NASA TECHNICAL MEMORANDUM · George C. Marshall Space Flight Center ... EPS Block Diagram ... where STS performance limitations, Van

NASA TECHNICAL

MEMORANDUM NASA TM 83435

I -

Accommodations Assessment: 6 t

(*I I A +O tn n Spaceborne Doppler Lidar .-, .. . F

a0 a 3 , Wii;d Measuring System

By Program Development Preliminary Design Office

August, 1981

NASA

George C. Marshall Space Flight Center Marshall Space Flight Center, Alabama

https://ntrs.nasa.gov/search.jsp?R=19810023201 2018-08-02T13:34:13+00:00Z

ACCOMMODATIONS ASSESSMENT SPACEBORNE DOPPLER LIDAR

WIND MEASURING SYSTEM

CONTENTS

PAGE -

A. BACKGROUND AND SCOPE--------------------------------- 6 B. ASSESSMENT CONCLUSIONS/RECOMMENDATIONS--------------- 7 C. MISSION OBJECTIVES: WIND MEASUREMENT NEEDS----------- 10 D. MISSION TECHNIQUE: PRINCIPLES OF WIND MEASUREMENT---- 1 4

1II.CONFIGURATION AND MASS PROPERTIES------------------------ 24

A, REQUIREMENTS AND ASSUMPTIONS------------------------- 25 8. CONCEPTUAL DESIGN--------------------------L--------- 25 C. S Y S T ~ SIZING---------------------------------------- 27

A. ELECTRICAL POWER------------------------------------- 3 0 B. COMMUNICATIONS AND DATA MANAGEMENT------------------- 35 C. ATTITUDE DETERMINATION AKD CONTROL------------------- 53

V. MISSION OPERATIONS AND PERFORMANCE----------------------- 5 8

A. WIND ACCURACY REQUIREMENT: ORBIT SELECTION----------- 5 9 B. SHUTTLE PERFORMANCE: DESIGN AND OPERATIONAL

IM~LICATIONS----------------------------------------- 63

KEY CONTilIBUTORS (CONTRIBUTOR, OFFICE SYMWL, DISCIPLINE

The assessment activities summarized here were performed by the below-named members of the MSFC Preliminary Design Off ice (PDO1). However, it is appropriate to acknowledge the assistance of Mr. J. Bilbro (EC32) relative to the characteristics and principles of doppler lidar wind measuring systems. Also, Mr. R. Beranek (PS06) provided guidance to the over-all assessment of which these PDOl activities were a part.

J. STEINCAMP, PD34, ENGINEERING STUDY MANAGER

J. HOWELL, PD12, SUBSYSTEMS LEAD ENGINEER/ATTITUDE DETERMINATION AND CONTROL

W. FINNELL, PD14, COMMUNICATIONS 61 DATA MANAGEMENT

W. JOHANSON, PD24, MASS PROPERTIES

D. MERCIER, PD33, STS PERFORMANCE

J. PEOPLES, PD22, THERMAL ANALYSIS

R. ROOD, PD14, ELECTRIC9L POWER

T. WHEELGR, PD32, ATMOSPHERIC ENVIRONMENT

A. YOUNG, PD32, FLIGHT PERF3RYANCE

D. WALTON, PD32, DOCUMENTA'FION

LIST OF FIGURES

I .A-1

1.B-1

I. D-1

1I.A-1

1I.B-1

1I.B-2

1II.A-1

1V.A-1

IV. A-2

1V.A-3

IV. A-4

1V.A-5

1V.A-6

1V.A-7

IV. A-8

IV. B-1

1V.B-2

Doppler Lidar Wind Measuring System

Programmatic Schedule

Lidar System Concept (Laser 6 Optical Subsystems)

SNR vs . Altitude Molecular Absorption Coefficient

IR Backscattering Coefficient as a Function of Altitude

Launch Configuration

EPS Block Diagram

Solar Array Block Diagram

Suntime Fraction (Sun-Synchronous Orbit)

Suntime Fraction

EPS Efficiency Block Diagram

Solar Array Cross Section

Solar Array Sizing

LIDAR EPS Summary

LIDAR Pulse Temporal and Spectral Characteristics

TDRSS Coverage Geometry

IV. B-3 TDRSS Dead Zone

1V.B-4 Tape Recorder Design Life

1V.B-5 CDMS Block Diagram

1V.B-6 CDMS Summary

1V.C-1 Spacecraft Coordinate System

1V.C-2 Local Vertical Misalignment Geometry

1V.C-3 ACDS Block Diagram

1V.C-4 ACDS Summary

3

, C ' s l - - L *-,. - - --.

LIST OF FIGURES (CONTINUED)

V.A-1 Wind Accuracy Variation with SNR

V.A-2 Wind Accuracy Sensitivity to Signal-to-Noise Ratio

V.A-3 Comparison of Orbital Altitudes and Inclinations

V. B-1 Shuttle Performance (VAFB Launch)

V.B-2 Augmented Shuttle Performance (VAFB, Unofficial)

V.B-3 Shu1:tle Performance (KSC, Unofficial 1

I. INTRODUCTION

I. INTRODUCTION

I.A. BACKGROUND AND SCOPE

This dccument summarizes the activities carried out by the MSFC Preliminary Design Office as part of the NASA assessment of a spaceborne doppler lidar wind measuring sys-

, tem concept. The assessment was requested by Dr. Ron Greenwood, Director of the Environmental Observation Division, Office of Space and Terrestrial Applications, of NASA Headquarters, who designated MSFC as the lead center, to be supported by JPL, LaRC, and GSFC in selected disciplines augmenting MSFC's experience in ground and airborne pulsed C02 doppler lidar wind measuring systems.

The assessment activity was organized to concentrate on three areas vital to concept feasibility:

o Laser and optical systems (Lidar S Y S C ~ hardware

o Atmospheric characteristics

o Spacecraft Accommodations

with corresponding assignments of responsibilities to the MSFC Optical and RF Systems Division, Atmospheric Sciences Division, and Preliminary Design Office. Overall assessment coordination was provided by the MSFC Advanced Studies Office.

The system concept under consideration was defined in three reports:

1) "Feasibility Study of Satellite - Borne Lidar Global Wind Monitoring System", NOAA Tech. Memo ERL WPL-37 (1978)

2 ) "Feasibility Study of Satellite - Borne Lidar Global Wind Monitoring System, Part II", NOAA Tech. Memo ERL WPL-63 ( 1980)

3 ) NOAA LMSC WINDSAT Study, Final Briefing Charts (Sep 1980) Contract NA 79 RAC 00127

The L&c study was managed by the NOAA ~nvironmental Resources Laboratory and funded by the USAF Space Division. All three studies emphasized a Shuttle-borne system, which was considered an evolutionary step in the development of an operational system.

A "clean sheet" approach to spacecraft conceptual design was taken to provide the greatest flexibility in accommodating the mission equipment requirements. This seemed particularly appropriate in view of the electrical power requirements, which would require significant modification of existing spacecraft. Figure 1.A-1 is an artist's concept of the operational system.

I.B. ACCOMMODATIONS ASSESSMENT CONCLUSIONS/ RECOMMENDATIONS

Subject to the three caveats following, the principal conclusion of this assessment is that a spacecraft with the capabilities needed to support the operation of the doppler lidar wind measuring system is technically feasible. Cost estimates prepared for a dedicated, new-development spacecraft (using standard, available, and/or existing technology components) suggest economic feasibility of such a spacecraft. Such a new development may be unnecessary: the present assessment did not include detailed consideration of the potential use of existing and "in-development" spacecraft and platforms. The caveats are that

1) cryogenic cooling of the photodetector will not be required in the operational system,

2 ) detailed structural/pointing analyses will establish feasibility of short-terra optical axis stability for efficient lidar heterodyning, and

3 laser power requirements will not increase beyond those now anticipated by more than a factor of three.

The structural and pointing specialists consulted were of the opinion that an acceptable short-term stability could be realized, but that proof of their opinion would require analyses beyond the scope of the assessment. The laser power requirements are a critical design driver: the spacecraft conceptualized in this assessment will not be adequate if the laser power requirement increases much beyond the assumed range of 2140-2340 W. Since the laser power requirement ir most particularly determined by atmospheric backscatter, better 'knowledge of this parameter is required to adequately scope the spacecraft definition activity.

The uncertainty relative to atmospheric backscatter, together with the need for demonstration of adequate laser lifetime and chirp (intra-pulse frequency stability) characteristics, were key considerations in the formulation of the schedule recommendation shown in Figure 1.B-1. In

FIG

URE I. 0

-1

PRO

GRA

MM

ATIC

SCH

EDUL

E

AS

SE

SS

ME

NT

TE

CH

NO

LO

GY

PH

AS

E

DE

MO

NS

TR

AT

ION

PH

AS

E

.U

OP

ER

AT

ION

AL

DE

MO

NS

TR

AT

ION

PH

AS

E

KE

Y D

EC

ISIO

N P

OIN

TS

WIT

HIN

18

MO

NT

HS

: P

RE

LIM

INA

RY

BA

CK

SC

ATT

ER

CR

ITE

RIA

DE

FIN

ITIO

N

CH

IRP

RE

QU

IRE

ME

NT

S D

EM

ON

ST

RA

TIO

N

LAS

ER

CO

NF

IGU

RA

TIO

N S

EL

EC

TIO

N

PR

EL

IMIN

AR

Y S

YS

TE

M L

IFE

TIM

E F

EA

SIB

ILIT

Y D

ET

ER

MIN

AT

ION

WIT

HIN

3-5

Y

EA

RS

: B

AC

KS

CA

TTE

R C

RIT

ER

IA F

INA

LIZ

AT

ION

S

YS

TEM

LIF

ET

IME

DE

MO

NS

TR

AT

ION

W

IND

ME

AS

UR

EM

EN

T T

EC

HN

IQU

E D

EM

ON

ST

RA

TIO

N

particular, a phased hardware development program should be initiated after technology studies yield adequate definition of system requirements.

System studies should continue at an appropriate level to provide guidance to the technology studies and to evolve system requirements. The needs for ground truthfng and observations of localized meterological phenomena should be defined as early as possible, as these needs may become 3 .sign drivers.

1.C MISSION OBJECTIVES: WIND MEASUREMENT NEEDS

The ultimate objective is that of improving the accuracy of weather forecasting, particularly numerical forecasting. A spaceborne doppler lidar wind measuring system will contribute to this end by providing direct measurement of w i ~ d profiles. Two mission objectives have been identified through discussions with NOAA, USAF, and NASA personnel.

Objective I : Provide frequent, accurate, and extensive measurements

Objective 11: Provide intensive local observations

The second objective is required to some extent for ground truthing; apart from this, a desirable aim is the observation of localized meteorological phenomena.

Table I.C.l provides seven goals which may be used to quantify the mission objectives.

TABLE I . C . l MISSION OBJECTIVES/GOALS

GOAL - OBJECTIVE I OBJECTIVE 11 QUANTIFICATION QUANTIFICATION

1. OBSERVATION INTERVAL 3-12 HRS ON DEMAND-?

2. AREAL COVERAGE TROPICS-GLOBAL -GLOBAL

3. VERTICAL RANGE 10-20 KM UNSPECIFIED

4 . RESOLUTION, VERTICAL 1 KM UNSPECIFIED

5. RESOLUTION, HORIZONTAL 100-500 KM 1 0 KM

6. ACCURACY : SPEED 1-2 M/S UNSPECIFIED

7. ACCURACY : DIRECTION - + 1 0 DEG UNSPECIFIED

The f i r s t g o a l relates t o t h e f r e q u e n c y w i t h w h i c h t h e wind f i e l d s a re u p d a t e d , a n d t h i s , i n t u r n , d e t e r - m i n e s t h e number o f s p a c e c r a f t i n t h e o p e r a t i o n a l s y s t e m . To v i s u a l i z e t h i s , i m a g i n e a s i n g l e Lidar s p a c e c r a f t i n a p o l a r o r b i t , b e i n g o v e r a p o i n t X o n t h e e q u a t o r a t l o c a l dawn. Twelve h o u r s l a te r t h i s p o i n t w i l l be b e n e a t h t h e s p a c e c r a f t i f i t h a s c o m p l e t e d a h a l f - i n t e g r a l number o f o r b i t a l r e v o l u - t i o n s - a n d t h e w i n d s a t p o i n t X w i l l b e o b s e r v e d a g a i n . F o r t h i s s i t u a t i o n , t h e f i r s t p h y s i c a l l y r e a l i z a b l e s o l u t i o n o c c u r s a t 7 . 5 r e v o l u t i o n s per h a l f - d a y , c o r r e s p o n d i n g t o a n o r b i t a l a l t i t u d e o f 570 K M . A lesser number o f h a l f - i n t e g e r . - e v o l u t i o r s p z r h a l f - d a y would r e q u i r e a s p a c e c r a f t a l t i t u d e ex,-eeding 1200 K M , w h e r e STS p e r f o r m a n c e l i m i t a t i o n s , Van A l l e ~ b e l t r a d i a t i o n a n d t h e s t r e n g t h o f t h e r e t u r n i n g L i d a r p u l s e become e x p e n s i v e d e s i g n d r i v e r s . I f a s e c o n d s p a c e - c r a f t i s now p l a c e d i n a p o l a r o r b i t a t t h e same a l t i t u d e , o r i e n t e d a n d p h a s e d t o c a u s e t h e s p a c e c r a f t t o p a s s o v e r t h e p o i n t X a t local noon , t h e n t h e s y s t e m o f t w o s p a c e c r a f t c a n u p d a t e t h e w i n d s a t a p o i n t X e v e r y s i x h o u r s . S i m i l a r l y , t h r e e s p a c e c r a f t a re r e q u i r e d t o u p d a t e a t f o u r h o u r i n t e r - v a l s , s i x are r e q u i r e d f o r t h r e e h o u r u p d a t e s , a n d so o n .

The s e c o n d g o a l , c o v e r a g e , d e s c r i b e s t h e p o r t i o n of t h e E a r t h ' s s u r f a c e o v e r wh ich w i n d s are t o be m e a s u r e d . To e n s u r e t h a t w i n d s a r e m e a s u r e d a t a l l p o i n t s a r o u n d t h e e q u a t o r - a c c o r d i n g t o t h e o b s e r v a t i o n i n t e r v a l c h o s e n , e a c n s p a c e c r a f t m u s t m e a s u r e w i n d s a l o n g a s w a t h e x t e n d i n g some d i s t a n c e t o e i t h e r s i d e o f i t s g r o u n d t r a c k . To d e d u c e t h e t o t a l s w a t h w i d t h , c o n s i d e r a g a i n t h e case o f a s i n g l e s p a c e - c r a f t i n a p o l a r o r b i t , a n d n o t l e t p o i n t X be a t t h e w e s t e r n e d g e o f t h e measu remen t s w a t h . I t is e a s y t o see t h a t measu remen t o f w i n d s a l o n g t h e e q u a t o r w i l l b e complete i f t h e p o i n t X is o n t h e e a s t e r n e d g e o f t h e s w a t h o n e o r b i t a l

p e r i o d l a t e r . Us ing t h e o r b i t a l p e r i o d a l r e a d y c a l c u l a t e d , t h e total . s w a t h w i d t h is r e a d i l y shown t o b - a b o u t 2700 KM. Note t h a t t h i s v a l u e r e s u l t s f rom a s s u m i n g t h a t t h e r e are n o measurement g a p s a l o n g t h e e q u a t o r , a n d so is i n d e p e n d e n t o f t h e o b s e r v a t i o n i n t e r v a l . F u r t h e r n o t e t h a t , a l t h o u g h w i n d s a l o n g t h e e q u a t o r are m e a s u r e d e x a c t l y o n c e d u r i n g t h e o b s e r - v a t i o n i n t e r v a l , w i n d s i n t h e temperate a n d a r t i c r e g i o n s are m e a s u r e d more f r e q u e n t l y d u e t o t h e o v e r l a p p i n g measu remen t s w a t h s i n t h e s e z o n e s . I n t h i s c o n n e c t i o n , t h e r e is a cir- c u m s t a n c e wh ich may be e x p e c t e d t o f i g u r e i n a n y f u t u r e d o p p l e r l i d a r wind measu remen t s y s t e m t r a d e s t u d i e s : t r o p i - cal w i n d s c a n n o t b e d e d u c e d r e l i a b l y f rom t e m p e r a t u r e a n d p r e s s u r e measu remen t s owing t o t h e s m '-1 g r a d i e n t s o f t h e s e p a r a m e t e r s i n t h e t r o p i c s . E l s e w h e r e , d i r e c t w i n d m e a s u r e - m e n t s c a n improve o n p r e s e n t m e t h o d s . I n t h e p o l a r z o n e s , a b o v e a f e w k i l o m e t e r s , t h e s i g n a l - t o - n o i s e r a t i o (SNR) o f t h e r e t u r n i n g l i d a r p u l s e i s less t h a n a t o t h e r l a t i t u d e s , w h i c h t e n d s t o r e d u c e t h e a t t a i n a b l e a c c u r a c y . However, t h e o v e r l a p p i n g s w a t h s j u s t described p r o v i d e more m e a s u r e m e n t s , a n d t h i s s l i g h t l y o f f s e t s t h e a c c u r a c y r e d a ~ c t i o n .

The t h i r d a n d f o u r t h g o a l s , v e r t i c a l r a n g e a n d r e s o l u t i o n , re la te t o t h e a l t i t u d e s t o w h i c h , a n d v e r t i c a l s p a c i r l g s a t w h i c h , w i n d s c a n b e m e a s u r e d . A 1 KM ve r t i c a l r e s o l u t i o n means t h a t t h e a v e r a g e h o r i z o n t a l wind i a t o b e e s t i m a t e d i n e a c h 1 KM "sLabl', up t o t h e v e r t i c a l r a n g e . The v e r t i c a l r e s o l u t i o n i s l i m i t e d by t h e l i d a r p u l s e l e n g t h , t h e " least r e a s o n a b l y c o n c e i v e a b l e v a l u e " b e i n g a b o u t 400 KM ( t h i s d o e s n o t s u g g e s t t h a t so f i n e a v e r t i c a l r e s o l u t i o r i is e c o n o m i c a l l y a t t a i n a b l e , or e v e n m e t e o r o l o g i c a l l y u s e f u l ) . The s t r e n g t h o f t h e r e t u r n i p 3 l i d a r p u l s e d e p e n d s o n t h e b a c k s c a t t e r f rom a t m o s p h e r i c a e r o s o l s - w h o s e c o n c e n t r a t i o n must. u l t i m a t e l y d e c r e a s e w i t h a l t i t u d e a n d v a r y w i t h g e o - g r a p h i c l o c a t i o n , s e a s o n , t i m e o f d a y , and o c c u r r e n c e o f a e r o s o l - p r o d u c i n g e v e n t s ( e . g . , d u s t storms a n d v o l c a n i c a c t i v i t y ) . T h i s i m p l i e s t ! i a t t h e maximum a l t i t u d e fron! w h i c h u s e f u l r e t t r n s w i l l be o b t a i n e d w i l l v a r y w i t h a l l t h e s e f a c t o r s . Thus , t h e lower f i g u r e o f 1 0 KM g i v e n f o r t h e v e r t i c a l r a n g e s h o u l d b e t a k e n a s i n d i c a t i v e o f a n e e d f o r measu remen t o f w i n d s u p t o t h e t r o p o p a u s e (whose h e i g h t v a r i e s w i t h s e a s o n a n d l a t i t u d e ) , w h i l e t h e u p p e r f i g u r e s u g g e s t s t h e u t i l i t y o f s t r a t o s p h e r i c wind m e a s u r e m e n t s .

The fifth goal, horizu~tal resolution, derives from two circumstances. First, present numerical weathsr models operate with qeographic grids in which the mean wind over the grid (in each vertical resolution element) is used. Second, only the radial component of wind along the lidar pointing direction can be measured from each pulse (again, in each vertical resolution element). Thus, it is necessary to direct several lidar pulses at a given qeographical area, from different directions, to estimate the mean wind by vector resolution. The area, or grid, is usually taken t,o be square, and the horizontal resolution is the length of the grid side. The lower figure of 100 KM would be representative of small area (fine mesh) modeling. The upper figure of 500 KM would be associated with hemispheric or global model- inq. This discussion implies that the flow of data from a doppler lidar wind-measuring spacecraft is not well matched to conventional large-area numerical forecasting models whose world- view requires regular grids and synoptic measurements. This mis-match poses no insurmountable mathematical or physi- cal problem, but does indicate the need for model "velop- ments to effzctively exploit the wind data. The relationship of the data produced and the using models should be factored into future systems engineering studies.

The sixth and seventh goals, accuracies, de- scribe the accuracy with which mean grid winds are to be measured. The values given are slightly more stringent than the accuracies attained when winds are calculated from temperature and pressure measurements.

As noted at the start of this Section', a space- born2 doppler lidar is expected to improve numerical weather forecasting by providing direct wind measurements on a global scale. It seems likely that the flow of data from such a system, with some additional ground and/or onboard process- ing, would permit inference of other atmospheric dynamic parameters (e.g., cloud cover characteristics). The information so gained would not only enhance the contribution to forecasting, but also serve to advance understanding of atmospheric processes. This assessment has focused on the practicability of, and requirements for, accommodating a lidar system on a free-f lyinq operational spacecraft, with some limited consideration of other accommodations posoibili- ties. Not addressed was the possibility of accommodating additional meteorological sensors which, operated in conjunc- tion with a doppler lidar, would provide an even greater contribution to the ultimate objective and, possibly, understanding of atmospheric processes. Both possibilities-fuller exploitation of doppler lidar data and complementary sensors - should be considered in future studies.

Certain simplifying assumptions were made above to illustrate the concepts of fleet sizing and swath widths. The precise assumptions (e.g., polar orbits) and parameters based thereon (e.g., 570 KM orbital altitude) are not necessarily required in practice. The conclusions about the required fleet size remain valid.

I.D. MISSION TECHNIQUE: PRINCIPLES OF WIND MEASUREMENT - The word "lidarn was originally an acronym for

"Light Detection and Ranging" - i.e., a lidar is a radar operating at optical frequencies. In the system under consideration, a pulse of light of a few microseconds duration, at an essentially constant frequency, is directed at a geographically-fixed volume of the atmosphere. Particles (aerosols) in the volume reflect (backscatter) a portion of the pulse energy back along the direction of propagation of the pulse. This energy is collected at the system and its frequency compared to the frequency of the emitted pulse by optical heterodyning. The relative velocity between the system and the aerosols (or, more precisely, the relative velocity component along the direction of propagation) is defined by the frequency difference, caused by the doppler effect.

The motion of the Earth in inertial space is calculable, and the motion of the spaceborne lidar system can be determined by use of the global positioning system ( G P S ) . The attitude of the spacecraft will be maintained by the onboard inertial reference subsystem. With this ixformation and knowledge of the beam pointing direction, the doppler frequency shift due to spacecraf t-Earth motion can be factored out of the total shift, with a remainder due only to the component of aerosol motion along the beam pointing direction. If the aerosols are assumed to move with the wind, then the component of wind motion along the beam pointing direction has been determined. If the wind field ic the atrnospberic volume is uniform, then directing several pulses into the volume from different points on the orbit permits determination of the wind components by vector resoluti~n along the different beam pointing directions. Although the assumption of wind field uniformit;? over the atmospheric volume does not always hold, particularly when the volume corresponds to the larger horizontal resoluticn elements, the nonuniformity does not compromise the concept, as only the mean wind in the volume is sought. In the usual case, it may also be assumed that vertical winds are negligible: the ex- ceptions are generally localized and/or transient phenomena.

Figure I.D.1, adapted from the NOAA/LMSC WINDSAT final briefing package, illustrates a doppler lidar wind- measuring system concept. The primary "TEA" (Transversely Excited, Atmospheric)laser output pulses are reflected by an optomechanical switch in the "transmit" positon through a beam alignment subsystem consisting of a fold mirror, a two- axis gimbaled mirror, and a LAC (Lag Angle Compensating) mirror. The pulse is then reflected from the tertiary mirror into the cassegrainian telescope, which expands the beam and directs it at the atmosphere. The pointing direction in this concept is controlled by rotating the telescope about the local vertical at an offset scan angle. In principle, this rotation can be either ctntinuous or stepped.

The returning pulse of backscattered light is collected by the cassegrainian telescope and reflected by the tertiary mirror through the beam alignment subsystem, which directs it through the "receive" position of the switch. The beam alignment subsystem corrects the misalignment due to Bradley abberation caused by the spacecraft motion and, if the rotation is continuous, for the rotation of the telescope (lag angle) during the pulse roundtrip time (typically between 5 and 10 milliseconds). After passing through the switch, the returning light is mixed with a frequency- offset beam from a local oscillator laser by a combining beamsplitter. The offset permits the sense of the doppler shift (red or blue) to be determined. The frequencies of both the local oscillator and the TEA laser are controlled by the technique of injection locking.

The mixed light from the combining beam-splitter forms an interference pattern on the photo-detector. This pattern varies with the doppler-shifted frequency of the returning light. Thus, the detector output is an FM-modulated electrical sign31 whose modulation frequency corresponds to the doppler shift. The predictable shift due to the spacecraft-Earth moti-n will vary by about +1.7 GHz, depending on whether the beam pointing directron is along or against the spacecraft velocity vector. The function of the frequency synthesizer in the removal of this "grossn doppler shift. The result.ing signal is then digitized and range- gated (corresponding to the vertical resolution elements) and, after further preprocessing to extract signal characteristics, recorded for later downlinking.

FIGU

RE I. D

-1

LIDA

R SY

STEM

CON

CEPT

(L

ASER

& O

PTIC

AL S

UBSY

STEM

S)

11. SYSTEM PERFORMANCE ANALYSIS

I1 .A. SIGNAL-TO-NOISE RATIO

As noted in Section I.C., a primary system performance goal is wind speed measurement accuracy. It is intui- tively reasonable that this will improve with increasing values of the signal-to-noise ratio (SNR) of the return pulse. Further, it is reasonable to expect an expression for the SNR of the return from a volume of atmosphere to contain terms involving 1) the output pulse characteristics (energy, duration, and wavelength), 2 ) atmospheric effects (backscatter, attenuation, and turbulence) and 3 ) geometry of the collecting optics.

The wind measurement accuracy analyses of WPL-63 are based upon the expression

J7T - 2 + R KD*

SNR = - 315. pe 8%

where

h = Planck's ccnstant

J = pulse energy

= doppler-shifted wavelength

Z = pulse duration

f l = atmospheric backscatter coefficient

,& = atmospheric attenuation coefficient

R = range from spacecraft to atmospheric volume

7 = overall detector-optics efficiency

D = telescope diameter

5 = turbulence-induced transverse coherence radius

= 0.069 2615 (~~:)-3/5

C: = refractive index structure parameter

f = focal length

In fact, the SNR should be written as a function of the altitude from which the return occurs: letting

a = altitude of the sampled volume

A a = vertical resolution

d = scan angle

4 , spacecraft altitude

= range from spacecraft to altitude at scan angle

The expression fie-4UR is more accurately written as ~ R ( R N C . & ; a)

da 11 . B ~ ~ , a ) O l r {-z J,'s> ( l , l ~ d ~ ( t i , ~ ,Q+#

where the form of the integral indicates that the attenuation is evaluated along the line connecting the spacecraft and the volume. A similar modification for the turbulence radius is implied (cf. WPL-37, P.203!.

For the present purpose of assessing the accommodations required for a wind measuring doppler lidar system, a simpler form of the SNR expression will suffice, and the following expression has been used: k*

- , I s ~ r l [ g ~ a . ~ ~ { - z J , 4 4 ( ~ ) ~ ~ ( ~ 5 ~ . ~ ; 4 ] ! * ? [ .aU,M;o S N R ( 2 ) - a

Figure 1I.A-1 displays the result of one calculation using this expression. The assumptions relative to atmospheric attenuation and backscatter are discussed in the next section.

I I. B. ATMOSPHERIC UNCERTAINTIES

Ultimately, the feasibility of accommodating the proposed wind-measuring system will depend on atmospheric properties, particularly attenuation and backscatter. The total attenuation can be expected to be only slightly greater than the molecular absorption, particularly above about 7 KM. Figure 1I.B-1 plots the molecular absorption coefficient at the 9.1145,~ wavelength for the AFGL model atmospheres. Al- though the absorption at a specific altitude and place will vary with time-of-day, season, and weather, the vari- ability is reasonably well bounded. The form of the attenuation term in the SNR expression implies that long light paths through the atmosphere should be avoided.

The uncertainty associated with the backscatter is mucb less well understood. Figure 1I.B-2 is a compendium of backscatter estimates from various wavelengths other than the ones of interest, with considerable reliance on the aerosol size distribution models. In this report, calculations have been based on the assumed "worst case" curve. The presence of Lidar returns does indicate the basic feasibility of the doppler lidar wind measuring technique.

FIG

URE IZ. A

-1

S IG

NAL-

TO-N

O1

SE R

ATIO

VS.

AL

TITU

DE

OR

BIT

AL

AL

TIT

UD

E =

650

(K

MI

OR

BIT

AL

INC

LIN

AT

ION

= 5

7 (D

EG

) O

PTI

CS

DIA

ME

TE

R =

1.2

5 (M

I W

AV

EL

EN

GT

H =

9.1

145 (M

ICR

O M

)

AT

TE

NU

AT

ION

= A

FG

L T

RO

PIC

AL

AB

SO

RP

TIO

N

BA

CK

SC

AT

TE

R=

MS

FC

WO

RS

TC

AS

E

SA

MP

LED

AL

TIT

UD

E, k

m

Clearly, an operational system will not be able to measure winds at all altitudes, at all times: clouds are one inhibiting factor, and there will be occasions when the concentrations of aerosols in some regions and at some altitudes will be insufficient to provide a useful return signal.

FIG

URE II.

B-1

MO

LECU

LAR

ABSO

RPTI

ON

COEF

FIC

IENT

(km

-1)

ATX-

9.1

145

p m

- TRO

P

- ....... *

..- M

LS

--

----

M LW

-- S A

S

---

SA

W

10al

0 1

2 I 4 I

6 I 8 I

10

I 12

I

14

I 16

I

18

1 20

I

AL

TIT

UD

E (km)

111. CONFIGURATION AND MASS PROPERTIES

111. CONFIGURATION AND MASS PROPERTIES

1II.A. REQUIREMENTS AND ASSUMPTIONS

The following assumptions and requirements determined the configuration design approach

(i) Beam pointing shall be realized by a scanning Cassegrainian telescope with a maximum diameter of 1.25 m. The scanning axis shall be aligned with the local vertical, and the optical axis of the telescope shall bt cffset by approximately 56 deg (the scan angle). The scan rate shali be approximately 4 rpm.

(ii) The Shuttle shall be tht: ;lunch vehicle, and the spacecraft . ; 1 be retrievable

(iii) There shall be two solar array wings of approximately 28m2 individual area, capable of both deployment and retraction.

(iv) There shall be two deployable/ retractable radiators (both sides active) with total area of approximately 5.7m2

The mission equipment (laser, scan mechanism, and telescope) was assumed to be similar to the continuously scanning, one- meter, Cassegrainian configuration reported in the NOAA/LhSC WINDSAT Final Briefing Charts: a fritted glass mirror was assumed rather than the Beryllium mirror recommended by LMSC.

1II.B. CONCEPTUAL DESIGN

Figure 1II.A-1 illustrates the features of the conceptual configuration, as it would be stowed in the orbiter cargo bay. The cylindrical spacecraft bas the maximum diameter compatible with the orbiter cargo ba:. dynamic envelope and a height of 1.5 m. The resulting internal vo1:une 'easily accommodates the mission and subsystems equipment, as determined by a trial layout. This "pancake" configuration is the result of a natural tendency to effectively utilize the cargo bay volume: it does -- not, in this instance, result from the present commercial payload STS charge policy. The driving cost factor under this policy according to performance calculations, would be the weight delivered to orbit

FIGU

RE I

n. A

-1

- DO

PPLE

R LI

DAR

WIN

D M

EASU

RING

SPA

CECR

AFT

12

5 M S

CA

NN

ING

LA

UNCH

CON

F1GU

RATI

ON

CA

SS

EG

RA

INIA

N T

ELE

SC

OP

E -

RA

DIA

TO

R

L.O

. A

. -4

.46

M (1

4.6

FT

) D

:A =

4.4

7 M

(176

IN)

MA

SS

= 2

260

KG

S.

A P

OW

ER (

BO

L) - 8.5

KW

DE

PLO

YM

EN

T W

NG

E

11 I

RA

DIA

TO

R

y.b

K

EE

L

FIT

TIN

G

SlL

L

FIT

TIN

G

(1 O

F 4

)

,

AN

TEN

NA

-.- --

-

PD

IZ3

ST

RID

ER

r a t h e r t h a n p a y l o a d l e n g t h . The a n t e n n a f e e d h o r n is s t o w e d i n t h e body o f t h e s p a c e c r a f t when t h e a n t e n n a is retracted: a f o l d a b l e h o r n c o u l d a l s o be u s e d t o m i n i m i z e t h e s t o w e d p a y l o a d l e n g t h .

The k e e l arid s i l l f i t t i n g s a t t a c h t o t h e p a y l o a d p r i m a r y s t r u c t u r e , o b v i a t i n g t h e n e e d f o r a s e p a r a t e l a u n c h c r a d l e . T h i s i n t e g r a t i o n of t h e l a u n c h c r a d l e LnLo t h e s p a c e c r a f t is n o t o n l y p r o g r a m m a t i c a l l y s i m p l e r , it a l so m i n i m i z e s t h e t o t a l p a y l o a d - c h a r g a b l e cargc mass. I n d o i n g so, it p e r m i t s c o n s i d e r a t i o n o f d i r e c t d e l i v e r y t o o r b i t ( a t 57 d e g ) w i t h o u t u s e o f a n u p p e r s t a g e or a n i n t e g r a l p r o p u l - s i o n s y s t e m f o r o r b i t a l t r a n s f e r ( a t l e a s t o n e , a n d l i k e l y two, orbi ter OMS k i t s would b e r e q u i r e d o t h e r w i s e ) . T h i s i s n o t t o s u g g e s t t h a t t h i s w i l l u l t i m a t e l y b e t h e p r e f e r r e d a p p r o a c h , o n l y t h a t t h i s is o n e o p t i o n . The s i l l f i t t i n q s are s p a c e d 59 i n c h e s a p a r t (as o n a s t a n d a r d ESA p a l l e t ) t o p r o v i d e t h e maximum number o f p o s s i b l e a t t a c h p o s i t i o n s i n t h e c a r g o bay . A g r a p p l e f i x t u r e ( n o t shown) is p r o v i d e d t o a l l o w d e p l o y m e n t / r e t r i e v a l u s i n g a remote m a n i p u l a t o r arm.

The t e l e s c o p e , r a d i a t o r s , a n d f o l d e d solar a r r a y s are mounted o n o n e f l a t s i d e o f t h e " p a n c a k e " s p a c e - c ra f t a n d s e c u r e d by l a u n c h l o c k s . The r a d i a t o r s ro ta te 1 8 0 d e g d u r i n g d e p l o y m e n t i n t o t h e p o s i t i o n i l l u s t r a t e d i n t h e a r t i s t s ' c o n c e p t i n F i g u r e 1 . A - 1 . S i m i l a r l y , t h e so la r a r r a y w i n g s a re mounted o n s h o r t a-ms a n d ro ta te 1 6 0 d e g a b o u t t h e i l l u s t r a t e d h i n g e p o i n t s before u n f o l d i n g t o t h e i r o p e r a - t i o n a l p o s i t i o n . The a n t e n n a is mounted on t h e r e m a i n i n g f l a t s i d e o f t h e s p a c e c r a f t .

1 I I . C . s Y S T P l S I Z I N G

S u b s y s t e m a n d s p a c e c r a f t mass estimates are g i v e n i n t h e f o l l o w i n g T a b l e : t h e s e masses a re b a s e d o n t h e m i s s i o n e q u i p m e n t masses t a k e n from t h e NOAA/LMSC WINDSAT S t u d y F i n a l B r i e f i n g C h a r t s , s u b s y s t e m mass estimates, a n d t h e a s s u m p t i o n o f a n a l l - aluminum s p a c e c r a f t s t r u c t u r e .

TABLE 1 I I . C - 1

DOPPLER L I D A R WIND MEASURING SYSTEM

PAY LOAD MASS SUMMARY

L I D A R M I S S I O N EQUIPMENT

STRUCTURE

A T T I T U D E CONTROL AND DETERMINATION

COMMUNICATIONS AND DATA MANAGEMENT ELECTRICAL POWER

THERMAL CONTROL

1V.A. ET-ECTRICAI, POWER SUBSYSTEM

IV. A . 1 REQUIREMENTS AND ASSU-WTIONS

Conceptual design and analysis of the electrical power sp~hsys~em requires definition of the loads (power and voltage), orbit (altitude and inclination), and mission fac- tors such as mission duration and spacecraft pointing.

For the doppler lidar wind measuring system spacecraft, the loads were estimated as:

Guidance and Navigtion 227 -- 247 W

Communications and Data Manrgement 110 - 115 W

Thermal Control 360 W

Mission Equipment 2140 - 2346 W Totals 2827 - 3068 W

The subsystem estimates are based on representative equipment selections by subsystem engineers. The thermal control power requirement is highly conservative, being based on the startup power for a gas pump. The mission equipment power estimates were taken from the September 198G NOAA/LMSC WINDSAT Final Briefing Charts.

A design margin of 15-25% is customary in preliminary design activity to allow for uncertainties which almost invariably lead to growth in power requirements as th= design matures. Using 3500 W as the total electrical power requirement yields a 24% and a 14% margin relative to the low and hig3 estimates, respectively, of the last paragraph. On balance, this estimate is slightly more conservative than usual. The EPS design is not sensitive to small departures trom this value, whose adopt.ion obviates the unenl ightening production of many tables and equipment lists tailored to minor differences i ~ , the ,-.ases considere ;.

The choice of orbit has two primary effects upon EPS design: first, the solar array and battery system must be sized to supply exlipse load?; secund, the altitude, inclination, and launch date (within the solar cycle) affect the particulate radiation, which degrades the solar array. A lower orbit gives a lower light/ dark ratio (requiring a larger array and battery system) whereas a higher orbit yields greater radiation damage due to geomagnetically trapped particles. For this study, a three--year mission coinciding

w ~ t h a peak period of solbr activity was assumed, so that flare protons were a majar contributor to the array degradation.

Five orbits were analyzed to scope the characteristics of a conceptual EPS design:

ORBIT ALTITUDE INCLINATION NCDE # (km; (deg) - (hr)

1 800 sun-synch 0600



The fSrst and third of these represent the best- and worst-cases, respectively, for a high-inclination, 800 kin orbit (as per the NOAA WPL-37 and -63 reports). The 0600 node maximizes the day/night ratio (yielding the "minimum" EPS) while the 1200 node has the opposite effect. The second oroit was analyzed to assess the sensitivity of the EPS to orbital altitude at the most favorable nodal position. The fourth a-d fifth orbits were analyzed to bound the EPS char- acterist-cs in orbits potentially accessible from KSC: nodal ~osition is not an EPS driver at the 57 deg inclination.

1V.A. 2. CONCEPTUAL DESIGN

The relatively large power requirement of 3500 W and the local vertical orientation of the telescope scan axis necessitates an oriented (sun-tracking) solar array. Two rotational degrees of freedom are used to maximize the array output: roll about t!le scan axis and rotation of the solar arrays. The first of these is ccmpatible with the need to minimize the sunlight incident on the radiators. Rotation of the solar axrays implies use of solar array drive mechanisms and power transfer devices (such as slip rings or flex cables). These solar array components have been developed for many programs, and there is ample experience for the present application.

A variety of power distribution and control schemes are possible: the concept described here is not necessarily optimum, but is adequate to provide estimates of the realizable efficiencies, weights, and dimensions.

F i g u r e 1 V . A - 1 i l l u s t r a t e s t h e c o n c e p t . S o l a r a r r a y power is t r a n s f e r r e d v i a s l i p r i n g s or f l e x c a b l e s t o t h e ma in power b u s a t a n o m i n a l 3 4 Vdc ( v a r y i n g w i t h a r r a y t e m p e r a t u r e ) . The power c o n t r o l u n i t (PCU) p r o v i d e s command, c o n t r o l , a n d p r o t e c t i o n f u n c t i o n s . D u r i n g t h e d a y l i g h t p o r - t i o n o f t h e o r b i t , t h e PCU r o u t e s solar a r r a y power t o t h e b a t t e r y c h a r g e c o n t r o l l e r s ( C C s wh ich p r o v i d e a c h a r g e / d i s c h a r g e p a t h f o r t h e N i C d b a t t e r v a s s e m b l i e s . The PCU a l so r o u t e s power f r o m t h e s o l a r a r r a y (or b a t t e r y , a t n i g h t ) t o t h e m u l t i p l e l o a d b u s e s w h i c h , i n t u r n , d i s t r i b u t e power t o t h e v a r i o u s u s e r s . The l o a d b u s e s i n t e r f a c e l o a d s t h r o u g h r e m o t e l y l o c a t e d d i s t r i b u t o r s w h i c h p r o v i d e l o c a l s w i t c h i n g , i s o l a t i o n , a n d p r o t e c t i v e f u n c t i o n s . No te t h a t t h i s d i s t r i - b u t i o n scheme d i s t r i b u t e s u n r e g u l a t e d power a t 22-34 d c ( d e p e n d i n g o n a r r a y t e m p e r a t u r e a n d b a t t e r y s t a t e o f c h a r g e ) . The v a r i o u s l o a d s mus t t h e n p r o v i d e t h e i r own power s u p p l i e s ( d c - d c ) , wh ich is n c t t r u e f o r a d i s t r i b u t i u t ~ scheme ta i lored t o t h e case w h e r e a l l l o a d s r e q u i r e a s p e c i f i c , r e g u l a t e d ( u s u a l l y 28 Vdc) s o u r c e . I n t h e p r e s e r t a p p l i c a t i o n , t h e l i d a r is t h e m a j o r power u s e r , r e q u i r i n g h i g h power , h i g h v o l t a g e s u p p l i e s . The r e g u l a t 2 d b u s a p p r o a c h wou ld h a v e t h e d i s a d v a n t a g e o f r e g u l a t i n g t h e power t w i c e , y i e l d i n g a l o w e r o v e r a l l c o n v e r s i o n e f f i c i e n c y .

F i g u r e 1 V . A - 1 i n d i c a t e s o p t i o n a l c h a r g e c o n t r o l - lers a n d b a t t e r i e s . T h i s r e f l e c t s t h e d i f f e r e n t e n e r g y s tor- a g e r e q u i r e m e n t s of t h e c a n d i d a t e o r b i t s . An o ~ e r a t i o n a l s y s t e m w i t h s e v e r a l s p a c e c r a f t a t d i f f e r e n t n o d e s c o u l t , h a v e a cominon power s y s t e m w i t h d i f f e r e n t e n e r q s t o r a g e c a p a c i t i e s .

F i g w e IV.13-2 i l l u s t r a t e s t h e s o l a r a r r a y con- c e p t , h a v i n g two i d e n t i c a l w i n g s e q u i p p e d w i t h d e p l o y / r e t r a c t mechanisms a n d power t r a n s f e r / d r i v e mechanisms f o r s u n o r i e n - t a t i o n a n d power t r a n s f e r . Each wing c o n s i s t s o f a number o f subn iodu le s f o r ease o f m a n u f a c t u r e , a s s e m b l y , a n d l a u n c h . The s u b m o d u l e s a re a s s e m b l i e s o f solar ce l l s , v a r i o u s t e r m i - n a l s , d i o d e s , a n d h a r n e s s w i r i n g f e a t u r e s . Some s u b m o d u l e s also i n c o r p o r a t e t h e r m a l a n d c u r r e n t s e n s o r s f o r a r r a y d i a g - n o s t i c s . Sun s e n s o r s o n e a c h wing p r o v i d e t h e s o l a r c r i e n t a - t i o n s i g n a l s n e e d e d by t h e so lar a r r a y d r i v e e l e c t r o n i c a s s e n l b l y mounted w i t h i n t h e s p a c e c r a f t .

Power a n d s i g n a l s a re r o u t e d t o t h e s p a c e c r a f t by a "power s y s t e m h a r n e s s " .

IV.A.3. SYSTEM SIZING

The f i r s t s t e p i n s y s t e m s i z i n g is t h e d e t e r m i - n a t i o n o f t h e e n d - o f - l i f e s o l a r a r r a y power r e q u i r e m e n t a n d

FIGU

RE IP. .4

-1

SOLA

R

AR

RA

Y

WIN

G

Ll DA

R EL

ECTR

ICAL

POW

ER S

YSTE

M

SYST

EM B

LOCK

DIA

GR

AM

r T

SOLA

R

AR

RA

Y

-

BA

TTE

RY

1

A8 T

WIN

G

CH

AR

GE

I

L

J b

1

BA

T 2

CO

NTR

OL

1

7

- r-'

SIG

NA

L

DR

IVE

POW

ER

W

w

CO

MM

AN

D

DR

IVE

-

b C

Cl CC

3

U

0

>

3

M

I L I

-

I

CC

2

I r-

----

-7

I J

I cc

4 !-

-J

\

I

.------*

I I

r--------I

! B

AT

4 1

r---

---7

--d

I C

C5

LII!

WL

- - - - - - --

I I

J L-

----

- J

(OPT

ION

AL

CC

) I

r------- 1

\ I

I t

ELEC

T. ,

STA

TUS

4

I

OlS

T RIB

UTO

RS

SIG

NA

LS, C

OM

MA

ND

S LO

AD

S

s

m 5 I

-

- O

RlV

k f

1

POW

ER

S-

-

c

COH

LOA

DS

V)

W

U) 3

10

o

a

0 4

POW

ER

CO

NTR

OL

UN

IT

4

BA

T b

I

L ,,------

J

IOP

IIO

NA

L B

ATT

ERIE

S)

OIS

TRIB

UTO

RS

EX

P

LOA

OS

FIG

URE 19.

A-2

WlN

G 2

(SIM

ILA

R T

O W

lNG

1)

POW

ER S

YSTE

M H

AR

NES

S

OLA

R AR

RAY

BLO

CK D

IAG

RAM

m

-

'N"

STR

ING

S IN

PA

RA

LLE

L

POW

ER -

RET

UR

NS

CO

MM

AN

D

CU

RR

ENT

SUN

A

ND

STA

TUS

SENS

OR

T EM

P

y SE

NSO

RS

y SE

NSO

RS

I I,

111 111

II 111

111

11 111

111

. 4

- O

EIL

OY

/RE

TRA

CT

MEC

HA

NIS

M

DEP

LOY/

RLT

RA

CT

MEC

HA

NIS

M

----------

----

A _ - - - - - - -

SLIP

RIN

Q/D

RIV

E

SOLA

R A

RR

AY

7

-

-

DR

IVE

SL

IP R

ING

IDR

IVE

A

SSEM

BLY

1

ASS

EMB

LY

ELEC

T RO

NlC

S

energy storage capacity requirement. Apart from relatively standard allowances for conversion and transmission efficiencies, the variation in the suntime fraction during the year must be accounted for. This fraction is shown in figures 1V.A-3 and 1V.A-4 as "percent time in sunn (PCTIS). For an 800 km, sun-synchronous, orbit with an 0600 hr node, Figure 1V.A-3 shows that the spacecraft is continuously in the sun except for a period of about 75 days centered around the summer solstice, when the minimum orbital suntime percent is 83.6%. In contrast, Figure 1V.A-4 shows that a spacecraft in an 800 km, 57 deg orbit is occulted on almost every orbital revolution, frequently by as much as 35%. Figure 1V.A-5 defines the end-of-life solar array power requirements associated with the candidate orbits.

Figure 1V.A-6 shows a cross-section of the solar array, indicating the typical components of a lightweight, rigid honeycomb core type panel. Detailed test data on high- efficiency solar cells (from JPL) were used to select the baseline solar cell and to establish the electrical and thermal operating points (The cell conversion efficiency is sensitive to the cell temperature: at high temperatures, the voltage drops rapidly and, while the current increases slightly, the next effect is decreased output power at the operating point.). Application of estimated degradations due to irradiation and thermal cycle and micrometeorite damage, together with allowances for Earth orbit eccentricity, assembly losses, interconnections, panel layout and accessory components, yields an estimate of the beginning-of-life (BOL) array output requirement. This estimate of the array size also allows estimation of the total wire length/mass connecting the array to the spacecraft. Figure 1V.A-7 defines the BOL power requirement for the various orbits considered. Figure 1V.A-8 defines the corresponding total EPS masses: the primary differences are due to

(1) additional battery capacity for increased night times, and

( 2 ) additional solar array area to charge the extra battery capacity.

1V.B COMMUNICATIONS AND DATA MANAGEMENT

IV :B. 1 DATA RATES

The maximum data rate requirements for the CDMS can be derived straightforwardly by considering the temporal and spectral characteristics of the doppler-shifted pulse backscattered from the atmosphere.

FIG

URE

IP.

A-3

SU

NTIM

E FR

ACTI

ON

(SUN

-SYN

CHRO

NOUS

ORB

IT)

AL

TIT

UD

E - 80

0 KM

INC

LIN

AT

ION

- 98.6

DE

G

NO

DE

- 060

0 H

R

TlN

lE F

RO

M V

ER

NA

L E

QU

INO

X (HR)

FIGU

RE 19. A

-5

LICA

R EL

ECTR

ICAL

SYS

TEM

- EF

FIC

IENC

Y BL

OCK

DIAG

RAM

r..-.-..--.-.--.-.-.-.-.--..--.....q I

I

'SA

I

I t

I ?

I '

I p

~'3

50

0W

I

I

SO

LAR

I

ww

tn

- 91s S

~L

R - I

7 0

A

RR

AY

I

I

A L

t

LO

AD

S

I

I I

. I

I I

1 I

! I I

I 'I

c ! 'I

D

I I

I I

I I

I I

1 I

I I

I I

I I

I

I CA

SE 1

CA

SE 2

CA

SE 3

CA

SE 4

CA

SE L

I

I I

I A

LTIT

YO

E

1 M

AX

. MIG

HT

?SA

800

/8:0

0 99

2 50

50

4591

47

93

61 27

60

95

6468

I

WA

TTS

650/

6.00

11

43

4718

800/

12:0

0 21

05

3845

8001

57~

2087

39

63

5001

57~

2145

35

29

Km

SE

C SE

C

SO

LA

R A

RR

AY

R

AT

ING

FIGU

RE a. A-7

SO

LAR

ARRA

Y S

IZIN

G

SO

LAR

AR

RA

Y D

EC

RA

OA

TIO

NS

S

OLA

R A

RR

AY

PO

WE

R

RA

DIA

TIO

N

MIC

RO

ME

TE

OR

ITE

U/V

. A

SS

EM

BLY

TH

ER

MA

L C

YC

LE

PSA

R (

BO

L A

RR

AY

PO

WE

RI'

*PSA

R I

ST

HE

BE

GIN

NIN

G O

F L

IFE

IBO

L) S

OLA

R A

RR

AY

, O

WE

R O

UTP

UT

AT

OP

ER

AT

ING

TE

MP

ER

AT

UR

E A

ND

IL

LU

MIN

AT

ION

t

OR

llT

PS

AR

CA

SE 1

800/

6:0O

601 2

CA

SE 2

660/

6.00

6200

CA

SE 4

8001

57~

1908

CA

SE 3

8001

1 210

9

8040

CA

SE 6

I

500/

57O

K

M

8488

I

WA

TTS

FIG

URE IE.

A-8

LID

AR E

PS S

UM

MAR

Y

I ALT

ITU

DE

(KM

) O

RB

IT

INC

LIN

ATI

ON

N

O D

E (H

R)

SOLA

R A

RR

AY

SUBS

YSTE

M

SOLA

R A

RR

AY

DR

IVEI

DEP

LOYM

ENT

DR

IVEI

DEP

LOYM

ENT

ELEC

T.

DIO

DE

ASSE

MBL

Y SU

B TO

TAL

POW

ER C

ON

TRO

L &

EN

ERG

Y ST

ORA

GE

POW

ER D

ISTR

IBU

TIO

N

POW

ER C

ON

TRO

L CH

ARG

E C

ON

TRO

L B

ATT

ERY

SUB

TOTA

L

POW

ER D

ISTR

IBU

TIO

N

EQU

IPM

ENT

HARN

ESS

PAYL

OA

D H

ARNE

SS

SUB

TOTA

L

TOTA

L, K

G

850

SUN

SYNC

06

00

121 68

18

14

22 1 23

45 7 98

173

125 68

193

587

800

SUN

SYNC

06

00

11 5 68

18

14

21

5 23

45 7 98

173

125 68

193

58 1

800

SUN

SYNC

12

00

155 80

18

18

27 1 23

55

11

250

339

136 68

2W

814

800

57

AN

Y

1 54 80

18

18

27

0 23

55

11

250

339

136 68

500

57

AN

Y

163 80

18

18

27

9 23

55

11

250

339

136 68

I

Figure 1V.B-l(a) illustrates the tzmporal characteristics, using representative numerical values taken from the NOAA/LMSC study final presentation charts. A laser pulse of about 5 y s duration is transmitted from the spacecraft into the atmosphere, from which backscat.tering occurs until the trailing edge of the emitted pulse reaches the surface. Thus, the duration of the return pulse is

where

AID = Atmospheric Interaction Distance

c = Speed of light

r = Emitted pulse duration

or about 350 ,us for the parameters assumed in the NOAA/LMSC study. This can be chopped illto 20 "range bins" of 17.5 ,us duration, each corresponding to a 1 km vertical resolution.

Figure 1V.B-l(b) illustrates the essentail spectral characteristics. The greatest frequency shift is due to the spacecraft motion relative to the Earth (about - + 1.5 GHz, depending on whether the laser "firing direction" is along or against the space craft motion). The doppler shift due to a + 1 m/s wind is about 220 kHz - leading to a 44 MHz bandwidth for measuring +lo0 m/s winds. The sampling rate is then 88 MHz, or about 150 MHz. Based on the NOAA and LMSC studies, a 4 bit quantization has been selected Thus, each pulse generates 140 kbit of phase information. At 8 pulses evcry second, this yields 1.12 Mbps of phase data. Quantiz- ing the amplitude data at 10 bits, for each range bin, and at a 1 MHz sample rate, yields an additional 30 kbps.

The total raw data rate is then about 1.15 Mbps. Chpability for downlinking this data is necessary, at least initially, for evaluation and special studies. Operation- ally, such a flood of dat3 is neither necessary nor desirable, and some form of onboard preprocessing must be considered.

At the extreme, this preprocessing mcy reduce the data for each range bin to three quantities indicative of the radial component of the wind, the dispersion due to shear and turbulence, and the signal strength. With an austere allowanc.: for data defining the orientation of each "shot" relative to the Earth, 5 Kbps emerges as a lower limit for the preprocessed data downlink rate.

These downlink data rates (raw and preprocessed) could be halved if the number of range bins is halved (either

FIGURE

TRANSMITTED LASER PULSE DURATION - 5 p S

EARTH SURFACE

ROUNDTRIP TIME " 7 pS

\ 1 km RANGE BIN

I-- 350 irs RETURN SIGNAL

(a) TEMPORAL CHARACTERISTICS OF LlDAR PULSE

RETURN SIGNAL TRANSMITTED LASER PULSE (1 km RANGE BIN)

200 kHz TRANSFORM LIMITED PULSE

0 - T 1.5 GHz WIND GROSS DOPPLER DOPPLER DUE TO SPACECRAFT - EARTH RELATIVE MOTION TI. * 200 DUE kHz TO + WIND Af SHEAR

(b) FREQUENCY CHARACTERISTICS OF LlDAR PULSE

by halving the vertical ranae or increasing the vertical resolution to 2 km). Conversely, if the quantization of phase data were increased from 4 to 8 bits, the downlink rates would be doubled.

1V.B. 2. TDRSS UTILIZATION

Figure 1V.B-2 illustrates the positioning of the two geostationary TDRSS spacecraft. One consequence of this positioning is the presence of a "dead zonen in which the spacecraft will be out of contact with both TDRSS spacecraft. Figure 1V.B-3 depicts the geographic location and extent of this dead zone for a spacecraft with an orbital altitude of 500 km. There are three categories of TDRSS service to be considered:

o Multiple Access (MA) - Up to 20 users/TDRS - Up to 50 kbps downlink rate

o S-band Single Access (SSA) - Twc user/TDRS - Up to 3 Mbps downlink rate

o Ku-band Single Access (KuSA) - Two user/TDRS - Up to 300 Mbps downlink rate

The multiple access service would limit the system to transmission of data ~rocessed to less than 50 kbps (less, to allow for housekeeping, attitude, and position data). Although the TDRSS operational modes have not been fixed at this time, there is reason to believe that MA chan- nels may not be available to a single user for years 9f operation. The multiple access service does not provide a useful capabilit~? for transmissi~n of unprocessed data.

The SSA service appears to be the preferred service for this application: A limited amount of unprocessed data can be downlinked in real time and processed data can be handled with a store-and-dump mode using NASA standard 108 tape recorders. The principal issue relative to the store-and-dump mode is tape recorder life, the m ~ i n limits on which are the numbers of tape znd negator spring reversals (there being two of the latter for each of the former 1 . The recorder is designed for 20,000 tape passes and 50,000 spring reversals. Tv~enty-eight transports have been flown, and have accumulated 51,000 hours of operation without a major failure. If the maximum recorder dump rate (2.5 Mbps) is used - a parallel track dump - then parallel recording must be used. Figure 1V.B-4 describes the relationship between the record rate and the recorder tape/negator spring ,life limits for two cases: A single recorder and dual recorder

FIG

URE E. 0-

2 TD

RSS

COVE

RAGE

GEO

MET

RY

TO

RS

EA

ST

T

OR

S W

ES

T L

ON

G 1

71°

W

TO

RS

EA

ST

SH

AD

OW

ZO

NE

FIG

UR

EIY.

0-4

TA

PE R

ECOR

DER

DESI

G

LIFE

N

ASA

STAN

DARD

5 X

1 $ B

ITS

60 U

NIT

S B

UIL

T

28 T

RA

NS

PO

RT

S F

LO

WN

51

,000

HO

UR

S O

F O

PE

RA

TIO

N I

N S

PA

CE

(N

O M

AJO

R F

AIL

UR

ES

)

DE

SIG

N L

IFE

FO

R T

AP

E P

AS

SE

S

RE

CO

RD

ED

DA

TE

RA

TE

IN

K b

ps

operated in a flip - flop qode. Limiting the record rate to no more than 50 kbps not only provides a 50% margin on tape passes, but permits an alternate mode usable in the event of tape recorder failure: processed data would be downlinked continuously using MA service to the extent allowed by com- peting traffic (while not in a dead zone).

The technical requirements of the doppler lidar wind-measuring system are satisfied by the SSA service: this - is also the most economical approach, based on the current "softn TDRSS use charges (the charge rates used are not official but are accepted figures for planning purposes). For the MA service, assuming continuous telemetry of data processed to less than 50 kbps, the three-year mission use charge would come to $6.57 M. In contrast, the corresponding use charge for SSA with a store-and-dump mode at a 50 kbps record rate / 2.5 Mbps dump rate is $2.63 M. The S-band communication equipment is the same in either case, and three NASA standard recorders at approximately $500 K each are required for the SSA mode (two to ensure no loss of data, and an additional recorder to allow one failure without degradation of system performance). The SSA use charge may be significantly lower if the recorders can be dumped to exploit the TDRSS "as available" rates.

No need for downlinking large quantities of unprocessed data have been identified: if such a need does arise, a store-and-dump mode utilizing the KuSA service will need to be considered. The limiting factor in this instance will be the tape recorder dump rate. The Spacelab High Data Rate Recorder (HDRR) could be used; its maximum dump rate of 32 Mbps would result in a KuSA utilization rate of about 7%. At a charge of $5,00O/hr, the resulting use charge for a three-year mission is $9.2 M. The cost of the Ku-band communications equipment and adaptation of the HDRRs would likely equal or exceed this figure.

1V.B. 3. CDMS CONCEPT

Figure 1V.B-5 illustrates the basic CDMS concept for the doppler lidar wind measuring spacecraft. Since a nominal 3-years mission is appropriate for a free-flying spacecraft, redundant hardware has been provided. The A/D converters and the preprocessor are "special purpose" items and will probably not be considered a part of the CDMS.

For routine operations, the preprocessor reduces the 1.15 Mbps stream of amplitude and phase data to a rate between 14 and 48 kbps; this preprocessed data is multiplexed with telescope and spacecraft navigation and housekeeping data to form a stream of no more than 50 kbps. This final stream is either recorded for later playback or downlinked

FIG

URE

DL B

-5

DOPP

LER LI

DAR

SPAC

ECR

An

COM

MUN

ICAT

ION

AND

DATA

MAN

AGEM

ENT

BLOC

K D

l AG

RAM

r - 1

GN

LC

,

CM

OS

DAC

S .

SUBS

YS

CE

NTR

AL

C C

3

F1

DIS

H

DA

TA

C

ON

TRO

LLE

R

A

UP

LIN

K

1 KB

PS

4

OM

Nk

2.5 M

BB

50

A

MP

LITU

DE

jg

KBP

S O

R K

BPS

L

DA

TA

10

BIT

A/D

I)

60 K

BPS

I

OR

r 26

MB

PS

O

R

TAPE

,

,-T

2.6 M

BPS

1.1 S

MBP

S 0 R

1.1

6 M

BPS

AU

GM

EN

T ED

RA

W D

AT

A

L

t 4

PHAS

E 1.

12 M

BPS

+

DA

TA

C

MO

S D

ATA

CM

DS

1 I

i

r

4 B

IT A

ID

POW

ER

AM

PLI

FIE

R

20W

RF

2 KB

PS

.

A

1 I

r - I

r >

DAC

S - DETE

CTO

R

I PR

EPR

OC

ESSO

R

TELE

ME

TRY

48

KB

PS

MU

X

L A

- 1

A

through the TDRSS MA channel. The MA service utilization capability is intended to satisfy currently unanticipated needs and to provide a means of circumventing tape recorder problems/failures. Normally, a full recorder is dumped at 2.5 Mbps while another is recording. One of the recorders can fail with no effect. After the second recorder failure, no data can be recorded while the survivor is in the dump mode - about 2% of the time if data is recorded at 50 kbps.

A secondary mode routes unprocessed amp1 i tude and phase data, augmented by telescope and spacecraft navigation data, through the TDRSS SSA in near-real-time. Such data would not be transmitted while in the TljRSS dead zone. In this mode, the preprocessor also acts as a multiplexer.

Figure 1V.B-6 is the equipment list for the CDMS. The totals for mass and power, conservatively estimated, are 68 kg a-.d 187 watts. If a mission profile were run, the total power to support the CDMS would be less. For instance, ane transponder was considered to be on and trans- mitting at all times, which would not be the case f ~ r a store-and-dump mode. The conservatism should have little effect on the overall spacecraft, as the CDMS power requirement is less than 10% of the power needed for the mission equipment.

1V.C. ATTITUDE CONTROL AND DETERMINATION

1V.C. 1. ATTITUDE CONTROL REQUIREMENTS

Two distinct sets of attitude control requirements flow from the mission objectives.

For large-area wind estimation, the telescope scanning mechanism provide one element of beam pointing; the attitude control functions are then to

o align the telescope scan axis (nominally the spacecraft z-axis 1 with the local vertical (The spacecraft coordinate system is illustrated in Figure 1V.C-1) and

o roll the spacecraft about the z-axis to orient the radiators and solar arrays.

Radiator and solar array orientation is relatively undemanding: with the :pacecraft z-axis aligned with the local vertical, the spacecraft is rolled until the sun is in the plane of the radiators (i.e., the yz-plane). Normally prudent design practice will oversize the radiators to allow 1-3 degrees of misorientation (in any axis). This roll

FIG

URE 19.

B-6

CDM

S SU

MM

ARY

ITE

M

1.

A/D

CO

NV

ER

TE

R

2.

PR

EP

RO

CE

SS

OR

3.

DA

TA

AC

QU

ISIT

ION

&

CO

MM

UN

lCA

TlO

N S

YS

TE

M (

DA

CS

)

4.

TAP

E R

EC

OR

DE

R

5.

TR

AN

SP

ON

DE

R

6.

TDR

SS

DIS

H A

NT

EN

NA

7.

OM

Nl W

HIP

AN

TE

NN

A

8.

PO

WE

R A

MP

LIF

IER

NU

MB

ER

R

EQ

UIR

ED

4 2 4 3 2 1 2 2

UN

IT S

IZE

(CM)

10

X1

5X

5

30.5

X 2

5.5

X 7

.5

20 X

18

X 1

1

15

X2

0X

20

20

x2

1 X

9

91 D

IA. X

15

30

11.5

X6

.5 X

2

UN

IT M

AS

S

(kg

)

1 .O

6.0

3.2

6.4

3.0

10.0

1 .O

0.9

UN

IT P

OW

ER

(W

) O

YE

RA

TIO

NA

LB

T~

ND

BY

511

2511

0

13/4

8W (R

EC

OR

D)

17W

(P

LA

YB

AC

K)

2W (S

TA

ND

BY

)

1015

60/1

0

FIGURE TJZ. C-1 S PACECRAFT COORDINATE SYSTEM

T ELESCOPE SC rN AXIS

FIGURE IX. C-2 LOCAL VERTICAL MISALIGNMENT GEOMETRY

S: SPACECRAFT POSITION

N : NADIR POINT

h: SPACECRAFT ALTITUDE

a : SCAN ANGLE h

I : PULSE IMPACT POINT WITH NO LOCAL VERTICAL MISALIGNMENT ERROR

t /

A s : ALLOWABLE MISS DISTANCE N

I +: PULSE IMPACT POINTS ON GREAT CIRCLE N I HAVINC; MIS DISTANCES A8

Acu 5 : ASSOCIATED LOCAL VERTICAL MISALIGNMENT ERROR

maneuver also orients the solar arrays: the rotational capability (beta-angle conpensating) provided for these will then maximize the output power. Again, a misorientation of 1-3 degrees will have no significant effect.

No firm requirement for the alignment error of the z-axis and local vertical has been derived from the large area wind measurement objective. Obviously, the error should nat be so great that some of the lidar pulses miss the Earth entirely. This consideration leads to an upper bound of 5-10 degrees for the allowable error. A tighter bound will he necessary, for the return signal from pulses which reach the horizon will have too low a signal-to-noise ratio to be use- fuT (at least at the lower altitudes).

A tentative criterion for the alignment error can be based on the observdtion that the reference pulse rate and scan rate irnply an "average" Earth surface sample area of about 50 km x 50 km. This suggests that each pulse should strike the surface within 25 km of the point it would strike if there were no error. Figure 1V.C-2 illustrates the geometry. If the allowable miss distance is 25 km, the corresponding allowable angular misalignment is 5 mrad at a spacecraft altitude of 800 km, 3.9 mrad at 550 km. A constant, or slowly varying, error of this magnitude may not be important, depending on the orientation of the error. For example, referring to Figure 1V.C-2, suppose the spacecraft velocity vector is in the plane of the paper, directed to the right. The error the^ lies in the orbital plane, and the effect is to shift the scan pattern forward or backward along the ground track. On the other hand, suppose the velocity vector is perpendicular to the plane of the paper. The error is now normal to the orbital plane, and the effect is to shift the ground track left or right. The effect of the pattern shift of this magnitude along the ground track poses no problems in itself; the same is true of a crosstrack shift so long as no large coverage gaps are introduced. The overlapping of coverage swaths away from the equator suggests that a constant or slowly -varying crosstrack local vertical alignment error of 4-10 mrad is tolerable.

A ftrther tightening of the local vertical alignment error requirement may arise from consideration of the lag angle miscompensation resulting from the error. This miscompensation will likely be dependent on the design of the lag angle compensator/beam steering subsystem. In any event, the resulting error will reduce the SNR of the return.

As noted in Section I.C., the requirements for the second mission objective ii.e., a capability for intensive local observation) are poorly defined at this point, with ground truthing and localized meterological phenomena being the obvious considerations. The technique c~irently

envisioned for realizing this objective entails halting the telescope scan, locking the telescgpe, discontinuing lag angle compensation, and pointing the telescope axis at an atmospheric targic by using the attitude control system. The definition of the 10 km horizontal resolution (i.e., 10 km x 10 km surface-level target) mentioned in the WPL-37 and -63 reports seems to have been chosen with "ground truthingm in mind: such a target might be reasonably studied by ground- based, balloon-, and aircraft-borne sensors, and the results compared to observations from space.

There are some obvious p: ints to be considered in connection with localized obsertrations. If the atmospheric volume to be observed is small (a few tens of kilometers) and located in the orbital plane, the crosstrack cdmpocent of the wind cannot be measured reliably. If the voluz~e is not in the orbital plane, and the angle between the local vertical from the spacecraft and the spacecraft-to-target line-of- sight is greater than the operational scan angle, then the SNR will be reduced, with accompanying reduction in wicd measuring accuracy. Also, a given small atmosphere target will not be viewable "on demand": if located at the equator, as much as twelve hours could elagsz between viewing oppor- tunities (assuming a single spacecraft).

To avoid having +'*is secondary mission objective become a design driver, fur Lher def ini tion of observational needs is required, particularly considering the potential for ground truthing by inference.

The WPL-37 report discusses a pointing jitter requirement of 2 prad/5 msec - which would permit efficient heterodyning, by ensuring overlap of the transmitter and receiver fields of view. The 5 msec here is intended to be representative of the pulse roundtrip time, and the 2prad is to include errors Cce to structural vibrations, lag-,angle miscornpensation, and scan mirror jitter. Although a requirement of this order is considered achievable, the detailed structural analysis needed for verification is beyond the scope of thj s assessment activity.

1V.C. 2. ATTITUDE DETERMINATION REQUIREMENTS

Knowledge of spacecraft attitude and posi~ion is required for control purposes (e.g., to calculate the orientation of the local vertical in spacecraft coordinates), for gross doppler removal via the frequency synthesizer, and for vector resolution of the horizontal wind. The allowable uncertainty in local vertical alignment corresponding to a horizontal wind uncertainty A W is given by the approximation

E = A W tand - vc

where Vc is he orbital velocity and oL is the scan angle. For nW=l m/s and atmospheric coverage at 550 km altitude, E m 220prad. At 800 km altitude, E M 300prad.

The gross doppler induced by spacecraft - Earth motion will also depend on the beam azimuth relat.tve to the spacecraft velocity vector, according to

D = 2 V c COSB 3-

whence AD = 2 Vp A #

h for # = 900 - i.., the sensitivity to the gross doppler error is greatest when the firing direction is at right angle tc the ground track. At 800 km altitude, the uncertainty

A# corresponding to a 220 kHz doppler uncertainty (1 m/s wind) is 130prad.

IV. C .3. ACDS CONCEPT

Figure 1V.C-3 illustrates the concept. A system of biased reaction vrheels is used to cancel the momcnteum of the rotating telescope, and the resulting zero-momenteum system rotates inertially once per orbit to maintain the telescope scan axis along the local vertical. The reaction wheels also provide the torques to control attitude excur- sions. The magnetic torquers react against the Earth's magnetic field for momenteum desaturation of the wheels. This field can be sensed by the magnetometers, or calculated. The sun sensor is used to define the relationship of the radiators +o the sun.

Two-axis star trackers placed on the anti-sun of the spacecraft permit frequent updating of the D R I R l J - I 1 reference gyro assembly which serves as the primary attitude reference. The attitude information is combined in the digi- tal processor assembly (DPA) with position and velocity information derived from the global positioning system (GPS) to calculate the orientation of the local vertical in spacecraft coordinates. The D P A then issues commands to the reaction wheels to align the spacecraft scan axis with the local vertical.

Figure 1V.C-4 defines the ACDS mass and power requirements.

FIG

URE ISL

. C

-3

Alll

TUD

E C

ONTR

OL A

ND D

ETER

MIN

ATIO

N BL

OCK

DIA

GR

AM

AN

TE

NN

A

C-a S

IGN

AL

s

SE

NS

CR

AS

SE

MB

LY

TRA

CK

ER

S

IGN

AL

AS

SE

MB

LV

El-

M

AG

NE

- TO

ME

TER

S

CO

NTR

OL

INT

EG

RA

- T

ION

A

SS

EM

BLY

I_ C

OM

MA

ND

S

MA

GN

ET

lC

TOR

QU

ER

S

CO

NT

RO

L

I

FIG

URE 19.

C-4

AT

TITU

DE A

ND P

OSI

TIO

N D

ETER

MIN

ATIO

N SU

MM

ARY

SE

NS

OR

S

- S

TA

R SENSOR

- S

UN

SH

AD

E

- S

UN

SE

NS

OR

- G

YR

O P

AC

KA

GE

- G

PS

- G

PS A

NT

EN

NA

S

- M

AG

NE

TO

ME

TE

RS

CO

NT

RO

LL

ER

S

- R

EA

CT

ION

WH

EE

LS

- M

AG

TO

RQ

UE

RS

PR

OC

ES

SO

R

NO

. RE

QU

IRE

D

2 2 2 1 1 2 2 4 3 1

PO

WE

R (

W)

22

2 1 20

40

25 2 36

SIZ

E (C

M 1

17.8

X 1

9.1

X 4

0.6

- - 30

.5 X

30.

5 X

22.

9

30.5

X 3

0.5

X 3

0.5

20.1

AC

RO

SS

.

-

35.6

DIA

X 2

0.3

2.5

DIA

X 1

00.1

24.1

X 2

4.1

X 2

6.7

WE

IGH

T (K

G)

9.1

2.3

15.9

16.8

2.0

3.0

15.9

5.9

9.1

V . M I S S I O N O P E R A T I O N S AND PERFORMANCE

V. MISSION OPERATIONS AND PERFORMANCE

V . A WIND ACCURACY REQUIREMENT: O R B I T SELECTION

Consideration of WPL-37 and -63 yields the following conclusions:

o The sensitivity of achievable wind accuracy to orbital altitude and inclination has not been established.

o This sensitivity is dependent, probably in a vital manner, on atmospheric attenuation and backscatter.

Additionally, whereas prior analyses have consistently considered an 800 km, high-inclination operational orbit

o Simplified analyses indicate that lower orbital hltitudes and inclinations may be possible witt- out important degradation of wind measurement accuracy; some enhancement may be possible.

o Future trade studies of orbital altitude and inclination versus wind accuracy should be based on simulations of the type described in the WPL reports. Even with scch an approach, definite conclusions may be difficult to achieve, due to the uncertainties introduced by imperfectly understood atmospheric phenomena, particularly backscatter.

The remainder of this section interprets wind accuracy as a system figure of merit and traces its (concept- ually straightforward) relationship to orbital altitude and inclination. Also, a sun-synchronous and a 57 degree orbital inclination are compared, using various simplifications.

The primary filnction of the proposed doppler lidar wind measuring system is that of estimating the mean (= average) wind in an atmospheric volume called a resolution element - typically, a horizontal "slab" 1 km deep and 100- 500 km on a side. This is done by calculating a (suitably weighted) average of the estimates from a number of pulses (shots)' directed into the resolution element from different points along the orbit. From theory, the error in the estimate of the average should decrease as the number of shots is increased, i.e.,

where

v avg = standard deviation of the wind estimate (m/s)

r = standard deviation of the estimate of a single shot (m/s)

N = number of shots

The wind accuracy requirement is then expressed as a required value of ravg. There are two approaches to satisfying the requirement: increase the number of shots, or decrease (r. The improvement by the first approach must ultimately reach a ~ractical limit, imposed by the available power: note that halving the error Cavg requires a quadrupling of the number of shots, and this translates directly into a quadrupling of power.

To understand what is involved in decreasing r, consider in its place a typical formula for the standard deviation of the radial wind estimate:

I

where

?r = doppler-shifted laser wavelength (ml

5 = rms velocity width of received spectrum (m/s)

Z: = pulse duration (s)

p = / 2 Vmax where

= maximum velocity to be measured (m/s) vm.*

SNRw= wideband signal-to-noise ratio = m r SNR Figure V.A-1, taken from the WPL-63 report, illustrates the variation in the pulse-derived wind accuracy,, , as a function of the signal-to-noise ratio. Figure V.A-2, also taken from WPL-63, carries this one step further, and suggests that the wind measurement accuracy is relatively insen- sitive to the SNR. However (and as noted), this insensitiv- ity depends upon a number of assumptions, particularly the backscatter profile: the predicted inser~sitivity is there- fore suspect.

If the atmosphere is assumed to he a homogeneous shell of 20 km depth, with an attenuation coefficient of .002 m-1 and a backscatter coefficient f l = 3 x 10-8 rn-l.sr-l, a simplified SNR expression is

FIGURE '51. A-1 W I N D ACCURACY V A R I A T I O N W I T H SNR

2.5

2.0

1.5

q (mi ' )

IBO

0.5

0 ,

, Q

,

,

-

- V k w

I ! 1 I 0 10 20 30 40

SNRw (dB)

Figuec 7.4 Rndial v c t o c i ~ ~ l l ~ ~ n 60n one d h O t v ~ . n a w u h d rig&-@- m ~ r uti. w v b c i . t g l i d l k c p u t t. 0 . 5 mi 1 . 0 m 4.'. V--10 rn r-' .

FIGURE Y . A - 2 WIND ACCURACY S E N S I T I V I N TO S IGNAL-TO-NOISE RATIO

A

20

IS

- E t i 10 cn Q, I

5

0 0 1 2 3 4 s

RMS Error (m/s)

Fqune 1 . 1 3 RMS L n d m u 604 o p W d b a e w e p a h ~ ~ e t ~ h d at Spical SNR and doubte SNR.

I

where

AID = Atmospheric interaction distance, i.e., distance from ground to top of the atmosphere, along the beam

TSR = total slant range, i.e., distance from ground to spacecraft, along the beam

Figure V.A-3 displays two sets of SNR curves, one set corresponding to a sun-synchronous orbital inclination; the other, corresponding to a 57 degree inclination. In both cases, the SNR was calculated for the return from the bottom kilometer of the atmosphere, the scan angle was selected to provide contiguous swaths at a height of 20 km at the equator, and the attenuation coefficient was varied by - + 50% to indicate sensitivity.

The first conclusion to be drawn from Figure V.A-3 is that there is an optimum orbital altitude lying between 400 and 600 km. In fact, while an optimum altitude does exist, more detailed calculations accounting for altitude-varying attenuation and backscatter suggest that the optimum may lie above 800 km. However, the variation of the SNR with altitude in the range 500-800 km appears to be less than 3 db, according to these more refined calculations.

The second obvious conclusion is that the SNR is about 2 db greater at the 57 degree inclination than at the sun-synchronous inclination. This has so far been borne out by the more detailed calculations. More extensive and detailed analyses are needed for confirmation; these must carry the SNR results on into the wind accuracy calculation, and account for the various model atmospheres.

V. B. SHUTTLE PERFORMANCE: DESIGN AND OPERATIONAL IMPLICATIONS

Straightforward use of the Shuttle requires placing the doppler lidar wind-measuring spacecraft into its operational orbit and revisiting the spacecraft for retrieval or for servicing and orbital reboost. No other launch vehicle. is likely to be available. The basic Shuttle capabilities are inadequate for the orbital altitudes and inclinations of interest (inclination 57 deg, altitude 500 km).The possibilities for achieving these orbits are

o OMS kits

o Shuttle augmentation

FIGURE % A-3 COMPARISON OF ORB ITAL ALTITUDES AND INCLINATIONS

L I I 1 I I 1

300 400 m 600 700 800 ORBITAL ALTITUDE tkm)

0 Future upper stages

o Integral propulsion

Figure V.B-1 describes the current Shuttle performance capabilities (JSC 07700, Vol. XIV, Rev. GI for a VAFB launch to a 98 deg-inclinatian. Even with OMS kits the performance capability is quite limited. Shuttle augmentation provides some relief: Figure V.B-2 is one (unofficial) projection of augmented Shut.tle performance capabilities. Even with augmentation, OMS kits are required, and the maximum achievable altitude is only about 680 km.

The current official Shuttle performance projection for a KSC launch does not address inclinations other than 28 deg. Thus, an earlier (again, unofficial) performance projection, given by Figure V.B-3, was consulted. With this projection, the operational spacecraft can be delivered (using OMS kits) to a 57 deg orbit, to an altitude in excess of 800 km, with a considerable performance margin.

Since the Shuttle/OMS kits capabilities to a 57 degree orbital inclination permit a simpler spacecraft design, this should be considered a desirahle inclination, subject to coverage and accuracy considerations. The results of this assessment indicate a minor, perhaps even favorable, accuracy effect, as ccmpared tc the 800 km orbit considered by NOAA/WPL. Coverage necessarily suffers: with a scan angle selected for a vertical range of 20 km, there is no coverage within about 22 deg of either pole. The impact of this loss of coverage on global prediction models has not been assessed.

FIG

URE

V.

B-l

NE

AR T

ERM

CAR

GO

WEI

GHT

VER

SUS

CIR

CU

LAR

O

RB IT

AL A

LTIT

UDE-

VAFB

LAU

NCH,

DEL

IVER

Y O

NLY

SH

UTT

LE P

ERFO

RM

ANC

E IS

A F

UN

CTI

ON

OF

PRO

GR

AM

VA

RIA

BLE

S (

LAU

NC

H D

ATE,

EN

GIN

E TH

RU

ST,,I

ND

IVID

UAL

O

RBI

TER

WEI

GH

T,

ETC

) AN

D M

US

T B

E D

ETE

RM

INE

D O

N A

CA

SE

BY

CA

SE

BA

SIS

CAR

GO

WE

IGH

T IN

CLU

DE

S A

LL

PA

YLO

AD

ITE

MS

AN

D

PA

YLO

AD

SU

PPO

RT

SE

RV

ICE

S

SHAD

ED A

RE

AS

RE

PR

ES

EN

T P

RIM

AR

Y R

EGIO

NS

AFF

EC

TED

BY

PR

OG

RAM

VA

RIA

BLE

S.

PA

YLO

AD

S

REQ

UIR

ING

PER

FOR

MAN

CE

IN T

HE

SE

REG

ION

S S

HO

ULD

V

ER

IFY

CA

PA

BIL

ITY

WIT

H J

SC A

S E

AR

LY A

S P

OS

SIB

LE

PO

TEN

TIA

L E

T D

ISP

OS

AL

PR

OB

LEM

S E

XIS

T FO

R I

NC

LIN

ATI

ON

S B

ELO

W 7

00

FIR

ST

OM

S K

IT

10

6

20

0

30

0

40

0

50

0

60

0

70

0

80

0

i

CIR

CU

LAR

OR

BIT

ALT

ITU

DE

, N

MI

I

I I

I I

I 2

00

4'0

0

60

0

80

0

10

00

1

20

0

CIR

CU

LAR

OR

BIT

AL

ALT

ITU

DE

, K

M

NA

SA

-S-7

9-49

FI

GUR

E P. B

-2

lw

F( '0

3

MA

XIM

UM

[-I PE

RFOR

MAN

CE O

F TH

E AU

GMEN

TED

SHUT

TLE

I VA

FB L

AUNC

H 40 x

1s

INC

LIN

AT

ION

, D

EG

UNO

FFIC

I AL .~~

O=~T

IME

PE

RIO

D

.OMS

A V

RE

SE

RV

ES

= 3.396

.OMS

A V

LA

UN

CH

(1Q

O N

MI)

= 4

76 F

PS

INT

EG

RA

L O

MS

.N

O

RE

N3E

zVO

US

a2 o

FR

PO

N A

OA

~

~]

vE

~~

~c

LE

P

RO

JEC

TIO

N

NA

SA

RE

FE

RE

NC

E M

lSS

lON

4 M

AN

iFE

ST

FIR

ST

OM

S

(32K

LB

P/L

+ F

UL

L O

MS

+ F

UL

L)

RC

S T

O 9

80 I

NC

LIN

AT

ION

I 1

.- I

i C

IRC

ULA

R O

RB

ITA

L A

LTIT

UD

E -

- K

M

FIG

URE

31.

0-3

NEAR

TER

M (P

RE 1

985)

CAR

GO

WEI

GHT

VER

SUS

C IR

CULA

R 08

0 IT

AL A

LTIT

UDE

lea x

lo3

KSC

LAUN

CH,

DELI

VERY

ONL

Y

I U

NO

FFIC

IAL

40 x

103

M

AX

IMU

M R

OU

TIN

E P

ER

FO

RM

AN

CE

E

ST:

19

82-8

5 T

IME

PE

RIO

D

OV

-99

BLO

CK

II

ET

F

UL

L R

CS

3

FU

LL

EP

ST

AN

KS

O

OM

S A

V R

ES

ER

VE

= 3.396

OM

S A

V L

AU

NC

H (T

O 1

00 N

MI.

CIR

C)

= 20

0 FP

S

INC

LIN

AT

ION

, D

EG

F

EB

79

VE

HIC

LE P

RO

JEC

TIO

N

NO

PE

ND

EZ

VO

US

S

TS R

ES

ER

VE

INC

LUD

ED

FIR

ST

OM

S K

IT A

DD

ED

SE

CO

ND

OM

S K

IT A

DD

ED

TH

IRD

OM

S K

IT A

DD

ED

I I

I 10

0 200

; 300

a;

500

600

700

I

I

CIR

CU

LAR

OR

BIT

AL

: A

LTIT

UD

E,

N.

MI.

:

I

I I

I

I #

1 I

1

200

430.

600

800

loo

0

1200

11

00

CIR

CU

LAR

OR

BIT

AL

ALT

ITU

DE

, K

M

APPROVAL

ACCOMMODAT I ONS ASSESSMENT :

SPACEBOJNE DOPPLER L IDAR

WIND PiEASURING SYSTEM

The i n f o r m a t i o n in th i s r e p o r t h a s been r e v i e w e d f o r s e c u r i t y c l a s s i f i c a t i o n . Rev iew of a n y i n f o r m a t i o n c o n c e r n i n g D e p a r t m e n t of Defense o r Atomic E n e r g y C o m m i s s i o n p r o g r a m s h a s b e e n made by t h e h l S F C S e c u r i t y C l a s s i f i c a t i o n Off ice r . T h i s r e p o r t , in i t s e n t i r e t y , h a s been d e t e r m i n e d t o be unc lass i f i ed .

Th is document h a s a l s o t ~ e e n rev iewed and a p p r o v e d f o r t e c h n i - c a l a c c u r a c y .

CHARLES R . DARWIN D i r e c t o r , Preliminary Des ign Oif ice

b~ Qb'm -- L-??~-.

WILLIAM R. MARSHALL D i r e c t o r , P r o g r a m Deve lopment

Documents

NASA TECHNICAL MEMORANDUM · George C. Marshall Space Flight Center ... EPS Block Diagram ... where STS performance limitations, Van