters are given in Table 111-14. It shows the parachute deployment pressure varying from 41 to 102 millibars while the pressure at the first descent measurement varies from 51 to 122 millibars. The final pressure for a descent designed at 44 minutes varies from 3.1 bars in the Uranus warm to 19.2 bars in the Uranus cool temperature. However, the probe falls 238 km in the Uranus warm, while descending only 92 km in the Uranus cool, atmosphere. The distance in the Saturn warm atmosphere corresponding to the same 44 minutes is 320 km.
The entry and descent times, instrument sampling times, and resulting bit rates are shown in Table IV-4. The transmission bit rates for descent are slightly higher than the collection bit rates because the descent data collected during probe acquisition by the spacecraft must be stored and later interleaved with real-time data for transmission. The difference between this table and Table VI-2 of Volume II is that the mass spectrometer sampling time has been reduced to 50 seconds. Both the entry and descent times are slightly longer because of consideration of other model atmospheres; thus the bit rate to telemeter the stored accelerom-eter data is coincidently 3.4 bps in both tables despite the fact that more data is collected here (8750 bits plus formatting, etc).
The descent mission measurement performance is given in Table IV-5 for Saturn and in Table IV-6 for Uranus. The first column gives the instrument, its sampling interval, and the measurement to be made. The second column gives the criteria set forth in Volume II. The remaining columns give the performance numbers in the units specified by the criteria. The minimum values are those at the highest point in the atmosphere that the measurement is to be evaluated, and are specified individually for each atmosphere to allow determination of worst case atmospheres. The maximum value given in the last column is the maximum performance of all the atmospheres at the end of the design mission (44 min from parachute deployment).
IV-11
First, note that the mass spectrometer makes two full 1 to 40 amu sweeps inside the methane cloud in the Uranus warm atmosphere at the 50-second sampling time. At 60 seconds, it was only completing one sweep and a partial sweep inside this extremely low-pressure cloud. Since definition of the atmospheric clouds is an important objective, it was necessary to reduce the sampling time to the lower value in order to adequately sample this cloud.
Secondly, note that in most cases, the performance is lowest in the warm atmospheres. In fact, the pressure and turbulence ac-celerometer measurements are below the criteria at the highest evaluated point in both warm atmospheres. However, since the number of measurements per kilometer increases with descent and an adequate number of measurements is obtained within the highest modeled cloud (26 to 38), this does not necessitate a modification of the instrument sampling time.
Table IV-4 Instrument Sampling Times and Data Rates for Common Saturn/ Uranus Probe with Exploratory Pay load
PHASE
ENTRY
(84.5 sec MAXIMUM)
DESCENT
(2640 sec CONSTANT)
INSTRUMENT
ACCELEROMETERS
AXIAL
LATERAL
LATERAL
TEMPERATURE GAGE
PRESSURE GAGE
NEUTRAL MASS SPECTROMETER
ACCELEROMETERS
TURBULENCE
STORED
SAMPLING TIMES, sec
0.2
0.4
0.4
4
4
50
8
0
COLLECTION BIT RATE, bps
50
25
25
2.5
2.5
8.0
7.5
0
TRANSMISSION BIT RATE, bps
0
0
0
2.6
2.6
8.3
7.8
3.4
SCIENCE TOTAL 24.7
ENGINEERING 0.5
FORMATTING 2.5
TOTAL 27.7 bps
IV-12
Tab
le
TV
-5
Satu
rn
Des
cent
M
easu
rem
ent
Per
form
ance -
Exp
lora
tory
P
aylo
ad
INSTRUMENT
MASS
SPECTROMETER
TEMP
ERAT
URE
GAGE
PRESSURE
GAGE
ACCELEROM-
ETERS
SAMPLING TIME
50 sec
4 sec
4 sec
8 sec
MEAS
UREM
ENT
MINOR
CONSTITUENTS
CLOU
D COMPOSITION
H/He RATIO
\ ISOTOPIC RATIO
|
MOLECULAR
WEIGHT
>
TEMP
ERAT
URE
PROFILE
CLOU
D STRUCTURE
PRESSURE PROFILE
TURB
ULEN
CE
CLOU
D STRUCTURE
TURB
ULEN
CE
CRITERIA
2 PER SCALE HE
IGHT
*
2 INSIDE EACH CL
OUD
4 ME
ASUR
EMEN
TS
1 PER °K
2 INSIDE EACH CL
OUD
2 PER km )
1 PER km ) *
2 INSIDE EACH CL
OUD
1 PER km
MINIMUM
PERFORMANCE
COOL
5.8
10.4 in NH3
52
1.5
137 in NH3
2.7
137 in NH3
1.3
NOMINAL
7.7
4.0 in NH3
16 in H20
52
1.2
46 in NH3
196 in H20
2.0
46 in NH3
196 in H20
1.0
WARM
6.5
3.0 in NH3
10.5 in H20
52
1.3
38 in NH 3
47 in HjO
1.5
38 in NH3
47 in H20
0.7
MAXIMUM
PERFORMANCE
37
8.1
7.8
3.9
*BEL0W CL
OUD
TOPS
.
Tab
le
IV-6
U
ranu
s D
esce
nt
Mea
sure
men
t P
erfo
rman
ce -
Exp
lora
tory
P
aylo
ad
INSTRUMENT
MASS
SPECTROMETER
TEMPERATURE
GAGE
PRESSURE
GAGE
ACCE
LERO
M-ETERS
SAMPLING TIME
50 sec
4 sec
4 sec
8 sec
MEAS
UREM
ENT
MINOR
CONSTITUENTS
CLOU
D COMPOSITION
H/He RATIO
\
ISOTOPIC RATIO
>
MOLE
CULA
R WEIGHT )
TEMPERATURE
PROFILE
CLOU
D STRUCTURE
PRES
SURE
PROFILE
TURB
ULEN
CE
CLOUD
STRUCTURE
TURB
ULEN
CE
CRITERIA
2 PER SCALE HE
IGHT
*
2 INSIDE EACH CL
OUD
4 ME
ASUR
EMEN
TS
1 PER °K
2 INSIDE EACH CL
OUD
2 per
km J
1 PER
km )
2 INSIDE EACH CL
OUD
1 PER km
MINIMUM
PERFORMANCE
COOL
5.6
18.3 in C
H„
52
3.0
230 in CH M
4.8
230 in
CH,,
2.4
NOMINAL
6.1
4.4 in CH,,
12 in NH3
52
3.0
48 in CH,,
113 in NH 3
2.5
48 in CH„
113 in NH3
1.4
WARM
4.6
2.0
in C
H„
5 in NH3
52
3.1
26 in CH,,
63 in NH3
1.6
26 in CH„
63 in NH3
0.9
MAXIMUM
PERFORMANCE
38
10.9
14.4
7.2
*BELOW CL
OUD
TOPS
.
c. SYSTEM DESIGN AND INTEGRATION
1. Functional Sequence
Figure IV-4 shows a pictorial sequence of events with emphasis on the probe activity. It depicts the major probe and spacecraft events and shows the relationship between the probe and spacecraft after probe separation. Pre-entry times are omitted because they are different for each planet due to entry arrival uncertainty. Tables IV-7 and IV-8 show the detailed sequence of events for Saturn and Uranus, respectively, and are similar to the sequence for the probe-dedicated Jupiter mission shown in Table V-33 of Volume II. These sequences are based upon the nominal atmospheric models but show the variation of the various parameters (time, Mach No., pressure level, and altitude) for the cool and warm atmospheric models. As an example, the main parachute is deployed at Saturn at 5 g (decreasing) plus 15 seconds. This event occurs at entry (E) plus 59.5 seconds in the nominal atmosphere, E plus 44 seconds in the cool atmosphere, and E plus 84.5 seconds in the warm atmosphere. This event concludes the entry phase and starts the descent phase. Since accelerometer data is stored during the entry phase at a high data rate, the warm atmosphere duration, being the longest, is used for sizing the data storage.
E+44 min 59 sec Link Established ] End of Mission start Data Transmission
(Approx 8 bars)
Figure IV-4 Sequence of Events (Task 1)
IV-14
Table IV-7 Saturn Sequence of Events (Task I)*
ITEM
1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
11.
12.
13.
14.
15.
16.
17.
18.
19.
20.
21.
22.
23.
24.
25.
26.
27.
28.
29.
30.
31.
32.
33.
*ASSUM NOMIN
+ (C) =
TIME
L = 0
L + 2h t o L + 1221.8d
S - 7 h , 30m, Os
S - 7 h , 17m, 0s
S - Oh, 17m, Os
S - Oh, 2m, Os
S = 0 (L + 1 2 2 2 . I d )
S + Om, 0 .5s
S + Om, 30s
S + 4m, Os
S + 15m, 30s
S + 17m, 36s
S + 18m, Os
S + 33m, Os
L + 1222.12d t o L + 125.79d
E - 43m, 57 .5s
E - 23m, 57 .5s
E - 23m, 27 .5s
E - 8m, 9 .5s
E - 7m, 39 .5s
E - 5m, 59 .5s
E = 0
E + Om, 5s
E + Om, 17.5s
E + Om, 44 .5s [5g ( R e f ) ]
E + Om, 56 .5s (5g + 12s)
E + Om, 59 .5s (5g + 15s)
E + l m , 9 .5s (5g + 25s)
E + l m , 14.5s (5g + 30s)
E + 2m, 39 .5s (5g + 115s)
E + 44m, 59 .5s (5g + 2655s)
E + . 2 h , 12m (L + 1258d)
IS AN ENTRY ARRIVAL UNCERTAINTY M ATMOSPHERE.
COOL ATMOSPHERE; (W) = WARM ATf
EVENT
LAUNCH (DEC 3 , 1980)
SEPARATE SPACECRAFT FROM LAUNCH VEHICLE
CRUISE
SPACECRAFT POWER TO PROBE; EJECT ENVIRONMENTAL COVER
START PROBE CHECKOUT
PROBE CHECKOUT COMPLETE; START SPACECRAFT ORIENTATION ( 1 0 . 6 ° )
SPACECRAFT ORIENTATION COMPLETE; ACTIVATE SEPARATION SUBSYSTEM
SEPARATE PROBE FROM SPACECRAFT
START PROBE SPINUP TO 5 rpm
START SPACECRAFT ORIENTATION FOR AV ( 5 9 . 1 ° )
PROBE SPINUP COMPLETE; DEACTIVATE PROBE SYSTEM
COMPLETE SPACECRAFT ORIENTATION FOR AV
APPLY SPACECRAFT AV (96 m/sec)
START SPACECRAFT ORIENTATION TO EARTH LOCK ( 6 9 . 7 ° )
SPACECRAFT ORIENTATION COMPLETE
COAST
START BATTERY ACTIVATION PYRO CAPACITOR CHARGE
ACTIVATE BATTERY
POWER ON (DATA HANDLING SYSTEM & PYRO)
TURN ON SCIENCE & TRANSMITTER (WARM-UP)
START PROBE ACQUISITION
COMPLETE PROBE ACQUISITION; START DATA TRANSMISSION
ENTRY (536 km ABOVE 1 a t m ; 1 x 1 0 - 7 atm)
RF POWER AMPLIFIER OFF ( 0 . 1 - g SENSING); START DATA STORAGE
ENABLE PROBE DESCENT PROGRAM (100 -g SENSING)
START DESCENT PROGRAM ( 5 - g SENSING); OPERATE BASE COVER BAND CUTTERS
DEPLOY BASE COVER QUADRANTS
DEPLOY MAIN PARACHUTE (M = 0 . 7 ; ^65 m b ) ; RF POWER AMPLIFIER ON; START ACQUISITION
EJECT AEROSHELL
RELEASE MAIN CHUTE; DEPLOY TEMPERATURE GAGE; UNCOVER NEUTRAL MASS SPECTROMETER
PROBE ACQUISITION COMPLETE; START DATA TRANSMISSION
END OF DESIGN MISSION (n-8 b a r s )
SPACECRAFT PERIAPSIS ( 3 . 8 R s ) ; MAY 9 , 1984
OF 4 . 4 m i n u t e s , A DESCENT BALLISTIC COEFFICIENT OF 11
10SPHERE.
COMMENTS+
ENGINEERING DATA AT i 4 bps
297 km ( C ) ; 9 6 8 . 5 km (W)
1.7s ( C ) ; 12s (W)
8 .5s ( C ) ; 35s (W)
29s ( C ) ; 69 .5s (W)
41s ( C ) ; 81 .5s (W)
( 4 4 s ( C ) ; 84 .5s (W) }M = 0 .56 ( C ) ; M = 0 .88 (W) ( 1 0 2 mb ( C ) ; 44 mb (W)
28 bps
45m, 24 .5s (W)
18 bars ( C ) ; 4 bars (W)
3 kg /m 2 ( 0 . 7 s l u g / f t 2 ) , AND A
IV-15
Table 1V-8 Uranus Sequence of Events (Task I)*
ITEM
1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
11.
12.
13.
14.
15.
16.
17.
18.
19.
20.
21 .
22.
23.
24.
25.
26.
27.
28.
29.
30.
31.
32.
33.
*ASSUM NOMIN
+ (C) -
TIME
L = 0
L + 2h
L + 2h t o L + 2816 .7d
S - 1 3 h , 30m, 0s
S - 1 3 h , 17m, 0s
S - Oh , 17m, 0s
S - Oh, 2m, 0s
S = 0 (L + 2817 .3d )
S + 0m, 0 .5s
S + Om, 30s
S + 4m, Os
S + 15m, 30s
S + 17m, 36s
S + 18m, Os
S + 33m, Os
L + 2817.32d t o L + 2824.6d
E - l h , 8m, 39 .5s
E - 48m, 39 .5s
E - 48m, 9 .5s
E - 32m, 51 .5s
E - 32m, 21 .5s
E - 30m, 41 .5s
E = 0
E + Om, 4s
E + Om, 15s
E + Om, 38 .5s [5g ( R e f ) ]
E + Om, 50 .5s (5g + 12s)
E + Om, 53 .5s (5g + 15s)
E + l m , 3 .5s (5g + 25s)
E + l m , 8 . 5 s (5g + 30s)
E + 2m, 33 .5s (5g + 115s)
E + 44m, 53 .5s (5g + 2655s)
E + l h , 51m (L + 2826 .9d )
ES AN ENTRY ARRIVAL UNCERTAINTY AL ATMOSPHERE.
COOL ATMOSPHERE; (W) = WARM AT
EVENT
LAUNCH (DEC 3 , 1980)
SEPARATE SPACECRAFT FROM LAUNCH VEHICLE
CRUISE
SPACECRAFT POWER TO PROBE; EJECT ENVIRONMENTAL COVER
START PROBE CHECKOUT
PROBE CHECKOUT COMPLETE; START SPACECRAFT ORIENTATION ( 2 1 . 4 ° )
SPACECRAFT ORIENTATION COMPLETE; ACTIVATE SEPARATION SUBSYSTEMS
SEPARATE PROBE FROM SPACECRAFT
START PROBE SPINUP TO 5 rpm
START SPACECRAFT ORIENTATION FOR AV ( 4 6 ° )
PROBE SPINUP COMPLETE; DEACTIVATE PROBE SYSTEM
COMPLETE SPACECRAFT ORIENTATION FOR AV
APPLY SPACECRAFT AV (86 m/sec)
START SPACECRAFT ORIENTATION TO EARTH LOCK ( 6 7 . 4 ° ) ,
SPACECRAFT ORIENTATION COMPLETE
COAST
START BATTERY ACTIVATION PYRO CAPACITOR CHARGE
ACTIVATE BATTERY
POWER ON (DATA HANDLING SYSTEM & PYRO)
TURN ON SCIENCE & TRANSMITTER (WARM-UP)
START PROBE ACQUISITION
COMPLETE PROBE ACQUISITION; START DATA TRANSMISSION
ENTRY (532 km ABOVE 1 a t m ; 1 x 1 0 - 7 atm)
RF POWER AMPLIFIER OFF ( 0 . 1 - g SENSING); START DATA STORAGE
ENABLE PROBE DESCENT PROGRAM ( 1 0 0 - g SENSING)
START DESCENT PROGRAM ( 5 - g SENSING); OPERATE BASE COVER BAND CUTTERS
DEPLOY BASE COVER QUADRANTS
DEPLOY MAIN PARACHUTE (M = 0 . 7 2 ; ^47 m b ) ; RF POWER AMPLIFIER ON; START ACQUISITION
EJECT AEROSHELL
RELEASE MAIN CHUTE; DEPLOY TEMPERATURE GAGE; UNCOVER NEUTRAL MASS SPECTROMETER
PROBE ACQUISITION COMPLETE; START DATA TRANSMISSION
END OF DESIGN MISSION ( ^ 7 . 3 b a r s )
SPACECRAFT PERIAPSIS ( 2 R U ) ; AUGUST 2 4 , 1988
OF 29 m i n u t e s , A DESCENT BALLISTIC COEFFICIENT OF 110
MOSPHERE.
COMMENTS1"
ENGINEERING DATA AT ^ 1 bps
200 km ( C ) ; 1074 km (W)
0 .5s ( C ) ; 14s (W)
4 . 5 s ( C ) ; 36s (W)
18.5s ( C ) ; 65s (W)
30 .5s ( C ) ; 77s (W)
( 3 3 . 5 s ( C ) ; 80s (W) { M = 0 .59 ( C ) ; = 0 . 8 4 (W) ( 4 1 mb ( C ) ; 49 mb (W)
28 bps
19 bars ( C ) ; 3 . 1 bars (W)
kg /m 2 ( 0 . 7 s l u g / f t 2 ) AND A
IV-16
Functional Block Diagram
The functional block diagram is the same as Figure V-76 of Volume II with the four subsystem areas deleted or reduced as follows for spacecraft deflection mode: propulsion and ACS electronics are deleted, ACS propulsion is reduced by 75%, and separation power is reduced by 50%.
System Data Profile
The data profile for the Saturn/Uranus probe shown in Figure IV-5 is very similar to Figure V-101 of Volume II for the probe-dedicated Jupiter mission. As in the earlier data profile, the separation phase is very short (approximately 4 min) because the probe is deployed by the spacecraft, and because of the short duration, data is not collected during this interval. Due to the entry arrival uncertainty of 4.4 minutes for Saturn and 29 minutes for Uranus, the pre-entry activity is different for each planet and therefore requires a different sequence to be stored onboard the probe and selected by Earth commands to the spacecraft. The entry and descent data profile is the same for both planets, i.e., total data collected is 70,300 bits, maximum storage on the probe is 11,637 bits, and the probe-to-spacecraft data transmission rate is approximately 28 bits/sec.
System Power Profile
Figure IV-6, showing the power profile for the Saturn/Uranus probe, is similar to that for the probe-dedicated Jupiter (Volume II, Chapter V, Section C3) except that the "worst case" entry arrival uncertainty of 29 minutes for Uranus is shown. The figure shows the effect of this arrival uncertainty on the power requirements. The late arrival condition requires 133 w-hr of power at the equipment. Late arrival is the condition where the probe has not yet arrived at the planet when the timer (preset before separation) runs out and starts the sequence before it is really needed.
System Weight Summary
Table IV-9 shows the weight breakdown for the Saturn/Uranus probe. Ejected weight is 89.02 kg (196.08 lb), entry weight is 89.01 kg (196.05 lb), and the final descent weight is 40.46 kg (89.11 lb). Compared with the Saturn probe weight shown in Table VI-6 of Volume II, this probe ejection weight is lighter because of the absence of a AV motor, but the entry weight is heavier due to higher decelerations, longer entry and descent time, which increases the battery weight, and to the small amount of propulsion weight that is carried to entry instead of being ejected with the service module as was the previous case.
SEPARATION
& COAST
*35.8
days
(S
ATUR
N)'
9.5
days (URANUS)
DATA
, 10
3 bi
ts' 70
60
50
40
30
20
10 0
PREENTRY
ENTRY*
DESCENT
START TRANSMISSION AT URANUS
576
BITS STORED
DATA +
0.55 bps
REAL-TIME
DATA
(7 w RF POWER)
*ENTRY DATA BASED ON
SATURN W
ARM ATMOSPHERE
START TRANSMISSION AT SATURN
1048 BITS
STORED
DATA +
0.55 bps
REAL-TIME
DATA AT <\-l bps
(7 W RF POW
ER)
START
STORING -|
URANUS DATA AT
0.55 bps
I START
STORING
SATURN DATA AT
0.55 bps r
_L
_L
-60
-40
-20
TIME FROM EN
TRY,
minutes
20
40
Fig
. IV
-5
Dat
a P
rofi
le
for
Satu
rn/U
ranu
s P
robe
(T
ask
I)
ENTRY
POWER,
w
100
90
80
70
60
50
40
30
20
10
0 SE
PARATION
-*
— 10
w
0.67 w-hr
1
COAST
35.9 days
(SATURN)
9.5
days
(URANUS)
\h
<
0 2
TIME AFTER
SEPARATION, minutes
•A)
PREENTRY
LATE
ARRIVAL
(LA)
NOMINAL
ARRIVAL
(NA)
SCIENCE &
1
TRANSMITTER ON
(7 w RF POWER)
DHS &
PYRO ON
9.7 w
1
EARLY
ARRIVAL
(EA)
54.3 w
RF
POWER
AMPLIFIER-
OFF
DESCENT
100 w
TRANSMITTER ON
(25 w RF POW
ER)
EA =
79.2 w-hr
NA = 106.5 w-hr
LA =
132.8 w-hr
30.6 w
I Ji
J L
-60
-40
-20
TIME FROM ENTRY, minutes
20
40
Fig
. IV
-6
Pow
er
Pro
file
fo
r Sa
turn
/Ura
nus
Pro
be
(Tas
k I)
Table IV-9 Weight Summary for Saturn/Uranus Probe SAG Exploratory Pay load
with
SCIENCE
POWER & POWER CONDITIONING
CABLING
DATA HANDLING
ATTITUDE CONTROL SUBSYSTEM (DRY)
COMMUNICATIONS
PYROTECHNICS
STRUCTURES & HEAT SHIELD
MECHANISMS
THERMAL
PROPELLANT (ACS)
ENGINEERING INSTRUMENTATION
15% MARGIN
EJECTED WEIGHT
ENTRY WEIGHT
DESCENT WEIGHT (FINAL)
WEIGHT
kg
8.08
4.70
3.86
2.43
2.84
3.86
6.38
34.00
6.03
5.22
0.01
0.00
11.61
89.02
89.01
40.46
lb
17.80
10.35
8.50
5.35
6.25
8.50
14.05
74.89
13.28
11.50
0.03
0.00
25.58
196.08
196.05
89.11
IV-20
D. TELECOMMUNICATIONS SUBSYSTEM
Results of the parametric study performed to arrive at a common Saturn/Uranus probe design were used to define the telecommunications subsystem. The Saturn/SU 80 mission trajectory generates the worst-case communications geometry due mainly to the large communications range of 2 x 10 km at entry. The cool Saturn atmosphere model has the greatest microwave absorption at 0.86 GHz (see Figures 111-43 and 111-44). Therefore, the Saturn/SU 80 trajectory, as developed in Chapter III, Section D.l as configuration 4-S, is used to define the telecommunications subsystem for the Task I science payload.
The system noise temperatures were reevaluated for the follow-on effort. Revised curves are shown in Figures 111-37 and 111-38 for Saturn and Uranus, respectively. The revised curves indicate a slightly higher system noise (due to cable attenuation) and its impact on receiver front-end noise temperature. The antenna noise temperature is unchanged for Saturn. A revised antenna noise temperature curve for Uranus was used based on data from a monograph published by NASA (see Ref III-2). Feedline noise was also considered for the revised Uranus., system noise temperature calculation. Values for receiver noise temperature were also investigated and found to be in agreement with the predicted receiver state of the art by several suppliers. Therefore, the original values in Vol II, Figure V-ll, p V-21 were used in the updated analysis.
As discussed in detail in Chapter III, Section D.l, trajectory adjustments were made in order to optimize the communications geometry for Saturn encounter from the SU 80 trajectory and at the same time develop common trajectories for Saturn and Uranus. Probe-to-spacecraft communications range is shown in Figure IV-7 for the Saturn/SU 80 mission. Increasing range before entry is caused by the relative motion of the probe and spacecraft with the spacecraft moving ahead of the probe as the probe begins descent into the atmosphere. Maximum range occurs at entry and decreases by approximately 0.2 R (12,000 km) at mission completion. The spacecraft reaches periapsis after the mission is over (E+2.3 hr).
The probe aspect angle as a function of mission time is shown in Figure IV-8. The angle steadily increases during descent from 7.5° at entry to 24° at mission completion. Three-sigma variations in the angle are 3°, based upon a 100-sample Monte Carlo
IV-21
Rs = 59,800 km
REJ = 3 x 107 km
1.98
1.96
1.94
1.92 COMMUNICATIONS RANGE, 105 km
1.90
1.88
1.86
1.84
Yr = -30°
PERIAPSIS Rp = 3.8 P
E + 2.3 hr
COMMUNICATIONS RANGE, Rc
-0.4 0 0.2 0.4 TIME FROM ENTRY, hr
Fig. IV-7 Probe-to-Spaaecraft Communications Range for the Saturn/SU-1980 Mission
IV-2 2
LOOK VECTOR
DESCENT ANTENNA
DESCENT PROBE
MAIN BEAM AXIS
h-PRE-ENTRY » >W
ACQUISITION ENTRY (E-O.l) I I
50
40
30
PROBE ASPECT ANGLE, i>, deg
20
10
DESCENT -
-0.4 -0.2
CONDITIONS:
Rp = 3.8 R$
YE R EJ
= -30°
= 3 x 107 km
= 2.1 hr 'L EOM = 8 bar (E + 45 minutes] COOL ATMOSPHERE MODEL
END OF MISSI0N(E+0.75)
1_ ±3° 3-a VARIATIONS
T ±20° PARACHUTE SWING
0 0 . 2 0 . 4
TIME FROM ENTRY, hr
Fig. IV-8 Satum/SU 1980 Mission Probe Aspect Angle
0.6 0.8
IV-2 3
analysis. Also included is 20° to account for parachute swing resulting from atmosphere turbulence and lateral winds. The resulting probe descent antenna beamwidth is 100° with circular polarization.
Relative probe positions during the mission are shown in Figure 111-52. The ellipses account for navigation uncertainties and execution errors used with a 100-sample Monte Carlo error analysis. Nominal positions are shown by the triangles. Whether the ellipse major axis is aligned with the cone or clock axes is determined by the direction of rotation of the planet relative to the spacecraft trajectory. For Saturn, elevation angle (cross-cone angle) dispersions are small with the major differences occurring in cone angle. Decreasing cone angle during descent results from the spacecraft pulling ahead of the descending probe. A 20° beamwidth probe receiving antenna on the spacecraft covers the relative probe positions including uncertainties for Saturn encounter from the SU 80 trajectory.
Probe transmitter output power requirements as a function of mission time are shown for the three missions in Figure IV-9 for Task I data rates. The curves are the result of considering all the RF link parameters using binary FSK modulation. RF power requirements for Saturn are increasing rapidly at mission completion due to increasing atmosphere attenuation (Figure 111-44) and decreasing spacecraft antenna gain (Figure 111-52). Two frequencies are shown for the worst-case mission. As seen in the figure, a small increase in the RF operating frequency results in large increases in RF power required. Also indicated on the curves are descent pressures for the two worst-case atmosphere models. The curves indicate that the worst-case mission is a Saturn encounter from the SU 80 trajectory.
Parameters of the RF link are depicted in Table IV-10 for the Saturn/SU 80 mission at mission completion which is the worst-case point. An RF power of 25 w is required at this point with the Task I science payload. This is the power level that defines the probe transmitter output power requirement and is the worst-case trajectory, atmosphere, and descent point.
IV-2 4
Table IV-11 depicts design details of the RF components that comprise the telecommunications subsystem for the common Saturn/ Uranus probe with Task I science payload. The probe transmitter has two RF power levels available for the mission, 7 and 25 w. Thermal control at Uranus prior to entry requires 7 w. The transmitter operates at 0.86 GHz and uses binary FSK modulation with a tracking tone. At the present state of the art, a mechanical RF switch may be used, but the possibility exists that a lighter, solid-state switch may be available by 1976. Identical pre-entry and descent turnstile/cone antennas are used on the probe, and the probe receiving antenna on the spacecraft is positioned at a single point in space to cover the three missions (Figure 111-52). Polarization of probe and spacecraft antennas are right-hand circular .
IV-25
URANUS ACQUISITION
SATURN ACQUISITION ENTRY
45
40
35
30 PROBE RF POWER REQUIRED, w
1 .PRE-ENTRY. 1 bps —
LEGEND:
S/SU 80
S 79
U/SU 80
.DESCENT 28 bps
NOTE: 1. COOL ATMOSPHERE MODEL FOR SATURN. 2. NOMINAL ATMOSPHERE MODEL FOR URANUS.
END OF MISSION
20
18
-0.2 0 0.2
TIME FROM ENTRY, hr 0.8
Fig. IV-9 Task I Probe RF Power Requirements
IV-2 6
Table IV-10 Probe Telemetry Link Design Table for the Saturn/SU 1980 Mission with Task I Science*
PARAMETER
1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
11.
12.
13.
14.
15.
16.
17.
18.
19.
20.
21.
22.
23.
24.
25.
26.
27.
TOTAL TRANSMITTER POWER, dBw
TRANSMITTING CIRCUIT LOSS, <JB
TRANSMITTING ANTENNA GAIN, dB
COMMUNICATIONS RANGE LOSS, dB
PLANET ATMOSPHERE & DEFOCUSSING LOSS, dB
POLARIZATION LOSS, dB
ANTENNA PATTERN RIPPLE LOSS, dB
RECEIVING ANTENNA GAIN, dB
RECEIVING CIRCUIT LOSS, dB
NET CIRCUIT LOSS, l (2 - 9), dB
TOTAL RECEIVED POWER (1 + 10), dBw.
RECEIVER NOISE SPECTRAL DENSITY, dBw/Hz
TRACKING TONE
TONE POWER/TOTAL POWER, dB
RECEIVED TONE POWER (11 + 13), dBw
TRACKING THRESHOLD BANDWIDTH, dB
THRESHOLD SIGNAL/NOISE RATIO,, dB
THRESHOLD TRACKING POWER (12 + 15 + 16) , dBw
TRACKING PERFORMANCE MARGIN (14 - 17), dB
DATA CHANNEL
DATA POWER/TOTAL POWER, dB
RADIO SYSTEM PROCESSING LOSS, dB
FADING LOSS, dB
RECEIVED DATA POWER (11 + 19 + 20 + 21), dBw
DATA BIT RATE, dB
THRESHOLD Eb /N Q, dB
THRESHOLD DATA POWER (12 + 23 + 24), dBw
PERFORMANCE MARGIN (22 - 25), dB
NOMINAL LESS ADVERSE VALUE (26 - 26 adv), dB
NOMINAL VALUE
14.0
-1.1
6.0
-196.5
-4.1
-0.3
0
17.9
-1.0
-179.1
-165.1
-196.2
-4.6
-169.7
12.4
10.0
-173.8
4.1
-1.6
-1.0
-1.0
-168.7
14.5
8.9
-172.8
4.1
0
•CONDITIONS: 1. S/SU 80 COMMON PROBE MISSION (WORST-CASE
2. WORST-CASE (END OF MISSION) CONDITIONS Al
3. CONVOUJTIONAL ENCODER, M = 2 , V = 2 , Q =
4. BER = 5 x 10"5 FOR BINARY FSK WITH K = 8
ADVERSE TOLERANCE
0
0.1
1.8
0.3
0.2
0
0.2
1.0
0.1
3.7
3.7
0.4
0
3.7
0
0
0.4
4.1
0
0
0
3.7
0
0
0.4
4.1
TRAJECTORY)
" 0.86 GHz.
8.
code.
REMARKS
25.1 w at 0.86 GHz
CABLE & SWITCH
100° BEAMWIDTH
1.8 x 105 km
COOL, 18 bar
20° BEAMWIDTH
CABLE
Ts = 1750°K
NF = 8.5 dB
Pt = 8.7 w
17.5-Hz BANDWIDTH
Pd = 17.4 w
28 bps
IV-27•
Table IV-11 Telecommunications RF Subsystem for the Task I Mission*
COMPONENT
TRANSMITTER
RF SWITCH
PROBE PRE-ENTRY & DESCENT ANTENNAS
SPACECRAFT ANTENNA
. SPACECRAFT RECEIVER
CHARACTERISTIC
RF POWER OUT OVERALL EFFICIENCY MAXIMUM dc POWER IN AT 28 vdc TOTAL WEIGHT VOLUME
TYPE INSERTION LOSS WEIGHT VOLUME
TYPE MAIN BEAM ANGLE BEAMWIDTH MAXIMUM GAIN SIZE (DIAMETER x HEIGHT) WEIGHT
TYPE BEAMWIDTH MAXIMUM GAIN SIZE (DIAMETER) WEIGHT
DESPIN POSITION SEARCH FREQUENCY ACQUISITION CLOCK ANGLE, 6 CONE ANGLE, <(,
NOISE TEMPERATURE NOISE FIGURE dc POWER IN AT 28 vdc WEIGHT
•CONDITIONS: 1. PLANETS: SATURN & URANUS.
2. MARINER SPACECRAFT.
3. FREQUENCY = 0.86 GHz.
4. BIT RATE = 28 bps.
VALUE
7 AND 25 w 45% 55.5 w 2.73 kg (6 lb) 3440 cm3 (210 in.3)
MECHANICAL SPDT 0.3 dB 0.23 kg (0.5 lb) 443 cm3 (27 in.3)
TURNSTILE/CONE 0 deg 100 deg 6.5 dB 15.1 x 8.25 cm (6 x 3.25 in.) 0.45 kg (1.0 lb)
PARABOLIC DISH 20 deg 18.3 dB 128 cm (50.4 in.) (4.2 ft) 4.54 kg (10 lb)
NO ONE 31 sec 257 deg 116 deg
300°K 3.1 dB 3.0 w 0.9 kg (2.0 lb)
IV-2 8
DATA HANDLING SUBSYSTEM
The data processing electronics are essentially unchanged from the previous study, as indicated in Chapter III, Section D2 of this report. The modifications necessary to provide the inflight target planet selection consist of approximately ten ROMs for the additional formats and a latching relay to implement the selection. The increase in weight associated with these additional elements is offset by a decrease in power conditioning and thermal dissipation equipment requirements. As discussed in Chapter III, Section D2 of this volume, COS/MOS data storage design has been employed for this mission. The physical and electrical characteristics are approximated by the following equations:
V = volume = 443 + 16.0 M (10~3) cm3
W = Weight = 0.2 + 0.715 M (10~5) kg
Power (stand-by) = P = 0.2 + 11.0 M (10-6) w
Power (operate) = P = 1.0 + 5.50 M (10~6) w
M = memory = number of bits stored
The assumption is made that twelve 1024-bit chips (M = 12,288 bits) will be required to store the blackout data. This results in the following values for the memory electronics, physical and electrical interface characteristics.
V = 640 cm3 W = 0.29 kg P = 0.34 w P = 1.7 w s o
Comparible characteristics for data processing electronics are assumed to be identical to the previous study (Vol I thru III).
V = 2330 cm3 w = 2.13 kg P = 6.9 w
The DHS begins the initial descent sequence shortly after the descent battery is energized. The initiation of the sequence and the initial state of the DHS will be established by a timing circuit working off battery voltage. This will ensure that the turn-on battery transient has subsided and will eliminate the need of an additional event from the coast timer. The required sequence of events is illustrated in Table IV-7 (Saturn Sequence of Events Task I). The pre-entry sequence occurs from item 19 to item 22 inclusive. The actual time with respect to entry of these events will vary (±4.4 min, Saturn; ±29 min, Uranus) due to trajectory
IV-29
uncertainties. The control state established at item 22 will remain fixed regardless of the actual time before entry until 0.1 g sensing (item 24). The functions required of the DHS formats with respect to the Task I Saturn sequence are tabulated.
From To Item Item Format
19 22 Store engineering data
22 24 Interleave stored and real-time engineering data
24 30 Store entry accelerometer science and engineering data
30 31 Store descent science and engineering data
31 32 Interleave stored and real-time science and engineering data
During periods 22 to 24 and 31 to 32, the interleaved data is sequenced to the encoder and RF transmitter.
IV-30
POWER AND PYROTECHNIC SUBSYSTEM
The configuration of the power and pyrotechnic subsystem for Task I is unchanged from the previous study (Vol I thru III) as indicated in Chapter III, Section D3 of this volume. Remote activated Ag-Zn batteries provide power during the brief post-separation activity and the entry and descent phase, Hg-Zn batteries provide power for the coast timer (1.3 volts, 7 u amp, 35 days max) and the initial entry pyro event (36 volts, 25 ma peak, 5 ma average, 1.0 hr).
The selection of the post-separation battery (EP GAP 4384; 0.25 kg, 93.5 cm3) is based on an available battery rather than an optimum battery. During the previous study, it was assumed that the DHS (6.9 w) and the pyrotechnics subsystem (0.5 w) would be operating at this time for approximately 4.0 minutes. Assuming some engineering functions, the total load was estimated to be 10.0 w. The battery is designed to deliver 1.0 amp (28 w) for 16 minutes. At lower loads, the capability of the battery may be expected to be greater. For purposes of decreased sequencing complexity, consideration of direct pyrotechnic initiation from the battery has been considered. A small RC timer may be used for the timing function, eliminating the need for operating the DHS and the pyrotechnic subsystems. Assuming this battery can provide the initiation currents for the dual squibs, the complexity of the separation activity is significantly reduced. The battery peak current capability remains to be fully validated(Chapter III, Section D.3); however, since the original option (activation of DHS and pyrotechnics) is available and there is no significant difference in the physical description of the two approaches, the original size and weight specification remains the same.
The entry and descent battery specification is based on the power required for subsystem components plus..losses and margins and the sequence of events for a late arrival probe. The maximum transmitter power at Saturn and a late arrival probe at Uranus provide the worst case requirement for the battery. The nominal power demands of the various subsystems are tabulated.
Subsystem Power (w) Subsystem Power (w)
Transmitter 25/69.4 NMS 14.0
DHS (pre-entry/descent) 7.2/8.6 Pressure 1.3
Temperature 1.4
Engineering Housekeeping 2.0 Accelerometer 2.8
Distribution losses 9.7 (max)*
^Assumes 5% of total power. IV-31
The margins that are applied consist of ±29 minutes trajectory uncertainty at Uranuss ±5% descent time uncertainty, and 10% overall uncertainty in power levels. These calculations result in a total required energy storage of 160 w-hr and a peak demand of 105 w at 28 volts. Using an energy density of 13.6 w-hr/kg and 0.138 w-hr/ cm s the entry/descent battery physical specification is 2.41 kg and 1200 cm3.
The physical specification of the power and pyrotechnic subsystem is based on a summation of the estimated required equipment as listed.
SCR - two per (dual bridgewire) event
Relays - one per event
Capacitor banks - six discharges allowed per bank; sufficient banks must be allowed to provide separate banks for simultaneous (less than 15 sec apart) discharges
Electronics - engineering estimate.
The number of capacitor banks (8) is controlled by the number of simultaneous events (4) for these missions. The total of 15 events requires 30 SCRs and 15 relays. The physical characteristics of the pyrotechnics is listed .
Number Component Size/Weight Required Total Size/Weight
Capacitor bank 164 cm3/81.9 g 8 1310 cm3/0.66 kg
Relays 14 cm3/34.1 g 15 210 cm3/0.51 kg
SCR 13.6 cm3/14.1 g 30 410 cm3/0.42 kg
Electronics 1230 cm3/0.91 kg 1 1230 cm3/0.91 kg
The size and weight of each element includes 10% to 15% for packaging
IV-32
ATTITUDE CONTROL
The attitude control dynamics are minimal for the design configuration considered for the Saturn/Uranus probe. The errors involved are:
1) initial error due to spacecraft pointing;
2) drift error accumulation during period between separation and spinup due to tipoff rate;
3) error due to tipoff rate combined with spinup;
4) error due to spin>-up jet misalignment.
Only momentum vector errors are of significance in this analysis since nutation errors are damped out and the spin axis (principal moment of inertia) will align with the momentum vector. The momentum vector errors involved are discussed in the following paragraphs.
1) Drift error - Assuming an initial tip off rate of one-half deg/ sec about a nominal X-axis, the spin axis is displaced in the minus y direction.
6T = WT 'D 9T = d r l f t e r r ° r
W = tip off rate = 0.5°/sec t = drift period =0.5 sec
6T = 0.25°
o
3) As indicated in Chapter III, Section D4 of this volume, the error may be calculated from knowledge of the spin momentum, transverse moment of inertia and spin torque.
W = 0.52 rad/sec I =4.9 kg-m2
s
m = 3.4 N-m I = 4.2 kg-m2
q = p2/ml = 0.465 P (o)/P = 0.795 (deg)
e (p) < 0.802 e (p) < -0.19° x - y
IV-33
Since the combination of the drift and spinup/tipoff error is not random, the error due to (2) and (3) combined is;
9 (P) < 0.802°
9 (P) < -0.44° = -0.25° -0.19° y
4) The analysis of Chapter III Section D4 of this volume indicated the spin torque vector offset 6 = 0.136°. The total rss momentum vector error is then
6 = 0.92°.
IV-34
STRUCTURAL AND MECHANICAL SUBSYSTEMS
The approach to design of the structural and mechanical subsystems for the Saturn/Uranus baseline common probe is essentially the same as that presented in Volumes I and II of the report for the earlier outer planet atmospheric probes. This probe design, however, uses a spacecraft-deflect mode for the trajectory deflection maneuver rather than a probe-deflect. Therefore, this configuration deletes the solid propellant deflection motor and the spin-despin-precess service module of prior configurations. The probe utilizes an aeroshell/heat shield structure to withstand the high heating of planetary entry at either Saturn or Uranus, and to provide an entry ballistic coefficient satisfying science mission requirements for initiation of atmospheric descent. The aeroshell structure and heat shield remaining after entry are jettisoned after completion of entry and before planetary atmospheric descent. This approach rids the descent probe of the heat shield and its absorbed heat load, and eliminates contamination of the descent probe mass spectrometer by carbon and other elements originating from the heat shield. The descent probe is separated from the aeroshell after entry by means of a separation parachute, and then descends on a second parachute through the planetary atmosphere to accomplish the scientific mission.
The entire probe is designed to withstand the entry loads encountered during planetary entry at Saturn and Uranus. For the selected missions, the entry heating is most severe for Saturn entry, while structural loadings are more severe for Uranus entry. The structural and mechanical design for the common probe are described in the following paragraphs.
Configuration and General Arrangement
The common Saturn/Uranus probe is a two-stage configuration using one configuration for the cruise/coast/entry phases of flight and a second configuration for planetary atmospheric descent. These configurations are depicted in Figure IV-10. The cruise/coast/ entry configuration is depicted on the right hand side of the figure, and the descent configuration on the left. The total weight breakdown for the entire probe before entry, and the sequential configuration weights are presented in Table IV-12. The descent probe contains all the scientific instrumentation plus the supporting electrical and electronic components necessary to successfully complete the descent science mission. That probe equipment required prior to entry, but not for descent, is located within the aeroshell/heat shield structural assembly.
IV-3 5
The descent probe has a diameter of 47.0 cm (18.5 in.) and weighs 40.5 kg (89.1 lbm) as it descends on the parachute. The equipment within the descent proba is protected from the hostile cold environment after planetary atmospheric entry by a layer of low density foam insulation within the descent probe shell structure. In addition, the equipment support deck is thermally isolated from the cold outer structure by a thermal standoff isolator ring at the periphery of the deck located between the deck and the outer structure.
The mass spectrometer inlet is located at the center of the lower cone" of the entry.probe such that it is exposed to the incoming uncontaminated atmosphere of the planet. The temperature gage is located on the cylindrical side of the probe, with the temperature probe extended from the shell after separation of the descent probe from the entry aeroshell. The pressure ports for the pressure instruments are also on the side of the probe, and these ports are continuously open. The descent antenna and a dual parachute system, for descent probe separation from the aeroshell and for atmospheric descent, are located on the aft end of the descent probe.
The entry probe consists of the descent probe contained and encapsulated within the aeroshell/heat shield entry structure. The nose of the entry probe has been baselined as a 60° half angle conical shape. The nose cone is capped with a spherical segment having a ratio of nose radius to cone base radius (R^/R ) of 0.2. The
aft closure of the probe is a spherical segment. The entry probe has a diameter of 86.8 cm (34.2 in.) and a weight of 89.0 kg (196.1 lbm). The ballistic coefficient of the entry probe is 102 kg/m (0.65 slugs/ft2), meeting science entry requirements. The total weight breakdown for the entire probe is presented in Section H4.
A graphite forebody heat shield protects the aeroshell and the descent probe during planetary entry heating. An ESA 3560 ablator protects the afterbody structure from entry heating. The entire aeroshell/heat-shield structure and its contained equipment is jettisoned after planetary atmospheric entry. This is accomplished by opening the three-segment afterbody cover, releasing the clamp band retaining the descent probe, and deploying a separation parachute to drag the descent probe from the aeroshell. The aeroshell then follows a separate trajectory from that of the descent probe, eliminating interference during descent.
IV-3 6
Page intentionally left blank
PRECEDING PAGESBLANK MOT FILMED
Table IV-12 Baseline Common Saturn/Uranus Probe Weight Breakdown
SCIENCE
TEMPERATURE GAGE PRESSURE TRANSDUCER ACCELEROMETER SENSOR
ELECTRONICS (CONVERTER) NEUTRAL MASS SPECTROMETER ANALYZER
ELECTRONICS PUMP BALLAST TANK
POWER & POWER CONDITIONING
POWER CONDITIONER POWER DISTRIBUTION BOX POWER FILTERS ENTRY BATTERIES POST-SEPARATION BATTERIES COAST BATTERY
CABLING
INNER PROBE EXTERNAL STRUCTURE
ATTITUDE CONTROL SYSTEM (LESS PROPELLANT)
ACS SYSTEM & TANKS NUTATION DAMPER COAST TIMER
COMMUNICATIONS
PRE-ENTRY ANTENNA POST-ENTRY ANTENNA RF TRANSMITTER RF ANTENNA SWITCH
PYROTECHNIC SUBSYSTEM
PYRO ELECTRONICS PYRO CAPACITORS (PROBE) PYRO CAPACITORS (EXTERNAL) PYRO RELAYS (PROBE) PYRO RELAYS (EXTERNAL) SCR (PROBE) SCR (EXTERNAL) PYRO SQUIBS PYRO THRUSTER
DATA HANDLING
DATA HANDLING SUBSYSTEM MEMORY BANKS
STRUCTURES AND HEAT SHIELDS
DESCENT PROBE STRUCTURE EQUIPMENT SUPPORT DECK BASE COVER AEROSHELL (2 l b FOR PAYLOAD RING) FORWARD HEAT SHIELD ( 1 6 . 6 l b ABLATED
DURING ENTRY) AFT HEAT SHIELD (BASE COVER)
' MECHANISMS
PIN PULLER LATCHES & BANDS MAIN PARACHUTE SECONDARY PARACHUTE (DESCENT) CLAMP SEPARATORS
WEIGHT
kg
0 . 4 5 0 . 6 8 0.59 0.91 1.82 2.73 0.45 0.45 8.08
0.22 0.50 0.91 2.41 0.25 0.41 4.70
2.36 1.50 3.86
1.63 1.09 0.12 2.84
0.45 0.45 2.73 0.23 3785"
0.91 0.33 0.33 0.17 0.34 0.14 0.28 0.27 3.61 6.38
2.13 0.30 2.43
3.92 3.54 4.04 4.99
16.43 1.08
34.00
0.66 0.91 3.45 0.54 0.47 6.03
lb
1.00 1.50 1.30 2.00 4.00 6.00 1.00 1.00
17780
0.50 1.10 2.00 5.30 0.55 0.90
10.35
5.20 3.30 8.50
3.60 2.40 0.25 6.25
1.00 1.00 6.00 0.50 8.50
2.00 0.72 0.72 0.37 0.75 0.31 0.63 0.60 7.95
14.05
4.70 0.65 5.35
8.65 7.80 8.90
11.00
36.16 2.38
74.89
1.44 2.00 7.60 1.20 1.04
13.28
THERMAL
EXTERNAL INSULATION BLANKET ) (FORWARD HEAT SHIELD) (
EXTERNAL INSULATION BLANKET 1 (BASE COVER) >
PROBE HULL INSULATION (INTERNAL) ISOTOPE HEATERS ENVIRONMENTAL TANK & N2
PROPULSION
ACS PROPELLANT
TOTAL
15% CONTINGENCY
1 . PRE-ENTRY WEIGHT
ACS PROPELLANT 15% CONTINGENCY
(196.08 - 0.03 = 196.05 lb ) 89.02 - 0.01 = 89.01 kg
2. POST-ENTRY WEIGHT
FORWARD HEAT SHIELD (ABLATED) AFT HEAT SHIELD (ABLATED) = 0 FORWARD INSULATION BLANKET I AFT INSULATION BLANKET i PRE-ENTRY ANTENNA
15% CONTINGENCY
(196.05 - 23.20 = 172.85 lb) 89.01 - 10.63 = 78.38 kg
3. INITIAL WEIGHT ON PARACHUTE
SEPARATION PARACHUTE 15% CONTINGENCY
(172.85 - 2.99 = 169.86 lb ) 78.48 - 1.36 = 77.12 kg
4 . WEIGHT ON PARACHUTE
PYRO THRUSTERS PYRO CAPACITORS PYRO RELAYS PYRO SQUIBS AEROSHELL FORWARD HEAT SHIELD (NOT ABLATED) BASE COVER HEAT SHIELD & BASE COVER SEPARATION PIN PULLERS
(BASE COVER S PARACHUTE) LATCHES & BANDS ISOTOPE HEATERS EXTERNAL CABLING NUTATION DAMPER ACS SYSTEM SEPARATION BATTERY CLAMP SEPARATORS
15% CONTINGENCY
(169.86 - 80.75 - 89.11 lb ) 77.12 - 36.66 = 40.46 kg
WEIGHT
kg
1.27
1.59 1.82 0.54 5727
0.01
77.41
11.61 89.02
0.01
—
7.44
1.27
0.54 9.25 1.38
10.63
1.18 0.18 T73B"
3.61 0.33 0.34 0.18 4.99 8.98 5.12 0.66
0.91 1.82 1.50 1.09 1.63 0.2b 0.47
31.88
4.78 36.66
lb
2.80
3.50 4.00 1.20
11.50
0.03
170.50
25.58 196.08
0.03
—
16.37
2.80
1.00 20.17 3.03
23.20
2.60 0.39 2.99
7.95 0.72 0.75 0.40
11.00 19.79 11.28 1.44
2.00 4.00 3.30 2.40 3.60 0.55 1.04
70.22
10.53 80.75
o
rt a4
H-
CO
-a
i-i n>
CO
CO
c M ro
M o to
CX
s H- rt P4
to
a ro
era
M
H- cr
i-1 ro
-a
ro
a to
i-1
rt
v;
H-
3 S3
ro
H-
era
a4 r
t •
H-
CO
(X
H-
CO
o a CO
CO
ro
CX
H-
3 CZ)
ro
n rt
H-
O 3 C-i
• H
3"
ro era
to
CO
ro
CO
• H
3"
H-
CO
H-
(0 to
o o o 3 -a
h-1
H-
CO
a4
ro
a- to
CO
CX
Td
ro
CO
n ro
a rt
id
i-i o a4 ro
CO
rt i-t a o r
t C
i-i ro
n to
3 to
n o ro
Td
r
t
to
H
rt o Mi
-a
i-t
o
ro
O o *• o o o a
"•
-^
3 N
•~s
ro
cr
to
^ CO
^^
(_i.
a CO
rt to
rt ro
3 rt
i-t
^ rt o CX
ro
(-• to
cr >
•<:
ro
rt a*
ro
M 3 to
M o o 3 rt i-l
O
H*
•* to
3 C
X
ro
3 rt i-t
v: o Ml
•a
H-1
to
a ro
rt to
i-i
^
CX
ro
CD
o ro
3 rt
M H
-r
t
H-
CO
rt ro
3 Td
o •-i to
i-i
M
^ -a
f-s
ro
CO
CO
c H
H-
N ro
rx
rt O to
M ro
< ro w
<
ro
a r
t P
4
O a era
a4
rt a4 ro
ex
ro
CO
o ro
3 rt •a
H
O cr
ro
H-
CO
< ro
a rt ro
cx
cx
a H
H-
3 era
h"
0
O
Mi
to
•a
Td
i-
i o •A
H- 3 to
rt ro
M
vj
M to
a ro
rt to
i-i
V! to
r
t 3 o CO
•a
p4 ro
ric
ro
3 rt
i-t
"<!
O
Mi
rt a4 ro
•a
CO
rt
l-t e o rt a i-t to
H- o o 3
i-"a
to
a ro
r
t to
i-t
•-=: to
rt 3 o CO
•a
p4 ro
if ro o a ro
a r
t to
H-
to
l-t era
ro
i-1
•<!
CO
a Td
•a o i-t rt
rt P4 ro
H-
i-t O -3
a •3
ro
H-
era
a4 r
t
CX
a i-i H-
3 era D
O
3 -a
M
ro
r
t ro
rt a4 ro
CX
ro
co
o ro
a rt •a
it o a4 ro
CO
rt
l-f e o rt C
H
to
H1
O
O 3 l-h
H
-0
Q
C
i-i to
rt
H-
O 3 H a4 ro
CD
ro
rt -3
o
CO
a4 ro i-t
H
-3
H'O
Q
i-1 • >
Hi o l-t •a
to
i-t a.
3 o CO
ro
Mi
to
H-
i-t H-
3 era
to
3 cx
to
Ml
H-1
to
rt P4 o 3 ro
"-<
o o 3 cr
to
Ml
rt
O
h->
O
CD
C
l-t ro M
l l-t
to
3 ro
• H
3*
H-
CO
Ml o I-
i •3
to
n CX
H
H-
3 cm
Ml
>-i to
3 ro
H-
3 rt C
M 3 3 to
rt ro
CO
•3
H-
rt P4
rt a4 ro
to
ro
f-t o i
cr
^ 3 c i-1
rt
H-
Td
M
ro
o o 3 -a
i-t ro
CO
CO
H- o 3 M o 3 era
ro
i-t
o 3 CO
Ml
l-t o 3 r
t a4 H
-CO
l-(
H-
3 era
rt o rt 3* ro
Ml o M S
to
i-i
(X
(X
ro
CO
n ro
3 rt
Td
H
o cr
ro
r>
^ M
H-
3 rx
>-i
H- n to
M S
to
i-1
M
CO
w •3
H
-r
t a4
rt P4 ro
H- a ro
M
rt
H- to
I-1
o to
rx
CD
rt t to
3 CO
Ml ro
H
l-t
ro
(X
(X
H-
H
CD
n rt (-1
^ o 3 to
o H-
K n a 3 Ml
ro
i-t ro
3 rt
H- to
I-1
(X
(D
O ?r
i CD
a •a
•a
o ii
rt
f-i H-
3 era
to
rt
rt to
o a4 ro
rx
rt O
rt a4 ro o 3 O
i-t
CO
a CD
T3
ro
3 cx
ro
(X
Ml
i-t o 3 r
t P4
H-
CO
CX
ro
n W
H
P4 ro
(X
ro
o •^
CO
rt •1
a o rt a i-i ro
H-
3 rt a i-i 3 i-i ro
CO
rt
CO
to
CO
o M
H-
CX
-a
i-1
to
rt ro
• >
i-1
i-1 o o 3 1
3 O 3 ro
3
3 ro
3 rt
CO
a •a
-a
o i-t
rt
(X
ro
n f-4
H-
CD
to
3 H-
3 rt ro
rt e
ra
CO
»*
33
H-
rt P4 3 H-
3 O
K ro
X o ro
-a
rt
H-
O P
CO
to
i-t ro
3 o a 3 rt ro H
to
h
-1
1 H
H- cr
I CO
rt
H-
Ml
Mi
ro
a ro
rx
cx
H-
CD
O 3 to
o a4 H
-3 ro
tx
H
P4 ro
CO
rt i-t c <->
rt a i-t to
h-> o
o 3 l-h
H
-
era
c •-i
to
rt
h>- O
3 H-
CO
CD
P4 o 5!
3 H-
3 •T|
H-
era
a •-{ ro
M < i M
M • H
P4 ro
ro
Ml
,0
l-t o
rx
3 a H
-
-a I
H
P4 ro
(X
ro
(0 n ro
3 rt -a
i-t o cr
ro
H-
CD
cx
ro
CD
H-
era
3 ro
cx
ro
3 rt
H-
l-t ro
i-1
^ o l-h
--J
O
~J
Ul 1 i-3
ON
to
M a 3 H-
3 a 3 to
M
M o ^
to r
o ro
3
H
rt
O
i-i
CO
•-<
P4
ro to
M
3
(-•e
ra
M
H-
ro
CD
•
CX
)
• H
C
J\
3*
CD
><!
o M
O
O
i-
i c
ni-
i ro
rz;
co
-^•a
3
o K
>3 C
X
•-s
H-
M
3 oo
era
v o
cx
o^
O
3 to
•a
3 CD
H
-
MI
n
- -a
to
i-t
rt r
o CO
>a
co
ro
a to
i-t
fc
ro
i-1 t
o o
o to
rt
rx
p.
• a era
o a rt a4
ro r
t P4 ro
-a
ro
to
-^
cx
ro
o ro
t-1 ro
H to
rt
H-
O 3 H-
CD
00
ON
O
l
era
*•
H-
3 O
h-1
a cx
H-
3 era
to
Cn
o
CX
H
-CO
-a
ro
I-I
CO
H- o a H-
3 rt 3* ro r
t P
4 ro
ro
3 rt i-t <
to
3 em
H> ro
CO
CD
i-1 ro
n rt ro
cx
^-v
O
i
o o
>~
»»
to
3 cx
rt P4 ro
n o o t-> to
rt 3 o CO
•a
P
4 ro
K
H-
O 3 o cx
ro
M
*
Ml
o l-t ro
3 rt i-i
^ o Ml
rt P4 ro
•a
(-•
to
3 ro
rt <=:
•t
to
3 a CO
u rt
S4 ro
3 o CD
rt
CO
ro
< ro
I-I ro
ro
3 <
H-
H
o 3 3 ro
3 rt • •n
O
t-
t
C
ro
H-
era
p4
rt
-a
ro
3 to
•3
H-
rt a4
CO
rt to
3 CX
H
-3
HW
r
t H
-CD
CD
• H
P4 ro
CO
rt
i-i a o rt a n
ro
o
ro
X
rt ro
i-t
3 to
M* to
rt 3 o CO
-a
P4 ro
i-t
H-
O
MI
-c)
rt a4 ro
ro
3 rt
H-
f-i ro
>a
M
O a4 ro
H-
CD
cx
ro
CO
H-
era
3 ro
cx H
ro
CO
CO
a i-. ro
CO
to
3 cx
rt P4 ro
o o 3 CD
ro
XI a ro
3
rt a i-t ro
CD
to
i-t ro
<
ro
3 rt ro
cx
w o cr
<
H
- to
rt
H-
3 era
rt P4
CD
3 ro
ro
cx
Mi
o i-i
CO
rt
i-l a o rt a I-I
to
M o to
Td
to
cr
H-
M1
H-
rt
rt
^ O
Ml
rt
H-
O 3 to
I-1
H-
3 3 o CO
rt H ro
CO
-a
ro
n rt
CD
• td
o rt a4
rt P4 ro
ro
3 rt
i-l
^ to
3 cx
cx
ro
CD
o ro
3 rt
-a
I-I o cr
ro
CO
rt
M a o i
n p4 ro
CO
rt ff a n
rt a n to
M
(X
ro
CO
H-
era
3 o Ml
rt P4
ro
w
to
rt a i-t
3 -
-•
^
C3 »i
to
3 a CO
o o 3 e o 3 Td
i-
i o a4 ro
H-
CO
n o 3 <
ro
3 I
CO
r
t H
e o rt a M to
M o
ro
CO
H-
TO
3
3 H
-ro
3
O
CD
ro
a CD
M
co
to
to
rt
l-t
H
-
^ o
• 3 •a
H
H- o l-t
rt o ro
3 rt
i-i
V! >*•
P4
O
•3
ro
<
ro
i-t
<#
rt P4
H-
CO
H-
CO
3 o rt cr
ro
M
H- ro
<
ro
cx
rt o cr
ro g
3 ro
r
t ^ H- n to
M
cx
I-I
to
era
• > 3 to
i-1
rt ro
I-I
3 to
rt
H- <
ro
to
•a
-a
M o to
o a4
•3
o a i-1 rx
cr
ro
rt o
C_
l.
ro
rt
rt
H-
CD
O 3 rt
•a
i-i o o- ro
rt
rt
CO
a4 "a
ro
xt H
O
p"-a
n
o
a era
p
4 ro
3 rt H
^ to
3 CX
CD
p4 o a i-1
CX
to
cr
M to
rt ro
to
•3
to
S •3
H-
rt P4 o a r
t ro
X o ro
CO
CO
H- <
ro
to
CO
P4
^:
CD
1
o CO
ro
rx
cx
ro
CO
H-
era
3 «*
rt P4 ro
H- a CO
a H" to
rt
H-
O a !-••
CO
to
M
I-1 o -3
ro
cx
rt o l-t § to
p- 3 o 3 rt P4 ro H
-3 a o M
N
h
-1
ro
CO
« to
3 cx
rt
p4
CD
ro
3 rt i-l
^ to
3 rt ro
3 3 to
•a
ro
3 CD
rt
M to
rt ro
rt S4 ro
H-
3 CO
a (-• to
rt H-
O 3 • M 3
Ml
M
H-
era
P4
rt • O
3 1-"
"<
rt P4 ro
•a
•f o cr
ro
to
rt
rt to
o S4
3 ro
a rt CO
rt O
rt P4 ro
o to
n
H
H- ro
ii
CO
Td
to
n ro
o n
to
Ml
rt «•
rt a4
ro o M
i
rt P4 ro
Td
K
o a- ro
to
3 cx
H-
rt
CO
n o 3 Td
o 3 CD
3 rt
CO
CX
a H H-
3 era
o H a H-
CD
ro
—
o o to
CD
rt
Td
a4 to
CO
ro
CD
o Mi
M to
^ ro
i-t
p4
H-
era
P4 i
Td
ro
^ Ml o l-t 3 to
3 o ro
H-
3 CO
a i—•
to
rt
H-
O 3 o4
H
P4 ro
o n a H-
CO
ro
-^
o o to
CD
rt *-
ro
3 r
t i-
t •<!
Td
H
O
cr
ro
H-
CD
D
Q
3 !-
• T
d
to
3 •v
ro
rt
rt
O to
CD
CD
H-
CO
rt
rt P4 ro
i-t 3 to
M
O
O 3 rt rol
h-1
ro
rt
CD
(-•
V! ro
a n (-
• o CO
ro
cx
53
H-
rt
P4
H-
3 to
3 a M ti-
Page intentionally left blank
The planetary entry aeroshell configuration is a ring frame stiffened monocoque structure with an end closure box frame, incorporating a "saddle" ring frame at approximately the mid-diameter of the conical surface to accept the descent probe. The analysis of this structure is presented in Appendix 0 of Volume III. The structure has been designed of titanium alloy 6A&-4V for this study. A more refined analysis of the temperature-time history may show that an aluminum aeroshell will be compatible with the anticipated temperatures. If so, a weight saving of some 1.8 kg (4 lbm) may be achieved.
The afterbody cover is a three-segment 7075-T6 aluminum structure hinged at the base of each segment. The three segments are shell structures reinforced along the free edges and locally at the points of attachment of the development thrusters. These are opened after entry for descent probe separation.
i
Mechanisms
The primary mechanisms of the Saturn/Uranus probe are the spacecraft/probe separation mechanism, the aft closure opening thruster system, the descent probe retention/release mechanism, the parachute deployment system, and the parachute release mechanism. The latter two are described in the parachute system description of Section H5; the first three are depicted in Figure IV-11.
The spacecraft/probe separation mechanism is simply a bolted joint, three-point attachment to the spacecraft. Pyrotechnically actuated nuts retain the attachment of the probe to the spacecraft and are activated for probe release. Three matched-performance springs push the probe away from the spacecraft at separation, imparting a delta velocity of 1 m/s (~3 fps) .
The base cover is opened in three segments, after completion of atmospheric entry, to permit separation of the descent probe from the spent aeroshell/heat shield assembly. The opening is accomplished by activation of three pyrotechnic thrusters, one attached to each of the three base-cover segments. The thrusters open the segments wide enough to permit separation of the descent probe, reach the end of their stroke, and lock the aft cover segments in the open position. Before opening the base cover segments, the segments are held in the closed position by pyrotechnic pin pullers that are activated just prior to opening the base cover segments. Not shown in Figure IV-11 is the pre-entry antenna. This unit is
IV-43
attached to one of the aft cover segments, and after entry heating separates with the cover segment as it is opened.
The descent-probe retention/release mechanism is simply a clamp ring around the descent probe with tension links attaching the clamp ring to the aeroshell structure, and thereby holding the descent probe in the aeroshell. The descent probe is released by dual pyrotechnic clamp separators releasing the tension of the clamp ring. A separation parachute then pulls the descent probe from the aeroshell.
Mass Properties
The weight breakdown for the probe is presented in Table IV-12. This table presents the weights tabulation for the baseline configuration Saturn/Uranus probe by subsystems, which are summed to establish the total weight of the probe. The last portion of the table reduces the weight sequentially, in keeping with the normal mission sequence of events to indicate the probe weight for the various phases of the mission.
The moment of inertia data for this probe has been computed for use in attitude control propulsion propellant calculations and as an indicator of stability about the spin axis. The moment of inertia data are tabulated.
Descent Probe Entry Probe
I . 0.612 slug-ft2 3.61 slug-ft2
spin 6 &
I .„ , , 0.475 slug-ft2 3.07 slug-ft2
pitch/yaw b B
It is seen that the moment of inertia about the spin axis is approximately 17.5% larger than that of pitch and yaw as presently configured, and should be adequate. This value can be increased to approximately 20%, if desired, by rearrangement of equipment in the aeroshell.
Parachute Subsystem
The parachute subsystem consists of two parachutes that are deployed in sequence. A large main parachute is deployed after completion of atmospheric entry and is used to separate the descent probe from the spent heat shield/aeroshell structure. This parachute is then jettisoned and a descent parachute is deployed and used for the scientific mission descent through the atmosphere.
IV-4 4
The size of the main parachute is large enough to assure prompt separation of the descent probe from the entry probe assembly. The aeroshell/heat shield assembly with its contained equipment weighs 36.7 kg (80.8 lbm) after entry (with the descent probe removed). It has a ballistic coefficient of approximately 59.5 kg/m2 (0.38 slugs/ ft2) when traveling subsonic (Mach 0.7). To assure separation of the descent probe, the main parachute selected gives the descent probe a ballistic coefficient of 18.8 kg/m2 (0.12 slug/ft2). This value is somewhat arbitrary, but will assure prompt separation. The resulting parachute diameter for a 40.5 kg (89.1 lbm) descent probe is 2.29 m (7.5 ft) in a disc-gap-band parachute configuration and weighs 3.4 kg (7.6 lbm). The parachute is exposed to dynamic pressures between 1707 and 2546 N/m2 (35.7 to 53.0 psf) at approximately Mach 0.7 at deployment, depending upon the planet and the atmospheric model at entry. The descent probe after separation is decelerating at a rate of 3.9 to 6.6 g relative to the spent aero-shell, also dependent on the planet and the atmosphere.
The descent parachute is a circular parachute, selected to provide a ballistic coefficient of 110 kg/m2 (0.7 slug/ft2) to meet the atmospheric descent science requirements. The resulting parachute has a diameter of 0.73 m (2.4 ft) and a weight of 0.54 kg (1.2 lbm).
The main parachute is deployed by a pyrotechnic mortar which ejects the entire parachute canopy by means of a sabot. The parachute is released by a pyrotechnic pin puller in the parachute riser system and in being released, pulls out the descent parachute. Both parachutes are stowed in fiberglass canisters for RF transparency to the descent antenna.
Both parachutes are designed of Dacron fabric and are maintained at "room" temperature until deployment. It appears that the Dacron material, which has been tested to liquid nitrogen temperature (77°K), will suffice in spite of the extreme low temperature. However, sufficient doubt exists to warrant a test program of the material at some time in the future.
6. Heat Shield
The forebody heat shield is designed of ATJ graphite, and is based on the work performed by M Tauber of NASA/AMES for the Outer Planets studies. The heat shield mass fraction is established by entry of the atmosphere of the planet Saturn, which is more severe than the entry at Uranus. The heat shield mass fractions requirements for Saturn and Uranus were presented versus entry angle in Chapter VI, Section B.8.e of Volume II of this report. It can be seen from
IV-45
C-3
this data that, for an entry angle of 30° plus a dispersion error of 5°, the heat shield mass fraction is 0.21. This results in a heat shield weight of 16.4 kg (36.2 lbm) for a probe having an entry weight of 89.0 kg (196 lbm), including heat shield.
In keeping with the heat shield design studies performed by M. Tauber, the heat shield mass fraction and initial thickness is based on maintaining a final ablator thickness of 4 mm (0.16 in.) after planetary entry and providing a low density carbon felt insulator between the hot backface and the aeroshell. In this concept, the heat shield must be rigidly supported to resist launch loads, and rigidly supported (but not restrained against thermal expansion) or released to rest against the carbon felt insulation during entry. In computing weights data for the probe, the released heat shield approach has been assumed. However, all the mechanical details of this approach have not been evaluated.
The afterbody heat shield is essentially the same as that described for the Saturn probe in Chapter VI, Section B.8.e of Volume II of this report. It is based on the analytical approach for afterbody heat shields described in Chapter V, Section A.8.c and tabulated data from Chapter B.8.3 of Volume II. This analysis assumes the amount of ablator to be used shall protect the understructure to a temperature of 422 °K (300°F). The resulting ablator weight is 1.37 kg/m2 (0.28 lbm/ft2) using ESA 3560 ablator material.
The afterbody ablator is pre-split along the parting lines of the afterbody cover segments to facilitate opening of these segments after entry. This is not a new concept. A modified Gemini reentry vehicle was launched by Martin Marietta Corporation in November 1966 as part of the MOL/HSQ (Manned Orbiting Laboratory/ Heat Shield Qualification) program, in which the re-entry heat shield had an access door pre-cut through the heat shield.
7. Alternative Antenna Approach
Since two identical antennas are used for pre-entry and descent phases of the mission, it is logical to question whether one antenna could handle both jobs. As presently conceived, the probe design places one antenna external to the multilayer insulation blanket to transmit pre-entry data, and one antenna within the afterbody cover to transmit during atmospheric descent. This descent antenna is exposed for transmission after separation of the descent probe from the entry aeroshell. A comparison was made using a single antenna; the descent antenna on the probe was retained and the pre-entry antenna was eliminated.
IV-46
For the single antenna configuration, it is necessary to provide RF transparency to the afterbody of the entry probe. If an RF transparent afterbody cover and ablator is used, then a portion of multilayer insulation approximately 0.37 cm2 (4.0 ft 2), prior to data transmission. The time of separation of the insulation would be of the order of one hour prior to entry for late planetary arrival. It it is assumed that the RF transparent ablator on the afterbody has an emissivity of £ = 0.70, which is probably conservative, then the heat loss from the afterbody by radiation is 165 watts, or approximately 165 w-hr for one hour. This value would be reduced to approximately 80 w-hr by use of a low-conductivity afterbody material.
The loss of heat could be accommodated by a decrease in temperature of the probe (undesirable), or by adding extra battery weight and using resistance heaters to compensate for the loss. To compensate for this heat loss using batteries and heaters would result in a weight penalty of approximately 1.4 kg (3.0 lbm). Thus, a first-cut comparison of the penalties of the two configuration options is as follows:
1) a weight increase of 1.4 kg (3.0 lbm) of batteries for the single antenna, compared to 0.7 kg (1.5 lbm) for the extra antenna and antenna switch of the two-antenna configuration;
2) an extra event for the single-antenna configuration (jettisoning of the insulation blanket);
3) a potential antenna interference problem for the single antenna due to the proximity of the undeployed dense-packed parachute;
4) RF transparency requirement for the base cover of the single-antenna configuration. (The weight comparison of RF transparent afterbody and ablator versus non RF transparent configurations has not been made.)
In keeping with the philosophy of maintaining the common Saturn/ Uranus probe similar to the previous design approaches used in this study, the two-antenna configuration has been retained although it is recognized that the single-antenna warrants further study on future design activities.
IV-4 7
I. PROPULSION SUBSYSTEM
With deletion of the requirement for the probe deflection maneuver, the propulsion system becomes quite simple. The solid propellant deflection motor is deleted from the probe and with it, the need for a high rate of spin to stabilize the probe during motor burn. Thus, this common Saturn/Uranus probe needs only a small attitude-control spin-up system to provide the necessary probe attitude stabilization from probe/spacecraft separation until planetary entry. This system is small and light enough to be included in the entry probe. The propulsion system for probe attitude control is a cold gas (GN2) system, basically the same as that reported in Volume II for the other planetary atmospheric probes. The spin rate to be imparted to the probes for attitude stabilization has been selected as 0.52 rad/sec (5 rpm). This spin rate is somewhat arbitrary, but will provide pointing attitude stability prior to entry, with minimum interference with dynamic forces of aerodynamic stabilization at entry.
The selected system, depicted in Figure IV-12, uses a stored high pressure gaseous nitrogen supply that is vented through a pair of spin nozzles to provide the desired spin torque. The system is activated by opening of a parallel pair of pyrotechnic valves that vent nitrogen through a pressure regulator to the spin nozzles. The spin-up torque is terminated by activating two series-connected pyrotechnic valves. This system provides a predictable spin rate even if some of the stored nitrogen gas leaks away.
For a probe having a moment of inertia about its spin axis of 5.02 kg-m2 (3.7 slug-ft2), the amount of gas required to spin the probe to 0.52 rad/sec (5 rpm) is approximately 0.015 kg (0.033 lbm) including 50% reserve. The pressure bottle weight is conservatively estimated to be 0.09 kg (0.2 lbm).
IV-48
—D
3<H
?
O
C)
-©
SPIN NOZZLES
1.0 1bx
PROBE
SPIN AXIS MOM
ENT OF INERTIA
PROBE
COAST
SPIN RATE
SPIN NOZZLE MOMENT ARM
SPIN IMPULSE
PROPELLANT SPECIFIC IMPULSE
PROPELLANT REQUIRED (5
0% EXC
ESS)
SYSTEM HARDWARE WEIGHT
TOTAL
SYSTEM WEIGHT
= 5.02 kg-
m2 (3.7 sl
ug-f
t2)
= 5 rpm
= 38.2 cm (1.25 ft)
= 6.89 N-sec (1.55
lbf-s
ec)
= 707 N-sec/kg
(72 sec)
= 0.015 kg (0.033 lb )
= 1
.54
kg
(3.4
1b
m)
= 1
.63
kg
(3.6
3 lb
)
LEGEND
© i $
X-D
-e-
: PR
ESSURE TRANSDUCER
FILTER
NORMALLY CLOSED PYROVALVE
PRESSURE REGUALTOR
NORMALLY OPEN PYROVALVE
< I F
ig.
IV-1
2 A
ttit
ude
Con
trol
P
ropu
lsio
n Sy
stem
The characteristics of the system are depicted in Figure IV-12 and the components weights and quantities are listed.
Component
GN2 bottle (1)
Fill valve (1)
Pressure transducer
Squib valves (4)
Filter (1)
Pressure regulator
Thrusters (2)
Lines
Total
•s (2)
(1)
Weight
kg
0.09
0.09
0.23
0.45
0.18
0.18
0.18
0.23
1.63
lbm
0.20
0.20
0.50
1.00
0.40
0.40
0.40
0.50
3.60
IV-50
J. THERMAL CONTROL SUBSYSTEM
The thermal control subsystem design outlined in Chapter III is used in Task I. To compensate for the cold atmospheres encountered, elevation of the probe coast temperature to a maximum before entry allows optimum leeway for internal temperature decrease during descent. For•consistency of the common Saturn/ Uranus probe design, the warm Saturn atmospheric encounter and the warm Uranus atmospheric encounter, both with late arrival uncertainty, were analyzed to determine a maximum probe equipment coast temperature for entry and descent. Because of the criticality of probe entry temperature, positive heating is achieved by radioisotope and a louver system design is recommended to control probe temperatures before entry. On the basis of the design for descent thermal control, a probe equilibrium arrival temperature of 302.0°K was chosen, and an uncertainty temperature band of ±10CK assumed. Calculations have shown that approximately 18 watts of radioisotope heater power supply with an Iridite 14-2 external finish would achieve the desired entry condition.
Thermal control for the atmospheric descent portion of the mission relies on the probe coast/entry temperature, internal probe foam insulation, and sufficient probe thermal inertia to survive the very cold descent environment encounter. On the basis of the thermal control design studies performed and previously outlined in Chapter III, results show that an environmental purge system using 2.5 bars (0.127 kg) of neon gas would be sufficient to offset electrical heating of the probe payload required for the nominal and cool Uranus atmospheres. For all Saturn atmosphere encounters anticipated, a totally passive system would be sufficient for descent thermal control requiring no gas purging or electrical heating.
The parametric results presented in Chapter III indicate the high probe heat losses accompanying vented probe design configurations. For the Uranus cool atmosphere encounter, a completely vented Task I probe design would experience a total heat loss of approximately 390 w-hr to the atmosphere, whereas the neon purged probe would experience a heat loss of only approximately 220 w-hr. Neon gas was chosen over previously identified nitrogen gas due to the cooler probe shell temperature anticipated in Uranus cool atmosphere encounters, and the increased probability of condensation and freezing of the nitrogen gas at the insulation/probe shell boundary.
IV-51
The worst-case atmospheric encounters and uncertainties were analyzed for the Science Advisory Group exploratory payload common probe design. The warm atmospheric encounters were analyzed with late arrival uncertainties and cool atmospheric encounters with early arrival uncertainties. Table IV-13 presents the minimum and maximum equipment temperatures experienced during the mission while figures IV-13 and IV-14 present plotted thermal analysis results. The worst-case conditions establish the probable operative temperature bands to be experienced at each planet. Assuming a 302°K probe arrival temperature, these bands are as follows:
Saturn, °K Uranus, °K
RF transmitter 291-319 274-327
Insulated primary batteries 294-307 290-310
Aggregate internal equipment 280-304 261-310
The baseline arrival temperature uncertainty of ±10°K was established as the required probe design margin and exists at all limits for the Task I design. Note that the large arrival uncertainty at Uranus establishes higher probe temperatures although Uranus atmospheres are cooler than Saturn's atmospheres. For this reason, the large Uranus arrival uncertainties became an important consideration for establishing acceptable pre-entry transmitter power levels. Although simpler design is afforded by using a constant power transmitter for pre-entry and descent, preliminary analyses indicate that over-temperature transmitter conditions would exist before atmosphere pressurization could occur. An additional item or probable constraint identified during the analyses was the effect of the cold atmosphere on the externally mounted RF antenna. Current intelligence indicates that potential thermal stress problems could exist at very low temperatures requiring possible radome protection with auxiliary heaters.
IV-5 2
Tabl
e IV
-13
Min
imum
and
Max
imum
Pro
be T
empe
ratu
res
for
Wor
st-C
ase
Atm
osph
eric
E
naou
nter
s an
d A
rriv
al
Unc
erta
intie
s
PROB
E CO
AST
TEMP
ERAT
URE
RF TR
ANSM
ITTE
R
INSU
LATE
D PR
IMAR
Y BA
TTER
IES
AGGR
EGAT
E IN
TERN
AL.E
QUIP
MENT
PRE-
ENTR
Y TE
MPER
ATUR
ES
RF TR
ANSM
ITTE
R
INSU
LATE
D PR
IMAR
Y BA
TTER
IES
AGGR
EGAT
ED INTERNAL EQ
UIPM
ENT
MAXI
MUM
DESC
ENT
TEMP
ERAT
URES
RF TR
ANSM
ITTE
R
INSU
LATE
D PRIMARY
BATT
ERIE
S
AGGR
EGAT
ED INTERNAL EQ
UIPM
ENT
MINI
MUM
DESC
ENT
TEMP
ERAT
URES
RF TR
ANSM
ITTE
R
INSU
LATE
D PRIMARY
BATT
ERIE
S
AGGR
EGAT
ED INTERNAL EQ
UIPM
ENT
mrn
mrn MAXIMUM
PROBE TEMPERATUR
\\\\\\\\\\\\\ MIN
IMUM
PR
OBE
TEMPERATURI
EQUI
PMEN
T TE
MPER
ATUR
E LI
MIT, °K
225/
350
225/
325
240/
325
250/
345
275/
325
255/
320
345
325
320
250
•275
250
IS
IS
SATU
RN
WARM
AT
MOSP
HERE
(LATE
ARRI
VAL
UNCE
RTAI
NTY)
302°
K
308°
K
304°
K
304°
K
319°
K
307°
K 304°
K
308°
K
304 °VK
297°
K
COOL AT
MOSP
HERE
(EARLY AR
RIVA
L UN
CERT
AINT
Y)
302°
K
305°
K
302°
K
303°
K
314°
K
302°
K
304°
K
291°
K
294°
K
280°
K
URAN
US
WARM
AT
MOSP
HERE
(LATE
ARRI
VAL
UNCE
RTAI
NTY)
302°
K
320°
K
309°
K
310°
K
!327
8KS:
i310
°K:|
:;3io°K:i:i
320°
K
307°
K
294°
K
COOL ATMOSPHERE
(EARLY ARRIVAL
UNCE
RTAI
NTY)
302°K
305° K
302°
K
303°
K
313°K
302°
K
303°K
^274
V°K^
^290°K^
^261°K%
LEGEND:
RF TRANSMITTER TEMPERATURE
— PRIMARY BATTERY WITH 1-CM FOAM INSULATION
•AGGREGATE INTERNAL EQUIPMENT
TEMPERATURE, °K
275
LAUNCH SPACECRAFT SEPARATION
ENTRY NEAR SATURN ARRIVAL
a. WARM ATMOSPHERE ENCOUNTER WITH LATE ARRIVAL UNCERTAINTY (MAXIMUM ANTICIPATED TEMPERATURES)
END OF MISSION
LAUNCH SPACECRAFT SEPARATION
NEAR SATURN ARRIVAL
ENTRY END OF MISSION
b. COOL ATMOSPHERE ENCOUNTER WITH EARLY ARRIVAL UNCERTAINTY (MINIMUM ANTICIPATED TEMPERATURES)
Figure IV-1S Saturn/Uranus Probe Thermal History for Saturn Mission (SAG Exploratory Payload)
IV-54
TEMPERATURE, °K
275
LAUNCH
WARM ATMOSPHERE ENCOUNTER WITH LATE ARRIVAL UNCERTAINTY (MAXIMUM ANTICIPATED TEMPERATURES)
END OF MISSION
LAUNCH SPACECRAFT SEPARATION
END OF MISSION
COOL ATMOSPHERE ENCOUNTER WITH EARLY ARRIVAL UNCERTAINTY (MINIMUM ANTICIPATED TEMPERATURES)
Figure IV-14 Saturn/Uranus Probe Thermal History for Uranue Mission (SAG Exploratory Payload)
IV-5S
PROBE-TO-SPACECRAFT INTEGRATION
Mechanical
The Saturn/Uranus probe integration with the Mariner Spacecraft bus is very similar to that described in Volume II, Chapter VI, Section V.ll of the final report. The physical arrangement is depicted in Figure IV-15. From the figure, it can be seen that the probe is canted at an angle with respect to the spacecraft center-line to minimize interference with the spacecraft rocket engine. The probe is contained within an environmental enclosure to protect the probe from meteoroid impingement and other space environment during the cruise phase. The Mariner rocket engine is moved aft to minimize plume impingement of the engine on the probe enclosure. The engine must control the spacecraft with the probe attached and jettisoned. The line of action of the engine thrust is therefore nominally pointed 2°.outboard such that the thrust vector is pointed halfway between the spacecraft-only and the spacecraft-plus-probe center of gravities.
The probe separation-spring thrust vector is directed to act through the center of gravity locations of both the probe and spacecraft to minimize imparted tumbling motions to each at probe separation.
Thermal
For thermal control, reduced probe equilibrium temperatures are required during cruise. The internal probe temperature is designed warm-biased to offset atmospheric heat losses during descent, but cooler probe equipment temperatures are desired for long-term equipment storage. Since probe heating is supplied exclusively by radioisotope heaters whose output cannot be changed during the mission, the design of the spacecraft mounts must be utilized to reject excessive heating during cruise. Preliminary calculations indicate that approximately 8 watts of excess heat must be dissipated and high conductance spacecraft/probe hard mounts should accommodate this load. Disconnect of these heat paths at separation will then allow the probe to attain its desired warm-biased temperature before planetary encounter.
Electrical/Electronic
All probe checkout commands will be channeled through the spacecraft command/control subsystem. Diagnostic data will return through the spacecraft flight data subsystem. Checkout and
IV-56
Page intentionally left blank
.£>
M
M
M
•* M < * CO
3 a- < o
Mi
ft
CT
P
-(A
<
O
P1 e g
<t>
•
o
Hi
rt
ro
3 rt
f(
3*
VI
ro
o
o
g g c 3 p- o
Co
rt
H-
O
3 t->
H-
3 5^
Co
f(
ID
a- P-
CO
n e CO
CO
ID
o- P-
3 Cfl
(0
o
rt
P-
O
3 a o
Ml
o
3"
Co
XJ rt
ro
ft
Cfl
Co
f|
ID
(D
Cfl
Cfl ro
3 rt
P-
CD
t->
P"
VJ CO
p
. g
P-
M
0)
ft
rt
O
rt
3*
CD
XJ H ro
<
P-
O
c CO
CO
rt
C
H
3*
(D
Cfl
XJ CO
n
CD
o
ft
Co
ft
i rt
Ml
c 3 o
rt
P-
O
3 CO
Co
3 a.
<-t n>
X> c H-
i-t
ro
g ro
3 rt
CO
Cu
iC
ft
P
-3 00
(XX
J ^ • H
M
O
V ro
3*
XI
CD
Cu
(l>
rt
P>
P
-
i- CO
3 CD
.rt
CO
>-f
H^
Cfl
o o
g g c 3 P-
O
CO
rt
P-
O
3 r*
P-
3 Z?
Mi
o
ft ro
x>
M
CO
3 ro
rt
Co
1 VJ ro
3 rt
ft
^ •
rt
P- g
ro
ft
Ml
c 3 o
rt
H-
O
3 s:
a- c n
P-
3 00 rt
3
* ro
xt ft ro
i Cfl ro
3d
XI
P-
O
CT
g c CO
rt
rt
Co
7^ ro
X) P
1
Co
n
ro
Co
rt
Co
Ml
P- X
ro
cu
rt
P- g
ro
P-
3 rt
ro
1-1
< CO
P1
o- ro
i
Co
K
Co
rt
P- o
3 X) ro
H
P
- o
Cu
H-
Cfl
rt
3" ro
P-
3 P-
rt
P-
CO
rt
P-
O
3 O
Ml
rt
3" ro
n o
Co
Cfl
rt
cu n
H
-V;
CO
rt
ro
X) ro
ft
H
-o
cu
a*
ro
Ml
o
H ro
CO
ro
XJ Co
i-S
CO
rt
P
-O
3 • H
3
J ro
o
3 P"
v; o
l-i
P-
rt
P-
O
Co
P"
rt
H- g
P-
3 0Q
ro
M
12
g
ro
3 rt
n
P1
ro
Cfl
*#
rt
3* ro
I-I ro
Cu
f|
Co
P
-3 O
3 rt
3
* ro
CO
p.
X)
CO
3 o
ft ro
XI c P-
i-i ffi
g
ro
3 rt
rt
O
O o
3 ftl
P-
3 ro
n
3" ro
n
7?
o c rt
rt
O
rt
3* ro
P- g
g
ro
i
Co
o
ro
n n
Co
Mi
rt • W
P-
3 n
ro
rt
3* ro
XI n
o
cr
ro
P-
CO
Cu
ro
Co
o
rt
p
- <
Co
rt
ro
Cu
D" ro
rt
S3
ro
ro
3
Co
3 cu
rt
3" ro
X) ft o
cr
ro
S
H- (-<
M &
ro
Cu
ro
Co
o
rt
H
- <
Co
rt
ro
a- cr
ro
rt « ro
ro
3 o
•xJ n
i-1 ro
CO
rt
O
H ro
a.
e r>
ro
X) o
£ ro
H
« H-
M1
h-1
rt
to
?**
ro
CO
i-i
ro
!-»
Co
rt
H- <
ID
I-1
^ CO
3*
0 n
rt
rt
H- g
ro
o
0 g
X) Co
>-t
ro
d
-
S3
p-
rt
3"
rt
3" ro
rt
O
rt
Co
M
O
"<
n
H1
ro X
) H
O !»!
H-
g
Co
rt
31
ID
h-1
O
O
rt
X)
ro
h-1
V! Ul • OJ
3" « 0)
rt a H
Co
3 C
CO
to
3 a-
a- ro
M
P
V! O
Mi
rt
S4 ro
10
X) Co
a CD
n
M
Co
M
l rt
•*—
rO
CT
O
• ~J
D*
H
Co
rt
C/3
Co
rt e i-t
3 •T3 « O
o*
ro
n- 3* ro
n
?r
o e rt
H
O c 3 a- CO
rt
Co
rt
P-
O
3 P"
P-
3 !^
*J
El*
P
-O
3
*
P-
Cfl
Co
X) 1
K
p-
rt
3*
0Q
n
o
c 3 a.
CO
rt
Co
rt
P-
O
3 C
X)
a*
ro
CO
rt
Co
M
rt
ID
CL
(0
C
Ml
Ml
P- r>
p- ro
3 rt
r-1
tx
vj
Co
rt
P-
3 0Q • H
3*
ID
rt
P- g
ID
Ml
o
>"i ro
Co
o
3" n
v; n
M
ID
p-
3 n
M c a.
ro
CO
ro
Co
H
M
V! rt
O
Co
rt
o
Co
n
ID
D- c o
ID
a*
Cu
Co
rt
Co
•-( ro
rt
C
H
3 g
p-
Cfl
Cfl
p- o
3 «*
MX
) M
O
«:
co
ro
<
ro
M
Co
M
O
3" ro
o ^ o c rt
O
V! o
P1
ro
CO
fi ro
i CO
ro
X) CO
M
Co
rt
P
-O
3 n
3
* ID
r>
?r
o
c rt
CO
3"
O c M
Cu
n
Co
X) (U
a*
p
- P»
p-
rt
V! Ml
o
•i
.ro
M
p. g
p-
3 Co
rt
P-
3 CW
Ml
Co
P-
M
CD
a- CO
c cr
CO
vj CO
rt
ro
g
CO
S
3*
P- n
3*
Co
K. ro
3 o
rt
ID
Cfl
CO
ID
3 tia
Cu
c n
ro
x»
ID
Co
?c
X) o SJ
ID
ft
Cu
ro
g
Co
3 C
u
o
3 rt &
ID
CO
XI Co
n
ro
n
i-t
Co
M
l rt
• C/3
p-
3 n
ro
rt
S4 ro
i-t
ro
p-
CD
X) i-t
o
cr
ro o
M
i
XJ ft o
cr-
ro
CO
c cr1
CO
V! CO
rt
ro
g
CO
p-
3 ft
ID
Cu
C
O
ID
C
L g
o
Cu
ID
CO
P-
CO
p-
3 CL
P- n
Co
rt
ID
Cu
P-
3 O
H
Cu
ro
i-t
rt
O
ft
ID 1
P-
rt
CO
H
o
Cu
p
> ro
•
rt
ID
ft
H
g
3-
H-
ro
3 ro X
J C
u
O
sd
a*
ro
vj
ft
rt
ft
3'
ro
ro x
i c X
) P
-ft
I-t
o
ro
cr
cu
ID
Ml
a ft
re
o
w
g
Co
rt
3 &
cu
ro
en
Cfl
c xt
a
4 Co
CO
O
^
ID
CD
O
rt
ft
ID
Co
g
MI
rt
c 3 C
u
Cu
C
ro
ft
i-t
p
-3
rt
0Q
ID
CO
rt
rt
3*
• P
-Cf
l
OX
) X
J ro
ID
ft
ft
p
-(B
O
rt
C
u
ion
rt
3^
ID
XJ ft
O a*
ro
CO
C
cr
co
•<:
CO
rt ro
g
Cfl
Co
3 Cu
XJ ro
ft
M
l 0 ft
g
p-
3 00 rt
ro
CO
rt
CO
c 3 Cu
ro
ft
CO
XI Co
o
ro
n
i-t
Co
M
l rt
O
O
3 1
Cu
XJ
c l-t
p-
3 oo
n
H c p-
CD
ID
• -T)
ft
ID 1 CO
ID
X) Co
ft
Co
rt
P
-O
3 o
X
?
ro
o
7T
O c rt
-3
p-
ft ro
CO
CO
c ft ro
s3
H-
rt
cr
o
n
o
Co
Cfl
p- o
3 Co
M
O
XJ ro
•t
Co
rt
P
-O
3 O
M
i
f-" o
-3
1-.
XJ
P1
ID
3 rt
Co
P-
M
X) o
•3
ro
i-t
p-
3 oo
c X)
o K
ro
ft cr
ID
Co
rt
ID
f(
Cfl
P.
CO
ro
X
X) ID
n
ted
g
p-
3 O
ft
ID
3 OO
p
-3 ID
ID
ff
P
-3 00 P-
3 CO
rt
ft
C g
ID
3 rt
Co
rt
p- o
3 CO
C
ft <
ID
P-
M
P1
Co
3 O
ID
O
Ml
rt
(B
g
xt ro
ft
Co
rt
c I-t
ro
P
3 Cu
ID
3 ro
X
rt
X
J ft
V
! Co
O n
ro
M
ID
ft o g
ID
rt
ID
ft
CO
S3
H>
M
P1
cr
ro
X) o
s3
ro
ft ro
Cu
Cu
c ft
p
-3 OO
rt
cr
ro
P1
Co
c 3 O cr
XJ cr
Co
CO
ro
-••
ID
n
rt
ID
Cu
rt
cr
Co
rt
rt
cr
ID
XJ ft o
cr
ID
Cu
CO
rt
Co
cr
Co
3 Cu
P
1
P-
3 oo
CO
c a- CO
V! CO
rt
ro
g
Co
3 Cu
Cfl n
p- ro
3 o
ID
to
3 Cu
rt
cr
cr--
-!
o c 0
0 3*
P1
Co
c 3 O
3*
Co
3 Cu
n
ft c p-
CO
ro
Ml c 3 O
rt
p-
O
3 CO
cr
CO
< ID
3 O
rt
cr
ID
ro
3 Cu
ro
M
l P
-3 ro
C
u
» p-
rt
P-
Cfl
Ml
P1
P-
rt
00
3*
ID
Cfl
XJ Co
o
ro
o
ft
Co
M
l rt
Mi
o
n
Co
p-
p-
XI ft o
cr
ro
Co
r>
rt
p- <
p-
rt
V! cr
ro
Mi
o
ft
ID
CO
ro
XJ Co
ft
Co
rt
P
-O
3 • >
P
" 1
3*
rt
Cu
P
-CO
o
o
3 3 ro
o
rt
C g
cr
p-
P1
p- o
CU
P1
•^ o
p-
ft n c p-
rt
ft
V! • fd
o
33
ro
ft
S3
p-
P»
P1
cr
ro
CO
c XJ
XI P
1
p- ro
r-
Cu
p
-Co
0
0 3 o
CO
rt
H-
O
Ml
e 3 o
rt
P-
O
3 Cfl
to
ft ro
rt
ft
Co
3 Cfl g
p-
rt
rt ro
Cu
rt
O
rt
cr
ID
X) ft
O cr
ro
cr
^ cr
Co
ft
Cu
S3
p
-ft
ro
*~\
p-
3 1
w
< 1 h
-1
rt
3*
H-
CO
<
o H e 3 (T) •
pi 3 CL
rt 3*
CD
CD
X
X) P 3 CL
CO
C
A
TS
0)
V! M
O
0) CL
H- 3 CO
rt
i-i
C § (0
3 rt
CO
H- 3 n 3"
P
•3
rt
(0
i-t
M
H
M
>#
CAS
(D
O
r
t H
- o 3 W
o Hi
P
H
ft)
Cu
H
-CO
O
c CO
CO
CD
CL
H-
3 CL
CD
r
t P
M
-r-
1
H- 3 n s4
pi
•3 rt
CD
i-i
3 c CO
rt
P>
M
CO
O w
fl> 0)
D
O
O c 3 IT-
CD
CL
Hi
o H • H
3^
CD
CD
X
i-
cd
H
M
>*
CO
CD
O
M
O
H
P
rt
O
i-i
rt
VJ
H- o 3 o • M
O
Hi < o I-1 c 3 CD
M
H
T) 03
V
J M
O
P
CL
H> 3 co
rt
H
3 3 CD
3 rt
CO
03
H
(D
Hi
O
H
rt
3*
CD
H- 3 CO
rt
H
C 3 CD
3 rt
CO
O 3 H-1
^ **•
o 3*
03
3 OQ
CD
CO
H-
3 rt
3*
CD
03
CO
CD
O
O
H-
03
rt
CD
CL
CO
•<!
CO
rt ffi
3 CO
H-
CO
^J
rO
S^ 3 o H
CD
rt D'
03
3 Hi
o H
rt
rt
M
3*
O
ro
.> £
rt
CD
n*
H-
(D
S-9
0Q
S 3
>
0 I-i
CD
(D
CD
3 r
t CO
O
O
r
t i-
i 03
CO
!-
>
V
< CD
O
p
. (_
i
c G
0
Q
3 CD
3- B
-(D
CD
X
"3 M
O
i-
i 03
r
t O
H
^ •3 15
V
I M
O
03
a • H
3*
CD
CO
CD
"3 CD
K
O
(D
3 rt
03
OQ
CD
CO
O
C
VO
3 u
: r
t Sv
S
(D
Cu
O
H
i r
t 0 K
a1 r
t H
-3
" O
34
T3
CD
H
H-
-j
3 k
O
Q3
5-9
l-(
- "<
3 0
0 H
O
ro
3 -3
r
t 0
3-
3 03
CD
3
3 rt
rt
CO
3* a> t
>
3*
CD
CD
X
i-i
tJ
CD
M
O
*3
H
O
03
CD
rt
CD
O
H-
H
cr
CD
3"
^ M
O
H-
C
CO
rt
CO
H-
H-
3 C
L
CO
CD
H-
. a
. CD
H
rt
V
V
CD
CD
TJ
*3
O
H
S
0 ID
tr
" H
CD
H-
CT
3
O
O
CL
i-i
^!
CD
" 03
C
O
CD
CD
•3
• 03
V
! M
H
O
Br1
03
fD
Cu
03
fi 3
*3
CL
CO
3 H
- C
L
rt
CD
CL
O
n
>
3 O
03
3
V!
13
(-J
H-
O
CD
03
co
a* H
- r
t CO
H
-r
t O
Q
H-
03
O
rt
CO
CD
O
rt
Hi
p*
ro
rt
3*
C
CD
"3
CO
"3
CD
CD
i-i
H-
3 0
) co
r
t r
t 3
H
O
C
CO
3 X
) CD
3
" 3
CD
rt
H
CO
CD
03
03
H
3 CD
O
-
00
H-
H-
O
<!
3 CD
O
3
CD
X)
H-
Cf
3 (D
i-
i H
CD
03
.
V
M
CD
H
34
< CD
| H
<r
03
i-t
co
CD
H-
CD
O
O
M
O
< 3
* CD
03
an 03
H
- O
3
rt
rt
CD
O
H 1
M
O
O
03
rt
H-
O 3 a- CD
Hi
H-
3 H-
rt
H-
O 3 03
3 a- 03
CO
CD
rt
O
Hi
T3 H
CD
1 (D
3 rt
i-i
vi H- 3 co
rt
i-i
C 3 CD
3 rt
CO
rt
O
H-
3 < CD
CO
1
CD
X
•3 M
O
H
03
r
t O
i-
i V
!
Xi P
VJ M
O
03
O-
H-
3 CO
rt
i-i c 3 CD
3 rt
CO
v 03
3 CD
"3 3
* (D
(-1 O
3 CD
rt
CD
H
Hi
O
i-i &.
(D
CO
O
CD
3 rt
O (-•
O
3 CA
.
H
ff
CD
CD
X
-3 03
3 a
, CD
a- CO
0 H-
CD
3 0 CD
•3 03
V
j I-1
O
03
(A
I-1-
3 O
M
C
CU
CD
CO
V
* H- 3 03
CU
C
L
IJ-
rt
H-
O 3 rt
O
rt
3^
CD
C/l
>
0
CO
n H
W a 0 w
M a CO
H
g
|^
W a H >
H
M
O 2 P>
b
7]
O
•TJ
W S
O
td
pg
|
^ a 0 M
rt
3*
H-
en
i-i
CD
.Q
C
H-
i-i
CD
3 fD
3 rt >
^•
N
CO
03
r
t 3 i-
i 3 ~
-j
<D
u
CO
03
r
t C
i-
i 3 ^^
C
O
a 00
0 •#
03
3 D-
C
i-i
03
3 3 CD
CO
c
!
OO
0 03
3 03
h
-1
v;
N
CD
3-
CD
03
rt
H-
CO
Hi
H-
(V
CD
O
3 rt 3"
CD
CO
c 3
"3
O
1-3
03
i-i
tr
n CD
r
t 03
H-"
3 O
r
t H
« H
i CD
CO
l-
i CO
CD
CD
H
-
x a
o r
-"
3 3
H-
rt
CO
H- a- CD
O
Hi
rt V
CD
X) r-
1
03
3 CD
rt • M
3 rt
M
^ Hi
O
i-i
rt
3*
CD
rt
3*
i-i
CD
CD
3 H-
CO
CD
H-
O
3 CO
03
CT
" C
O
3 ^
&.
(S3
CD
rt
CD
&
3"
CD
(A.
CD
(D
O
CD
Hi
CD
"3
3 3
03
CD
O
3 P
-CD
&
CD
H
-0
a.
3 3
co
n 3
C)
D1
H.
H-
03
CD
CD
"3
CD
3 r
t
H-n
ID
O
CD
H
3 O
H
C
L
O
< CD
3
-CD
X
) H
- t-
1 C
O
era
CD C
D
3 3
0 CD
r
t H
- 3
H-
CD
rt
O
• 3
rt 3*
> 03
H
r
t p
* 03
CD
f«
CD
CD
3
"3
rt
i-i
3 i-
i H
- O
V
! 3
rt
03
0 1
Q,
n
V)
p-
O
03
C
H-
3 i-
i 3
OQ
CO
1
CD
CL
> <
s I-I CO
C
O
M
O a 0 M N
M a H
H
H
O a
Hi
r-1
CD
O
rt
H-
O 3 3 0 Cu
CD
•
3 H-
CO
CD
H-
O 3 CO
V r
t 3
* CD
^ c 0 f-t
CO
rt
O
03
CD
(D_
03
rt 3 O
CD
•3 cr
CD
H
H
-O
3 0 CL
CD
h
-1
CO
V 03
3 CL
CO
•3 03
O
CD
O
i-
t 03
H
i r
t de-
rt
31
CD
cr
03
CO
H-
O
CO
rt e CL
V
J 03
3 CL
fi
rt
CD
3 CL
CD
C
L
rt
O &
CD
O
O 3 -
3 03
rt
H- &
M
CD
£ H*
rt
3^
rt
3"
CD
rt
3*
i-i
CD
CD
3 §
rt
CD
i-t
M <
• H
CD
O
CD
3 rt
03
t-1
0 3 P
'OQ
CD
CL
CD
H
i H
- 3 H-
rt
O 3 C
CO
CD
CD
O
O 3 O
CD
X) rt
CO
rt
3 H>
rt &
rt S4
CD
<—. M
a"<
! 8 >
>
-•
• H
3"
CD
CO
O
H-
CD
3 O
CD
CO
*
3
>
O fi
•3 h-1
O
H
03
rt
O
H
V!
•3 03
p*v
s 03
r
t S
CD
i-t
CD
OQ
CD
3 CD
i-i
03
rt
CD
CL
CL
3 i-
i 3 OQ
M
O
03
CL
«*
Cu
03
O
7?
03
OQ
CD
03
H
CO
O
H-
3 O
M
3 Cu
CD
CO
03
3 CD
H-
>3
CD
n 3 CD
CO
CD
CL
H- 3 rs
3*
03
X) 1
3^
CD
H1
O 3 CD
rt
CD
H
Hi
O
H de-
N
CD
i-t
•~v
H S
>
>-•
V
03
3
CD
c 0 3^
P
CO
rt
3^
CD
tr1
P
3 0Q
C
u
3
a CD
3 rt
H
03
r-1
•tj
03
H
rt
H-
O
h-1
CD
(=0
CD
rt
03
I-t
CL
H- 3 O
Q
3 H-
H
•d
I-i
0 cr
CD
•—v
f hd
v
-x
V H
O
3 !*
CD
rt
P
i-t
CL
H
-3
•nO
Q
O
It-
CD
3 rt
H-
03
M > 3 P
•d
O
rt
CD
3 rt
H-
P
M >
h-'
3
VI N
CD
H
P 1
CL
H
CD
&
H
i p
. H
- CD
3 H-
CO
rt
CD
H-
O
O
rt
3 H
-O
/-,
3 H
p
CO
co
3 K
I
H
P
H
H
v_
^ p
. N
S
! CD
|y
co
H
-O
r
t C
T-
B*
CD
3 CO
O
CD
O
CD
3 § T
J O
H
3
CD
1 C
O
CD
p
3 r
t r
t e
i-t
i-t
vj
3^
co
cj
O
H
p.
03
CD
3
3 C
O
CO
CD
X
) H
- ft
3
O
CO
CT1
rt
CD
H 3
co
3 v
j CD
0
1 3
rt
rt
(D
CO
3
Table V-l Instrument Charaateristios
INSTRUMENT/COMPONENT
EXPLORATORY PAYLOAD TEMPERATURE GAGE PRESSURE TRANSDUCERS (2) ACCELEROMETER SYSTEM
TRIAD SENSOR PRC & ELECTRONICS
NEUTRAL MASS SPECTROMETER ANALYZER ELECTRONICS PUMP INLET SYSTEM
EXPANDED PAYLOAD ADDITIONS NEPHELOMETER (DESCENT) LANGMUIR PROBES (PRE-ENTRY) ION RPA (PRE-ENTRY)
SENSOR ELECTRONICS
NEUTRAL RPA (PRE-ENTRY) SENSOR ELECTRONICS
TOTALS
WEIGHT, kg
0.45 0.68
0.60 0.91
1.83 2.72 0.44 0.45
1.13 1.36
0.23 1.36
0.44 1.83
14.43
POWER, w
1.4 1.3 2.8
14.0
3.0 3.0 3.0
5.0
33.5
VOLUME, cm3
426 246
262 656
935 3,444 443 492
1,312 1,200
400 1,800
400 2,100
14,116
Pre-entry altitude profiles for both planets are given in Figure 111-27 along with associated velocity time histories. The pressure descent profiles are given in Figure 111-31 and are specified for the six model atmospheres. The descent ballistic coefficient is 110 kg/m2 for all profiles. Values for descent parameters are given in Table 111-14, showing that parachute deployment pressure varies from 41 to 102 millibars while the pressure at first descent measurement varies from 51 to 122 millibars. The final pressure for a descent designed for 44 minutes varies from 3.1 bars (-96 km alt) in the Uranus warm atmosphere to 19.2 bars (-60 km alt) in the Uranus cool atmosphere.
The pre-entry, entry, and descent times, instrument sampling times, and resulting bit rates are shown in Table V-2. The transmission bit rates for descent are slightly higher than the collection bit rates because the descent data collected during probe acquisition by the spacecraft must be stored and later interleaved with realtime data for transmission. The pre-entry transmission of 235 bps is real time; thus the collection equals the transmission.
V-2
Tab
le
V-2
In
stru
men
t Sa
mpl
ing
Tim
es
and
Dat
a R
ates
fo
r C
omm
on
Satu
rn/U
ranu
s P
robe
w
ith
Exp
ande
d P
ay l
oad
< i
PHASE
PRE-ENTRY
(200 sec)
-
=•
ENTRY
(84.5 sec
MAXIMUM)
DESCENT
(2640 sec)
INSTRUMENT
LANGMUIR PROBES
IRPA
NRPA
ACCELEROMETERS
AXIAL
LATERAL (2)
TEMPERATURE GAGE
PRESSURE GAGE
NEPHELOMETER
MASS SPECTROMETER
ACCELEROMETERS
TURBULENCE
STORED
STORED PRE-ENTRY
SAMPLING
TIMES, sec
0.5
2.0
3.0
0.2
0.4
4 4 3
50
8 0 0
COLLECTION
BIT RATE, bps
60
60
93.3
ENGINEERING
FORMATTING
TOTAL
50
50
2.5
2.5
3.3
8.0
7.5
0 0
SCIENCE TOTAL
ENGINEERING
FORMATTING
TOTAL
TRANSMISSION
BIT RATE, bps
60
60
93.3
213.3
0.5
21.3
235.1
0 0
2.6
2.6
3.4
8.3
7.8
3.4
16.8
44.9
1.0
4.6
50.5
The descent bit rate includes the nephelometer at 3.4 bps, which does not appear in Table IV-1 with the exploratory payload. Also included in the design is a capability for replaying the pre-entry data from storage, shown by the last full line in Table V-2. A rate of 16.8 bps is required to telemeter the stored pre-entry data. The redundancy mode for the pre-entry data is allowed by the fact that the total power requirements for the pre-entry realtime data transmission at 235 bps using PSK exceeds that for descent using FSK until the descent bit rate reaches about 53 bps. The use of the excess power during descent takes about 26 w-hr of energy at the bit rate of 50.5 bps. This is a minor penalty to pay for the extra data.
Other options are also available; one, of course, is not to utilize this excess power. In that case, the data rate would be 31.5 bps as shown in Table 111-18. A more viable option is to use the excess to make more detailed descent measurements, for instance making the mass spectrometer sweeps more detailed. The present sampling is 400 bits/sample, which is one voltage setting and current reading per amu. For a total of only 48.8 bps, the sample size could be increased to 1200 bits which would be three readings per mass number, giving a more detailed mass spectrum. The excess power could also be used to obtain more sample points per kilometer with any instrument. Decreasing the sampling times on the simultaneous temperature and pressure measurements to one second would increase the amount of data by a factor of four at a total bit rate of about 48 bps.
The descent mission measurement performance is given in Table V-3 for Saturn and in Table V-4 for Uranus. The first column gives the instrument, its sampling interval, and the measurement to be made. The second column gives the criteria set forth in Volume II. The remaining columns give the performance numbers in the units specified by the criteria. The minimum values are those at the highest point in the atmosphere at which the measurement is to be evaluated, and are specified individually for each atmosphere to allow determination of worst-case atmospheres. The maximum value given in the last column is the maximum performance of all the atmospheres at the end of the design mission, which is 44 minutes from parachute deployment.
The primary difference from Tables IV-2 and IV-3 is the addition of the nephelometer performance. It exceeds the one measurement/ kilometer criteria in all atmospheres and obtains a minimum of 34 measurements in all modeled clouds. The performance of all measurements is generally lowest in the warm atmospheres and improves in the cool.
V-4
Table V-Z Saturn Descent Measurement Performance - Expanded Pay load
INSTRUMENT
MASS SPECTROMETER
TEMPERATURE GAGE
PRESSURE GAGE
ACCELEROMETERS
NEPHELOMETER
SAMPLING TIME
50 sec
4 sec
4 sec
8 sec
3 sec
MEASUREMENT
MINOR CONSTITUENTS
CLOUD COMPOSITION
H/He RATIO ]
ISOTOPIC RATIO /
MOLECULAR WEIGHT )
TEMPERATURE PROFILE
CLOUD STRUCTURE
PRESSURE PROFILE
TURBULENCE
CLOUD STRUCTURE
TURBULENCE
CLOUD LOCATION
CLOUD STRUCTURE
CRITERIA
2 PER SCALE HEIGHT*
2 INSIDE EACH CLOUD
4 MEASUREMENTS
1 PER °K
2 INSIDE EACH CLOUD
2 PER km)
1 PER km) *
2 INSIDE EACH CLOUD
1 PER km*
1 PER km''
2 INSIDE EACH CLOUD
MINIMUM PERFORMANCE
COOL ATMOSPHERE
5.8
10.4 IN NH3
52
1.5
137 IN NH3
2.7
137 IN NH3
1.3
1.4
183 IN NH3
NOMINAL ATMOSPHERE
7.7
4.0 IN NH3
16 IN H20
52
1.2
46 IN NH3
196 IN H20
2.0
46 IN NH3
196 IN H20
1.0
1.3
66 IN NH3
275 IN H20
WARM ATMOSPHERE
6.5
3.0 IN NH3
10.5 IN H20
52
1.3
38 IN NH3
47 IN H20
1.5
38 IN NH3
47 IN H20
0.7
1.4
50 IN NH3
62 IN H20
MAXIMUM PERFORMANCE
37
8.1
7.8
3.9
10.4
*BEL0W CLOUD TOPS.'
tBELOW 100 mb.
Table V-4 Uranus Descent Measurement Performance - Expanded Pay load
INSTRUMENT
MASS SPECTROMETER
TEMPERATURE GAGE
PRESSURE GAGE
ACCELEROMETERS
NEPHELOMETER
SAMPLING TIME
50 sec
4 sec
4 sec
8 sec
3 sec
MEASUREMENT
MINOR CONSTITUENTS
CLOUD COMPOSITION
H/He RATIO
ISOTOPIC RATIO
MOLECULAR WEIGHT J
TEMPERATURE PROFILE
CLOUD STRUCTURE
PRESSURE PROFILE
TURBULENCE
CLOUD STRUCTURE
TURBULENCE
CLOUD LOCATION
CLOUD STRUCTURE
CRITERIA
2 PER SCALE HEIGHT*
2 INSIDE EACH CLOUD
4 MEASUREMENTS
1 PER °K
2 INSIDE EACH CLOUD
2 PER km i
1 PER km 1*
2 INSIDE EACH CLOUD
1 PER km*
1 PER km+
2 INSIDE EACH CLOUD
MINIMUM PERFORMANCE
COOL ATMOSPHERE
5.6
18.3 IN CH4
52
3.0
230 IN CH4
4.8
230 IN CH4
2.4
2.4
306 IN CH4
NOMINAL ATMOSPHERE
6.1
4.4 IN CH4
12 IN NH3
52
3.0
48 IN CH4
113 IN NH3
2.5
48 IN CH4
113 IN NH3
1.4
3.2
76 IN CH4
174 IN NH3
WARM ATMOSPHERE
4.6
2.0 IN CH4
5 IN NH3
52
3.1
26 IN CH4
63 IN NH3
1.6
26 IN CH4
63 IN NH3
0.9
1.8
34 IN CH4
86 IN NH3
MAXIMUM PERFORMANCE
38
10.9
14.4
7.2
19.1
*BELOW CLOUD TOPS.
+BELOW 100 mb.
V-5
Table V-5 Pre-Entry Measurement Performance
INSTRUMENT (At)
LANGMUIR PROBES (0.5 sec)
IRPA (2 sec)
NRPA (3 sec)
KILOMETERS PER MEASUREMENT
AT 5000 km
SATURN
9.7
38.8
58.2
URANUS
10.1
40.2
60.3
AT ENTRY
SATURN
9.1
36.4
54.6
URANUS
10.3
41.3
62.0
Table V-5 shows the performance during pre-entry in kilometers/ measurement at the beginning and end of the pre-entry phase. As can be seens the values change only a small amount during this period. The number of actual measurements was discussed together with the upper atmospheric models in Chapter III, Section B.3.C. According to the electron number density equation given in the Monograph (Ref III-l), the electron scale height is 250 km. This means that with the performance given in Table V-5, the Langmuir probes are making approximately 25 measurements per scale height. Since electrons are probably most abundant and other particles have a smaller scale height, this value is the largest number of measurement/scale height obtainable. The minimum number is an unknown based upon the number density profile of the heaviest neutral particle measurable.
SYSTEM DESIGN AND INTEGRATION
Functional Sequence
The functional sequence for the Saturn/Uranus probe with the expanded complement is similar to that for the SAG exploratory pay-load except for the added events required for the pre-entry science and the added descent nephelometer. During pre-entry, the science and engineering data is transmitted in real time and also stored for descent transmission as discussed in Subsection 3 herein. This dual pre-entry data mode was incorporated because the pre-entry and descent RF power requirements are almost identical as discussed in Chapter IIIS Section D.l herein. Detailed sequences of events are shown in Tables V-6 and V-7 for Saturn and Uranus, respectively.
V-6
Table V-6 Saturn Sequence of Events (Task ID* ITEM
1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
11.
12.
13.
14.
15.
16.
17.
18.
19.
20.
21.
22.
23.
24.
25.
26.
27.
28.
29.
30.
31.
32.
33.
34.
35.
36.
TIME
L- = 0
L + 2h
L + 2h TO L + 1221.8d
S - 7h, 30m, Os
S - 7h, 17m, Os
S - Oh, 17m, Os
S - Oh, 2m, Os
S = 0 (L + 1222.Id)
S + Om, 0.5s
S + Om, 30s
S + 4m, Os
S + 15m, 30s
S + 17m, 36s
S + 18m, Os
S + 33m, Os
L + 1222.12d TO L + 1257.9d
E - 40m, 14s
E - 20m, 14s
E - 19m, 44s
E - 9m, 54s
E - 9m, 24s
E - 8m, 14s
E - 8m, 9.5s
E - 7m, 44s
E - 3m, 20s
E = 0
E + Om, 5s
E + Om, 17.5s
E + Om, 44.5s [5 g (REF)]
E + Om, 56.5s (5 g + 12s)
E + Om, 59.5s (5 g + 15s)
E + lm, 9.5s (5 g + 25s)
E + lm, 14.5s (5 g + 30s)
E + 2m, 39.5s (5 g + 115s)
E + 44m, 59.5s (5 g + 2655s)
E + 2h, 12m (L + 1258d)
•ASSUMES AN ENTRY ARRIVAL UNCERTAINTY C AND A NOMINAL ATMOSPHERE. +(C) = COOL ATMOSPHERE; (W) = WARM ATMC
EVENT
LAUNCH (DEC 3, 1980)
SEPARATE SPACECRAFT FROM LAUNCH VEHICLE
CRUISE
SPACECRAFT POWER TO PROBE; EJECT ENVIRONMENTAL COVER
START PROBE CHECKOUT
PROBE CHECKOUT COMPLETE; START SPACECRAFT ORIENTATION (10.6°)
SPACECRAFT ORIENTATION COMPLETE; ACTIVATE SEPARATION SUBSYSTEM
SEPARATE PROBE FROM SPACECRAFT
START PROBE SPINUP TO 5 rpm
START ORIENTATION FOR AV (59.1°)
PROBE SPINUP COMPLETE; DEACTIVATE PROBE SYSTEM
COMPLETE SPACECRAFT ORIENTATION FOR AV
APPLY SPACECRAFT AV (96 m/sec)
START SPACECRAFT ORIENTATION TO EARTH LOCK (69.7°)
SPACECRAFT ORIENTATION COMPLETE
COAST
START BATTERY ACTIVATION PYRO CAPACITOR CHARGE
ACTIVATE BATTERY
POWER ON (DATA HANDLING 8, PYRO SYSTEMS); TURN LPs ON WARM-UP; START STORING ENGINEERING DATA
TRANSMITTER ON WARM-UP
START ACQUISITION; REMOVE NRPA COVERS
RPAs ON WARM-UP; CONTINUE STORING ENGINEERING DATA
DESCENT; SCIENCE ON WARM-UP
LPs & IRPAs READY TO MEASURE; ACQUISITION COMPLET
LPs 8 IRPAs SENSE SIGNAL TO START STORING & TRANSMITTING DATA
ENTRY (536 km ABOVE 1 atm; 1 x 10"7 atm)
RF POWER AMPLIFIER OFF (0.1-g SENSING); START DATA STORAGE; PRE-ENTRY SCIENCE OFF
ENABLE PROBE DESCENT PROGRAM (100-g SENSING)
START DESCENT PROGRAM (5-g SENSING); OPERATE BASE COVER BAND CUTTERS
DEPLOY BASE COVER QUADRANTS
DEPLOY MAIN PARACHUTE (M = 0.7; %65 mb); RF POWER AMPLIFIER ON; START ACQUISITION
EJECT AEROSHELL
RELEASE MAIN CHUTE; DEPLOY TEMPERATURE GAGE; UNCOVER NEUTRAL MASS SPECTROMETER & NEPHELOMETER
PROBE ACQUISITION COMPLETE; START DATA TRANSMISSION
END OF DESIGN MISSION ( 8 bars)
SPACECRAFT PERIAPSIS (3.8 P.A; MAY 9, 1984
F 4.4 minutes, A DESCENT BALLISTIC COEFFICIENT OF 110 kg/m2 (0.7
SPHERE.
COMMENTSt
RF POWER = 7 w (TRACKING TONE).
4300 km ABOVE 1 atm.
297 km (C); 968.5 km (W).
1.7s (C); 12s (W).
8.5s (C); 35s (W).
29s (C); 69.5s (W).
41s (C); 81.5s (W).
44s (C); 84.5s (W); M = 0.56 (C); M = 0.88 (W); 102 mb (C); 44 mb (W).
47.7 bps.
(45 m, 24.5s (W); 118 bars (C); 4 bars (W).
lug/ft2),
V-7
Table V-7 Uranus Sequence of Events (Task ID*
ITEM
1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
11.
12.
13.
14.
15.
16.
17.
18.
19.
20.
21.
22.
23.
24.
25.
26.
27.
28.
29.
30.
31.
32.
33.
34.
35.
36.
*ASSU AND /> f(C)
TIME
L = 0
L + 2h
L + 2h TO L + 2816.7d
S - 13h, 30m, Os
S - 13h, 17m, Os
S - Oh, 17m, Os
S - Oh, 2m, Os
S = 0 (L + 2817.3d)
S + Om, 0.5s
S + Om, 30s
S + 4m, Os
S + 15m, 30s
S + 17m, 36s
S + 18m, Os
S + 33m, Os
L + 2817.32d TO L + 2824.6d
E - lh, 4m, 50s
E - 44m, 50s
E - 44m, 20s
E - 34m, 30s
I - 34m, Os
E - 32m, 51.5s
E - 32m, 50s
E - 32m, 20s
E - 3m, 20s
E = 0
E + Om, 4s
E + Om, 15s
E + Om, 38.5s [5 g (REF)]
E + Om, 50.5s (5 g + 12s)
E + Om, 53.5s (5 g + 15s)
E + lm, 3.5s (5 g + 25s)
E + lm, 8.5s (5 g + 30s)
E + 2m, 33.5s (5 g + 115s)
E + 44m, 53.5s (5 g + 265s)
E + lh, 51m (L + 2826.9d)
MES AN ENTRY ARRIVAL UNCERTAINTY C NOMINAL ATMOSPHERE.
= COOL ATMOSPHERE; (W) = WARM ATMC
EVENT
LAUNCH (DEC 3, 1980)
SEPARATE SPACECRAFT FROM LAUNCH VEHICLE
CRUISE
SPACECRAFT POWER TO PROBE; EJECT ENVIRONMENTAL COVER
START PROBE CHECKOUT
PROBE CHECKOUT COMPLETE; START SPACECRAFT ORIENTATION (21.4°)
SPACECRAFT ORIENTATION COMPLETE; ACTIVATE SEPARATION SUBSYSTEM
SEPARATE PROBE FROM SPACECRAFT
START PROBE SPINUP TO 5 rpm
START SPACECRAFT ORIENTATION FOR AV (46°)
PROBE SPINUP COMPLETE; DEACTIVATE PROBE SYSTEM
COMPLETE SPACECRAFT ORIENTATION FOR AV
APPLY SPACECRAFT AV (85 m/sec)
START SPACECRAFT ORIENTATION TO EARTH LOCK (67.4°)
SPACECRAFT ORIENTATION COMPLETE
COAST
START BATTERY ACTIVATION PYRO CAPACITOR CHARGE
ACTIVATE BATTERY
POWER ON (DATA HANDLING & PYRO SYSTEMS); START LP DECONTAMINATION; START STORING ENGINEERING DATA
TRANSMITTER ON WARM-UP
START ACQUISITION; REMOVE NRPA COVERS
DESCENT SCIENCE ON WARM-UP
RPAs ON WARM-UP; CONTINUE STORING ENGINEERING DATA
LPs 8. RPAs READY TO MEASURE; ACQUISITION COMPLETE
TRANSMIT AT HIGH POWER; STORE DATA (LPs & IRPA SENSING)
ENTRY (532 km ABOVE 1 atm; 1 x 10"7 atm)
RF POWER AMPLIFIER OFF (0.1-g SENSING); START DATA STORAGE; PRE-ENTRY SCIENCE OFF
ENABLE PROBE DESCENT PROGRAM (100-g SENSING)
START DESCENT PROGRAM (5-g SENSING); OPERATE BASE COVER BAND CUTTERS
DEPLOY BASE COVER QUADRANTS
DEPLOY MAIN PARACHUTE (M = 0.72; - 47 mb); RF POWER AMPLIFIER ON; START ACQUISITION
EJECT AEROSHELL
RELEASE MAIN CHUTE; DEPLOY TEMPERATURE GAGE; UNCOVER NEUTRAL MASS SPECTROMETER & NEPHELOMETER
PROBE ACQUISITION COMPLETE; START DATA TRANSMISSION
END OF DESIGN MISSION ( 7.3 bars)
SPACECRAFT PERIAPSIS (2 R^; AUGUST 24, 1988
F 29 minutes, A DESCENT BALLISTIC COEFFICIENT OF 110 kg/m2 (0.7 si
SPHERE.
COMMENTS1
RF POWER = 7 w (TRACKING TONE).
4300 km ABOVE 1 atm.
200 km (C); 1074 km (W).
0.5s (C); 14s (W).
4.5s (C); 36s (W).
18.5s (C); 65s (W).
30.5s (C); 77s (W).
33.5s (C); 80s (W); M = 0.59 (C); M = 0.84 (W); 41 mb (C); 49 mb (W).
47.7 bps.
19 bars (C); 3.1 bars (W).
ug/ft2)
V-8
Functional Block Diagram
The functional block diagrams for the Saturn/Uranus probe with the SAG exploratory payload and the probe with the expanded pay-load are the same except that the additional science instruments are added for the expanded configuration.
System Data Profile
Figure V-l shows the data profile for the Saturn/Uranus probe with the expanded science complement. Included in the figure are the total data collected for Saturn and Uranus entries and the impact of the entry-arrival uncertainties. The entry and descent data profiles for each planet are the same. Compared with the basic Saturn/Uranus probe, discussed in Chapter IV, Section C.3, the descent data rate for this application is increased by 3.7 bps for the added nephelometer and approximately 18.5 bps for the stored pre-entry science data. The RF power requirement at pre-entry of 41 watts for transmission of 235 bps using PSK modulation, is almost the same as storing the pre-entry and entry data and transmitting at 50.5 bps.
The figure shows the total data storage requirement is 60,200 bits and the total data collected is 128,145 bits. See Sections B and D for further discussion of the science requirements and RF link analysis, respectively.
System Power Profile
Figure V-2 shows the "worst case" power profile for the Saturn/ Uranus probe using the expanded science complement. The separation phase is identical to that for the earlier Saturn/Uranus probe. The pre-entry power profile is based upon the Uranus arrival uncertainty of 29 minutes, and the entry time is based upon Saturn warm atmosphere. The late arrival condition requires 191 w-hr of power to the equipment. Late arrival is described as the condition in which the probe has not yet arrived at the planet when the timer (preset before separation) runs out and starts the sequence before it is really needed.
System Weight Summary
The system weight summary is shown in Table V-8. The ejected weight is 103.22 kg (227.36 lb), entry weight is 103.21 kg (227.33 lb), and the final descent weight is 50.32 kg (110.82 lb). Compared to the basic Saturn/Uranus probe, this probe ejected weight is 13.36 kg heavier and its descent probe is 9.15 kg heavier.
V-9
Preentry Entry Descent
140
120
Combined — ^ ^ Realtime Transmit at
Data, 100- ** ^ atn d §
Bits x 1000 I Storage at W\
80
60
40
20
QL-V
Storage 235 bps
0 10 20
Time from Entry, min
Fig. V-l Data Profile for Saturn/Uranus Probe (Task II)
Power, Watts
160
140
120
100
80
60
40
20
0
Separation
' lO W 0.67 W-hr
2 i
Coast
35.8 days (Saturn)
9.5 days (Uranus)
\ V
Preentry Entry
T 161.3 Transmitter On
(RF Power 41 W H
Late Arrival (LA)
Nominal Arrival (NA)
Early Arrival (EA)
n 68.9 W
Science On
40.4 W [transmitter On |(RF Power 7 W! • I C O H I
I 1 Power O n -
J L
15.3 W
Preentry Science &—r
J RF Power r Amp Off J
I ["
Descent
150.3 W
—Transmitter On (RF Power 41 W)
End of Mission
\
(EA) • (NA) = (LA) •
36.3 W
124.16 W-hr 157.47 W-hr 190.77 W-hr
Time from Separation, min
-70 -60 -50 -40 -30 -20 -10 0 10 20 30 40
Time from Entry, min
Fig. V-2 Power Profile for Saturn/Uranus Probe (Task II)
V-IO
Table V-8 Weight Summary for Saturn/Uranus Probe with Expanded Science Complement
SCIENCE
POWER & POWER CONDITIONING
CABLING
DATA HANDLING
ATTITUDE CONTROL (DRY)
COMMUNICATIONS
PYROTECHNICS
STRUCTURES 8 HEAT SHIELD
MECHANISMS
THERMAL
PROPELLANT (ACS)
ENGINEERING INSTRUMENTATION
15% MARGIN
EJECTED WEIGHT
ENTRY WEIGHT
DESCENT WEIGHT (F INAL)
WEIGHT
kg
1 4 . 4 4
5 . 7 9
4 . 5 0
2 . 7 8
2 . 8 4
3 . 9 5
6 . 4 4
3 7 . 1 2
6 . 6 7
5 . 2 2
0.01
0.00
13.46
103.22
103.21
50.32
l b
31.80
12.75
9.92
6.12
6.25
8.70
14.19
81.76
14.68
11.50
0.03
0.00
29.66
227.36
227.33
110.82
D. TELECOMMUNICATIONS SUBSYSTEM
The Task II expanded science payload impacts the telecommunications subsystem only from the standpoint of increased data bit rate. The three mission trajectories are unchanged from Task I. Table V-9 depicts the RF power required by the probe transmitter for Task II.
Table V-9 Task II Telecommunication Subsystem Comparisons for the Saturn/SU 80 Mission
MODULATION
FSK + TONE
• PCM/PSK/PM
FSK + TONE
FSK + TONE
DATA BIT RATE, bps
235
235
31.5
50.5
MISSION PERIOD
Pre-Ent ry
Pre-Ent ry
Descent
Descent
MAXIMUM PROBE RF POWER REQUIRED, w
116
41
28.5
40
V-ll
The expanded science package includes upper atmosphere instruments that increase the pre-entry data bit rate to 235 bps, which includes engineering data and formatting. RF power of 116 watts is required during pre-entry for the Saturn/SU 80 mission, using binary FSK modulation with a tracking tone. High power is required since a high E /N ratio of 8.9 dB is required for the noncoherent system
(Vol II, Chapter V, Section A.4.e, p V-30).
PSK modulation is more efficient than FSK and was considered for the pre-entry portion of the mission. For a bit error rate of 5 parts in 105, the required E, /N ratio is 4 dB, as seen in Figure
V-15 of Volume II, p V-31. With this improvement, the RF power required, using 235 bps, reduces to 41 watts, as indicated in Table V-9. Convolutional coding was selected as the preferred approach indicated in Chapter III, Section D.3. A half-rate, constraint-length 8 convolutional code was chosen with a Viterbi algorithm decoder using soft decision and 8-bit symbol quantization with the output delayed three constraint lengths.
The exploratory payload with an added nephelometer (Task II) in-, creases the descent data rate to 31.5 bps and can be handled with maximum RF power of 28.5 watts. Since a 41-watt transmitter is already required to handle the high pre-entry bit rate, it was concluded that better use of the available RF power during descent could be achieved. It was found that data transmission reliability would be improved if the pre-entry data were transmitted to the spacecraft in real time and also stored on the probe. During descent, the pre-entry data will be interleaved with the realtime descent science and engineering data and transmitted to the spacecraft a second time to enhance the probability of pre-entry science data return to the spacecraft. This increases the descent data rate to 50.5 bps and RF power of 40 watts is required which fully utilizes the 41-watt probe descent transmitter capability.
Parameters of the RF link are depicted in Table V-10 for the .. Saturn/SU 80 mission with Task II science payload and PSK modulation for pre-entry transmission. Maximum RF power is required at entry with the high data rate. The probe RF power requirements for Task II are shown in Figure V-3. The probe is acquired for either planet from the tone only. The 7-watt probe transmitter level determined from Task I requirements during pre-entry is sufficient for Task II with a tone only.
V-12
Table V-10 Probe Pre-Entry Telemetry Link Design Table for the Task II Mission*
PARAMETER
1. TOTAL TRANSMITTER POWER, dBw '
2. TRANSMITTING CIRCUIT LOSS, dB
3. TRANSMITTING ANTENNA GAIN, dB
4. COMMUNICATIONS RANGE LOSS, dB
5. PLANETARY ATMOSPHERE & DEFOCUSSING LOSS, dB
6, POLARIZATION LOSS, dB
7. ANTENNA PATTERN RIPPLE LOSS, dB
8. RECEIVING ANTENNA GAIN, dB
9. RECEIVING CIRCUIT LOSS, dB
10. NET CIRCUIT LOSS, s(2-»-9), dB
11. TOTAL RECEIVED POWER (1 + 10), dBw
12. RECEIVER NOISE SPECTRAL DENSITY, dBw/Hz
CARRIER TRACKING
13. CARRIER POWER/TOTAL POWER, dB
14. RECEIVED CARRIER POWER (11 + 13), dBW
15. CARRIER THRESHOLD LOOP NOISE BANDWIDTH, dBw/H
16. THRESHOLD SIGNAL/NOISE RATIO, dB
17. THRESHOLD CARRIER POWER (12 + 15 + 16), dBw.
18. PERFORMANCE MARGIN (14 - 17), dB
DATA CHANNEL
19. DATA POWER/TOTAL POWER, dB
20. RADIO SYSTEM & DOPPLER OFFSET LOSS, dB
21. SUBCARRIER DEMODULATION & BIT SYNC/ DETECTION LOSS, dB
22. RECEIVED DATA POWER (11 + 19 + 20 + 21), dBw
23. DATA BIT RATE, dB
24. THRESHOLD Efa /NQ FOR Pg = 5 x 10"5, dB
25. THRESHOLD DATA POWER (12 + 23 + 24), dBw
26. PERFORMANCE MARGIN (22 - 25), dB
27. NOMINAL LESS ADVERSE VALUE (26 - 26 ADVERSE), dB
•CONDITIONS: 1. S/SU 80 COMMON PROBE MISSION (WORST 2. PRE-ENTRY TASK II SCIENCE PAYLOAD A 3. PCM/PSK/PM MODULATION WITH CONVOLUT 4. V = 2, Q = 8, K = 8. 5. DECODER - VITERBI ALGORITHM, BER =
NOMINA*.!. VALUE •
16.1
-1.1
3.9
-197.0
0
-0.3
0
17.9
-1.0
-177.6
-161.5
-196.2
-8.0
-169.5
12.4
10.0
-173.8
4.3
-0.7
-1.0
-1.0
-164.2
23.7
4.0
-168.5
4.3
0
-CASE TRAJ T ENTRY CO IONAL ENCO
5 x 10-5.
ADVERSE TOLERANCE
0
0.1
1.8
0.3
0
0
0.2
1.2
0.1
3.7
3.7
0.4
0.2
3.9
0
0
0.4
4.3
0.2
0
0
3.9
0
0
0.4
4.3
ECTORY). EDITIONS. DER.
REMARKS
40.5 w AT 0.86 GHz
CABLE & SWITCH
100° BEAMWIDTH
1.97 x 105 km
ENTRY
20" BEAMWIDTH
Ts = 1750"K
NF • = 8.5 dB
Pc = 6.4 w
17.5 Hz
Pd = 34.1 w, (J> = 66
235 bps
5°
V-13
URANUS ACQUISITION
PROBE RF POWER REQUIRED, w
45
40
35
30
25
20
15
10
i SATURN ACQUISITION ENTRY
J_J END OF MISSION
n r •PRE-ENTRY, PSK MODUL •
-TONE ONLY-
LEGEND:
S/SU 80
S 79
U/SU 80
-H2 3 5
*p!a
i 20 DESCENT, FSK MODUL, 50.5 bps
NOTE: 1. COOL ATMOSPHERE MODEL FOR SATURN. 2. NOMINAL ATMOSPHERE MODEL FOR URANUS
-0.6 -0.4 -0.2 0 0.2 TIME FROM ENTRY, hr
0.4 0.6 0.8
Fig. V-3 Task II Probe RF Power Requirements
V-14
At approximately E-3.5 minutes, when the Langmuir probe or IRPA senses the presence of the upper atmosphere, the data rate is increased to 235 bps and the maximum RF power requirement of 41 watts occurs at entry for the Saturn encounter from the SU 80 mission. During descent, the modulation is changed to FSK with a data rate of 50.5 bps. The worst-case trajectory is again the Saturn/SU 80 mission, as seen in the figure, and the maximum RF power required occurs at mission completion and is 40 watts. These values are also summarized in Table V-9. Also indicated on the curves are descent pressures for the two worst-case atmosphere models.
An RF link analysis for the worst-case descent condition for Task II is shown in Table V-ll. FSK modulation is used with a data rate of 50.5 bps at 0.86 GHz. RF power of 40 watts is required at the end of mission in the cool atmosphere model for a Saturn encounter from the SU 80 trajectory.
Table V-12 depicts design details of the RF components that comprise the telecommunications subsystem for the common Saturn/ Uranus probe with Task II science payload. The probe transmitter has two RF power levels of 7- and 41-watts available for the mission; 7-watts is sufficient for pre-entry during probe acquisition when only the tone is being transmitted; 41 watts is sufficient to handle the high pre-entry data rate for Task II using PSK modulation, and descent using FSK modulation with a lower data rate. Other components that comprise the subsystem remain unchanged from Task I as discussed in Chapter IV, Section D. For the increased transmitter RF power level, a mechanical antenna switch on the probe will most certainly be required to provide more reliable operation over a solid-state switch. A switch is necessary to connect the 41-watt transmitter output terminal from the pre-entry to descent antenna during planet entry. The transmitter output will be reduced to zero during the blackout period; only the oscillator and modulator will be energized during this time.
Design details of the spacecraft receiver are beyond the scope of the study. In general, a PSK receiver with two additional data filters for FSK could be used. Acquisition takes place with the logic technique described in Volume III, Appendix C. During pre-entry, the PSK receiver performs in a normal manner. During descent, the receiver provides the tracking function only, and the additional data channel filters demodulate the FSK data. The fact that PSK (235 bps) and FSK (50.5 bps) signals pass through different receiver channels simplifies the bit synchronization acquisition problem because each channel can be designed for an appropriate bit rate.
V-15
Table V-11 Probe Descent Telemetry Link Design Table for the Task II Mission*
PARAMETER
1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
11.
12.
TOTAL TRANSMITTER POWER, dBw
TRANSMITTING CIRCUIT LOSS, dB
TRANSMITTING ANTENNA GAIN, dB
COMMUNICATIONS RANGE LOSS, dB
PLANETARY ATMOSPHERE & DEFOCUSSING LOSS, dB
POLARIZATION LOSS, dB
ANTENNA PATTERN RIPPLE LOSS, dB
RECEIVING ANTENNA GAIN, dB
RECEIVING CIRCUIT LOSS, dB
NET CIRCUIT LOSS, l(2-*~9), dB
TOTAL RECEIVED POWER (1 + 10), dBw
RECEIVER NOISE SPECTRAL DENSITY, dBw/Hz
TRACKING TONE
13.
14.
15.
16.
17.
18.
DATA
19.
20.
21.
22.
23.
24.
25.
26.
27.
TONE POWER/TOTAL POWER, dB
RECEIVED TONE POWER (11 + 13), dBw
TRACKING THRESHOLD BANDWIDTH, dB
THRESHOLD SIGNAL/NOISE RATIO, dB
THRESHOLD TRACKING POWER (12 + 15 + 16), dBw
TRACKING PERFORMANCE MARGIN (14 - 17), dB
CHANNEL
DATA POWER/TOTAL POWER, dB
RADIO SYSTEM PROCESSING LOSS, dB
FADING LOSS, dB
RECEIVED DATA POWER ( 1 1 + 1 9 + 2 0 + 2 1 ) , dBw
DATA BIT RATE, dB
THRESHOLD Eb/NQ, dB
THRESHOLD DATA POWER (12 + 23 + 24), dBw
PERFORMANCE MARGIN (22 - 25), dB
NOMINAL LESS ADVERSE VALUE (26 - 26 ADVERSE), dB
NOMINAL VALUE
16.0
-1.1
6.0
-196.5
-4.1
-0.3
0
17.9
-1.0
-179.1
-163.1
-196.2
-6.6
-169.7
12.4
10.0
-173.8
4.1
-1.1
-1.0
-1.0
-166.2
17.0
8.9
-170.3
4.1
0
•CONDITIONS: 1. S/SU 80 COMMON PROBE MISSION (WORST-CASE TRAJ 2. DESCENT TASK II SCIENCE PAYLOAD AT MISSION CO 3. BINARY FSK MODULATION, CONVOLUTIONAL ENCODER. 4. M = 2, V = 2, Q = 8. 5. DECODER - VITERBI ALGORITHM, BER = 5 x 10"5 K
ADVERSE TOLERANCE
0
0.1
1.8
0.3
0.2
0
0.2
1.0
0.1
3.7
3.7
0.4
0
3.7
0
0
0.4
4.1
0
0
0
3.7
0
0
0.4
4.1
ECTORY). MPLETION.
ITH K = 8 COC
REMARKS
39.8 w AT 0.86 GHz.
CABLE & SWITCH
100° BEAMWIDTH.
1.8 x 105 km
COOL, 18 bar-
20° BEAMWIDTH
CABLE.
Ts = 1750°K
NF =8.5 dB
Pt = 8.8 w
17.5 Hz
Pd = 31 w
50.5 bps
E.
V-16
Table V-12 Teleaommunioations RF Subsystem for the Task II Mission*
COMPONENT
TRANSMITTER
RF SWITCH
PROBE PRE-ENTRY & DESCENT ANTENNAS
SPACECRAFT ANTENNA
SPACECRAFT RECEIVER
•CONDITIONS: 1. P 2. M 3. F 4. B
CHARACTERISTIC
RF POWER OUT OVERALL EFFICIENCY MAXIMUM DC POWER IN AT 28 vdc TOTAL WEIGHT OF DUAL-OUTPUT POWER & MODULATOR
VOLUME
TYPE INSERTION LOSS WEIGHT
VOLUME
TYPE MAIN BEAM ANGLE BEAMWIDTH MAXIMUM GAIN SIZE (DIAMETER X HEIGHT)
WEIGHT
TYPE BEAMWIDTH MAXIMUM GAIN SIZE (DIAMETER)
WEIGHT
DESPIN POSITION SEARCH FREQUENCY ACQUISITION CLOCK ANGLE, 9 CONE ANGLE, <j>
NOISE TEMPERATURE NOISE FIGURE DC POWER IN AT 28 vdc WEIGHT
LANETS: SATURN & URANUS: ARINER SPACECRAFT; REQUENCY = 0.86 GHz; IT RATE = 235 & 50.5 bps.
VALUE
7 AND 41 w 45% 91 w
2.82 kg (6.2 lb)
3540 cm3 (216 in.3)
MECHANICAL SPDT 0.3 dB 0.23 kg (0.5 lb)
443 cm3 (27 in.3)
TURNSTILE/CONE 0 deg 100 deg 6.5 lb 15.2 x 8.25 cm (6 x 3.25 in.)
0.45 kg (1.0 lb)
PARABOLIC DISH 20 deg 18.3 dB 128 cm (50.4 in.) (4.2 ft)
4.54 kg (10 lb)
NO ONE 31 sec 257 deg 116 deg
300°K 3.1 dB 3.0 w 0.9 kg (2.0 lb)
V-17
E. DATA HANDLING SUBSYSTEMS
The data processing electronics for Task II are essentially unchanged from the previous study (Chapter III, Section D.2). The modifications necessary to provide the inflight target planet selection consists of approximately 12 ROMs for the additional formats and a latching relay to implement the selection. As previously discussed, COS/MOS data storage design has been employed for this mission. The physical and electrical characteristics are approximated by the following equations:
V = volume = 443 + 16.0 M (10~3) cm3
W = weight = 0.2 + 0.715 M (10_3) kg
Power (standby) = P = 0.2 + 11.0 M (10~6) w
Power (operate) = P =1.0+5.50 M (10-6) w
M = memory = number of bits stored
The assumption is made that fifty-nine 1024-bit chips (M = 60,416 bits) will be required to store the blackout data. This results in the following values for physical and electrical interface characteristics of the memory electronics.
V = 1410 cm3 W = 0.65 kg P = 0.88 w P = 4.4 w s o
Similar characteristics of the data processing electronics are assumed the same as in the previous study (Volume I thru III)•
V = 2330 cm3 W = 2.13 kg P = 6.9 w
The DHS begins the initial descent sequence shortly after the descent battery is energized. The initiation of the sequence and the initial state of the DHS will be established by a timing circuit working off battery voltage. This will ensure that the turn-on battery transient has subsided, and will eliminate the need of an additional event from the coast timer. The required sequence of events is illustrated in Table V-6 (Saturn Sequence of Events Task II). The pre-entry sequence occurs from Item 19 to Item 22 inclusive. The actual time with respect to entry of these events will vary (+4.4 min, Saturn; +29 min, Uranus) due to trajectory uncertainties. The control state established at Item 22 will remain fixed regardless of the actual time before entry until science data is detected (Item 27). The functions required of the DHS formats with respect to the Task II Saturn sequence are tabulated.
V-18
From To Item Item
19
25
25
27
33
34
25
27
27
33
34
35
Store engineering data
Interleave real-time pre-entry science and engineering data with stored engineering data
Store pre-entry science and engineering data
Store entry accelerometer and engineering data
Store descent science and engineering data
Interleave real-time descent science and engineering data with stored science and engineering data in this order: (a) Item 27 to 34, (b) Item 25 to 27.
The interleaved data (Item 25 to 27, and Item 34 to 35) is sequenced through the encoder to RF transmitter.
POWER AND PYROTECHNIC SUBSYSTEM
The configuration of the power and pyrotechnic subsystem for Task II is unchanged from the previous study (Volume I thru III), as indicated in Chapter III, Section D.3. The discussion on post-separation power, coast timer power, and initial entry/descent pyrotechnic power of Task I (Chapter IV, Section D.3) applies identically to Task II.
As in Task I, the entry and descent battery specification is based on the power required for subsystem components plus losses and margins and the sequence of events for a late arrival probe. The maximum transmitter power at Saturn and a late arrival probe at Uranus provides the worst-case requirements for the battery. The nominal power demands of the various subsystems are listed.
V-19
Subsystem
Transmitter
DHS
Eng Hskpg
Distribution
Losses
Nephelometer
Power (watt)
25/114
7.7/10.7
2.0
Subsystem
NMS
Pressure
Temperature
Accelerometer
Power (watt)
14.0
1.3
1.4
2.8
15 (max)*
3.0
IRPA
NRPA
3.0
5.0
Langmuir Probe 3.0
*Assumes 5% of total power.
The margins that are applied consist of a +29-min trajectory uncertainty at Uranus, +5% uncertainty descent time, and 10% overall uncertainty in power levels. These calculations result in a total required energy storage of 227 w-hr and a peak demand of 169 watts at 28 volts. Using an energy density of 31 w-hr/lb and 0.138 w-hr/cm3, the entry/descent battery physical specification is 3.31 kg and 1640 cm3.
The physical specification of the power and pyrotechnic subsystem is based on a summation of the estimated required equipment as indicated in the following:
SCR - two per (dual bridgewire) event;
Relays - one per event;
Capacitor bands - six discharges allowed per bank; sufficient banks must be allowed to provide separate banks for simultaneous (less than 15 sec apart) discharges;
Electronics - engineering estimate.
V-20
The number of capacitor banks (8) is controlled by the number of simultaneous events (4) for these missions. The total of 16 events requires 32 SCRs and 16 relays. The increase in the number of events over Task I is due to the required removal of the NMS cover. The physical characteristics of the pyrotechnics are tabulated.
Total Size/Weight, Number Size/Weight,
Component cm3/g Required cm3/g
Capacitor bank 164/81.8 8 1310/0.66
Relays 14/34.1 16 224/0.55
SCR 13.6/14.1 32 360/0.45
Electronics 1230/0.91 1 1230/0.91
The size and weight of each element includes 10% to 15% for packaging.
G. ATTITUDE CONTROL
The attitude control dynamics are minimal for the design configuration considered for the Saturn/Uranus probe. The errors involved are:
1) initial error due to spacecraft pointing;
2) drift error accumulation during period between separation and spinup due to tip-off rate;
3) error due to tip-off rate combined with spinup;
4) error due to spin-up jet misalignment.
Only momentum vector errors are of significance in this analysis since nutation errors are damped out and the spin axis (principal moment of inertia) will align with the momentum vector. The momentum vector errors involved are discussed in the following paragraphs:
V-21
1) Initial error is a function of the spacecraft and will not be discussed.
2) Drift error - assuming an initial tip-off rate of 0.5 deg/sec about a nominal X-axis, the spin axis is displaced in the minus y direction.
°T = WT 'D 8T = d r i f t e r r ° r
W = tip-off rate = 0.5 deg/sec t = drift period =0.5 sec
9T = 0.25 deg
3) As indicated in Chapter III, Section D.4, the error may be calculated from knowledge of the spin momentum torque, P, transverse moment of inertia, 1^ and spin torque, m.
W = 0.52 rad/sec
m = 3.4 N-m
q = q2/ml = 0.654
G (P) < 0.81° xv —
Since the combination of the drift and spin-up/tip-off error is not random, the error due to (2) and (3) is,
6 (P) < 0.81° 6 (P) < -0.51° = -0.25° - 0.26° x — y
4) The analysis of Chapter III, Section D.4 indicated the spin torque vector offset, 6 = 0.136°. The total rss momentum vector error is then
e_ = 0.97°
I = 6 . 9 kg-m2
s b
I = 5.6 kg-m2
P (0) /P = 0.795 (deg)
6 (P) < -0 .26°
V-22
H. STRUCTURAL AND MECHANICAL SUBSYSTEM
The addition of the pre-entry science instruments of Task II to the Saturn/Uranus probe results in a general increase in: its size and weight over that of the baseline probe. However, the design configuration is unchanged except for probe volume, and attachment provisions for the added science instruments and their support components. The changes to the probe are described briefly in this section.
1. Configuration and General Arrangement
The configuration of the entry and descent probes is shown in Figure V-4. The pre-entry science sensors require location on the forward (flight direction) face of the entry probe, such that they may sense undisturbed incoming particles from space. To accommodate this requirement, the NRPA and IRPA sensors are mounted directly on the forebody heat shield, projecting through the multilayer insulation so that the inlet to the sensors is actually ahead of the nose of the heat shield. In like manner, the Langmuir probe sensors are located on the forebody heat shield. The electronic black boxes associated with these sensors are mounted internal to the aeroshell and around its periphery so as to maximize the spin moment of inertia of the entry probe. Wiring joining the sensors to the electronic boxes is routed between the heat shield and the multilayer insulation blanket.
The physical attachment of the sensors to the heat shield is by means of a thermosensitive adhesive, such that the attachment will be severed by initial entry heating, permitting the sensors to separate. A high-thermally-conductive, low-mass attachment flange on the sensor will expedite separation of the sensors at entry prior to high heating.
An alternative approach for sensor separation would be the use of pyrotechnic devices to sever the structural connection. The final choice will involve a total system study of the heating, structural loading, and dynamic stability of the probe. However, for purposes of the system level definition, it appears that the bonded attachment is a practical approach.
The nephelometer is located inside the descent probe with a viewing window in the side. Although it is not shown in Figure V-4, a muff-type cover, attached to the descent-probe-retaining clamp band, protects the window from contamination until descent probe separation after entry. The window cover remains with the aeroshell structure at separation.
V-23
Page intentionally left blank
The nephelometer views horizontally through the side window of the descent probe. The remainder of the descent science instruments are located in the same relative orientation as the Task I instruments described in Chapter IV, Section H.l.
The Task II descent probe has a diameter of 48.9 cm (19.25 in.) and weighs 50.3 kg (110.8 lbm), The Task II entry probe has a diameter of 92.1 cm (36.2 in.) and weighs 103.2 kg (227.3 lbm). The configuration of the probe is the same as that for Task I except as noted herein.
Structural Design
The structural design of the Task II science probe is identical to that described in Chapter IV, Section H.2 for the Baseline (Task I) configuration except for physical size. The planetary entry decelerations are the same for either configuration, resulting in the same relative loads.
Mass Properties
The weights breakdown for the Task II configuration is presented in Table V-13. The weights are first presented by subsystem to arrive at a total cruise configuration weight, and are then presented at the end of the table for the various phases of flight.
The moment of inertia data for this probe has been computed about two axes. These data are as tabulated.
Descent Probe Entry Probe
I . 1.19 kg/m2(0.88 slug/ft2) 6.80 kg/m2(5.00 slug/ft2)
Vtch/yaw 1'11 k s / m 2 ( 0 , 8 2 slug/ft2) 5'87 kg/m2(4.32 slug/ft2)
Parachute Subsystem
The added science complement of Task II results in an increased size and weight of both the descent and the entry probe configurations. Since the requirements imposed by science for entry and descent ballistic coefficients are the same as for Task I, the parachute sizes for Task II are increased. The separation parachute increases in size from 2.29 m (7.5 ft) to 2.50 m (8.2 ft),
V-27
Table V-13 Taak II Saturn/Uranus Probe Weight Breakdoim (Added Science)
SCIENCE
TEMPERATURE GAGE PRESSURE TRANSDUCER ACCELEROMETER SENSOR
ELECTRONICS (CONVERTER) NEUTRAL MASS SPECTROMETER ANALYZER
ELECTRONICS PUMP BALLAST TANK
NEPHELOMETER IRPA SENSOR IRPA ELECTRONICS NRPA SENSOR NRPA ELECTRONICS LANGMUIR PROBES (2) LANGMUIR PROBE ELECTRONICS
POWER 8 POWER CONDITIONING
POWER CONDITIONER POWER DISTRIBUTION BOX POWER FILTERS ENTRY BATTERIES POST-SEPARATION BATTERIES COAST BATTERIES
CABLING
INNER PROBE EXTERNAL STRUCTURE
DATA HANDLING
DATA HANDLING SYSTEM MEMORY BANKS
ATTITUDE CONTROL SYSTEM (LESS PROPELLANT)
ACS SYSTEM S TANKS NUTATION DAMPER COAST TIMER
COMMUNICATIONS
PRE-ENTRY ANTENNA DESCENT ANTENNA RF TRANSMITTER RF ANTENNA SWITCH
PYROTECHNIC SUBSYSTEM
PYRO ELECTRONICS PYRO CAPACITORS (PROBE) PYRO CAPACITORS (EXTERNAL) PYRO RELAYS (PROBE) PYRO RELAYS (EXTERNAL) PYRO SCR (PROBE) PYRO SCR (EXTERNAL) PYRO SQUIBS PYRO THRUSTER
STRUCTURES 8 HEAT SHIELDS
DESCENT PROBE STRUCTURE EQUIPMENT SUPPORT DECK BASE COVER AEROSHELL (2 lb FOR PAYLOAD RING) FORWARD HEAT SHIELD (18.98 lb ABLATED DURING ENTRY) AFT HEAT SHIELD
WEIGHT
kg
0.45 0.68 0.59 0.91 1.82 2.72 0.45 0.45 1.14 0.23 1.36 0.45 1.82
1.36
14.44
0.23 0.68 0.91 3.31 0.25 0.41
5.79
2.77 1.73
4.50
2.13 0.65
2.78
1.64 1.09 0.11
2.84
0.45 0.45 2.82 0.23
3.95
0.91 0.33 0.33 0.17 0.38 0.14 0.31 0.26 3.61
6.44
4.02 3.79 4.04 5.08 19.07 1.12
37.12
lb
1.00 1.50 1.30 2.00 4.00 6.00 1.00 1.00 2.50 0.50 3.00 1.00 4.00
3.0 31.80
0.50 1.50 2.00 7.30 0.55 0.90
12.75
6.10 3.82
9.92
4.70 1.42
6.12
3.60 2.40 0.25
6.25
1.00 1.00 6.20 0.50
8.70
2.00 0.72 0.72 0.37 0.83 0.32 0.68 0.60 7.95
14.19
8.85 8.34 8.9 11.20 42.01 2.46
81.76
MECHANISMS PIN PULLER LATCHES S BANDS MAIN PARACHUTE (3.51 lb IN CHUTE 8 RISER) SECONDARY PARACHUTE CLAMP SEPARATORS
THERMAL
EXTERNAL INSULATION BLANKET (FORWARD HEAT SHIELD)1 EXTERNAL INSULATION BLANKET (BASE COVER) 1 PROBE HULL INSULATION (INTERNAL) ISOTOPE HEATERS ENVIRONMENTAL TANK 8 N2
PROPULSION
ACS PROPELLANT
TOTAL
15% CONTINGENCY
1. PRE-ENTRY WEIGHT
ACS PROPELLANT
103.22 - 0.01 = 103.21 kq (227.36 - 0.03 = 227.33 lb)
2. POST-ENTRY WEIGHT
FORWARD HEAT SHIELD (ABLATED) FORWARD INSULATION BLANKET I AFT INSULATION BLANKET I PRE-ENTRY ANTENNA
15S CONTINGENCY
103.21 - 11.89 « 91.32 kq (227.33 - 26.20 = 201.13 lb)
3. INITIAL WEIGHT ON PARACHUTE
PARACHUTE (SEPARATION)
15K CONTINGENCY
91.32 - 1.83 = 89.49 kq (201.13 - 4.04 = 197.09 lb)
4. FINAL WEIGHT ON PARACHUTE
PYRO THRUSTERS PYRO CAPACITORS PYRO RELAYS PYRO SCR PYRO SQUIBS AEROSHELL FORWARD HEAT SHIELD (NOT ABLATED) BASE COVER 8 AFT HEAT SHIELD SEPARATION PIN PULLERS LATCHES 8 BANDS ISOTOPE HEATERS EXTERNAL CABLING NUTATION DAMPER ACS SYSTEM SEPARATION BATTERY CLAMP SEPARATORS
153 CONTINGENCY
89.49 - 39.17 = 50.32 kq (197.09 - 86.27 = 110.82 lbs)
WEIGHT
kg
0.65 0.91 4.00 0.64 0.47
6.67
1.27
1.59 1.82 0.54
5.22
0.01
0.01
89.76
13.46
103.22
0.01
0.01
8.62 1.27
0.45
10.34
1.55
11.89
1.59
0.24
1.83
3.61 0.33 0.38 0.31 0.18 5.08 10.45 5.16 0.65 0.91 1.82 1.73 1.09 1.64 0.25 0.47
34.06
5.11
39.17
lb
1.44 2.00 8.80 1.40 1.04
14.68
2.80
3.50 4.00 1.20
11.50
0.03
0.03
197.70
29.66
227.36
0.03
0.03
18.98 2.80
1.00
22.78
3.42
26.20
3.51
0.53
4.04
7.95 0.72 0.83 0.68 0.40 11.20 23.03 11.36 1.44 2.00 4.00 3.82 2.40 3.60 0.55 1.04
75.02
11.25
86.27
V-28
and weighs 4.00 kg (8.8 lbm); the descent parachute increases in size from 0.73 m (2.4 ft) to 0.89 m (2.9 ft) and weights 0.64 kg (1.4 lbm). The remainder of the parachute description is the same as presented in Chapter IV, Section H.5.
Heat Shield
The planetary entry conditions are the same for the probe with Task II instrumentation as for Task I instrumentation, and consequently the fofebody heat shield mass fraction is the same for both. The probe is heavier for Task II, however, resulting in a forebody heat shield weight of 19.07 kg (42.01 lbm). The afterbody ablator is ESA 3560, the same as in Task I, and is assumed to be evenly distributed over the afterbody at 1.37 kg/m2 (0.28 lbm/ ft2).
PROPULSION SUBSYSTEMS
As in Task I, a stored gas (nitrogen) propulsion system is used for probe spin at probe/spacecraft separation to stabilize its separation pointing attitude. The spin-axis moment of inertia for the Task II probe has increased to 6.80 kg/m2(5.00 slug/ft2) from 5.02 kg/m2(3.7 slug/ft2) for the Task I probe. The moment arm of the spin nozzles has increased to 41.9 cm (16.5 in.) from 38.6 cm (15.2 in.). Thus, the amount of gas required increases by 24%. However, since the weight of the gas is less than 1% of the total ACS spin-up system, the total weight is the same for both probes.
THERMAL CONTROL SUBSYSTEM
The thermal design developed and used for the SAG exploratory payload is directly applicable for the expanded science complement. The additional science results basically in only increased probe size, mass, and internal power, and the design trades and parametrics presented in Chapter III are, therefore, applicable for the probe design of Task II. Since the additional probe mass and power dissipation are desirable from the standpoint of thermal control, minor adjustment of the probe arrival temperature can be accomplished. Analysis showed, however, that only
V-29
a 2°K influence would result; therefore no adjustment in the probe coast temperature was considered. Perhaps the largest influence of the expanded science payload is on the probe coast thermal control. The increased penetrations in the multilayer blanket for the pre-entry science measurements result in more radioisotope heating to maintain positive probe heat balance. Again, due to the criticality of the probe entry temperature for descent thermal control, a louver system design is recommended to control probe coast temperatures prior to entry and reduce the thermal uncertainty.
Using identical thermal subsystem design as outlined for the SAG exploratory payload (i.e., 2 cm internal foam insulation with 2.5 bars of neon gas environmental control), the expanded science complement probe design was analyzed for worst-case atmospheric encounters and arrival uncertainties. Table V-14 presents the minimum and maximum equipment temperatures experienced, and Figures V-5 and V-6 present plotted thermal analysis results. As before, the warm atmospheric encounters with late arrival uncertainty and the cool atmospheric encounters with early arrival uncertainty establish the probable operative temperature bands to be experienced at each planet. Assuming a 302°K probe arrival temperature, these bands are as listed:
Saturn, °K Uranus, °K
RF transmitter 297-331 291-341
Insulated primary batteries 295-307 281-309
Aggregate internal equipment 283-305 265-312
Again, the large Uranus arrival uncertainties are an important consideration for establishing acceptable pre-entry transmitter power levels. In addition, since the RF power requirement for Task II is 41 watts as compared to 25 watts for Task I, the RF transmitter becomes the pacing item that controls the allowable probe entry temperature. For Task II, therefore, the use of louvers to reduce the probe arrival temperature uncertainty, or the implementation of phase change material within the transmitter, presents a design trade to be carefully evaluated as a more detailed probe design evolves.
V-30
«-1222 days- -35.8 days-
SPACECRAFT CRUISE & CHECKOUT
PROBE COAST
28 minutes-
UPPER RF TRANSMITTER OPERATIONAL LIMIT (345 °K k
PRE-ENTRY ACTIVATION
10 20 30 40 ATMOSPHERIC DESCENT, minutes
LAUNCH SPACECRAFT NEAR ENTRY SEPARATION SATURN
ARRIVAL
a) WARM - ATMOSPHERE ENCOUNTER WITH LATE ARRIVAL UNCERTAINTY (MAXIMUM ANTICIPATED TEMPERTAURES)
350
325
300
TEMPERATURE, °K
275
250
225
-1222 days- -35.8 days-
SPACECRAFT CRUISE & CHECKOUT
^19 minutes*
LEGEND:
RF TRANSMITTER TEMPERATURE
PRIMARY BATTERY WITH 1-cm FOAM INSULATION
AGGREGATE INTERNAL EQUIPMENT
END OF MISSION
UPPER RF TRANSMITTER OPERATIONAL LIMIT (345
LAUNCH SPACECRAFT SEPARATION
NEAR SATURN ARRIVAL
ENTR\ END OF MISSION
b) COOL - ATMOSPHERE ENCOUNTER WITH EARLY ARRIVAL UNCERTAINTY (MANIMUM ANTICIPATED TEMPERATURES)
Fig. V-S Saturn/Uranus Probe Thermal History for Saturn Mission (Expanded Science Pay load)
V-31
TEMPERATURE, °K
275
-2817 days-77
-minutes-
SPACECRAFT CRUISE & CHECKOUT
PROBE COAST
UPPER RF TRANSMITTER OPERATIONAL LIMIT (345 °K)
PRE-ENTRY 0 ACTIVATION
LAUNCH SPACECRAFT SEPARATION
NEAR URANUS ARRIVAL
10 20 30 40 ATMOSPHERIC DESCENT, minutes
ENTR>
a) WARM - ATMOSPHERE ENCOUNTER WITH LATE ARRIVAL UNCERTAINTY (MAXIMUM ANTICIPATED TEMPERATURES)
-2817 days-e-U— 9.5 days 19
"minutes
LEGEND:
•RF TRANSMITTER TEMPERATURE
•PRIMARY BATTERY WITH 1-cm FOAM INSULATION
•AGGREGATE INTERNAL EQUIPMENT
END OF MISSION
UPPER RF TRANSMITTER OPERATIONAL LIMIT (345 °K)
- UPPER EQUIPMENT OPERATION LIMIT (320 »K)
SPACECRAFT CRUISE & CHECKOUT
PROBE COAST
PRE-ENTRY 0 ACTIVATION
10 20 30 40 ATMOSPHERIC DESCENT, minutes
LAUNCH SPACECRAFT NEAR ENTRV SEPARATION URANUS
ARRIVAL
b) COOL - ATMOSPHERE ENCOUNTER WITH EARLY ARRIVAL UNCERTAINTY (MINIMUM ANTICIPATED TEMPERATURES)
Fig. V-S Saturn/Uranus Probe Thermal History for Uranus Mission (Expanded Science Pay load)
END OF MISSION
V-32
Table V-14 Table V-24 Minimum and
and Arrival Maximum Probe Temperatures for Worst-Case Atmospheric Encounters Uncertaintiesj Expanded Science Payloads
P&JBE COAST TEMPERATURE
RF TRANSMITTER
INSULATED PRIMARY BATTERIES
AGGREGATE INTERNAL EQUIPMENT
PRE-ENTRY TEMPERATURES
RF TRANSMITTER
INSULATED PRIMARY BATTERIES
AGGREGATED INTERNAL EQUIPMENT
MAXIMUM DESCENT TEMPERATURES
RF TRANSMITTER
INSULATED PRIMARY BATTERIES
AGGREGATED INTERNAL EQUIPMENT
MINIMUM DESCENT TEMPERATURES
RF TRANSMITTER
INSULATED PRIMARY BATTERIES
AGGREGATED INTERNAL EQUIPMENT
EQUIPMENT TEMPERATURE LIMIT, °K
225/350
225/325
240/325
250/345
275/325
255/320
345 325 320
250 275 250
vmmmMAXIMUM PROBE TEMPERATURES
\\\\\\\\\\\\\MINIMUM PROBE TEMPERATURES
SATURN
WARM ATMOSPHERE (LATE ARRIVAL UNCERTAINTY)
302°K .
314°K
304° K
305°K
331°K
307°K
305° K
314°K
304°K
300" K
COOL ATMOSPHERE (EARLY ARRIVAL UNCERTAINTY)
302° K
310°K
302"K
304 °K
323°K
302° K
304° K
297°K
295°K
283°K
URANUS
WARM ATMOSPHERE (LATE ARRIVAL UNCERTAINTY)
302 °K
322° K
308° K
312°K
gm&fes •28mm zMmii
322°K
308° K
296° K
COOL ATMOSPHERE (EARLY ARRIVAL UNCERTAINTY) ,
302°K
310°K.
302° K
304° K
323°K>
302°K'
304° K
HHl M|iii Hl
K. PROBE-TO-SPACECRAFT INTEGRATIONS
Integration of the probe and spacecraft is the same for Tasks I and II, as reported in Chapter IV, Section K.
V-33
