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List of Tables........................................ 4


University of Virginia

USRA/IIASA Advanced Design Program Summer Conference June 1988

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Report 2of 3

A Proposal For a manned Orbital Transfer Vehicle For the 21st Century

Submitted by WWSR Incorporated

TABLE OF CONTENTS

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Lkt of Figures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2 List of Tables . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4 Important Abbreviations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5 Foreword . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6 Project Orion Team Members . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7 Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8 Chapter 1 . Design of the OTV . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12 Chapter 2 .The Aerobrake . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 21 Chapter 3 . Engine Specifications . . . . . . . . . . . . . . . . . . . . . . . . . . . 46 51 Chapter 4 . Fuel System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Chapter 5 . Materials and Structures . . . . . . . . . . . . . . . . . . . . . . . . . 54 65 Chapter 6 . Ambient Heat Transfer . . . . . . . . . . . . . . . . . . . . . . . . . . Chapter 7 . Electrical Power System . . . . . . . . . . . . . . . . . . . . . . . . . 67 Chapter 8 . Environmental Control and Life Support . . . . . . . . . . . . . . . 70 Chapter 9 . Guidance Navigation and Control . . . . . . . . . . . . . . . . . . . . 82 Chapter 10 . Data Management System . . . . . . . . . . . . . . . . . . . . . . . . 84 Chapter 11 . Communication System . . . . . . . . . . . . . . . . . . . . . . . . . . 86 Chapter 12 . Satellite Repair and Recovery System . . . . . . . . . . . . . . . . . 88 95 Chapter 13 . Cost Analysis . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Chapter 14 . Managing Project Orion . . . . . . . . . . . . . . . . . . . . . . . . . 98 Chapter 15 .Mission Planning . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 107 Conclusion . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 108 Appendix 1 . System and Subsystem Weight and Power Estimates . . . . . . . .115 Appendix 2 . Orbital Mechanics . . . . . . . . . . . . . . . . . . . . . . . . . . . . .119 Appendix 3 . OTV Servicing Aboard the Space Station . . . . . . . . . . . . . . .124 Appendix 4 . Section Authors . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .127

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LIST OF FIGURES
Figure 1-1: 1-1: 1-2: 1-3:
1-4:

1-5: 1-6: 1-7: 1-8: 1-9:
2-1:

2-2: 2-3: 2-4: 2-5: 2-6: 2-7: 2-8: 2-9: 2-10: 2-11: 2-12: 2-13:
2-14:

2-15: 2-16: 2-17:
2-18:

3-1:

3-2: 3-3:
4-1: 4-2:

5-1: 5-2: 5-3:
5-4:

5-5:
5-6: 5-7:

7-1: 8-1:

10 An Artist’s Rendition of the OTV . . . . . . . . . . . . . . . . . . . . . . . . Detailed Drawing of WWSR’s OTV . . . . . . . . . . . . . . . . . . . . . . . 14 General Configuration of WWSR’s OTV (Side View) . . . . . . . . . . . . . 15 General Configuration of WWSR’s OTV (Front View) . . . . . . . . . . . . 16 General Configuration of WWSR’s OTV (Top View) . . . . . . . . . . . . . 17 Detailed Drawing of Interior Layout . . . . . . . . . . . . . . . . . . . . . . . 18 View of Cockpit as Seen From Interior . . . . . . . . . . . . . . . . . . . . . 19 View of Interior as Seen From Cockpit . . . . . . . . . . . . . . . . . . . . . 19 20 Diagram of CCM & EVAM (Side View) . . . . . . . . . . . . . . . . . . . . Diagram of CCM & EVAM (Top View) . . . . . . . . . . . . . . . . . . . . 20 Schematic of Aerobrake Maneuver . . . . . . . . . . . . . . . . . . . . . . . . 22 23 Altitude History . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Aerobraking Velocity Decrements . . . . . . . . . . . . . . . . . . . . . . . . . 24 Aerobrake Geometry . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 25 Geometric Construction of Raked-Cone . . . . . . . . . . . . . . . . . . . . . 27 Flight Path Angle . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 28 NASTRAX Model . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 29 Construction of Heat Shield . . . . . . . . . . . . . . . . . . . . . . . . . . . . 31 Aerobraking Heating Rates . . . . . . . . . . . . . . . . . . . . . . . . . . . . 32 33 Minimization of Heating Rates . . . . . . . . . . . . . . . . . . . . . . . . . . . Heating Rate History . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 34 Ballistic Coefficient Effects . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 35 Reduction of Heating Rate . . . . . . . . . . . . . . . . . . . . . . . . . . . . 37 Heat Flux and Pressure Distribution . . . . . . . . . . . . . . . . . . . . . . . 38 Schematic of Thermal Protection System . . . . . . . . . . . . . . . . . . . . 39 40 Thermal Protection System on Aerobrake . . . . . . . . . . . . . . . . . . . . 41 Gap Filler Configuration . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . After-body Impingement . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 42 RLlO Derivative Engine . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 48 RLlOO Engine Flow Schematic . . . . . . . . . . . . . . . . . . . . . . . . . . 49 Aerobrake Engine Configuration . . . . . . . . . . . . . . . . . . . . . . . . . 50 Fuel System Schematic . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 51 Pressurization System of Tank Pair . . . . . . . . . . . . . . . . . . . . . . . 52 Tank Pressure and Weight . . . . . . . . . . . . . . . . . . . . . . . . . . . . 56 Schematic of MLI Thermal Protection System . . . . . . . . . . . . . . . . . 58 OTV and Modular Tanks . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 59 Tank Support Structure and Connectors . . . . . . . . . . . . . . . . . . . . 60 Longitudinal Tensile Strength of Graphite/Epoxy . . . . . . . . . . . . . . . 61 Tank Support Structure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 62 Semi-monoque Structure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 64 The Electrical Power System . . . . . . . . . . . . . . . . . . . . . . . . . . . 69 OTV ECLSS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 71

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OTV Atmospheric Revitalization System . . . . . . . . . . . . . . . . . . . . 72 75 OTV Water Loop . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . OTV Freon Loop . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 75 Incidence of Radiation Sickness . . . . . . . . . . . . . . . . . . . . . . . . . . 80 Manned Maneuvering Unit . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 91 Extravehicular Mobility Unit . . . . . . . . . . . . . . . . . . . . . . . . . . . 92 Trunion Pin Attachment Device . . . . . . . . . . . . . . . . . . . . . . . . . 92 Serviceable Satellite Configuration . . . . . . . . . . . . . . . . . . . . . . . . 93 94 "Stinger" Device in Use . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 94 Satellite Grasping Arm . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Management and Manufacturing Structure . . . . . . . . . . . . . . . . . . .100 A2-1: Trajectory Schematic . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 120 A2-2: Location of OTV's Delta V's . . . . . . . . . . . . . . . . . . . . . . . . . . . 123 8-2: 8-3: 8-4: 8-5: 12-1: 12-2: 12-3: 12-4: 12-5: 12-6: 14-1:

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LIST OF TABLES

Table 2-1:

..

Characteristics of the Aerobrake . . . . . . . . . . . . . . . . . . . . . . . . . 26 2-2: Thickness and Density of TPS . . . . . . . . . . . . . . . . . . . . . . . . . . 39 3-1: 1987 OTV Engine Goals and the RLlOO Engine . . . . . . . . . . . . . . . . 47 5-1: Properties of 2219 and 2090 Aluminum Alloys . . . . . . . . . . . . . . . . . 57 5-2: Fracture/Tensile Properties of 2090-T8E41 Aluminum Alloy . . . . . . . . . 57 8-1: Space Radiation Dose Rates (rad/day) . . . . . . . . . . . . . . . . . . . . . . 79 79 8-2: Space Radiation Dose Rates (rem/day) . . . . . . . . . . . . . . . . . . . . . 8-3: Duration and Apogee Altitude . . . . . . . . . . . . . . . . . . . . . . . . . . 80 8-4: Effects of Acute Radiation Doses . . . . . . . . . . . . . . . . . . . . . . . . . 81 8-5: Radiation Thresholds for Certain Materials . . . . . . . . . . . . . . . . . . . 81 12-1: Satellite Repair Missions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 88 13-1: Project Orion Costs . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 97 102 1 5 1 : Mission A Profile . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 15-2: Mission B Profile . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 104 15-3: Mission C Profile . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 106 118 Al-1: Total System Weights (Worst Case) . . . . . . . . . . . . . . . . . . . . . . . Al-2: M s and Power Analysis of OTV ECLSS . . . . . . . . . . . . . . . . . . . as 116 A1-3: Electronic and Power Systems Analysis . . . . . . . . . . . . . . . . . . . . . 117 Al-4: Structural Component Weight Estimates . . . . . . . . . . . . . . . . . . . . 118 122 A2-1: Summary of Delta V’s . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

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Important Abbrevat ions
AOTV CCM c.g. ECLSS
1 . .

EMU
EPS EVA EVAM GEO GNC LEO LH2 or LH, LO2 or LO, MLI

MMC'
OMV OTV PIB RCS TPS

Aerobraked (Aeroassisted) OTV Crew Command Module center of gravity Environmental Contol & Life Support Systems Extra-vehicular Mobility Unit Electrical Power Systems . Extra-vehicular Activity EVA Module Geosynchronous Orbit Guidance, Navigation and Control Low Earth Orbit Liquid Hydrogen Liquid Oxygen Multi-layered Insulation Manned Maneuvering Unit Orbital Maneuvering Vehicle Orbital Transfer Vehicle Phased Injection Burn Reaction Control System Thermal Protection System

5

Foreword

The following paper is the final report in fulfillment of the requirements for the undergraduate design sequence in aerospace engineering, AE 441-442. manned orbital transfer vehicle. Even though Group W did not work on a project for the AIAA design competition, design. we did attempt to present a proposal that would meet the requirements of the competition if a request for proposal had been made for our This meant that we needed not only to design an OTV but to address such concerns as costs, manufacturing, and management. For this reason, the paper is written to be a proposal from an aerospace corporation that is to be presented to NASA. fictious corporation. It is the culmination of nine months of work completed by Group W on its design for a

WWSR Inc. was created to be this
Any

WWSR is a composite of many aerospace corporations.

similiarities to an actual corporation is purely coincidential.

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PROJECT ORION TEAM MEMBERS

Gregory Weigand
Project Orion Group Leader, Management, and Mission Planning

Michael Doheny
Rocket Engines and Heat Transfer

Richard Franck
Materials, Structures, and Aerobraking

Steven Hollo

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Avionics, Control, Power Systems, Orbital Mechanics, Mission Planning, and Aerobraking

I

Kenneth Ibarra Life Support and Orbital Mechanics

William Nosal
Economics, Aerobraking, and EVA Activities

Thomas Redd
Design and Space Station Intergration

7

Introduction

Since the. late 1970s to early 1980s, there has been considerable research into the deployment of an American space station. The proposed Space Station will

It will also permit numerous spaced-based missions that may not have been practical in the past. One of these
allow for a permament manned settlement in space. missions is the deployment of an orbital transfer vehicle (OTV). an OTV is to make excursions from one orbit to another. also a major difference between an OTV and its The purpose of There is orbital More specifically, it is counterpart the

to be capable of going into higher Earth orbits than the Space Shuttle.

maneuvering vehicle (OMV) in that an OMV is only designed for orbit changes of a few hundred miles while the OTV is designed for orbit changes of thousands of miles. For the most part, current OTVs have been designed to be able to go, at the very least, from Low Earth Orbit (LEO) to a geostationary orbit (GEO).

NASA has been investigating several proposals from other areospace firms for
OTVs.

A few proposed OTVs have been ground-based, but most have been

designed to be permamently based at the Space Station. In Pioneering the Space Frontier:

The Report of the National Commission

on Space, the Commission states that:
A high priority exists for this vehicle [an OTV], which will greatly lower the cost of access to geostationary orbit and to the Moon for crews and The transfer vehicle will be payloads ranging from 10 to 20 tons. modular, single-stage, fueled by liquid oxygen and liquid hydrogen, and outfitted with an aerobrake to conserve fuel by allowing the vehicle to slow down through the drag of Earth’s atmosphere ... With appropriate modification the transfer vehicle could be used as a lunar lander [l, p. 1221.

In response to the need for an OTV expressed in the report, WWSR has created a proposal for a manned OTV that meets the criteria selected by the Commission. criteria: The design that WWSR is proposing will also meet the following

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1. Be based at the Space Station.

2. Have the capability of supporting 3 people for a mission lasting no longer than 14 days.
3. Be a b l e ' t o perform multiple missions between LEO and GEO with a minimum amount of servicing.
4. Carry a maximum payload of 24,000 pounds between LEO and GEO. 5 . Support EVA.

The primary mission of the OTV is to support manned excursions to GEO to service a satellite in orbit without needing to return it to the Space Station or to Earth. WWSR realizes, however, that it may not be possible due to some unique For this reason, the OTV has been It is This eliminates failure of a satellite to repair it at GEO.

designed to be capable of bringing the satellite back to the Space Station. also capable of returning the same (or another) satellite to GEO. Centaur). WWSR has based its design on a "worst case" scenario. mission that consists of the following:
1. Leaving the Space Station, going to GEO, and returning.
2. Carrying a 24,000 pound payload to GEO.

unnecessary missions to GEO by other payload delivery systems (such as the PAM

This scenario is a

3. Carrying a full crew of 3.
4. Lasting for 14 days.

This worst case scenario may never be realized within in the first few years of deployment. One reason is that current satellites rarely weigh over 10,000 lbm.

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Another reason is that if the mission is simply to repair a satellite, it is highly unlikely that a crew of three will be required or that they will need 14 days to complete the mission. However, the Project Orion team has designed its OTV in NASA is quite intent on creating other platforms Our OTV will be used to realize anticipation of future missions.

in addition to the Space Station based at LEO.

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this goal. It satellite repair.

WGINAL PAGE 1s OF POOR QUALlN
has been designed to be capable of transporting the heayy components of a platform without being unsuitable for its primary mission of It has been designed to support a three person crew for a duration Other missions
of time that will allow them to work on assembling the platform.

that may be possible because of the constraints of our worst case scenario will be manned missions to the Moon. longer duration missions (with lighter payload requirements). higher orbit missons. or missions with more personnel (this would be accomplished by adding an additional crew module).

Figure 1-1: .in arrist’s rendition of an O T I - similiar to \Y\I-SR * c proposed design. .source: Pion e c riri g thc S p a c e Frontir r

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With the aforementioned criteria and worst case scenario, the Project Orion team embarked on its design process. The goal of the team is to create an OTV Project It is the general that will maximize performance by using the most up to date technologies. Orion will not- use systems that have not been fully proven. than risk the vehicle or the crew in order to cut costs.

philosophy of the team that it is better to stick with “tried and true” methods We also feel that using One major Early in our proven state-of-the-art systems will actually cut costs in the long run. abberation of this philosophy may be the use of an aerobrake. on its return to LEO. highly feasible.

decision process, we selected the aerobrake as our choice for slowing down the OTV It is not a totally proven system, but it has been substantially investigated by WWSR and other companies and has shown to be Even so, our choice for an aerobrake is similar to the method used successfully for the Apollo missions. The following chapters of this report consist of Project Orion’s design for the OTV and its subsystems. investigation. may be more efficient. This design has been chosen after eight months of Other designs for OTVs that use electrical, solar, or nuclear power We feel, however, that our design is the most optimal

possible to meet the National Commission on Space’s demand for a chemicallypowered, aerobraked, manned OTV as well as the design scenario selected by WWSR and MOVERS.

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Chapter 1 Design of the OTV
The final design of the OTV involved the integration of many different systems. The main design parameter was the aerobrake, after that the propellant tanks, and then the cargo, and crew module areas. The aerobrake is the most important part of the design. The the type of shield that was selected is called a raked sphere cone lifting brake. trip to the Space Station. This shape was chosen so that maximum propellant savings would be obtained during the return The brake will be shipped up tho the Space Station via The the Space Shuttle. It will assembled and attached to the OTV at the station. in the Space Shuttle cargo bay. Chapter 2 of this report. The design of the propellant tanks was chosen with ease of construction and delivery in mind. The propellant tanks are modular and consist of two tanks (LH2 The tank modules will be The tanks will then be The The OTV has been The tanks and L 0 2 ) , the required support .structure, and piping.

brake will be made of numerous sections each of which will be small enough to fit The aerobrake is covered in more detail in

built on Earth and flown up empty on the Space Shuttle.

moved to the OTV area of the Space Station for integration to the OTV. tank modules are designed to be identical and interchangable.

designed to carry anywhere from two to six sets of tanks depending on the mission.

These tanks are attached radially around the central command module. are put into place by cranes in the OTV servicing area. diagonal supports are connected by attending astronauts.

The fuel lines and The modular design

shortens the time needed for servicing the OTV, thus reducing costs. Since the OTV has to travel in space as well as through a portion of the atmosphere. the placement and design of the crew command module (CCM), EV-4 module (EVAM), and cargo area are very important. central axis. The semi-spherical design of the aerobrake made it necessary to put the manned portions of the craft along the The interior components needed to positioned as symmetrically as

12

possible to ensure the center of gravity was near the central axis. from the atmospheric heating during the aerobraking maneuver.

The central The area

location of the manned portions also means that this area will be better protected protected by the aerobrake will form a cone above the brake.

In order to keep the

components of the OTV as well as its payload within this cone of protection meant that the central structure must be narrow but not excessively tall. The CCM and EVAM are designed to be transported in the shuttle cargo bay. The CCM is 22 ft in length and 12 ft in diameter. The CCM contains all
of the supplies, perishables, computers, controls, and facilities needed for a 14 day

mission. Interior components of the CCM are broken down into hexagonal sections that fit within the circular cross section of the main pressure walls. shower, and head are in the extreme rear of the CCM. area. The galley, The computers and

avionics are placed in front of these sections so that they are closest to the cockpit The life support, electrical power, and air revitalization systems are located Unlike the rest of the CCM, the The two The controls are placed in a Below the cockpit is the This in modules place in the "floor" and "ceiling."

cockpit area maximizes space by returning the circular cross section. pilot's seats are located side by side facing forward. hatch to the EVAM. manner similiar to that of the Space Shuttle's cockpit.

The third crew member will have a seat underneath and

behind the cockpit such that he would be facing the hatch to the EVAM.
entering the EVAM.

seat will fold up when not in use. The area then can be used to prepare for

The EVAM is where the MMU, equipment, and tools for satellite repair will be stored. The EVAM contains an airlock that will be used to transfer between The rest of EVAM will be normally left evacuated. space. EVAM, and CCM.

Outside the EVAM is the robot arm that will be used to grapple satellites and

.MMUs.

The main EVA hatch will also double as the hard docking hatch when The EVAM can be detached from the CCM. The EVAM

the OTV is at the Space Station.

This allows the versatility of adding any sort of module such as another crew module or space laboratory that might be needed for a given mission. is is 8 ft in length and 10.5 ft in diameter.

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ORIGINAL PAGE I S OF POOR QUALITY

The engines are placed centrally for several reasons. the most important. of which is stability. The engines' center of thrust will be in line with the center of The central placement will also reduce the number of
Two

gravity of the whole OTI'.

lines needed from the propellant tanks and simplify servicing the OTI'.
for a safer and more reliable system.

engines acting redundantly were chosen over one main engine since this provided

.-. .
'.

\

\

1-1: Detailed Drawing of I171YSR's O T I -

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ORIGINAL PAGE E S OF POOR QUALITY

0

I TATELLITE SUPPORT I I

.-

-

EVA MODULE

CREW COMMAND

1LIQUID H,J I i I

ySUm

b2-

T.

1-2: General Configuration of WWSR’s OTV (Side View)

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1-3: General Configuration of WWSR’s OTV (Front View)

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1-4: General Configuration of WWSR’s OTV (Top View)

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1-5: Detailed Drawing of Interior Layout

ORIGINAL PAGE G S OF POOR QUALITY
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11AIRLOCK HATCH -

1-6:View of Cockpit as Seen from Interior

ACCESSWAY

1-7: View of Interior from Cockpit

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1-8: Diagram of CCM & EVAM (Side View)

loft

T
.-’

1
t-

Nl
’ t V A 9,UIT

8 ft

+-

- --- -_I-.__..- __ 22 f t

-

--_.---

1-9: Diagram of CCM & EVAM (Top View)

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Chapter 2 The Aerobrake
...

The Aerobraking Maneuver The WWSR OTV was designed to carry large payloads to geosynchronous orbits. In order to maximize the weight of this payload
F

hile keeping the amount

of fuel needed low, the OTV will use the drag produced by passing through the Earth’s atmosphere to dissipate its excess velocity as it returns from GEO. Aerobraking,
as this

process is called, results

in a

arge fuel savings over

propulsively slowing the craft using retrorockets. configured all-propulsive craft [ 121.

In fact, it has been shown that

an aerobraking OTV can carry twice the roundtrip payload to GEO as a similiarly Aerobraking is, however, a very complex maneuver, creating many important vehicle design considerations. the added complexity of a thermal protection system. criteria. The aerobraking maneuver is initiated at GEO where the OTV’s engines are fired to produce the necessary plane change and inject the vehicle into a transfer trajectory that will take it into Earth’s atmosphere. through the atmosphere. For most WWSR OTV missions (no returning payload), aerobraking will be performed in two passes

As the vehicle

passes through the atmosphere it experiences severe aerodynamic heating, requiring Additionally, since the craft is essentially flying, aerodynamic configuration and control become prime design

A schematic diagram of a two pass maneuver, as

compared to a one pass, is shown in Figure 2-1 [lo]. The first pass will last only
5 minutes and will take the OTV to within 85 kilometers of the Earth’s surface.

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The deceleration of the vehicle due to the drag on the aerobrake will place it in a intermediate orbit with an apogee midway between LEO and GEO. Slight corrections in this orbit will take the OTV through the atmosphere for a second time, at approximately the same altitude but for 11 minutes (due to the already reduced velocity of the OTV). delta-v).
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This pass will place the craft in an orbit that can

be circularized at LEO with a relatively small propulsive burn (less than 200 m/s

GEOSYNCHRONOUS TWO P A S S

AEROBRAKING

V O U J AEROBRAKING ORBIT Cowparison of AOTV trajectories.

RENDEZ

Figure 2-1: Schematic drawing of a one and two pass aerobraking maneuver.
The altitude and velocity of an unloaded (nominal return configuration)

WWSR OTV versus the time into the aerobraking maneuver are shown in Figures
2-2 & 2-3, respectively.

As can be seen, the first pass is a quick dip into the

atmosphere that reduces the OTV’s excess velocity by approximately 730 meters per second. The second pass takes the OTV down into the atmosphere almost as However, because of the reduced velocity of the OTV the It is during quickly as the first.

time for the vehicle to climb out of the atmosphere is much longer. aerobraking occurs.

this climb out period that the major portion of the velocity decrease due to This second pass reduces the excess velocity of the craft by The second pass leaves the OTV in an orbit that can be
1610 m/s while producing a maximum heat transfer rate that is slightly less than

that of the first pass.

22

1
I # I 1 I a I
circularized at LEO by a small propulsive burn (200 m/s versus a burn of approximately 2400 m/s needed for an all-propulsive return to LEO). Both graphs The were constructed using data obtained from a computer program used to solve the differential equations of motion of the OTV through the atmosphere [4]. graphs shown are for an unladen OTV returning from GEO. three passes in order to keep the heating rates low.
Two Pass Asrobrake Maneuver

For an OTV

returning a heavy payload to low Earth orbit (LEO) the option exists to make

J
1 1 I 1 1
.2
Y

0

0)

7550-

3

-First

25 -0 1

- - Second Pass
: : : : : : : : : : : : : : ; :

Pass

Time into Aerobrake Maneuver (sec)
Figure 2-2: Altitude history of A WWSR OTV during aerobraking.

I
0

In order to increase the safety and lower the heating rates of the aerobraking maneuver the OTV flies through the atmosphere with a negative

L/D

[6].

I I

Essentially, the vehicle is flying upside down, using the lift produced by the brake to pull the craft towards the Earth. This has two distinct advantages. If the OTV were to encounter higher than expected densities, which could catastrophicaily slow the vehicle sending it crashing to the Earth, the vehicle can rotate around its axis to produce a positive lift. reduce the deceleration. Guidance and Navigation. This will increase the altitude of the OTV and This is discussed further in the section on Aerobraking The second advantage of flying with a negative

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L/D is

23

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I
that it reduces the maximum heating rate by allowing the vehicle to make longer, but shallower passes through the atmosphere. Because the vehicle remains in the This effect is atmosphere longer, it can pass through at higher altitudes to produce the same deceleration. -Higher pass altitudes result in lower heating rates. discussed further in the Aerobrake Heating Section.

d
I

Two Pass Aerobrake Maneuver Velocity Decrement
10.5 7

1

9.5

out of Atmosphere

\

\

\\
0.5

\

-First Pass - Second Pass

\
\

-

1

----i : : : : :

7 . 5 f : :

i

:

z

:

:

:

0

100

200

300

400

500

600

700

4
800

Time into Aerobrake Maneuver (sac)
Figure 2-3: Graph showing velocity decrement of OTV during each pass of the aerobrake maneuver.

The two pass aerobrake maneuver was chosen for a number of reasons. importantly, it provides a margin of safety. disastrous.

Most

Aerobraking the OTV in one deep

pass, a maneuver called aerocapture, is possible, however, slight errors could prove If the OTV were to encounter a higher than predicted air density on By such a deep pass into the atmosphere, the velocity decrement due to drag would be
so large that the vehicle may not be capable of pulling out of the atmosphere.

making two, shallower passes the effect of this type of variation can be reduced and easily counteracted. The total aerobrake maneuver, from GEO to injection into LEO, will take only 8.6 hours; only 15 minutes of which is actual aerobraking.
24

The maximum

deceleration due to aerobraking will be approximately 1.5 g’s (for a two pass maneuver). This is below the maximum accelerations that will be encountered during other phases of the mission such as engine firings. Aerobrake Design The design of the aerobraking device for the proposed OTV has proven to be the basis upon which the majority of the other systems have been based. the control systems, and the treatment of heating effects. chosen a raked sphere-cone (see Figure 2-4). factors lead to the selection of this aerobrake. maneuvering where heating effects are small. The aerobrake design affects the orbital mechanics of the OTV, the materials required, For our OTV, we have Several This design has a blunt nose The raked sphere-cone has a low

configuration, similar to but not the same as the Apollo space capsule.

ballistic coefficient (W/CDS = 10 lb/ft2) which makes it ideal for high altitude

In addition, it is flexible enough to

require only a one to three pass aerobraking maneuver through the Earth’s atmosphere during the return phase of the mission from GEO to a low Earth parking

e
(a) Side view.

Figure 2-4: Aerobrake Geometry
25

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We have designed our brake to be a permanent appendage to the main body of the OTV. Although the brake will only have to be removed if severe damage occurs, the aerobrake has been designed to allow for easy servicing. must be by EVA so time constraints are important. and stability.

All servicing,

The OTV has been designed

with the engines protruding through the brake in order to provide better control Some characteristics are given in Table 2-1.

Table 2-1 Characteristics of the Aerobrake

2800 lbm 0.28 2
29000/(1.6)(25)2~

= 10 lb/ft2

The design of the aerobrake is derived from work by Park [9] and Bragg [l]. The fuel tanks and payload are arranged symmetrically around a reference force line (not axially). The aerobrake geometry is derived by raking-off a circular cone, blunting the apex with a spherical cap, and faring the frustrum by a fourth-order
polynomial. This is shown in Figure 2-5.

L/D equals 0.28, when the vehicle flies

at an angle of attack of the angle of attack.

- s o , with respect to the cone axis. By shifting the LO2

from tank to tank, the c.g. can be shifted in the yaw and pitch plane, changing Using this control the vehicle can even remain stable after the The cut-out openings for the loss of one engine [9]. The engines have extendable nozzles that are stored flush with the heat shield during atmospheric flight. lower. engines are at an off-stagnation point location where the heat- transfer rates are In the back side of the aerobrake, the tanks and payload are covered by a shroud which provides protection from solar flux, the heat of aerobraking, and the impacts of meteoroids and space debris.

26

I 4 I

...

Cone rake angle = 73.

1
I

Caontric conrtruction of blunted, ralud-off cone.

1 I
I

Figure 2-5: Geometric Construction of Raked-Cone
The reference force line (the X-axis, the axis of symmetry) represents the net aerodynamic force vector originating from the center of pressure at the desired flight angle of attack. The relation between the X-axis and and direction of travel
ag

is seen in Figure 2-6. As long
of attack will not be affected.

the c.g. moves along the X-axis, the trim angle Also, the c.g. can easily be shifted to bring The c.g. can be shifted by moving LH, This alone can control the

This means that changes in cargo and fuel loadings

do not affect the trim in this design.

the aerobrake to any desired trim angle. and LO,,

and also by gimballing the engines.

navigation of the OTV or the RCS rockets can also be used to roll the entire vehicle, achieving a time average angle of attack. The aerobrake is to be constructed of an inflexible heat shield material, cemented on metallic panels. These panels are supported by a system of beams and struts, as seen in Figure 2-7. Weight is distributed over the aerobrake, and

27

the structure of the aerobrake is integrated into that of the entire vehicle, thereby minimizing the total structural weight.
.-

-

c

angle = go.

va c-

'j"
R a k e angle = 7. 3

Figure 2-6: Flight Path Angle

a

28

A

0

I I

I

0

Figure 2-7: NASTRAN Model

I
29

The skeletal construction of the heat shield is shown in Figure 2-8. structure consists of aircraft-type skin, stringer, rib, and frame construction. riveted onto the structure. rolled into a circular shape. the ring as shown.

The The

skin, which serves as the inner mold line of the thermal protection system, ie In section A-A (see Figure 2-8b) notice that the ring is This ring has a flange for the purpose of riveting an The peripheral bulkheads are riveted onto The ring,

annular closeout plate at the bottom.

These bulkheads have a flange to which the wraparound edge

panels are to be riveted as shown ,in section E E (see Figure 2-8c). periphery of the heat shield.
is stiffened every 5 " .

annular plate skin, and bulkheads form enclosed structural boxes around the In effect, this provides a stiff outer hollow ring that This structural ring then serves to support brackets to

attach the heat shield to the OTV. The aerobrake must be transported by parts and assembled in space, because it is too large to fit in the shuttle or the aft- compartment of the external tank. Aerobrake Heating One of the most problematic aspects of the aerobraking maneuver is the heating of the aerobrake due to drag as it passes through the Earth's atmosphere. There are two methods of reducing the maximum heat transfer rate of the aerobraking maneuver; making multiple passes through the atmosphere and flying at a negative lift-to-drag ratio. Multiple passes allow the OTV to make shallower dives into the atmosphere. The heating rate of the brake is reduced because the aerodynamic slowing of the OTV is performed gradually over a greater time period. Figure 2-9 shows the heating rates of a two pass maneuver relative to that produced by single pass aerobraking [4]. Making two passes results in a decrease of the maximum heating rate by as much as 30%. As seen in Figure 2-10, the two pass maneuver can be optimized to give a minimum mission heating rate. Minimizing the heating rates results in a As it turns out optimization greater than minimum results in both passes being of approximately the same depth into the atmosphere. slightly deceleration on the second pass, however, this deceleration is well within the structural and physiological limits of the OTV and crew.
30

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I.) 01.r.11 .L.I...l
h. .l

ORlGfNAL PAGE 1 5

OF POOR QUALITY 2

.,,.I,.

J 1

11 *.&I

7.-

U

Figure 2-8: Construction of Heat Sheild

31

Two Pass Aerobrake Maneuver Aerobra ke Heating
1.2 1 .o
-u 0
Q)
-m

. *

..

m c .-u

0.8--

-First Pass - -Second Pass - - Single Pass
*

I
Q)
Z *

0
>
Q)

0.6
0.4

-I

0 -

QC

02 .
0.0

Figure 2-9: Graph showing the relative heating rates of a two pass aerobraking maneuver compared to that of a single pass.

32

80

-HEAT RATE - - - - ACCELERATION
PASS 2
\

60
d

PASS 1 ,-PASS 2

40
0

/

20
0

/ /

-.
\

-'PASS
I

1

0

I

I

I

I

1

zoo0

4ooa

6oob

8QIo

lorn

APOCEE AFTER FIRST PASS, n fd

Figure 2-10: Graph showing minimization of a two pass aerobraking maneuver. Note optimization of heating rates results in higher than optimal decelerations.
Figure 2-11 shows the heating rate during aerobraking for an OTV of similar configuration as the WWSR OTV. This graph was constructed for an OTV with a ballistic coeffiecent of 11.9 lb/ft2 making a one pass aerobraking maneuver [l].
The WWSR OTV has a slightly lower ballistic coefficient (10 lb/ft2) and will

therefore encounter lower heating rates than shown. of the ballistic coefficient on the heating rate.
30%.

Figure 2-12 shows this effect

Additionally, the WWSR OTV will

be performing a two pass maneuver that will reduce these rates by approximately A conservative estimate of the total (convective and radiative) maximum Compared with other heating rates encoutered by the WWSR OTV, as compiled from numerous sources
[1,8,9,10), has been calculated as 25 Btu/ft2-sec (28 W/cm2).

braking configurations, such as a lifting body or aerobraking tug, the WWSR OTV will produce relatively low heating rates. The relationship of these rates to the thermal protection systems of the OTV will be discussed in a following section.

33

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t

60

50

W/CDA

cy,

0.)

53 0 *
c
a

Convection

qMAX

Convection

11.9 Ib/ft2

= 34.6 8tu/ft2-sec

T
Q

M

=~ 25800 F (C = 0.85) = 5295 8tu/ft2

: 20
10

J

1

radiation without collision

1
0
50
100

150

200 250 300 Time from 400 OOO ft, sec

350

400

450

500

Figure 2-11: Graph showing the heating rate history of an aerobraking OTV of similiar configuration to the WWSR OTV.

In order to reduce the heating rates further, the WWSR OTV will fly through the atmosphere with its lift vector pointing towards the Earth. 0 This allows the OTV to make a shallower pass into the atmosphere because the lift produced by the vehicle will hold it down in the atmosphere longer producing the necessary deceleration. This longer but shallower pass produces the same deceleration as a quick, deep pass but with much lower heating rates since the densities encountered in the long, shallow pass are lower.

34

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NUMBER OF PASSES

I

2

3
4 5

EXTRAPOLATED

I
0

I

1

I

1

I

4 0

80

120 W/COA, psf

160

200

Figure 2-12: Graph showing the effect of vehicle ballistic coefficient and number of atmospheric passes on maximum heating rates.
Because of the large size of the WWSR OTV’s aerobrake and the use of a multiple pass, negative lift aerobraking maneuver, the heating rates produced will be well within the limits of the aerobrake’s heat shield and thermal protection system. Thermal Protection System Several studies [1,8,9,10] have shown that a one pass return trip from GEO to LEO, the raked sphere-cone with a ballistic coefficient of 15 lb/ft2 will experience between a 35 and 40 w/cm2 heating rate and 2 g’s of decceleration. Our computer simulation and other studies (8,101 have shown that thermal and mechanical stress can be reduced by 50% for a three pass return with a negative

lift vector and 30% for a two pass return with a negative lift vector.
in Figure 2-13.

This is seen

For a one pass mission the thermal protection system (TPS)
A two pass mission with negative lift effectively

would weigh 2300 lb and the supporting structure would weigh 2000 lb for a total aerobrake weight of 4300 lb.

35

I
I I I
reduces the total weight of the aerobrake to less than 2800 lb. designed for 2800 lb to give a large safety margin. hours for a three pass mission [8]. and heat reduction.
As noted in the preceeding section, 28 W/cm2 is the maximum heating rate

Our aerobrake is

Menees has shown that the

time required for the return trip from GEO to LEO is 6 hours for one pass and 14 This time difference is insignificant for a 14 day mission; therefore, a multiple pass return is advantageous for weight savings

encountered.

This heating rate at only one location on the aerobrake and for only Figure 2-14 shows the drop in heat flux Notice that the heating rate is small across

a few seconds of the re-entry maneuver. and pressure across the aerobrake. most of the aerobrake.

The heart of the thermal protection system is the high-temperature reusable surface insulation (HRSI) such as that used on the Space Shuttle. view of the HRSI is shown in Figure 2-15. refractory composite insulation (FRCI-12) reinforced with silicon carbide fibers. consisting of sintered

A cut-a-way
silica fibers

This material is a 12 lb/ft3 fibrous

The exposed surfaces of the tiles are coated

with reaction-cured borosilicate glass with SiB4 included as an emittance agent
[1,2]. The tiles are bonded with a 0.0075 inch thick layer of RTV-560 adhesive to

a 0.16 inch thick strain isolation pad (SIP) made of felted aromatic polyamide fibers (NOMEX) which is bonded to the aluminum skin with RTV-560. to 5 5 0 ° F and the temperature of the inner bond line to 3 5 0 ° F . and density of each material is given in Table 2-2. The thickness of the FRCI-12 is designed to limit the temperature of the outer bondline The thickness

36

1 I

250 r APOGEE AFIER FIRST PASS = l o a ~n mi ,

t

1
I
I I
/

'

---- ACCELERATION

9

I
9

5*r
400

'\

\

---- ACCELERATION
\
\ 0

HEAT RATE

,PASS

2

I

I

I

I

1

Figure 2-13: Reduction in heating rate and deceleration due to multiple passes.

37

a

9 r

u !
O!IWJ

Y

0

xnli IweH

Figure 2-14: Heat Flux and Pressure Distribution

38

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I 4 D
RCG coating

--

1
D
-

FRCl 12

-

'n I
Nomex felt SIP
Density ( Ib/ft3)
0.13
0.11

Bond adhesive: RTV 560

Aluminum skin /

Figure 2-15: Schematic Diagram of the Thermal Protection System Table 2-2 Thickness and Density of TPS
Material Thickness (inch)
0.015
0.0035 each

Tile coating
RTV-550 cement (2 layers)

SIP FRCI-12 tile

0.08
0.43

0.072 0.873

The skirt of the aerobrake, which is a region of high curvature, is covered by an array of rectangular tiles arranged in four circumferential rings. in the side view in Figure 2-16. Figure 2-16). This is shown The large, shallow cone area and the ellipsoidal

nose area of the aerobrake is covered with an array of hexagonally shaped tiles (see The hexagonal shape of these tiles has several advantages over The hexagon has a smaller perimeter-to-area ratio than a Also, rectangular tiles.

rectangular or square, which results in fewer or shorter gaps between tiles.

39

there are no long running gaps that tend to augment tile edge heat flux. Gaps are provided between tiles to accommodate the difference in thermal expansion between the tiles and the aluminum substrate, and thus prevent tile-t+tile contact. pressure gradient to prevent high tilegap heating. in Figure 2-17. Tilet= tile gap fillers- of woven ceramic cloth are used in regions of high entry-surface The gap filler fabric is shown

I

Figure 2-16: Thermal Protection System on Aerobrake

40

/

CERAMIC FABRIC

t
FI8ROUS INSULATION.

Figure 2-17: Gap Filler Configuration

The hexagonal shape also results in reduced stresses in the tile, in the tile coating, and at the tile bondline. To decrease the cost of the tiles, fewer and larger tiles are assumed rather then many small tiles. Often a problem exists about the convective heat-transfer frustrum edge.
[Ill.

rates at the

A circular frustrum produces high convective heat-transfer rates

Such Occurrences of high heat-transfer rates are avoided by contouring the

frustrum such that the surface curvature increases gradually toward the edge [5]. Another problem to avoid is after-body flow impingement, a narrow region around and extending behind the aerobrake where convective heat-transfer becomes very large. The base turning angle [9] is the angle between the free-stream flow This Shih has shown that this angle is about 1 5 " [ll]. vector and the line connecting the frustrum edge with the reattachment. angle is visible in Figure 2-18.

The best protection against this heating is to keep the structure of the OTV and payload within the "cone of protection" provided by the aerobrake, as measured by the base turning angle.

41

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0

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a

( :1

1

AOTV BLUFF BODY FLOW
1.00.m

Schlieren photo of Mach 13 flight of AOTV model in ballistic range at NASA h m o R e s e a h Center (courtesy of Intrieri).

1
I 1 I i I

0.8

-

0.r
0.a

0.1-

0.4-

7

1.1

I

X
I."

1.9

1

1

V * ? l 0c i - Y Vector

A

1
a

I I

.I

1.3

1.0

1.1

1.a

1.1

1.4

1.s

J
I 4

Figure 2-18: Base turning angle of 1 5 " is shown for the after-body impingement.

ORIGINAL PAGE I S OF POOR QUALITY
42

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Guidance and Navigation [ 1 1 In all past space missions requiring reentry into Earth’s atmosphere, only the destination coordinates, expressible with three parameters, have been specified. For the AOTV these three parameters need to be specified; plus the velocity vector, expressible also with three parameters, and the time when the velocity should be attained, are specified at the end of the atmospheric flight. thus renders it one of the most critical of all technological issues. One must assume that there may be errors in the time, position, and the velocity vector of the vehicle at the time of atmospheric entry caused by unforeseen events. The functional relationship between the position and the velocity vector of the vehicle at the completion and those at the beginning of the atmospheric flight indicates that such an error tends to be amplified: the exit parameters are a sensitive function of the entry parameters. atmospheric flight can be defined arbitrarily. (The beginning and the end of an Typically, the altitude of 150 km is The additional requirement makes the guidance and navigation problem very difficult to solve, and

I

1
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JI
I J I 1 I

considered to be the border between the atmosphere and the vacuum of space.) Therefore, any such errors must be corrected early during the atmospheric flight. Moreover, the density of the atmosphere at high altitudes, as determined from the Shuttle’s flight data, tends to deviate considerably; that is, typically by from the standard values.

+/-

25%

In order to reach the specified position with the

specified velocity despite the fluctuations in the atmospheric density, the vehicle must have a capability for controlling the flight path. The raked sphere-cone design provides two degrees-of-freedom control by varying lift. There are two methods of controlling lift. In the first method, the angle of attack is fixed, and the direction of the lift vector with respect to the

I
I

direction of vehicle’s motion is changed by varying the bank angle of the vehicle through the use of the Reaction Control System (RCS) engines. between two bank angles, the vehicle can achieve a time-averaged smaller than the L/D of the vehicle. used in all the pre-Shuttle space missions. By oscillating

J
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L/D which is

This method of control is similar to that In the second method, the angle of

43

attack is varied as well as the bank angle. ' T h e angle of attack can be varied by shifting the c.g.(ie. moving LO,).

A lifting vehicle induces a coupling between the directional and the lateral
motions. Therefore, a roll motion requires use of directional'RCS engines as well as lateral RCS engines. This problem has been solved by Gamble in Reference 3.

Because of the complexity of the navigation constraints, it is impossible to define uniquely the most optimum algorithm for guidance and navigation of a lifting AOTV during its atmospheric flight. However, general guidelines can be given:
I. It is advantageous to fly with a negative L/D (with the lift toward the Earth) because this increases the perigee height in exchange for a lengthened duration of the atmospheric flight and thereby lowers the peak dynamic pressure and heat-transfer rates.
2. The crosskange travel (orbital plane change) should be made mostly

during the descent phase; the ascent phase should be reserved for correcting for the errors caused by the fluctuation of the air density.
3. During ascent the vehicle should fly near maximum L/D so that if the atmospheric density is too large, the vehicle could roll 180" to produce a positive L/D which will raise the flight path and shorten the flight duration and avoid catastrophic loss of velocity.
4. When the navigational errors and fluctuations in density are such that the vehicle cannot reach the destination orbit, effort should be m a d e to

insert the vehicle into the correct orbital plane, sacrificing accuracy in apogee height and phase angle (longitudinal). The vehicle should then execute in- plane rendezvous maneuvers propulsively to correct for the errors. The worst situation for fluctuation in density is a lower than expected density on descent and then a higher density than expected on ascent. This guides the vehicle into a deeper dive in order to decelerate enough. And when the vehicle ascends it will encounter a very large density, resulting in excessive deceleration. However, calculations show that an L/D of 0.15 would be large enough to lift the vehicle out of the atmosphere on a worst case situation of density fluctuation of

e

+/- 25%

191.

44

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An idea by Menees to aid in navigation is to launch a small projectile in front of the aerobrake shortly before atmospheric entry [7]. trajectory, the density of the atmosphere can be deduced. then fed into the flight computer as an input to produce prediction and maneuver strategy.
a

By analyzing its

The density data is

more accurate trajectory

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45

Chapter 3 Engine Specifications

Over the past two decades major research has been conducted in an effort to produce a rocket propulsion system capable of reliable and efficient transportation
of payloads into and from Earth orbit. Until recently, the mainstay of NASA was

the RL10A-3-8. This engine was defined in 1967 as the engine for an imporved Centaur. The RL10A-3-8 was used in the early Shuttle Upper Stage Studies, and
it was found t o be lacking in several areas [3]. In an attempt to produce a state-

of-the-art high performance engine, study contracts were awarded to Aerojet, Pratt
& Whitney, and Rocketdyne to determine what could be done to improve upon

current designs.

This action instigated independent research into the development

of modern light-weight high performance engines. The engine type which came out of this research was the Category expander cycle engine.

IV

This engine was the first expander cycle engine specifically

designed for the OTV mission requirements. Many of the features designed for this engine have been carried through multiple design iterations to the present Pratt & Whitney advanced engines. materials (31. The advent of modern turbomachinery design in the 1980s has permitted the stresses acceptable to modern engine chamber designs to be nearly twice that of earlier engines.
As a result of this advance in technology, NASA has re-evaluated

At the time. of its design, the Category IV engine

maintained the highest chamber pressure (915 psia) thought possible for existing

the requirements it is placing on the technology goals of the OTV engine.

To this The

date, no engine design has met all of the requirements set out by NASA.

Pratt & Whitney 1985 Advanced Expander Cycle Engine, specified the RL100, shows the most promse in fulfilling the mission requirements currently set down for a manned OTV mission. Table 3-1 shows the 1987 updated goals for the OTV engine in comparison with the specifications of an unmodified stock RLlW engine
Ill.
46

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Table 3'-1
C o m p a r i s o n of 1987 OTV Engine Goals and the RLloo Engine Parameters Man- r at ing Fuel Oxidizer Vacuum Thrust Engines per Vehicle Mixture Ration O / F Mixture Ratio Range Inlet TemperatureHydrogen Oxygen Aerobraking Design Criteria Vacuum Specific Impulse NPSHHydrogen Oxygen Weight Length Reliability Operational Life Service Free Life
'87 NASA Goals

RLlOO Yes Hydrogen Oxygen 7500 lbf (per)
2

Yes Hydrogen Oxygen 7500 lbf (per) 2 Minimum
6.0

5- 7

6.0 5.5-6.5

37.8 R (TBD) 162.7 R (TBD) The engine must be compatible with aeroassist return of the vehicle to low-Earth orbit. 490 lbf-sec/lbm 477 lbf-sec/lbm
O

15 ft-lbf/lbm 2 ft-lbf/lbm 360 lbm (TBD) .9997 20 hours 4 hours

15 ft-lbf/lbm 2 ft-lbf/lbm 290 lbm 60 in. (TBD)
25 missions

(TW

An unmodified RLlOO meets or exceeds most of the requirements stipulated

I

by NASA for the technology goals of the OTV engine.

The chamber pressure

(1210 psia) and the vacuum specific impulse of the RLlOO are limited by the reduction in efficiency inherent in using small pumps [I]. Research is currently being conducted in an effort to alleviate the limitations of the smaller pumps by improving the purity of the materials used in the production of the pump shaft, seals, bearings, gears and thrust chamber. minimum
of

Advancements and innovations in this

area can be expected to raise the overall performance of the stock RLlOO by a at least five percent. In an attempt to compensate for the performance limitations experienced by the RL100, several design innovations have been incorporated.

47

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transportation.

OR1GiNAL PAGE I S OF POOR QUALITY
An extendable nozzle is incorporated into the engine design to allow a large expansion area ratio without the corresponding length requirements for storage and The extendable 'nozzle of the RLlOO produces an increase in Figure 3-1 shows how the extendable nozzle of a specific impulse of approximately 20 lbf-sec/lbm over the same engine equipped with a stationary nozzle (3). RLlO derivative engine functions to increase the expansion area ratio without increasing the overall length of the engine.

C.205

1

7
71 dil

Figure 3-1: RLlO Derivative Engine

The extendable nozzles produced by Pratt & Whitney are composed of carbon/carbon fibers coated with silicon carbide. the engine while also decreasing its weight. pumping system is reduced. 0 The turbomachinery of the RLlOO will be manufactured using state-of-the-art technology to permit the pumps to perform at a maximum output of 150,000 rpm
[3]. Figure 3-2 shows the flow schematic of the RLlOO engine at full thrust.

The use of these modern thermal By being a radiation-cooled nozzle

resistant materials over traditional nozzle materials increases the operational life of rather than a dump-cooled nozzle, the complexity and size of the engine and

The

performance requirements of the gears and turbines are shown here to be well above any engine with similar performance ratings. compared with similar engines.
48

By using high speed pumps,

the overall mass and displacement of the RLlOO is reduced by one-third when

J '

ORIGINAL PAGE I S OF POOR QUALITY

Figure 3-2: RLlOO Engine Flow Schematic

Unlike comparable engines, the RLlOO is self-contained and modular.

This

allows the engine to be easily separated from the OTV for inspection and maintenance [ 5 ] . As seen in Figure 3-2, the simplicity of the fuel transfer system requires that only two valves be shut to isolate the engine from the fuel delivery system [3]. While in free- fall, the engine can be removed from its support structure in the aerobrake and disconnected from the avionics of the OTV in approximately 3 hours (51. For the reasons stated above and the stipulated mission requirements, it was determined that two RLlOO engines with a combined thrust of 15,000 lbf would be the best main propulsion for a man-rated OTV mission.
to
Two engines were chosen

give

the

OTV

single-engine-out

propulsion

capability.

Current

research

conducted by Pratt & Whitney and Aerojet has shown that a reliability of 99.6% can be expected on a vehicle with two engines. failure rate between 0.03 and 0.05 (61. This data gives a nonindependent The fuel efficiency obtained by using two

e

engines is less than that obtained using a single engine, but the reliability and safety gained from a two engine design increases the expected life of the OTV. During the aeroassisted deceleration, the lift versus drag characteristic of the aerobrake will be changed by rotating the oblate aerobrake about its center of gravity. Having the engine nozzles extended would generate problems with the Figure 3-3 aerodymanics and cause severe deterioration of the nozzles themselves.
49

J *
I 4 I 1 I
shows how the engines are retractable flush to the aerobrake.

The engines are

inherently capable of tolerating the temperatures at the stagnation point in front of the aerobrake without oblation. This additional factor makes the RLlOO engine the ideal main propulsion system for W WSR’s man-rated OTV.

/
Figure 3-3: Aerobrake Engine Configuration

a

50

J *
Chapter' 4

Fuel System

The fuel system for the OTV will consist of six pairs of fuel/oxidizer tanks each with independent delivery and pressurization systems. Check valves will be incorporated into the delivery systems to allow isolation of each tank and permit pressure relief when necessary. and the entire system. Figure 4-1 is a schematic of the fuel system showing the check valves, delivery systems and pressurization systems for each tank

U

Figure 4-1: Fuel System Schematic
The fuel/oxidizer delivery system will independently draw and pump from each tank. The fuel/oxidizer will then be pressurized at a second pumping station just prior to entering the engine. This allows the fuel/oxidizer to be brought from its containment pressure of 7 psia to the inlet pressure of 17 psia for oxygen and
16 psia for hydrogen.

The internal pumps of the RLlOO then increase the pressure

of the reactants to over 1200 psia before they reach the combustion chamber. It will be necessary to have the fuel tanks pressurized at a constant level to 51

J
simplify the pump and turbine requirements 'of the delivery and propulsion system. The reactants that boil off, and are usually vented, will be used to maintain a constant pressure with the tanks. This will be accomplished by computer Any excess controlled venting and recycling of the gaseous reactants. Any excess oxygen will be shunted to the ECLSS to be used inside the manned module. hydrogen will have to vented to space. Since it' is unsafe to mix the reactants

during storage, separate pressurization systems will have to exist for each tank. Figure 4-2 is a schematic of the pressurization system for one pair of tanks with it connection to ECLSS.

Figure 4-2: Pressurization System of Tank Pair

The reactants used in the OTV's fuel system will be liquid oxygen and liquid hydrogen. Typical specific impulse values for the RLlOO using these reactants The reactants efficiency will be The would be between 470 and 485 seconds (61.

improved by addition of metallic aluminum suspended within the liquid hydrogen and the addition of a extendable nozzle to the main engine structure. efficiency obtained after these modifications will increase the specific impulse of the RLlOO to approximately 502 seconds (11 not assuming any design improvements during its production.

In determining the fuel requirements to fulfill the mission objective, a dry
mass for the OTV of 32,670 lbm was used. The requirements of the mission

52

J
I ,4
I

(worst case scenario) are to transport a 24,ObO lbm payload form LEO to GEO and then return to the Space Station without a payload. In this case the payload is For this mission, the This considered to be part or all of gn orbiting space platform.

engines are required to fire multiple times, assuming instanteous acceleration, to facilitate orbital transfer, course corrections and aeroassisted deceleration.

1 I a I

information is covered more in depth in Appendix 2. The fuel requirements for the mission were determined by using the rocket equation and iterating backwards through the required velocity changes. mission.

A total

fuel mass of 121,616 Ibm is required to produced the necessary transfers for this For safety the tanks will be filled to 125,000 Ibm. However, if necessary, The fuel is separated in 18,857 the tanks could be filled to capacity, 132,000 Ibm.

lbm of liquid hydrogen and 113,143 Ibm of liquid oxygen.

53

Chapter 5

Materials and Striietures

The basic ideas in designing the material and structure for the OTV is to make it lightweight, strong enough to withstand the stress of aerobraking, and heat and radiation resistant. Lightweight - If the structure is reduced by one pound, then the savings in fuel will be four pounds. Therefore, the structure will be made as light as possible in order to save fuel or to increase the payload. Strength

-

Since the OTV is space based, the structure will not have to be as The acceleration experienced during

heavy (strong) as a ground based structure.

the ascent t o GEO is at least an order of magnitude lower than the 1.5 g

S. deceleration of aerobraking I ] The structure will be under maximum stress during
aerobraking, not the ascent to GEO. from the heat shield. The stress caused by thrusting the engines is also in the same direction as the stress of aerobraking because the engines protrude The structure is designed to withstand the stress of aerobraking (1.5 g) and a safety factor of 1.4, unless stated otherwise. Heat resistance - The structure will be exposed to thermal cycling and high temperatures. During aerobraking the structure will be exposed to various high temperatures depending on the location of the part. during the exposure at that temperature.

All exposed areas must retain

sufficient strength a t the maximum temperature to withstand the stress that occurs Because the

OTV will be used many

times, the structure must be able to cycle between maximum and minimum temperature without losing a critical amount of strength. Thermal cycling will also occur due to exposure to solar radiation on one side, while the other side is shaded. Temperatures could cycle from - 1 7 5 " to +500'F if the OTV is not constantly rotating or if the solar radiation is not reflected. Radiation resistance - Solar radiation tends to weaken materials. The

54

structure must be designed to retain the required strength during the OTV's entire lifetime, or the'parts must be easy to replace.

0

I 1 I I

Structure command

-

The structure includes the engines, engine quick disconnect plate, EVA module,
docking/service/equipment/avionics assembly,

thrust structure, connectors, tanks, tank support rings, struts and supports, module, payload attachments, robot arm and aerobrake. A description of each follows. Enpines - The engines have extendable nozzles that protrude out the heat shield. During re-entry the nozzles will be retracted so that they are flush with the heat shield. Refer to Chapter 3 for a more detailed description of the engines. Mass = 580 Ib. Engine quick disconnect plate - This aluminum plate enables the engines to be disconnected quickly for repair or replacement. Thrust structure Mass = 100 lb.

c

l

-

The thrust structure transmits loads from the engine to The assembly consists of a cone-

the rest of the structure and to the payload.

frustrum-shaped composite structure consisting of honeycomb sandwich skin panels
(0.01 inch graphite/epoxy face sheets on a 0.079 inch thick nomex core of 0.91

lbm/ft3 density), a thrust distribution ring, and thrust beams. connectors.
Six tubular thrust beams

The assembly

begins directly below the command module and attaches to the tanks through the

8

(2

inches in diameter) are attached to the Total mass = 210 Ib.

aerobrake to uniformly distribute the load across the brake. Propellant tanks

1
0

-

A spherical design has been chosen because it is simple,

I

has good pressurization characteristics, and has maximum volume-to-mass ratio. The tanks can be spin formed and then chem milled to the correct thickness [7]. The tanks will be insulated by multi-layered insulation (MLI) which is described in the next section. Unlike ground-based vehicles, a space-based OTV is designed to The propellant will be held at a low operate solely in the vacuum of space and does not require that propellant tank pressures be maintained above 14.7 psia. pressure, 7 psia, to reduce the load on the tank structure, therefore making the structure lighter [7]. Figure 5-1 shows that the weight of the propellant tanks
55

l J
I

J '
decreases as the tank pressure decreases'. Reducing tank pressures below atmospheric requires that propellant saturation conditions be lowered so that the fluids remain in the liquid ph&e

[?I.

The LO,

tank is not pressure cycled The tank interiors are

(purged) between missions, but the LH, tank must be.

designed to support slosh baffles, inner bladder and a liquid acquisition device.

9001
700 a

1

I

i

TOTAL TANK STRUCTUR'E WEIGHT

I -

z
3

n

o

o

k

1

WEIGHT LHZTANK

I

I (3 w

300 200
100

PRESSURE (PSIA)
6039-4

Figure 5-1: The relation between tank pressure and structural weight.
The selection of material for the propellant tanks is important because many materials are sensitive to LO, and many can be embrittled by hydrogen. aluminum is often used for cryogenic tanks and works well. properties than 2219 [7].
2219 and 2090 alloys.

Also, low
2219

temperatures can reduce ductility and fracture toughness of some metals.

A new alloy, 2090-

T8E41 aluminum alloy ( Al-Cu-Li-Zr ), has been developed that has better

Table 5-1 is a listing of mechanical properties of the The Also, the 2090

The 2090 alloy has higher strength and less density.

higher strength-to-weight ratio will enable the tanks to be lighter.

alloy's tensile strength increases at lower cryogenic temperatures [SI. Table 5-2 lists the mechanical properties of 2090 at 298, 77, and 4 " K showing this increase in tensile strength.

56

Table 511

Table 5-2

Fracture Toughness and Tenrlle ?raprrtlas o f
~~ ~

ZOYO-lEEll a t 298 K a d 77 K

Temperature

Fracture Toughness S-T S-1 L-T Lt45 (MPdm)

Yteld Stress
535
600 615

U n i a x i a l Tensile Properties UTS X Elong. ( 1 (HPa) (on 25.4 m)

298 K

36 35

29
47*

16 13

17 15

5 65 715
820

11 14
18

77 K
41:

51' 511 641

-

-

-

The tanks were designed so that six pairs will carry the fuel necessary for maximum payload (worst case scenario). and holds 18856 lb of LO,.
3144 lb of

Each LO, tank is 4.2 ft radius, 100 lb,

Each

LH, tank is 5.8 ft radius, 250 lb, and holds

LH,.
The Kapton MLI is composed of layers of 3.75 micron aluminized See These are held together by widely spaced plastic pins. The tank is

e

MLI Figure 5-2.

kapton plastic (30 for LH,, 20 for LO,) each separated by a silk-net layer. built like a Thermos flask, with an evacuated double wall. way for heat to be exchanged between the layers.

There is no efficient

If the heat follows a tortuous

57

J
path by way of the plastic pins, the low heat conductivity of the plastic allows very little to get through. And if it radiates from layer to layer, the aluminum This insulation allows liquid MLI has coating on each sheet reflects nkarly all of it. hydrogen and oxygen to be store for extended periods of time in space. used polyurethane foam insulation to reduce volume and weight. some protection from meteoroids/debris.

been tested in a vacuum at NASA Lewis [I]. MLI was chosen over the commonly MLI also offers

I I

I
Figure 5-2: Schematic of MLI Thermal Protection

Connectors - The connectors attach the LH, tanks to the thrust structure and the LO, tanks to the command module. panels that allow the tanks to be modular. to the OTV. These connectors contain the disconnect
Up to six pairs of tanks can be added

This is shown in Figure 5-3 and 5-4.

An aluminum alloy is used
This handling might

because the tanks will be connected and disconnected often. damage a composite material and cause delamination. and LO,) below the plumbing and interface with the propulsion system.

For each pair of tanks (LH, Male connectors are located

two polygonal frames of aluminum support all the propellant system

LH, tank and above the LO, tank.

Female connectors are located at

the thrust structure below the command module and at the top of the command module. Mass = 75 lb per set. Tank support rings, struts, and supports - These components will be RCA-2606114 graphite/epoxy. tape/low microcracking epoxy. This is an ultra-high modulus graphite unidirectional The material’s resin solid content is 40% with a

nominal prepreg thickness of 0.0025 inch (31. The low thermal conductivity of this
58

ORIGINAL PAGE T S OF POOR QUALITY

1
I
material will prevent heat
loss

ORGINAL PAGE I S

OF POOR QUALITY
from the cryogenic tanks. The ceramic

graphite/epoxy also has a much greater strength-teweight ratio than metals, allowing the structure to be lighter.

of the G/E.

--

Figure 5-5 shows the ultimate tensile strengtb

Figure 5-3: Drawing of OTV with two modular tank sets connected to the central structure.

59

ORIGINAL PAGE IS OF POOR QUALITY

6039-2

Figure 5-4: Tank support structure and connections with the centr a1 structure.
RCA has performed thermal cycling tests, radiation tests and combinations of

the two tests resembling 10 years at GEO orbit, and has found that this graphite/epoxy retains much of its strength. The radiation environment was The dose rate during simulated by exposing the test coupons of the materials to an electron beam of energy 12 MeV to a total ionization dose of 3x108 rads. irradiation was 3x108 rads/hour. space radiation environment [3]. This dose rate is about four orders of magnitude The thermal environment was simulated by

higher than the space dose rate and represents the worst-case simulation of the thermal cycling (3000 cycles) between temperature extremes of -300 and 160' F.

A

transition rate of about 11' F per minute was used for thermal cycling [3]. The results showed that beginning of life tensile strength = 135 ksi, end of life tensile strength = 110 ksi [3]. The tensile strength of aluminum is closer to 50 or 60 ksi.
60

J
I I I 1 I e I
ORlGtMAL PAGE 8% OF POUR QUALIN

F,
,

'"(1

J
I J I 1 I
Beginning-of-Life and End-of-Life Longitudinal Tensile Strength Of U CraphitelCpoxy " Composite

Figure 5-5
Graphite/epoxy tubular struts (2 inch diameter) are used to attach the tanks
to the tank frames.

Figures 5-4 and 5-6 show the skeleton structure of the OTV.

Twenty-four struts are used for each tank (see Figure 5-6). To prevent buckling of the tank wall, strut angles must be selected such that the tank does not experience negative deformations or compressive stresses. is 220 lb ( for a pair of tanks). A G / E tubular support ring ( 5 inch The mass of these components diameter) will support and separate the two tanks.

I
0

I I

J
I 4
61

MULTI- LAYER INSULATION

/

DISCONNECT PANELS / 6039-3

Figure 5-6:

Tank Support Structure

62

Command/EVA module structure will be

-

The modules will be semimonocoque. The command module also

The outer The holds the

skin will be stiffened with ring frames and skin stringers (see Figure 5 7 ) .

2090 aluminum.

graphite/epoxy tube ring support and the female connector for the propellant tanks. The mass for the structure of the command and EVA modules are 10700 and 1500 lb, respectively.
300 lb.
Docking/service/equipment/avionicsassembly

The modules also contain three hatches with a combined mass of

-

This assembly will be attached

to the side of the EVA module.

The assembly provides for external mounting of A peripheral latch/release system for payload The

equipment and avionics, a universal docking system, and service connector panels for fluids, gases, and electric power. accommodation and robot arm are attached to the top of the EVA module. arm is discussed in Chapter 12. Mass = 180 lb (excluding the robot arm). Aerobrake - The aerobrake will have to repeatedly withstand very high temperatures and heating' rates for a short period of time, and keep the temperature of the structure below 3 5 0 " . reach 35 to 40 W/cm2. For a one pass mission the maximum

temperature on the surface may reach 2000" and maximum heating rate could Therefore, the OTV will conduct multiple pass missions, The aerobrake is discussed in detail in the aerobraking thus reducing the maximum surface temperature to below 1000° and maximum heating rate to 20 W/cm2. chapter. Mass = 2800 lb.

Heat and debris protection - The OTV structure and payload need to be protected from the heat of aerobraking and collisions with meteoroids/debris. and payload. space debris.

A

very thin aluminum foil extendable blanket will be used to surround the structure The high reflectivity of the aluminum foil will reflect most incident The probability of puncture by micrometeoroids is low and could be solar or heat flux away from the OTV, and will provide some protection from substantially reduced if the OTV were to be stationed within a depot when not in use [Z]. And even if a micrometeoroid did puncture a fuel tank, the tank would leak but would not fail catastrophically [4]. Therefore, a heavy meteoroid protection shield will not be used.
63

Surface coating radiative heat

-

All surfaces that will be exposed to solar radiation and maneuver will be painted white White paint (293 and S13GLO) has the best The absorptance is 0.18 and the emittance is

transfer from the aerobrake

(excluding the heat shield tiles). reflectivity and lowest absorption.
0.9.

This reduces the amount of solar radiation that is changed into heat.

Skin
txme (0

Figure 5-7: Semi-monocoque Structure t = 1.0 in t, = 0.5 in b = 10.0 in b, = 1.5 in

64

Chapter 6 Ambient Heat Transfer

c
The ambient radiation received by an object at one astronomical unit from Sol is known as the solar constant and has a value of 1353 W/m2. the incident radiation. surroundings. this value assumes the object to be located at the equator of the Earth and perpendicular to The actual value received by the OTV will be within 15% The OTV will also radiate excess heat to the
of the solar constant [2].

Conductiion and convection can not occur into a vacuum, therefore,

radiation will be the only way for the vehicle to lose heat. The energy flux lost to the surroundings by radiation can be determined by using the equation:

In this instance the ambient temperature of the surroundings is approximately 4 " K not including the Earth. The surface temperature of the OTV is limited by Assuming conduction from the temperatures of the fuel and manned module.

engine and crew quarters through the support structure of the vehicle, the surface temperature of the vehicle would be at a maximum of 295°K for GEO conditions.
This value can reach as high as 350 ' k during aerobraking [3]. The average

emissivity value for the OTV materials is 0.89 [2]. times as much energy as it is radiating.

Assuming this value, the

radiation flux to the surroundings is 80.740 W/m2. The OTV is receiving 16.75

This influx of energy will cause a loss in fuel due to boil-off.

To partially The

alleviate this problem, the OTV will be coated with materials that overall have a low transmissivity and absorptivity while maintaining a high reflectivity. relationship of these three values can be seen in the following equation:
p + a + t = l

[z]

Where p is the reflectivity, a is the absorptivity, and t is the transmissivity.
65

Polished aluminum, aluminum coatings, or gold will be used to insulate areas (fuel tanks) where radiation absorption is to be kept to a minimum. Those areas (aerobrake and. exhaust nozzle) - where radiation emission is. required will be coateh with silicon carbide and ceramic tiles similar to those used by the Space Shuttle. The manned module will be constructed of aluminum with a white metallic coating. Since this module is surrounded by six sets of fuel tanks, this coating will be all that is required to maintain a minimum absorption of energy. for the crew compartments and the fuel tanks. The combination of these materials. will allow the vehicle to maintain a relatively constant temperature From data already obtained, the expected surface temperature of the OTV will be approximately 2 0 0 ° K [I].

66

Chapter 7 Electrical Power System

The Electrical Power System (EPS) produces electrical power for the OTV during all mission phases. The EPS onboard the WWSR OTV will consist of two hydrogen (H2) oxygen (02)fuel cells and one bipolar nickel-hydrogen battery. The fuel cells will be United Technologies’ latest version of Shuttle-technology power plants (which are thirty percent lighter than current cells).

[I] These fuels

cells are extremely reliable and provide the most effecient means of production of electricity for the OTV’s mission (two week duration at 20 kilowatts maximum). The Ni-H battery represents state-of-the-art technology in energy storage. lightest, most reliable, and most powerful of all spacecraft battery systems. It is the

I
I I 1 I

The fuel cells produce direct current electrical power through a controlled chemical reaction of the hydrogen and oxygen. The hydrogen and oxygen reactants will be cyrogenically stored in the main tank sets. system. Proper reactant gas pressure is maintained in the tanks by small heaters controlled by the onboard computer Additionally, the oxygen tanks will double as the storage tanks for the life The fuel celIs will simultaneously produce 28 volts of direct The total maximum, onboard power support systems.

current at a maximum power of 10 kilowatts. payloads if needed.

requirements are 7.5 kilowatts; the extra capacity is available to power the OTV’s

The cells will be actively redundant, as each cell is capable of Power production A

providing full mission power in the event that one goes o f line. f

is controlled by the Electrical Control Unit (ECU) which is part of the fi.:4 cell. The ECU controls the reactant flow rate as determined by the power demand.
is pure water.

by-product of the production of electricity by the reaction of hydrogen and oxygen This water, on the order of 6 kg per hour, will be stored in one of two water tanks and can then be used for thermal control or human consumption. Total mission energy is expected to be approximately 2000 kW-hours thus requiring about 1390 lbs kg of oxygen and 210 Ibs of hydrogen [l]. A single nickel-hydrogen battery will provide emergency power backup, line
67

transient supression, and autonomous startup capability. fuel cells during a mission.

The Ni-H battery will be

fully charged at the Space Station and will be actively recharged by the onboard The battery will be capable of providing reduced emergency power for approximately two hours in the event of a catostophic failure of the fuel cell system (a source of electricity is needed to start or restart power production in the fuel cells). It's main function, however, is to provide a source to smooth power surges caused by major subsystems coming on line [3]. Electrical power distribution is controlled by the Electrical Power Distribution System (EPDS). The EPDS converts and controls the flow of electricty to the Additionally, the EPDS monitors and controls the subsystems of the OTV. battery.

i 1 1

reactant gas levels and pressures, surge supression, and charging of the Ni-H The EPDS is connected to the Data Management System for status output and crew control.

68

Figure 7-1: The Electrical Power System

ORIGINAL PAGE I S OF POOR QUALITY
69

Chapter 8 Environmental Control and Life Support Systems

The ECLSS for the OTV will be largely based on present technology used on the Space Shuttle [l]. This was decided to eliminate the cost of research and development on new systems. systems [4]:
1. Atmospheric Revitalization

Also, present ECLSS technology on board the Space The life support will have the following

Shuttle has proven to be highly reliable.

2. Thermal Control
3. Crew Systems

A system integration flow chart o the above systems is shown in Figure 8-1. f
Each of these systems will be discussed in more detail below. Atmospheric Revitalization This system is given the task of providing fresh air to the crew members and is therefore the most important system. It is illustrated in Figure 8-2.

Air is drawn into the system by fans located strategically throughout the crew and command modules particularly around the cockpit area. After passing through the intake ducting, the air is filtered by a debris trap to remove dust and foreign particles. The exiting air is then divided into several other air streams which are individually processed. One stream enters a unit of canisters containing lithium hydroxide, copper sulfate, and activated charcoal. The lithium hydroxide extracts carbon dioxide and the charcoal removes air impurities for odor control. The copper sulfate extracts ammonia.

70

I

!

__

.-

-

-I-+---

c

-.-

-t

J

I

I

1
1.
I

Ii I
I I

I

I

-------c--

--PI-'

4

i i
t
I
I

I.

J

1

Figure 8-1: OTV ECLSS

ORIGINAL FAGE k ' 3

OF PO8R QUALfTY
71

...-_ ORIGINAL PAGE ES OF POOR QlsALlrV

Figure 8-2: OTV Atmospheric Revitalization System
Only two canisters can fit into the system. other is a reserve. One is actually used and the When the components in the active canister are consumed, the Consequently, the canisters Canister life, with

system automatically switches to the reserve canister. three astronauts, will last 32 hours. day mission.

must be changed by the astronauts to insure system operation.

Thus, 11 canisters will be needed for a 14

These will be stored above the system in one of the package

compartments for quick and easy access.

A temperature sensor in the crew and command module activates a valve that divides the air. A portion enters the air bypass duct where micro-organisms are filtered, and the other portion enters
the cooling system. This cooling system is actually a condensing heat exchanger

The purified air then rejoins the main airflow.

72

that cools the air below the dew point.

The heat build-up that occurs is reduced The air Fresh oxygen is

by the water_'cooling loop (this will be discussed in the next section). exits the heat exchanger and is then rejoined with the bypass air. new mixture is vented into the crew and command modules.

immediately added to the mixture from oxygen in the propulsion system, and the It is estimated that

an air flow rate of 353 ft3/min is needed to operate the system.
This system should maintain an air temperature between 5 5 ' F to 70'F and an air pressure of 14.7 psia. mix. Nitrogen will be stored in separate tanks adjacent or across from the system so that the atmosphere will have a 20% 0, and an 80% N,

A control in the crew module will permit desired selection of the

temperature.

A repressurization airlock will be needed in the EVA module.
a cylinder whose diameter is 4 feet and whose height is 7 feet. outer edge of the EVA module. into the OTV from the Space Station. During repressurization entrance door. EVA operations, the fully suited astronaut

This airlock is

It is placed on the

This airlock will facilitate crew exit and entrance

will

enter

the Upon

port or airlock from the command/crew module and seal the Exit from the module may then be achieved accordingly.

completion of EVA, the astronaut reenters the port, seals the exit hatch, and repressurizes the port. command/crew module. The air lock is repressurized by air that is bled from the It is estimated that the airlock will require about 6.65

lbm of air. This amount of air is not expected to effect the amount needed in the crew/command module, whose air requirements are about 202.1 lbm. Additionally, the EVA module will not be pressurized at all, thus eliminating the need for a separate repressurization system. of required 0, and N,. vacuum environment. This will also reduce the amount The astronauts will perform their necessary work in a

Entrance and exit into this portion of the module is made

through a door in the airlock. Since the OTV will be pressurized with and docked alongside the Space
73

Station, a full-scale repressurization system is unnecessary for the entire vehicle before mission operation. The OTV, before its severence with the docking bay, will activate its ECLSS. A safety factor of 1.25 has been included for the metab06 requirements of 0, and N, to account for leaks in the system. A monitor system will also be included to measure the oxygen, nitrogen, and carbon dioxide levels. carbon dioxide removal. case of malfunction. This system will control oxygen and nitrogen supply and Information from the system will also alert the crew in Table A1-2 in Appendix 1 gives the mass and power Most of the power

requirements to operate the complete air revitalization system. exchanger. Thermal Control

will be needed to operate the ventilation system, the fans, and the condensing heat

A thermal control system is needed to remove excess heat away from the
command and crew modules of the OTV. astronauts. This system is illustrated in Figure 8-3 and 8-4, and is comprised of a water and a Freon cooling loop. two heat exchangers. Water, cooled from the Freon interchanger, is routed to The These heat exchangers cool the crew’s drinking water. This excess heat originates from the electronic equipment on board, the fuel cells, the windows, and body heat from the

water is then fed into the condensing heat exchanger (humidity control heat exchanger in the diagram) of the air revitalization system. The water passes into the inertial guidance heat exchangers which cool the guidance system of the OTV.

74

J
I I I 1 I I I J I 1 I

0

A r i O n l C s bay

cooling assembly

J
Figure 8-3: OTV Water Loop [I]

I
e

I I
Figure 8-4: OTV Freon Loop [I]

J
I I

75

Then, the partially heated water is routed to a water pump that returns a portion of the water back to the Freon interchanger and feeds another portion to both a cold plate and to the avionics bay cooling assembly which cools the avionics, in the cockpit area. The water from this assembly goes into a cold plate where From here, the water is routed through the water temperature is partially lowered. heating.

window and hatch passages to cool these structures from sunlight and aerodynamic The water then returns to the Freon interchanger.

It has been estimated

that a water flow rate of 221 lbm/hr is needed for adequate heat removal. The Freon loop receives all the heat from the water cooling loop through an interface heat exchanger and cools the water to about 4 1 OF. The flow rate must be at 780 lbm/hr for proper operation. tremendous heat absorption. The Freon then flows into the water flash evaporator where it is cooled to
38.8' F.

A pump circulates
These systems are
O F

the Freon as shown which flows to the fuel cell and power system heat exchangers. Freon cooled accordingly in which the Freon now has been heated to 158
due to

This evaporator vaporizes water to the outside of the vehicle and uses the The heated Freon is piped This Water is then sprayed

heat of vaporization of the water to cool the Freon.

into a low pressure chamber through minute passages in the chamber walls. pressure chamber is equipped with a vent to the outside. the Freon.

onto these walls where it evaporates, and this evaporation extracts the heat from The extracted heat is later vented to the outside in the form of steam. The water that is needed for this operation should come as a by product from fuel cell operation. After this process, the Freon is returned to the interchanger. directly into the crew module. alert the crew of leaks. Due to its toxic

nature to humans, the Freon loop must be adequately sealed since it will be placed Sensors must be installed around this location to The mass and power requirements for the thermal control

system can be found in Appendix Table AI-2. Power will be mainly needed to operate the pumps found throughout the water and Freon loops.

76

Crew Systems The crew systems for the OTV will facilitate eating, drinking, sleepink hygiene, and liquid/solid waste disposal. diet for the astronauts. variety. The requirements for this system are given in Appendix Table A1-2 Dehydrated and frozen foods will compose the main At first, dehydrated foods were only considered, but owing to the rather long mission duration of 14 days, frozen foods were added for food The dehydrated food is. rehydrated by adding water (hot or cold, Drinking water depending on preference) from the potable water system. The frozen foods are stored in a small freezer and prepared in a small microwave oven. by the water- cooling loops before its actual use. will be furnished from water produced from fuel cell operation, which will be cooled Potable water can also be obtained from the condensation that forms from the cold plates in the thermal control system and from condensation that forms from the condensing heat exchanger in the air revitalization system. An emergency water storage tank will also be provided in case of system failure or malfunction. Human wastes are handled with a toilet that separates the solid and liquid wastes which are placed into individual chambers by pressurized air. waste (which also contains air odors) is injected into a separator. The solid wastes are stored until the OTV docks with the Space Station, where as the liquid This device uses The air odor
a rotating shell to force the liquid to the outer perimeter where it is removed and

piped to the waste water tank for eventual ejection to outer space. the cabins.

is directed through a charcoal filter to remove the odors and then is returned to

Hygiene will be provided through towel wipes laced with an antiseptic and compact shower bags like the ones found on the Space Shuttle. components will be taken from fuel cell operations. Water for these The water from the fuel cells

will be at a temperature of 160' F and will be maintained at this temperature until

it is used to prevent the growth of bacteria.
water cooling loop to about 110' F.

Prior to use, it will be cooled via the

77

The crew will use compact sleeping bags that will suspend freely from the sides of t h e - crew module interior to sleep and rest. Three such bags will be included so that all crew members may sleep or rest simultaneously. Radiation Dose limits for radiation workers on earth are currently set at 5 Rem/year. Such limits are unrealistically low for astronauts [2]. Astronauts will be exposed to the danger of radiation unless they are protected with heavy radiation shielding. But space travel is a hazardous undertaking, and reducing the possibility of mission failure due to one type of hazard significantly below other types of hazards may be undesirable. Increasing the radiation shielding may in turn reduce the safety margin in propulsion or life support by adding too much weight, and may increase the overall risk of mission failure. The amount of radiation that the astronauts of the OTV will receive during normal orbiting is negligibly small, even after 14 days. As seen in Tables 8-1 and

-

*

8-2, the OTV will receive 0.8 RAD per day (0.9 REM per day) at GEO and less
than 0.1 RAD per day at LEO. Most danger comes from solar flares and the van Allen Belt. The time spent in the van Allen Belt on re-entry is very small, even As shown in Table 8-3, Menees calculates that even for with multiple pass entry.

a 3 pass mission the OTV will graze the lower edge of the van Allen Belt only on the first pass, because the belt extends between 2.5 to 7 earth radii [3]. Solar flares, on the other hand, could cause significant radiation exposure. The protection that is afforded in the OTV is the structure of the OTV, the structure of the tanks, the propellant in the tanks, and the astronaut’s space suits. The astronauts also have the option of turning the aerobrake to block radiation from solar flares if no pertinent operations are being performed at the time. During our worst neglecting solar flares. mission. case mission, the OTV will receive only 7.2 REM, As seen in Figure 8-5 and Table 8-4 this is a negligible

e

amount and will not cause illness nor decrease the astronauts ability to perform a Table 8-5 also demonstrates that the radiation will have an insignificant

78

ORIGINAL PAGE IS OF POOR QUALITY

effect on the electronic components, lubricants, hydraulic fluids, glass, ceramic, and

I

structural metals.

Table 8-1
SPA^ RADIATION DOy R A (RAD/DAY)EXPECTED ORBITAL ~ FOR Musio@
1.0gm/cml 10 qn:cm’

1
0

Orbital ahitude (km)
300 Equator Polar 400 Equator Polar 600 Equator Polar 1000 Equator Polar 3000 Equator Polar 10,000 Equator Polar 31,000 Equator Polar

0.1 gm/Cml Van Allen 3 x 102 5 x 10’ 2 x 10’ 4 x 10’ 1 x 10’ 3 x 103 1 x lo5 3 x IO’ 3 x 105 I x 105 1 x 106
4 x 10’ 4 x 10’ I x 105

Other‘ <0.1 90
<0.1

Van Allen
<0.1 0.1

Other

Van Allen
<0.1 <@.1 0.3
<0.1

Other

<O.l
3 <0.1 3
<0.1 4

<O.l
0.2 <0.1 0.2 co.1 0.2
<0.1 0.2 qo.1

1

100 <o.i <0.1

11 x

I 0.2 5 I 50

2
0.4

<O.l
4

15
5

200 ~o.1

MO <O.l
400
16

I5 lo00 300
30 10 3 0.6

800

<0.1 5 <0.1 6 16 16

300 100 10 3 0.5 0.I

0.3 <0.1
0.4

0.8 0.8

‘ AI1 entries have la limits of & a factor of 3. Van Allen dose rates calculated for orbiu in 1970, active Sun. usuming no more high altitude nuclear detonations. Galactic and flue do# aC culated for solar maximum, 1% flare probability, averaged over 6 m o n t h Other: i n d u d a flare and galactic radiation



Table 8-2
S ? ARADIATIONDasr RATES (REMIDAY) X P E C ~ FOR ~ E ~D
Orbital altitude
(km)
ORBITAL MISloNs’

0.1 gm/cm’
Van Allen 3 x 102 5 x 10’ 2 x 10’
4 x

1.0 gmtcm’

10gmlcm’
Van Allen
Other

Other‘

Van Allen <0.1
<0.1

Other

300 Equator Polar so0 Equator Polar 600 Equator Polar 1000 Equator Polar 3000 Equator Polar 10.000 Equator

<O.l
250 <0.1

<0.1
6

<O.l
<0.1

IO’

300

1.3 0.3
6.5

<0.1
8

I x 10‘ 3 x 103 I x 105 3 10. 3 x 105 I 105 1 x 106
4 x 10s 4 x 105 1 x 10’

<O.I
500

1.3
65

<O.I 800
c0.1
<0.1

20

<0.1 10 <0.1 12
<0.1

0.3 <O.l 2
0.4

<0.1 0.2 <0.1
0.2

<0.1 0.2
<0.1

16 5

0.2

1200
103

1300 400 35
12

I5
<0.1 ia
50 50

Polar 31,000 Equator Polar

2

330 I10 10 3
0.5

<0.1 0.3
<0.1 0.4 0.9 0.9

50
4 x 103

3
0.6

0.1

All entries have la limits of f a factor Of3. Van Allen dose rates calculated for orbits in 1970, active Sun. assuming no more high altitude nuclear detonations. Galactic and flare do= calculated for solar maximum, I s; flare probability. averaged over 6 months. Other: includes flare and galactic radiation.



79

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ORIGINAL PAGE IS OF POOR QUALITY

Table 8-3

Duration and apogee altltudc for multiprrr . .oris aeroass1 1-

Nuder o f atmospheric

CEO to- Shuttlcorbit

passes
1

hr

I e I

5:
I1
3 #2

6.1 10.0 14.1

km 400 11,661 400 16,773 7.670

Alt,

NO

INCAPACITATION

c n

0

5

is

1

Y

4 z

Y

r

750
ACUTf WOLf DODY DOSf

(MM)

Figure 8-5: Incidence of sickness and death from acute radiation.

80

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1 I I 1 I I
Table 8-4

ORIGINAL PAGE IS OF POOR QUALITY

EXPECTED EFFECTS ACUTE of WHOLE-BODY RADIATIOS DOSES Acute Dose (roentgens) Probable cflect

-

e5 -0
80- 120
130-170
180-220

0

270-330

I I I

c

400-500

550-750

1000

So00

No obvious cflect, except possibly minor blood changes Vomiting and nausea for about I day in 5 to 10 percent of exposed personnel; fatigue, but no serious disability Vomiting and nausea for about 1 day, followed by other symptoms of radiation sickness in about 25 percent of personnel; nodeathsanticipated Vomiting rnd nausea for about 1 day, followed by other symptoms of radiation sickness in about 50 percent of personnel; no deaths anticipated Vomiting and nausea in nearly all personnel on first day, followed by other symptoms of radiation sickness; about 20 percent deaths within 2 to 6 weeks after exposure: survivors convalescent for about 3 months Vomiting and nausea in all personnel on first day, followed by other symptoms of radiation sickness; about 50 percent deaths within 1 month; survivors convalescent for about 6 months Vomiting and nausea in all personnel within 4 hours from exposure, followed by other symptoms of radiation sickness; up to 100 perant deaths; few survivors convalescent for about 6 months Vomiting and nausea in all personnel within 1 to 2 hours; probably no survivors from radiation sickness Incapacitation almost immediately; all personnel will be fatalities within I week

Table 8-5

RADIATION DAMAGE THRESHOLDS CERTAIN CLASSES OF MATERIALS FOR Electronic components Polymeric materials Lubricants, hydraulic fluids Ceramic, glasses Structural metals. alloys
10'-10' rad

10'-109 rad rad 106-108 rad 109-10" rad

l'I' O-O

a

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Chapter 9 Guidance, Navigation, and Control

The main purpose of the GNC system is to:
1. Determine the position of the vehicle.

2. Determine the magnitude, direction, and change of vehicle velocity.
3. Calculate and control manuveurs rendezvous will a target satellite.

to

reach

specified

position

or

The position and the velocity of the OTV will be determined by information received from the planned Global Positioning System. This system, which will be composed of satellites positioned in 12-hour orbits, will produce signals that can be used to acurately determine vehicle position and velocity (time rate of change of position). Onboard, autonomous GNC will be provided by a combination of stellar The, stellar tracker is tracker and laser-gyro inertial measurement units (IMU’s). from precise angular measurement of selected stars.

an opto-electrical device that is used to obtain vehicle attitute and position data The stellar tracker onboard the WWSR will have three axis imaging capability and a larger star catalogue than the Space Shuttle providing much higher accuracy and longer on-time [6]. The IMU provides vehicle attitude and velocity data from internal laser gyros and accelerometers. This part of the GNC system will play an important role during the aerobraking maneuver when the stellar tracker is unuseable and reception from the GPS system may be hampered by ionization of the air flow around the OTV. The GNC system will be controlled by the general purpose computer systems. The computers will perform position and velocity determination from the various GNC sensors, will calculate needed maneuvers, and control the main engines and the attitude control system (ACS) to carry out the necessary changes. Initially the WWSR OTV will be equiped with Ku-Band Rendezvous radar. This radar, which will also double as a communications link, will provide automatic 82

target detection and tracking to provide the range, velocity, roll, and pitch of a target satellite. This system will greatly reduce target location errors by allowing The Ku-Band radar can track a satellite with an pre-rendezvous flight corrections. miles [I]. Reaction Control System (RCS) The Attitude Control System will respond to flight software commands and GNC inputs via the Data Management System to control the OTV’s attitude, trajectory, rendezvous maneuvers. The ACS jets will use N,H, hydrazine fuel and will each produce a thrust of 111 Newtons at a specifc impulse of 220 seconds (51. There will be a total of 36 jets arranged in 8 locations to provide complete translational and rotational control of the OTV during rendezvous, docking, Four stations, each with four thrusters, are trajectory correction and aerobraking. located around the EVA module.
for these four stations.

active transponder at a range of 400 miles and a dead satellite at a distance of 14

A tank within the EVA module supplies the fuel

The remaining four stations are attached along the rim of Each of the stations has its own hydrazine fuel tank.

the aerobrake.

These stations have five thrusters each, with some firing through

the edge of the brake itself.

The OTV will carry a maximum of 2900 lbs of hydrazine fuel.

83

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Chapter 10

Data Management S y s t e m

The Data Management System (DPS) will control and monitor the OTV during the course of each mission. Some of these functions include:

1. Support of GNC system, including calculation and control of vehicle position and trajectory.
2. Monitoring and control of vehicle subsystems including electrical power,

environmental control, and main engines.
3. Processing vehicle data for radio transmission and responding to uplinked

vehicle commands. The DPS systems will consist of 3 onboard computers, 2 mass memory units,
3 crew input/output stations, and the data bus network.

The onboard computers

will be IBM’s new 1750A (Air Force Standard) avionics system [l]. These high speed, high capacity machines were choosen because of the enormous computing power needed during the aerobraking maneuver. highest computing speed in the smallest box. write-once optical discs. The IBM system provides the The mass memory units will be

Each of the two memory units will contain copies of the

flight software and star catalogue for the stellar tracker and will provide memory

for mission data storage.
The forward flight deck will consist of three flat screen plasma displays, two keypads, and the numerous controls and switches that operate all of the subsystems
of the OTV.

All phases of operation of the OTV are controlled from the flight

deck, either automatically though the computers or manually. used as a work station o f the flight deck. f

The remaining

display and keyboard, attached within the avionics component compartment, can be

The data bus network provides a means of communication between each of the vehicles subsystems and the DMS. The data buses will be high density optical

84

cable

to

reduce

weight,

size,

and

electromagnetic

interference.

The

multiplexer/demultiplexer systems will convert DMS and subsystem signals to coded
light signals for transmission over the data bus network. multiplexer systems will be tripley redundant [4]. The data bus and

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Chapter 11

Communication System

The communication system provides direct voice and data links between the OTV and the space station, ground control, and EVA astronauts. The communications system onboard will consist of S-Band, Ku-Band, and UHF radio frequency links. The S-Band phase modulation system will be used to transmit The (two and receive voice and data to and from the space station and ground control. Data Relay Satellites (TDRS). The S-Band system will be redundant

system can either be used in direct link mode or relayed through the Tracking independent systems) as it is the most versitle and important communications link. The Ku-Band system (same device as rendezvous radar) will be used to transfer data at rates much higher than the S-Band system. The Ku-Band system can only The relay data through TDRS and is not operational during aerobraking (antenna will be stowed) or when being used as rendezvous radar.

UHF system will be

used for voice communication between the OTV and EVA astronauts and during docking procedures with the space station [3]. The entire communications system will be interfaced with the Data Management System to control reception, transmition, command execution and data telemetry. The antennas for the S-Band and UHF radios will be flush mounted on the structure of the OTV. Four sets of redundant S-Band antennas, spaced at 90 degree intervals around the EVA module, will provide complete transmission and reception coverage with the space station and ground control either directly or through TDRS. Three UHF strip antennas, one near the docking berth, one inside module, and one inside the EVA capsule, will provide the command

e

communications with and between astronauts before and during EVA and with the space station during docking. In addition, small headset radios can be used inside the command module to allow all of the astronauts to communicate with each other as well as be linked into the entire comm net. The Ku-Band intergrated radar and communications system antenna is a
86

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deployable %foot parabolic dish [l]. It will be stowed down along the side of the

EVA module during the aerobraking maneuvers to protect it from drag forces and
aerodynamic heating.

87

Chapter 12

Satellite Repair and Recovery System
f

The satellite repair and recovery system is designed to satisy the objectives of the mission

-

to attempt a repair (or refueling) of a dysfunctional geostationary This system will reduce the costs of satellite
a

satellite and, if unsuccessful, dock with the satellite and return it to the space station at LEO for further servicing. operation.

As the cost of replacing a satellite far exceeds the cost of

repair

mission, significant savings can be gained. repair missions [2].

These savings are evidenced by past

Table 12-1: Satellite Repair Missions
Satellite Estimated cost Repair Mission Comments

Palapa Solar Max

200 million

10 million

Resold for 60 million Redeployed

270 million

43 million

The satellite recovery and repair system consists of 6 items:
1. Manned Maneuvering Unit (MMU)

-

see Figure 12-1

2. Extravehicular Mobility Unit (EMU)
3. Manipulator arm
4. Grappling device 5 . Repair Tools

- see Figure 12-2

6. Docking Ring

The above items function collectively to create an integrated system for repairing or recovering the satellite. The following typical mission employing the system serves to describe the characteristics and functions of each component.

88

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Upon rendezvous of the OTV with the satellite to within 150 feet by means of RCS, an astronaut dons the EMU and enters the EVA module through the passage provided. gravity environment Although the suit is heavy (approximately 300 lbs.) the zera allows relatively easy manipulation of the EMU. The

astronaut then proceeds to put on the externally mounted MMU, which is also stored in the EVA chamber. the relatively awkward MMU. The MMU is a self propelled backpack device for maneuvering the astronaut through space as a free flier. The MMU is equipped with twenty-four futed of thrust Additionally, the MMU is It is designed gaseous nitrogen thrusters each capable of delivering of delivering 1.7 lbs. and allowing six degree of freedom maneuverability. The escape hatch to the outside is sufficiently large to allow the astronaut to move away from the OTV without complication due to

equipped with an automatic altitude hold which provide sufficient control to damp out the motion induced by the movement of the astronaut’s limbs. to be failsafe subsystems.

-

fully redundant controls in electrical, electronic and propulsion The dimensions of the MMU are approximately 50 in. deep with The 340 lb. unit fully charged with 26 lbs.

Electrical power is supplied by two batteries, each with an energy

capacity of 752 watt-hours.

high, 33.3 in. wide, 27 in. deep with arms in launch position and 48 in. arms in the extended flight position.

of propellant can function for a six hour EVA and has a range of 3000 ft.

A t full

charge, the two aluminum pressure tanks with Kevlar ovenvrap (pressurized to

3000 psia) can induce a propulsive delta v of 66 fps to the 800 lb. combination of

man, MMU and EMU. reliably in free flight [l]. 0

This device has performed flawlessly on three previous

misssions and has proven its goal to move an astronaut easily, accurately, and

The MMU configuration proceeds to the disabled satellite and matches angular velocity. At this point, several options are presented. The astronaut may attempt to repair the satellite by attaching the MMU to the satellite by a means determined by the specific satellite. Previous missions such as that to repair the Simple operations such as replacing a Solar Maximum Mission Satellite used a device known as a trunion pin adapter (see Figure 12-3) to make this attachment.

89

satellite module may be accomplished in this fashion. satellite may require more sophisticated servicing.

More likely however, the

Therefore, the astroaut will need

to prepare the satellite for returu to the OTV by means of the manipulator a r a This necessitates the use of grappling device attached to the Satellite to which the manipulator arm may secure itself. Unfortunately, there are, presently, no universal grappling devices for satellite repair. promote easy repairability. An optimum solution to this

problem would be the standardization of all future satellites (see Figure 12-4) to Then a universal “stinger” device such as that used to retrieve the Westar VI satellite (see Figure 12-5) may be connected to the satellite and the astronaut-MMU configuration could propel the satellite to within reach of the OTV’s manipulator arm. Without this optimum satellite standardization, however, a number of

alternatives arise to continue the mission. Instead of using the MMU to propel the satellite to the OTV (which can only be accomplished reliably if the mass of the satellite is sufficiently low), the MMU may be used to attach a device to to the satellite to which the manipulator arm may attach itself. greatest potential. The manipulator arm of Figure 12-6 may be used to grasp the satellite and lower it to the docking berth on the outside of the EVA module. The ability of the docking berth to be adapted to properly fit and securely hold the satellite is essential and unfortunately, subject to the same limitations of the grappling device described above. Once this problem has been overcome, however, the manipulator
As

By maneuvering the

OTV to within 15 feet of the satellite the manipulator arm may be employed to its

arm assumes another role as a “cherry picker” [5]. To this extent, the arm serves to maneuver an astronaut around the satellite for further satellite servicing. camera mounted just behind the end of the grasping arm. from various angles. The above components compromise the satellite repair and recovery system of visibility from within the OTV is limited, the manipulator arm is teleoperated by a This arm will require six degrees of freedom to successfully attach to the satellite and permit approach

90

ORIGINAL PAGE I S OF POOR QUALITY

the

WWSR orbital transfer vehicle.

As there are great variations in current
Modifications may The ability of the space station

satellite design, the non- rigidity of proposed system is obvious. be necessary aa dictated by the individual mission. the functioning of the OTV.

to stock a sufficient supply of repair and recovery system components is essential to

I

The M M U is a selfcootined backpack for propelling UI astronaut during EVA. Twenty-four fixed gaseous nitrogen thrusters. each dclivenng 1.7 Ib of thrust, allow SIX d c g r m of freedom maneuvering ability. .-

Figure 12-1: The Manned-Manuevering Unit

a

91

I

1

!

i

I

!
I

i

I

I

I
i
!

Figure 12-2: The Extravehicular Mobility Unit

Figure 12-3: Trunion Pin Attachment Device

92

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I 4 I 1 1

I
!
I J 1 i I

,
II

'.
!
I

I
0

1
\ \
\

I I

J
I I

Figure 12-4: Serviceable Satellite Configuration

ORIGINAL PAGE I S OF POOR QUALITY

93

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I I

I
Figure 12-5: "Stinger" Device in Use
ORCGlNAL PAGE I S OF POOR QUALITY

I
I

/

n
~

I
I I
LCAMERA

Figure 12-6: Satellite Grasping Arm

I
94

Chapter 13
Cost Analysis

.
a

The contract on any new type of design is inevitably determined, in part, as a function of its cost. Although the

RFP for this project never specified

requirement to consider cost efficiency, the WWSR project team sought to integrate the most cost effective, usable components whenever possible.

To this extent, many
technology.

0

I

tried and tested devices are included in the design to avoid the comparatively large costs associated with research and development of stateof-the-art Nevertheless, responsible for a large portion of the overall cost. these costs could not be eliminated in all instances and were The cost of the computer

hardware and software necessary to successfully complete the complex aerobraking maneuver, for example, compromised nearly 13% of the overall cost of the OTV. However, the seemingly large expenditure on these computer systems is justified by the argument that the proposed aerobrake configuration will produce a dollar savings of over fifty percent as compared to existing orbital transfer vehicle concepts using all-propulsive methods of transfer between approximately ten “typical” missions, this savings will

LEO and GEO.
compensate for

In the

undeniably large research, development and systems testing costs which necessarily accompany the installment of any new technology. The approximate costs of the majority of the systems, structures, components are provided on the following page. the current figure. and

Wherever possible, the costs of

previously used components were researched and economically scaled to determine In some instances, such as the determination of the computer software cost and the aerobrake research and development cost, some fundamental concepts of engineering cost estimation and analysis were employed to determine a numerical figure. The cost breakdown on the following page does not include the cost of the fuel itself or the cost of transporting the OTV or the fuel which it requires to the

95

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space station. One should not assume, however, that these costs are negligible. In Based on the overall ''dry". fact, current NASA figures indicate that the cost of transporting a m s aboard the as shuttle to low earth orbit is approximately $1500/lb. weight of the OTV, we can expect an additional expense of nearly $40 million dollars just to transport the OTV in pieces to its berth at the space station. expense of approximately $70 million. the space shuttle. The results of the cost analysis are open to a variety of interpretations. million including transportation costs). The

In

order to fully fuel the OTV for the worst possible case would require yet another This figure is based on the contention that the fuel is put in orbit by a mote cost effective means than in the cargo bay of

final cost of development of a single 0 1 V was determined to be $850 million ($970

In light of the fact that the modern version

of the shuttle costs approximately $1 billion and space station cost projections

waver around $9 billion, we can conclude that the cost of this project is, by no

1

means, insignificant. abandoned. mission

Nor can any realistic cost decrement for the specified design This is not to say that such a design project should be However, some alterations of the design and/or appropriate. To this extent, the man-rated

parameters be expected.

The 3-man crew capability provides great opportunity for the repair of

malfunctioning or dead satellites. specifications are very

functioning, coupled with the capability of the OTV to deliver and/or recover a
30,000 lbm object from

LEO and GEO impose significant weight additions to the
to modify the mission

mission which, consequently, boost both mission and design costs tremendously. Therefore, it is the recommendation of WWSR, Inc. requirements. function of the design. The manned OTV will be of great value to the satellite repair However, when considering the satellite deploy and recovery

function of the design, consideration of other options such as an unmanned OMV may prove to be more cost effective.

96

Table 13-1 Numerical Breakdown of Project Orion Costs
Item Cost (In Millions o Dollars) f

Aerobrake

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

.150
20

Fuel Tanks . . . 6 liquid 0 , 6 liquid H , Avionics Hardware Software

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

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

20 .130 50 15

Pratt & Whitney Engines (2)

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

Power Generator . . . . . . . . . . . . 2 United Technology Fuel Cells Battery

EVA Module (with docking mechanism) Reaction Control System (RCS)

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

25 5

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

Satellite Recovery System . . . . . . . . . . . . Manipulator Arm & Grappling Device MMU & EMU Berthing Device Tools Main Cabin Structure and Components Pressurization and Temperature Control System

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

5

. . . . . . . . . . . . . . . . . . . .200
45

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

Program Development and Management Research, Development, and System Testing . . . . Unaccounted Incidentals Summation of Costs

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

75

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

.250

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

75

...............................
97

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Chapter 14
Managing Project Orion

This section is a brief discussion of how Project Orion should be managed. Managing Project Orion will be a joint effort between WWSR and NASA.
as well as monitoring the work of these corporations.

WWSR

will be responsible for establishing the contractors and subcontractors of the project NASA in turn will monitor WWSR as well as manage the deployment of the OTV. funding of our corporation.

In monitoring WWSR,

NASA’s responsibilities will consist of approving the decisions, selections, and NASA will have the power to override decisions made responsibilities with respect to its contractors and WWSR will be responsible for by WWSR. WWSR’s

c

subcontractors will be similiar to that of NASA’s.

the distribution of funding from NASA to the contractors as well as approving major decisions and designs developed by the contrators. similiar fashion. WWSR itself is a relatively new corporation in the space market.
We are,

It is expected that the

relationships between contractors and their subcontractors will be managed in a

however, one of the oldest airframe manufacturers in the country and have enjoyed a very successful partnership with the government in ensuring the defense of this country. Fifteen years ago, WWSR went through a major restructuring to assure
It was decided then that

W W S R would continue its work on development of civilian and military aircraft as well as devote
viability into the twenty-first century.

a substantial amount of capital into research and development of space systems an area we felt confident would provide us with many exciting and challenging projects. Our goal was to be prepared to make a bid on a major space contract WWSR then began to merge and aquire several firms active in WWSR Inc. is now divided into six fairly automonous in ten years.

e

developing space systems.

“companies” : W WSR Aircraft, Sunnex Controls, Airprop Engines, Vitel Electronics, WWSR Space Systems (Spacsys), WWSR Space Analysis Division (Spacad). The development of the OTV in this proposal was primarily the responsibility of Spacad with appropriate input from Spacsys, Sunnex, and Vitel.
98

Spacad will be

the company responsible for monitoring the work on Project Orion.

Spacsys will Sunnex,

manufacture the crew module, EVAM, propellant tanks, and support structure. This will be done at our recently converted airframe facilities in California. control systems for the OTV. will be responsible for developing and manufacturing the electronic and mechanical Vitel will manufacture most of the electonic Some of the systems and components All the components needed by the other companies.

of the OTV that will be contracted out will be the main engines, thermal controls,

B
0

communication systems, RCS engines, and the aerobrake and thermal tiles. components not directly manufactured by Spacsys will be installed at company’s plant. Figure 14-1 illustrates Project

Orion’s manufacturing and

management structure.

I

Once the completed system is delivered to NASA, WWSR’s responsibilities will be to provide replacement components for the OTV and to consult NASA through Spacad in mission planning. for managing the Space Station. Systems Operations. It is Spacad’s opinion that NASA should employ the same system of management for Project Orion that it proposes to use Assuming NASA uses the management system proposed by Granville Paules [2], Project Orion will be a subsystem of Space Space Systems Operations controls space system activities The subsystem User Operations Each of User concerning the Space Station that occur in orbit or on the ground. which will monitor Project Orion will consist of six divisions:

Support, Mission Planning, Predeployment/Postdeployment Operations, Integrated Logistics Support, Market Research, and Cost and Financial Managment. these divisions will consist of members from NASA, WWSR, and users. the allocation of the OTV. deployment
of

Operations Support will be responsible for assisting users in planning and directing Mission Planning will create the optimal strategy for set up by User Operations. Predeployment/Postmissions

deployment Operations will manage the functions of final servicing, integration, and processing of subsystems just before and after the OTV leaves and returns to the Space Station. Integrated Logistics Support will delegate the logistic requirements Market Research will serve as a catalyst for developing new Cost and Financial Mangament will of the various users.

areas in which the OTV can be employed. promote cost-effective operations.

99

PROJECT

onion

r3NASAq
I

management Consulting Funding

- ----

Finished Product - - Figure 14-1: Management and Manufacturing Structure

100

Chapter 15

Mission Planning

The purpose of this section is to present three scenarios for possible missions for WWSR's OTV. The description of the missions include mission objective, OTV configuration and weight estimates, fuel requirements, time of various actions, and delta v's and fuel consumed for various manuevers. Mission A: Worst Case Scenario Mission Objectives: The OTV will leave the Space Station carrying components for constructing a platform at GEO (payload, 24000 lbm). The OTV will also carry provisions for a full crew of 3 for a 14 day mission. Eight days on station will be anticipated for construction of the platform. Upon completion of construction, the OTV will returned unloaded to the Space Station. Configuration: 6 pairs of propellant tanks, 2 MMUs, 3 crew. Weight Estimates: System Weight (lbm)

ECLSS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3560 Tanks and Supporting Structure (6 pairs) . . . . . . . . . . . . . . . . . . . . 3660 Engine System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1050 Crew Cabin, EVA, and Components . . . . . . . . . . . . . . . . . . . . . . .13260
Aerobrake

Electronics EPS . . . RCS . . . MMU (2) Crew (3)

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .985 ....................................... 2215 ....................................... 3350 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1280
........................................
510

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

2800

Total (Dry) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .32670 Payload (Out) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .24000 Payload (Return) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0 Total Propellant . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 125000 Mission Profile: The mission profile for the delivery and setup of a 24,000 lbs space platform to geosynchronous orbit is shown in Table 15-1. Following separation from the Space Station and subsequent systems checkout, the OTV performs a phasing orbit injection burn (PIB). The phasing orbit is designed to bring the 101

J
OTV to the transfer orbit injection point at the proper time so it will arrive at the correct location in GEO. The transfer injection burn places the OTV in a Hohmann elIiptica1 transfer to GEO, which lasts approximately five hours. Following circularization at GEO, the OTV can remain on station for eight days to deploy (i.e. possibly construct) the space platform. After deployment is completed, an injection burn places the OTV in a GEO-LEO transfer orbit that will take it through the Earth’s atmosphere. The first aerobraking pass, dipping the OTV to a height of 85 kilometers above the Earth, lasts only five minutes and leaves the vehicle in an intermediate orbit. Based on the results of the first pass, correction burns take the OTV through the atmosphere a second time. This time the maneuver lasts about 11 minutes and places the OTV in an orbit that can be circularized at LEO by a small propulsive bum. Note that the main fuel tanks are not full to capacity and that there is still fuel in reserve. This indicates that the OTV could carry even heavier payloads than 24,000 lbm.

0

I

Table 15-1
Profile of Mission A: GEO Delivery of 24,000 Ibm Payload
Event Separate Phase Injection Coast Transfer Burn Coast & Correct GEO Circularization
Trim

Duration (hrs)

AV (m/s)
3 1400 5 1006 10 1826 5
10

Prop. (lbml

4.O 02 . 3.0 0.1 5.O 0.1
12.0 160 9.

251 44793 315 25112 512 34168 176

(RCS) (RCS) (RCS) (RCS)

0

Deliver Payload Phase Transfer Burn Coast & Correct Aerobrake Manuever Coast Aerobrake Manuever LEO Circularization Rendezvous & Dock Launch Mass: Return Mass: Total Elapsed Mission Time: Total H,-0, Prop. Used: Total RCS Fuel Used:

10.0 01 . 5.0 0.1 3.2 0.2 01 . 60 .
181,270 lbm 33,798 lbm 240 hrs 121,616 lbm 2,375 lbm

-

240 (RCS)

-

1845 10 10 5 10 200 20

16155 164 164 81 162 1392 310

(RCS) (RCS) (RCS) (RCS) (RCS)

(3384 lbm reserve) (525 Ibm reserve)

102

Mission

B: Satellite Repair

Mission Objectives: The OTV will leave the Space Station travelling to GEO and carrying no payload. The OTV will also carry provisions for a crew of 2 fdi a 6 day mission. At GEO, the crew will service two satellites. It will be anticipated that servicing will take one day for each satellite. Upon completing service of the first satellite, the OTV will make a epoch change of 30° to rendezvous with the second satellite. Upon completing service of the second satellite, the OTV will returned unloaded to the Space Station.
Configuration: 4 pairs of propellant tanks, 2 MMUs, 2 crew. Weight Estimates: System Weight (lbm)

2335 ECLSS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tanks and Supporting Structure (4 pairs) . . . . . . . . . . . . . . . . . . . . 2440 Engine System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1050 Crew Cabin, EVA, and Components . . . . . . . . . . . . . . . . . . . . . . .13260 Aerobrake . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2800 Electronics . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .985 EPS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1615 RCS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3250 MMU (2) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1280 C r e w ( 2 ) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 340
Total (Dry) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .29445 Payload (Out) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0 Payload (Return) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0
Total Propellant

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

.88000

Mission Profile: The profile for a mission to service two geosynchronous satellites is shown in Table 15-2. After the OTV is fitted with three fueled tanks, it separates from the space station and preforms full systems checks. The OTV then uses the same sequence of phase injection and transfer orbit injection burns to arrive at the proper location in GEO as detailed for Mission A. At GEO the RCS engines are used to maneuver the OTV to retrieve the first satellite. Depending on its configuration, the satellite may be recovered using either the robot arm or with the assistance of an astronaut in a Manned Maneuvering Unit (MMU). The satellite is berthed to the OTV where the EVA astronauts can effect repairs. The robot arm is particularly useful for moving an astronaut around the satellite, providing a mobile work platform. After repairs are completed, the satellite can be deployed and fully tested to assure proper operation before the OTV moves to the next satellite.

103

To change its placement in GEO, the OTV performs an epoch change burn which
places the vehicle in an orbit slightly smaller (and more elliptic) than GEO. This brings the OTV back to GEO, the epoch change orbit’s apogee, in 21.6 hours (lcss than the 24 hour period of GEO. By recircularizing, this effectively moves t h s OTV forward by about 30° in GEO. The same procedure outlined above is used to recover and repair the second satellite. After completing the second round of repairs, the OTV will perform a transfer orbit injection burn which will take it through the Earth’s atmosphere twice and return it to LEO.

I
Event Separate Phase Injection coast Transfer Burn Coast & Correct GEO Circularization Rendezvous Repair Unload Payload Epoch Change Burn coast GEO Circularization

Table 15-2 Mission B Profile: GEO Servicing of Two Satellites Separated by 30°
Duration (hrs)
4.O 0.2 30 . 01 . 50 . 0.1 6.0 24.0 30 . 01 . 2. 16 01 . 60 . 24.0 30 . 01 . 50 . 0.1 3.2 02 . 0.1 6.0 117,455 Ibm 28,300 lbm 115 hrs 85,907 lbm 2,846 lbm

AV (m/s)
3 1400 5 1006 10 1826 25

Prop. llbm)
163 28925 203 16216 331 22063 566

(RCS) (RCS) (RCS) (RCS)

-

-

I

Rendezvous

5 200 5 200 25

Repair Unload Payload Transfer Burn Coast & Correct Aerobrake Manuever Coast Aerobrake Manuever LEO Circularization Rendezvous & Dock Launch Mass: Return Mass: Total Elapsed Mission Time: Total H,-0, Prop. Used: Total RCS Fuel Used:

-

112 (RCS) 1929 107 (RCS) 1848 513 (RCS)

-

5 1845 10 10 5 10 200 20

101 13740 140 139 69 138 1184 263

(RCS) (RCS) (RCS) (RCS) (RCS) (RCS)

(2093 lbm reserve) (54 lbm reserve)

104

Mission C: 15,000 lbm Payload Up and Back Mission Ob-jectives: This mission is used to compare the performance of WWSR’s OTV to that of MOVERS’. Essentially, the mission consists of carrying a paylo& of 15,000 lbm from the Space Station to GEO and back. This payload might be some sort of experiment assembly used for SDI testing. The OTV will carry a crew of 3 for a total mission time of 7 days. Configuration: 6 pairs of propellant tanks, 2 MMUs, 3 crew. Weight Estimates: System Weight (Ibm1

ECLSS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2500 Tanks and Supporting Structure (6 pairs) . . . . . . . . . . . . . . . . . . . . 3660 Engine System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1050 Crew Cabin, EVA, and Components . . . . . . . . . . . . . . . . . . . . . . .13260 Aerobrake . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2800 Electronics . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .985 EPS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1730 RCS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3350 MMU (2) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1280 Crew (3)‘ . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ..510 Total (Dry) . . . Payload (Out) . Payload (Return) Total Propellant

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .31125 .................................. .15000 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .15000 .................................. 132000

Mission Profile: The mission profile for the delivery to GEO and return to the Space Station of a 15,000 Ibs payload is shown in Table 15-3. Following separation from the Space Station and subsequent systems checkout, the OTV The transfer injection burn places the OTV in a Hohmann performs a PIB. elliptical transfer to GEO, which lasts approximately five hours. Following circularization at GEO, the OTV can remain on station for five days to perform the necessary experiments.

e

After completing the experiments, the OTV will return to the Space Station with the payload following similar procedures for returning to LEO as described in the Mission A profile.

105

Table 15-3
Mission C Profile: 15,000 lbm Payload Up and Back

1
Event Separate Phase Injection Coast Transfer Burn Coast & Correct G E 0 Circularization Trim Station Keeping Phase Transfer Burn Coast & Correct Aerobrake Manuever coast Aerobrake Manuever LEO Circularization Rendezvous & Dock Launch Mass: Return Mass: Total Elapsed Mission Time: Total H,-0, Prop. Used: Total RCS Fuel Used: Duration (hrs)
4 .O 0.2 3 .O 0.1 5.0 0.1 12.0 120.0 10.0 0.1 5.0 0.1 3.2 0.2 0.1 6.0 178125 lbm 47803 Ibm 168 hrs 127474 lbm 2844 lbm

c.V (m/s)
3 1400 5 1006 10 1826 5 10

Prop. (lbm)
247 44015 310 24675 503 33575 173 345 23210 236 235 117 233 2000 445

(RCS) (RCS) (RCS) (RCS) (RCS)

-

1845 10 10 5 10 200 20

(RCS) (RCS) (RCS) (RCS) (RCS)

(4526 lbm reserve) (56 lbm reserve)

106

Conclusion

WWSR has presented what it feels is the most optimal design for a chemical propellant, manned OTV that fulfills the previously described constraints. WWSR’s OTV is designed to be versatile and modular. component changes. Even though this is the end of our report, we feel that much more research can be done. Many more missions other than the ones described in this proposal may be possible with minor design or We especially feel confident that with a small amount of development, our OTV would be capable of performing missions to the Moon. This could include orbiting to retrieve payloads or landing on the lunar surface. Because of its modular design, WWSR’s OTV will truly be the orbital transfer vehicle for the 2lSt century.

I

107

References

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Chapter 2

1. Bragg, B.J. A Design Study for Houston, Texas: NASA, June 1985.

an Aeroassist

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Experiment.

2. Dotts, R. L. “Shuttle Orbiter Reusable Surface Insulation Thermal Performance, Entry Vehicle and Thermal Protection Systems: Space Shuttle, Solar Starprobe, Jupiter Galileo Probe.” Progress in Astronautics and Aeronautics Vol. 85, 1983.
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7. Meness, G.P. “Determining Atmospheric Density Using a Space-Launched Projectile.” Journal of Spacecraft and Rockets Vol. 23 No. 3, May-June 1986.
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94

. :

Chapter 3

1. Brown, J.R. System Requirements Impact on Advanced OTV Engine Design. AIAA Paper 85-1340,July 1985.
2. Charhut, D.E. and W.J. Ketchut. “Future Requirements and Applications for Orbital Transfer Vehicles.” General Dynamics Convair Division. 3. Foust, R.R. RL 10 Derivative Engines for the OTV. AIAA Paper 851338, July 1985.
4. MacConochie, I.O., J.J. Rehder, and

E.P. Brien. “Preliminary Design for
3,

a Space-Based Orbital Transfer Vehicle.” J. Spacecraft Vol. 17 No. November 1979.

5 . Maloney, J.W. and L.R. Pena. Maintaining and Servicing a Space-Based Orbital Transfer Vehicle ( O T V ) at the Space Station. AIAA Paper

8 6 2 332, Septern ber 1986.

6. Redd, L. Main Propulsion System Recommendations for an Advanced Orbital Transfer Vehicle. AIAA Paper 85-1336, July 1985.

Chapter 4

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-4.jia”, & .”
4. Kaplan,
t

M.H. Modern Spacecraft Dynamics?a+ -

Control. New York:

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Fracture Analysis of Surface and Through Crach in Cylindrical Pressure Vessels. NASA TN D-8325, December 1976.

5. Rao, K.T. “On the Fracture Toughness of Aluminum-Lithium Alloy 2090-TSE41 at Ambient and Cryogenic Temperatures.” Scripta Metallurgica Vol. 22 No. 1.
6.

Rehder, John J. “Multiple Pass Trajectories for an Aeroassisted Orbital Transfer Vehicle.” Thermal Design of Aeroassisted Orbital Transfer Vehicles. H.F. Nelson ed. New York: AIAA, 1984.

7. Torre, C.N. Low-Pressure/Lightweight Cryogenic Propellant Tank Design for the Space-Based Orbital Transfer Vehicle. AIAA Paper 86-0915, 1986.

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2 Incropera, F.P. and D.P. Dewitt. Fundamentals . Transfer, New York: John Wiley and Sons, 1985. ‘
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of Heat

and Maaa


I

3. Menees,

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Comparison of the Nickel-Cadmium Battery, Bipolar Nickel-Hydrogen Battery, and Regenerative Hydrogen-Oxygen Fuel Cell Energy Storage Subsystems for Low Earth Orbit.” lgth Intersociety Energy Conversion Engineering Conference, San Francisco, 19-24 August 1984.
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Workbook. Houston: NASA, 1982.
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Transfer Vehicle.” Thermal Design of Aeroastristed Vehicles. H.F. Nelson ed. New York: AIAA, 1984.
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Station and Emergency Rescue Vehicle.” Rescue and Emergency Aerospace Capabilities Team (REACT). University of Virginia, 29 April 1987.

111

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--

1. Byington, L., and D. Theis. “Data-Processing Hardware for Spacecraft Air Force Standard 1750A-ISA is New Trend.” Computer November 1986: 50-59. 2. Carney, P. “Selecting On-Board Satellite Computer Systems.” Computer April 1983: 35-41. 3. Greenberg, E. et al. “Survey of Spacecraft Memory Computer March 1985: 29-39.

Technologies.”

4. National Aeronautics and Space Administration. Data Processing System Overview Workbook. Houston: NASA, 1984. 5. Ross, Cindy. State-of-the-Art Report: Data Management Systems. Dept. of Mechanical and Aerospace Engineering, University of Virginia, 1986.
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Survey.”

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“Orbiter Ku-Band Integrated Radar and 1. Cager, Ralph H. et al. Communications Subsystem.” IEEE Transactions on Communications COM- 26 (1978): 1604-1619.
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1. Ballendock, Walter W., et al. “MMU Technology.” Aerospace America Vol. 23, May 1985: 56-62. 2. Cowen, Robert C. “Doctoring Satellites: A Success Story.” Technology Review Vol. 87, July 1984: 4-5.

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4. Minor, Ronald. Design of a Mechanical System for Orbital Recovery of a Satellite. University of Virginia, May 1988.



7

5. “New Cherry Picker Set for Orbit.” Science Digest Vol. 91, April 1983: 22-23.

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1 3

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and

Thomas Reister

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Chapter 14

1. Goldman, Nathan C. Space Commerce: Free Enterprise o n the High Frontier. Cambridge, MA: Ballinger Publishing, 1985. 2. Paules, Granvill. “Space Station Overall Management Approach for Operations.” AIAA Space Station in the 2lSt Century Conference, Reno, NV, 3-5 September 1986.

3. Steiner, George A. and William G. Ryan. Industrial Project Management. Toronto: Macmillian, 1968.
4. Stekler, Herman 0. The Structure and Performance of the Aerospace

Industry. Los Angeles: University of California Press, 1965.

Appendix 2

1. Bate, R., D. Mueller, and J. White. Fundamentals of Astrodynamics. New York: Dover Publications, 1971.
2. Future

Orbital Transfer Vehicle Technology Study Vol. Contractor Report 3536, Boeing Aerospace Company.

2.

NASA

113

Appendix 3

1. General Dynamics. "Definition of Technology Development Missions for

Early Space Station Orbit Transfer Vehicle Servicing." NASA-CR-170863, 1983. 2. Martin Marietta Aerospace. "Definition of Technology Development Mission for Early Space Station Satellite Servicing." NASA-CR-171229, 1984. 3. Natqatomo, N. "Orbital Operation of Co-orbiting Space Station." AIAA TIS 3/13. Spacecraft with a

114

Appendix 1 System and Subsystem Weight and Power Requirement Estimates

The following pages are tables of our estimates for the weights and power requirements of various subsystems. scenario. These estimates were based on our worst case The For missions other than worst case, our weights may be lower.

total weights of various subsystems are as follows:

Table Al-1 Total System Weights (Worst Case)

ECLSS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Main Fuel Tanks and Supporting Structure (6 pairs) . . . . . . . . . . . . . . . . . . . . . . . . . . Aerobrake . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Command/EVA Module and Components . . . . . . . . . . . . . . . Electronics . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . EPS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . RCS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Engine System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Crew (3) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . MMU (2) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Robot Arm . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Flight Chairs (3) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Total (Dry)

3560 lbm 3660 2800 12680 .985 2215 3250 1050 .510 1280 .330 .250 32670

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

lbm lbm lbm lbm lbm lbm lbrn lbm Ibm lbm lbm lbm

Propellant (Total) Total (Fueled) . . Payload (Max.) Grand Total .

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

.1250o0 Ibm .157670 lbm
24000 Ibm .181670 lbm

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

115

Table A1-2
Mass and Power Analysis of OTV ECLSS

System

Weight (lbm)

Power (Watts)
900 900

Air Revitalization System Thermal Control

. . . . . . . . . . . . . . . . . 650
880

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

Crew Systems (Worst Case) 0, (metabolic - 2.25 lbm/man-day)

. . . . . . . . . . 96 N, . . . . . . . . . . . . . . . . . . . . . . . . . . . . 193
H,O (drinking - 8.0 lbm/man-day) . . . . . . . . . 336 H,O (hygiene . 15.0 lbm/man-day) . . . . . . . . . 630
Food (2.50 lbm/man-day) . Waste (1.00 lbm/man-day)

-

-

. . . . . . . . . . . . . . 105 . . . . . . . . . . . . . . .42

700

Other Components Freezer and Microwave : . . . . . . . . . . . . . . . . 6 0 LiOH/contaminant removal cannisters . . . . . . . 220 Sanitation and Hygiene . . . . . . . . . . . . . . . . 200 Galley . . . . . . . . . . . . . . . . . . . . . . . . . . 150 Totals

-

2600

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

.3560

116

.
Table A1-3 Electrical. Avionics. and Communications
Subsystem Weights and Power Consumption

Subsystem

Weight (lbm)

Power (Watts)

GNC

GPS Receivers (2) . . . . . . . . . . . . . . . . . . Stellar Tracker . . . . . . . . . . . . . . . . . . . . . . IMU (2) . . . . . . . . . . . . . . . . . . . . . . . . . . Ku-Band Radar . . . . . . . . . . . . . . . . . . . . Total GNC . . . . . . . . . . . . . . . . . . . . . . .

. .40
40 40

60 20 320

NA
160

NA
400

DMS Computers (3) . . . . . . . . . . . . . . . . . . . . . . 63 Mass Memory (2) . . . . . . . . . . . . . . . . . . . . 31 Displays (4) . . . . . . . . . . . . . . . . . . . . . . . . 16 Keyboards (3) . . . . . . . . . . . . . . . . . . . . . . 15 Data Bus Network . . . . . . . . . . . . . . . . . . . 100 Instrumentation . . . . . . . . . . . . . . . . . . . . 100 Total DMS . . . . . . . . . . . . . . . . . . . . . . . 325 Communications S-Band PM Radio (2) . . . . . . . . . . . . . . . . . 200 UHF Radio . . . . . . . . . . . . . . . . . . . . . . . . 4 0 Ku-Band Radio/Radar . . . . . . . . . . . . . . . . 260 Total Communications . . . . . . . . . . . . . . . . 500

300 20 80 10 20 50 490

700 25 590 1315

EPS

Fuel Cells (2) . . . . . . . . . . . . . . . . . . . Ni-H battery . . . . . . . . . . . . . . . . . . . . EPDS (2) . . . . . . . . . . . . . . . . . . . . . . . . Total Reactants . . . . . . . . . . . . . . . . . . . . Total EPS . . . . . . . . . . . . . . . . . . . . . . .

. . 350 . . 165
100 1600 2215

NA NA
200

NA
200

0

RCS Reaction Control System RCS fuel . . . . . . . . . Total RCS . . . . . . . . Grand Total

. . . . . . . . . . . . . . . 450 . . . . . . . . . . . . . . . 2900 . . . . . . . . . . . . . . . 3250
7710

300

NA
300 2705

1
I

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

117

Table A1-4 Structural Component Weight Estimate

Structural Component

Weight (lbm)

Engine Quick Disconnect Plate . . . . . . . . . . . . . . . . . . . . . . . . . -100 Thrust Structure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 210 Connectors (6 sets) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 240 LO, Tanks (6) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 600 LH, Tanks (6) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .1500 Tank Support Rings and Support Struts (6 sets) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 220 Command Module . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10700 EVA Module . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . : 1500 Aerobrake . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2800 Hatches (3) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 300 Docking/Service . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 180 Total

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 18850

e

118

Appendix 2 Orbital Mechanics

Great emphasis was spent on determining the proper trajectory of the OTV. The biggest problem was to determine a successful rendezvous and intercept trajectory with a target satellite in geosynchronous orbit (GEO). one launch window was necessary. Potential launch windows were investigated, but after careful analysis, it was discovered that only To achieve this result, a Phase Injection Burn (PIB) was proposed [I]. Below is a thorough explanation of this maneuver. PIB is used to phase the OTV with the GEO satellite so that an

approximation to the Hohmann transfer from LEO to GEO can be executed. Consider FIGURE A2-1. What is done is that the OTV's time of flight from LEO Then, this value is used to ascertain the angular Thus, in order for a to GEO is first determined.

displacement of the GEO satellite from the intercept point.

successful rendezvous to occur at GEO, the target satellite must be at an angle of
79.2" from the line of nodes at the commencement of the OTV Hohmann transfer.

Note that the OTV transfer can only be initiated at the line of nodes intersecting GEO and LEO. By the time that the OTV reaches the intercept point, the displacement. satellite will have traveled the 79.2

The main problem is that the position of the OTV (point A) and the position of the satellite (point B) at the beginning of LEO-GEO transfer rarely occurs, if at all. PIB rectifies this situation.

What happens is that the OTV is launched into a round trip elliptical path from A when the satellite is found at any point on arc PC. The time of flight of the OTV's PIB corresponds to the time it wll take the satellite to reach point B. Thus, by the time the OTV returns to its original location (point A), both spacecraft are perfectly phased for intercept through the Hohmann transfer.

119

ORIGINAL PAGE IS OF POOR QUALITY

L

GEO

LI

Figure A2-1: Trajectory Schematic

120

Under no circumstances is PIB to be executed if the satellite is on arc CB. This is because the corresponding PIB will take the OTV inside LEO where it will encounter significant drag . Of course, it is possible to perform a PIB if the satellite is found on the arc

ZC.

However, such a launch would mean a waste of propellent because one

complete revolution at LEO corresponds to 22.9' of GEO displacment. passed point A about 3 times.

In other

words, by the time the satellite reaches C from 2, the Space Station will have Using the same argument, by the time the satellite reaches P from Z, the Space Station will have passed point A about two times. Arc PC was chosen for the delta V analysis because this displacement corresponds to one complete revolution on LEO. launch opportunity for PIB. other end of the nodal line. In essence, it is the optimum PIB can also be executed for an intercept at the Thus, 2 rendezvous intercept opportunities are

guaranteed within a 24 hour period from the Space Station. Since LEO is inclined 28.5 relative to GEO, it was found that a

simultaneous plane change and circularization maneuver at GEO involved the least delta V. Furthermore, the aerobraking maneuver was thoroughly investigated. Details of this maneuver are discussed in Chapter 2.
Table A2-1 is a chart outlining the required delta V's which the OTV will

need to execute for a typical mission. assist trajectory is not included.

See Figure A2-2.

For simplicity, the aero-

121

J
I I I
Table A2-I Summary of Delta V’s
AV(km/s)
1.5166 0.8893

LOCATION
1
2

REASON

PIB
Injection transfer from LEO to Hohmann transfer ellipse Circularization and plane change at GEO Plane change for LEO return and to shorten perigee height for aerobraking Aerobraking at 80 km altitude maximum. 2 passes through Earth atmosphere. (Free velocity ’ decrement of 2.250 km/s)

1.8258

3

1.8437

4

0.0000

5

0.4513

6

LEO circularization

TOTAL

AV’s = 6.5268 km/s

It is important to note that the sum of t..e

delta V’s at locations 1 and 2

(2.4059 km/s) is invariant. This means that no matter what the PIB and the

transfer injection delta V’s are, their sum will always equal 2.4059 km/s. Also note that the total propulsive delta V’s to GEO is approximately equal to the ones needed to return to LEO (aerobraking velocity increments included).

122

/
I

'1
\

I

I i
I

i

0

Figure A2-2: Location of OTV's Delta V's

ORIGINAL PAGE I S
OF POOR QUALtlY
123

Appendix 3

OTV Servicing Aboard the Space Station

8 1

When the Space Station becomes operational sometime in the mid-l990s, there will be a need to service OTVs. the time between missions. WWSR’s OTV will reqiire some servicing during This repair and refurbishment will take place in a This does not mean the repair area

special area aboard the Space Station. This area will need to be separated from the main portion of the station by some distance. areas. The servicing area will mainly consist of a large hangar. This hangar will consist of several distinct areas. These areas will be: fuel depot, engine bay, fuel tank storage, cargo handling, avionics repair, heat shield repair, command module repair and refurbishing, and ship integration area. The fuel depot will consist of several large cryogenic storage tanks for the
LH2 and L02. These fuels will be stored in insulated thermos-like tanks that will

will be free-flying, only out on a boom away from the living quarters, and scientific

have to be small enough to carry up in the shuttle cargo bay. being punctured by meteorites.

These tanks will

have to be protected from the rays of the sun, as well as have protection from The protection from the Sun will consist of The moveable shades that will move as the direction of the Sun changes. will stop all but the largest meteorites.

protection of the tanks from puncture will consist of a honeycomb structure that If a tank happens to get punctured by a large object, it will have to be jetisoned immediately so that the escaping gases do not create a force imbalance on the Space Station. The engine bay of the servicing area will be the place where spare engines for the OTV are stored and repaired. The OTV engines will be modularized, so they will only need to be snapped on and off between missions. The engines may be taken out after each sortie to make sure that no malfunctions happen during a mission. In the event that the engine cannot be repaired in space, it will have to

124

J
be brought down to Earth on the Space Shuttle. Part of the engine bay will consist of the storage empty spare fuel modules for the OTV. For the moving of these tanks, as well as the engines and other large portions of the OTV, the hangar will have a large servicing crane. assembled OTV. The cargo handling area of the hangar will consist of a place to store the satellites before they are loaded on the OTV for transfer to GEO. The satellites may have to wait long periods of time before they can get a flight out to GEO. This means that the satellites are able to be checked out and serviced in this waiting area. The cargo area will also need a means of transferring the retrieved satellites from the OTV area to the satellite repair area. The repair area of the hangar is probably the most important. The station must be able to repair all but the most severe malfunctions without having to send portions of the OTV back to Earth. This will mean that there must be astronauts on the station that are knowledgeable in all areas of the OTV, and that the repair area will be equipped well enough for repair of all major portions of the OTV including: avionics, life support, reaction control engines, fuel handling, cargo handling, and heat shield. Finally, the ship integration area is where the whole OTV will be assembled. This area will need to be large enough to contain an entire assembled OTV. The integration area will need to have several cranes, as well as robot arms for the astronauts to stand on while putting the OTV together. This is also where the This crane will be on a track that will run down the length of the hangar, and will have enough power to move the whole

OTV will be stored between missions.

The reason for storing the OTV inside the

hangar is to protect it from damaging radiation, micrometeorites, and random debris that will be floating around the Space Station. The cost of this repair satation has yet to be determined because the area has yet to be fully designed. The current estimates are that the area will cost about $700 million. This does not include the cost of sending up parts on the

125

shuttle.

This repair .-angar and all attached areas will take about three shuttle

flights to lift to

LEO. At current costs, this means another $300 million to the

price for a grand total of $1.0 billion. This price is only preliminary and will n9 doubt increase as production of the pieces moves ahead.

I
a

126

Appendix 4
Section Authors

Michael Doheny

-

Chapter Chapter Chapter

3 4 6

Richard Franck

-

Chapter 2 Chapter 5 Chapter 8 Chapter 2 Chapter 7 Chapter 9 Chapter 10 Chapter 15

Steven Hollo

-

Kenneth Ibarra

-

Chapter 1 Chapter 8 Appendix 2 Chapter 12 Chapter 13 Chapter 1 Appendix 3

William Nosal

-

Thomas Redd

Gregory Weigand (Editor)

Foreword Introduction Chapter 1 Chapter 14 Chapter 15 Conclusion

127



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