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RRS-027

Reusable Reentry Satellite (RRS),

Summary Report

Thermal Control Trade Study

(NASA-CR-186528) REUSABLE REENTRY

SATELLITE (RRS): THERMAL CONTROL

TRADE STUDY Summary Report

(Science Applications International

Corp.) 44 p

G3/I8

N93-16719

u.c i as April 1990

0277562

Contract HAS9-18202

DRL 02

Prepared for:

National Aeronautics and Space Administration

Lyndon B. Johnson Space Center

Houston, Texas 77058

_ Science 21151 Western Applications Avenue International • Torrance, California Corporation90501 • (213) 781-9022

TOR4.6/23.3
RRS-027

L

Reusable Reentry Satellite (RRS)

Summary Report

Thermal Control Trade Study

April 1990

L

Contract NAS9-18202

DRL 02

Prepared for:

M

National Aeronautics and Space Administration

Lyndon B. Johnson Space Center

Houston, Texas 77058

N

L

_ 21151 ScienceWestern Applications Avenue International • Torrance, California Corporation90501 • (213) 781-9022

TOR4.6/23.3
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FOREWORD

The Reusable Reentry Satellite (RRS) Thermal Control Subsystem (TCS) 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 three times a year for a period of 10 years. The expected

principal use for such a system is research on 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 documents the work performed by Science Applications International

Corporation (SAIC) in the assessment and design of the TCSs necessary to maintain the

temperatures of applicable reusable reentry satellite hardware within prescribed limits during onorbit operations. Specifically, this report describes the efforts to define and verify TCS designs for

maintaining the satellite's Payload Module (PM) environmental control system heat exchanger,

propellant, and water supply within required temperature limits.

The study was performed under the contract technical direction of Mr. Robert Curtis, SAIC

Program Manager. The analyses described herein were performed by Mr. Clark Wallace, SAICTorrance, Thermal Sciences Division. Mr. Michael Richardson, JSC New Initiatives Office,

provided the RRS objectives and policy guidance for the performance of this study under the NAS

9-18202 contract.

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CONTENTS

Section Page

Foreword ........................................................................................... i

Table of Contents .................................................................................. ii

o.o

List of Figures ..................................................................................... m

List of Tables ................................................................................... iv

List of Abbreviations and Acronyms ............................................................ v

Executive Summary ............................................................................... vi

1.0 INTRODUCTION .......................................................................... 1

1.1 Background ........................................................................... 1

1.2 NASA JSC Statement of Work Task Definition ................................. 1

1.3 Scope .................................................................................. 3

2.0 STUDY APPROACH ...................................................................... 3

2.1 Organization .......................................................................... 3

2.2 Document Format .................................................................... 3

3.0 PURPOSE ................................................................................... 4

4.0 GROUNDRULES AND ASSUMPTIONS ............................................. 4

5.0 PAYLOAD MODULE ECLSS HEAT EXCHANGER

THERMAL CONTROL SUBSYSTEM ................................................. 5

5.1 Introduction ........................................................................... 5

5.2 Modeling Tool ........................................................................ 10

5.3 Model .................................................................................. 10

5.5 Assumptions .......................................................................... 13

5.6 Results ................................................................................ 14

5.7 Conclusions .......................................................................... 21

6.0 EXTENDED MODULE PROPELLANT THERMAL RESPONSE ........... . ..... 21

6.1 Introduction ........................................................................... 21

6.2 Modeling Tool ........................................................................ 22

6.3 Model .................................................................................. 22

6.4 Assumptions .......................................................................... 27

6.5 Results ................................................................................ 28

6.6 Conclusions .......................................................................... 29

7.0 PAYLOAD MODULE WATER SUPPLY THERMAL RESPONSE ................ 34

8.0 REFERENCES ............................................................................. 36

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Figure

5-1

5-2

5-3

5-4

5-5

5-6

5-7

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5-9

5-10

5-11

5-12

5-13

6-1

6-2

6-7

6-8

6-9

7-1

LIST OF FIGURES

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DRM-3 Orbital Geometry and RRS Vehicle Orientation Considered in

Forward Module Aft Cover and Aeroshell Surface Heat Balance Calculations... 7

Net Heat Flux on Aft Cover and Aeroshell (180" Conic Station) for

DRM-3, Vernal/Autumnal Equinox, Gravity Gradient Vehicle Orientation ....... 8

Payload Module ECLSS Heat Exchanger TCS Coolant Loop and

Space Radiator Concept. ................................................................ 9

Payload Module ECLSS Heat Exchanger Thermal Control Subsystem ........... 9

Subdivision of Aeroshell Radiator Into 12 Equivalent Panels ...................... 11

Detail of Aeroshell Radiator Panel Model ............................................. 12

ECLSS Heat Exchanger TCS Coolant Loop Fluid Network Model ............... 12

ECLSS Heat Exchanger TCS Pump Characteristic Curve .......................... 14

Orbital Geometry and Vehicle Orientation and Rotation Modes

Considered in Total Absorbed Flux Computations ................................... 15

Total Absorbed Flux (Solar + Albedo + Earth IR) at Radiator Panel 1

Surface During One Orbital Period for DRM-3, Vernal/Autumnal Equinox

and DRM-5, Summer Solstice With Vehicle Normal and Axial Rotation

Rates of 7.0 and 0.14 rpm, Respectively .............................................. 17

Fluid Network Hydraulic Response ................................................... 18

Coolant Thermal Response .............................................................. 18

AerosheU Surface and Coolant Thermal Response, DRM-3 and DRM-5 ......... 20

RRS Extended Module Model .......................................................... 23

DRM-3 Orbital Geometry and RRS Vehicle Orientation Considered in

Extended Module Solar Array Thermal Response Analysis ......................... 24

Extended Module Solar Array Thermal Response for DRM-3 orbit. .............. 25

Extended Module Solar Array Thermal Response for DRM-3 orbit

With and Axial Rotation of 0.14 rpm .................................................. 25

Extended Module Model Electrical Network Analog ................................. 26

Total Absorbed Flux (Solar + Albedo + Earth IR) at Extended Module

Aft Cover Surface During One orbital Period for DRM-3, VemaYAutumnal

Equinox, Normal Rotation and Gravity Gradient Vehicle Orientations ............ 30

Total Absorbed Flux (Solar + Albedo + Earth IR) at Extended Module

Aft Cover Surface During One Orbital Period for DRM-5, Summer

Solstice, Normal Rotation and Gravity Gradient Vehicle Orientations ............. 31

Predicted Temperature Response of Extended Module for DRM-3,

Vernal/Autumnal Equinox, Normal Rotation and Gravity Gradient

Vehicle Orientations ...................................................................... 32

Predicted Temperature Response of Extended Module for DRM-5, Summer

Solstice, Normal Rotation and Gravity Gradient Vehicle Orientations ............. 33

Water Storage Tank Locations Within Forward Module ............................ 34
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Table

4-1

5-1

5-2

LIST OF TABLES

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

ECLSS Heat Exchanger TCS Coolant Loop Fluid Network Components ........

DRM-3, Vernal/Autumnal Equinox and DRM-5, Summer Solstice

Orbit Def'mitions ..........................................................................

PM ECLSS TCS Weight Budget .......................................................

Extended Module Model Electrical Network Analog Nodal Descriptions .........

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16

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E LIST OF ABBREVIATIONS AND ACRONYMS

W

DRM

ESM

HYSTAM

IR

NASA

PM

RRS

SAIC

SRD

TASC

TCS

Design Reference Mission

Silicone Elastomer

Hypersonic-Vehicle Structural, Thermal, and Acoustic Management

Infrared

National Aeronautics and Space Administration

Payload Module

Reusable Reentry Satellite

Science Applications International Corporation

System Requirements Document

Thermal Analyzer for Systems Components

Thermal Control Subsystem

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EXECUTIVE SUMMARY

This report documents the design and assessment work performed in defining the on-orbit

TCS requirements for the RRS. Specifically, it describes the hardware and design measures

necessary for maintaining the PM Environmental Control Life Support System (ECLSS) heat

exchanger, the hydrazine propellant, and PM water supply within their required temperature limits.

A review of the Ames RRS design study and a preliminary evaluation of the SAIC RRS

configuration indicated that a liquid coolant loop in conjunction with a space radiator, was a viable

thermal control concept for the ECLSS heat exchanger. Ethylene glycol was identified as an

appropriate coolant and the vehicle aeroshell was determined to have sufficient radiative capacity

for dissipating the anticipated PM heat load.

A thermal-fluid response assessment of the fluid network and radiator was performed

considering the proposed RRS orbits and vehicle orientations that subjected the aeroshell to the

most thermally stressing environments. The analysis indicated that eight, 0.889 diameter coolant

tubes, in conjunction with a coating of low absorptivity, high emissivity white paint on the

aeroshelrs surface, were adequate space radiator design measures for effectively cooling the

ECLSS heat exchanger. The estimated power requirement for the subsystem is a continuous 7.4

watts and the projected mass is 38.4 lbs (17.4 kg).

A conservative assessment of the on-orbit thermal response of the RRS extended module

indicated that the propellant housed within its interior should remain between 280 and 295 K for all

anticipated RRS missions. Shielding the propellant tanks from direct extraterrestrial exposure and

coating the exterior surfaces of the module's forward and aft covers with a high absorptivity black

paint will maintain the propellant above its minimum allowable of 275 K. The thermal assessment

of the extended module also indicated that an 0.14 rps vehicle axial rotation is necessary to alleviate

excessive cell temperatures experienced by the _uie's exterior solar array. The rotation induces

accelerations below the maximum allowable of 10-5 g.

The PM water supply tanks will experience radiant exchange primarily with the inner

surface of the aeroshell substructure and the external surface of the PM cannister. Since the

temperatures of these surfaces will fall between 281 and 291 K, which is above the 273 K

minimum allowable, it was concluded that no active heating of the water will be necessary.

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1.0 INTRODUCTION

1.1 Background

As currently conceived, the RRS will be designed to provide investigators, in several

biological disciplines, with a relatively inexpensive method to access 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 reflight with new and varied payloads. The

RRS system will be capable of three reflights 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 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 maintained, and economically refurbishable for the next flight. Some systems may

be designed to be replaced rather than refurbished if cost effective and capable of meeting the

specified turnaround time. The minimum time between recovery and reflight 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 f'wst launch in

1993.

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

viable RRS concept that satisfies the requirements. The National Aeronautic and Space

Administration (NASA) has contracted with SAIC to perform a Phase B study to provide the RRS

concept definition. The Thermal Control 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 TCS 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 Para_aph 3.1.2.7 RRS Thermal Control System - Conduct tradeoff studies related

to the choice and characteristics of the thermal control system for the RRS (and associated rodent

module (RM) and other modules).

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(a) Examine thermal control concepts of the RRS vehicle subsystems and the RM and

other modules through appropriate ground support equipment (GSE) during

integration to the RRS vehicle, integration of the RRS system to the launch vehicle,

and pre-launch preparations.

(b) Examine thermal control concepts of the RRS system and the RM and other modules

during the launch ascent trajectory, orbital flight (both drifting at less than 10 -5 g and

spun about the longitudinal axis for artificial gravity), the deorbit exoatmospheric

trajectory, atmospheric flight to landing, the post-landing period prior to the

attachment of GSE, and the post-landing period prior to demating of the experiment

payload.

Specific:

SRD Paragraph 3.3.5 Thermal Control Subsystem

3.3.5.1 RRS Subsystems. The RRS Thermal Control Subsystem shall

maintain the temperature of all RRS subsystem within their required nonoperating

and operating temperature ranges throughout all mission phases.

3.3.5.2 PM. The RRS Thermal Control Subsystem shall provide the

appropriate thermal control at the RRS/PM thermal interfaces to satisfy the PM

thermal requirements (TBD) for all phases of a mission.

3.3.5.3 Pre-Launch Phase. The RRS will be maintained at the desired

temperature range by use of GSE connected to the internal RRS cooling system.

The GSE shall not impede late access to the PM.

3.3.5.4 Orbit Phase. The RRS cooling system shall be designed to remove

the maximum payload heat rate, TBD, for orbit conditions from sun-synchronous

orbit to low-inclination orbits, plus a reasonable margin for growth.

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3.3.5.5 Recovery Phase. The RRS Thermal Control Subsystem shall be

designed to minimize the reentry heat soak into the internal cavity of the vehicle and

to minimize the increased RRS PM temperature. The design shall allow thermal

control via GSE to be applied to the PM within TBD minutes of ground

touchdown.

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 TCS 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.

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, tradeoff studies and system concept definitions.

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

accommodate a compendium of analyses that are published in several separate documents.

Because of the synergistic nature of one tradeoff study across the entire RRS system design, it is

suggested that the reader review all summary reports to get a complete picture of the SAIC RRS

design.

2.2 Document Format

Individual analyses and studies are not necessarily amenable to documentation in exactly

the same topical arrangement; 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

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• Analysis Results

• Conclusions

• Recommendations

3.0 PURPOSE

The purpose of the RRS Thermal Control Trade Study was to assess the design of the

TCSs necessary to maintain the temperatures of applicable RRS hardware within prescribed limits

during on-orbit operations. Specifically, this report describes the efforts to define and verify TCS

designs for maintaining the sateUite's PM environmental control system heat exchanger, propellant,

and water supply within required temperature limits.

4.0 GROUNDRULES AND ASSUMPTIONS

The overall requirements of the RRS TCS were discussed previously in Section 1.2 as

stated in the RRS SRD. These requirements and the design reference missions (DRMs) defined

early in the study, and shown in Table 4-1, were the only top-level groundrules and assumptions

used in the thermal trade system studies.

Table 4-1. RRS Design Reference Missions

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Definition DRM-I

Parameter

Character

Inclination

Orbit Type Circular

Orbit Altitude 350 km

(189 nm)

Launch Site

Recovery Site

land Recovery

33.83*

Eastern Test

Ranse fETR)

White Sands

Missile Range

0VSMR)

Design Reference Mission Set

DRM-2 DRM-4 DRM-5

High Altitude Water Recovery

33.83 °

Circular

Integer Orbits

35.65 °

DRM-3

High Inclination

98_

Circular,

Near-Integ_

897 km

(4,84 nm)

WTR

WSMR

Circular,

Integer

479 km

(259 nm)

28.5°

Circular

900 km 350 km

(486 nm) (189 nm)

ETR ETR ETR

WSMR WSMR Water (ETR,

Gulf of

Mexico, WTR)

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5.0 PAYLOAD MODULE ECLSS HEAT EXCHANGER THERMAL CONTROL

SUBSYSTEM

5.1 introduction

The RRS PM contains a complete ECI.,SS capable of providing the required atmosphere for

sustaining 18 rodents and a suite of related electronics and hardware for up to 60 days. Within the

PM, the air supply's humidity, chemical content, and temperature are monitored and adjusted

continually. The humidity and content are controlled by subsystems completely internal to the PM.

The temperature is maintained by regulating the flow of air through the ECLSS heat exchanger.

During satellite launch, ascent, reentry, and recovery periods, the heat exchanger functions

autonomously to cool the PM airflow through the incorporation of a fusible wax within its

construction. The wax undergoes a solid/liquid phase transition at 278.8 K and the quantity

contained within the heat exchanger has been sized to provide 1055 kilojoules of thermal storage.

This translates into approximately 3 hours of independent cooling capability. A complete

description of the ECLSS heat exchanger design is provided in the SAIC RRS Payload Module

Trade Study. While the satellite is in orbit, however, heat removal and consequent dissipation

must be provided by a TCS coupled to the heat exchanger, but external to the PM.

The heat generated within the PM is attributable to the rodents and electronics/hardware

which produce, on average, 45 and 75 watts, respectively. The external TCS must, therefore, be

capable of transporting and dissipating approximately 120 W of heat. In addition, the minimum

anticipated PM air temperature is 291 K which requires the TCS to maintain the PM ECLSS heat

exchanger at or below this temperature.

The Ames Reusable Reentry Satellite Design Study (Ref. 1), confronted with a similar set

of capability requirements, went a long way in identifying the character of a potential TCS. Two

preliminary observations were instrumental in shaping the basic concept and defining their final

TCS design. First, the RRS concept places maximum emphasis on the reusability of onboard

subsystems and, therefore, an expendable TCS was not permissible. Secondly, the satellite's

geometry, in conjunction with the transient nature of the PM heat loads, did not bode well for

passive TCSs such as heat pipes. Therefore, it was concluded that an integrated, active TCS

consisting of a fluid network coupled with a space radiator was the most efficient alternative. The

SAIC RRS configuration is no exception to the above considerations and therefore this study

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addressedthe

concept.

design and correspondingassessmentof a fluid network/spaceradiator TCS

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To ascertain the validity of this concept in relation to the SAIC RRS vehicle design,

potential locations for the required space radiator were examined. The aft cover of the forward

module was considered fast. This location provides approximately 1.5 square meters of relatively

contiguous radiator area. However, when the vehicle is orientated to induce a microgravity

environment (gravity gradient orientation - no vehicle rotation in the orbital plane - nose facing

outward) the aft cover is Earth-facing over the entire orbit. Such an orientation will severely

degrade this location's dissipative efficiency.

To assess the actual impact this orientation has on the aft cover's efficiency as a radiator, a

heat balance calculation was performed at its surface for the DRM-3 orbit at Vernal/Autumnal

Equinox. Figure 5-1 depicts the DRM-3 orbit and corresponding vehicle orientation used in the

analysis. The calculation accounted for direct solar insolation, Earth-reflected solar insolation, and

Earth infrared (IR) radiation as a function of the cover's time variant orbital position and

orientation. Surface reradiation to the Earth (Te = 289 K) and deep space were also included. The

aft cover's surface was assumed to have a solar absorptance of 0.2 and a hemispherical emissivity

of 0.9. The Earth albedo was set at 0.35 and the cover's surface temperature was held constant at

270 K.

The results from this heat balance calculation are presented in Figure 5-2 where the net heat

flux on the aft cover's surface is plotted as a function of time. The net flux is always positive

(surface is absorbing energy) and the average over the entire orbit is 19 W/m 2. Since the surface

must experience an average negative net heat flux of at least -80 W/m 2 to effectively dissipate the

PM heat load, the analysis indicates that the aft cover is not an acceptable location for the radiator.

Having eliminated the aft cover as a candidate location for the space radiator, the remaining

alternative was to use the vehicle's aeroshell. The aeroshell provides approximately 4.4 square

meters of unobstructed radiator area, excluding the aft skirt and nose of the vehicle. To verify the

dissipative potential of this location, a heat balance calculation, as described above, was performed

at the aeroshelrs surface for the 180 ° circumferential station (see Figure 5-1).

Before the results from the above calculation are discussed, it is important to note that an

assessment of the extended module's on orbit thermal response (Section 6.0) indicated that a

vehicle axial rotation rate of 0.14 rpm was necessary (see Figure 5-1). This requirement was

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imposedto alleviatethetheexcessivetemperaturesexperiencedby the sun-facingportion of the

module's solar array that resulted in a significant degradation in cell power production efficiency.

Polar orbits, such as DRM-3, in conjunction with the microgravity vehicle orientation (no rotation

in the orbital plane) resulted in the most severe temperature excursions. An axial rotation rate of

0.14 rpm was found to effectively smooth the highly nonuniform heating about the module's

circumference, consequently lowering the bulk temperature of the power producing portion of the

solar array.

5eL_,n: Vemal/Autumn_d Equ m,ox

Orbital InclinJtion: 98.0"

(relative to ecliptic

OrbitJ| Period: 6176.0 sec

Occultauon P_od: 0.0 sec

Altitude: 89'7.0

Axial Rotation: O. 14 RPM

Aft Cover

--O°station

of Eclipitic

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Figure 5-1. DRM-3 Orbital Geometry and RRS Vehicle Orientation Considered in

Forward Module Aft Cover and Aeroshell Surface Heat Balance Calculations.

This axial rotation rate was found to induce an acceleration of 0.8x10 5 g at a radius of 35.6

cm, which is the maximum radius of the PM rodent cages. This acceleration falls below the

maximum allowable level of 10 -5 g set forth in the RRS System Design Study Statement of Work,

Section 4.2.5.2 (Ref. 2) and is therefore acceptable. The impact of this rotation on vehicle stability

is described in the SAIC RRS Guidance, Navigation, and Control Trade Study.

Figure 5-2 shows results from the heat balance calculation at the 180 ° circumferential

station. The periodic nature of the net heat flux is due to the vehicle's 0.14 rpm axial rotation rate.

The average net heat flux over one orbit is -158 W/m 2. This translates into almost 700 W of

dissipation power for the entire aeroshell for this orbit if one assumes a construction from a thin

wall, high conductivity material at a constant temperature of 270 K. Since the radiator will, in fact,

consist of almost 3 centimeters of low conductivity foamed silicone elastomer (ESM), the

anticipated efficiency will be much lower. Nevertheless, the above result indicates that the

aeroshell is a viable location for the TCS space radiator.

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400

DRM-3, Vernal/Autumnal Equinox, Gravity Gradient

%= 0.2,fir= 0.9, albedo = 0.35, Te = 289 K, Tr_f= 270 K

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z

300

2O0

100

o

-100

-200

-300

Conic, Axial = 0.141 RPH

Aft Cover

_ _ L_

-400 .... I , - J _ , , L = • ' .... =

o 1ooo zooo 3ooo 4000 5000 6000

TIME (sec)

Figure 5-2. Net Heat Flux on Aft Cover and Aeroshell (180 ° Conic Station) for DRM-3,

Vernal/Autumnal Equinox, Gravity Gradient Vehicle Orientation.

Having identified a location for a space radiator on the SAIC RRS, a TCS concept for the

ECLSS heat exchanger was developed. The resulting subsystem is presented in Figure 5-3.

Coolant is pumped through the ECLSS heat exchanger and then distributed into one of two sets of

cooling tubes, each spanning 180 ° sections of opposing aeroshell circumference. The cooling

tubes are integrated directly into aeroshelrs aluminum substructure. During transit, heat is

transferred from the coolant to the aeroshell backface and then conducted to the aeroshell surface

where it is radiated to the environment. The cooled fluid streams are then merged and pass through

a reservoir before returning to the heat exchanger.

A schematic of the complete ECLSS heat exchanger TCS is presented in Figure 5-4.

Valves are incorporated to permit the network to interface with a ground support cooling system

during the RRS prelaunch period. Additional valves are included to permit the coolant loop to

bypass the ECLSS heat exchanger and/or space radiator and thereby provide some degree of

coolant temperature control. An auxiliary pump serves as a backup in the event of primary pump

dysfunction or failure. A thermal control unit interfaces with the various valves providing fluid

network flow control as dictated by the environmental control unit processor.

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ECLSS HEAT

EXCHANGER PUMP RESERVIOR

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Figure 5-3. Payload Module ECLSS Heat Exchanger TCS Coolant Loop and Space

Radiator Concept.

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RADIATOR I

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BYPASS (_ )®

f

._ % / PAYLOAD. _ MODULE

HEAT EXCHANGER I AIR SUPPLY

qin

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THERMAL

CONTROL

UNIT

ENVIRONMENTAL

CONTROL

UNIT

Figure 5-4. Payload Module ECLSS Heat Exchanger Thermal Control Subsystem.

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Having arrived at a generalTCSconcept,thedesignissueof identifying the appropriate

coolantwasaddressed.For this concept,it wasdecidedthat a low systemoperatingpressureof

somewherebetween10to 20 psia would provide a good balanceof safety,cost, complexity,

thermalefficiency, powerrequirements,andweight. At thesepressuresandanticipatedECLSS

heatexchangertemperaturesrangingbetween270to 300K, a single-phasesystemusinga liquid

coolantsuchasethyleneglycol appearedappropriate.Oneinherentdisadvantage,however,with

usingthis coolantis thatits viscosity displaysa strongdependenceontemperature.As aresult,

theviscositybecomesrestrictivelyhighasthetemperatureapproachesits freezingpoint of 258K.

Therefore,in additionto amaximumallowablecoolanttemperatureof 291K imposedby thePM

payload,aminimumallowablecoolanttemperatureof 273K was artificially imposed to prevent the

viscosity of the glycol from becoming unacceptably large.

5.2 Modeling Tool

W

±

z 2

= :

= =

Ill

i

To assess the performance of the proposed TCS concept, a modeling tool capable of

handling in depth thermal response of multi-material structures, coupled with the thermal-fluid

response of a coolant network, was required. Such a capability is achieved in the SAIC-developed

Hypersonic-Vehicle Structural, Thermal, and Acoustic Management (HYSTAM) computer code

(Ref. 3). HYSTAM was originally developed to assess the performance of candidate cooling

concepts for hypersonic vehicles. The code consists of three coupled modules that predict

aeroheating and acoustics loads, structure thermal response, and cooling network thermal-fluid

response. For the application at hand, the aeroheating and acoustics module was replaced by a

module that predicted the solar insolation, Earth-reflected solar insolation, and Earth IR flux on an

arbitrarily oriented surface in Earth orbit. Included within the routine were algorithms that tracked

the orientation of each radiator panel as a function of vehicle rotations and orientation relative to the

orbital plane.

5.3 Model

The philosophy adopted in this assessment was to model as thoroughly as possible the

thermal-fluid response of the coolant, the transport of heat through the aeroshell thickness, and the

consequent radiant heat balance of the aeroshell surface with the extraterrestrial environment. The

conceptual construction began with the subdivision of the aeroshell into 12 equivalent rectangular

panels, approximately 80 by 46 cm on a side as shown in Figure 5-5. Each panel represented 30 °

of vehicle circumference and the combined areas matched that of the actual conic's midsection

minus the aft skirt and nosecap. Each panel consisted of a 2.286 cm lay-up of ESM (the aeroshell

-10-
r_

iE_r

J

w

thermal protection material) over a 0.145 cm substrate of aluminum into which coolant passages

were integrated as shown in Figure 5-6. This was an approximation of the actual substrate that

will be aluminum honeycomb with aluminum coolant tubes bonded to the inner surface of the

outboard facing sheet. However, the mass of the solid aluminum plate modeled replicated the

anticipated mass of the actual substrate. Although the number and size of the coolant tubes was

arrived at iteratively as part of the assessment process, it is noted at this time that the TCS design

requires eight coolant tubes, 0.899 cm in diameter, spaced 5 cm apart.

The aeroshell ESM surface was coated with a low absorptivity white paint to enhance the

dissipative efficiency of the system. A solar absorptance of 0.2 and a hemispherical emissivity of

0.9 were used in the assessment.

iliil

E

0 o 30 ° 60 ° 90 ° 120 ° 150 ° 180"

I I I I I I I

0 ° 330 ° 300 ° 270 ° 240 °

I I I I I

21 O* 180 °

I I

Figure 5-5. Subdivision of Aeroshell Radiator Into 12 Equivalent Panels.

z=

Iml

7

u

Following the HYSTAM modeling approach, a fluid network model of the cooling loop

and its various components was constructed. A schematic of the network and a list of the

component descriptions are presented in Figure 5-7 and Table 5-1, respectively. The network

model consisted of a 1 liter reservoir, a pump, the ECLSS heat exchanger, an inlet manifold which

branched the flow, two sets of six panels representing 180 ° of vehicle circumference, a collection

manifold which merged the flow, and 3 meters of miscellaneous tubing. Loss coefficients were

applied to the heat exchanger and both manifolds. The pump characteristic curve used in the

analysis, which was postulated, is presented in Figure 5-8.

-11-
I

0

0

' "/"_ 0.x_$ c_

_Channel

"" _ _" "" • Aluminum

Figure 5-6. Detail of Aeroshell Radiator Panel Model.

L

M

J

m

-±

I

w

u

J

M

M

RESERVIOR

qin

PUMP EXCHANGER

1 2

q out q =0

I II Ik____ll I

3 4 5 6

RADIATOR PANELS

7 8 9 10 11 12

4

PLUMBING

Figure 5-7. ECLSS Heat Exchanger TCS Coolant Loop Fluid Network Model.

-12-
5.5 Assumptions

In developing the TCS concept model, several assumptions and simplifications were

necessary to adhere to the multi-one-dimensional modeling limitation in HYSTAM. The surface

area of each aeroshell radiator panel was modeled exactly; however, the additional surface area

provided by the aft skirt and nose sections was neglected. The spherical geometry of the nosecap

and the complex radiant heating and exchange experienced by the skirt's backface present

situations HYSTAM could not model. Intuitively, one may surmise that the additional area will

enhance the overall efficiency of the radiator.

Table 5-1. ECLSS Heat Exchanger TCS Coolant Loop Fluid Network Components.

w

7

u

m

=

W

w

= =

M

D

Reservoir

Pump

Heat Exchanger

Radiator Panel

(see Figure 5-6)

Coolant Tube

(see Figure 5-6)

Plumbing

(assorted tubing)

Coolant

Volume

Pressure

Material

See Figure 5-8

120 W continuous

Surface:

Backface:

Number

Diameter

Material

Total Length

Diameter

Material

Ethylene Glycol

= 0.001 m3

= 14.7 psia

= Aluminum

eh,= 0.9

as= 0.2

Adiabatic

= 8 / panel

= 0.889 cm

= Aluminum

= 304.8 cm

= 1.27 cm

= Aluminum

In the model, the backface of the aeroshell substructure was treated as a perfectly adiabatic

boundary. A low density insulation will cover the substructure surface to minimize heat transfer

with vehicle's internals, however, some radiant exchange will still take place. As noted above,

HYSTAM is not configured to model the complex radiant exchange between the insulation and the

satellite's internals. However, the temperature differences between the various internal surfaces

and the substructure insulation will be small and, therefore, the consequent radiant transfer of heat

should be negligible.

-13-
With regard to the ECLSS heat exchanger, no attempt was made to model the time variant

nature of the heat load or the transport of heat from the incoming PM air to the coolant. For this

analysis, the heat exchanger control volume served as a continuous 120 W source term in the

fluid's thermal energy balance.

Lastly, since the HYSTAM approach is multi-one-dimensional, no structural thermal

communication was permitted between panels. However, the heating loads about the aeroshelrs

circumference will be essentially uniform because of the vehicle's constant axial rotation. This will

result in a fairly uniform temperature distribution, consequently negating circumferential

conduction.

5O

4O

3O

t_

<1 20

10

N

o.ooo 0.025 0.050 0.0?5 o.ioo o.tz_ o.tso o.t75 0.200

MASS FLOW RATE (kg/sec)

Figure 5-8. ECL_S Heat Exchanger TCS Pump Characteristic Curve.

H

1

M

1

5.6 Results

Prior to exercising HYSTAM, a preliminary set of computations was conducted to ascertain

which of the 30 possible DRM/vehicle orientation/season combinations (five DRMs, three seasons,

two vehicle orientations) subjected the aeroshell's surface to the most stressing thermal

environments. The goal was to economize the required number of HYSTAM runs by

concentrating on the scenarios that produced the highest and lowest aeroshell heating rates and,

-14-
E_

s.1

W

m

i

r_

m

_ H

N

w

m

w

therefore, the hottest and coldest aeroshell surface temperatures. The orbital geometry and vehicle

orientation and rotation modes considered in the analysis are portrayed in Figure 5-9. The same

orbital heat loads model used for the aft cover and aeroshell heat balance computations described

previously was used. For this investigation, however, surface reradiation was not computed.

For the analysis, total absorbed flux computations were performed for DRMs 1 through 5,

considering both normal rotation (7 rpm, see Figure 5-9) and gravity gradient vehicle orientations

with a constant axial rotation rate of 0.14 rpm. Orbit inclination angle relative to the ecliptic was

adjusted to account for Vernal/Autumnal Equinox, Summer Solstice, and Winter Solstice seasonal

variations. The analysis location corresponded to the aeroshelrs 15 ° circumferential station which

translated to the center of panel 1 in the coolant loop fluid network model.

Based on the results of this preliminary investigation for the 30 possible combinations, the

average absorbed flux at the aeroshell's surface during one orbital period ranged between 76

and 103 W/m 2. These extremes corresponded to the DRM-5, Summer Solstice and the DRM-3,

Vernal/Autumnal Equinox orbits with the vehicle in a 7 rpm normal rotation mode. Table 5-2

depicts the parameters that define these two orbits. The total absorbed flux witnessed by aeroshell

panel 1, as a function of time for these two orbits, is presented in Figure 5-10.

Vehiclelo_tionand

oflentatlon at analysis

Time = 0 seconds

Axial Romti_

Plm_¢ol F:cllpiZtc

\

Figure 5-9. Orbital Geometry and Vehicle Orientation and Rotation Modes Considered in

Total Absorbed Flux Computations.

-15-
m

Having defined a TCS network model and identified the most stressing orbits, the

thermal/hydraulic response of the fluid network and space radiator were assessed using HYSTAM.

Figure 5-11 shows the predicted hydraulic response of the subsystem at steady-state conditions for

both orbits. The computed pump power, assuming an efficiency of 65%, is a continuous 7.4 W.

The pump produces approximately 5 psia of head, a coolant mass flow rate of 0.153 kg/sec, and a

coolant velocity of 0.137 m/sec within the radiator coolant tubes.

m

z

w

m

Table 5-2. DRM-3, Vernal/Autumnal Equinoxand DRM-5, Summer Solstice Orbit Definitions.

Designation

Orbit Inclination

(relative to ecliptic)

Orbital Period

Occultation Period

Altitude

Normal Rotation

Axial Rotation

DRM-3

Vernal/Autumnal

Equinox

98°

6176 see

0sec

897 km

7 rpm

0.14 rpm

DRM-5

Summer

Solstice

o

5492 sec

2177 sec

350 km

7 rpm

O.14 rpm

The predicted thermal response of the coolant for both orbits is presented in Figure 5-12.

The temperatures correspond to that of the fluid as it enters the ECLSS heat exchanger. The initial

system temperature for the DRM-3 orbit was set at 300 K and was designated the hot orbit/hot start

condition. The initial system temperature for the DRM-5 orbit was set at 280 K and was

designated the cold orbit/cold start condition. The intent was simply to derive some insight into

how initial temperature influenced the time required for the subsystem to reach a steady-state

operating condition.

The hot orbit/hot start condition requires the coolant loop to bypass the heat exchanger for

approximately 3600 seconds. The bypass permits the residual heat within the subsystem to

dissipate allowing the temperature of the coolant to drop below the 291 K maximum allowable for

ECLSS heat exchanger. For the analysis, the bypass was simulated by switching the 120 W

source term in the ECLSS heat exchanger control volume on at 3600 seconds into the calculation.

This is evidenced in Figure 5-12 by the burp in the coolant temperature response curve. For the

DRM-3 orbit, the coolant temperature stabilizes at 284 K.

m

_=

u

-16-
!

= =

= =

w

DRM-3, NR = ZO RPM, Axial = 0.14 RPM

_= 0.2, eu= 0.9, albedo = 0.35, Te = 289 K, Tret-.-=270 K

400 _ , _ _ _

350

300

_ 250 !

200

t50

 ,oo

5(1

0 '

o t ooo zooo 3ooo 4000 5000 8000

TIME (sec)

--4

1

m

m

W

S •

1

[mmml

L

4O0

350

3OO

250

_-_ 200

150

0

03

.<

100

5O

DRM-5, NR = 7.0 RPM, Axial = 0.14 RPM

as= O.2,elr= 0.9, albedo = 0.35, Te = 289 K, Tref = 270 K

I I , T I I I

0

o mooo 2000 3000 4000 5000

TIME (sec)

N 3

Kind

s

I

Figure 5-10. Total Absorbed Flux (Solar + Albedo + Earth IR) at Radiator Panel 1 Surface During

One Orbital Period for DRM-3, Vernal/Autumnal Equinox and DRM-5, Summer Solstice With

Vehicle Normal and Axial Rotation Rates of 7.0 and 0.14 rpm, Respectively.

-17-
i

[]

m

m

m

m

[]

n

i

E

w

=

84

22

2O

_// 16

14

12

lO

0

3tO

305

300

205

200

r._

z85

280

275

270

I ' I I . , , I I I I

Pump Power: 7.4 W (continuous)

Total Ap : 5.1 psla

rh : 0.153 kg/sec

Vradiator : O.[37 m/sec

tube

• ""'-..•..%.

........................... ..........................

, DP,M - 3

i

1 2 3 4 5 6

DISTANCE (m)

Figure 5-11. Fluid Network Hydraulic Response

Fluid Temperature at Heat Exchanger Inlet

I i

i i , - I _ • 1

_oooo zoooo _oooo 4oooo _oooo soooo

TIME (sec)

Figure 5-12. Coolant Thermal Response

-18-
= ,

Considering the cold orbit/cold start condition, the prediction indicates that the coolant

temperature quickly equilibrates to a pseudo-steady-state condition between 282 and 284 K. The

periodic oscillation in the coolant temperature response is the result of the 2177 second occultation

period that occurs during the DRM-5 orbit. Note that the coolant temperatures achieved for both

orbits fall between the maximum and minimum allowable coolant temperatures of 291 and 273 K

identified previously.

The required number and diameter of coolant tubes in the space radiator were arrived at

iteratively by performing successive sets of HYSTAM computations. The above coolant

temperatures were achieved using eight, 0.889 cm diameter tubes per panel spaced 5 cm apart.

Figure 5-13 permits a comparison of the thermal response of the coolant relative to that of

the aeroshell surface. The temperature response at the surface of radiator panels 1 and 4 (15 ° and

105 °) are shown in relation to the coolant temperature as a function of time for both DRM-3 and

DRM-5. The aeroshell surface temperature ranges between 260 to 280 K. In addition, the results

indicate that the maximum temperature difference between adjacent panels does not exceed 3 K.

Therefore, circumferential conduction about the aeroshell is negligible as assumed originally.

Having arrived at a viable concept, a weight estimate of the proposed TCS was made

assuming a construction entirely of aluminum. The complete breakdown for the subsystem is

given in Table 5-3. The total weight calculated, excluding the ECLSS heat exchanger and the

aeroshell and its substructure, is 38.4 lbs (17.4 kg).

m

u

w

u

L_

w

m

Table 5-3. PM ECLSS TCS Weight Budget.

Mass

Component Ibs (kg)

Coolant

Tubing

Pumps

Reservoir

Valves and fittings

Structure

TOTAL

10.8 (4.9)

15.6 (7.1)

5.5 (2.5)

3.1 (1.4)

2.2 (1.0)

1.1 (0.5)

38.4 (17.4)

-19-
i

310

300

_" 39O

[.,.3

0::;

380

Or:;

[.a.]

0..,

[.a..]

[._, 270

260

250

0

DRM-3, Vernal/Autumr_al Equinox (Ti = 300 K)

Normal Rotation = 7.0 RPM, Axial Rota{ion 0.14 RPM

i i _ I i

10000 ::'0000 30000 40000 50000 60000

TIME (see)

i

==

w

310

300

290

t_

280

0..,

270

DRM-5, Summer Solstice (_i = 280 K)

Normal Rotation = 7.0 RPM, Axial Rotation = 0.14 RPM

i i i I i

260

250

0 10000 20000 30000 40000 50000 60000

TIME (see)

Figure 5-13. Aeroshell Surface and Coolant Thermal Response, DRM-3 and DRM-5.

w

u

-20-
_i

m

L

w

w

w

m

m

r

m

m

L_

5.7 Conclusions

Based on results reported in the Ames Reusable Reentry Satellite Design Study and the

current trade study, it was concluded that an active TCS, consisting of a fluid network coupled

with a space radiator, was a viable thermal control concept for the SAIC RRS Payload Module's

ECLSS heat exchanger. A preliminary thermal response calculation indicated that the RRS

aeroshell would serve as an acceptable location for the TCS space radiator. Based on the

subsystem's projected operating pressure and temperature, the liquid coolant ethylene glycol was

chosen as the TCS working fluid.

A thermal-fluid response assessment of the proposed TCS was performed using the

HYSTAM computer code. The analysis considered the two RRS orbit/vehicle orientation/seasonal

combinations that subjected the aeroshell to the most stressing thermal environments. The aeroshell

surface was coated with a low absorptivity white paint to enhance its dissipative efficiency. The

coating was assumed to have a solar absorptance of 0.2 and a hemispherical emissivity of 0.9.

The results indicated that eight, 0.889 cm diameter coolant tubes integrated into the aeroshelrs

substructure effectively dissipated the heat exchanger's 120 W heat load while maintaining the

coolant within the mandated temperature range of 273 to 291 K.

The pump power requirement for the TCS was estimated to be a continuous 7.4 W. The

projected weight of the complete subsystem is 38.4 lbs (17.4 kg).

6.0 EXTENDED MODULE PROPELLANT THERMAL RESPONSE

6.1 Introduction

The RRS extended module contains six spherical tanks fabricated from titanium in which

hydrazine propellant is stored. Current system requirements dictate that the temperature of the

propellant must be maintained between 275 and 330 K to ensure its safe containment. The tanks

are almost completely enclosed by surrounding surfaces and thereby shielded from direct

extraterrestrial exposure, however, because of the wide array of orbit types and vehicle orientations

the RRS might encounter, thermal exchange with these externally exposed structures will still take

place. Therefore, it is probable that under certain conditions the temperature of the tanks and their

accompanying propellant will violate the above requirement if corrective measures, such as

insulations, spectrally selective coatings, or heaters, are not imposed. This section describes the

-21-
w

li-a

m

i

w

thermal assessment of the extended module that was performed to ascertain which corrective

measures, if any, were necessary.

6.2 Modeling Tool

The extended module presents a relatively complex three dimension geometry that

experiences radiant exchange with its extraterrestrial environment as well as radiant and conductive

exchange between its component parts. The SAIC-developed Thermal Analyzer for Systems

Components (TASC) computer code is well suited for this application (Ref. 4). TASC is a

lumped-capacitance, electrical resistor-capacitor network analog-based code for analyzing timedependent thermal response of arbitrarily configured, multi-material, one, two, and three

dimensional systems. A routine for computing direct solar insolation, Earth-reflected solar

insolation, and Earth IR radiation on an arbitrarily oriented surface was integrated into the code

which computed the magnitudes of these radiant components exactly as a function the extended

module's time variant orbital position and orientation. This routine was an adaptation of the model

used in the ECLSS TCS heat balance and HYSTAM computations described in Section 5.0.

6.3 Model

A pictorial representation of the extended module model developed for this assessment is

presented in Figure 6-1. The model consisted of a truncated cone composed of a solar array over a

substrate of aluminum honeycomb, two circular aluminum honeycomb disks serving as forward

and aft covers, and six tanks, half full of hydrazine (nominal condition), suspended within the

resulting enclosure. The entire model experienced rotations as prescribed by the particular orbital

mission under consideration. The solar array and the external surfaces of the covers were exposed

to direct solar insolation, Earth-reflected solar insolation, and Earth IR radiation and experienced

radiant exchange with the Earth and deep space.

Before developing a TASC electrical network analog for the extended module model, a set

of preliminary calculations was undertaken to assess the thermal response of the module's solar

array. The motivation was to obtain an estimate of temperature gradients about the module's

circumference and hence guide the model nodalization process.

As described above, solar panels cover the entire exterior surface of the module's truncated

cone perimeter. A simple TASC model of the array was developed by dividing the truncated cone

into six equivalent circumferential sections. Each section consisted of a lay-up of quartz over a

-22-

W
zm

u

X...A

r .

Solar Array / ._ '_

Subs_ucuJrc

• Solar • Earth Albedo Normal Rotation = 7.0 RPM

• Earth IR Gravity Gradient = 0.0 RPM

'_ _ _ Axial Rotation = 0.5 RPM

Forward Cover

_ "" Propell_t Tank

/

/

/

AftCover

m

v

m

i

i

Figure 6-1. RRS Extended Module Model.

substrate of aluminum honeycomb. Each panel was permitted to communicate with its neighbors

via conduction, however, no radiant exchange across the cone's interior was considered. The

thermal mass and response of the remaining extended module's structure was ignored.

The TASC code was exercised for the DRM-3 orbit at Vernal/Autumnal Equinox with the

vehicle in the gravity gradient orientation (no vehicle rotation, vehicle's axis is perpendicular to the

Earth's surface, and vehicle nose facing outward). From analyses discussed in Section 5.0, this

scenario was determined to impose the highest heat loads on vehicle conic surfaces. Figure 6-2

presents a pictorial representation of the orbital geometry and vehicle orientation considered.

m

Lw

= ,

m

b

The results are presented in Figure 6-3 which plots the temperature response at the centroid

of each panel versus time. The temperatures experienced by the sun facing panels (-120 ° - 180 ° -

-240 °) cause a substantial degradation in the array's overall efficiency and a consequent loss in

power production that is unacceptable.

-23-
-2

m

U

N

w

w

m

m

E

Sease_: Vcrrml/Aulumnal rc.qu_x

(re_v¢ m _lipti¢

OrbJ_tlPeriod: 6176.0 sec

Oc_[mt/onP_iod: 0.0 _._

Al_mae_ 19"/.0 k_

Plane at Ecl/pi_

Vehicle location a.nd

orientation it inalys_

Exle_led Module

tS

Figure 6-2. DRM-3 Orbital Geometry and RRS Vehicle Orientation Considered in

Extended Module Solar Array Thermal Response Analysis.

An obvious solution for eliminating the large nonuniformity in temperature about the

module's circumference is to impose an axial rotation on the entire vehicle. However, the loads

induced by such a rotation must not produce an acceleration greater than 10 -5 g as set forth in the

RRS System Design Study Statement of Work, Section 4.2.5.2 (Ref. 2). A simple calculation sets

the maximum allowable axial rotation rate at 0.14 rpm. This rotation produces an acceleration of

10 -5 g at the PMs 44.5 cm outer radius. This results in an acceptable 0.8x10 -5 g at the maximum

rodent cage radius of 35.6 cm.

i

Repeating the previously described TASC analysis with a vehicle axial rotation rate of 0.14

rpm results in the circumferential temperature distribution presented in Figure 6-4. The rotation is

sufficient to effectively smooth the nonuniformity and results in an array temperature of

300K+/- 10 K.

m

m

=

The above results were used to develop the electrical network analog for the extended

module model shown in Figure 6-5. Because of geometric and thermal symmetry, the solar array,

array substructure, and forward and aft covers were modeled as single nodes. In addition, the six

propellant tanks and their resident propellant were also treated as single nodes. The surface areas,

masses, and conductive contact areas between each of these structures, however, were all

computed based on the extended module's actual geometry. The radiation view factors between

the array substructure, forward cover, aft cover, and propellant tanks were calculated assuming no

intervening obstuctions.

-24-
z :

w

m

M

m

=

m

D

m

V

450

4OO

35O

300

250

200

150

DRM-3, Vernal/Autumnal Equinox, Gravity Gradient

No Vehicle Axial Rotation

I l _ I I

180 °

/

240 °

................ 0oo

0 °

100 = i , I , , , I , , , l I l , , ,

0 2000 4000 6000 8000 10000 12000

TIME (see)

Figure 6-3. Extended Module Solar Array Thermal Response for DRM-3 Orbit.

DRM-3, Vernal/Autumnal Equinox, Gravity Gradient

Vehicle Axial Rotation = 0.14 RPM

450 , ,

400

350

300

250

200

150

i00 , , , I = ,= , I , , I I I

0 2000 4000 6000 BOO0 10000 12000

TIME (sec)

Figure 6-4. Extended Module Solar Array Thermal Response for DRM-3 Orbit With and Axial

Rotation of 0.14 rpm

-25-
m

Solar Array

/I

_(t)

C :

R :

_'k_Forward Cover

C/R

C R C

Substructure pellant Tank

e3_._ t C°ver

_l_t) _t)

Radiation to environment (Deep Space, Earth)

Radiation from environment (Solar, Earth Albedo, Earth IR)

Conduction

Radiation

Propellant

L

U

L

w

E

N

m

Figure 6-5. Extended Module Model Electrical Network Analog

Table 6-1 describes the material, mass, and radiant properties for each of the six network

nodes. The external surfaces of the forward and aft covers were coated with a high absorptivity

black paint. The coating was assumed to have a solar absorptance and hemispherical emissivity of

0.9. For the DRM-3 polar orbit, the mass of the node representing the propellant was set to zero.

Since the heating environments induced by this orbit are relatively uniform, the temperatures of the

various surfaces which constitute the module's exterior will not fluctuate broadly over each orbit.

Since the temperatures of these surfaces in turn drive the thermal response of the propellant tanks,

the tanks themselves will experience even smaller fluctuations and as a consequence equilibrate to a

steady-state condition. As a result, the additional mass provided by the propellant only serves to

lengthen the system's equilibrium time constant. By eliminating the mass of the propellant when

considering the DRM-3 orbit, the system reaches its equilibrium condition more rapidly thereby

reducing the required computation time interval.

=

m

-26-
Table 6-1. Extended Module Model Electrical Network Analog Nodal Descriptions

Node Material Modeled Nodal Mass Radiant

Surface Properties *

Solar Array Quartz 3.15 as = 0.86 e. = 0.80

Substructure Aluminum Honeycomb 20.10 a = 0.80 _ = 0.80

**Forward Cover Aluminum Honeycomb 8.23 cq = 0.90 _ = 0.90

a =0.80c =0.80

"*Aft Cover Aluminum Honeycomb 15.41 cq = 0.90 r.n = 0.90

a =0.80¢ --0.80

Propellant Tank tTitanium 44.91 a = 0.80 _ ffi0.75

Propellant Hydra_a¢ DRM-3 0.0

DRM-5 154.20

• a s mad_. correspond to vehicle external sm'faces, a and r. correspond to vehicle internal surfaces

• * External surfaces coated with black paint

t External surfaces coated, but uninsulated

6.4 Assumptions

As is evident from the inspection of Figures 6-1 and 6-5, the model representing the

extended module resulted from several simplifying assumptions that assisted in economizing the

analysis without sacrificing the integrity of the predictions. Several of these assumptions have

been discussed previously. The remainder are addressed here.

Since struts and/or flanges with relatively small cross-sectional areas will be used to secure

the tanks within the module, it was assumed that radiative rather than conductive exchange would

be the dominant mechanism of heat transfer between the module's shell and the tank surfaces.

Also, it was assumed that the tanks are completely shielded from direct exposure to the external

environment. This, in fact, is not the case since the forward cover is penetrated by the portion of

the PM which protrudes from the forward module when the vehicle is in the retracted position.

This requires an 88.9 cm diameter opening in the cover that exposes a portion of the tankage

surface to the extraterrestrial environment. However, the placement of a cylindrical radiation shield

of a diameter equivalent to that of the opening and extending into the module (to the aft cover if

necessary) would suffice in completely isolating the tanks.

L

-27-
! !

I

u

Additional assumptions were adopted in order to impart some degree of conservatism into

the overall analysis. The tanks were assumed to be coated with a surface preparation, but were left

uninsulated. The analysis neglected the 67 W of heat rejected from the various electronics,

batteries, and hardware located within the module (a conservative estimate when considering

minimum predicted temperatures). Also, as noted previously, the view factors between the tanks

and the inner surfaces of the solar array and covers were computed assuming that there were no

intervening obstructions.

6.5 Results

u

As was the case in the ECLSS heat exchanger TCS thermal response assessment described

in Section 5.0, prior to executing the TASC code, a set of preliminary calculations was performed

to determine which DRM/vehicle orientation/season combinations subjected the module's surfaces

to the most stressing thermal environments. As before, the goal was minimize the number of code

computations necessary to adequately assess the module's thermal performance. Figure 5-9 shows

the orbital geometry and RRS vehicle orientation and rotation modes considered in the analysis.

The orbital heat loads routine described in Section 5.2 was used to perform the computations in

conjunction with the surface radiative properties given in Table 6-1.

w

i

_w

The results from the above investigation indicated that the DRM-3, Vernal/Autumnal

Equinox and DRM-5, Summer Solstice orbits in combination with a vehicle normal rotation rate of

7 rpm produced the lowest and highest heat loads on the module. The forward cover was the

exception. This surface received the lowest and highest absorbed fluxes for the above orbits, but

in combination with the gravity gradient vehicle orientation. Examples of the computed total

absorbed fluxes at the surface of the extended module aft cover are presented in Figures 6-6 and

6-7 for the above scenarios.

Having identified the most thermally stressing orbits, the TASC code was exercised on the

previously described extended module model. Figure 6-8 shows the predicted temperature

responses of the solar array substructure, the forward and aft covers, and the propellant tank for

DRM-3. Since the solar array was treated as a single node and the thermal heat balance was

performed at only a single circumferential station on the conic, the 0.14 axial rotation produces an

oscillatory temperature response for this surface. For both the normal rotation and gravity gradient

vehicle orientations, the propellant tank equilibrates to approximately 280 K.

-28-
z_:

ii

Figure 6-8 shows the predicted temperature responses of the solar array substructure,

forward and aft covers, the propellant tank, and propellant for the DRM-5 orbit. For this orbit, the

nodal temperature fluctuations are more pronounced because of the 2177 second occultation

period. Although the temperature variations of the module's external surfaces fall well below the

minimum propellant allowable of 275 K, the radiant coupling between these surfaces and the tank

is sufficiently damped to prevent the tank temperature from dropping below 280 K. The bulk

temperature of the propellant remains relatively steady, equilibrating to about 295 K for both the

normal rotation and gravity gradient vehicle orientations.

;___

r

6.6 Conclusions

An assessment of the on-orbit thermal response of the SAIC RRS extended module has

shown that the propellant housed within its interior should remain between 280 to 300 K for all

anticipated RRS missions. The analysis indicates that shielding the tanks from exposure to the

external environment and coating the exterior surfaces of the forward and aft covers with a high

absorptivity black paint are adequate measures for maintaining the propellant above its minimum

allowable of 275 K. At present, it appears that no active thermal control measures such as heaters

will be necessary.

w

L ,

=

w

w

m

-29-
_J

= =

= =

w

X

tl)

t_

O

5"}

m

1400

1200

1ooo

800

600

400

2o0

DRM-3 , NR = 7.0 RPM,

as= 0.9#,= 0.9, albedo = 0.35, T_ = 289 K, T_ = 270 K

t I I I f

tooo eooo aooo 40o0 sooo eooo

TIME (sec)

w

E

E

k_

I

w

0

<

1400

1200

tO00

800

6O0

400

200

DRM-3 , Gravity Gradient,

%= 0.9,ell= 0.9, albedo = 0.35, T_= 289 K, T_f = 270 K

0 . . I .... l I .... l . . • • I

0 2000 3000 4000 5000 6000

TIME (see)

• I , .

1000

Figure 6-6. Total Absorbed Flux (Solar + Albedo + Earth IR) at Extended Module Aft

Cover Surface During One Orbital Period for DRM-3, Vernal/Autumnal Equinox, Normal

Rotation and Gravity Gradient Vehicle Orientations.

-30-
1400

i200

DRM-5 , NR = 7.0 RPM,

as= 0.9,ell= 0.9, albedo = 0.35, 're= 289 K, T_= 270 K

¢'4

0

0

09

lO00

800

600

400

2O0

=

=

= :

Ill,,=.l

Iiil

w

tooo 2000 3ooo 4000

TIME (sec)

5000

1400

1200

"" 1000

M 0oo

E:I 600

r,=.l

0

03

/_ 400

2OO

DRM-5 , Gravity Gradient,

as= 0.9,ell= 0.9, albedo =0.35, Te = 289 K, Tre r= 270 K

I ! I I

• • • I | i i • • I , . l • ] i i , , I . ,

tO00 2000 3000 4000 5000

TIME (sec)

Figure 6-7. Total Absorbed Flux (Solar + Albedo + Earth IR) at Extended Module Aft

Cover Surface During One Orbital Period for DRM-5, Summer Solstice, Normal Rotation

and Gravity Gradient Vehicle Orientations.

-31-
z

360

340

320

_" 300

280

t_

F-,-I

C_

z_o

[..o

E--,

240

220

200

0

DRM-3, Vernal/Autumnal Equinox

Normal Rotation = 7.0 RPM, Axial Rotation = 0.14 RPM

I I I ' • t

Array Backface

Propellant Tank

Forward Cover

Aft Cover

. , , , I , , , , 1

10000 20000

I ..... I I

30000 40000 50000

TIME (see)

60000

= =

w

h-

=

i

i

360

340

32O

"-" 300

fa_

280

_6o

E240

220

200

0

DRM-3, Vernal/Autumnal Equinox

Gravity Gradient, Axial Rotation = 0.14 RPM

I I I I I

i'i!!ii_i i .._!_i!_. :!_!. . :!i, ;:_h , i_: ,_:: .i _: .: ii .,:_

::::: ":t':":::>'.:::_:" "',.:_::,::::: :_ ' ':,'.:_::_z_:;_'"_ _::::_.:::::::',_ :','::_z _::t " ._::_:::'::':'_ ,_.:_: _:," "',','::.z:.::_::.'. "':::.z:_

: ": :: :: :::::_:_:: ::" :::::::::::::::::::::::: :::_ _: _:: _= : :::: ":_ ::::::: :::_:::::_::: ¢_iV :_;;_:_. :: : _:_:::::: :::: ::

Propellant Tank

Array Back face

Forward Cover

Aft Cover

10000 20000 30000 40000 50000 60000

TIME (see)

i

Figure 6-8. Predicted Temperature Response of Extended Module for DRM-3, Vernal/Autumnal

Equinox, Normal Rotation and Gravity Gradient Vehicle Orientations.

-32-
DRM-5, Summer Solstice

Normal Rotation = 7.0 RPM, Axial Rotation = 0.14 RPM

b

L

360

340

3zo A

I

/J

240

22O

200

0

v 300

280

z6o

r._

[.-,

I I I I I

r_ ^

I

\

_ _A_ _

i._--I:-.

1

I

\

Array Backface

ProPe]!a.nt Tank...

Propellant

Forward Cover

Aft Cover

i i 1 i [ i i i i I

tO000 20000

i I .... •

30000 40000

TIME (see)

50000 60000

w

==

i

_

=

E :

360

340

320

_-_ 300

280

zeo

F.,

240

220

200

k

/\,

t_

II

DRM-5, Summer Solstice

Gravity Gradient, Axial Rotation = 0.14 RPM

': I.

._iL ...... :..__.

/I

_\tl

Array Backface

..P_ggeHant- .Tank....

__Ptoe_j!=__m_.......

Forward Cover

\

\

Aft Cover

A

\

!I "

I\

' I

i__ _-I!._I

IV .,

\

If\ I

It \1

\.

. ._____ ,.:___

60000

.... I .... I .... L .... t .... 1 , ,

_oooo zoooo 30000 _oooo 50000

TIME (see)

Figure 6-9. Predicted Temperature Response of Extended Module for DRM-5, Summer Solstice,

Normal Rotation and Gravity Gradient Vehicle Orientations.

-33-
w

it

w

W

i:7 i

_:= !

7.0 PAYLOAD MODULE WATER SUPPLY THERMAL RESPONSE

Three cylindrical tanks reside within the RRS forward module, each containing

approximately 20 kilograms of water. The tanks are situated concentrically about the PM as shown

in Figure 7-1. The total quantity of water stored is sufficient to sustain a population of 18 rodents

for 60 days. Since the water is stored in liquid form, it must remain above 273 K throughout all

anticipated RRS missions. The following addresses the design issue of whether or not active

heaters are required to maintain the water above this temperature.

m_

• i

b

= -

h_

Aft Skirt Aeroshell

\ /

j.S '_ ,,,.

• / " Water Tank

Aft Cover

Figure 7-1. Water Storage Tank Locations Within Forward Module.

While in orbit, the tanks experience both radiative and conductive exchange with their

surroundings. However, conductive heat transfer will be minimized because of the relatively small

cross-sectional areas of the brackets used to secure the tanks to the aeroshell substructure.

Therefore, it may be assumed that radiative, rather than conductive exchange, will be the dominant

mode of heat transfer for the water tanks.

An examination of Figure 7-1 shows that the primary radiating surfaces, which constitute

the forward module's internal cavity and experience some form of thermal forcing function,

-34-
z

whether it be internal (rodents and electronics) or external (solar insolation), are the exterior of the

payload module and the insulated backface of the aeroshell substructure. The oxygen and air tanks

that bound each water tank can be ignored since they experience the same thermal environment as

the water tanks and will, therefore, equilibrate to the same temperature, neutralizing radiant

exchange.

Mission requirements dictate that the temperature of the PM interior range between 299 and

291 K. The analysis described in Section 5.0 indicates that the temperature of the aeroshell

substructure will fall somewhere between 284 and 281 K. Since radiative exchange is dominant

and the PM and aeroshell substructure surfaces essentially encompass the water tanks, the

temperature of the tanks will equilibrate between these two limits. Since the lower limit, the

substructure temperature, is well above 273 K, it is concluded that no active heating of the water

tanks win be necessary.

E_

--,,-

w

F

F_

L

-35-
7 : W

m

°

*

o

*

References

"A Conceptual Design Study of the Reusable Reentry Satellite," NASA Technical

Memorandum TM 101043, NASA Ames Research Center, Moffet Field, California, October

1988.

"Statement of Work for a Phase B Study of a Reusable Reentry Satellite", NASA Request for

Proposal 9-BE2-23-8-48P, NASA Johnson Space Center, Houston, Texas, January 10, 1989.

Wassel, A. T., D. C. Bugby, A. L. Laganelli, and C. E. Wallace, "Hypersonic-Vehicle

Structural, Thermal and Acoustic Management (HYSTAM) Program," SAIC Report No.

SAIC-90/1208, SAIC, Torrance, California, 1989.

Duncan, T. C., J. L. Farr, F. T. Makowski, and A. T. Wassel, "Satellite Laser Vulnerability

Model: Thermal Analysis Module," SAIC Report No. SAI-060-80R-073-LA, SAIC,

Torrance, California, 1980.

=

L_

E

u

m

i--

w

-36-