report section 5

Project Bellerophon 207

Author: Nicole Bryan

A.2.0 Avionics

A.2.1 Avionics Introduction The avionics group is responsible for the design of the electrical components onboard our launch

vehicle. A basic avionics system is comprised of navigation, communication, and multi-system

management equipment to allow for accurate control of the vehicle. The employment of an

active navigation system is critical for mission success and to ensure the safety of people,

property, and our vehicle. Communication equipment is necessary for the implementation of the

range safety system and for adherence to FAA regulations. The main pieces of equipment we

identified as part of the multi-system management category include the inertial measurement unit

(IMU), computer processing unit (CPU), sensors, and battery. Communication equipment, self-

destruct hardware, and wiring make up the remainder of the components that will be analyzed in

this section.

Besides component selection, the avionics group performed an analysis on the power and

communications systems employed in our vehicle design. The communications analysis is

necessary to ensure that the signal coverage is large enough and strong enough to handle the

conditions of the launch. The power systems analysis is necessary to aid in sizing a battery that

will be able to support our vehicle’s power usage.

Historically, avionics systems are extremely expensive and contribute largely to the overall cost

of the launch vehicle. For this reason, investigating different ways of reducing this cost was a

vital part of our project. By eliminating excessive system redundancy, using mostly commercial

grade components, and choosing only the individual ‘off-the-shelf’ components we deemed

necessary, we aim to lower the overall cost of an effective avionics system.



A.2.2 Design Methods A.2.2.1 Component Ratings According to the NASA Electronics Parts Assurance Group, there are 3 basic ratings of electrical

components. They are:

Level 1: parts are inherently low risk and are suitable for use in all applications including life

support and mission critical systems. Level 1 active parts are reviewed for radiation hardness1.

Level 2: parts have inherently higher risk than level 1 and are considered moderate risk. Level 2

parts are suitable for most general purpose space flight applications but are not recommended for

life support or mission critical applications unless there is on-orbit reparability. Level 2 active

parts need to be evaluated for radiation hardness1.

Level 3: parts are inherently high risk because there is little dependable data or history available

for them and changes in their materials, designs, and processes may occur continuously without

notification. Level 3 parts are intended for mission applications where the use of high risk parts

is acceptable. Level 3 parts should not be used in single-point failure or single-string

applications unless a very high risk for failure or malfunction is acceptable. Level 3 parts must

be evaluated for radiation hardness and radiation testing is recommended1.

Level 1 parts are commonly called ‘space-rated’ or ‘S-Class’. The term ‘Class-B’ is

synonymous with level 2 parts and level 3 components are known as ‘commercial’ electronics.

Each different class has specified requirements for testing procedures. Table A.2.2.1.1 outlines a

few of these parameters.



Table A.2.2.1.1 Electronic component rating comparison 1

Commercial requirements S-Class requirements Infinite production lot size Built one at a time

No burn-in testing Extended burn-in time (up to 240 hours)

No detailed failure analysis Extended failure analysis No product lifetime testing 1,000 hour product lifetime testing on every

lot Production method allows for no

traceability Traceable to raw material selection

>1% failure per 1,000 hours is considered acceptable

0.0001% failure rate per 1,000 hours is considered acceptable

The difference in testing procedures and type-ratings accounts for a large price difference in

electronic devices. The avionics group has decided upon the use of non-space-rated

(commercial) electronics with the goal of keeping the overall cost of the vehicle as low as

possible.

References 1 “NEPAG Info.” NASA Electronic Parts Assurance Group [online],

URL: http://nepp.nasa.gov/nepag/info.htm [cited 05 March 2008].


Author: Justin M. Rhodes

A.2.2.2 Telecommunications Analysis One of the primary goals of the telecommunication equipment on our launch vehicles is to allow

the vehicle to ‘talk’ with the ground. The telecommunications system keeps mission controllers

informed of the vehicle’s health and its current position as it rises to its destination. The

selection of a signal frequency for the telecommunications system dictates how far the signal can

travel, how much the signal is altered by outside forces, what hardware is needed to build the

system, and the rate at which the system can transmit data. We iteratively evaluated these factors

to determine the signal frequency for our design.

The signal frequency is speed at which the radio waves in the signal will travel. Most

communications frequencies in modern communication exist in the ultra high frequency (UHF)

or the super high frequency (SHF) ranges. The UHF band consists of signal wavelengths 100

millimeters to 1 meter in length and frequencies ranging from 300 Megahertz (MHz) to 3

Gigahertz (GHz). We chose to primarily investigate this frequency range because of its common

use in today’s society for communication. Mobile phones, television signals, two-way radios,

and GPS navigation systems all run on signals in the UHF band, making the technology to

support it common and well-developed.1

We began our investigation by examining the frequency allocations put in place by various

political powers throughout the world. In order to prevent a sort of radio chaos, it is common for

governments and world organizations to divide up the frequency spectrum into segments,

allocating certain divisions for specific uses such as the FM radio range or restricted military

frequencies. In the United States, the Federal Communications Commission (FCC) is in charge

of regulating the use of the radio spectrum. By examining the allocations in use by the FCC, we

were able to select several potential signal frequencies. A listing of some of our options appears

in Table A.2.2.2.1, below.



Table A.2.2.2.1 Potential Signal Frequencies2

Signal Frequency (MHz) Current Use

15.005-15.010 space research

18.068-18.168 amateur satellite

30.005-30.010 satellite identification and space research

137.00-137.025 space operation

272-273 space operation

400.15-401.00 space operation

1427-1429 space operation, telemetry

1525-1530 space operation, mobile satellite





As the table above shows, there are many potential frequencies to choose from. To compare the

available frequencies, we prepared a link budget using data from Purdue University’s Cube-Sat

project and varied different parameters to see their effect on the signal strength’s final margin.

In order for the signal make it to its destination, the final margin must be greater than 3 decibels

(dB). The process of constructing a link budget and calculating the system’s final margin is

described in Section A.2.2.3.

The first signal frequency analysis we performed looked at changing the maximum path length

the signal would have to travel and the resulting link budget’s final margin. A representative set

of frequencies was run through this analysis to determine how far we could send the signal

before it would no longer be strong enough to read. Figure A.2.2.2.1, below, shows this process.



Figure A.2.2.2.1: The final link margin as the path length and signal frequencies change.

(Justin Rhodes)

Figure A.2.2.2.1 shows that as the path length the signal travels increases, the final link margin

decreases. The figure also indicates that as the signal frequency increases, the maximum path

length of a usable signal decreases. Higher frequency signals are more susceptible to

interference and disruption as they pass through space. This decreases the distance they can

successfully travel, suggesting that we can achieve better distance performance by using a lower

frequency signal.

One of the properties of signal frequency is how fast the signal travels through space. High

frequency signals travel very fast, whereas low frequencies move much slower. Due this signal

velocity, a signal being transmitted from a moving vehicle experiences a phenomenon known as

the Doppler shift. To an observer on the ground, the signal will appear to be travelling at a

different speed. This effectively causes the signal to have a different frequency than intended.

The change in the frequency at a given velocity is calculated using Eq. A.2.2.2.1 on the

following page.

0 1000 2000 3000 4000 5000 6000 7000 80000

5

10

15

20

25

Path Length (km)

Link

Mar

gin

(dB

)

401 MHz1.45 GHz2.2 GHz3.0 GHz3 dB Margin



Δ (A.2.2.2.1)

where Δ is the observed Doppler shift in kHz, f is the signal’s original frequency in kHz, v is

the velocity of the vehicle in meters per second, and c is the speed of light in meters per second.

By varying the vehicle’s velocity, we predicted the Doppler shift for our representative set of

potential signal frequencies, as Fig. A.2.2.2.2, below, shows.

Figure A.2.2.2.2: The Doppler shift effect as the vehicle velocity and original signal frequency change.

(Justin Rhodes) In the early design phases of our launch vehicles, it was suggest that the vehicle may reach up to

over 18,000 meters per second during its ascent. As Fig. A.2.2.2.1 shows, at this velocity a high

frequency signal can shift by almost 180 kHz, whereas the 401 MHz signal barely changed 20

kHz. Even at the more moderate velocities, the Doppler shift can mutate the signal by a large

margin. This shift in the frequency will have to be accounted for in ground tracking hardware,

suggesting that a radio system running on a lower frequency may be easier to construct.

Radio communications equipment designed to operate at high frequencies are generally more

complex than their counterparts using lower frequencies. This complexity is caused by the need

to more carefully construct, transmit, and receive the signal to account for higher frequency

signals being more delicate. Factors such as the Doppler shift and the signal degradation as it

0 2000 4000 6000 8000 10000 12000 14000 16000 180000

20

40

60

80

100

120

140

160

180

200

Vehicle Velocity (m/s)

Freq

uenc

y S

hift

(kH

z)

401 MHz1.45 GHz2.2 GHz3.0 GHz



travels through space must be accounted for in the hardware. The cost of a high frequency radio

system is therefore higher than that of a low frequency radio ssystem.3 We are primarily

concerned with cost in the design of our launch vehicles, so we again find an advantage in using

a low frequency signal.

The rate at which the telecommunications system can transmit data from the vehicle to the

ground is also a function of the signal frequency. There are many ways in which a signal can be

changed such that data can be transmitted, and the exact method of data transmission is beyond

the scope of this project. The data rate plays a role in establishing the final link margin of the

system, and we can use it to evaluate the ability of different frequency signals to support the

amount of data the vehicle will need to transmit. Using the same iterative approach as the path

length analysis, the link margin at a range of data rates was calculated for a set of signal

frequencies, as shown in Fig. A.2.2.2.3, below.

Figure A.2.2.2.3: The change in final link margin as more bits per second are transmitted

across a set of different frequencies. (Justin Rhodes)

It can be seen that as more bits per second are pushed through the telecommunications link, the

link margin falls. At a set data rate, as the signal frequency rises, the link margin again falls.

This suggests two things. First, the lower the data rate, the stronger the link margin will be. The

Purdue Cube-Sat project is designed to support data rates up to 8x107 bps. If we can limit the

0 1 2 3 4 5 6 7 8 9

x 107

0

5

10

15

20

25

Data Rate (bps)

Link

Mar

gin

(dB

)




amount of data our vehicle needs to transmit, we can greatly increase the strength of our

telecommunications system. Second, lower frequencies are able to sustain better link margins at

higher data rates than higher frequencies. While similar performance may be possible at

extremely lower data rates, as the data rate requirements increase, lower signal frequencies will

enable better signal margins for the system. We note that it is common for systems that require

high data rate capacity, such as communications satellites, to use higher frequency signals.

These systems must support different capacities than our vehicle will need, thus the logic that the

lower frequency signal will be more robust as the data rate increases holds true.

In each parameter of the telecommunications system’s performance, we find that the lower

frequency signals will better support our final link margin. From this analysis, we choose the

401 MHz frequency to support our vehicle telecommunications system. By selecting the signal

frequency, we can then begin a much more detailed link budget analysis and establish other

system design specifications.

References 1 Pozar, David. Microwave Engineering, Wiley, New York, NY, 2004. 2 "FCC Radio Spectrum Homepage." Federal Communications Commission [online], 2007. URL: http://www.fcc.gov/oet/spectrum [cited 26 February 2008]. 3 Filmer, David. “Link Budget Analysis,” AAE 450 Lecture, Department of Aeronautical and Astronautical Engineering, Purdue University. 23 January, 2008.



A.2.2.3 Link Budget Analysis In order for our vehicle to effectively communicate with a tracking station, a communications

“link” is established. This link is made between two radio systems: a transmitter on the launch

vehicle and a receiver on the tracking system. It should be noted that the tracking system can be

in the form of a ground station, an aircraft in flight, or a satellite in orbit – the link will be similar

across all of these systems. Within the link, a number of factors are analyzed in a linear format,

each contributing a signal strength component. Each component is calculated in decibels (dB), a

dimensionless unit related to the strength of the signal. By converting signal strengths to

decibels, these components can be added together to form the link budget for the vehicle. The

final result from the link budget is the link margin. The link margin describes the final signal

strength once the signal has completed its full path from the transmitter to the receiver. If this

signal strength falls below 3 dB, the signal is considered to be below half strength, making signal

reception and analysis unreliable. Our design has a final margin of 39.22 dB, well above the

3 dB limit. The process below outlines the methodology for constructing and evaluating a link

budget analysis for one of our launch vehicles.

The first step in establishing a link budget is to select a signal frequency. The frequency plays an

important role throughout the link budget. It is a major part in determining the physical system

design and evaluating items such as space and propagation signal loss. For our vehicles, a base

frequency of 401 MHz was selected, as described in section A.2.2.2 of this report.

Once the signal frequency has been selected, the transmitter power, in Watts, is put into the link

budget. The transmitter power helps establish the initial signal strength and is largely dictated by

the vehicle’s size and power limits. More power will allow for a more powerful signal, but that

power will require a larger battery to support it. This causes an undesirable increase in the mass

of the vehicle.

The vehicle itself is made up of physical systems that must be manufactured and assembled

together in the final design configuration. These physical systems will always have certain

imperfections inherent in the materials and construction. In order to account for the

imperfections in the radio system, a line loss factor is added into the link budget for both the



transmitter and receiver. This is an estimated value based on historical data from other radio

communication systems. For our link budgets, an estimate of 1 dB was chosen to represent the

line loss on both the transmitter and receiver.1

The next step in the link budget process is to select the transmitter antenna’s beamwidth. Radio

antennas are essentially pieces of metal that radiate radio waves. The most basic antenna will

emit radio waves of equal strength in every direction. This is inefficient, as a lot of power is

wasted sending waves in directions that are not needed. A more efficient solution is to alter the

construction of the antenna such that the radiation pattern is no longer spherical but is instead a

series of lobes. Most of the signal strength will be focused in the main lobe, and the design will

attempt to minimize the size and strength of the subsequent side lobes. The antenna beamwidth

is a property of the main lobe. This allows for a stronger signal to be sent across a smaller area,

but it uses less power than its omnidirectional equivalent. By knowing the beamwidth, an

approximation of the maximum antenna pointing offset angle is also factored into the link

budget. This approximation allows the analysis to account for errors due to the antenna pointing

in a slightly different direction than intended. Based on other communications systems, we

chose a maximum offset of 27 degrees for our transmitting antenna.1

Knowing the antenna’s beamwidth and the signal frequency, the peak antenna gain for the

transmitter is calculated. The gain value represents the change in the signal radiation pattern

caused by selecting the desired beamwidth. The gain is recorded in decibels for addition in the

link budget and is calculated using Eq. A.2.2.3.1, below.

44.3 20log (A.2.2.3.1)

where G is the peak antenna gain in decibels and is the antenna beamwidth in degrees.

At this point, a feasibility check can be made by calculating the transmitter’s required antenna

diameter. The diameter is directly related to the signal frequency and the antenna beamwidth,

and is calculated from Eq. A.2.2.3.2, below.

(A.2.2.3.2)



where D is the required diameter in meters, f is the signal frequency in GHz, and is the antenna

beamwidth in degrees.

The launch vehicle design constrained the antenna diameter to be within the diameter of the

second stage of the vehicle, but it is assumed that the antenna diameter calculated is that of a

dipole antenna. A dipole antenna is essentially two poles of metal stacked end to end. We

assumed that a more advanced antenna design, one that would have a capacity equivalent to that

of our dipole calculations, could be used instead, but the design of such a system is beyond the

scope of this report. For the purposes of signal coverage, we assumed that this more advanced

design would provide a nearly spherical coverage pattern.

If the antenna sizing is acceptable, the antenna pointing loss is then calculated. This calculation

represents the loss in signal strength due to the antenna pointing error and the chosen beamwidth,

as shown in Eq. A.2.2.3.3.

12 (A.2.2.3.3)

where is the antenna pointing loss in decibels, e is the antenna pointing offset in degrees, and

is the antenna beamwidth in degrees. The net gain of the antenna is then found by adding the

negative pointing loss to the peak antenna gain.

The transmitter system and its antenna are now complete. As another feasibility check, the

equivalent isentropic radiated power (EIRP) of the antenna can be found by adding the

transmitter power, the transmitter line loss, and the net antenna gain together. Calculating the

EIRP compares the antenna to the “ideal” antenna which radiates at equal power in every

direction. This also allows for comparison against other communications systems.

In order to reach the desired tracking station, the signal that the transmitter emits must now fly

through space to reach its intended receiver. Space is an unfriendly place for signals, filled with

pesky items such as radiation and an atmosphere filled with water molecules that do their best to



degrade our poor little signals. In order to account for this in our link budget, a space loss value

is calculated using the empirical formula shown in Eq. A.2.2.3.4, below.

147.55 20 log 20log (A.2.2.3.4)

where is the signal loss due to space, S is the maximum signal path length in meters, and f is

the signal frequency in Hz. A more detailed path length analysis is needed to directly account

for signal refraction and bending through the atmosphere. For the purposes of this analysis, a

maximum path length of 5,000 kilometers was assumed.

In addition to the signal loss due to the signal’s trip through space, there is an additional loss due

to signal propagation and polarization as it travels toward its destination. In our analysis, we

estimated this to be -0.3 dB based off of other communications systems. We feel that this

assumption is reasonable, as the loss is typically low for low frequency signals such as the one

we selected. As the signal frequency gets above the 20 GHz range, the propagation and

polarization losses can become much more significant.1

If we are following the signal as it progresses through the communications link, it will now

arrive at the tracking station, entering the receiver as it is picked up through the receiving

antenna. The evaluation of the receiving antenna is very similar to that of the transmitting

antenna above, but now we are allowed the freedom of choosing the antenna size. This allows us

to balance weaknesses in the transmitter by making a more robust receiver. A large antenna

using more power can make up for a weaker signal strength coming from the vehicle.

By choosing the receiving antenna diameter, its peak gain can be calculated using the empirical

formula given in Eq. A.2.2.3.5, below.



159.59 20 log 20 log 10log (A.2.2.3.5)

where is the peak receiving antenna gain in decibels, is the receiving antenna diameter in

meters, f is the signal frequency in Hz, and is the antenna gain towards the satellite. In this

case, is assumed to be 0.55, based on other communications systems.

The receiving antenna beamwidth can be found by solving Eq. A.2.2.3.2 for using the desired

receiving antenna diameter. Likewise, the receiving antenna pointing loss can be calculated

from Eq. A.2.2.3.3, assuming e to be the same as the transmitting antenna. The net gain for the

receiving antenna comes from the peak gain calculated in Eq. A.2.2.3.5 added to the negative

pointing loss, giving us a complete evaluation for the tracking station.

The signal has now travelled from its origin to its destination, but there are still several factors to

be considered before the final margin can be calculated. The system noise temperature is a

property of the atmosphere in which the antennas exist at a certain point in time. The exact

calculation of this property is beyond the scope of this report, and as such an estimate of

135 Kelvin was taken based off of other communications systems.1

One of the key properties of a communications system is how much data it can send across the

link in a set amount of time. Assuming a digital system, this data rate is measured in bits per

second. Our vehicles are designed to be as simple as possible, which allows us to severely limit

the required data rate for the communications link. Sources suggest that telemetry and tracking

data can easily be communicated on the order of one thousand bits per second.2 Allowing for

small amounts of additional data and the possibility of multiple streams of telemetry data being

sent, we selected a data rate of 9,600 bps.

Having the data rate, the energy per bit to noise power ratio ( ) can be calculated. This is a

normalized signal to noise ratio that will form the first main component of the final link margin.

The calculation of the ratio is given by Eq. A.2.2.3.6, below.



228.6 10 log 10 log (A.2.2.3.6)

where is the energy per bit to noise power ratio, P is the transmitter power, Lf is the

transmitter line loss, Gt is the net transmitter gain, Lpr is the receiving antenna pointing loss, Ls is

the loss due to space, La is the propagation and polarization loss, Gr is the net receiver gain, T is

the system noise temperature in degrees Kelvin, and R is the data rate in bps. With the exception

of the system noise temperature and data rate, all values are in units of decibels. From this, the

carrier to noise density ratio can also be calculated for use in further analysis. This ratio comes

from Eq. A.2.2.3.7.

10 log (A.2.2.3.7)

where is the carrier to noise signal ratio, in decibel-Hertz, is the energy per bit to noise

power ratio in decibels, and R is the data rate in bits per second.

Using data from other communications systems, the required value is equal to 9.6 decibels.

For the final margin, this required value is subtracted from the found above. The last piece of

the margin to be considered is an additional implementation loss due to imperfections in the

system as a whole. Based on other systems, we allow a loss of 2 dB due to implementation

issues. The final margin of the system can now be concluded as the value from Eq. A.2.2.3.6

minus the required and the implementation loss. The link budget listed in Table A.2.2.3.1,

below, follows this procedure.



Table A.2.2.3.1 Vehicle Link Budget Analysis

Item Symbol Value Units frequency f 401 MHz transmitter power P 5 Watts transmitter line loss Lf -1 dB-W antenna gain towards vehicle η 0.55 dB transmit antenna beamwidth θt 60 degrees peak transmit antenna gain Gpt 8.74 dB-i transmit antenna diameter Dt 0.87 m transmit antenna pointing offset et 27 degrees transmit antenna pointing loss Lpt -2.43 dB transmit antenna gain (net) Gt 6.31 dB-i equiv isotropic radiated power EIRP 12.30 dB-W propagation path length S 5,000 km space loss Ls -158.49 dB propagation & polarization loss La -0.30 dB receive antenna diameter Dr 10 m peak receive antenna gain (net) Grp 29.88 dB-i receive antenna beamwidth θr 5.24 deg receive antenna pointing error er 0.20 deg receive antenna pointing loss Lpr -0.02 dB receive antenna gain Gr 29.86 dB-i system noise temp Ts 135 K data rate R 9,600 bps Eb/No Eb/No 50.82 dB carrier to noise density ratio C/No 90.64 dB-Hz bit error rate eb 1x10-5 bps required Eb/No (Eb/No)required 9.60 dB implementation loss -2 dB final margin 39.22 dB

The link budget analysis above represents the final link budget for our launch vehicles. Each

vehicle will be carrying the same avionics package onboard, thus only one link budget is

required. The final margin of 39.22 dB is well above the absolute minimum of 3 dB required.

We feel this will allow the vehicle to survive a large margin of error potentially caused by using

cheaper components and less professional design methods.



References 1. Filmer, David. “Link Budget Analysis,” AAE 450 Lecture, Department of Aeronautical and Astronautical Engineering, Purdue University. 23 January, 2008. 2. Wertz, James R., and Wiley J. Larson, eds. Space Mission Analysis and Design, 3rd ed., Torrance, Kluwer Academic and Microcosm Press, 1999.


Author: Timothy Lorenzana

A.2.2.4 Range Safety Considerations The range safety subsystem in our launch vehicle exists to ensure public safety during the launch

and flight of our vehicle. Public safety is defined as preventing damage or harm to people or

property. If the launch vehicle behaves erratically, fails to remain within its flight corridor, or

malfunctions in a way which prevents it from reaching orbit, then the range safety system can be

used to terminate the flight and recall the launch vehicle to Earth.

The term ‘Range Safety Officer’, or RSO, describes an individual who is responsible for the

remote destruction and consequent flight termination of the launch vehicle should it be deemed a

hazard1. To ensure public safety, the RSO and the personnel working under him/her may

evaluate the launch vehicle design as well as oversee the manufacturing and design process. The

RSO monitors both the launch vehicle and environmental conditions prior to launch for anything

that may necessitate the postponement of the launch. Finally, the RSO monitors and tracks the

launch vehicle during flight. The RSO’s job duration ends when the launch vehicle is destroyed

or reaches orbit2. In our case, we prefer the latter and define ‘orbit’ as an altitude of 300km, as

described in the mission statement.

To maintain public safety, the launch vehicle must fly within a predetermined flight corridor.

The flight corridor is an imaginary boundary that exists in a space above the launch range. The

flight corridor is implemented so that if the launch vehicle loses power during flight, the vehicle

will fall to the ground in an uninhabited area. However an engine failure outside the flight

corridor means it may fall on people or property. Therefore if the launch vehicle exits the flight

corridor, the RSO must destroy the launch vehicle to ensure public safety. Destruction of the

launch vehicle will be done remotely using explosive charges placed aboard the launch vehicle,

due to their high reliability1.



The manner in which the launch vehicle is destroyed is highly circumstantial. In most cases, it is

common to terminate the launch vehicle’s propulsion system prior to destruction. Afterwards,

two real options are available. The first option is to detonate the explosive charges aboard the

launch vehicle while it is still in flight. Detonating the charges allows the debris to dissociate

through the air, lessening the force of impact. The second option is to leave the launch vehicle

intact. Not detonating the charges allows for a localized debris field2. Choosing either option is

the responsibility of the RSO and depends on the current situation. In a study presented in the

book “Streamlining Space Launch Range Safety”, it was shown that the malfunctioning rocket is

often destroyed from dynamic forces before ground personnel reacted.

Flight tracking of the launch vehicle is also monitored internally and flight termination may be

initiated automatically, without input from the RSO. The avionics of the launch vehicle will use

sensor data that pertains to positioning, such as the inertial measurement unit, to initiate flight

termination. The data is used by the flight computer to determine the vehicle’s current position in

relation to its intended flight path. If the computer determines that it is outside of the designated

flight corridor, then the computer will initiate termination automatically. Automated termination

is used as a compliment to the RSO and is not intended to replace him/her.

The Federal Aviation Regulations, or FARs, require that the launch vehicle has at least two

adequate and independent components relaying tracking information from the vehicle to the

ground. According to the FARs, it is acceptable to lose one of the two components needed for

tracking. However, if both components are lost, it is immediate cause to terminate the launch

vehicle. The launch vehicle must be destroyed even if there is no indication that it is straying

from the intended flight path. The launch vehicle is destroyed in this situation to ensure public

safety. It is understood that requiring two components is a form of redundancy used to ensure

public safety and mission success rather than for functionality. This point is used as a guideline

when designing a complete subsystem.

We found through research that companies such as Honeywell and L3 Communications

manufacture and sell complete range safety packages. Honeywell has a system called Ballistic



Missile Range Safety Technology. They claim their system has everything required for a total

mobile range safety solution3. Likewise, L3 Communications has their version of a range safety

system. However, Honeywell’s and L3 Communications packages are too complicated and

considered out of scope for this project. Also, the costs of Honeywell’s and L3 Communications

systems significantly over expand the design budget and are disregarded.

Many components from our existing avionics package will be used to create a partial range

safety subsystem. Using existing components will save on complexity, weight, and cost. The

avionics flight computer (central processing unit) is incorporated into the design and used to

perform the task of autonomous vehicle termination. The subsystem will also use the existing

inertial measurement unit and sensor package. The inertial measurement unit is needed to relay

the tracking data to the flight computer so it can perform its assigned task. The existing sensor

package is needed to relay vehicle health data back to the RSO, which will aid in his task.

The range safety subsystem also requires components that are not in our existing avionics

package. These components include a transmitter, receiver, antenna, global positioning unit,

explosive charges, and associated safe/arm devices.

A separate transmitter and receiver can be combined into one unit called a transceiver. The

transceiver and antenna will be used as a dedicated relay for the destruct signal. It is also a good

idea for the transceiver to be encrypted. The encryption will protect the launch vehicle from

anyone other than the RSO implementing the destruct command. Aerocomm manufactures an

adequate transceiver called the AC4424. The AC4424 transceiver operates on an S-band

frequency, so the chosen antenna must also match this operating frequency in order to function.

Using the I-Fuze S-band antenna from Syntronics will be adequate. Finally, using a global

positioning unit, or GPS, will act as a second, redundant unit working with the inertial

measurement unit.

Price estimates for the explosives and associated safe/arm devices are not included in this report.

This was done because no research material was found on the subject and thus, an analysis could

not be performed.



References 1“Range Safety.” Wikipedia [online], URLhttp://en.wikipedia.org/wiki/Range_safety [cited 25 March 2008]. 2Committee on Space Launch Safety (author), Aeronautics and Space Engineering Board (author), National Research Council (author). Streamlining Space Launch Range Safety, National Academics Press, 2000. 3“Range Safety.” Honeywell Aerospace [online], URL: http://www.honeywell.com/sites/aero/Range-Safety.htm [cited 27 March 2008].



A.2.2.5 Sensor Selection An important part of fulfilling system requirements and reducing the system failure rate includes

the implementation of a variety of different sensors. We have identified pressure sensors,

thermal sensors, force sensors, and flow sensors as being the four most common types. Each of

these senses and transmits data, allowing for an accurate assessment of the overall health of the

launch vehicle. While sensor prices and features are included in this section, control

mechanisms for managing the systems they are used on were determined to be beyond the scope

of this project.

The avionics group has decided to use non-space-rated sensors in the design of our launch

vehicle due to the analysis and information presented in Section A.2.2.1. Table A.2.2.5.1 shows

typical price values for these non-space-rated sensors.

Table A.2.2.5.1 Non-space-rated sensors 1 2 3 4

Sensor type Part number Manufacturer Unit price Thermal TS3-85 Cantherm $10.29 Pressure 2000260 Measurement Specialties Inc. $104.06 Flow 26616 Gems Sensors, Inc. $154.95 Force FSS1500NSR Honeywell Sensing and

Control $46.56

We have budgeted $789.65 to account for the sensors on our vehicle. Included in this estimate

are 2 non-space-rated sensors of each type (thermal, pressure, flow, and force) and a safety factor

of 1.25 to account for any inaccuracies.

For comparison, Table A.2.2.5.1 shows available pricing information for some typical space-

rated sensors.



Table A.2.2.5.2 Space-rated sensors 3 Sensor type Part number Manufacturer Unit price Thermal S311-641/04 Honeywell Sensing and

Control $800

Pressure PPTR Honeywell Sensing and Control

$970

Honeywell Sensing and Control is currently one of the largest manufacturers of space-rated

avionics. Most other companies, however, note that they can manufacture a space-rated sensor if

given complete specifications for the part. Price estimates for specialty parts are not included in

this study due to the lack of attainable information.

References 1"Liquid Level and Interface Sensors." Fluid Components International LLC [online], 2006. URL: http://www.fluidcomponents.com/Aerospace/Products/LiquidLevel/A_ProdLiquid.asp [cited 26 February 2008]. 2"Sensors by Measurement Specialties." Measurement Specialties, Inc. [online], 2007. URL: http://www.meas-spec.com/myMeas/default/index.asp [cited 26 February 2008]. 3"Space Applications." Honeywell Aerospace [online], 2007. URL: http://www.honeywell.com/sites/aero/Space_Solutions.htm [cited 26 February 2008]. 4“Cantherm Thermal Sensoring Products.” Cantherm-Canadian Thermostats and Control Devices, Inc [online], 2008. URL: http://www.cantherm.com/products/control_thermostats/index.html [cited 26 March 2008].



A.2.2.6 Inertial Measurement Unit Selection An inertial measurement unit (IMU) is the main component of an inertial guidance system used

in aerospace vehicles.3 An IMU works by sensing its own rate and direction of motion using a

combination of accelerometers and gyroscopes. These measurements can, then, allow a

guidance computer to track its position. In accordance with an IMU, many vehicles employ

additional systems to correct inaccuracies such as ‘drift’. Due to the short nature of the use of

our particular IMU and the goal for cost-effectiveness, we have chosen not to include any of

these corrective systems in our avionics configuration.

Inertial measurement units are available in many configurations and have a vast amount of

specifications associated with them. The main specifications we are concerned with are its

ability to handle high g-loading and have a sampling rate greater than 100Hz to comply with

requirements specified by the dynamics and control group. The Landmark 10 MEMS IMU,

manufactured by Gladiator Technologies, proved to meet these requirements, however, it is not

considered ‘space-rated’. It can handle approximately 12 g’s of acceleration and has a data

sampling rate of 200 Hz.2 This g-loading is sufficient because is higher than the requirement of

6gs provided by the trajectory group. Some of the critical features of this item are noted in Table

A.2.2.5.1.

Table A.2.2.5.1: LMRK10IMU-300-12 Data Sheet 2

Parameter Value Units Range (accel axis) 300 deg/sec Range (rate axis) 12 g's Bias (in run stability) 10 to 100 deg/hr Temperature Range Operating -40 to 85 deg C Non-operating -55 to 85 deg C Update rate 200 Hz Power Consumption 280 at 3.3 V MW Size 5 x 4.5 x 3 cm Weight 108 g



According to an official price quote by Mike Abbott (Manager Sales & Quality Assurance,

Gladiator Technologies), this item has a unit price of $3,495. However, because Gladiator

Technologies recognizes Purdue University as an accredited institution, a discount is given that

reduces the unit price to $2,695. Including shipping costs, the price for a commercial IMU is

approximately $3,300.2

According to Honeywell’s website, the unit price for a space-rated IMU is approximately

$15,000-$20,000.1 These systems are often configured in a redundant manner for use in a launch

vehicle, making the total estimated price even higher. Due to the fact that the IMU from

Gladiator Technologies possesses the critical and necessary specifications that we require, we

have decided that it is acceptable to employ a commercial IMU in our design.

References 1"Space Applications." Honeywell Aerospace [online], 2007. URL: http://www.honeywell.com/sites/aero/Space_Solutions.htm [cited 26 February 2008]. 2 “Digital MEMS IMU.” Gladiator Technologies [online], 2007. URL: http://www.gladiatortechnologies.com/product_LANDMARK10_IMU.htm [cited 26 March 2008]. 3Wertz, James R., and Wiley J. Larson, eds. Space Mission Analysis and Design, 3rd ed., Torrance, Kluwer Academic and Microcosm Press, 1999.



A.2.2.7 Central Processing Unit Selection A central processing unit (CPU) describes a class of logic machines that can execute computer

programs. The CPU is one of the most crucial components of any electrical system and, because

of its importance, typical launch vehicle CPUs are extremely expensive. For our application, the

CPU will function as our primary flight computer.

The RAD6000 radiation-hardened single board computer was manufactured by IBM for BAE

Systems and has been used onboard numerous NASA spacecraft. In addition to 77 satellites (as

of 2003), the processor is/was used in the Spirit and Opportunity Mars rovers, the Mars

Pathfinder lander, the Deep Space 1 probe, and the Genesis and Stardust sample return missions.

The computer has a maximum clock rate of 33 MHz and a processing speed of about 35 MIPS.

In addition to the CPU itself, the RAD6000 has 128 MB of ECC RAM. The flight boards in this

system have switchable clock rates of 2.5, 5, 10, or 20 MHz. The unit cost for a RAD6000

Central Processing Unit is approximately $250,0001. Because of the high cost of this system,

other options were considered.

Commercial CPUs are not radiation-hardened and do not undergo the extensive testing

procedures that the space-rated CPUs do. These CPUs are mass produced with a variety of

possible specifications and have a very large price range. Based upon the current cost of a

commercial computer, we are able to estimate this cost to be approximately $1,000. We

budgeted $10,000 for this system to account for redundant hardware, an extended software

development cycle, and radiation hardening. We think this system will be sufficient for use in

our launch vehicle due to the relatively low altitude of our flight and the minimal processing

power required.

References 1Scott, Tim. "RAD 6000 Processor." BAE Systems [online], 2008. URL: http://www.baesystems.com/ProductsServices/bae_prod_s2_rad6000.html [cited 26 February 2008].



A.2.2.8 Power Budget Analysis To aid in battery sizing, a power budget analysis was performed using the silver-zinc battery

specifications obtained from BST Systems. Due to the lack of information regarding system

power usage on a launch vehicle, an estimation of 200 Watts was assumed. This value portrays

the typical power usage for a small satellite according to the Space Mission Analysis and Design

text [1] . Because our system is so simple and lacks many of the complicated systems aboard a

small satellite, we believe this estimate is reasonable. Other parameters in the analysis included

the vehicle flight time and balloon rise time. These were based off of the trajectory code’s

output for the nominal trajectory and are in agreement with the final trajectory created by D&C.

Next, we determined the breakdown of percent operating power based upon seven primary

categories that require electrical power. These included payload, propulsion, attitude control,

communications, command and data handling, thermal control, and power management1. By

multiplying the percent of operating power by the total system power of 200 Watts, we were able

to determine the approximate amount of power (in Watts), for each system. Dividing the

wattage of each system by the flight or rise time, we obtained the energy consumption in Watt-

hours. Using an energy density of a typical silver-zinc battery (approximately 110 Wh/kg)2, we

were able to divide the energy consumption by this value and obtain a mass, in kg, for a battery.

Included in our analysis is a safety factor of 1.5. This allowed for the analysis to be valid for a

total system power of 300 Watts. Tables A.2.2.8.1, A.2.2.8.2, and A.2.2.8.3 show the details of

this analysis for each payload.



Table A.2.2.8.1 Power budget analysis for a 200g payload

Flight time Rise time Vehicle System % operating

power Power (W)

Energy consumption (Wh)

Mass (kg)

Power (W)


Mass (kg)

Payload 5 10 1.49 0.01 10 30 0.27 Propulsion 35 70 10.43 0.09 0 0 0.00 Attitude Control 15 30 4.47 0.04 0 0 0.00 Communications 10 20 2.98 0.03 20 60 0.55 Command and Data Handling

5 10 1.49 0.01 10 30 0.27

Thermal 5 10 1.49 0.01 0 0 0 Power Management

25 50 7.45 0.07 10 30 0.27

Total 100 200 29.80 0.27 50 150 1.36 Total with Safety Factor (1.5)

300 44.70 0.41 75 225 2.05

Assumptions: 200 g payload, 3 hour rise time, 536.4 second flight time. Total system power of 200 W, silver-zinc battery with energy density of 110 Wh/kg.

Table A.2.2.8.1 Power budget analysis for a 1kg payload


power Power (W)


Mass (kg)

Power (W)


Mass (kg)


5 10 1.43 0.01

10 30 0.27


25 50 7.16 0.07

10 30 0.27


300 42.94 0.39

75 225 2.05

Assumptions: 1 kg payload, 3 hour rise time, 515.3 second flight time. Total system power of 200 W, silver-zinc battery with energy density of 110 Wh/kg.



Table A.2.2.8.3 Power budget analysis for a 5kg payload


power Power (W)


Mass (kg)

Power (W)


Mass (kg)


5 10 1.57 0.01

10 30 0.27


25 50 7.87 0.07

10 30 0.27


300 47.19 0.43

75 225 2.05

Assumptions: 5 kg payload, 3 hour rise time, 566.3 second flight time. Total system power of 200 W, silver-zinc battery with energy density of 110 Wh/kg.

The mass of the battery that supports the vehicle during flight is much smaller than the one used

during the balloon’s rise, as expected, due to the launch configuration. The validity of this

analysis is based on the fact that the smallest battery BST Systems makes for launch vehicles

weighs approximately 2kg2. Table A.2.2.8.4 shows a summary of the results.

Table A.2.2.8.4 Battery mass summary Payload size Battery mass for

flight time Battery mass for rise time

Units

200 g 0.27 1.36 kg 1 kg 0.26 1.36 kg 5 kg 0.29 1.36 kg

Note: These results are nominal and do not include a safety factor. References 1Wertz, James R., and Wiley J. Larson, eds. Space Mission Analysis and Design, 3rd ed., Torrance, Kluwer Academic and Microcosm Press, 1999. 2“BST Products.” BST Systems, Inc.[online], URL: http://www.bstsys.com [cited 25 March 2008].


Author: Danielle Yaple

A.2.2.9 Battery Selection The specifications for the battery we are using to support the launch vehicle were chosen by

examining batteries that were historically used in space applications and then finding the one that

best fits our needs. Our first source of information was a NASA document titled, “Battery

Selection Practice for Aerospace Power Systems”. This document outlined features to consider

when picking a battery for an aerospace purpose. The best type of battery for use in our vehicle

is a primary battery. Primary batteries are not rechargeable and are used for short durations

similar to our mission. Secondary batteries are used more for long-term missions and satellites.

We limit the primary battery choices to three options: Lithium/Monoflouride (Li/CF),

Lithium/Thionyl Chloride (Li/SOCl2), and Silver/Zinc (Ag/Zn).

Table A.2.2.8.1 Primary Battery Selection Criteria

LiCF Li SOCl2 AgZn

Watt-Hours/Kilogram 130 185 110

Watt-Hours/Liter 160 240 200

Discharge Rate Low Moderate High

Failure Tolerance Low Low High

Cell Voltage 2.95V 3.1V 1.5V

Experience Level High Moderate High

Costs Low Low Low

As shown in the table above, the Silver/Zinc battery has the highest failure tolerance. The

Silver/Zinc batteries have been used on all the Apollo missions and the Delta Rockets. The

Lithium/Monoflouride and Silver/Zinc have similar energy densities (Watt-hour/kg and Watt-

hour/liter). The Lithium/Thionyl Chloride battery is a fairly new battery with limited experience.

Thus, we determined that the Lithium/Thionyl Chloride would have had too great a start up cost

and failure rate for just a slight gain in performance. Because Silver/Zinc batteries have a high

discharge rate, a high failure tolerance, high experience level, and relatively low cost, we chose

to use a Silver/Zinc battery as our primary power source.

Originally we placed one battery on the top stage of the launch vehicle. This design idea was

quickly discarded due to the fact that our battery outweighed the two smaller payloads.


Author: Danielle Yaple

Currently, the design consists of two batteries, one that will be on the gondola to control the

ignition of the first stage and another one which will be located with the CPU and Telecom

systems in the skirt between the second and third stage. The additional cost of having a second

battery is outweighed by the cut in cost due to having less weight on the rocket.

A lot of considerations were taken into account during our battery selection process. However

we were not able to make select a specific battery. Due to unknown parameters in the engines

and other systems, the full power needed was only approximated and not exact. This is a

limitation in our analysis. Also, we only analyzed traditional space-capable batteries. Though

there are significant assumptions made in our avionic analysis, our avionics package is a good

basis for a small payload launch system.

References 1NASA SP-172, "Batteries for Space Power Systems," NASA 1968.

2 “Nickle-cadmium battery.” Wikipedia [online], URL: http://en.wikipedia.org/wiki/Nickel-cadmium_battery [cited 25 March 2008].

3 “Silver-oxide battery.” Wikipedia [online], URL: http://en.wikipedia.org/wiki/Silver-oxide_battery [cited 25 March 2008].

4 “Lithium battery.” Wikipedia [online], URL: http://en.wikipedia.org/wiki/Lithium_battery [cited 25 March 2008].

5Ritchie, A.G. Military Applications of Reserve Batteries, Philosophical Transactions: Mathematical, Physical and Engineering Sciences, 2006.

6 “Silver-sinc battery designs principal characteristics.” Eaglepitcher Technologies, Inc. [online], URL: http://www.eaglepicher.com/NR/rdonlyres/E1F7CBAC-733A-45DC-A2E5-0604EF16737A/0/AgZnPrincipalDesign.pdf [cited 25 March 2008].

7 “BST Products.” BST Systems, Inc.[online], URL: http://www.bstsys.com [cited 25 March 2008]. 8 “Saft lithium battery cells.” Rathbone Energy [online], URL: http://www.rathboneenergy.com/batteries/battery_cells_by_mfg/saft/saft_lithium_cells.htm [cited 25 March 2008].



A.2.2.10 Power Distribution We designed the power distribution system for our launch vehicles based off of information

taken from the Vanguard family of launch vehicles. Our design uses wiring that matches the

MIL-W-16878 specification, which was used on the Vanguard vehicles and is still used in

modern aerospace manufacturing. The power distribution accounts for 2 kg in the first stage, 2

kg in the second stage, and 1 kg in the third stage of the launch vehicles. This mass distribution

is calculated using 10-gage and 22-gage wiring. These two wire sizes were chosen to reflect

average sizes used in the Vanguard vehicles.1 Using cost data from a modern manufacturer, we

budgeted $500 for the power distribution materials.2 Installation costs were budgeted at $7,500

using the average salary of an aviation electrician over a four week time period.3

One third of the wiring is assumed to be the heavier 10-gage wiring. This is used to transport

power from the main battery located in the second stage to devices in the first and third stages.

The rest of the wiring is made up of 22-gage material, reflecting smaller power channels to

individual devices after the bulk of the power has been delivered via the 10-gage wiring. We use

a simplified power distribution system in this analysis as an exact design requires detailed

information about every electronic component in the vehicle, such as the sensor packages,

ignition systems, and fuel supply requirements. Other detailed components of a power

distribution system include power relays and timers, switches, and electrical connectors. These

are neglected in our high level analysis.

The exact power distribution system is unique to each launch vehicle. Due to the similarities in

our three launch vehicles, we have applied the same approximate design across all of the designs.



References 1. Klewans, B., “The Vanguard Satellite Launching Vehicle”, The Martin Company, Engineering Report Number 11022, Baltimore, MD, April 1960. 2. TexWire Wire and Cable, personal communication, 04 March, 2008. 3. "Average Aircraft Electrician Salary." Salary.com [online], Jan. 2008. URL: http://swz.salary.com/salarywizard/layouthtmls/swzl_compresult_national_SC16000294.html [cited 21 February 2008].



A.2.2.11 Tracking Considerations We used the ground tracking code for this analysis. The code provides an understanding of our

signal projection capabilities for our telecommunications system design. The tracking code,

AAE450_Ground_Tracking.m, works by using trajectory data and a simplified signal theory in

order to calculate a ground signal path and coverage area for the system. The signal analysis will

tell us where the receiving ground stations need to be located in order to pick up the signal that is

being broadcast from the launch vehicle.

For the simplified signal theory we make a few initial assumptions. First, we assume that the

Earth is spherical in shape. Second, we assume that the area where the signal intercepts the Earth

is circular. Third, we neglect any antenna mounting errors or design issues. Fourth, we neglect

any atmospheric anomalies which may refract the signal.

Figure A.2.2.11.1 shows the propagation of a directional and omnidirectional antenna. Our

telecommunications system design uses a cluster of directional antennas. However, as stated in

section A.2.2.3 of this report, we can assume the cluster of directional antennas to propagate as

an omnidirectional antenna. Therefore, we will be using the omnidirectional antenna to perform

the ground tracking analysis. Finally, we assume the signal propagation from an omnidirectional

antenna to be spherical in shape. Figure A.2.2.11.2 defines the signal geometry that we use to

perform our analysis.



Figure A.2.2.11.1: Antenna assumptions.

(Timothy Lorenzana)

Figure A.2.2.11.2: Signal cone geometry.

(Timothy Lorenzana)



After defining the signal geometry, we derive Eq. (A.2.2.11.1) through Eq. (A.2.2.11.7) to model

the propagation.

Eq. (A.2.2.11.1) defines the variable θ in Fig. A.2.2.11.2

( )Sh1cos−=θ (A.2.2.11.1)

where h is the altitude of the launch vehicle and S is a line from the origin of the signal to the

point of contact where the maximum path length of the signal touches the Earth. The path length

is a constant and was determined from the link budget analysis in section A.2.2.3 of this report.

Eq. (A.2.2.11.2) defines the radius of the circle where the signal touches the Earth

( )θsinSr = (A.2.2.11.2)

where r is the radius of the circle where the signal touches the Earth, and S is a line from the

origin of the signal to the point of contact where the maximum path length of the signal touches

the Earth.

Combining Eqs. (A.2.2.11.1) and (A.2.2.11.2), we get Eq. (A.2.2.11.3), defined below

( )( )ShSr 1cossin −= (A.2.2.11.3)

where r is the radius of the circle where the signal touches the Earth, S is a line from the origin of

the signal to the point of contact where the maximum path length of the signal touches the Earth,

and h is the altitude of the launch.



In order to plot the signal propagation where the signal intercepts the Earth, a conversion

constant is needed. The conversion constant converts between kilometers and degrees. the

conversion constant is calculated using Eq. (A.2.2.11.4) defined below

3602 EarthR

cπ

= (A.2.2.11.4)

where c is the conversion constant in kilometers per degree, and REarth is the radius of the Earth

in kilometers.

The longitude coordinate of the signal position is calculated using Eq. (A.2.2.11.5) below

crx /)cos(γ= (A.2.2.11.5)

where x is the longitude coordinate in degrees, r is the signal coverage radius in kilometers, γ is a

vector which ranges from 0 to 360 degrees, and c is the conversion constant in kilometers per

degree.

The latitude coordinate of the signal position is calculated using Eq. (A.2.2.11.6) below

cry /)sin(θ= (A.2.2.11.6)

where y is the latitude coordinate in degrees, r is the signal coverage radius in kilometers, γ is a

vector which ranges from 0 to 360 degrees, and c is the conversion constant in kilometers per

degree.

The longitude and latitude coordinates of the signal position are then added to the ground track

longitude and latitude coordinates. Adding the two coordinates together results in a signal path

that will change as the launch vehicle ascends. The ground distance covered during vehicle

ascent is then calculated by subtracting the position when the vehicle reaches orbit from the

position when the vehicle is released from the balloon.



Fig. A.2.2.11.3 below defines the geometry used to approximate the signal coverage area.

Figure A.2.2.11.3: Ground track signal geometry.

(Timothy Lorenzana)

The total signal coverage area is approximated using a combination of two half circles, two

rectangles, and two triangles as stated in Eq. (A.2.2.11.7) below.

( )⎥⎦⎤

⎢⎣⎡ −+++= rRdrdrRA2122

2222 ππ (A.2.2.11.7)

where A is the approximated signal coverage area in square kilometers; R is the initial signal

coverage radius in kilometers; r is the final signal coverage radius in kilometers; and d is the sum

of the balloon drift distance and the ground distance covered, in kilometers.

The ground track for the 200g payload launch vehicle is shown in Fig. A.2.2.11.4. A ground

track for the 1kg payload launch vehicle is shown in Fig. A.2.2.11.5. A ground track for the 5kg

payload launch vehicle is shown in Fig. A.2.2.11.6.



Figure A.2.2.11.4: Track from balloon release to orbit for 200g payload launch vehicle.

(Timothy Lorenzana)

Figure A.2.2.11.5: Track from balloon release to orbit for 1kg payload launch vehicle.

(Timothy Lorenzana)



Fig. A.2.2.11.6: Track from balloon release to orbit for 5kg payload launch vehicle.

(Timothy Lorenzana)

The blue outline in Fig. A.2.2.11.4, Fig. A.2.2.11.5, and Fig. A.2.2.11.6 represents the North, the

South, and the Central Americas. The red line in Fig. A.2.2.11.4, Fig. A.2.2.11.5, and Fig.

A.2.2.11.6 represents the ground track from balloon release until the launch vehicle reaches

orbit. The distance between the Florida coast (blue) until the start of the launch vehicle ground

track (red) in Fig. A.2.2.11.4, Fig. A.2.2.11.5, and Fig. A.2.2.11.6 represents the drift distance

covered by the balloon. The green circles in Fig. A.2.2.11.4, Fig. A.2.2.11.5, and Fig. A.2.2.11.6

represent the signal coverage areas at arbitrary instances in time. Since the signal coverage areas

are quite large and overlap extensively, the final coverage area for each launch vehicle is

approximated as the initial coverage area.

In Fig. A.2.2.11.4, Fig. A.2.2.11.5, and Fig. A.2.2.11.6, the final signal coverage area contains

the entire ground track. This means that only one ground station is necessary to cover the entire

ascent of the launch vehicle. Existing ground tracking station locations were researched.

However, due to time constraints, no optimization concerning the ground station location was

performed. In order to receive transmissions from the vehicle during ascent, any location within

the initial green circle is assumed to an adequate ground station location. Costs concerning the

tracking station are discussed in section A.9.2.1 of this report.



Table A.2.2.11.1 below gives the balloon drift distance from ground release until each launch

vehicle separates from the balloon.

Table A.2.2.11.1 Balloon Drift Distance

Launch Vehicle Payload Drift Distance Units 200 g 120.74 km 1 kg 120.45 km 5 kg 121.74 km

Balloon drift is discussed in section A.4.2.1.2.2 of this report.

Table A.2.2.11.2 below gives the ground distance covered for each of the launch vehicles during

ascent.

Table A.2.2.11.2 Ground Distance Covered

Launch Vehicle Ground Distance Units 200 g 236.10 km 1 kg 355.16 km 5 kg 215.12 km

The 1kg payload launch vehicle takes longer to reach orbit, thus the ground distance covered is larger than its counterparts.

Table A.2.2.11.3 below gives the total signal coverage area for each of the launch vehicles.

Table A.2.2.11.3 Signal Coverage Area

Launch Vehicle Signal Area Units 200 g 1.6036e+008 km2 1 kg 1.6155e+008 km2 5 kg 1.6016e+008 km2

The 1kg payload launch vehicle takes longer to reach orbit, thus the signal coverage area is larger than its counterparts.



A.2.3 Avionics Closing Comments Historically, the cost for a complete and sophisticated avionics package is well over $1,000,000.

Throughout the vehicle design process, the goal of the avionics group was to determine the

differences in available systems and to eliminate redundant or unnecessary components. We

accomplished this goal by employing commercial components into our final launch vehicle

design.

In a subsequent design, the team members should start with a broad knowledge of electrical

systems. The avionics team spent a good amount of time learning the most basic concepts of

power systems and this left less time for design. The group should further investigate the power

budget to come up with a better estimated value for the total system power, or find a way to

derive it themselves. The basic concept and structure for this analysis is provided in section

A.2.2.8. Also, the team should investigate the electrical systems behind engine ignition and

launch vehicle separation, as these systems were not included in our analysis due to their

complicated nature. The material presented in this report provides a basic overview of the

avionics components or systems we have chosen to employ in our launch vehicle design. The

main goal of a subsequent vehicle design would be to expand on and strengthen each of these

sections to the best of their knowledge.


Compiled By: Justin M. Rhodes

A.2.4 User’s Guide to Running the Avionics Code Compiled by Justin Rhodes The avionics code collection consists primarily of small scripts designed to analyze and visualize different aspects of the telecommunications systems onboard the launch vehicle. Due to the nature of the topic, much of our analysis was performed at a high level, leaving fewer direct calculations than other areas of the project. Below you will find our efforts to quantify and support our chosen design. Index fnc_link_budget.m J. Rhodes page 2 margin_vs_distance.m J. Rhodes page 4 bps_eval.m J. Rhodes page 6 doppler_shift.m J. Rhodes page 8 AAE450_Ground_Tracking.m T. Lorenzana page 9



User’s Guide for fnc_link_budget.m Written by Justin M. Rhodes

Revision 1.0 – 30 January 2008 Description: The function fnc_link_budget.m accepts a series of input parameters that describe the basic design of a telecommunications system. These inputs are then processed through the link budget calculations to determine the link budget margin and several other system design parameters. Assumptions: The following variables were assumed to have certain values. An explanation of these assumed values can be found Section A.2.2.3.

Variable Name Description Assumed Value

transmit_line_loss Signal line loss on the transmitted [dB] -1

propagation_polarization_lossSignal loss caused by propagation and polarization effects [dB]

-0.3

receive_pointing_offset Pointing error on the receiving antenna [degrees] 0.2

system_noise_temp Noise temperature level of the path through which the signal travels [Kelvin]

135

required_signal_to_noise Required level of signal to noise ratio for the system [dB] 9.6

implementation_loss Signal loss due to issues resulting from putting the whole system together [dB]

-1

Call line: The function can be called with the following line of code: [margin, transmit_diameter, eirp, receive_beamwidth] = fnc_link_budget(frequency, transmit_power, transmit_beamwidth, transmit_pointing_offset, path_length, receive_diameter, data_rate)



Inputs: All of the variables that are passed into the function are described below: Variable Name Description frequency Signal frequency [GHz] transmit_power Power supplied to the transmitter [Watts]

transmit_beamwidth Beamwidth angle of the transmitting antenna’s main lobe. This is a property of the antenna design. [degrees]

transmit_pointing_offset Pointing error in the transmitting antenna [degrees]

path_length Maximum path length the signal is required to travel [km]

receive_diameter Diameter of the receiving antenna, assuming a dipole antenna [m]

data_rate Required data rate across the telecom link [bps] Outputs: All of the variables that returned by the function are described below: Variable Name Description margin Final signal link margin [dB]

transmit_diameter Diameter of the transmitting antenna to support the design, for a dipole antenna [m]

eirp Equivalent isotropically radiated power – the amount of power an isotropic antenna would have to emit for this design. Design comparison tool. [dB-Watts]

receive_beamwidth Pointing error in the transmitting antenna [degrees]



User’s Guide for margin_vs_distance.m Written by Justin M. Rhodes

Revision 1.0 – 30 January 2008 Description: The script margin_vs_distance.m calculates the final signal margin for a range of maximum path lengths. Multiple signal frequencies can be evaluated at once. The resulting data is then plotted for visual analysis. Uses: This script calls the following function: fnc_link_budget.m User Inputs: All of the variables that are initially set in the script are described below: Variable Name Description transmit_power Power supplied to the transmitter [Watts]




data_rate Required data rate across the telecom link [bps]

max_path Maximum path length the signal is required to travel. Path lengths from 0 to max_path are evaluated. [km]



Sample Output: This script produces a figure in the following format:

0 1000 2000 3000 4000 5000 6000 7000 8000

0

5

10

15

20

25

Path Length (km)

Link

Mar

gin

(dB

)




User’s Guide for bps_eval.m Written by Justin M. Rhodes

Revision 1.0 – 30 January 2008 Description: The script bps_eval.m calculates the final signal margin for a range of required data rates. Multiple signal frequencies can be evaluated at once. The resulting data is then plotted for visual analysis. Uses: This script calls the following function: fnc_link_budget.m User Inputs: All of the variables that are initially set in the script are described below: Variable Name Description transmit_power Power supplied to the transmitter [Watts]




path_length Maximum path length the signal is required to travel [km]

max_datarate Required data rate across the telecom link. Values from 0 to max_datarate are evaluated. [bps]



Sample Output: This script produces a figure in the following format:

0 1 2 3 4 5 6 7 8 9

x 107

0

5

10

15

20

25

Data Rate (bps)

Link

Mar

gin

(dB

)




User’s Guide for doppler_shift.m Written by Justin M. Rhodes

Revision 1.0 – 30 January 2008 Description: This script doppler_shift.m calculates Doppler shift in kHz at a given vehicle velocity. Multiple signal frequencies can be evaluated at once. The resulting data is then plotted for visual analysis. User Inputs: All of the variables that are initially set in the script are described below:

Variable Name Description

delta_v Row array of velocity values to evaluate in the form of [startvalue:increment:endvalue]. [m/s]

frequency Row array of signal frequencies to evaluate. [MHz] Sample Output: This script produces a figure in the following format:

0 2000 4000 6000 8000 10000 12000 14000 16000 180000

20

40

60

80

100

120

140

160

180

200

Vehicle Velocity (m/s)

Freq

uenc

y S

hift

(kH

z)

401 MHz1.45 GHz2.2 GHz3.0 GHz



User’s Guide for AAE450_Ground_Tracking.m Written by Timothy Lorenzana

Revision 9 - 27 March 2008 Description: The purpose of this code is to have a graphical representation of the signal projection from the launch vehicle to the ground. The computational computer language used was MATLAB. The code works by using nominal trajectory data and simplified signal theory, as discussed in section A.2.2.10 of this report, in order to calculate a ground signal path and coverage area. Assumptions: The code assumes the balloon drifts in only one direction, to the East, of its initial launch point. The code also assumes the Earth only rotates about its North Celestial Pole. Input Section: The code has two input sections. The first input section is titled “User Defined Inputs”. The only input here is the maximum path length for the telecommunications system being used. The second input is a load file titled “ephemeris.text”. This file is an output of the trajectory code, and its contents can be found in trajectory portion of the appendix. The inputs used in this code are listed below. Variable Name Description path_l Maximum antenna path length [km] t = ephemeris(:,1) Time vector [seconds] x = ephemeris(:,2) Position vector in the Cartesian coordinate system [meters] y = ephemeris(:,3) Position vector in the Cartesian coordinate system [meters] z = ephemeris(:,4) Position vector in the Cartesian coordinate system [meters]

Output Section: The code outputs a figure with the Earth plotted in blue, the ground track plotted in red, and the signal coverage area plotted in green. The code also outputs the total distance covered by the ground track, and an approximate signal coverage area.



The variables given as outputs of this code are given in the table below.

Variable Name Description ground_track Total ground distance covered during ascent [km] signal_area Approximate signal coverage area [km2]

Sample Output: This code produces a figure in the following format:

Documents

report section 5