PROJECT LOCOST

(Laser Or Chemical hybrid Orbital Space Transport)

June 19, 1990

VPI & SU Aerospace Senior Design Team

Department of Aerospace and Ocean Engineering

Virginia Polytechnic Institute & State University

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Abstract

Project LOCOST

A potential mission in the late 1990s is the servicing of spacecraft assets located in GEO. The

Geosynchronous Operations Support Center (GeoShack) will be supported by a space transfer vehicle based at the Space Station (SS). The vehicle will transport cargo between the SS and the

GeoShack. A proposed unmanned, Laser Or Chemical hybrid Orbital Space Transfer vehicle

(LOCOST) can be used to efficiently transfer cargo between the two orbits.

A preliminary design shows that an unmanned, laser/chemical hybrid vehicle results in the fuel

savings needed while still providing fast trip times. The LOCOST vehicle receives a 12 MW laser

beam from one earth orbiting, solar pumped, iodide Laser Power Station (LPS). Two Energy

Relay Units (ERU) provide laser beam support during periods of line-of-sight blockage by the

earth. The baseline mission specifies a 13 day round trip transfer time. The ship's configuration

consist of an optical train, one hydrogen laser engine, two chemical engines, a 18 m by 29 m box

truss, a mission-flexible payload module, and propellant tanks. Overall vehicle dry mass is 8,000

kg. Outbound cargo mass is 20,000 kg, and inbound cargo mass is 6,000 kg. The baseline mission

needs 93,000 kg of propel/ants to complete the scenario. Fully fueled, outbound mission mass is

121,000 kg.

A regeneratively cooled, single plasma, laser engine design producing a maximum of 768 N of thrust

is utilized along with two traditional chemical engines. The laser and chemical engines fire independently, though not simultaneously, to allow the vehicle to make an in-plane spiral trajectory to

the desired destination. The chemical engine is primarily used for maneuvers such as plane changes

and recircularizations. Collection and transmission of the laser beam to the single laser engine is

accomplished via a 14 m diameter, primary collection mirror, a secondary convex parabolic refocusing mirror, two redirection mirrors, and a final, focusing fifth mirror.

The payload module is designed to hold 40,000 kg of cargo, though the baseline mission specifies

less. The payload module is supported by a space truss--either the truss or the payload module is

capable of detaching independently, adding greater cargo fle,'dbility. The chemical rockets are

gimballed about their supports through +/- 7 ° while the laser rocket is attached to a rocket basket

which moves perpendicular to the long axis of the main truss. By moving the rocket basket, the

line of thrust can be made to pass thru the instantaneous center of mass.

The maximum cargo mass that can be transferred is 40,000 kg. A 40,000 out/40,000 back cargo

transfer scenario is considered. For this scenario, round trip time is 23 days and total propellants

used is 220,00 kg.

A proposed design of a laser/chemical hybrid vehicle provides a trip time and propellant efficient

means to transport cargo from the SS to a GeoShack. Its unique, hybrid propulsion system provides safety through redundancy, allows baseline missions to be efficiently executed, while still allowing for the possibility of larger cargo transfers.

Abstract ii
Preface

This report was prepared by the Virginia Tech aerospace senior design team to fulfill their senior

design proiect requirement (AOE Design 4066). Dr. Jakubowski was the coordinating and guiding

faculty member who oversaw the entire project. The team members included

Alan Dixon

Alicia Kost

Gregory Lampshire

Rob Larsen

Bob Monahan

Geoff Wright

Because of the small size of the design group, many subsystems were not fully explored. This report

gives details for the key subsystems but only attempts to discuss some specifications and ordinary

choices for other subsystems.

Special thanks to Dr. John Costain for letting the group utiliTe his computer resources, and to Scott

Frame for CAD help.

Preface ill
Table of Contents

Introduction ............................................................ !

Background ............................................................ 1

Project Objectives ........................................................ 1

The General Mission Scenario ............................................... 2

Redundancy of the Propulsion System ....................................... 3

Extended Mission Capability .............................................. 3

Vehicle Configuration ..................................................... 4

Design Rationale ........................................................ 4

Configuration Evolution ................................................... 5

Configuration 1 ........................................................ 5

Configuration 2 ........................................................ 5

Selected Configuration .................................................... 6

Discussion of the Selected Configuration ....................................... 6

Main Truss ........................................................... 6

Propulsion ........................................................... 6

Chemical Propulsion ....................................... '. .......... 6

Laser Propulsion ..................................................... 6

Propellant Storage ..................................................... 10

Optical System ....................................................... 10

Optical Train ....................................................... 10

Opto-Electronic Feedback System ........................................ 10

Payload Module ...................................................... 13

Attitude Control ...................................................... 13

Rocket Control System (RCS) ............................................ 13

Communications System ................................................ 14

Electrical Power ...................................................... 14

Configuration Analysis and Mass Breakdown ................................... 14

Orbital Mechanics ...................................................... 17

Introduction ........................................................... 17

Mission Assumptions .................................................... 17

Program OM ........................................................ 17

Station Solar Occultation ................................................ 19

The Baseline Mission .................................................... 19

The Maximum Cargo Scenario ............................................. 21

Other Scenarios ........................................................ 21

Conclusions ........................................................... 22

(Laser Or Chemical hybrid Orbital Space Transport) ............................. 22

(Laser Or Chemical hybrid Orbital Space Transport) ............................ 22

Optical System ......................................................... 23

Optical System Evolution ............................................... 23

Optical Train Geometry .................................................. 23

Primary Mirror ....................................................... 25

Secondary and Third Mirror ............................................. 27

Fourth Mirror ........................................................ 27

Fifth Mirror ......................................................... 27

Table of Contents iv
MirrorAssembly...................................................... 27

OpticalTrainControl .................................................. 27

MotorsandPowerRequh'ements.......................................... 30

Dielectrics........................................................... 30

Cooling ............................................................ 34

Propulsion ............................................................ 39

Laser Propulsion ....................................................... 39

Introduction ......................................................... 39

Methods of Energy Transfer .............................................. 39

Assumptions ......................................................... 39

Laser Propulsion Parameters ............................................. 40

Chemical Propulsion ..................................................... 43

Introduction ......................................................... 43

Chemical Propulsion Parameters .......................................... 43

Chemical Rocket Cooling ............................................... 45

Materials ............................................................ 46

Injectors ............................................................ 46

Pumps and Turbines ................................................... 46

Expander ........................................................... 46

Structures ............................................................ 48

Introduction ........................................................... 48

Truss Structure ......................................................... 48

Environmental Effects .................................................... 49

Joint Design ........................................................... 50

Structural Analysis ...................................................... 52

Thermal Analysis ....................................................... 54

Laser Rocket Basket Design ............................................... 56

Cargo Bay ............................................................ 56

Docking .............................................................. 58

Propellant Storage ...................................................... 62

Introduction ........................................................... 62

Tank Material ......................................................... 62

Tank Insulation ........................................................ 63

Tank Support Structures .................................................. 66

Propellant Pumps ....................................................... 69

Dynamical Control System ................................................ 70

Introduction ........................................................... 70

Discussion of Gyro Technology ............................................. 70

Discussion of Gyroscopes ............................................... 73

Specification of CMGs ................................................... 74

Desaturization Rocket System .............................................. 76

Placement of RCS ..................................................... 76

Acquisition, Tracking and Pointing .......................................... 79

Introduction ........................................................... 79

The ATP System ....................................................... 79

Tracking Laser ...................................................... 80

Optical Receiver and Transmitter ........................................ 81

Mirror Alignment System .............................................. 83

Typical ATP Scenario .................................................... 85

ATP Phase I--Aiming of the LOCOST vehicle ................................ 85

ATP Phase III--Acquisition of Pilot Beam ................................... 85

ATP Phase IV--Tracking of the Power Beam ................................. 85

Communications ........................................................ 87

Introduction ......................................................... 87

Table of Contents v
Communication system requirements ....................................... 87

System Protocol and Coding ............................................. 87

Antennas ........................................................... 89

Summary ........................................................... 89

Electrical Power Sytems .................................................. 90

Introduction and System Comparison ....................................... 90

Components ......................................................... 90

Preheater System and Cooling ............................................ 92

Water Removal System ................................................. 92

Fuel Cell Housing ..................................................... 93

Appendix A. Mirror Cooling Calculations ...................................... 94

Appendix B. NOTS Input and Calculations .................................... 96

Appendix C. Thermal Balance Calculations .................................... 98

Appendix D. Propellant Tank Calculations .................................... I00

References ........................................................... 103

Table of Contents vi
List of Illustrations

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1. Configuration 1 ................................................. 5

2. Selected Contiguration--Top View ................................... 7

3. Selected Configuration--Side View ................................... 8

4. Selected Configuration--Rear View ................................... 9

5. Primary Collection Dish-Side View .................................. 11

6. Incident Laser Beam Path ........................................ 12

7. Location of LaserCom Equipment .................................. 13

8. Station-solar Occultation and Earth Blockage .......................... 18

9. Basic Geometrical Description of Vehicle Vectors ....................... 19

10. Optical System Configuration Side View .............................. 24

11. Optical System Configuration Rear View ............................. 25

12. Primary Mirror View 1 .......................................... 26

13. Primary Mirror View 2 .......................................... 28

14. Primary Mirror Off Isometric View .................................. 29

15. Secondary Mirror .............................................. 30

16. Third and Fourth Mirrors ........................................ 31

17. Different Positions Due to Rotations of Primary Mirror Assembly ........... 32

18. Fifth Mirror .................................................. 33

19. Linear Actuator Used to Move Fifth Mirror ........................... 34

20. Dielectric Layering Scheme ....................................... 33

21. Feedback Control System ........................................ 36

22. Radiator Placement and Heat Pipe Front View ......................... 37

23. Radiator and Heat Pipe Placement Side View .......................... 38

24. Laser Rocket Engine ............................................ 41

25. Laser Rocket Engine Mounting .................................... 43

26. Chemical Engine Geometry I ...................................... 45

27. Gimbal Actuator ............................................... 45

28. Scheme to Transport the Maximum Cargo ............................ 49

29. Titanium End Node Connector .................................... 51

30. Inertial Loads on Main Truss ...................................... 53

31. Thermal Contributors ........................................... 55

32. Laser Rocket Basket Top View and Side View ......................... 57

33. Cargo Bay .................................................... 59

34. Cargo Bay Stringers ............................................. 60

35. Attachment Mechanisms for Cargo Bay Shock Absorbers ................. 61

36. Spherical Oxygen Tank .......................................... 63

37. Insulation .................................................... 65

38. Propellant Tank Support Structure 1 ................................ 67

39. Propellant Tank Support Structure 2 ................................ 68

40. LOCOST Vehicle's Body Axes ..................................... 71

41. Fiber Optic Gyroscope (FOG) Schematic ............................. 75

42. CMG and RCS Placement on the vehicle ............................. 77

43. Interrelationships of ATP with Other LOCOST Subsystems ............... 80

44. Optical Trasceiver Labelled Schematic ............................... 82

45. Polarizing Beam Splitter .......................................... 83

46. Mirror Alignment System for Primary and Secondary Mirrors .............. 84

47. Mirror Alignment System for Fifth Mirror ............................ 85

48. Laser Rocket Mirror Alignment System .............................. 86

List of Illustrations vii
Figure 49. Block Diagram of Sub-System Interactions of the ATP System ............. 86

Figure 50. PIP Sample ................ : .................................. 88

Figure 51. Basic Components of a Fuel Cell Power Generator ...................... 92

.°°

List of Illustrations vm
List of Tables

Table I. Basic Requirements and Assumptions ................................. 2

Table 2. Important Design Issues for the LOCOST Vehicle ........................ 4

Table 3. Mass Breakdown ............................................... 15

Table 4. Main Engines .................................................. 16

Table 5. Center of Mass Locations ......................................... 16

Table 6. Baseline Mission Assumptions ...................................... 20

Table 7. Baseline Scenario Trajectory Analysis Results ........................... 20

Table 8. Mass Ratios ................................................... 21

Table 9. The Maximum Cargo Scenario--40,000 kg Out and Back .................. 21

Table 10. Maximum Cargo Scenario--Performance Ratios ......................... 21

Table 11. Mirror Reflectivities ............................................. 35

Table 12. Optical System Thermal Analysis .................................... 35

Table 13. Optical Train Masses ............................................ 36

Table 14. Basic Assumptions .............................................. 40

Table 15. Laser Engine Parameters .......................................... 42

Table 16. Chemical Propulsion-NOTS Input .................................. 44

Table 17. Chemical Propulsion-NOTS Results ................................. 44

Table 18. Chemical Propulsion Rocket Geometry ............................... 44

Table 19. Chemical Propulsion Individual Component Masses ...................... 46

Table 20. Possible Materials for Space Truss ................................... 48

Table 21. Possible Materials for Joint Design .................................. 50

Table 22. Main Truss Mass Breakdown ...................................... 52

Table 23. Mass Breakdown of the Joints ...................................... 52

Table 24. Simplifying Assumptions For Thermal Analysis ......................... 54

List of Tables ix
Table 25. Stress Analysis of Loads .......................................... 55

Table 26. Material Choices for Actuator Wheels ................................ 56

Table 27. Mass Savings for Cargo Container ................................... 58

Table 28. Comparison of Aluminum Alloys ................................... 62

Table 29. Tank Data .................................................... 62

Table 30. Foam Comparison .............................................. 64

Table 31. Insulation Summary ............................................. 66

Table 32. Comparison of Composite Materials ................................. 66

Table 33. Mass Breakdown of the Tank Support Structures ........................ 69

Table 34. Fluid Line Diameters ............................................ 69

Table 35. Possible Momentum Storage Devices ................................. 70

Table 36. Low Frequency Spacecraft Disturbances .............................. 71

Table 37. Tradeoffs for Different Momentum Exchange Actuators ................... 72

Table 38. Main Points with Momentum Exchange Devices ........................ 72

Table 39. Momentum Exchange Device Selection Heuristics ....................... 73

Table 40. Advantages of a Fiber Optic Gyroscope ............................... 74

Table 41. Moments of Inertia .............................................. 75

Table 42. Maximum Precession Angle of Gyros and Angular Rates .................. 75

Table 43. Selected Baseline CMG ........................................... 76

Table 44. CMG and RCS Summary ......................................... 78

Table 45. Major Components of ATP System .................................. 79

Table 46. Possible Lasers for the LCOTV ..................................... 80

Table 47. Vehicle Power Requirements ....................................... 91

List of Tables x
1990 LOCOST Senior Design Project

Introduction

Background

One of the NASA goals is to establish a low earth orbiting space station. Establishing this station

provides an excellent foothold for further space exploration and industrialization. However, to

provide the longterm logistical support for such an endeavor, instruments and equipment will need

to be provided to the station in a cost efficient manner. Also, it is likely that a significant amount

of scientific and engineering activities performed in the geosynchronous orbit (GEO) will lead to

an increase in Low Earth Orbit (LEO) to GEO traffic.

An attractive location to store valuable assets and to service a wide range of spacecraft and platform

designs is a depot situated in a geosynchronous orbit. A Geosynchronous Operations Support

Center (GeoShack) orbiting at GEO would require less orbit maintenance and provide an ideal

storage environment.

To provide logistical support for the GeoShack, it will be necessary to ferry cargo from the Space

Station (SS) to the GeoShack. Manned flights will be made, but a large percentage of the cargo

can be moved by an unmanned space transfer vehicle.

Project Objectives

Specitically, the project objective is an efficient two-way transfer of cargo between LEO and GEO.

By "efficient", it is meant that the mission:

• Minimizes propellant usage

• Minimizes round trip time

• Maximizes cargo mass transfer capability

• Insures adequate safety through redundancy

The type of propulsion for the vehicle greatly influences the characteristics of a mission. The propulsion for an orbital transfer vehicle can be either low thrust or high thrust. Possible high thrust

propulsion devices are:

1. Nuclear

2. Chemical

Possible low thrust propulsion devices are:

1. Electrical

2. Laser

The high thrust propulsion devices give quick trip times but require high propellant usage. Low

thrust devices provide propellant efficiency but result in long trip times. These two types of propulsion devices have strengths and weaknesses. In order to capture the strengths of each, a hybrid

vehicle can be designed. This vehicle would exhibit efficient propellant usage due to its low thrust

propulsion system. The vehicle would also utiliTe a high thrust propulsion system which would fire

opportunistically to yield shorter trip times than would be possible with a completely low thrust

vehicle. The design of a hybrid vehicle presents several design challenges-- many of which will be

met with the design presented in this report.

Introduction 1
1990LOCOSTSeniorDesignProject

Among the high thrust propulsion devices, only the chemical rocket provides any possibility of

utilization. Nuclear propulsion units will not be sufficiently developed for use by the year of technology assumed by this report.

Electric propulsion is conceivable but a large mass penalty is imposed because an electric propulsion vehicle must carry its own electrical power generation equipment. Laser propulsion, however,

does not suffer from this drawback. With a laser propulsion unit, a continuously tracked laser

power beam is received by the vehicle. The laser beam is converted into thrust by heating a

propellant. The propellant heats and expands, then exits a nozzle providing thrust. Laser propulsion has been sufficiently studied so its utilization is feasible.

The hybrid laser-chemical concept does not imply that the two systems are coupled together

physically. They are instead used opportunistically whenever one device offers substantial advantages over the other for a given orbital maneuver. The two types of propulsion devices can be fired

independently, although not simultaneously.

Basic requirements and assumptions for the LOCOST vehicle are listed in Table 1.

Table 1. Basic Requirements and Assumptions

Orbit Transfer Time 2-3 weeks

Cargo Mass 20,000 - 40,000 kg

Laser Type Direct Solar Pumped Iodide

Laser Wavelength ~ 1.315 micrometers

Laser Power I 0-12 MW

Isp Chemical Rocket 480 seconds

Isp Laser Rocket 1500 seconds

Level of Technology 2010

It is assumed that the SS is capable of assembling the LOCOST vehicle after it is lifted from earth.

It is to be lifted to orbit with no more than four (4) shuttle sized launches. Also, it is assumed that

the Laser Power Station (LPS) and the Energy Relay Units (ERU) will be firmly established in

their orbits ready for operation.

The single LPS will provide all of the laser power for the LOCOST vehicle during the duration of

a mission. There axe two ERUs; one each positioned 120 ° lead and 120" lag. The vehicle can

operate without the ERUs, but the ERUs provide a substantial improvement in propellant usage

and trip time if used because they eliminate the line-of-sight problem between the vehicle and the

LPS (see "Orbital Mechanics" on page 17).

The General Mission Scenario

The general mission begins in LEO plane with orbital inclination of 28.6 °. It is assumed that a

cargo booster or the space shuttle has lifted a payload from earth, or that the space station has

payload intended for GEO. The cargo will be transported to the GeoShack. The outbound

mission will assume a cargo mass of 20,000 kg. The LOCOST vehicle loads the cargo and initiates

an outbound, spiral trajectory characteristic of a low thrust propulsion device. The laser rocket

will be the only propulsion device in operation during the entire orbit transfer. Once the orbit

transfer is complete, the vehicle will recircularize using the chemical rocket at GEO, then instigate

a plane change. The plane change will use the chemical rocket. The plane change is performed

at GEO because the AV requirement for the vehicle will be smaller than that needed at LEO, realizing significant propellant savings. After docking with the GeoShack and depositing the payload,

Introduction 2
1990LOCOSTSeniorDesignProject

thevehiclewill reloadwith 6,000 kg of cargo. The vehicle will initiate another plane change, followed by the orbit transfer, and fmaUy end the trip with another recircularization at LEO.

As mentioned in the scenario above, the outbound cargo is half the maximum the LOCOST vehicle

is designed to handle. It is important to analyze the scenario where the maximum cargo out and

back is transported. As seen in the trajectroy analysis chapter, this mission scenario requires a

modification of the truss to accomodate the propellant increase.

Redundancy of the Propulsion System

Because of the redundant propulsion system, the LOCOST vehicle can suffer the loss of one type

of propulsion device and still complete the mission. For example, it is conceivable that the LPS,

which is susceptible to failure from dangers that the vehicle might not experience, could fail during

the mission. The loss of access to the LPS will cause the vehicle's laser propulsion to become

useless. At this point, the chemical propulsion can activate and provide a safe return trip. The

ability to recall the vehicle in the event of a single propulsion failure is important when the payload

must be recovered within a given time constraint.

The redundant propulsion system increases the reliability. Although a reliability analysis is beyond

the scope of this report, it is clear that the hybrid vehicle is more reliable than a vehicle with a single

type of propulsion.

Extended Mission Capability

A hybrid vehicle provides extended mission capability. There axe certain missions that a dedicated

propulsion device vehicle can perform with better performance than a hybrid vehicle. If there axe

no time constraints on the completion of a payload transfer, a completely low thrust vehicle can

perform the mission with greater propellant efficiency. If, however, the payload transfer time is

critical, a high thrust propulsion device is best suited for the transfer.

Introduction 3
1990 LOCOST Senior Design Project

Vehicle Configuration

This section describes the overall LOCOST vehicle configuration. It also presents a design rationale, a short description of the configuration evolution, and a configuration component summary.

Design Rationale

The design of an aerospace vehicle must consider the traditional design criteria. Special features

of a vehicle merit additional study. Traditional design criteria are listed below.

• Low mass

• Low moments of inertia

• Simplicity in design

• Ease of integration of configuration components

• Efficiency in execution of mission

• Reliability

The above criteria must be balanced by a vehicle's special needs. For example, a technologically

advanced propulsion system can increase the vehicle mass and increase the complexity of design.

These concerns are balanced by the increase in overall performance of the vehicle. The advanced

propulsion system could also reduce overa// propellant consumption making the previous objections moot. Such is the case with the LOCOST vehicle.

The components of the vehicle which will provide the greatest design challenge are also the comportents which bend the traditional design criteria. Because the final design will contain both a laser

propulsion device and a chemical propulsion device, several issues arise. These issues are summarized in Table 2.

Table 2. Important Design Issues for the LOCOST Vehicle

• Placement of the laser engines

• Placement of the chemical engines

• Elimination of the interference between the two systems

• Collection of the laser beam for rocket use

• Line of thrust through the center of mass

• Placement of the cargo

• Placement of the fuel tanks

Overall Configuration 4
1990LOCOSTSeniorDesignProject

_upport

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Rocket

bay

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/_ Mirror

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___SupportN

f

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ri_mary -_ \_...Piv o t

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Secondary

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Figure 1. Configuration I

Configuration Evolution

Configuration 1

This configuration can be viewed in Figure I. It featured one laser rocket and one chemical rocket

in a symmetrical design. The laser rocket engine housing was composed of a multi-faceted lens.

This allowed the laser rocket to rotate through various angles but still capture the incident power

beam. Propellant storage was accomplished through six spherical tanks located near the primary

mirror. The primary mirror could rotate about a hinge support by extending a pivot arm. This

design was reiected for two reasons. First, there was not sufficient evidence that a multi facetted

lens configuration on the laser rocket housing would support the high incident power transfers.

Second, as the propellant tanks reduced their mass through propellant consumption, the center of

mass changes. The fixed position of the laser rocket did not allow for any variation of the line-ofthrust.

Configuration 2

The second configuration featured a large rectangular space truss and a non-symmetrical placement

of the primary dish. The primary mirror was mounted on top of the truss with two retractable

poles. Movement of the poles away or towards the main truss moved the primary mirror, The

Overall Configuration 5
1990LOCOSTSeniorDesignProject

mirror was also mounted on two gimbals which allowed the mirror to rotate. Two chemical rockets

were mounted symmetrically about the longitudinal axis of symmetry while a single laser engine

was mounted through the center. The most notable feature of this design was that the laser and

chemical rockets were mounted in a rocket basket. The entire rocket basket moved to compensate

for the change in the center of mass. The laser rocket face was changed to a piano-convex lens.

A multi-facet design was not needed.

This design was rejected for two reasons. First, the movement of the rocket basket (containing all

three rockets) was deemed inefficient and unnecessary. Second, the motion of the primary mirror

created innumerable difficulties with designing an optical system capable of transmitting the beam

to the laser rocket housing.

Selected Configuration

The selected configuration eliminated many of the problems associated with configuration 2. The

primary mirror is fixed and the chemical rockets are gimballed. The laser rocket, however, continues to move in a rocket basket for reasons which will be discussed. Figure 2, Figure 3, and

Figure 4 show the selected configuration for the LOCOST vehicle.

Discussion of the Selected Configuration

The following sections will discuss the main components of the selected configuration.

and Figure 3 point out the main features of the final configuration.

Figure 2

Main Truss

The main truss is 29.11 m in length, 6.3 m deep, and 18.9 m wide. It is a rectangular frame and it

is constructed with 6.3 m truss boxes. The mirror support structure extends 9.0 m up from the top

plane of the vehicle. Individual truss members range in length from under 1.0 meters to over 9.0

meters. The struts are made from graphite epoxy with a cylindrical shell design. The outer diameters are 5.4 cm and they have a wall thickness of 0.16 cm. The members are assembled using

titanium joint fittings. The total dry-mass which excludes all components except the struts and

nodes is 1730 kg.

Propulsion

Chemical Propulsion

A bi.propellant chemical rocket is used with hydrogen as the fuel and oxygen as the oxidizer. The

final configuration uses two rockets. Each chemical rocket is capable of delivering 25,000 N of

thrust. A rocket analysis program NOTS was used to calculate the flow in the chamber, throat and

nozzle. The rocket has an area ratio of 800 and a throat diameter of .04 m. The rocket has a

chamber pressure of I0 MPa and a mixture of 6:1. The maximum temperature occurs in the

throat--3380 K. The rocket was designed to have a specific impulse of 480 seconds. The chemical

rockets axe regeneratively cooled with hydrogen fuel. Each rocket has a mass of 427 kg.

Laser Propulsion

The laser energy will be coupled to the thermal energy of the hydrogen propellant via the inverse

Bremsstrahlung process. Particulate absorption by carbon particles will be used to initiate the

plasma upon which the inverse Bremsstralalung process will operate. The laser engine is 56 cm in

diameter and 1.25 m long. It will consist of a piano-convex sapphire window lens which will converge the laser beam within the engine chamber. Another main feature of the laser engine is the

mounting collar that allows the engine to be mounted inside the mobile rocket basket by

graphite-epoxy struts.

Overall Configuration 6
1990 LOCOST Senior Design Project

18.9

---- 6.3

CARGO

CONTAINER

OXYGEN_

TANK

RADIATION

SHIELD

SECONDARY.I

MIRROR

PRrHARY

HEAT MIRROR

PIPE CONPLEX

PIVOT

32.61 m

PRrNARY

_IIRROR

HYDROGEN

TANK

CHEMICAL

ROCKET

LASER

ROCKET

F

-- 6.3 m-- _

LASER

\ROCKET

POSITIONING

MECHANISM

Figure 2. Selected Configuration-Top View

Overall Configuration 7
1990 LOCOST Senior Design Project

1.] m--

/

SECONDARY

MIRROR

PRIMARY

PRIMARY/

MIRROR _

PIVOT

I"

5.7 m

/

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Figure 3. Selected Configuration-Side View

Overall Configuration 8
1990 LOCOST Senior Design Project

14.0

Figure 4.

ALL DIMENSIONS IN

Selected Configuration-Rear View

METERS

Overall Configuration 9
1990 LOCOST Senior Design Project

The rocket basket moves to adjust the line-of-thrust so that it always passes through the center of

mass. The laser rocket is situated along the longitudinal plane of symmetry of the vehicle.

Propellant Storage

The oxidizer is stored in two spherical tanks and the hydrogen fuel is stored in two cylindrical tanks.

The tanks are symmetrically placed.

The tank skins axe made out of an aluminum alloy. The mass of the tank skins axe 6 and 129 kg

respectively for the liquid oxygen and the liquid hydrogen. Internal pressure is maintained at 35

kPa. The tank interior supports slosh baffles and a "flow channel" type liquid acquisition device.

The tanks are covered with two types of insulation. Thin, alternating layers of metallized Kapton

Rim and Dacron net spacers (MLI) are used and an outer coating of Rohacell 31 foam is applied.

A thin layer of reflective paint, YB-71, is applied to the foam. The mass of the insulation for each

type of tank is 47 and 571 kg respectively for the liquid oxygen and hydrogen tanks.

Optical System

The optical system is composed of five mirrors, a cooling system and an opto-electronic feedback

system. The optical system will be described in detail in a later chapter. The following discussion

offers an overview of the optical system components.

Optical Train

The optical train is composed of five mirrors: a primary mirror, a secondary mirror, a third and

fourth mirror, and a final fifth mirror. The purpose of these mirrors is to collect the incident laser

beam and direct it to the laser rocket. The mirrors are labelled in the order that they are used. The

primary mirror is the largest mirror, 14 m diameter, and collects the beam slightly off-center. The

primary minor is mounted on two gimbals. These gimbals allow the mirror to rotate. The primary

mirror focuses the beam to the secondary mirror. The secondary mirror colimates the beam (to a

diameter of 1.1 m) and directs the beam towards a hole in the primary mirror. After passing

through this hole, the beam strikes a fixed elliptical third mirror and immediately reflects the beam

onto a fourth mirror. The fourth mirror can rotate about its axis. This feature gives the primary

mirror the ability to swivel about its gimbals and still direct the beam to the laser rocket. After

striking the fourth mirror, the beam strikes the fifth mirror. The fifth mirror moves along a track

in concert with the rocket basket.

The secondary, third, fourth, and fifth mirrors require cooling due to the intensity of the power

beam. The primary mirror has a low incident intensity therefore does not require cooling. The

secondary, third, and fourth mirrors axe cooled using a heat-pipe system. The heat pipes connect

to a primary radiator located around the edge of the primary mirror behind a radiation shield. The

fifth mirror is cooled with hydrogen fuel during operation of the laser rocket.

Opto-Electronic Feedback System

Continuous, precise alignment of the optical train during powered flight is difficult to achieve. An

opto-electronic feedback system has been designed to fine tune the optical system. The system is

primarily composed of a laser tracking communications system and optical feedback sensors. The

laser tracking communications system, dubbed the ATP (Acquisition, Tracking and Pointing system), uses a lasercom-like system to continuously track the incident laser beam. The lasercom

equipment is mounted on a boom located at one of the gimbal mounts of the primary mirror. A

laser system must be used in order to achieve precision tracking.

Avalanche photodiodes (APD) with dielectric fdters have been placed around the periphery of most

components of the optical train. These APDs provide the feedback sensor capability required for

the close-looped system.

Overall Configuration I0
1990LOCOSTSeniorDesignProject

14

Primary Mirror-Front View

R.

8.89

I

17 78

15.4

Figure 5.

\

8.75

,¥

1.5

J

ALL DIMENSIONS IN METERS

Primary Collection Dish-Side View

Overall Configuration 11
1990LOCOSTSeniorDesignProject

INCOMING

LASER

Figure 6. Incident Laser Beam Path

Overall Configuration 12
1990LOCOSTSeniorDesignProject

_....._ Omn_directlonal Antenna

Astro Star

1"racKer

Antenna

Boom Arm

Dlre tlonal Antenna

Figure 7. Location of LaserCom Equipment

Payload Module

The payload module is composed of a magnesium alloy cylinder which holds the payload during

the duration of the trip and a supporting space truss. The payload module is mission flexible.

Either the payload canister or the space truss may detach from the main truss. Special node connectors were designed to allow the space truss to snap together. This makes removing the payload

truss simple, and allows for tele-robotie operators to remove the payload at the GeoShack. The

mass of the entire payload module, including the space truss is 1550 kg.

Attitude Control

Attitude control is partially accomplished by four advanced control moment gyroscopes (CMG).

Attitude determination is achieved by fiber optic gyroscopes and charge-coupled star trackers.

Rocket Control System (RCS)

The rocket control system is used for desaturization of the CMGs, attitude and docking maneuvers.

The RCS system is composed of eight sets of three nozzles located at the comers of the main truss.

Overall Configuration 13
1990LOCOSTSeniorDesign Project

The RCS rockets are hydrazine/oxygen powered chemical rockets. Storage of the hydrazine will

be accomplished by locating small tanks near each RCS cluster. The mass for the total system is

1600 kg.

Communications System

The communications system is used for telemetry, control, and data acquisition. The main communications system uses a RF wavelength transmitter and receiver. One omni-directional antenna

and two directional antennas are required to start the power beam acquisition sequence. The mass

of the total system is 200 kg.

Electrical Power

Electrical power is provided by fuel cells which use oxygen and hydrogen in a chemical reaction to

produce electrical energy. The oxygen/hydrogen fuel scheme was selected because of the availability

of those elements aboard the LOCOST vehicle. The power system is composed of four fuel cells.

Each fuel cell has a mass of 100 kg. The undesirable by-product of the fuel cell chemical reaction,

water, will be vented into space.

Configuration Analysis and Mass Breakdown

Table 3 summarizes the mass breakdown for a vehicle satisfying the baseline mission.

Overall Configuration 14
1990LOCOSTSeniorDesignProject

Table 3. Mass Breakdown

StructKrss

Main truss

Total

Propulsion

Chemical rockets

Fuel Lines (Chemical)

Chemical mounting

Laser Rocket

Fuel Lines

Total

Optical System

Primary mirror

Support for primary

Secondary mirror

Mirrors 3,4,5

Total

RCSs & CMGs

Docking

Communications

Payload Module

Vehicle Dry Mass

Outbound (with cargo)

Inbound (with cargo)

Propellant usage for baseline mission.

Propellant

02

H2

Total

423

423

422

150

4

75

75

726

2,075

35

90

300

2,522

1,100

50

110

1,100

6,682

26,605

12,605

38,300

53,928

92,228

Outbound Total 118,833

Table 4 presents key characteristics of the propulsion system.

Overall Configuration 15
1990LOCOSTSeniorDesignProject

Table 4. Main Engines

Laser Rocket

Isp 1500 seconds

T=u 768 N

mdot=u .0522 kg/s

Chemical Rocket

Isp 480 seconds

T==, 25,000 N per engine

mdotm,_ 5.2 kg/s per engine

Table 5 summarizes the CG location for different conditions.

Table 5. Center of Mass Locations

Condition x (m) y (m) z (m)

No payload, 13.44 9.45 11.05

No propellant

No payload, 10.85 9.45 3.467

Full propellant

6,000 cargo, 19.44 9.45 6.24

No propellant

20,000 cargo, 13.00 9.45 3.413

Full propellant

40,000 cargo, 14.4 9.45 3.37

Full propellant

Overall Configuration 16
1990 LOCOST Senior Design Project

Orbital Mechanics

Introduction

Program OM was developed to evaluate different trajectories that involve space vehicles with two

disparate, primary types of propulsion. In this chapter, the evaluation of the baseline mission and

an evaluation of the maximum cargo trajectories will be discussed.

Mission Assumptions

The LOCOST vehicle will rendezvous with the space station (orbital inclination of 28.6"), then

transport the cargo to a GeoShack or a satellite cargo depot at GEO (0" inclination). The baseline

mission assumes two Energy Relay Units (ERU), but a non-ERU mission will also be considered.

One laser station cannot provide continuous power to the vehicle because of, station-solar

occultation and earth blockage (Figure 8).

Station-solar occultation occurs for 16.7 % of the laser station orbit. Earth blockage times are

dependent on the specific trajectory and thrusting rules. A large percentage of the LOCOST vehicle

in-flight time is spent on coasting.

Earth blockage occurs when

Rearth

R<

( sin 150" - 6)

where 6 equals the angular separation of the LPS and the LOCOST vehicle. ERUs are orbiting

platforms that relay the power beam around the earth horizon. In order to completely eliminate

the earth blockage problem, two ERUs axe needed. Each ERU is located in the plane of the LPS.

One ERU is positioned 120" behind the LPS and another 120 ° ahead.

Most low-thrust propulsion devices axe power limited, and hence, when the mision requirements

force the vehicle to conform to specific constraints, power-limited propulsion techniques must be

supplemented by other more powerful, high thrust devices. The purpose of this section is to explain

the strengths of a laser/chemical hybrid, how the hybrid vehicle utilizes the strengths of both types

of propulsion devices and to justify the hybrid's strength based on mission analysis.

Program OM

Program OM was developed to analyze the trajectory of a vehicle with more than one type of

propulsion device. The effects of earth blockage, station-solar occultation, and ERU placement

were coded in the form of firing rules. A mission scenario can be analyzed by linking together several flight modules, each of which was given a special maneuver number. Information about

payload mass and propellant starting parameters was obtained from an input file. A second input

file described instantaneous orbit sampling parameters and other run parameters which helped free

tune the numerical routines.

Program OM used functions from the IMSL library to integrate the equations of motion. A figure

giving some of the geometrical parameters is pictured in Figure 9. From this figure, equations of

Overall Configuration 17
1990 LOCOST Senior Design Project

Laser Station

J

\\

veh

To _ Sun Laser tation

Shadow _

Region [

Figure 8. Station-solar Occultation and Earth Blockage

motion were developed. Additional equations of motion were added to the set for the LPS and the

ERUs.

#= vCOSy

/'= v Siny

_t_ m

Tcos/_ /_sin?

m r2

Tsin fl

Yffi my

COS y

; (v 2/'- _lr _)

Where

Theta is the orbit path increment angle

v is the velocity vector

y is the angle between the local horizontal and velocity vector

Orbital Mechanics 18
r is the radius vector

T is the thrust vector

/_ is the angle between the thrust and velocity vector

m is the mass of the orbiting object

1990 LOCOST Senior Design Project

Space

Vehicle

Earth

Figure 9. Basic Geometrical Description of Vehicle Vectors

A Runge-Kutta sixth order method was used to integrate the equations used in program OM.

Subroutine IVPRT from the IMSL library was used. The time step used for this integration was

part of an input file. Runs were made with time steps and sampling parameters on the order of 30

seconds to 10 minutes.

Output was collected in several files. Separate output files for plotting, sampling, logging and maneuver statistics were generated.

Station Solar Occultation.

Figure 8 shows station solar occultation. Station solar occultation is a function of the LPS's angular velocity and the starting position of the LPS. The occultation figure shows the range of angles, for the orbiting altitude of the LPS, that the sun can reach the LPS. From the simple relation

OLp S = 0o + wl_.pst

it can be shown that the LPS is in the earth's shadow 16.67% of any trip TOF.

Because it is assumed that there is only one LPS, when station solar occultation occurs, it is impossible to fire the laser. During these times, the vehicle coasts.

The Baseline Mission

Table 6 shows some additional assumptions that were used when performing the trajectory analysis.

Orbital Mechanics 19
1990 LOCOST Senior Design Project

Table 6. Baseline Mission Assumptions

• Two ERU's are used, therefore, only sun occultation will prevent laser usage.

• Laser propulsion is used only for orbit transfers.

• Chemical propulsion is used only for recircularization and plane changes.

• The outbound cargo is 20,000 kg.

• The return cargo is 6,000 kg.

Using these assumptions, the baseline scenario was computed. It is presented in Table 7.

Table 7. Baseline Scenario Trajectory Analysis Results

Flight Segment AV Time Propellant Used

(kin/s) (days) (kg)

Transfer 5.16 9.63 36,393

Recirc. .00575 0.0011 1,018

Plane Ch. 1.52 0.025 22,879

Plane Ch. 1.52 0.0014 12,707

Transfer 3.17 1.74 6,522

Recirc. 1.52 0.008 7,414

Total Time 11.4 + 0.6 (Cargo Transfer) -- 12 days

Total Propellant's Used 87,221 kg

Laser Propellant 43,202 kg

Chemical Propellant 44,020 kg

For a 6: I oxidation ratio

LO2 propellant 38,310 kg

LH2 propellant 54,000 kg

Maximum acceleration .1 g's

Thrusting Times

Laser 83%

Chemical .4%

Coasting 17%

Residual Fuel ---3,255 kg (3.5% residual)

A noticeable feature about the baseline scenario is the coasting time. The coasting time in this

scenario is the lowest for all cases calculated. This figure indicates that the LPS was completely

utilized. If the traffic between LEO and GEO increases, more vehicles will be needed or less power

per vehicle and longer coasting times will result. This report assumes one dedicated LPS.

Table 8 shows various mass ratios characteristic of the vehicle.

Orbital Mechanics 20
1990LOCOSTSeniorDesignProject

Table 8. Mas_ Ratios

payload

propellant

0.28

propellant

vehicledrymass

11.6

propellant

1 -- mass

0.92

The Maximum Cargo Scenario

The maximum cargo scenario involves moving the maximum amount of cargo to GEO and back.

The cargo would be 40,000 kg out/40,000 kg back. It assumes two ERUs are in use.

Using program OM to analyze this scenario produces some interesting results. In order to move

the maximum amount of cargo in the manner described, the TOF increases as well as the propellant

consumption. Table 9 summarizes the extreme case.

Table 9. The Maximum Cargo Scenario--40,000 kg Out and Back

Total Trip Time 24 days (13 out/l 1 in)

Total Propellant Used 225,000 kg

Residual 3.5%

Performance ratios are presented in Table 10.

Table 10. Maximum Cargo Scenario-Performance Ratios

propellant/LOCOST dry mass = 28.1

total payload/propellant = .36

Even though the trip time doubled, the amount of propellant needed for given unit mass was lower

for the baseline scenario. The maximum cargo scenario provides better propellant efficiency at a

cost of time flexibility.

In order to accommodate the extra propellant, tanks must be added to the LOCOST vehicle. The

tanks attach towards the rear of the vehicle, and are placed in unit cells on the outer edges of the

truss. The section on structures will give specifics.

Other Scenarios

To understand the effect that the ERU usage assumption has on trip time and propellant consumption, another scenario was computed. In this scenario, whenever station solar occultation

occurrs, the chemical rocket is fired for ten minutes at 50% maximum thrust. Total propellants

Orbital Mechanics 21
1990LOCOSTSeniorDesignProject

consumedjump to 275,000kgandtrip timeincreasesdramaticallyto 32days. Payloadloadsare

maximum.

Whileit takes34timesthevehicle'sdrymassin propellantto movethe 80,000 kg total of payload,

the payload over propellant ratio remains nearly the same as that for the baseline mission, in this

case 0.29.

Conclusions

It was found that the assumption that there are ERUs in use can lead to significant time savings

over a non-ERU scenario. While the ERUs do not provide drastically superior payload to

propellant ratios, the ERU assumption introduces greater mission flexibility in the form of scheduling and diversity of payload transfer. The existence of ERUs is critical to the longterm logistical

support of a SS or GeoShack dependent on laser based vehicles. It is suggested that ERUs be included in any such endeavors.

From the traiectory analysis, it is apparent that a 40,000 kg cargo presents difficulties for low-thrust

vehicles. While the average mission can expect to carry masses at most half that size, the capability

for the maximum cargo case should exist. A low-thrust all laser vehicle cannot provide the support

for the extreme case. For the baseline case, the LOCOST vehicle yields significant savings in

propellant usage, but not as good as an all laser vehicle. The vehicle provides, however, better time

performance.

Orbital Mechanics 22
1990LOCOSTSeniorDesignProject

Optical System

The optical system provides a direct link between the power source and the laser engine. The objective of the optical system is to intercept the laser beam and redirect and focus the beam into the

thrust chamber for any vehicle attitude orientation.

Optical System Evolution

The main goal of the optical system is to capture and direct the laser beam into the laser rocket

chamber. A two laser rocket design was ftrst proposed. Although it allowed for a symmetrical

system it necessitated the use of a beam splitter and a larger primary mirror. The large increase in

mass of the primary mirror and that the beam splitter may not be able to split the beam exactly (to

provide equal power to the engines) motivated the search for other designs. The final design described below incorporates a one degree of freedom primary mirror and one laser engine.

The primary mirror was limited to designs which provided a very high reflectivity. This eliminated

an inflatable mirror which, although light and easy to deploy, does not have high reflectivities. A

polished metal mirror was investigated, but it, too, cannot achieve a desired reflectivity of over

99.0%. Finally, a dielectrically coated mirror was selected. This type of mirror offers very high

reflectivity along with relatively low mass. Also the dielectrics can be applied to the substrate material on earth and then assembled in space.

Optical Train Geometry

The selected optical train geometry is shown in Figure 10 and Figure 11. Figure 10 shows the

primary mirror assembly from a side view, while Figure 6 shows the path of the laser beam from

the rear and side of the vehicle.

The optical system employed for the vehicle is similar to a true Cassegrain system. A small (r =

.55 m) convex parabolic mirror is mounted on tripod legs extending from the large (r = 7m) concave parabolic mirror. The main mirror was sized to capture the beam at the maximum working

distance of the vehicle. The beam size at a given distance is a function of the diffraction limited

half angle, beam jitter, wave front error, and the separation distance.

S = J(1.3Sa) 2 + (Sj) 2 + (Sw) 2

Where

S is the divergence in radians

Sd is .4138 MD (diffraction limited half angle)

S t is 0.05 x 10-6 (beam jitter)

S,, = _/20 D (wave front error)

2 is 1.315 micrometers (wavelength)

D = 30 m (transmitting mirror diameter)

The spot size of the beam is found using the following relation

D=4SxR

Optical System 23
1990 LOCOST Senior Design Project

1 Secondary

Nlrror

\

\

\

/

Laser

Figure I0. Optical System Configuration Side View

Where R = 52,300 kin, the maximum separation distance. The spot diameter was determined to

be 11.6 meters. The primary mirror was sized up to a diameter of 14 meters to allow for off-center

capture of the beam (to reduce heating load on the secondary mirror housing) and to capture fringes

of the laser beam. This adds to the power captured and reduces radiation degredation on the vehicle

structure.

Optical System 24
1990 LOCOST Senior Design Project

Radiator

Heat

Pipe

Chemical

Ensines

Figure ! !. Optical System Configuration Rear View

Primary Mirror.

The placement of the primary mirror is based on several criteria. First the effects of thermal and

inertial loads on the structural members were considered. By analyzing these loads it was concluded

that the mirror must be placed no more than 10 meters out on the tripod legs (below this maximum

value, the vibrations and stability are well within the limits imposed by the power losses).

Using this maximum value and considering factors such as the desired final beam width and cooling

of the secondary mirror, a focal length of 8 meters was selected for the primary mirror.

The primary mirror can be seen in Figure 12 (viewed perpendicular to the gimbal mountings),

Figure 13 (viewed along the gimbal mountings), and Figure 14 (off isometric view).

Optical System 25
1990 LOCOST Senior Design Project

Figure 12. Primary Mirror View I

Optical System 26
1990LOCOSTSeniorDesign Project

14m

Figure 13. Primary Mirror View 2

Optical System 27
1990LOCOSTSeniorDesignProject

Secondary and Third Mirror

The primary and secondary mirrors share the same focal point, one-half meter beyond the second

mirror. By arranging the mirrors in this configuration one obtains a colimated laser beam with a

reduced diameter which is parallel and traveling in the same direction as the incident beam. The

beam travels through the small hole in the primary mirror and is redirected by a planar, rectangular

mirror which is fixed at a 45 degree angle to the common axis.

The secondary mirror is shown in Figure 15, Figure 12, and Figure 13. The combination of the

primary and the secondary mirror can best be seen in Figure 12 and Figure 13.

Fourth Mirror

The whole system described above can rotate with one degree of freedom about two pivot points

on either side of the primary mirror (Figure 12). The different positions of the mirror assemblies

due to rotation can be seen in Figure 17. To allow for this movement, the fourth mirror, a planar,

elliptical mirror which is also at a 45 degree angle to the incident beam, is able to rotate about its

center axis as shown in the diagrams. By rotating this mirror, the beam is always able to be directed

to the fifth mirror regardless of the position the larger system assumes (Figure 16).

Fifth Mirror

The fifth mirror (Figure 18), a curved elliptical parabolic mirror, has one degree of freedom in the

vertical plane in order to capture the beam regardless of the position of the mirror system and direct

it on a straight line into the laser rocket. This mirror focuses the beam down to the size of the

window opening in the laser rocket engine and then the window focuses the beam to a point in the

laser engine.

To determine the surfaces of the parabolic mirrors the equation

X 2 + Z 2 = 4P Y

was used with P = the focal point. From the geometry, it was determined that the 5th mirror

needed to focus the beam by 2°.

Mirror Assembly

The assembly of the primary minor is a concern in the design of the optical system. The method

chosen was to divide the mirror up into 12 equal area wedges. The wedges will be "snapped" together once in space. All 12 wedges stacked together--with allowances made for packing

materials--wiU fit into a rectangular box with dimensions 3.6x8x6.5 m. This box will fit into the

cargo bay of the space shuttle.

Optical Train Control

The performance of the optical train is critically important to the performance of the ship as a

whole. Small deflections in the optical train may cause a significant loss of power. For this reason,

almost every optical component will be equipped with electrically controlled feedback sensors

which will monitor the mirror at all times. The feedback system is described in the chapter on acquisition, tracking, and pointing. If the mirror deflections become to great in one direction these

devices will sense this and correct the problem.

A deflection analysis was performed on the small secondary mirror in order to analyze a "normal"

deflection. It was found that a power loss of 9.9 kW corresponds to a deflection of .00098 m from

Optical System 28
1990 LOCOST Senior Design Project

Figure 14. Primary Mirror Off Isometric View

Optical System 29
1990 LOCOST Senior Design Project

_ilOT _1 LILIR

Protective

houoln|

Figure 15. SecondaryMirror

the central axis. This deflection is much less than the deflection anticipated by the structural analysis. Therefore, the control system will most likely only be used for abnormal deflections.

Motors and Power Requirements

Some of the optical components require motors. In order to determine the power needed by these

motors a future technology value of 0.2984 kW per pound of motor weight was used. The primary

mirror uses two small motors at its pivot points in order to move the mirror to any desired position.

These motors together require approximately 150 W. The fourth mirror has a torque motor

mounted on the piece of the truss to which the mirror is attached. This motor requires 20 W to

rotate the mirror. The fifth mirror uses a linear actuator to move on its vertical axis (Figure 19).

This actuator is powered by a small 30 W motor mounted on the side of the actuator. Finally, the

heat pipe system needs 1.4 kW of power for the auxiliary heating units. The total power requirements for the optical system is approximately 1.6 kW.

Dielectrics

The multi-layer dielectric coated mirror was chosen over other types of mirror schemes. First, the

dielectric coatings offers a higher reflectivity (99.9%) with small mass compared to a metal (copper)

mirror. Also, the dielectrics are becoming easier to assemble in space, and the degradation over time

is reduced. Also, new techniques axe making the application of the dielectrics easier and more

precise.

Beryllium was chosen as the facesheet or substrate material onto which the dielectric coatings will

be applied. Beryllium was chosen over silica and other materials because of its many attractive

properties. These include, a low density which translates into a low mass per unit area, good

Optical System 30
1990LOCOSTSeniorDesignProject

T

Laser From 2

Figure 16. Third and Fourth Mirrors

thermal conductivity which makes cooling easier and good manufacturing tolerances which is important in applying the dielectric layers. The dielectric layers themselves were chosen to be silicon

and magnesium oxide. The absorption for this pair in the laser beams wave length range is about

.03%. This low absorption greatly reduces the cooling requirements for the mirrors. The

dielectric layers are applied to the substrate in layers of alternating high and low refractive indexes

with each layer .328 micrometers or 1/4 wavelength in thickness. The ftrst layer, silicon, is applied

directly to the beryllium facesheet. Then alternating layers of silicon and the low index material,

magnesium oxide (MgO), are applied. To achieve 99.9% reflectance a total of seven layers of

dielectric materials must be applied to the facesheet. The odd number of layers becomes necessary

because, in order for the reflections to be in phase, the high index material needs to be last and i'n'st

since its index is greater than beryllium. On top of the dielectrics, an outer protective layer, constructed from an oxide material, will also be added to provide environmental stabiLity, thus increasing the useful life of the dielectrics. This scheme can be seen in Figure 20.

Optical System 31
1990 LOCOST Senior Design Project

Figure 17. Different Positions Due to Rotations of Primary Mirror Assembly

Optical System 32
1990 LOCOST Senior Design Project

•------ 1.1 ---_t

i'. :.:.::: " :.:.':-"

Figure 18. Filth Mirror

-Two iaya_" altwn6tkflq diellKtr|c _ / $i

-_L.dLyer t.htr.knesl .0_ atcrDmmtwS

Protective

outer

layer

si Si Sx _i

/

// //'

-- /

j °-

/" /

Be /

/o / •

/

/

/.' /

Figure 20. Dielectric LayeringScheme

Optical System 33
1990 LOCOST Senior Design Project

BALL BEARINGS

40TOR

LOCKING NUTS

ROTATING CAM

_IXEO CAM

Figure 19. Linear Actuator Used to Move Fifth Mirror

A few areas of improvement for the dielectrics are anticipated in the future. This includes better

compensation for strain, improved reflectance and increased useful life. In order to accomplish this

the electric field inside the layers must be controlled. By deviating from the standard 1/4 wavelength

design this can be accomplished. If possible, the maximum electric field should be positioned in

the middle of the low index material (MgO). Also, the electric field should be minimized at the

interfaces of the dielectric materials. This will reduce the absorption and increase the damage resistance.

Cooling

The cooling of the optical components is a important aspect in designing the optical system because

of the high power of the laser. The first step in this process is to calculate the reflectivity of the

optical components. These are shown in Table I I with both beginning and end-of-life

reflectivities.

Optical System 34
1990 LOCOST Senior Design Project

Table !!. Mirror Reflectivities

Minor BOX EOL

First 99% 97.5%

Second 99.9% 98.4%

Third 99.9% 98.4%

Fourth 99.9% 98.4%

Fifth 99.9% 98.4%

Window 99.9% 98.4%

From this data, the energy of the laser absorbed by each component was calculated. Using the

absorbed power and accounting for the solar radiation in space, the Stefan-Boltzmarm law,

q - eaA T 4, was used to calculate the steady-state temperature of each component. The thermal

gradients were also calculated for the primary mirror. Using the thermal gradients and a thermal

expansion coefficient of 0.04 x 10-7 per *K, the expansion of the front surface of the mirror was

calculated (Appendix A). The expansion was on the order of 10-s% during the three hour cycle.

This does not represent a problem initially but, over time these stresses may cause degradation of

the mirror surfaces.

A maximum temperature of 600 K was chosen for the dielectric surfaces, in order to insure their

stability. So, as Table 12 indicates, the primary mirror does not need auxiliary cooling mainly because of its size and construction.

Table 12.

Optical Component

Primary Mirror

Secondary Minor

Third Minor

Fourth Minor

Optic_ System'Thermal An_ysis

Steady State Temp.(K)

430.08

1370.33

1179.76

1279.3

Energy Absorbed (KW)

255.25

161.5

159.8

156.2

Maximum temperature for dielectric surfaces is 600 K

However, the other components do need to be cooled. The design chosen specifies that the mirror

closest to the engine and the engine window to be cooled by using the hydrogen fuel. This will

eliminate the need of using heat pipes in the 5th mirror. The other mirrors will be cooled using a

heat pipe system. A variable conductance heat pipe with an arterial wick (stainless steel) and active

electrical feedback control was chosen (Figure 21).

The feedback design was chosen because of its ability to maintain a device mounted at the

evaporator at a near constant temperature, independent of the power being generated by the device.

The arterial wick is necessary for heat pipes in space because it provides a low pressure path for

transporting a liquid. Other notable features of this heat pipe are: a heat source control, easy adjustment, minimum storage volume, and no moving parts. The only drawback to this system is

that it needs auxiliary electrical power. This design also offers good reliability as it has been used

on satellites in the past.

The first step in designing the heat pipe system was to determine the total amount of cooling capability needed. Considering all three mirrors to be cooled this value is 477.5 kW. Using this and

Stephan-Boltzman law with _= .85, the minimum area of the radiator, 76.5 m 2 , is obtained (see

Appendix A). The radiator is placed around the primary mirror and radiates away from the mirror

surface (Figure 22). A heat pipe cross-sectional radius of .05 m was calculated as the radius needed

to facilitate the amount of cooling which was needed. The choice of a working fluid for the heat

pipe system involved checking many different figures of merit. The two fluids that work well at

high temperatures (1200 *K) are Li and Na. Li was chosen because it performs better in the heat

transport category and it will not boil as quickly (see Appendix A). The fluid flow of the heat pipe

can be seen on the mirror in Figure 22 and Figure 23.

Optical System 35
1990 LOCOST Senior Design Project

Table 13. Optical Train Masses

Material Mass (kg)

Beryllium 577.2

Dielectrics 1.447

Gr/Ep 1497.0

(honeycomb)

Total Mirror Mass 2075.6

Truss Structure:

Primary 3_lirror Truss:

Tripod r/_mbers:

TOTAL PRIMARY MIRROR MASS 2132.7

Secondary MirrorMass 120.0

Mirrors3&4Mass 200.0

Mirror 5 100.0

TOTAL OPTICAL SYSTEM MASS 2593

WICK gAS

RADIAT02

FIONT

RADIATION SHIELD

HEAT SOURCI P

TO AUX.

HEATER

HEATER CONTROL

Figure 21. Feedback Control System

Optical System 36
1990 LOCOST Senior Design Project

Figure 22. Radiator Placement and Heat Pipe Front View

Optical System 37
1990 LOCOST Senior Design Project

Radiator

Heat

Pipes

15m

eat

Pipes

Cooling and

Radiator Diagram

Figure 23. Radiator and Heat Pipe Placement Side View

Optical System 38
1990 LOCOST Senior Design Project

Propulsion

Laser Propulsion

Introduction

The basic concept of laser powered propulsion is to transfer laser energy directly to the hydrogen

propellant, thereby eliminating the need of carrying heavy oxidizers. There axe four main methods

of accomplishing this transfer. They are solid heat exchangers, molecular resonance absorption,

particulate absorption, and the inverse Bremsstrahlung process.

Methods of Energy Transfer

The solid heat exchanger utiliTes an absorber cavity to collect the focused laser energy and then

transfer the energy to the propellant through a heat exchanger. Maximum operating temperatures

of the engine limit the performance to well below the nominal 1500 second specific impulse assumed for this study. Thus this method of energy transfer was dropped from consideration.

In molecular resonance absorption, the hydrogen propellant is seeded with molecules. When these

molecules absorb laser light, they undergo a transition from their ground state to a higher

vibrational-rotational energy state. Molecular resonance absorption is very wavelength and seedant

sensitive, and not enough research has been completed to justify near-term application of this

method of energy transfer.

In the particulate absorption process, particles opaque to the laser light axe injected into the rocket

chamber. These particles absorb the laser energy, heat up, and form a plasma. They then transfer

their energy to the propellant by conduction and convection. The particulate absorption process

was chosen as the scheme to initiate the plasma at the beginning of each firing sequence of the

rocket.

In the inverse Bremsstrahlung process, photons from the laser beam transfer their energy through

inelastic collisions with the electrons, atoms, molecules, and ions of the plasma. In this process,

the plasma is not self starting and therefore a scheme to initiate the plasma is needed. The inverse

Bremsstrahlung process was chosen to be the main method of energy transfer for this vehicle.

Particulate absorption using carbon particles is chosen to initiate the plasma at the beginning of

each firing sequence. Once the plasma is initiated, the carbon additives will be gradually decreased

until the plasma is capable of sustaining itself by the inverse Bremsstrahlung.

Assumptions

Degradation effects of the reflective surfaces were considered to account for an approximate 6%

power loss through the optical train. Previous studies (Ref. Frisbee) suggest that a 1500 second

specific impulse is a realistic value for the laser engine. They also assume a nominal thermal conversion efficiency of 50%. The basic assumptions upon which the laser engine parameters were

established are listed in Table 14.

Optical System 39
1990LOCOSTSeniorDesignProject

Table 14. Basic Assumptions

Power of incoming laser (P)

Iodide laser wavelength

Optical train efficiency (OE)

Thermal conversion efficiency (TE)

Specific impulse (I)

12 MW

1.315 micrometers

94.14 %

50 %

1500 sec

Laser Propulsion Parameters

The thrust of the engine was determined from the energy equation

2P x OE x TE

T=

Ige

From this, the maximum thrust was calculated to be 768 Newtons.

The hydrogen mass flow rate is

rh= T =.0522kg/sec

lge

On the basis of the calculations performed in the aerospace department at Virginia Tech, increasing

the chamber pressure from Iatm to 10 atm leads to an increase in the absorption coefficient from

.5/cm to 23/cm. This higher absorption coefficient will allow for full energy absorption within a

shorter distance in the plasma and therefore a smaller plasma size. For this reason a chamber

pressure of 10 atm was chosen. To obtain a specific impulse of 1500 seconds with a chamber

pressure of 10 atm, the chamber temperature was established to be 4400 K. The nozzle area ratio

was set at 80.

Propulsion 40
1990LOCOSTSeniorDesign Project

. / LASTJ STA[ ,EO --\\ COOLING JACKET

x..___ H'fOP,06t ,I

INJECTION RiNG

Figure 24. Laser Rocket Engine

The frontofthe laserenginewillconsistof a plano-convexwindow lensmade of sapphire,which

was chosen becauseofitshightransrnissivity and strengthathigh temperatures.The window will

bc anti-reflectively coated on the outside,to allow formaximum transmitionof the laserbeam.

The window willalsobe reflectively coatedon the inside,to keep theradiationemittedfrom the

high temperatureplasma insidethe chamber. The spot sizeof the incidentlaserbeam on the

window was cstablishcdfrom a desireto keep the power densitybelow 25 kW/cm 2,which isthe

power densityat which the lasersupportedcombustion wave of the inverseBrcmmsstrahlung

processiscapableofbeingsustained(Rcf.Ffisbee).This allowsforlessdegradationofthewindow

surfaceand keeps theplasma from formingnext to the window. A spot sizeof 30 cm diamctcr

was determinedto be sufficiently large.The chamber diameterwas thenenlargedto35crntoallow

forsmalloffaxisdisturbancesof the laserbeam. The window lenshas a focallengthof 45 cm.

The largefocallengthadds to greaterplasma stability.

The throat characteristics were determined from a 1- dimensional model of the flowfield in the engine having a thermal efficiency of 50%, and producing a specific impulse of 1500 seconds. The

main laser engine characteristics are summarized in Table 15.

Propulsion 41
1990LOCOSTSeniorDesignProject

Table 15. Laser Engine Parameters

Laser power entering the engine

Thrust

Hydrogen mass flow rate

Chamber pressure

Nozzle inlet stagnation temperature

Average chamber flow velocity

Chamber diameter

Throat diameter

Nozzle area ratio

Window lens focal length

Overall engine diameter

Overall engine length

Total engine mass

11.3 MW

768 N

0.0522 kg/sec

10 atm

4400 K

20 m/see

35 cm

3.2 cm

80

45 cm

55era

1.25 m

150 kg

The hydrogen plasma has a core temperature of 20000 K. This leads to large radiative heat transfer

to the chamber walls, which must be cooled to maintain engine integrity. The engine will be cooled

by circulating the hydrogen propellant around it. After the hydrogen leaves the fifth mirror it will

proceed to the laser engine where it will enter a cooling jacket surrounding the engine. One fourth

of the total hydrogen mass flow will enter at the front of the engine to directly cool the window.

The rest of the hydrogen will circulate around the nozzle and chamber before being injected into

the chamber. The fuel will enter at the window port with a maximum temperature of 500 K and

enter the chamber with a maximum temperature of 2300 K. The rocket basket is movable, therefore the hydrogen must flow through a flexible hose and not a rigid conduit. The flexible hose is

attached to the adjacent strut by a spring. This keeps tension on the fuel line and keeps it out of

the way of the incoming laser beam. A schematic of the engine mount is shown in Figure 25.

Propulsion 42
1990 LOCOST Senior Design Project

I

£

il

//

i •

n

\ /

/

H

II

I

L

m

Figure 25. Laser RocketEngine Mounting

Chemical Propulsion

Introduction

The fu-st consideration in designing a chemical propulsion system is choosing a propellant. A bipropellant scheme was chosen with liquid oxygen as the oxidizer and liquid hydrogen as the fuel.

These were chosen because they offer high lsp, chemical stability, good pump properties and well

proven performance in general.

Chemical Propulsion Parameters

Based on the mission specifications, a thrust of 25000 Newtons per engine with two engines located

symmetrically on the vehicle was chosen. To calculate the properties of the flow in each engine,

the computer code "NOTS" was used. The required input was the entering enthalpies of the

reactants, the area ratio, the chamber pressure, and the oxidizer fuel ratio.

Propulsion 43
1990LOCOST Senior Design Project

Table 16. Chemical Propulsion-NOTS Input

H (O_) -4300.00 kcal/kg-mole

H (H2) -212.46 kcal/kg-mole

Area Ratio 800

Chamber P 10.377 MPa

MR 6:1

The output of the program provides a detailed analysis of the flow in the chamber, throat, and

nozzle of the engine. It uses a constant pressure heating then an isentropic expansion to the exit.

The results are shown in Table 17.

Table 17. Chemical Propulslon-NOTS Results

Exit Throat Chamber

Temperature (K) 702.43 3382.6 3602.1

Pressure (Pa) 507 5.7x 106 10.377x l0 s

Velocity (m/s) 4840 1616.1 91

Molecular Weight 14.1 13.6 13.4

K (Cp/Cv) 1.261 1.197 1.198

Max Thrust in Gs .51 g

Mass Flow Rate 5.17 kg/s

Combustion Products

Mole Percent of Gas

Name Formula Chamber Throat Exit

Hydoxyl HO 4.6645 3.5238 0

Dydroperoxyl H02 .0028 .0013 0

Water H20 65.954 68.4085 75.6

Hydrogen Peroxide H202 .0017 .0008 0

Monatornic Oxygen O .3982 .2536 0

Monatomic Hydrogen H 3.6311 2.9616 0

Diatomic Oxygen 02 .4124 .2779 0

Diatomic Hydrogen H2 24.9354 24.5724 24.4

The specifications of the engine are shown in Table 18. The calculations involved in obtaining the

dimensions of the engine are given in Appendix B.

Table 18. Chemical Propulsion Rocket Geometry

Diameter (m) Length (m)

Throat .0398 .076

Exit I. 126 ....

Nozzle .... 2.03

Chamber .1244 .3731

Total .... 3

Propulsion 44
1990 LOCOST Senior Design Project

_-_. 373m

4' T

124 =

AEBENERATIVELY

[OOLED

COrIBUSTION

CHamBER

J

L

" 2.479 m

/

PADtAT]D),,I COOLED

_17"7

1.12fi

tlI

"1

Figure 26. Chemical Engine Geometry !

Figure 27. Gimbal Actuator

Electric _,

driven pump ,_

22s',22:"/v, Actuator

A schematic of the engine is shown in Figure 26. The gimbal mount for each of the engines will

provide the capability to pivot the engine plus or minus six degrees to help k_p the center of mass

of the vehicle aligned. The motion of the gimbal is provided by a hydraulic actuator which uses

an electric oxygen pump. Some advantages gained through the use of this system are a high frequency response, low power consumption, self containment, and easy checkout. A few potential

problems are possible leakage, temperatur_ effects, and it is comparatively heavy. A schematic of

the actuator used can be seen in Figure 27.

Chemical Rocket Cooling

The engine is cooled using gaseous hydrogen. The engine is based on an expander cycle, which

will be explained in greater detail. In an expander cycle engine, the cycle power depends on the

turbine working fluid energy. Higher coolant exit temperatures are limited by the thermal

conductivity of the wall material and the hot wall temperature of the thrust chamber as well as

coolant velocity. The/42 -//2 regenerator utiliTes energy from the turbine discharge flow to preheat

the chamber coolant to approximately 705 °K providing hydrogen in its gaseous state for cooling.

Propulsion 45
1990 LOCOST Senior Design Project

The chamber regenerative coolant enters the inlet manifold located at the injector face plate and

flows into and through the nontubular combustion chamber to the throat area. The maximum

temperature of the engine is reached at the chamber wall on the interior. Gaseous hydrogen is used

as the coolant because of its high specific heat, good stability, and low vapor pressure. The coolant

passage in the nozzle is a double helix configuration. The average coolant velocity is 4.6 m/s, and

the flow rate varies with the thrust.

Materials

The chamber utiliTes electroformed nickel in the shell and copper- zirconium in the coolant barrier.

The nozzle is constructed of carbon compounds which are highly heat resistant and very strong and

light. The gimbal mount is made of an aluminum alloy. All materials used fall into the safe limit

for performance at the temperatures predicted to occur in the engine. The thickness of the chamber

and nozzle walls are approximately 3 ram, and the throat approximately 4 mm. A breakdown of

the masses of the individual components of the engine are presented in Table 19.

Table 19. Chemical Propulsion Individual Component Masses

Component Mass (kg)

Nozzle 69

Chamber + Cooling 101. I

Screw Jacks + Actuation 20.5

Gimbal Mount 4

Turbopumps 60.7

Heat Exchangers 54.3

Control Valves 54

Miscellaneous 63

Total 426.6

Injectors

The injectors used in this engine are multiple-hole impinging jet injectors, like the ones used in the

S-4 engine. The injector pressure drop of the oxygen injectors is 0.74 MPa and the hydrogen is 0.48

MPa. The number of oxidizer orifices is approximately 432 and the fuel orifices is 668. The final

velocity in the chamber after the propellants have been injected is approximately 91.4 m/s.

Pumps and Turbines

The pumps used in the engine are run off the gaseous hydrogen. To start the pumps and to get

them up to speed, an electric motor is used. Details of how the pumps and turbines are interconnected will be given. The power of the hydrogen pump is approximately 0.11 kW and the oxygen

is 4.7 kW.

The turbines used must achieve very high speeds of about 8376 rad/s. They are also very small,

having a blade tip diameter of 2 cm and a blade root diameter of 1.5 cm. The power of the hydrogen turbine is 287.277 kW and the oxygen is 172.7 kW. The turbine efficiency is 0.8 in both

cases and the pressure ratios are 7 and 3 respectively. The temperatures endured do not go much

above 1500 K.

Expander

In an expander cycle, most of the coolant is fed into a low-pressure- ratio turbine after having

passed through the cooling jacket where it picked up energy. Five to fifteen percent of the coolant

bypasses the turbine and rejoins the turbine exhaust flow before the entire coolant flow is injected

into the engine combustion chamber where it mixes and bums with the oxidizer.

Propulsion 46
1990LOCOSTSeniorDesignProject

The H2enters the engine through a ball type inlet shutoff valve mounted on the inlet of a low

pressure pump that is gear driven from the main oxidizer turbopump shaft. From the low pressure

pump, the fuel enters the first of two back to back shrouded impeller centrifugal stages. 5.8 percent

of the fuel is used as a thrust-piston balancing flow for the high pressure pump. This flow is taken

off at the second stage discharge, circulated to the thrust piston, and then injected back into the

propellant flowpath at the high pressure pump interstage. The flow then moves from the highpressure pump discharge and enters the H_-H2 regenerator which utilizes energy from the turbine

discharge flow to preheat the chamber coolant. The flow is then routed through the turbines to

provide the power to drive the turbopumps, and then through the hot side of the H2 -II2 regenerator. After leaving the regenerator, the turbine bypass flow re-enters the mainstream and is injected

into the throat chamber.

The oxygen enters the engine through an inlet valve similar to the fuel-side inlet valve and then

flows into a low pressure pump. The discharge from the pump enters a single-stage, shrouded,

centrifugal type, high pressure pump driven off a single stage, low reaction, full admission turbine.

The flow continues to the oxidizer control valve, which is preset to give the desired mixture ratio.

From the control valve, the flow enters the injector manifold and is injected into the combustion

chamber.

Propulsion 47
1990 LOCOST SeniorDesign Project

Structures

Introduction

Many factors were taken into consideration before choosing the final structural design. Taken under consideration were the large payload mass and its placement, the amount of propellant needed

for the mission, the placement of the mirror, time of flight, and space environmental effects. Even

with an optimized design, a material that could withstand high stresses for long periods of time and

numerous loadings was needed. The design and material must also resist large deflections to

maintain accuracy in the optical train. Reducing overall structural mass was also an important

design goal.

Truss Structure

The final main truss design is 29.11 m long logitudinally, 6.3 m deep (not including the mirror

support structure), and 18.9 m wide. Individual truss members range in length from under 1.0 m

to approximately 9.0 m with a characteristic length of 6.3 m. The members are made of graphite

epoxy. They are cylindrical shells with a 5.4 can outer diameter and a wall thickness of 0.16 cm.

The vehicle is adaptable to carry extra fuel tanks when a greater than baseline payload must be

transported (Figure 28).

Graphite/Epoxy (Gr/Ep) is one of the leading materials used in the construction of space vehicles.

The composite material chosen was P75S/934 Gr/Ep with a ply orientation of (0,_20,0). Linear

elastic behavior for tension loading and axial compression necessitates parallel arrangement of all

longitudinal fibers to avoid plastic deformation. Compared to other alloys (Table 20), Gr/Ep displayed the highest modula of elasticity and lowest densities. Along with its high stiffness-to-weight

ratio, it has an adequate crushing strength. Gr/Ep does not have a very high transverse strength,

but the truss is designed to take its highest loads in the longitudinal direction, therefore alleviating

the need for a large transverse Young modulus. These are not the only desirable properties of

Gr/Ep. It is highly durable to the space environment with a favorable thermal stability, and it is

easily available.

Table 20.

Ex(Gpa) 276

Ey(GPa) 6.9

Ult.Strength

in tension

long.(MPa) 660

tranv.(MPa) 18.7

Density (kg/m 3)

1165

Possible Materi_s _r Space Truss

Gr/Ep Boron/Al Al Magnesium/Al

230 70 45

160 70 45

570 110 380

380 70 165

2490 2710 1800

Structures 48
1990LOCOST Senior Design Project

\

\

\

t

f

\

\

/ I Nil

IINII

I lPql

\ 1 I/1 II

t

!

\

\

\ /

\/

/\ /

t \

Figure 28. Scheme to Transpo_ the Maximum Cargo

Environmental Effects

Atomic oxygen in LEO causes materials to deteriorate. Due to the fact that the vehicle orbits in

and out of the earth shadow, thermal cycling occurs with temperature changes ranging from 120 to

400 degrees Kelvin (Ref. Dutta). Charged particle radiation and space debris cause material properties to degrade. Gr/Ep has been shown to have a long lifetime in LEO, but to alleviate some of

the effects of the environment and to extend the lifetime of the structure material, a protective

Structures 49
1990LOCOSTSeniorDesignProject

coating is applied to the truss members. Aluminum foil with Ni electroplating, Teflon, and silver

were studied. Chromic acid anodized AI foil appears to be the best choice for the following reasons:

1. High reflectivity and optical tailorabilty.

2. Resistance to atomic oxygen degradation.

3. Adherent to Gr/Ep and easy handling.

4. Affected minimally by abrasion.

5. Easy manufacturability and low cost (Ref. Dursch).

The foil surface is tailored during tube fabrication to increase diffuse reflectance of visible light so

that the truss surfaces are not reflective. This makes working conditions more favorable for the

astronauts during extra vehicular activities while only increasing absorptance to a minor degree.

Joint Design

The node design is similar to standard design used in space structures. The node elements axe made

of titanium alloy (AI6 V4), a material with a strength of the same magnitude as the GrtEp truss

members, therefore ultimate load capability is not reduced by the joint construction.

Table 21. Possible Materials for Joint Design

Material Ex & Ey (GPa) Density(kg/m 3)

Titanium 114 4460

Aluminum 70 2710

Magnesium Alloy 45 1800

The titanium alloy is the only suitable choice for strut and node fitting design. Although using

aluminum or Mg/AI would have resulted in significant mass savings, 39% and 60% respectively,

titanium is the only material that tests up to the loads that the structure needs to withstand. In the

construction of the mirror, however, the loads axe significantly less so aluminum is used.

Some basic requirements in selection and design of the joints are:

• Adaptability to different payloads.

• Attachability to truss members from any angle.

• Maintainability of structural integrity.

The Gr/Ep members are fitted with two titanium end fittings. These are bonded using a cold

hardening adhesive system. The adhesive is applied to the tapered contour of the composite material and inside the mating titanium groove. The adhesive layer thickness is applied uniformly

except at the ends of the overlapping length. Increased adhesive thickness at the ends reduces the

stress peaks at these locations. The stress distributions along the bonded sections are designed to

be uniform (Ref. Franz). This bonded section is 50 mm long (Figure 29). After a destructive test

program was conducted to ensure structural integrity, an ultimate load capacity of 167,000 N was

found. The load capacity was not reduced by applied fatigue or temperature load cycles.

Seventy-five percent of the node elements are designed in an octagonal shape and are used in conneeting 90 and 45 degree angle truss joints (Figure 29). The others must be made for specific angles. The outer nodes have payload attachment planes so that all additional equipment, payload,

and propellant tanks can be mounted in various configurations. The sum of the combination of

the titanium node elements and end fittings added a mass to the main truss main truss of almost

1300 kg.

Structures 50
1990 LOCOST Senior Design Project

Figure 29.

l

Titanium I

Ib_ Connec tioe_ I

Bonded Section o£ Gr/Ep Strut

Shaft

I_yle_ Attachment Plane

/

|Se_ I

'/_/.I// ' " I

-,- IIZ_

?lt_lun

Node Element

J _ Payload

Attachment

__ Plane

/_( Top vie_)

/

Titanium End Node Connector

Structures 51
1990 LOCOST Senior Design Project

Table 22. Main Truss Mass Breakdown

Material Density Cross-section

Gr/Ep 1165 kg/m 3 0.00025 m 2

Strut length Mass # Members Total Mass

4.45 m 1.3 kg 166 215 kg

6.3 1.84 kg 100 184 kg

All Other 51 kg

Total Mass 450 kg

Table 23. Mass Breakdown of the Joints

Material: Titanium

Density: 4460 kg/m 3

Element Mass # members

Node Element 4.5 kg 93

End Fittings 1.5 kg 576

Total mass

Total Mass

419 kg

864 kg

1283 kg

Structural Analysis

With our material selected and its properties known, a structural analysis was performed. The

"Structural Software Analysis for Micros" (SSAM), a finite element computer program, performs

a static analysis of all members of the space structure. Basic assumptions made in applying the truss

to SSAM are as follows:

1. The nodes axe considered as ball and socket joints and can sustain no moment force

2. The structure is statically loaded, so acceleration is constant and in the direction of thrust

3. The nodes connected to the rockets are restrained from motion in any direction

4. Inertial loads due to the payload, fuel tanks, and the mirror are applied at the nodes to which

they are attached (Figure 30).

The program showed that the largest global loads occur in the members that support the rockets.

This is as expected because the members toward the rear must support the load of all the components forward of it. Moving from the rear of the vehicle to the front, loads and stresses became

progressively smaller. The forces and stresses in the rear members were on the order of 103 N and

10_ Pa, respectively. The ultimate longitudinal tensile stress for this material is 0.66 GPa. This

was not the restricting factor. The largest loads and stresses occurred in compression towards the

rear. Using Euler's Buckling equation, C_, := n:EIIL 2, Young's modulus, and the geometric properties of the strut, critical loads for each different length member were computed.

An analysis of the stresses was performed for the maximum acceleration case, 1.9 km/s 2 (during

recircularization). The loads were calculated for full propellant tanks and a cargo mass of 40,000

kg. This is the worst case scenario and will never actually occur. At recirculaxization 40% of the

fuel will already have been consumed which will cut the total mass and inertial loads by 35-40%.

For the baseline mission the actual mass at recirculaxization was approximately 10% of the mass

used for this analysis. Another analysis was performed for the actual recirculaxization case. In this

ease, the acceleration was 1.9 km/s _ and the total instantaneous remaining mass was slightly under

20,000 kg.

Structures 52
1990 LOCOST Senior Design Project

\

\ /

• -" :

.... _ _

AI[/I N _

,le_'

iiii

i i_: i /

::N /

:: -. :\ t

/ \

Figure 30. Inertial Loads on Main Truss

The smallest critical load occurs for the longest members because P= is inversely proportional to

the square of the length. In addition to axial loading, critical stresses must be calculated. SSAM

calculates the normal stress using loads and moments in all orientations to give the maximum

normal stress.

Structures 53
Thermal Analysis

1990 LOCOST Senior Design Project

The temperature of the truss members axe determined by a combination of three possible contributors (Ref. Van Vliet):

1. Absorbed heat from thermal radiation.

2. Heat radiated away by the exterior surfaces.

3. Heat produced by the space vehicle and its components.

The absorbed heat comes from three sources: the sun, the Earth, and the Earth's albedo

(Figure 31). The energy not absorbed by the members could possibly be reradiated towards another member. This reradiated energy must be added to the total heat flux acting upon each

member. The internal temperature determinant is in most cases negligible, tlowever, since our

mission involves the absorptance of a 12 MW laser beam, the heat due to the laser cannot be neglected. The sum of the heat which is absorbed by the primary mirror surface is reradiated out according to the mirror emmittance. Due to conduction, a percentage the radiated heat leaves the

back of the mirror. This will constitute a heat flux almost twice that of the sun.

In order to do an analysis of thermal deflections with SSAM, five simplifying assumptions were

made (Table 24) (Ref. Dursch).

Table 24. Simplifying Assumptions For Thermal Analysis

1. The top (mirror) side is shadowed by the mirror so external incident

radiation is less than on the bottom.

. The mirror radiates heat out of its backside which heats the top side

truss members. The bottom side is almost completely protected from this

heat flux by the fuel tanks, payload, and subsystems mounted inside the

frame.

3. The expansion of the titanium joints is negligible compared to the strut

elongation.

4. The orbits are considered circular.

5. The truss is assembled in LEO where the average temperature is 255 K

Where and at what temperature the structure is assembled influences the stresses and deflections the

members will encounter. If deflections are assumed zero during assembly, then the temperature

change each member undergoes is zeroed around the assembly temperature.

Using simplifying assumptions and the calculations in Appendix C, values for the extreme temperatures that the members would experience were found to be 435 K for the side exposed to the

mirror heat, 395 K for the side exposed only to external radiation, and a cold temperature of 130

K. Applying these thermal loads, the total stresses and deflections that the Gr/Ep struts would

experience due to heating was found.

As observed in Table 25 for the worst case, the maximum stresses are no greater than half the

critical stress in any members. However, when a combined thermal and inertial analysis was performed, the maximum stresses were much closer to the critical stresses. When the inertial loads are

reduced to realistic values and the truss is heated up to its most extreme temperatures, the maximum stresses are much greater. The thermal loads axe of greater concern than the inertial. But the

truss maintains its rigidity in these extreme cases. The highest strut temperatures are reached during

operation of the laser rocket because of the heat contribution from the primary mirror. This is,

however, when the vehicle is under the least inertial loading because the laser engine thrust is 98.5%

less than both the chemical engines combined thrust. Therefore, the maximum stresses encountered

in the combined analysis are greater than the stresses encountered during normal laser thrusting

with maximum truss heating.

Structures 54
1990 LOCOST Senior Design Project

Table 25. Stress Analysis of Loads

Case: Worst with No Thermal Loads

Length(m) o¢.,(MPa) Pmu(MPa)

9.0 11.8 4.5

6.3 23.5 9.3

4.45 47.1 13.6

1.26 568.0 9.5

Case: Actual Recircularization with Thermal Loads

Length(m) o,_,(MPa) P.,_(MPa)

9.0 I1.8 7.3

6.3 23.5 18.9

4.45 47.1 45.6

1.26 568.0 203.1

The deflection analysis with no thermal loads showed that the truss maintains its structural rigidity

with minor deflections. The maximum deflections occurred at the two nodes which connect the

mirror to the main truss. These deflections were 1.86 mm and 1.67 mm. All other deflections were

of the magnitude of 1 mm or less. For the case with thermal loading, the deflections were much

greater. The members not protected from the mirror radiation had a magnitude of deflection on

the order of 10-a meters. This is not enough to disrupt the optical train, though. The maximum

deflections of 3 cm occurred for members attached to the hydrogen fuel tanks which does not affect

the transfer of energy to the laser engine.

Solor

Solar

Radiation t -''/

Satellite thermal balance.

(ref. Goble)

Figure 31. Thermal Contributors

Structures 55
Laser Rocket Basket Design

1990 LOCOST Senior Design Project

The LOCOST vehicle was designed to have the laser rocket thrust through the longitudinal axis.

As fuel is consumed and the longitudinal centroid moves, the laser rocket must move. For this

purpose, multi- directional thrusting rockets are needed. Since the laser rocket cannot be gimballed

like the chemical rockets, another method to align the line-of-thrust with the center of mass was

sought. A moveable rocket basket was designed to meet these criteria (Figure 32).

The longitudinal center of gravity moves a total 3.3 m from one extreme to the other, which is from

full fuel tanks, full payload to empty fuel tanks, no payload. The laser rocket is mounted in a one

meter cubed truss. The truss is moved using linear actuation with rubber friction wheels which roll

along a truss member (Figure 32). The normal force between the wheels and the shaft must be

kept large in order to maintain sufficient friction for movement. The force is generated using high

tension springs connected to the axles of the rubber wheels (Figure 32). Natural rubber was chosen

over other elastomers for these wheels. Although rubber does not have a large operating temperature range, it maintains its mechanical properties within its mechanical properties best within its

range.

Table 26. Material Choices for Actuator Wheels

Material Operating Radiation Important

Range (*K) Resistance Properties

Natural 220-350 highest good mechanically,

durable in space

Silicone 190-560 poor

Butyl 230-420 lowest

brittle, undergoes

rapid degradation in

compression, excellent

heat resistance

softens until becomes

tarry fluid at high rad.

The inertial load of the vehicle is kept off the axles of the friction wheels with titanium rings. These

rings are connected to the basket joints. The inside of the rings carry all the normal force. The rings

are lined with molded nylon to reduce friction and abrasion to the shaft track. The total force of

the laser rocket (768 N) is distributed over the four tings. The rocket basket makes its trip of 3.3

m over the entire trip time. The power required to move this basket comes to less than 10-3 Watts.

The short analysis is given below.

P=Fx V

F=#kxN

Where #, is the coefficient of kinetic friction, N the laser rocket thrust, and V is 3.3 m/s (10 day trip

time). The power required is 3 Watts.

This work is generated by electric actuator motors which are powered by the main fuel cells. The

motors are connected to the rocket basket using aluminum bands attached to truss members with

an epoxy bonding.

Cargo Bay

The cargo bay is designed for loading and detaching convenience. It is completely removable from

the main space frame (Figure 33). Not only does this allow for the cargo to be dropped off quickly

and easily, but the discarded truss can be used in the erection of a GeoShack. The vehicle is also

not restricted to the single cargo configuration that is depicted in this report. It can carry any

Structures 56
1990 LOCOST Senior Design Project

1.0 m

Actuator

Motors

i

Connector

Ring

1.0_

Guide

3haft

Rubber

Friction

Wheels

Figure 32. Laser Rocket Basket Top View and Side View

Structures 57
1990 LOCOST Senior Design Project

payload that can be attached to nodes and and which does not change the center of mass past certain bounds determined by the gimballing of the chemical engines.

The payload is stored in a cylindrical shell cargo bay which is mounted in a 18.9 m x 6.3 m x 6.3

m rectangular truss flame. In addition, the entire cargo truss being removable, truss members can

be removed from the end so that the cargo bay cap unlatches and the payload can be unloaded.

The cargo bay could then be reloaded and returned. The entire cargo truss with cargo shell need

not be left behind at the GeoShack.

The longitudinal axis of the cargo bay is mounted transverse to the acceleration of the vehicle.

Referring back to the materials table, magnesium alloy (8.5% A1) displayed the most desirable

properties for construction of the cargo container. It can sustain the largest transverse loads and

has a low density resulting in considerable mass savings. The shearing stresses on the walls can be

greatly reduced using stringers which run along the longitudinal axis of the cylindrical shell. The

wall thickness of the shell can be very thin with the use of stringers, because the stringers will carry

the bending and shear forces. With a wall thickness of 2 mm and eight stringers (Figure 34), the

cargo bay is capable of supporting the maximum payload of 40,000 kg. The only drawback of using

Mg/AI is that it has low tolerance to the space environment. The outgassing losses are greatly reduced by applying a barrier coating that is only a few molecules thick of a material with a much

lower evaporation rate (Ref. Parcel). As discussed earlier, aluminum used as protection against the

space environment is suitable for this purpose.

The cargo bay has three Mg/A1 hoops bonded to its outside. These bands have integrated nodes

that are aligned with the titanium nodes used to assemble the surrounding truss frame. These nodes

connect the cargo shell to the cargo truss.

Table 27. Mass Savings for Cargo Container

Dcnfity (kg/m _) Mass (kg) % Savings

Mg Alloy 1800 1550 33.6

Bo/AI 2490 2150 8.1

AI 2710 2340 0

The attachment planes on the titanium joints need a supplementary latching device (Figure 35) to

make the connection complete. The cargo attachment devices (CAD) are not autonomously operated, but they are simple enough to be actuated by robotics or the gloved hand of an astronaut.

The CAD is a completely separable piece but can be left attached to either mating joint. It has a

button, which when pushed, disengages the beatings that kept the CAD from rotating. To unlatch,

the button must be pushed so the CAD can be turned 90 degrees in either direction. There are eight

mating nodes between the main truss and the cargo truss.

Docking

The docking mechanism has the versatility to dock anywhere. This is done by limiting the number

of interfaces, one being the smallest but more than one offering greater stability once locked in.

The docking mechanism uses a modified CAD with an extra device for absorbing the shock of

docking (Figure 35). These damping CADs are mounted on the eight joints of the comers of each

6.3 m square (Figure 33). This provides the flexibility to dock at a space station by latching on

with any or all of the nodes of the visiting LOCOST vehicle.

A system is alsoneeded to measure the distance and closing speed between the space station and

the approaching vehicle. Two common methods for measuring range and range rates are laser and

radar devices. Unfortunately, they both cause damage to the object to which the distance is being

measured. The radar also has a limited minimum range. For these reasons, a range and range rate

system was invented that would be suitable for docking and closing maneuvers. The video controlled solid state range finding system designed by Russell, Graham, and Epperly requires no radar

or high power laser, has an effective range from less than a meter to thirty meters, and needs no

sophisticated target on the object to which the distance is measured.

Structures 58
1990 LOCOST Senior Design Project

- Cameo atCachmen¢ devices used a¢ Chese nodes

- Docking mechanisms used a¢ these nodes

Figure 33. Cargo Bay

Structures 59
1990 LOCOST Senior Design Project

I

6.3m

I

End View

6.3m

2

Figure 34. Cargo Bay Stringers

The system has an opto-electric camera and a helium neon laser which emits a source beam of

coherent light. This light travels through a beam splitter which sends a reference beam back to the

camera and sends an outgoing beam to the first angularly variable reflector which is then redirected

to the distant object. An incoming beam is reflected from the object to the second angularly variable reflector which directs the incoming beam to the opto- electric camera by way of the beam

splitter. The first and second reflectors are set up so that the distance D travelled by the outgoing

beam from the beam splitter and the first reflector is equal to the distance travelled by the incoming

beam from the second reflector to the beam splitter. The reference beam produces a reference signal

in the geometric center of the camera. The incoming beam produces an object signal at the camera.

The difference between the reference signal and object signal is used by automatic means to vary

the angle A between the outgoing beam from the first reflector and the reference line between reflecting points of the first and second reflectors and the angle B between the incoming beam from

the second reflector and the reference line. Angles A and B arc kept equal. The difference between

the reference signal and object signal is then used as input by a rotator driven circuit to change the

angles of the first and second reflectors until the reference and object signals are coincident. The

Range R can then be found by R = D tan(A) (Ref. Graham).

Structures 60
1990 LOCOST Senior Design Project

I

J

/ /

T

3..5 cm

I

l

I

J

f J

f J

/ .f

f J

f J

f J

J J

/ /

/ /

/ J

m

Shock Absorbers

18.75 ca

Viscous

D',,penlnG

Pluld

\

\

15 ca

/

/

/

Figure 35. Attachment Mechanisms for Cargo Bay Shock Absorbers

Propellant Storage 61
1990 LOCOST Senior Design Project

Propellant Storage

Introduction

The scheme for the type, number and placement of the propellant tanks is based on the mass of

propellants used. For the baseline mission, the total mass of oxygen required is 40,230 kg and the

mass of the hydrogen required is 52770 kg. The vehicle configuration requires that the tanks fit in

a fixed truss section of 6.3x6.3 m. Considering the restrictions put on size, spherical tanks were

chosen for the oxygen and cylindrical for the hydrogen. The hydrogen tanks have a 6.3x6.3x 12.6

area to fit into.

Tank Material

The material chosen for the tank skin is an aluminum alloy. It was chosen based on its high specific

strength, high stiffness, low density, availability, low cost, and superior cryogenic toughness. A

comparison to another common alloy is shown in Table 28.

Table 28. Comparison of Aluminum Alloys

Metal AICu Alloy (2219 T87)

Yield Strength (N/rmr_) 370

Ultimate Strength (N/ram:) 440

Young's Modulus (x 10SN/mrr_) 0.72

Density(g/m3) 2.82

AILi Alloy (2090 T8E41)

535

565

0.758

2.55

The specifications for each individual tank are calculated in Appendix D. Relevant information is

shown in Table 29.

Table 29. Tank Data

Diameter (m) Thickness (ram) Mass(Shell-kg) m (fuel-kg)

LOX 3.23 .067 5.6 20,115

LH2 5.3 0.135 138.62 26,384

The pressure maintained inside the tanks is predetermined to be 34.475 kPa. This is the lowest

practical pressure that still provides a net positive suction head at the pump inlet. A NPSH is required to avoid cavitation at the outlet of the tank. This pressure is also desirable because it is

above the triple point of cryogenic fluids.

The tank interior is designed to support slosh baffles and a liquid acquisition device. Some of the

schemes examined to expel the cryogenic liquid in a zero g environment were to use the RCS, a

wick, or a liquid acquisition device. The RCS scheme maneuvers fuel into the correct position to

be siphoned out. Such a system was judged to be impractical for our vehicle. The use of a wick

to absorb the fuel and lead it to the outlet was also discarded because of lack of sufficient information. The liquid acquisition device was finally selected. A stainless steel 325x2300 mesh is used.

The most efficient way to provide total communication of the liquid is to use a device composed

of individual flow channels. The channels are bolted to the wall with non-metallic, teflon spacers.

Propellant Storage 62
1990LOCOSTSeniorDesignProject

The cylindrical configuration uses four vanes. The operation of the device depends upon the

interaction of the liquid/gas interface with the device. The device is configured to provide gas free

expulsion of the fluid. The mass of the liquid acquisition system is approximately 300 kg for each

oxygen system and 500 kg for the hydrogen.

The oxygen tank has an 800 micron screen at the exit to screen out impurities caused during storage. The hydrogen tank has a 400 micron screen. Hot wire sensors axe used for loading control

logic. Vent and relief valves in the nose cap of the tanks provide for gaseous propellant vent capability and for relief protection. Ullage transducers axe used by the actuation of control valves to

control ullage pressure in flight. Three to five percent space is left for ullage. The tanks stretch as

a function of static head and shrink when the cryogenic liquid is loaded. (See Figure 36)

DIFFUSER

PRESS.

VENT/RELIEF VALVE

PRE-CHILL SPRAY

NOZZLES

SPACE MUILTI-FEED

DISCOmlECT

POINT $EtiSOR$

_8AFFLES

FEED LINE

START $U/,tP

SIPIION ASSY

Figure 36. Spherical Oxygen Tank

Tank Insulation

The insulation chosen was a multi-layered insulation (MLI) layered with foam. This scheme has

the lowest density and has proven to be a capable insulation scheme. The MLI is composed of

alternating layers of metalliTed Kapton film and Dacron net spacers. Between the MLI and the

foam is a gas-impervious aluminized Kapton barrier film which is bonded over the inside wall of

the foam to keep degradation factors from reaching the MLI. The foam used is Rohacell 31. A

comparison to other foams is shown in Table 30. Rohacell 31 is superior in performance. It is

manufactured in sheets which are heat formed to the tank, then adhesively bonded to the surface

using vacuum bagging. A layer of paint follows the insulation. The paint is silicate bonded and

contains zinc othotitanate pigments (Zn2TiO4). An 8 mils thickness of YB-71 is then applied for

reflectivity. A final layer of mylar is applied to protect against micrometeorites.

Propellant Storage 63
1990 LOCOST Senior Design Project

Table 30. Foam Comparison

Type Material Density (kg/m3)

Rohacell 31 Polymethacrylimide 30

Stephan Foam Polyurethane 37

(BX25OA)

CPR 488 Polymetirc Isocyanate 36

A picture of the insulation and a summary of the characteristics are shown in Figure 37 and

Table 31, respectively. The outlet of the tanks is a contoured non-drop-out configuration with

anti-vortex baffles at the exit.

Propellant Storage 64
1990 LOCOST Senior Design Project

OXIDIZER TANK

6.3M

504 M

HYDROGEN TANK-END VIEW

.5176

2090 AL-LI

D=5.3 M

EPOXY

GIRTH

RING (6061-T6)

CHANNEL

(606 I-T6 )

Figure 37. Insulation

Propellant Storage 65
1990 LOCOST Senior Design Project

Table 31. Insulation Summary

Qvf(W) -- Venting Heat Flux

Qsr(W) = Solar Radiation Heat Flux

Ts(K) = Temperature On Tank Surface

Qvf Qsr Ts t(MLI-cm) t(Foam-crn) m (total-kg)

LOX 23.62 1374 196.9 1.86 .186 46.46

LH2 64.91 16901 196.49 1.86 .186 570.98

Tank Support Structures

The support structures for the tanks are made of composite materials. A comparison of basic

composite materials is shown in Table 32. The structures used were chosen to minimize boil off

from thermal conduction of heat to the tanks.

Table 32. Comparison of Composite Materials

Ultimate Tensile

Strength (MPa)

Modulus of Elasticity

106 KPa

S-Glass 1680 46.2

SP-250-52 1675 46.2

181 E-Glass 827 41.4

Optimum Composite 334 42.8

Lay-Up OF 334 42.7

S-Glass/E-Glass

The spherical symmetric oxygen tanks have S-glass epoxy members attaches to the tank in 6 places.

The S-glass attach to a graphite- epoxy frame which fits inside the basic truss. The frame is attached

by titanium clamps.

For the hydrogen tanks, three 6061-T6 girth rings attach to 24 S-glass epoxy members which in turn

attach to three 6061-T6 ring support channels. The ring support channels have longeron trunnion

fitting interfaces which attach the whole structure to the main truss. The support structures are

shown in Figure 38 and Figure 39. Mass breakdown of the tank support structures is given in

Table 33.

PmpeUant Storage 66
1990LOCOSTSeniorDesign Project

6.3M 1

SUPPORT CHANNELS

Figure 38. PropellantTank Support Structure I

Propellant Storage 67
1990 LOCOST Senior Design Project

/

tq_t (_ rxl_ g_o t_g:m_ Nm')

.o

RCI-LILmI_31 FO_

PAINT

._LINff.NIZED KAP'IX_

,_. BARRIXRF7224

TANK CR36S _CTICN

\

Figure 39. Propellant Tank

Propellant Storage 68
1990 LOCOST Senior Design Proiect

Table 33.

Oxygen-Spherical

S-Glass/Epoxy

Total Mass

Mass Breakdown of the Tank Support Structures

Density # Mass

(kg/m3) (kg)

2490 24 601

Hydrogen- Cylindrical

601

Density # Mass

(kg/m3) (kg)

S-Glass/Epoxy 2490 24 .7 .5176 241.29

6061-T6 2550 6 1 19.761 94.98

6061-T6 2550 6 1 18.2 87.433

Total Mass 397

Total Support Mass 1996

Propellant Pumps

Electric pumps used to transport the oxygen towards the engine weigh 40 kg and require 1.04 kW.

The Ap required for the oxygen is 1.9 MPa and that for hydrogen is 7.2 MPa.

The fluid flow lines have a total mass of 150 kg and are made of aluminum with insulation.

Table 34 contains the diameters of the outflow, inflow, and pressurization flow lines for the oxygen

and hydrogen tanks.

Table 34. Fluid Line Diameters

Outflow (cm)

Oxygen 9

Hydrogen 5

Inflow (cm) Pressurization (MPa)

6 1.5

6 4

Propellant Storage 69
1990LOCOSTSeniorDesignProject

Dynamical Control System

Introduction

A vehicle largely dependent on an incident power beam must have precise and accurate attitude

control. For over 15 years, momentum storage devices have been successfully used to rotate space

vehicles. Possible momentum storage devices are summarized in Table 35.

Table 35. Possible Momentum Storage Devices

* Fixed Momentum Wheels

• Reaction Wheels

• Gimballed Momentum Wheels

While these momentum storage devices all provide rotational ability, they are subject to saturation

limits. When they become saturated, a non-conservative device such a gas-jet must be used to return the system to the initial state where it can, once again, receive angular momentum. This feature does not subtract from the great strength of the devices, which is the ability to store large

momenta, and hence, instigate a high attitude stability.

To supplement the "flywheel" attitude control devices, a reaction control system (RCS) is utilized.

A RCS is composed of small thrusting, low I,, rockets. There are other methods of desaturizing

the momentum storage devices, such as magnetic torques and gravity-gradient torques. The space

station will use gravity-gradient torques (Wie, Byun et all). But these methods are too complex and

not appropriate for the relatively small size of the LOCOST vehicle, hence an RCS will be employed.

The Hubble Space Telescope employs reaction wheel assemblies with integrated desaturation

torquers and small direct drive motors to negate the effects of secular torques (see below). Additional study of this system could yield a scheme that is superior to the RCS system discussed below

(Sevaston et al). Of course, a RCS system would still be needed for docking.

Discussion of Gyro Technology

A schematic of the vehicle's body centered axis is shown in Figure 40. This body axis orientation

will be used in the remainder of the chapter. With this reference system, the yaw rate is rotation

about the x-axis, slew rate about the y-axis, and the roll rate about the z-axis.

The most significant low frequency disturbances affecting spacecraft axe summarized in Table 36

(Lademarm).

Dynamical Control System 70
1990 LOCOST Senior Design Project

\

i

\

Figure 40. LOCOST Vehicle'sBody Axes

Table 36. Low FrequencySpacecraft Disturbances

* Gravity-gradients

* Aerodynamic (especially near LEO)

• Solar radiation pressure

• Miscellaneous cyclic disturbance torques

Unless counteracted, these torques would cause the vehicle to tumble. Mass expulsion devices can

be used to counteract these torques, but these methods suffer the obvious disadvantages of extra

mission mass and mechanical accelerators. Momentum-exchange devices, such as flywheels and

CMGs, utiliTe a recoverable (electrical) energy source to maintain the vehicle's attitude.

Some of the tradeoffs associated with different momentum exchange actuators are summarized in

the Table 37 (Lademann). In this table the following acronyms are used: RWA for Reaction

Wheel Assembly, SGCMG for single gimbal CMG, DGCMG for double gimbal CMG.

Obviously, pointing and tracking are a high priority, however as can been seen in the table, some

desirable attributes are corttlicting. Tolerable compromises of performance must be made with the

intent of minimizing the impact on the design of any one particular parameter.

Note that the fixed momentum flywheel (RWA in the table) suffers the disadvantage of not being

able to completely transfer its momentum to the vehicle. Gimballing of the flywheel offers the

possibility of total momentum exchange between vehicle and flywheel (Schulz & Lange).

Some of the pros and cons of attitude control with momentum exchange devices are summarized

in Table 38.

Dynamical Control System 71
1990LOCOSTSeniorDesignProject

Table 37.

Device or

Parameter

Hardware

Complexity

Tradeoffs for Different Momentum Exchange Actuators

RWA SGCMGs DGCMGs

Least Intermediate Most

Comments

RWAs requh'e only speed

control circuits. CMGs

require additional gimbal servos

Control

Law

Complexity

Least Most Intermediate RWAs usually require pseudoinverse steering law. DGCMGs

have few singularity problems

while SGCMGs have severe

internal singularity problems.

Torque

Capability

Least Most Intermediate RWAs require large power for

high torque. DGCMGs

transmit torque through

tourquers and gear train while

SGCMGs transmit torque

directly vehicle.

Momentuln

Capability

Least Most Most Generating torque by changing

wheel speed at high speed

requires excessive power.

Pointing

Accuracy

Capability

Most Intermediate Least Since RWAs are not gimballed,

torque is controlled more

precisely. However,

hardware anomalies associated

with momentum excha.t_ge

devices are usually not the

main contributors to pointing

errors since rate sensors

close the loop around

actuator errors.

Table 38. Main Points with Momentum Exchange Devices

* Momentum exchange devices may be control momentum gyros (CMGs)

or reaction wheels (RWAs)

* Control System Reacts to External Disturbances Continuously

• High pointing accuracy is possible

• Momentum exchange device stores integral of external torques

• Since external torques have bias components, stored momentum

tends to grow, leading to saturation

• Saturation of RWAs is reached at maximum wheel speed

• Saturation of CMGs is reached when momentum vectors are pointing

in the same direction

Dynamical ControlSystem 72
1990 LOCOST Senior Design Project

Lademann offers three heuristics when selecting a momentum exchange device. They are summarize in Table 39.

Table 39. Momentum Exchange Device Selection Heuristics

1. Use RWAs if torque and momentum requirements are within RWA's

capability.

2. Use SGCMGs when very large torques axe required.

. Use DGCMGs for large space structures with moderate torque

requirements when momentum storage requirements may increase as

additional modules are added.

The last selection criteria in Table 39 is the most important for the LOCOST vehicle because of

the extra side propellant tanks to be added to the structure of the vehicle to fulfill the maximum

cargo scenario. Hence, double gimballed CMGs will be used. Also, these tanks will cause large

moments of inertias which must be counteracted by larger torques. As seen from Table 37, they

suffer the disadvantage of being more complex and sustaining the lowest pointing accuracy capability. It is assumed that by 2010 the DGCMGs will have increased their pointing accuracy. The

DGCMGs axe housed in metal boxes roughly 1 m on a side. Gimbals are essentially large boxes

with the rotor inside.

Nominally, three double gimballed CMGs are needed to provide rotation about all three body axes.

If more than one plane of symmetry had existed, it would have been possible to reduce the number

of CMGs. Each of the CMGs has an inner and outer gimbal ring. The outer ring allows rotation

about the angular momentum vector, while the inner ring allows a precession of the rotor axis

about the inner gimbal. The basic principle is that a torque (applied by an electric motor) about

one axis causes a precession about a mutually perpendicular axis. Through conservation of angular

momentum, the vehicle rotates. The precession rate is determined by the applied torque and the

moment of inertia of the vehicle.

While the vehicle rotates, it is necessary to have some feedback of the rotation rate. Possible ways

to gain information about the instantaneous angular rate of the vehicle include flywheels, star sensors, and gyroscopes. In most cases, a flywheel is a gyroscope, but the terminology used above

was used to differentiate between flywheel technology and more advanced gyroscopes such as ring

laser gyroscopes and fiber optic gyroscopes. Although the need for a star sensor will be established

independently of the CMGs, the star sensor will not be used to obtain angular rate information.

As a final note on high accuracy pointing applications of which the LOCOST vehicle is one, a

control law that is highly insensitive to errors must be used. Derivation of the control laws is an

important step in the implementation of the CMGs and overall dynamical control.

Discussion of Gyroscopes

Gyroscopes give angular rotation information. There are many ways to implement a gyroscope,

three of which will be presented.

Flywheel gyroscopes have the advantages of low noise, analog output, long life, high reliability, and

extremely low bias. The flywheel suffers numerous disadvantages. Recent advances in gyroscope

technology renders the flywheel gyroscopes obsolete.

Ring laser gyroscopes (RLG) appeared in the late 1960's (Lademann). RLGs do not have spinning

flywheels, but use lasers and the time of travel between two points to determine angular acceleration. RLGs contain an optical cavity with mirrors at each corner. A shining incident laser is split

into two beams. The beams axe then collected at photocells after reflecting off the mirrors. During

angular rotation, an apparent difference in path length is developed between the two beams. This

directly translates into a frequency shift which can be detected. A fringe pattern created by the inDynamical Control System 73
1990 LOCOST Senior Design Project

terference of the two beams, one shining through partially silvered mirrors, yields directional information, while the frequency shift yields scalar angular rate. RGLs do suffer lock-up rate, a

condition during low angular rotation rates that cause poor accuracy. Also, mirror surface erosion,

outgassing of materials within cavity, path length control precision and mirror optical axis stability

make RLGs an unsuitable choice.

A new type of gyroscope offers the potential for eliminating many of the problems associated with

RLGs (Stokes). The Fiber Optic Gyroscope (FOG) is the newest generation of rotation sensor

being developed. Both Bendix and Sperry have produced prototypes. The FOG has the advantage

of the complete absence of moving parts. Basically, the FOG is a non-resonant ring interferometer.

Only solid-state technology and fiber optics is used. This assures long life, high reliability and better

performance. A semiconductor diode provides the laser fight. The FOG uses a multi-turn fiber

loop--sometimes over a kilometer of fiber. The longer the fiber path the better accuracy and precision. Rotation of the FOG causes a phase shift between counter-rotation light waves. The phase

shift is given by the relation

2_rDL

AIk =--_ 2c

With A_b -- optical phase shift

D = loop diameter

L = fiber length

2 = wavelength of light

c = speed of light

a>= rotation rate

A schematic of the FOG is shown in Figure 41. Also, the main advantages of the FOG are summarized in Table 40.

Table 40. Advantages of a Fiber Optic Gyroscope

* High reliability (no moving parts)

* Low environmental sensitivity (high-g applications)

* Low cost (no critical machining or assembly operations)

* Small size and weight (= 1 kg.)

FOGs are not yet perfected. In fact, they suffer from more noise then their mechanical counterparts. It is expected, however, that by the projected year of the LOCOST vehicle, FOG technology

will have been reached a sufficient level to justify its use. The FOG requires 3 Watts per axis

measurement and uses only low frequencies (Stokes).

Specification of CMGs

Specification of the CMGs requires the determination of angular rates, vehicle moments of inertia

and the placement of CMGs relative to the the center of gravity (CG). If the CMG is not placed

directly at the CG, inefficiencies occur. It is assumed that there is a 100% efficiency of momentum

transfer for the following analysis. Angular rate information was obtained from the orbital mechanics calculations. The angular rotation rates were determined trigonometrically from vectors

from the vehicle to the LPS and the desired angular attitude relative to the body fixed axis system.

The torques necessary to generate the precessions are determined by multiplying the axes" moments

of inertia and the angular acceleration required to achieve the desired final angle. The values for

the moments of inertias used are summarize in Table 41.

Dynamical Control System 74
1990LOCOSTSeniorDesignProject

SOURCE LIGHT

DIRECTIONAL DIRECTIONAL

COUPLER COUPLER /

 OOULATOR j

t

PROCESSING SIGNAL i _I MODULATIONGENERATOR I

L'_ GYRO OUTPUT

Figure 41. Fiber Optic Gyroscope (FOG) Schematic

Table 41. Moments oflnertia

Moment Type

Ix

Iy

Iz

Value ( x 106 kg- trY)

.9

1.2

1.7

The maximum precession angles of the gyros and the angular rates that were used arc summarized

in Table 42.

Table 42.

Yaw

Slew

Roll

Maximum Precession Angle of Gyros and Angular Rates

Precession (°) Angular Rate

10 < 1 x 10=s

60 .001 (nominal)

15 < 1 x 10-s

In order to satisfy the torques involved a compromise was achieved. The torques needed for two

out of the three axes is small compared to the third, therefore an array composed of two CMGs

was used to satisfy the higher torque requirements. An analysis of single-gimbaUed CMG array's

momentum envelope topology is performed by Merle & Stocking. A similar analysis should be

done for the DGCMGs because it is possible that the non-linear effects of the CMGs will cause the

CMG's momentum contributions to be non-additive. Additionally, the topology should be explored because the arrayed CMGs might produce a surface where no additions of angular moDynamical Control System 75
1990 LOCOST Senior Design Project

mentum can be stored or transferred. The analysis of such topology envelopes is beyond the scope

of this design report. It is assumed that the CMG contributions to angular momentum can be linearly superimposed.

The CMG chosen is a Bendix developed, advanced CMG. It's parameters are summarized in

Table 43.

Table 43. Selected Baseline CMG

Parameter

Wheel Speed

Gimbal Rate

Angular Momentum

Rotor Inertia

Output Torque

Mass

Gimbal Rate

Command

Power (Zero G)

Quiescent

Peak Power Capability

Size _- 1 m on a side.

Capability

lOO(Hz)

2.0 rad

sec

2400 N-m-sec (minimum)

3.54 N-m-sec 2

4700 N-m (minimum)

100 kg

15 Bits

Including Sign

50 Watts

70 kWatts

The differences in the above specifications and those given by Lademann are due to predicted

technology advancements. Figure 42 shows the placement of the CMGs on the LOCOST vehicle.

DesatuHzation Rocket System

Only a small RCS system is needed because the vehicle already contains a powerful chemical propulsion system which can account for large orbital corrections. Therefore, only small

desaturizations and docking maneuvers must be performed via the RCS.

A 5 N thrust per rocket nozzle RCS system was chosen for the vehicle. Because of the impulsive

nature of the RCS system, the RCS should not be used during laser firing because the perturbations

developed can cause intolerable and disastrous errors in optical alignment.

For the baseline configuration, a hydrazine powered (oxygen oxidizer) rocket will be used. By 2010,

other propellants will be accessible and a conservative estimate for the I,, of the rocket is 260. The

use of a monopropeUant gives the system simplicity and reliability. Also, the small size of the

rocket means that radiative cooling is sufficient. The hydrazine propellant will be stored with the

clusters.

Placement of RCS

Clusters of three rockets are attached to each of the eight main comers of the vehicle. Figure 42

shows the RCS configuration chosen. Through this arrangement, it is possible to both translate

and rotate the RCS through its three degrees of freedom.

The mass flow rate can be determined for each of the RCS nozzles from,

Dynamical Control System 76
1990 LOCOST Senior Design Project

\ /

\/

Bend ix CMG

Bendix CMG

2 Bendix CHG's

/ _ = Placement of

i_. 2 RCS nozzles

Figure 42. CMG and RCS Placement on the vehicle.

T kg

= -- = .002

" t,pg,

To account for the propellant needed to power the RCS for the baseline mission, it was assumed

that desaturization could occur over a period of 800 seconds. This corresponds to the period in an

orbit where the LPS is located. During desaturization four engines would burn. For the baseline

mission, it is assumed that no more than 15 desaturizations will occur. For this, a total of I00 kg

of hydrazine is consumed.

Docking requires the careful manipulation of the vehicle. Because of its large mass size compared

to the total possible translational force in any of the three body axes directions, the RCS will need

to burn much longer during docking. A ten minute burn per nozzle is assumed for any particular

docking maneuver. For docking both at GEO and LEO, two ten minute bums are required. This

will consume 2.4 kg of propellant per nozzle. During docking, it is assumed that two directions

Dynamical Control System 77
1990 LOCOST Senior Design Project

of applied force will be needed, hence eight nozzles must fire. For all docking maneuvers then, only

19.2 kg of propellant will be needed. Including a 20% residual, the baseline mission requires 23

kg of propellant. The characteristics of the RCS and CMG system are summarized in Table 44.

Table 44. CMG and RCS Summary

Selected RCS Baseline Summary

Thrust per nozzle

Specific Impulse

Mass flow per rocket

Nozzles per cluster

No. of clusters

Propellant needed for both

desaturization and docking

in the baseline mission.

5N

260 s

.002

sec

3

8

123 kg

Dynamical Control System 78
1990 LOCOST Senior Design Project

Acquisition, Tracking and Pointing

Introduction

Acquisition, tracking and pointing is the process of acquiring the laser beam and subsequently

continuously tracking it by pointing an antenna towards the beam direction of propagation. Laser

powered vehicles require an accurate and dynamic system to insure the reception and control of the

power beam. The LOCOST vehicle ATP system must be able to locate and maintain the laser

beam link.

Pointing control is also needed to counteract the effect of external disturbances. Extemal disturbances include gravity gradient torques, solar pressure torques, and collector motion coupled through

the rotary joint. Also, changes in antenna attitude, thermal loading, aging effects, and structural

deflections necessitate dynamic control of the power beam (Ref. NASA CR-3350).

Similar problems with the ATP for satellites have already been addressed in the field of laser communications (lasercom). Much of the design analysis and selection presented in this chapter is

based on the design analysis for lasercoms.

The ATP system relies on three other subsystems of the vehicle: the communications system, the

control computer, and the RCS & CMG system. The interrelationships are shown in Figure 43.

The communication system establishes the initial link between the LPS and the vehicle, from which

the entire ATP cycle is initiated. The control computer coordinates the interactions from the ATPs

position and attitude demands to the appropriate subsystems. The RCS & CMG system fulfLUsthe

position and attitude demands of the ATP. All the systems work in concert.

ATP systems have also been developed for payload-experiments. Experiments often require continuous pointing, but on spin.stabiliTed satellites, continuous pointing presents many problems.

Payload pointing mechanisms are usually employed. The Hubble Space Telescope requires a free

pointing mechanism to continuously track stars. Its system employs traditional techniques of attitude control and attitude determination. Through good design practices, its system is capable of

high accuracy and precision (Sevaston et al).

The A TP @stem

The ATP system is composed of five major components. They are summarized in Table 45.

Table 45. Ma]or Components of ATP System

• Tracking Laser

• Optical Receiver

• Optical Transmitter

• Fine Pointing Gimbal Assembly

• Mirror Alignment System

The mirror alignment system includes alignment control for the primary, secondary and fifth mirrors.

Acquisition, Tracking and Pointing 79
1990 LOCOgT Senior Design Project

Tracking Laser

There are four types of lasers to choose from for the LOCOST vehicle tracking laser. The tracking

laser must be power efficient, low in mass, and resistive to the space environment. The four types

of lasers are presented in Table 46.

Table 46. Possible Lasers for the LCOTV

Laser Type ,1 (micrometers)

AIGa.As 0.89

Nd::YAG 1.06

doubled

Nd::YAG 0.532

CO2 10.6

The AIGaAs diode lasers are efficient but of low power. Hence, a large array must be used. Large

arrays, however, introduce beam.combining problems. Also, these diode laser suffer appreciable

wavelength shifts from aging. The COs lasers are reliable, but problems arise due to their sealed

RadLo Links

Optical System

Feedback

Fiber Optic

Gyroscope Structur

Feedback

m

I PropeI1ant System

Pumps, Valves, and

Pressures... Highly Paral)el

Processor

8ange/rat

System

ar Sensor

CHG's

.&set Communlcationsl

Figure 43. Interrelationships of ATP with Other LOCOST Subsystems

Acquisition, Tracking and Pointing 80
1990 LOCOST Senior Design Project

tubes containing the gases. Also, the carbon dioxide laser would need cryogenic cooling equipment.

Both of these lasers were dropped from consideration.

The Nd:YAG (Neodymium: Yttrium, Aluminum, Garnet) lasers have become the workhorse laser

for lasercom (Ref. Katzman). By using the doubled Nd:YAG laser, a direct detector can be used

(discussed below). The Nd:YAG will be pumped with AlGaAs diode lasers. Because of the

problems associated with the A1GaAs diode lasers, the design calls for a modular AlGaAs diode

array. As the diode lasers suffer age effects, the pumping unit can be replaced with a new one.

The laser power can be significantly reduced by pulsing the beam. Pulsed power is 200 Watts at

peak power for 1 microsec. The laser has a maximum of 1000 pulses per second. No modulation

techniques are needed because no information needs to be coded into the beam. The presence,

absence, or spatial shift of the beam from the receiver communicates the required information.

Optical Receiver and Transmitter

A receiver can utiliTe an independent antenna or share an antenna with the transmitter. If the receiver and transmitter antenna axe independent, added complexity and mechanisms for accurate

boorsighting must be utilized to maintain two lines-of-sight. Therefore, the transmitter and receiver

will share the same antenna. This realizes a weight and volume savings and allows for easy alignment. However, this approach requires internal optics to separate the two signals. Also, a practical

limit on the data transmission rate is introduced (<0.1 Mbit/s), but this does not pose a problem

because the tracking system relies on the location and existence of the pulsed beam--not on the

information modulated within (Ref. Katzman).

For lasercom, an antenna is actually a telescope. For the LOCOST vehicle, a Schmidt-Cassegrain

telescope setup is used. The Cassegrain setup minimizes the effects of pointing jitter and misalignment bias (Ref. Katzman). The system is mounted on a two degree of freedom yoke and attached to a boom. The boom is located at the pivot point of the primary minor. Figure 7 shows

the configuration. Figure 44 shows a labelled schematic of the transceiver.

The telescope has several subcomponents which are discussed below.

The Telescope Shelh(A). The telescope shell is made from Lightweight aluminum formed into a

cylinder. The cylinder dimensions are 45 cm high with a 40 cm diameter. These dimensions are

sufficient to catch a pilot beam from the LPS that has dispersed slightly. This aperture size also

makes a sufficiently sized beam at the vehicle, so that a reasonably sized LPS telescope can capture

it. The telescope shell is attached to a yoke. The gimbal axis of the yoke and the gimballing of the

entire yoke comprise the two degrees of freedom needed for easier pilot beam acquisition. This

yoke comprises the t'me gimbal pointing assembly.

Both the azimuth and elevation are controlled by induction motors contained within the gimbal

assembly. Each motor consumes 12 W of power (Ref. LPIV).

The Dielectric Filter:(B). The dielectric fdter is placed over the opening of the telescope to reduce

possible environmental damage to the multi-dielectric surfaces (discussed in element F). The fdter

also reduces the false alarm rate. There were two choices for the dielectric filter: fused silica or BK7

glass (Ref. Katzman). Both have good transmission in the wavelength being used (0.532 micrometers). However, fused silica experiences a significant decrease in transmittance under radiation

bombardment. The BK7 glass was chosen.

Polarizing Beam Splitter:(C). The polarizing beam splitter splits a beam according to its

polarization. As discussed above, the send and receive beam propagate at the same wavelength but

have different circular polarizations. The utility of this scheme rests on the stable polarization

characteristics of the Nd:YAG laser. Also inherent in this scheme is the assumption that free space

will not alter the initial polarization.

The beam splitter is made of quarter-wave dielectric stacks, similar in design to the primary mirror

of the vehicle. Figure 45 shows a schematic depicting the effects of the polarizing beam splitter.

Acquisition, Tracking and Pointing 81
1990LOCOSTSeniorDesignProject

i

Polarizing Beam Splitter

m

>

//

I

Figure 44. Optical Trasceiver Labelled Schematic

BE7 Glass dielectric

fLlter

_Quadrant

Photodector

Receiver

Receiver:.(D). The receiver sits at the bottom of the telescope, beneath the polarizing beam splitter.

Incident beams reflect off the small central obstruction, onto the beam splitter. The beam splitter

is configured to allow a preset polarization through.

For laser acquisition, the dominant requirement is signal detectability. For laser tracking, high

signal-to-noise ratios are needed for the angle output. A quadrant photodetector yields a high SNR

and will be used in the vehicle's optical receiver. The quadrant protector is a direct detection detector. This type of detector is most efficient when large intrinsic gain is available and when the

input wavelength is below 1 micrometer. Above one micrometer, thermal noise can dominate and

systems such as heterodyne or homodyne detectors must be used.

There are three types of direct detector elements available: photo-voltaic, biased-photodiode, and

avalanche photodiode (APD). The first has limited usefulness, but the last two are fast and accurate. The lower power requirements of the APDs and their smaller mass make it the device selected

for beam detection.

There are four APDs--one in each quadrant. The basic detection operation involves the computation of the intensity of the beam in the quadrants, normalized with respect to the incident beam.

By computing the change in intensity between the quadrants, it is possible to calculate the attitude

change of the laser beam. This information is translated into the vehicle's attitude motions by the

control computer.

Acquisition, Tracking and Pointing 82
1990 LOCOST Senior Design Project

X

Dielectric Stacks

Polarization of Transmit Beam

Polarization of Receive Beam

Figure 45. Polarizing Beam Splitter

Transmitter Path:The transmitter path is simply the path followed by the laser beam from the vehicle to the telescope. From the schematic of the telescope, it is obvious that the laser does not

have to be contained on the gimbal platform, rather, it can be off-platform. However, for simplicity, the laser will be located on the platform.

Mirror Alignment System

A network of mirror alignment system is distributed among the primary, secondary and fifth mirror.

Photodetectors are also placed about the circumference of the laser rocket. The purpose of this

system is two-fold. First, it allows free tuning of the optical path and second, it provides a redundant control (in addition to the laser tracking beam). The feedback system is described below for

each mirror and the laser rocket.

Primary Mirror:The primary mirror has a system of APDs built into the dielectric surface. They

axe located on three rings centered about the center of the primary mirror. The three rings have

radii of 6.0, 6.33 and 6.66 meters. The APDs axe angularly spaced every 20 °. The APDs contain

dielectric filter windows tuned to the incident wavelength of the power beam. The inner most circle

of rings corresponds to the actual beam spot size. The others are equidistant from the edge of the

primary mirror and the edge of the spot (Figure 46).

Upon movement of the laser power beam, the APDs will trigger and the control computer will

adjust the position of the main mirror via attitude mechanisms.

Secondary Mirror.The secondary minor has four photodetectors situated around the minor's periphery. These photodetectors detect focusing problems from the primary minor. Due to mirror

surface stresses, thermal gradients, and other surface altering disturbances, the surface of the primary

mirror will not focus squarely on the secondary mirror. If the beam does not converge evenly on

the secondary mirror, part of the main beam will be reflected back onto the primary mirror causing

further thermal damage and poor laser rocket performance. When this mis-focusing occurs, the

APDs will trigger, and slight focusing characteristics can be changed by moving four small, pneumatic actuators in contact with the backside of the mirror (Figure 46).

Acquisition, Tracking and Pointing 83
1990 LOCOST Senior Design Project

_APD

0 0

j o

Figure 46. Mirror Alignment System for Primary and Secondary Mirrors

Fifth Mirror.The third and fourth mirrors do not need any mirror alignment management system

because once the beam strikes the third minor properly (with the help of the secondary mirror

alignment system), the path from the third to the fourth is rigidly set. However, the fifth mirror

must insure that the beam strikes the laser rocket engine properly. To insure this condition, APDs

have been placed on both the fifth mirror and the laser rocket engine.

Four APDs have been placed on the fifth mirror on both the top and bottom edges and on the

sides. Also, a small rotational mechanical transducer has been attached to the fifth mirror's pivot

axis to allow for small angular rotations. This arrangement can be seen in Figure 47. When the

APDs trigger the fifth mirror will rotate about the gimbal axis.

Laser Rocket:The laser rocket also has four APDs situated about it's outer edge. This arrangement

is shown in Figure 48. If the APDs on the top or bottom of the rocket trigger, the laser rocket

basket can move to correct the alignment. If the APDs to the right or left trigger, then modifications will have 1o made to the fifth and secondary mirrors because the laser rocket has no

translational capability to move to the left or the right.

The block diagram shown in Figure 49 summarizes the basic interactions between different subsystems that comprise the ATP system.

Acquisition, Tracking and Pointing 84
1990 LOCOST Senior Design Project

LINEAR

ACTUATOR

Figure 47. Mirror Alignment System for Fifth Mirror

Typical A TP Scenario

A typical ATP scenario consists of four distinct phases.

ATP Phase I--Aiming of the LOCOST vehicle

The vehicle uses its omnidirectional antenna to initiate the ATP sequence. The LPS responds using

its RF antenna, The vehicle relays its attitude and spatial orientations to the LPS. The LPS sends

out its position information, The vehicle aligns itself so that it is now capable of using its directional antenna for telemetry tracking and control. The vehicle orients its lasercom telescope in the

direction of the LPS.

ATP Phase IIl--Acquisition of Pilot Beam

The LPS emits a defocused laser beam into the error volume. The optical receiver detects the defocused beam, The vehicle emits a return laser beam, The LPS detects the return laser beam. The

LPS narrows down the error volume until a narrowly focused pilot beam is achieved.

ATP Phase IV--Tracking of the Power Beam

The vehicle continues to track the laser pilot beam. When the beam becomes misaligned due to

vehicle movement, the attitude controllers onboard the vehicle activate. The directional antennas

are used to relay telemetry and vehicle=health information.

Acquisition, Tracking and Pointing 85
1990 LOCOST Senior Design Project

ON I..T POSSIBLE

HOVEHENT OF LASH

IOC[ET

Figure 48. Laser Rocket Mirror Alignment System

Figure 49.

Stufing _ Fil_ I_iminll Cotm'ol

H ¢7) l

Sowca

Alla'l'Y_)Iy

I

4too Tr,.,'_ in !

E kL-'Iron _,

,3, I

Coarw and F,ne Tr_ki_ll

D4P_K't O¢ Am_rr_lv

ll'

_t _ t

Sl_lCllCr a_l I ntlrt IlCll Tr inlm_l_ u_ll RKarvlnlil

Oatm L_ Data L_nt

Ol)t ¢,,III Path

E_c Sqlrel l_th

Block Diagram of Sub-System Interactions of the ATP System

Communications 86
1990 LOCOST Senior Design Project

Communications

Introduction

The communications system of any unmanned space vehicle is essentially the life line of the vehicle.

Although the data transfer requirements are lower than that of a manned vehicle, the communications link must provide sufficient communications for control of the entire vehicle. This includes

telemetry, vehicle control, and initial position tracking.

The communications system data load of the LOCOST vehicle is slightly reduced by the tracking

done through the APT system. However, the initial link between the vehicle and laser station must

be made through the communications system in order to provide the APT system with starting

position coordinates and alignment parameters.

Communication system requirements

The communication system on a spacecraft such as the LOCOST vehicle must have high reliability

low power consumption, high speed data transfer, and longevity in use. The system must transmit

and receive signals efficiently if it is to have a low power consumption. This plus the great distances

through which communication will be done necessitates the use of one or more highly directional

antennas. Since the communications system must also provide the initial link from a possibly unknown position, an omni-directional antenna is also required. Redundancy in the communication

system is important due to the dangers involved in a loss of communication. For this reason, two

transmitters and two directional antennas will be used. The communication system must also in=

dude RF subsystems for communications between the craft and Earth and/or the space station

(using S-band).

System Protocol and Coding

Although most of the sub-systems of the LOCOST vehicle will be controlled by the on-board

computer, the communication system must provide constant status for the subsystems. In the

event of a computer failure, the communication system must also be capable of handling the data

needed to control all of the vehicle's sub-systems.

Normally the optical system will be controlled through the APT system. Thus, all non-optical

systems will deliver system status and other information through the communications system.

Some of these systems include: position/stability mechanisms (RCS/CMG), electrical power control, propulsion control, fuel storage monitors, structural monitors, and the docking system. The

protocol of the communication system must handle the routine transfer of data from these systems

and be able to handle emergency situations. One way of handling so many systems is to use a

different radio frequency band for each sub=system. Because of the increasing number of communication links in space, and the possible crowding of radio frequencies in the future, the communication system of the LOCOST system was designed to use the least number of frequency bands.

Only a single band is needed to relay the data for all the vehicle's subsystems ff multiplexing is used.

To facilitate the multiplexing of normal sub-system information along with emergency status information and commands, a coding system was developed which has been named Priority Interrupt

Protocol (PIP). This system uses a digital signal carried by an analog wave. A sample of this

protocol is shown in Figure 50.

Communications 87
1990 LOCOST Senior Design Project

Sample of" Priority Interupt Protocol (P.I P.)

Coding With Check-Sum Error Handling

Next Stabllity System

Message Sensors Status:

IoloIolo o000ooo1 oo_ tlo111ol

IIIo11111oloI j/

Check-Sum

CMG's at Error

65% Saturation Handling " " "

ooioooom ooooT_oo

Binary

Coding

1 o ! ,

1

/

Carrier

Wave

Figure 50. PIP Sample

The multiplexing done by the P.I.P. system is accomplished by sending data for each subsystem in

an ordered sequence. The communications processor will send each set of data in the following

format: new message flag, subsystem identification, message type, data, and a check sum error

handling number. Since the entire message is coded in a binary digital code, the error handling is

done by summing up the total number of binary l's in the message and sending this number as the

last part of the message. This is called check sum error handling. In the event of an emergency, a

universal interrupt code can be sent by any sub-system to the communications processor. This

interrupt code will override the normal sequence of messages and the critical sub-system wil be

Communications _-, _.. 88
1990LOCOST Senior Design Project

given immediate communication priority. The actual modulation of the radio wave is done by

Frequency Shift Keying. This simple technique is basically a binary frequency modulation, as

shown in Figure 50. Each sub-system is linked directly to the communications processor, which

sequencially sends each system's status to a coding modulator. The encoded signal is then sent to

a transmitter, antenna switcher and then to one of the antennas. Signals sent back to the LOCOST

vehicle will be collected by an antenna, amplified through a receiver, interpreted by the coding

demodulator and sent to the communication processor.

Antennas

The size of the antennas on the LOCOST vehicle is primarily governed by the transmitter power

and wavelength used. The required transmitter power and antenna diameter are estimated to be

150 W and 1.35 m, respectively (2 = 0.06 m). The placement of the antennas and communications

system can be seen in Figure 50.

Since the vehicle will use the laser station for its communication link, the pointing of the directional

antennas will be done throught the APT system. This can be done because the primary laser

mirror and the directional antennas will be pointing in the exact same direction. This eliminates

the need of an additional tracking and pointing system. Spacecraft status and performance will be

monitored by various sensors. Data collected from these sensors will be transmitted to the ground

control station by the communication subsystem using 2.3 GHz downlink frequency. The uplink

frequency of 2.4 GHz will be used for command functions.

Summary

The communications system has a total mass of 200 Kg. The system uses two parabolic directional

antennas, each with a diameter of 1.345 m. One omni-directional antenna with a length of 12 m

is also used. The transmitter used has a power of 150 Watts, and operates at a frequency of 5 GHz

with a wavelength of .06 m. The total power consumption of the communications system is 200

Watts.

Communications 89
1990LOCOST Senior Design Project

Electrical Power Sytems

Introduction and System Comparison

The power systems considered for the vehicle were radioisotopic generators, solar collectors, and

secondary batteries, reactors, radioisotopic generators, solar collectors, and secondary batteries.

The necessary characteristics for the power system axe fight weight, dependability, high reliability

and power density, long life, and maintainability. The average power requirements for the vehicle

are summarized in Table 47.

For the length of mission specified, fuel cells and secondary batteries are the best alternatives.

Longer missions would make the volume and mass penalties of these two systems too great. Solar

cells must be connected in panels or arrays to provide a large enough power ratio to make them

feasuble for use. For the power-to-mass ratio necessary for our vehicle, this array would be much

too large. This system would also depend too much on solar energy since our vehicle is not always

illuminated. Secondary batteries have high energy densities but are not a viable alternative because

the capacity of a secondary battery is governed by the size and weight of electrodes which store

chemical energy, which accounts for much of the size and weight of the battery. Radioisotopic

generators do not have high enough power densities to be usefuLl because of the low efficiency of

direct conversion processes which this system utilizes.

The fuel cell is a form of storage battery in which the chemical energy is stored as a fuel in a reactant

tank outside the cell and is fed to, or removed from, the electrodes when required. The electrodes

are not changed in any way when the cell is operated. The capacity of a fuel ceLl is govemed only

by the size of the fuel tanks and the battery size is related only to the rate of conversion of power

output. Fuel ceLls were chosen because they are mass competitive with other systems, utilize readily

available technology, are highly reliable, offer redundancy, have a high conversion efficiency over

a wide output range and have zero fuel consumption during dormant periods. Four fuel cells will

be used to power the necessary systems on the vehicle. One cell will be placed below the primary

mirror on the support truss. Two will be in the truss section between the hydrogen and oxygen

tank sections. The last will be at the rear of the vehicle on the lower part of the truss between the

two chemical engines.

Components

The fuel cells that have been most extensively developed use hydrogen as their fuel and oxygen as

an oxidizer. The main components of a fuel cell are the fuel supply system, regulation system (

preheater and regulator ), reaction cells, and the water bleed off system, as seen in Figure 51.

Electrical Power 90
1990LOCOSTSeniorDesignProject

Table 47. Vehicle Power Requirements

System

Communications

Data Management

ATP: Laser Peak Power

Induction Motors (2)

Propulsion: Start-up Motor

H2 Turbopump

02 Turbopump

Gimbal Actuator

Electric Fuel Pumps (2)

RCS: CMGs Quiescient Power

Optics: First Mirror's Motors (2)

Fourth Mirror Tourque Motor

Fifth Mirror Linear Actuator Motor

Heat Pipe System Aux. Heating Units

Structures: Rocket Basket Linear Actuator Motor

Radiator Pumps

Distribution Loss

Gyroscopes

Power (kW)

.2

.5

.2

.012

.1

.22

9.4

.01

2.08

.5

.15

.2

.3

1.4

.003

.15

.05

1.2

Total 16.44

Electrical Power 91
1990LOCOST Senior Design Project

ra_alor

Figure $1.

liquid hydtonenn pump r--

,.:qu_ i

_0reheltlt

>

fuel ceil stacks

_o

coolut c_;culatJoa :ubm I lt.O ,,._ct

Basic Components of a Fuel Cell Power Generator

A "stack" of a number of cells connected in seriesis used in conjunction with other stacks in parallel

to give the required power. The maximum current output of a stack depends on the individual cell

membrane area and max. allowable current density of the membrane.

The reaction cell consists of an anode and cathode region, each filled with electrolyte, at which the

ion exchange occurs. Three plates are needed. The hydrogen metering, oxygen metering, and

sealed separation plates. The electrode support or current carrying plate is made of corrosion protected magnesium. This positions the electrodes, serves as a manifold to distribute the gas to the

electrode area, and conducts current externally.

Preheater System and Cooling

A percentage of the waste heat will be employed in a preheater system which raises the temperature

of the cryogenic fuel to the operating temperature of around 125 degrees celcius. A radiator will

be used to cool the fuel cell. The mass of the radiator is approximately 50 kg.

Water Removal System

The components of the water removal system are water transport membranes saturated with 45%

potassium hydroxide, two pourous support plaques, and a plastic water removal plate similar to the

support plate. Two end plates and tie bolts complete a single cell.

Electrical Power 92
1990LOCOSTSeniorDesignProject

Fuel Cell Housing

The fuel cell is housed in a double walled aluminum shell to provide protection from meteroid

strikes and radiation. The walls are. 15 cm thick and there is a separation distance of 10 an between

the inner and outer walls. The shell provides 99% particle protection and rejects or dissipates 97%

of incident radiation. The inner shell has a 1 m diameter, is I m tall and weighs 118 Kg. The outer

shell has a diameter of 1.2 m, is 1.2 m tall, and weighs 149.3 Kg. The mounting weighs 10 Kg.

This system will provide a very long lasting, efficient supply of power for the LOCOST vehicle.

Electrical Power 93
1990LOCOSTSeniorDesignProject

Appendix A. Mirror Cooling Calculations

The Heat Pipe

Figure of merit calculation for heat pipe fluids

Entrainment Limit:

2 N/2ippv°l cos 0 ent = nr vL ,t

Where

2 is the characteristic dimension of liquid/vapor interface for fine mesh (.036 m)

o_ is 0.086 N/m

Hence, Q,,t for Na is 423,086 kW and for Li it is 455,243 kW. These axe both well above the required heat transport capability.

Axial Heat Transport:

M =

Where

Pl is 430 kg/m 3

sigmal_ is .275 N/m

L is 20,000 kJ/kg

At a temperature of 1200 K, Mu is 10,282,608, while that for sodium is (MN,) is 1,371,994.

Boiling:

Boiling which may result in vapor blocking in arterial heat pipes--bubbles in artery is a problem.

Hence, a high superheat ( A T) is needed to reduce change of nucleation.

Ol

M=ATs- Lo v

At 1200 K, Na yields 0.0001 while Li produces 0.0049.

Radius of heat pipe

Qmax--.4= x .73L x _r 2

= 10 MW allowable heat input rate

Where

Pr is 9.59 Bar = 9.59x 105 kg/m _

L is 3,577,000 J/kg

Hence, r, is 0.045 m "-"5 cm.

Thermal gradients:

Sample calculations

Appendix A. Mirror Cooling Calculations 94
1990 LOCOST Senior Design Project

10MW times how long beam on mirror (3 hr.) = Q the total energy absorbed

Q

m (mass of mirror) = q sepecific energy absorbed

q = T temperature difference at front of mirror from original state

specific heat

Using the heat conduction equation,

Assume linear distribution of temperature through mirror. This gives thermal gradient through

mirror versus time.

E=oxaxT a

Radiator Size Calculations:

All cooling requirements (156.2 kW + 159.8 kW + 161.5 kW) -- q. Hence, q is 477.5 kW. Then

477500 = .85 x 5.67 x 10-8 x A x T a

= 76.5m 2

Support Calculations:

Maximum allowable deflection of 2.nd mirror. Assume

&4<.l%xA

Hence, zL4 is 0.00095033 ma

From

Z_,(avg) 2 + 2.rZkr(avg) - 2/_ -0

Hence, dLris 0.0012099 m.

Then, from

_'=dO

0 is 0.00014068 radians. The deflection can be expressed as 6 = Ox = 0.00014068. For a 9.9 kW

loss of power, 6 -- 0.000985 meters.

Appendix A. Mirror Cooling Calculations 95
1990 LOCOST Senior Design Project

Appendix B. NOTS Input and Calculations

The following equations and numbers are required input for the NOTS program.

The entering temperatures are:

Tentering 02 = 99.850K

Tentermg //2 = 267.07"K

The entering enthalpy:

H = Hf+ (SPAT + Hf_t

= O + .4(90 - 298) + (-50.9) = -4300

cal

k'gmole

The specific thrust:

T

lsp- rha

= 5835.59 rn

s

The maxmirnum thrust in g's where lg = 9.8 m/s:

gmax = WTotal/Empty mass = .51gs

The thrust coefficient:

2k 2 ? k+l po k-1 1 Pe Ae

CF=(-'k-Z-i-(-_--i-)k-IEl--(-_-# )T]IT+ Pc At - 1.955

Geometry calculations:

AT- T

6 F : Thl'ust Efficiency = .99

2500

AT= .99 x 1.955 x 10376975 =1.24x10 -am 2

°:44 T _o398m

A e = _AT where ,- = Area Ratio

Chamber Calculations:

TgRT c

hnjc_torPcVe MW

= .O12m 2

25000 x 9.8 x 3602 x 8314

4840 x 13.412 x 10376975 x 91.44

L'(typical) = 1.016m

Appendix B. NOTS Input and Calculations 96
Volumechamber = L'A,

Nozzle half magic = 15°

Lc = De(Length Ratio) = .373 lm

Wall thickness:

For a cylinder under radial pressure:

Ap D = 2twS

yield stress 25000

S = working stress = factor of safety 2

Pump parameters:

Turbopumps

rhH x SG

Power -

k

For turbopumps, k = 1 in metric units.

the,,= .76 kg/s

AP/= 9.9 MPa

H = 1.9 m = pumphead

P = .11 kW

rnox = 4.55 kg/s

APox = 9.6 MPa

H = .9m

SG = 1.139

P = 4.7kW

Electricpumps

P=rhHSG

APo; = 1.9MPa

APH2 = 7.16MPa

Po_ = 4.55 x .2 x 1.139 = 1.0365kW

PH2 = .759 x .72 x .0708 = 38.69W

- 125000psi.

1990 LOCOST Senior Design Project

Hence, t. = 3ram.

Appendix B. NOTS Input and Calculations 97
1990LOCOSTSeniorDesignProject

Appendix C. Thermal Balance Calculations

The thermal balance for the structure is calculated using an iterative process with the following

equations:

qA + ql + qR

qa = A s + qa + qe

The heat due to the sun:

(P- ts)

qs = S"sA s p

Where

S is the solar constant = 1350 w/m 2 (watts per square meter)

_, is the absorptance of the exterior surface for solar energy = 0.25

As is the cross sectional area exposed to sunlight

P is the orbital period

ts is the time in Earth shadow

The heat due to the albedo:

¢, a COS

qa = °a%_e _ p

IrR"

V_erc

So is the albedo = 0.345 x S

_ois the absorptance of the exterior surface for albedo radiation

A, is the cross-sectional area exposed to radiation from earth

cos a where _ is 22 °, the angle of the sateUite's orbital plane

R is the arbital radius in eaxth radii

The heat to the earth's radiation:

PWEarAr

qe - R2

Where

WE is the average flux of earth generated radiation = 240 W/m a

_o is the absorptance of exterior for Earth radiation - 0.20

ql is the heat radiated from the back of the mirror. The internal heat:

ql= P x M x am x Am

Where

M is the heat flux from back of mirror = 2390 W/m 2

A,, is the cross-sectional area exposed to radiant energy from mirror

a, is absorptance of the exterior surface for mirror radiation = 0.30

The dissipated heat:

Appendix C. Thermal Balance Calculations 98
qR=_x_xARxT a

1990 LOCOST Senior Design Project

Where

,, is the Stefan-Boltzman constant

is the material emmittance at temperature T =

AR is the "effective" area for heat radiation

T is the absolute temperature in degrees Kelvin

The area for heat absorptance for a cylinder (As.A.) is

A =DxH

0.20

Where

D is the diameter of cross-section of individual strut = 0.00025 ma

H is the sum of all strut lengths exposed to a particular radiation

The area for heat radiation for a cylinder (Aa)

(. x D 2)

A =(_rxDxH)+ 2

Where

H is the sum of all the lengths of all struts with incident radiation.

Appendix C. Thermal Balance Calculations 99
1990 LOCOST SeniorDesign Project

Appendix D. Propellant Tank Calculations

Mass of LH2 required = 52769 kg

Mass of LO_ required = 40231 kg

Density of LH2 = 70.8 kg/m 2

Density of LO_ = 1139 kg/m 3

Total volume required LH_ = 749 m 3

Total volume required L02 = 35.3 m 3

Spherical L02 tanks

V = 413_R 3

S = 4_rR2

Per tank:

V = 17.66 rn3

R = 1.62 m

D = 3.23m

S = 32.8 m

For A1 2090:

E = 7.584 x 104 MPa

a.lt = 2.237 x 104 MPa

e = 3547 kg/m _

Pr

m

F.S. 2t

Where

Om._= max stress

p = tank pressure

r = radius

t ffi thickness

The thickness is t = .067 ram.

From, mass of tank = surface area (t) e, the mass of the liquid oxygen tanks axe 5.6 kg.

Cylindrical LH2 tanks.

R = 3.14m

D = 6.28 m

S = 3.25 nD 2 = 402.67 rn_

Use the same calculation for the liquid oxygen tanks to obtain the mass of the hydrogen tanks.

Appendix D. Propellant Tank Calculations 100
1990LOCOSTSenior Design Project

Insulation Dimensions and Mass

The same calculations apply to both the liquid hydrogen and the liquid oxygen tanks. The liquid

hydrogen calculations are shown below.

Given the condition of I percent total boiloff from 26384.8 kg of fuel, the heat flux q for this boiloff

to take place is calculated using

q0

My f- hfz

Vv'her_

M,, is the mass of fuel vented

0 is the the length of mission -," 2 weeks

ht_ is the heat of vaporization of LH_ (499 kJ/kg)

q Is the total heat transfer

Then q = 64.857 W.

The heat flux due to solar radiation on the tank:

i! = Aq'%

'Where

is the heat flux

A is the area exposed to solar radiation (201.34 m_)

q" is the solar radiation constant (1400 W/m 2)

is the emissivity of tank (.06 since paint is 94% reflective)

Hence, q is 16.91 kW.

Use the Stefan-Boltzman law to fred the temperature on the surface of the tank:

= oAT 4

Where

q is the total heat flux (16973.9 W)

o is the Stefan-Boltzan constant (5.67x I0 -a W/m_K 4)

A is the area exposed to solar radiation

Hence, T = 196.4" Kelvin.

The MLI thickness can then bet determined from the equation:

kA(7"_ - 7"2)

q= t

Where

k is the thermal conductivity of insulation (.03 W/InK)

7"1is the temperature of the tank surface

Ta is the interface temperature of the MLl/foam (144 K).

Hence, tma is .0186 m.

From

tf°am = .1

tMLI

tt.,= is .186 era.

The densities of the MLI and foam are PMta = 35.08 kg/m 3 and Or.,= = 30 kg/m 3.

Appendix D. Propellant Tank Calculations 101
1990 LOCOST Senior Design Project

Hence, the total mass for the liquid hydrogen per tank is

m = S(tfoamProa m +/MLlPMLI)

= 570.98kg

The mass of the support structures can be obtained from the relation

Where Pmo/,0o_ = 1685 kg/m 3, and P60601.r6= 2550 kg/m 3.

Appendix D. Propellant Tank Calculations 102
1990LOCOSTSeniorDesignProject

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