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RRS-026

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Reusable Reentry Sateflite (RRS)

Summary Report

Propulsion System Trade Study

_--:_

April 1990

Contract NAS9-18202

DRL 02

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Prepared for:

National Aeronautics and Space Administration

Lyndon B. Johnson Space Center

Houston, Texas 77058

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_ Science Applications International Corporation

21151 Western Avenue • Torrance, California 90501 • (213) 781-9022

TOR4,6/23.3
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FOREWORD

The Reusable Reentry Satellite (RRS) Propulsion System Trade Study described herein

was performed during Part 1 of the RRS Phase B contract. This report is one of several that

describes the results of various trade studies performed to arrive at a recommended design for the

RRS satellite system. The overall RRS Phase B study objective is to design a relatively

inexpensive satellite to access space for extended periods of time, with eventual recovery of

experiments on Earth. The RRS will be capable of: 1) being launched by a variety of expendable

launch vehicles, 2)operating in low earth orbit as a free-flying unmanned laboratory, and

3) executing an independent atmospheric reentry and soft landing. The RRS will be designed to

be refurbished and reused up to 3 times a year for a period of 10 years. The expected principal

use for such a system is to research the effects of variable gravity (0-1.5 g) and radiation on small

animals, plants, lower life forms, tissue samples, and materials processes.

This summary report describes the RRS Propulsion System Trade Study performed to

select an appropriate propulsion system design for the RRS. The weight, performance, reliability

and complexity of a wide range of different propulsion systems were compared to select an

optimum system for the RRS based on its unique set of requirements.

The Propulsion System Trade Study was performed under the contract technical direction

of Mr. Robert Curtis, SAIC Program Manager. The study was directed by Mr. Steve Apfel who

was assisted by Ms. Chris Scheil, both of SAIC. Mr. Michael Richardson, JSC New Initiatives

Office, provided the RRS objectives and policy guidance for the performance of these tasks under

the NAS 9-18202 contract.

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CONTENTS

Section

FOREWORD ......................................................................................

CONTENTS .......................................................................................

LIST OF FIGURES ..............................................................................

LIST OF TABLES ................................................................................

LIST OF ACRONYMS ...........................................................................

EXECUTIVE SUMMARY ......................................................................

1.0 INTRODUCTION ..........................................................................

1.1 Background ......................................................................

1.2 NASA JSC Statement of Work Task Definition .................................

1.3 Scope ..................................................................................

2.0 STUDY APPROACH ..................................................................

2.1 Organization ..........................................................................

2.2 Document Format ....................................................................

3.0 PURPOSE ...................................................................................

4.0 GROUNDRULES AND ASSUMPTIONS .............................................

5.0 ANALYSIS METHODOI_X)GY ..........................................................

6.0 TOP-LEVEL SYSTEM TRADES ........................................................

6.1 Solid vs. Liquid System ............................................................

6.2 Jettisonable Propulsion Module ....................................................

6.3 Thruster Firing Direction ............................................................

7.0 LIQUID SYSTEM DESIGN DATA .....................................................

7.1 Propellant Consumption and System Sizing ......................................

7.1.1 Propellant Consumption ...................................................

7.1.2 Propellant Tank Sizing .....................................................

7.2 Propellant Management Devices ...................................................

7.3 Residuals ..............................................................................

7.4 Number of Engines and Layout ....................................................

8.0 PROPULSION SYSTEM DESIGN OPTIONS ........................................

8.1 Monopropellant System .............................................................

8.2 Bipropellant System .................................................................

8.3 Dual-Mode Bipropellant System ...................................................

8.4 _2 ..............................................................................

8.5 Advanced Technology Propulsion System .......................................

9.0 TRADE STUDY RESULTS ..............................................................

10.0 CONCLUSIONS / RECOMMENDATIONS ...........................................

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LIST OF FIGURES

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Monopropellant Propulsion System Schematic ....................................... 22

Bipropellant Propulsion System Schematic ........................................... 23

Dual Mode Propulsion System Schematic ............................................. 24

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Table

4-1

7-1

7-2

7-3

7-4

7-5

7-6

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LIST OF TABLES

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RRS Design Reference Missions ....................................................... 6

Monopropellant Propellant Budget - Blowdown ..................................... 11

Monopropellant Propellant Budget - Pressure Regulated ........................... 12

Bipropellant Budget - Blowdown ...................................................... 13

Bipropellant Budget - Pressure Regulated ............................................ 14

Dual Mode Propellant Budget - Blowdown .......................................... 15

Dual Mode Propulsion System - Pressure Regulated ................................ 16

Dual Mode Advanced Design Propellant Budget ..................................... 17

Gaseous Oxygen/Hydrogen Propellant Budget ....................................... 18

Propulsion System Dry Weights in Lbs ............................................... 26

Propellant Usage by System for Selected Deorbit Delta-Vs ......................... 27

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ACS

DRM

ETR

GN&C

JSC

PM

PMD

RRS

RRV

SAIC

SDI

SRD

TPM

WSMR

WTR

LIST OF ACRONYMS

Attitude Control System

Design Reference Mission

Eastern Test Range

Guidance, Navigation, and Control

Johnson Space Center

Payload Module

Propellant Management Device

Reusable Reentry Satellite

Reusable Reentry Vehicle

Science Applications International Corporation

Strategic Defense Initiative

System Requirements Document

Thrust Pulse Modulation

White Sands Missile Range

Western Test Range

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:: ; EXECUTIVE SUMMARY

The purpose of the RRS Propulsion System Trade Study described in this summary report

was to investigate various propulsion options available for incorporation on the RRS and to select

the option best suited for RRS application. The design requirements for the RRS propulsion

system were driven by the total impulse requirements necessary to operate within the performance

envelope specified in the RRS System Requirements Documents. These requirements were

incorporated within the Design Reference Missions (DRMs) identified for use in this and other

subsystem trade studies. This study investigated the following propulsion systems: solid rocket,

monopropellant, bipropellant (monomethyl hydrazine and nitrogen tetroxide or MMH/NTO), dualmode bipropellant (hydrazine and nitrogen tetroxide or N2H4]NTO), liquid oxygen and liquid

hydrogen (LO2/LH2), and an advanced design propulsion system using SDI-developed

components.

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Operational considerations such as jettisonable modules, thruster firing direction, and

propellant management systems were investigated in addition to performance, weight, cost,

complexity, and refurbishment trades for the systems listed above. Strawman system designs

were constructed for each system and systematically compared.

A liquid monopropellant blowdown propulsion system was found to be best suited for

meeting the RRS requirements and is recommended as the baseline system. This system was

chosen because it is the simplest of all investigated, has the fewest components, and is the most

cost effective. The monopropellant system meets all RRS performance requirements and has the

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capability to provide a very accurate deorbit burn which minimizes reentry dispersions. In

addition, no new hardware qualification is required for a monopropellant system. Although the

bipropeUant systems offered some weight savings capability for missions requiring large deorbit

velocities, the advantage of a lower mass system only applies if the total vehicle design can be

reduced to allow a cheaper launch vehicle to be used. At the time of this wade study, the overall

RRS weight budget and launch vehicle selection were not being driven by the propulsion system

selection. Thus, the added cost and complexity of more advanced systems did not warrant

application.

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1.0 INTRODUCTION

1.1 Background

As currently conceived, the Reusable Reentry Satellite (RRS) will be designed to provide

investigators in several biological disciplines with a relatively inexpensive method of access to

space for up to 60 days with eventual recovery on Earth. The RRS will be designed to permit

totally intact, relatively soft recovery of the vehicle, system refurbishment, and re-flight with new

and varied payloads. The RRS system will be capable of 3 re-flights per year over a 10-year

program lifetime. The RRS vehicle will have a large and readily accessible volume near the vehicle

center of gravity for the Payload Module (PM) containing the experiment hardware. The vehicle is

configured to permit the experimenter late access to the PM prior to launch and rapid access

following recovery.

The RRS will operate as a free-flying spacecraft in orbit and be allowed to drift in attitude

to provide an acceleration environment of less than 10 -5 g. The acceleration environment during

orbital trim maneuvers will be less than 10 -3 g. The RRS is also configured to spin at controlled

rates to provide an artificial gravity of up to 1.5 Earth g. The RRS system will be designed to be

rugged, easily maintainable, and economically refurbishable for the next flight. Some systems

may be designed to be replaced rather than refurbished if system replacement is cost effective and

able to meet the specified turnaround time. The minimum time between recovery and re-flight will

be approximately 60 days. The PMs will be designed to be relatively autonomous with experiments

that require few commands and limited telemetry. Mass storage if needed will be accommodated in

the PM. The hardware development and implementation phase is expected to begin in 1991, with a

first launch in 1993.

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Numerous trade studies and RRS functional design descriptions are required to define a

viable RRS concept that satisfies the requirements. NASA has contracted with Science

Applications International Corporation (SAIC) to perform a Phase B study to provide the RRS

concept definition. The Propulsion System Trade Study described in this report is one of the

supporting study analyses performed by the SAIC team.

1.2 NASA JSC Statement of Work Task Definition

The reentry propulsion system trade study was performed per the general direction of the

RRS Statement of Work and the System Requirements Document (SRD) as given in the following

paragraphs:

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General:

SOW Param'aph 3.1.2 Tradeoff Studies: "Conduct required tradeoff studies for each of the

areas and options listed below as well as others identified in the contractor's proposal or

which become apparent during the course of the study effort. The depth of analysis for each

individual option will vary as appropriate to clarify and document the viability of each

approach. In all of the following tradeoffs, particular attention should be given to effects on

the complexity, flexibility, or imposed constraints on the RRS design, RM design, or

mission operations. Also, special consideration should be given to system reliability and

operational safety as well as the reduction in program life cycle costs."

Specific:

SRD Para_aph 3.3.6 Attitude Control and Propulsion Subsystem: "The RRS Attitude

Control and Propulsion Subsystem shall provide the following capabilities:

(a) Perform attitude determination, stabilization, and control functions as required

throughout the orbital phase of a mission.

(b) Control RRS attitude rates as specified in Paragraph 3.2.3.3 (on-orbit acceleration

limits).

(c) Provide the necessary velocity changes for the deorbit maneuvers.

(d) Spin up and maintain the RRS at various rates to satisfy experiment requirements for

fractional g levels. Provide capability to despin for rmcrogravity levels in the same

flight.

(e) If required, spin up to TBD rpm for the deorbit thrust maneuver, and spin down to TBD

rpm for reentry.

(f) Correct launch errors in the orbital parameters and adjust the parameters to be compatible

with the recovery site requirements."

1.3 Scope

This NASA Phase B study is intended to provide definition of the RRS concept. The study

includes tradeoff studies with the depth of analysis as appropriate to clarify and document the

viability of each approach. The RRS system and operations are developed to the degree necessary

to provide a complete description of the designs and functional specifications. The propulsion

system trade described in this report was performed to ensure that the SAIC RRS design meets all

mission requirements yet remains as simple and cost effective as possible.

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2.0 STUDY APPROACH

2.1 Organization

The tradeoff analyses performed in Part 1 of the RRS Phase B study were organized to be

accomplished in a series of related but separate trade, off studies and system concept definitions.

Therefore, the documentation described in these summary reports has been formatted to

accommodate a compendium of analyses published in several separate documents. Because of the

synergistic nature of one tradcoff study across the entire RRS system design, it is suggested that

the reader review all summary reports in order to get a complete picture of the SAIC RRS design.

2.2 Document Format

Individual analyses and studies do not necessarily lend themselves to be documented in

exactly the same way; however, a general outline has been used where reasonable for all reports.

The guideline for preparing the individual study sections in this and all summary reports is

provided below:

• Purpose

• Groundrules and Assumptions

• Analysis Description

• Analysis Results

• Conclusions

• Recommendations

3.0 PURPOSE

The purpose of the RRS Propulsion System Trade Study was: 1) to investigate various

propulsion options available and 2) to select the option best suited for RRS application. During the

performance of this task, the weight, performance, reliability and complexity of a wide range of

propulsion systems were compared and analyzed Data for the selected system was then used in

other trade studies in order to develop the design definition of the SAIC RRS concept.

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4.0 GROUNDRULES AND ASSUMPTIONS

The overall requirements of the RRS propulsion system were discussed in Section 1.2 as

stated in the RRS SRD. In general, the RRS propulsion system will be used to supply all major

velocity increments and vehicle attitude adjustments after separation from the launch vehicle. The

system must be capable of providing highly accurate total impulse delivery in order to meet the

landing dispersion requirements. Since the propulsion system is one of the critical items for

reentry safety, the design must be fail operational / fall safe. The propulsion system, coupled with

the Guidance, Navigation and Control (GN&C) and power system redundancy, provides no

single-point failures in achieving a safe landing of the RRS with minimal safety hazard to the

public. The propulsion system must also deliver highly accurate total impulse as commanded by

the GN&C system, again to ensure that the landing dispersions are minimal.

The propulsion system must also provide impulses to spin up the vehicle for artificial

gravity missions. The limit of this impulse, determined by analyses discussed in the RRS

Configuration Summary Report, must be less than 1 lbf of thrust in order to prevent overstressing

the Astromasts in their extended configuration.

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From a science requirements perspective, the RRS propulsion system (and of course an

appropriate control system) must be capable of controlling on-orbit acceleration of the PM to

provide either a fractional gravity or a microgravity environment. The fractional gravity environment will subject experimental specimens to a steady artificial gravity acceleration of up to 1.5 g.

The artificial gravity level will be selectable at any value between 0.1 g and 1.5 g for any flight and

will be maintained within a range of +10%. The microgravity environment will subject

experimental specimens to an artificial gravity acceleration level which is less than 10 -5 g for at

least 95% of the Orbital Flight Phase. During the remainder of the flight, the artificial gravity

acceleration level shall not exceed 10 -3 g.

The propulsion system performance and tank sizing calculations discussed in this report

were performed using the design reference missions defined early in the study and shown in Table

4-1 as reference. The remaining general groundrule that was used in the propulsion system trade

was the desire to make the system as simple and cost effective as possible. To minimize development cost, space-qualified hardware was investigated and incorporated in designs studied

whenever possible.

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Definition

Parameter

Character

Inclination

Orbit Type

Orbit Altitude

Launch Site

Recovery Site

Table 4-1. RRS Design Reference Missions

Design Reference Mission Set

DRM-1

Land Recover},

33.83 °

Circular

DRM-2

High Altitude

33.83 °

Circular

DRM-3

Hi[h Inclination

98°

Circular,

Near-Integer

DRM-4

Integer Orbits

35.65 °

Circular,

Integer

DRM-5

Water Recovery

28.5 °

Circular

350 km 900 Ion

(189 nm) (486 nm)

Eastern Test ETR

Range (nTR)

White Sands

Missile Range

fWSMR)

WSMR

897 kin

(484 nm)

WTR

WSMR

479 km

(259 rim) .....

ETR

WSMR

350 km

(189nm)

ETR

Water (ETR,

Gulf of

Mexico, WTR)

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5.0 ANALYSIS METHODOLOGY

The analysis methodology used for the propulsion trade studies consisted of several phases

of analysis. The design requirements were determined by the DRMs (total impulse required) and

the design goals of the SAIC design. Top-level design goals were sufficient to perform a general

evaluation of potential types of propulsion systems that could be adapted for RRS application. The

purpose of this top-level screening was to determine the pros and cons of each type of system

relative to RRS system requirements. After evaluating these parameters and selecting a general

system type, a more detailed investigation of the option was performed. Specific issues were

investigated and traded such as performance, weight, cost, complexity, and refurbishment

potential. Operational considerations such as jettisonable modules, thruster f'wing direction and

propellant management systems were also investigated. Final recommendations for the baseline

RRS propulsion system were made at the conclusion of the study and discussed in this report.

6.0 TOP-LEVEL SYSTEM TRADES

6.1 Solid vs. Liquid System

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The first propulsion system trade performed was to select the overall type of system (solid

or liquid) best suited for RRS application. The RRS Ames Phase A study recommended a propulsion system that included three STAR 17 solid rocket motors for the main deorbit bum along with

a monopropellant hydrazine system for attitude control, spin control and general trim bums.

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The solid rocket motor deorbit system studied during Phase A provides relatively

inexpensive impulse for the deorbit maneuver; however, a number of performance limitations are

inherent in its use. The total delivered impulse accuracy for solid motors is _+0.5%. This number

is driven by motor temperature differences as well as motor to motor variations. Providing

accurate total impulse with this type of system requires a thrust termination system to shut off the

solids, which is not trivial. The incorporation of the hydrazine liquid system, required for spin

speed control and ACS, could be used to take out impulse errors associated with the solid rocket

system, but this solution requires two systems to perform a single function.

The motor-to-motor differences mentioned above also relate to differences in delivered

thrust. Averaging these differences requires either gimballing the solid motor exit cones or

spinning the vehicle to average out the differences. Gimballing of exit cones is performed on large

solids for launch vehicles and orbit transfer stages. However, it is expensive ($millions) and

would require design modifications be made to the STAR 17 motor. Spinning the vehicle, as well

as the solid motor bum itself, impose accelerations (2 to 3 g for the deorbit bum, 0.5 to 0.75 g

for the spinning maneuver at 30 rpm) in a direction different from other phases of the flight. The

SAIC design goal is to try to maintain the acceleration forces in the same direction for all phases of

flight.

From a reusability and refurbishment standpoint, solid motors have to be replaced after each

flight. The motor cases might be reusable, but the overall cost savings would be minimal. Solid

rocket motors are designed and loaded to deliver a total impulse. This makes a given motor

mission specific. Different reentry requirements, due to different orbital altitudes, would require

different motors for each mission. Operationally this would mean different motors must be kept on

hand at the refurbishment site to meet the 60-day re-flight capability. Finally, and most

importantly, the solid system does not meet the fail operational/fail safe design goal of the SAIC

design. A solid motor failure could greatly increase the landing dispersions (and it is questionable

if the vehicle would reenter at all) due to the loss of a large percentage of the total impulse and

coning of the vehicle induced by the thrust imbalance. Such occurrences may not be correctable,

thus creating a potential safety hazard.

Liquid propulsion systems, on the other hand, have higher initial cost compared to a solid

systems, but are reusable and provide much more flexibility. The liquid systems pulse capability

allows it to be tied into the attitude control system. This, coupled with accurate total impulse

accuracies (<__+0.01%) and various control techniques (e.g., active nutation), can minimize

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dispersion error sources inherent in the propulsion system. Such a system has the potential to

provide very accurate deorbit burns. A liquid system also has the potential to perform a trim burn

before reentry interface as an added safety backup. From a safety perspective, liquid systems for

satellites typically are designed with fall operational/fail safe capability. This takes the form of

redundant engines and dual coil valves.

For the above reasons, it was decided that a liquid deorbit propulsion system was much

more desirable for RRS application than a solid system postulated in Phase A. The inherent

capability to combine the attitude control system with the main system was also a benefit. Further

discussion on the evolution of the liquid system proposed for the RRS is given in subsequent

sections.

6.2 Jettisonable Propulsion Module

The proposal configuration for the SAIC design concept incorporated a jettisonable

propulsion module. The module served as the interface connection plane to the launch vehicle and

housed structural and subsystem components for the propulsion system. The module remained

attached to the RRS until completion of the deorbit burn, at which time it was jettisoned to reenter

separately from the RRS. The rationale for this design was that it allowed for easier structural

interfaces with the launch vehicle and was a simple way to implement forward firing thrusters.

Forward firing thrusters were desired in order to maintain a monodirectional acceleration force on

the vehicle at all times during the mission, as discussed in Section 6.1. The thrusters required to

perform in this manner were located outside the radius of the RRS heat shield with exit cones

oriented towards the nose of the RRS (i.e., forward firing).

This concept was subsequently modified for two reasons. The first was the inability to

positively quantify that there would not be a public safety hazard using this concept. The

propulsion module will contain components that may not burn up during reentry, thus creating a

potential safety hazard. A potential solution to this problem was investigated that entailed spinning

up the propulsion module after separation from the vehicle and f'Lring its main engines one or more

times to ensure that the propulsion module debris footprint falls over water only. This spun

configuration would require a simple control system, possibly as simple as a set of timers.

However, the major detractor from this approach was the unknown impact relative to the

propulsion system design. Different orbits and landing sites could have a large impact on the total

velocity increment required to perform the safe reentry of the propulsion module. The analysis and

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testing required to verify the propulsion module reentry footprint and to determine its effect on the

propulsion system design was felt to impact the RRS program too much to be considered further.

The second reason the original concept was modified was the cost associated with

replacement of the propulsion module for each flight. Depending on the type of propulsion

selected, the cost of the module could range from $2 to $4 million each. Life cycle cost impacts of

$50 to $110 million could be accrued over a lifetime of 30 flights. Even though this amount is

small compared with the launch costs of each flight, it was felt to be of importance nevertheless.

For these reasons, the decision was made to make the propulsion system integral with the vehicle.

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6.3 Thruster Firing Direction

As discussed briefly above, the proposal also called for deorbit thruster f'nings towards the

nose of the vehicle in order to maintain all accelerations in the same direction during all phases of

the mission. The forward f'ning thrusters also provided a means of safely de.orbiting the vehicle in

the event of an Astromast failure to retract. The deorbit maneuver could still be performed as the

masts would be in tension, "pulling" the main module. Since the masts do not have sufficient

bending stiffness to allow "pushing" the main module, this concept was preferred at the time. This

design also assumed a jettisonable propulsion module as discussed in the previous section. Since

early propulsion module trades determined that the module should be integral to the vehicle, direct

forward f'ning thrusters now presented a slightly different design problem.

Various design options were proposed and investigated to maintain the unidirectional g

feature in the design including thruster cavities (nozzles) in the aft end of the heat shield, retractable

booms and partial retraction of the Astromasts. In general, all concepts were found to be possible

but were considered high risk in terms of development and/or reliability. Subsequent discussions

with life science personnel about the desire to keep the unidirectional g feature vs design

complexities resulted in a compromise baseline of more conventional aft fining thrusters being used

for the deorbit burn. The scientists considered the reversal of g load direction during the deorbit

maneuver noncritical considering the time and magnitude of the burn (250 seconds maximum at

less than 0.3 g) proposed by the SAIC design. Thus a conventional approach of mounting the

main deorbit thrusters on the rear face of the vehicle was adopted.

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7.0 LIQUID SYSTEM DESIGN DATA

This section describes the kinds of data collected and provides scoping calculations

performed to devise various propulsion systems, described in Section 8, which were evaluated for

RRS application. Section 9 provides trade study results.

7.1 Propellant Consumption and System Sizing

7.1.1 Propellant Consumption

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The propulsion system was sized assuming a 350 m/sec (1148 ft/sec) deorbit bum from the

high altitude (DRM-3) case. This maneuver provides enough total impulse to bring the vehicle

down at a relatively steep angle from a 900 kilometer orbit in order to minimize dispersions. The

system was also designed to provide margins for nominal missions that may require orbit

adjustments to optimize the ground track for reentry. Miscellaneous velocity increments (other

than deorbit and spinup/spindown) were estimated based on drag, vehicle configuration, outgassing and orientation requirements. This maneuver history, along with estimated vehicle mass

properties and typical engine performance (Isp) values, was input to a simple spreadsheet model

that calculated total propellant load required for various types of propulsion systems investigated.

The calculations were performed for a monopropellant, bipropellant (monomethyl hydrazine and

nitrogen tetroxide or MMH/NTO), dual-mode bipropellant (hydrazine and nitrogen tetroxide or

N2H4/NTO), liquid oxygen and liquid hydrogen (LO2/LH2) and an advanced design propulsion

system using SDI-developed components. The propellant load calculations for a single case for

each system (pressurized and blowdown where applicable) are shown in Tables 7-1 through 7-8.

Table 7-1 shows the propellant budget for a monopropellant blowdown propulsion system.

This case documents a 350 rn/sec (1148 ft/sec) deorbit bum or a fully loaded case. The first

column gives the maneuver type, the next the magnitude in ft/sec for linear motion, RPM if it is a

spin up or down and degrees for reorientation of the vehicle. The next two columns give the

specific impulse, or Isp, and the duration for the maneuver. The total fuel consumed and launched

mass are given next with the final columns displaying moments of inertia in two planes and finally

comments. The bottom rows give details on the engine characteristics including the number of

engines and representative performance at pressure.

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Table 7-2 shows data similar to Table 7-1 only for a monopropellant system pressure

regulated at a constant 235 psia for the entire mission. Table 7-3 gives the budget for a bipropellant blowdown system. In this type of system, engine performance drops at the end of the mission

due to the lower inlet pressures. This lower engine performance evidences itself in the propulsion

budgets as lower average specific impulse for a blowdown system (Table 7-1) vs. a pressure

regulated system (Table 7-2). The deorbit velocity for this case is 100 rn/sec or 328 ft/sec, which

is representative of a DRM- 1 reentry. Table 7-4 gives the budget for bipropellant system pressure

regulated with a 350 m/sec or 1 i48 ft/sec deorbit burn. Again this calculation is representative for

DRM-3. Table 7-5 gives the budget for dual mode blowdown system with a 100 m/sec or 328

ft/sec deorbit burn. Table 7-6 gives the budget for a dual mode system pressure regulated with a

350 rn/sec or 1148 ft/sec de, orbit burn. Table 7-7 gives the budget for an advanced design dual

mode blowdown system assuming a 100 m/see or 328 ft/sec deorbit bum. This system uses 600

psi engine inlet pressures to achieve smaller and lighter componentry. Only a pressure regulated

system was investigated for this option. Table 7-8 gives the budget for GOX/GH pressure

regulated system assuming a 350 m/sec or 1148 ft/sec deorbit burn. This system used the smallest

propellant load for all DRMs. ACS engines were assumed to consist of 1 lbf cold gas engines and

0.1 lbf hydrogen resistojets for low thrust maneuvers (spin up and spin down).

7.1.2 Propellant Tank Sizing

Initially only existing tanks made of titanium were used in system design and weight

calculations. The tanks chosen from this somewhat limited list, however, turned out to be much

larger than required for RRS missions. This resulted in a larger volume and weight than would be

achieved with a more optimally sized system. Therefore, a second design iteration was performed

assuming custom designed carbon wrapped aluminum tanks for the He pressurization tanks, and

optimally sized main propellant tanks. The weights for this type of configuration were estimated

based on existing tank information and scaling up or down as required to match mission

requirements. Further discussions on this topic is provided in subsequent sections.

7.2 Propellant Management Devices

A number of different propellant management devices (PMDs) were studied for the various

propulsion systems but the two prime types considered were bladder and surface tension devices.

Bladders are thoroughly fiight-proven and are relatively inexpensive to use. Bladder tanks are also

insensitive to the direction of acceleration. They have a high expulsion efficiency, resulting in a

low propellant residual, and are easy to test. The negative for bladders, however, is that bladders

have a higher weight than other PMDs and are incompatible with many propellants.

-19-
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Surface tension devices have been used extensively in long-life satellites. They have a low

weight and are compatible with all propellants. However, surface tension devices have a higher

propellant residual, due to their lower expulsion efficiency, when compared to bladders. They are

also expensive, difficult to test, and are sensitive to the direction of vehicle acceleration.

A third option considered for propellant expulsion was the use of an auxiliary propulsion

system to produce an acceleration on the vehicle that will orient the propellant over the tank outlet.

After the propellant has settled, the engine being supplied must maintain the same vehicle

orientation. There is no tank volume penalty, and terminal draining of the tank determines the

propellant residuals. This method is an excellent option for propulsion systems such as the dualmode bipropellant, which already has a monopropellant attitude control system (ACS) system that

can easily be used to settle the oxidizer required by the main engines. If the propulsion system is

not originally equipped with this expulsion capability, the addition of a separate ACS system to

perform this function is usually excessively complex.

A final option considered for propellant expulsion was the use of rolling diaphragm tanks.

These incorporate a thin aluminum or stainless steel diaphragm that expels the propellant. The

advantages of this type of design are it can be used with any type of propellant. The tanks are not

cost effective for the RRV however, as they must be removed from the vehicle and a new

diaphragm installed after each flight. In addition, the tanks do not adapt to less than a 95% fraction

fill as the diaphragms are susceptible to breakage during launch vibrational loads if not filled.

7.3 Residuals

Propellant residuals are a function of the type of propellant, type of PMD used, temperature

extremes during operation of the system, expulsion efficiency of the propellant tanks and total

system line lengths. Bipropellant systems typically have a much higher residual than

monopropellant systems for a number of reasons:

The vapor pressure of the nitrogen tetroxide (NTO) is roughly 14.7 psi at 70 ° degrees F.

Temperature shifts from the the loading temperature causes the total pressure in the

oxidizer tank to rise or fall. This causes a fuel rich or fuel lean bum condition. For the

RRV, the main burn consumes -85% of the total mission propellant. Therefore the

temperature of the propellant at the end of the mission is critical for residuals. This

effect is exacerbated in a blowdown system as the vapor pressure of the propellant is a

higher percentage of the total pressure.

• The high vapor pressure and dense vapor means a high amount of propellant will be in

vapor phase in the tank.

• Twice the number of lines filled with propellant.

-20-
These factors must be taken into account when figuring the performance of a bipropellant

system. The residuals were assumed to be 20 lbs for the bipropellant system studied in this

analysis. These numbers assumed a 70 ° F loading temperature and an end of mission temperature

of 95 ° .

The monopropellant system investigated incorporated bladders for propellant management.

The bladders isolate the pressurant and the propellant minimizing any vapor residual. Therefore the

residuals were assumed to be 2 lbs for the monopropellant system.

B

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7.4 Number of Engines and Layout

All RRS propulsion systems evaluated were designed assuming 12 ACS engines in

redundant banks of 6 (2 pairs of 3 on opposite sides of the vehicle). The design of the main

system had 6 engines in 2 banks of 3 each. This provided for pairs, or couples, of engines for

performing all pitch roll and yaw maneuvers. All "banks" of engines were isolateable via bistable

latching valves. This provided for an engine-out capability with no loss in vehicle performance;

however, cross coupling would slightly increase propellant consumption for attitude control. The

implementation of the redundant set of engines could be either manual or autonomous depending

on the maneuver. The critical deorbit maneuver requires autonomous engine-out logic. A worst

case engine failure would be a valve stuck in the closed position. This effectively prevents this

engine from further operation. Open valve failure is handled by series redundant thruster valves.

The attitude control system would sense this by the vehicle attitude not responding to a command

thruster pulse. Depending on the location of the failed engine in respect to the vehicle center of

gravity, the deorbit burn would be completed with 3 to 5 of the remaining engines with minimal

loss in burn accuracy. The main maneuver engines were assumed to provide 80 to 110 lbs of

thrust depending on the system design. The weights for these elements, discussed below, were

taken from existing flight proven hardware.

8.0 PROPULSION SYSTEM DESIGN OPTIONS

The following sections describe point designs for various RRS propulsion systems which

are analyzed in more detail in Section 9.0. Propulsion system features are discussed, along with

schematics for each option, to determine the number of valves, filters, regulators, etc., required.

The weights of these components were taken from existing flight proven designs used on current

commercial satellite programs.

-21-
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8.1 Monopropellant System

A hydrazine (N2H4) monopropellant system was the simplest system considered in the

trade analysis. The propellant tanks were assumed to be made of titanium and use bladders as the

PMD because of their high expulsion efficiency and their proven effectiveness in zero-g

applications. The large propellant load for this system results in large tank volumes and has the

highest overall propulsion system weight of all those considered, for both the blowdown and

pressurized systems. The minimum operating temperature of 40 ° F for hydrazine engines may

require the use of heaters to keep the propellant above this, leading to an increase in the overall

power requirement. The hydrazine thrusters are capable of a large blowdown ratio. This allows

for smaller (relative to biprop systems) tankage with a lower starting ullage volume for blowdown

systems. The blowdown systems were designed with an initial tank pressure of 350 psia. The

maximum loading gives a final tank pressure of approximately 70 psia. The regulated or

pressurized system has the regulator set at 235 psia. A schematic of the monopropellant system is

shown in Figure 8-1.

T _ _ _ I OPTIONAL

(He _ ( He _ ( He _ I PRESSURE

' _ _ V I REGULATION

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L A T C H 'NG V AL VE '_l'lll_'jl_:lll ', ,A _, ,,_tr],

|00 LBFMONOPROPELLANT ENGINES

Figure 8-1. Monopropellant Propulsion System Schematic

-22-
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8.2 Bipropeilant System

The bipropellant system analyzed incorporates N204/MMH for both the deorbit and ACS

engines. Titanium propellant tanks were assumed using surface tension devices since N204 is

incompatible with bladders. The minimum propellant operating temperature for bipropellant

engines is 20 ° F. Because of this less stringent condition, the heating requirements for the

bipropellant system may be lower than those for the monopropellant system. The bipropellant

system has a starting pressure of 350 psia. The blowdown ratio is not as large as the

monopropellant system because the bipropellant engines do not perform well at pressures lower

than approximately 100 psia. The final end of mission pressure for a maximum load case (e.g.,

DRM-3) is 100 psia. This gives a fairly conservative 2.5:1 blowdown ratio. The pressurized

system had a regulator set point of 235 psia. A schematic of the bipropellant system is shown in

Figure 8-2.

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Figure 8-2. Bipropellant Propulsion System Schematic

-23-
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8.3 Dual-Mode Bipropellant System

The dual-mode bipropellant system assumed N2H 4 alone was used for the ACS system and

N204/N2H 4 for the divert engines. Such a system is less complex than a true bipropellant system.

The hydrazine tanks are equipped with bladders, and the ACS engines will be used to settle the

N204 before divert engine ruing. For this reason, no PMDs are required in the N204 tanks which

reduces system weight and cost. As with the monopropellant system, heaters (and their

corresponding power requirement increase) may be required to keep the hydrazine above its

minimum operating temperature. The dual mode system is limited in the blow down range by the

main engine minimum inlet pressure of 125 psia. The tanks are initially loaded to 350 psia,

identical to the aforementioned systems. The regulated pressure set point for this system is

235 psia. A schematic of the dual mode system is shown in Figure 8-3.

r %

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/" _ _ I PRESSURIZATION

C"-IXi t m IXi- Ij

I pvRoVALVE E -1

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I C_CK_AtVES [ _ REGULATORS I

I 10 LBF BIPROPELLANT ENGINES

Figure 8-3. Dual Mode Propulsion System Schematic

-24-
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8.4 LO2/LH 2

A propulsion system employing LO2/I.M 2 as its propellant was considered due to the very

high performance achievable with such a system (Isp=420 see). This performance was based on

the engine developed for the space station program. The engine burns gaseous hydrogen and

oxygen. Propellant storage could be accomplished either with high pressure tankage or cryogenic

storage to reduce the volume and weight requirements. High pressure gas storage was rejected for

the this system due to the excessive volume and weight. As an example, 350 lbs of gaseous

propellant would require almost 1000 lbs of carbon filament tankage for propellant storage.

Cryogenic storage would cut the tank weight in half. However, the use of cryogenics greatly

increases the complexity and cost of the system. Cryogenic propellants also introduce the major

problem of propellant boiloff. To combat this problem, either elaborate insulation is needed or an

additional complex cryo-cooling system is r_uired. These additional complexities are definitely

not desirable. Ground-hold ice formation can cause degradation of the insulation. Boiloff requires

an adequate venting system, including valves that are extremely sensitive to moisture. Materials

and components selected for a propulsion system using cryogenics must be chosen for their

performance at very low temperatures. In addition to the complexity of cryogenics, the extremely

low density of the LH 2 results in very large propellant tanks. For all of these reasons, an LO2/LH 2

system was not considered a serious option for the RRS propulsion system.

8.5 Advanced Technology Propulsion System

For comparison purposes, an N204/N2H 4 bipropellant system was investigated which uses

all advanced technology components. The propellant tanks were assumed to be made of aluminum

liners with carbon fiber overwrap, resulting in a 55% reduction in tank weight. The engines,

valves, filters, and regulators were assumed to be advanced technology designs being developed

for various SDI applications. The advanced technology system uses diaphragm tanks with stainless steel diaphragms for propellant expulsion. This lowers the residuals that would normally be

associated with a bipropellant system. The overall system weight is at least 50% lower than the

conventional propulsion system weights described above.

The dry weights of this system, along with each of the other systems discussed above, are

shown in Table 8-1.

-25-
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9.0 TRADE STUDY RESULTS

The selection of an optimal propulsion system for the RRS began with the elimination of

design choices that would require significant development, have major impacts on the vehicle

design or would require major ground support systems. The next step was to review the dry and

loaded weights for the remaining propulsion systems. The propellant usage for a range of deorbit

and on-orbit maneuvers (assuming a 1.5 g artificial gravity 60 day mission) for each system

defined in Section 8.0 is summarized in Table 9-1. The total mass for a given mission would be

the sum of the dry mass described earlier and the propellant loads shown in this table. These data,

balanced against the performance requirements necessary to perform the RRS mission, were used

to arrive at the selected system.

Table 9-1. Propellant Usage by System for Selected Deorbit Delta-Vs

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R

M

Delta Deorblt

Velocity

(ft/sec)

Monopropellant

Pressure

Regulated Blowdown

(Press) (BD)

328 (100 m/s) 236 241

574 (175 m/s) 340 348

820 (250 m/s) 447 458

1148 (350 m/s) 596 624

Blpropellent

CH3N2H3+N204 N2H4+N204 Advanced GOX/GO 2

Press BD Press BD Press Press

176 178 201 202 196 154

255 258 279 281 270 219

355 346 359 363 347 285

454 467 468 477 452 376

The LO2/LH 2 system and the advanced bipropellant system described in Section 8.0 would

require extensive thruster and tank qualification test programs before implementation. Thus, they

were ruled out on the basis of cost (i.e., expensive new development programs). Table 8-1 shows

that the dry mass difference between the remaining storable propellant systems is less than 35 lbs.

The difference in propellant consumption is directly proportional to the deorbit maneuver velocity

requirement. For the worst case (350 m/sec), the mass difference between the worst and best

system (monoprop blowdown and bipropellant pressure regulated respectively) was 170 lbs. This

difference decreased to 55 lbs with a 100 rn/sec deorbit maneuver.

-27-
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The bipropellant system has space heritage; however, no thruster of less than 2.2 lbf is

currently space qualified. The Astromast spin up requirement that the thrusters deliver less than

1.0 lbf total thrust implies that the maximum spin up engine be 0.5 lbf thrust. This ruled out the

bipropellant system as an option. This effectively left the dual mode and monopropellant systems

as options.

Pressure regulated systems save weight by delivering increased thruster performance over

the course of the mission and lowering the propellant tank volume. Performance is increased by

having the thrusters operate at a constant pressure. This is in lieu of the blowdown operation

where the thrusters operate at high pressures at the beginning of the mission and low pressures at

the mission end. A lower propellant tank mass results from a minimal ullage volume requirement

(typically 5% for a fully loaded system) for the pressurize regulated system (blowdown system

ullage typically 20-35%). This savings is offset, however, by helium tank and regulator masses.

The regulated systems offer larger savings as total propellant load increases. For systems in the

RRS class, a savings of only 15 lbs is achievable. These savings are at the cost of additional

complexity and cost of the regulated system. Thus the minimal weight savings and higher cost and

complexity eliminated regulated systems as an option for the RRS.

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10.0 CONCLUSIONS / RECOMMENDATIONS

A liquid monopropellant blowdown propulsion system was found to be superior in meeting

the requirements of the RRS and is recommended as the baseline system. This system was chosen

over all others considered for the following reasons:

1) The monopropellant system is the simplest of all investigated with the fewest number of

components. The components required for such a system are all fully space qualified

and are in production.

2) The monopropellant system was the most cost effective since it is relatively inexpensive

compared to the other systems. In addition, low unit cost implies that repair and/or

replacement costs will be low as well.

3) No new hardware qualification is required for a monopropellant system. All of the

other systems considered, except for the dual mode, required major component

qualifications.

4) Although the two bipropellant System offered some weight savings capability for

missions requiring large deorbit velocities, better spin balance (equal mass on fore and

aft sections of the deployed RRS) could be obtained with the monopropellant system.

The mass of propellant is consumed before reentry and is not detrimental to the reentry

stability. Thus the heavier propulsion system mass of the monopropellant system was

deemed an advantage for the SAIC RRS concept.

-28-
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6)

Similar to reason number 4, the advantage of lower mass system is only applicable if

the total vehicle design can be reduced to allow a cheaper launch vehicle to be used. At

the time of this trade study, the overall RRS weight budget and launch vehicle selection

was not being driven by the propulsion system selection. Thus, the added cost and

complexity of more advance systems do not warrant application.

The final reason a monopropellant system is recommended for the RRS is that storable

bipropellant main engines and bipropellant ACS engines are not as readily adaptable to

blowdown operation. The engines would run excessively hot at the beginning of the

mission (given a 350 psia inlet pressures) and are more susceptible to poor operation at

the lower inlet pressures as the mission progresses. Poor operation is defined as

chugging or induced off-mixture ratio combustion.

The recommended system design incorporates series redundant single coil valves for both

the attitude control and deorbit engines. This provides for engine-out capability. Dual-coil-dual

seat valves would prevent loss of an engine due to coil failure, but they were not felt to be

necessary considering the short mission duration, increased complexity, weight, and power

consumption of dual-coil valves. The peak power consumption for the propulsion system is

252 W. This equates to all six deorbit engines firing at once, with each engine requiring

1.5 Amps for both valves. Other power consumption items for the propulsion system include:

line, valve, and thruster heaters as well as pressure and temperature sensors. Heaters for the

propulsion tanks are not required based on thermal analysis to date, while valve and thruster

heaters would be used only before f'uings and are included in the power budget. There are two

pressure transducers for the system and sufficient temperature sensors to determine system

performance and control heater functions, if required.

w

-29-