CHAPTER TWO LAUNCH SYSTEMS - NASA · 2002-01-17 · Introduction Launch systems provide access to space, obviously a necessary com-ponent of all spaceflights. The elements of launch

CHAPTER TWO

LAUNCH SYSTEMS

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Introduction

Launch systems provide access to space, obviously a necessary com-ponent of all spaceflights. The elements of launch systems include thevarious vehicles, engines, boosters, and other propulsive and launchdevices that help propel a spacecraft into space and position it properly.From 1979 through 1988, NASA used both expendable launch vehicles(ELVs)—those that can be used only once—and reusable launch vehicles.This chapter addresses both types of vehicles, as well as other launch sys-tem-related elements.

NASA used three families of ELVs (Scout, Delta, and Atlas) and onereusable launch vehicle (Space Shuttle) from 1979 through 1988 (Figure2–1). Each family of ELVs had several models, which are described inthis chapter. For the Space Shuttle, or Space Transportation System(STS), the solid rocket booster, external tank, and main engine elementscomprised the launch-related elements and are addressed. The orbitalmaneuvering vehicle and the various types of upper stages that boostedsatellites into their desired orbit are also described.

This chapter includes an overview of the management of NASA’slaunch vehicle program and summarizes the agency’s launch vehicle bud-get. In addition, this chapter addresses other launch vehicle development,such as certain elements of advanced programs.

Several trends that began earlier in NASA’s history continued in thisdecade (1979–1988). The trend toward acquiring launch vehicles and ser-vices from the commercial sector continued, as did the use of NASA-launched vehicles for commercial payloads. President Reagan’s policydirective of May 1983 reiterated U.S. government support for commercialELV activities and the resulting shift toward commercialization of ELVactivities. His directive stated that the “U.S. government fully endorsesand will facilitate commercialization of U.S. Expendable LaunchVehicles.” His directive said that the United States would encourage useof its national ranges for commercial ELV operations and would “makeavailable, on a reimbursable basis, facilities, equipment, tooling,and services that are required to support the production and operation of

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CHAPTER TWO

LAUNCH SYSTEMS

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U.S. commercial ELVs.” Use of these facilities would be priced toencourage “viable commercial ELV launch activities.”1

The policy also stated the government’s intention of replacing ELVswith the STS as the primary launch system for most spaceflights.(Original plans called for a rate flight of up to fifty Space Shuttle flightsper year.) However, as early as FY 1984, Congress recognized that rely-ing exclusively on the Shuttle for all types of launches might not be thebest policy. Congress stated in the 1984 appropriations bill that “theSpace Shuttle system should be used primarily as a launch vehicle forgovernment defense and civil payloads only” and “commercial customersfor communications satellites and other purposes should begin to look tothe commercialization of existing expendable launch vehicles.”2 TheChallenger accident, which delayed the Space Shuttle program, also con-

NASA HISTORICAL DATA BOOK14

1Announcement of U.S. Government Support for Commercial Operations bythe Private Sector, May 16, 1983, from National Archives and Records Service’sWeekly Compilation of Presidential Documents for May 16, 1983, pp. 721–23.

2House Committee on Appropriations, Department of Housing and UrbanDevelopment-Independent Agencies Appropriation Bill, 1984, Report toAccompany H.R. 3133, 98th Cong., 1st sess., 1983, H. Rept. 98— (unnumbered).

Figure 2–1. NASA Space Transportation System (1988)

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tributed to the development of a “mixed fleet strategy,” which recom-mended using both ELVs and the Shuttle.3

Management of the Launch Vehicle Program

Two NASA program offices shared management responsibility forthe launch vehicle program: Code M (at different times called the Officeof Space Transportation, the Office of Space Transportation Acquisition,and the Office of Space Flight) and Code O (the Office of SpaceTransportation Operations). Launch system management generallyresided in two or more divisions within these offices, depending on whatlaunch system elements were involved.

The organizational charts that follow illustrate the top-level structureof Codes M and O during the period 1979–1988. As in other parts of thischapter, there is some overlap between the management-related materialpresented in this chapter and the material in Chapter 3, “SpaceTransportation and Human Spaceflight.”

Also during the period 1979 through 1988, two major reorganizationsin the launch vehicle area occurred (Figure 2–2): the split of the Office ofSpace Transportation into Codes M and O in 1979 (Phase I) and the merg-er of the two program offices into Code M in 1982 (Phase II). In addition,the adoption of the mixed fleet strategy following the loss of theChallenger reconfigured a number of divisions (Phase III). These man-agement reorganizations reflected NASA’s relative emphasis on the SpaceShuttle or on ELVs as NASA’s primary launch vehicle, as well as the tran-sition of the Shuttle from developmental to operational status.

Phase I: Split of Code M Into Space Transportation Acquisition (Code M) and Space Transportation Operations (Code O)

John F. Yardley, the original associate administrator for the Office ofSpace Transportation Systems since its establishment in 1977, continuedin that capacity, providing continuous assessment of STS development,acquisition, and operations status. In October 1979, Charles R. Gunnassumed the new position of deputy associate administrator for STS(Operations) within Code M, a position designed to provide transitionmanagement in anticipation of the formation of a new program officeplanned for later that year (Figure 2–3).

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3NASA Office of Space Flight, Mixed Fleet Study, January 12, 1987. TheNASA Advisory Council had also established a Task Force on Issues of a MixedFleet in March 1987 to study the issues associated with the employment of amixed fleet of launch vehicles and endorsed the Office of Space Flight studyresults in its Study of the Issues of a Mixed Fleet. Further references to a mixedfleet are found in remarks made by NASA Administrator James C. Fletcher onMay 15, 1987.

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The formal establishment of the new Office of Space Operations(Code O) occurred in November 1979, and Dr. Stanley I. Weiss becameits first permanent associate administrator in July 1980. Code O was theprincipal interface with all STS users and assumed responsibilities forSpace Shuttle operations and functions, including scheduling, manifest-ing, pricing, launch service agreements, Spacelab, and ELVs, except forthe development of Space Shuttle upper stages. The ELV program—Atlas, Centaur, Delta, Scout, and Atlas F—moved to Code O and wasmanaged by Joseph B. Mahon, who had played a significant role inlaunch vehicle management during NASA’s second decade.

Yardley remained associate administrator for Code M until May 1981,when L. Michael Weeks assumed associate administrator responsibilities.


- Engineering- Int. & Text- Rel. Qual. & Safety

Office of Space Transportation (Code M)John Yardley

Deputy Associate Administrator (Operations)Charles Gunn

SpacelabProgramD. Lord

Space ShuttleProgramM. Malkin

Expendable LaunchVehicle Program

J. MahonSTS Operations

C. Lee

Reliability,Quality & Safety

H. Cohen

Resource Mgmt/Administration

C.R. Hovell

AdvancedProgramsJ. Disher

- Engineering- Rel., Qual. & Safety- Sys. Operations

- Small & Med. LaunchVeh. Program• Atlas• Delta• Scout• Atlas F

- Upper Stages- STS Support Projects

- Mission Anal. & Int.- System Engr. &

Logistics- Integrated Ops.- Pricing, Launch

Agreement &Cust. Svc.

- Rel., Qual. & Safety

- Budget- STS Ops.- Spacelab Program

Budget & Control- Space Shuttle

Program Budget &Control

- ELV ProgramBudget & Control

- Adm. & Program Spt.

- Adv. Concepts- Adv. Studies- Adv. Development

Figure 2–3. Office of Space Transportation (as of October 1979)

Office of SpaceTransportation

(Code M)John Yardley

Phase ISplit of Code M, creatingnew Office of SpaceTransportation Operations(Code O) (November 1979)

Phase IIMerger of Codes M andO to create the Office ofSpace Flight (August 1982)

Phase IIIPost-Challenger1986 to return to flightSeptember 1988


Acquisition (Code M)John Yardley

James Abrahamson


Operations (Code O)Stanley Weiss

Office of SpaceFlight (Code M)

James AbrahamsonJesse MooreRichard Truly

Mixed Fleet Strategy

Figure 2–2. Top-Level Launch Vehicle Organizational Structure

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Two new divisions within Code M were established in May 1981. TheUpper Stage Division, with Frank Van Renssalaer as director, assumedresponsibility for managing the wide-body Centaur, the Inertial Upper Stage(IUS), the Solid Spinning Upper Stage (SSUS), and the Solar-ElectricPropulsion System. The Solid Rocket Booster and External Tank Division,with Jerry Fitts as director, was also created. In November 1981, MajorGeneral James A. Abrahamson, on assignment from the Air Force, assumedduties as permanent associate administrator of Code M (Figure 2–4).

LAUNCH SYSTEMS 17

Office of SpaceTransportation Systems

(Code M)John Yardley

Orbiter ProgramsM. Malkin (acting)

Ground Systems &Flight TestsE. Andrews

Reliability, Quality& SafetyH. Cohen

AdvancedProgramsJ. Disher

Engine ProgramsW. Dankhoff

(acting)

SystemsEngineering & Int.

LeRoy Day

Resource Mgmt/Administration

C.R. Hovell

ExpendableEqpt(a)

F. Van Renssalaer

- Electrical Systems- Engr. & Int.- Structural Spt.

- Flight Test- Launch & Landing

Syst.- Flight Systems

- Adv. Concepts- Adv. Development

- SystemsEngineering

- STS Integration

- Cost & ScheduleAnalysis

- Adm. & ProgramSpt.

- STS Program &Budget Control

- Solid RocketBooster

- Upper Stages- External Tank

(a) May 1981—Expendable Equipment Division disestablished.New divisions established:Upper Stages Division—Frank Van Renssalaer, Branches—Centaur, Solar Electric Propulsion Systems, IUS, and SSUSSolid Rocket Booster and External Tank Division—Jerry Fitts, Branches—Solid Rocket Booster and External Tank


Operations (Code O)Stanley Weiss

STS EffectivenessAnalysis

Expendable LaunchVehiclesJ. Mahon

Quality & SafetyH. Cohen

Operations &Systems RqmtsC. Gunn (acting)

Spacelab ProgramD. Lord

STS UtilizationC. Lee

Resource & Mgmt.Administration

W. Draper (acting)

- Atlas-Centaur- Delta- Scout- Atlas F

- Integrated Ops.- Systems Engr. &

Logistics

- Engineering- Integration & Test

- Mission Analysis &Integration

- Policy, Planning &Launch ServicesAgreements

- STS OperationsBudget

- Spacelab ProgramBudget

- ELV Program Budget- Adm. & Program Spt.- Resources

Integration

Figure 2–4. Code M/Code O Split (as of February 1980) (1 of 2)

Figure 2–4. Code M/Code O Split (as of February 1980) (2 of 2)

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Phase II: Merger of Codes M and O Into the Office of Space Flight

In preparation for Space Shuttle operations, Codes M and O mergedin 1982 into the Office of Space Flight, Code M, with Abrahamson serv-ing as associate administrator (Figure 2–5). Weiss became NASA’s chiefengineer. Code M was responsible for the fourth and final developmentalShuttle flight, the operational flights that would follow, future Shuttleprocurements, and ELVs. The new office structure included the SpecialPrograms Division (responsible for managing ELVs and upper stages),with Mahon continuing to lead that division, the Spacelab Division, theCustomer Services Division, the Space Shuttle Operations Office, and theSpace Station Task Force. This task force, under the direction of John D.Hodge, developed the programmatic aspects of a space station, includingmission analysis, requirements definition, and program management. InApril 1984, an interim Space Station Program Office superseded theSpace Station Task Force and, in August 1984, became the permanentOffice of Space Station (Code S), with Philip E. Culbertson serving asassociate administrator. In the second quarter of 1983, organizationalresponsibility for ELVs moved from the Special Programs Division to thenewly formed Space Transportation Support Division, still under the lead-ership of Joseph Mahon.

Jesse W. Moore took over as Code M associate administrator onAugust 1, 1984, replacing Abrahamson, who accepted a new assignment


Office of Space Flight (Code M)James Abrahamson

Safety, Rel., &Qual. Assurance

H. Cohen

CustomerServices

C. Lee (acting)

Space ShuttleOps. (b) (d)

L.M. Weeks (acting)Spacelab

J. Harrington

AdvancedPlanning (c)

I. Bekey (acting)

Resources &Institutions

M.J. Steel (acting)Special ProgramsJ. Mahon (acting)

Space StationTask Force (a)

J. Hodge

- STS Utilization- Systems Planning

& Effectiveness

- Orbiter Programs• Avionics & Electrical Systems• Engr. & Syst. Int.

- Engine Programs- Solid Rocket Booster

& External Tank• SRB• External Tank

- Ground Systems &Flight Test• Flight Test• Launch & Landing

- STS Systems Engr. &Int.• Systems Engr.• STS Integration

- STS Ops.

- Engineering- Int. & Test

- Adv. Concepts- Adv. Development

- Institutions & Adm.- Resources Mgmt

(Development)- Resources Mgmt

(Operations)

- ELVs• Atlas Centaur• Delta• Scout• Atlas F

- Upper Stages• Centaur• IUS• SSUS

- These divisions have an additional subsidiary organizational level also headed by Directors.

(a) The Space Station Task Force became the Office of Space Station (Code S) in August 1984.(b) In early 1983, the following changes took place in the Space Shuttle Operations Division:

- Propulsion Branch added- Flight & Turnaround Operations added- Engine Programs eliminated- SRB & external tank eliminated- STS Systems Engineering and Integration eliminated and replaced by Integration Office- STS Operations eliminated

(c) Advanced Planning Division added Advanced Transportation, Platforms and Services, and Requirements Definition; eliminated Advanced Concepts and AdvancedDevelopment.

(d) In the second quarter of 1983, organizational responsibility for ELVs moved from the Special Programs Division to the new Space Transportation Support Division,also under the leadership of Joseph Mahon.

(e) In late 1983, the Shuttle Propulsion Division was added. Within it were the Productivity Operations Support office, the Engine Program office, the Solid Rocket Programoffice, and the External Tank Program office.

(f) In early 1984, the Tether Satellite System office was added to the Space Transportation Support Division, and a Flight Demonstrations and Satellite Services and CrewServices office were added to the Advanced Programs Division.

(g) In 1986, the Orbital Maneuvering Vehicle office was added to the Space Transportation Support Division.

Figure 2–5. Code M Merger (as of October 1982)

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in the Department of Defense (DOD). Moore was succeeded by RearAdmiral Richard H. Truly, a former astronaut, on February 20, 1986.

Phase III: Post-Challenger Launch Vehicle Management

From the first Space Shuttle orbital test flight in April 1981 throughSTS 61-C on January 12, 1986, NASA flew twenty-four successful Shuttlemissions, and the agency was well on its way to establishing the Shuttle asits only launch vehicle. The loss of the Challenger (STS 51-L) on January 26, 1986, grounded the Shuttle fleet for thirty-two months. Whenflights resumed with STS-26 in September 1988, NASA planned a moreconservative launch rate of twelve launches per year. The reduction of theplanned flight rate forced many payloads to procure ELV launch servicesand forced NASA to plan to limit Shuttle use to payloads that required acrewed presence or the unique capabilities of the Shuttle. It also forcedNASA to recognize the inadvisability of relying totally on the Shuttle. Theresulting adoption of a “mixed fleet strategy” included increased NASA-DOD collaboration for the acquisition of launch vehicles and the purchaseof ELV launch services. This acquisition strategy consisted of competitiveprocurements of the vehicle, software, and engineering and logisticalwork, except for an initial transitional period through 1991, when pro-curements would be noncompetitive if it was shown that it was in the gov-ernment’s best interest to match assured launch vehicle availability withpayloads and established mission requirements.

The mixed fleet strategy was aimed at a healthy and affordable launchcapability, assured access to space, the utilization of a mixed fleet to sup-port NASA mission requirements, a dual-launch capability for criticalpayloads, an expanded national launch capability, the protection of theShuttle fleet, and the fostering of ELV commercialization. This last goalwas in accordance with the Reagan administration’s policy of encourag-ing the growth of the fledgling commercial launch business wheneverpossible. The Office of Commercial Programs (established in 1984) wasdesignated to serve as an advocate to ensure that NASA’s internal deci-sion-making process encouraged and facilitated the development of adomestic industrial base to provide access to space.

During this regrouping period, the ELV program continued to be man-aged at Headquarters within the Office of Space Flight, through the SpaceTransportation Support Division, with Joseph Mahon serving as divisiondirector and Peter Eaton as chief of ELVs, until late 1986. During this peri-od, the Tethered Satellite System and the Orbital Maneuvering Vehicle alsobecame responsibilities of this division. In late 1986, Code M reorganizedinto the basic configuration that it would keep through 1988 (Figure 2–6).This included a new management and operations structure for the NationalSpace Transportation System (NSTS). Arnold J. Aldrich was named direc-tor of the NSTS at NASA Headquarters. A new Flight Systems Division,still under the leadership of Mahon, consisted of divisions for ELVs andupper stages, as well as divisions for advanced programs and Space Shuttle

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carrier systems. The Propulsion Division was eliminated as part of theNSTS’s move to clarify the points of authority and responsibility in theShuttle program and to establish clear lines of communication in the infor-mation transfer and decision-making processes.

Money for NASA’s Launch Systems

From 1979 through 1983, all funds for NASA’s launch systems camefrom the Research and Development (R&D) appropriation. Beginning inFY 1984, Congress authorized a new appropriation, Space Flight,Control, and Data Communications (SFC&DC), to segregate funds forongoing Space Shuttle-related activities. This appropriation was inresponse to an October 1983 recommendation by the NASA AdvisoryCouncil, which stated that the operating budgets, facilities, and personnelrequired to support an operational Space Shuttle be “fenced” from the restof NASA’s programs. The council maintained that such an action wouldspeed the transition to more efficient operations, help reduce costs, andease the transfer of STS operations to the private sector or some new gov-ernment operating agency, should such a transfer be desired.4 SFC&DCwas used for Space Shuttle production and capability development, spacetransportation operations (including ELVs), and space and ground net-work communications and data systems activities.

Most data in this section came from two sources. Programmed (actu-al) figures came from the yearly budget estimates prepared by NASA’sBudget Operations Division, Office of the Comptroller. Data on NASA’ssubmissions and congressional action came from the chronological histo-ry budget submissions issued for each fiscal year.


4NASA, Fiscal Year 1985 Budget Submission, Chronological History, HouseAuthorization Committee Report, issued April 22, 1986, p. 15.

Figure 2–6. Office of Space Flight 1986 Reorganization

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Table 2–1 shows the total appropriated amounts for launch vehiclesand launch-related components. Tables 2–2 through 2–12 show therequested amount that NASA submitted to Congress, the amount autho-rized for each item or program, the final appropriation, and the pro-grammed (or actual) amounts spent for each item or program. Thesubmission represented the amount agreed to by NASA and OMB, notnecessarily the initial request NASA made to the President’s budget offi-cer. The authorized amount was the ceiling set by Congress for a particu-lar purpose. The appropriated amount reflected the amount that Congressactually allowed the Treasury to provide for specific purposes.5

As is obvious from examining the tables, funds for launch vehiclesand other launch-related components were often rolled up into the totalR&D or SFC&DC appropriation or other major budget category (“undis-tributed” funds). This made tracking the funding levels specifically des-ignated for launch systems difficult. However, supporting congressionalcommittee documentation clarified some of Congress’s intentions. In thelate 1970s and early 1980s, Congress intended that most space launcheswere to move from ELVs to the Space Shuttle as soon as the Shuttlebecame operational. This goal was being rethought by 1984, and it wasreplaced by a mixed fleet strategy after 1986. However, even though thegovernment returned to using ELVs for many missions, it never againtook prime responsibility for most launch system costs. From 1985through 1987, Congress declared that the NASA ELV program would becompletely funded on a reimbursable basis. Launch costs would be paidby the customer (for example, commercial entities, other governmentagencies, or foreign governments). Not until 1988 did Congress providedirect funding for two Delta II launch vehicles that would be used forNASA launches in the early 1990s. Although the federal governmentfunded the Shuttle to a much greater degree, it was also to be used, whenpossible, for commercial or other government missions in which the cus-tomer would pay part of the launch and payload costs.

In some fiscal years, ELVs, upper stages, Shuttle-related launch ele-ments, and advanced programs had their own budget lines in the con-gressional budget submissions. However, no element always had its ownbudget line. To follow the changes that took place, readers should consultthe notes that follow each table as well as examine the data in each table.Additional data relating to the major Space Shuttle budget categories canbe found in the budget tables in Chapter 3.

NASA’s budget structure changed from one year to the next dependingon the status of various programs and budget priorities. From 1979 through1983, all launch-related activities fell under the R&D appropriation.

LAUNCH SYSTEMS 21

5The term “appropriation” is used in two ways. It names a major budget cat-egory (for instance, R&D or SFC&DC). It is also used to designate an amountthat Congress allows an agency to spend (for example, NASA’s FY 1986 appro-priation was $7,546.7 million).

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Launch elements were found in the Space Flight Operations program, theSpace Shuttle program, and the ELV program. The Space FlightOperations program included the major categories of space transportationsystems operations capability development, space transportation systemoperations, and advanced programs (among others not relevant here).Upper stages were found in two areas: space transportation systems oper-ations capability development included space transportation system upperstages, and space transportation system operations included upper stageoperations.

The Space Shuttle program included design, development, test, andevaluation (DDT&E), which encompassed budget items for the orbiter,main engine, external tank, solid rocket booster (SRB), and launch andlanding. The DDT&E category was eliminated after FY 1982. The pro-duction category also was incorporated into the Space Shuttle program.Production included budget line items for the orbiter, main engine, andlaunch and landing.

The ELV program included budget items for the Delta, Scout,Centaur, and Atlas F. (FY 1982 was the last year that the Atlas F appearedin the budget.)

FY 1984 was a transition year. Budget submissions (which were sub-mitted to Congress as early as FY 1982) and authorizations were still partof the R&D appropriation. By the time the congressional appropriationscommittee acted, the SFC&DC appropriation was in place. Two majorcategories, Shuttle production and operational capability and space trans-portation operations, were in SFC&DC. Shuttle production and opera-tional capability contained budget items for the orbiter, launch andmission support, propulsion systems (including the main engine, solidrocket booster, external tank, and systems support), and changes and sys-tems upgrading. Space transportation operations included Shuttle opera-tions and ELVs. Shuttle operations included flight operations, flighthardware (encompassing the orbiter, solid rocket booster, and externaltank), and launch and landing. ELVs included the Delta and Scout. (FY1984 was the last year that there was a separate ELV budget category untilthe FY 1988 budget.) R&D’s Space Transportation CapabilityDevelopment program retained upper stages, advanced programs, and theTethered Satellite System.

Beginning in FY 1985, most launch-related activities moved to theSFC&DC appropriation. In 1987, NASA initiated the Expendable LaunchVehicles/Mixed Fleet program to provide launch services for selectedNASA payloads not requiring the Space Shuttle’s capabilities.

Space Shuttle Funding

Funds for the Space Shuttle Main Engine (SSME) were split into aDDT&E line item and a production line item from 1979 through 1983.Funds for the external tank and SRB were all designated as DDT&E.Beginning with FY 1984, SSME, external tank, and SRB funds were


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located in the capability development/flight hardware category and in thePropulsion System program. Capability development included continuingcapability development tasks for the orbiter, main engine, external tank,and SRB and the development of the filament wound case SRB. Congressdefined propulsion systems as systems that provided “for the productionof the SSME, the implementation of the capability to support operationalrequirements, and the anomaly resolution for the SSME, SRB, and exter-nal tank.”

Some Space Shuttle funds were located in the flight hardware budgetcategory. Flight hardware provided for the procurement of the externaltank, the manufacturing and refurbishment of SRB hardware and motors,and space components for the main engine; orbiter spares, includingexternal tank disconnects, sustaining engineering, and logistics supportfor external tank, SRB, and main engine flight hardware elements; andmaintenance and operation of flight crew equipment.

Tables 2–1 through 2–9 provide data for the launch-related elementsof the Space Shuttle and other associated items. Budget data for addi-tional Shuttle components and the major Shuttle budget categories arefound in the Chapter 3 budget tables.

Characteristics

The following sections describe the launch vehicles and launch-relatedcomponents used by NASA during the period 1979 through 1988. A chronol-ogy of each vehicle’s use and its development is also presented, as well as thecharacteristics of each launch vehicle and launch-related component.

In some cases, finding the “correct” figures for some characteristicswas difficult. The specified height, weight, or thrust of a launch vehicleoccasionally differed among NASA, contractor, and media sources.Measurements, therefore, are approximate. Height or length was mea-sured in several different ways, and sources varied on where a stage beganand ended for measuring purposes. The heights of individual stages weregenerally without any payload. However, the overall height of the assem-bled launch vehicle may include the payload. Source material did notalways indicate whether the overall length included the payload, andsometimes one mission operations report published two figures for theheight of a launch vehicle within the same report.

Thrust was also expressed in more than one way. Source materialreferred to thrust “in a vacuum,” “at sea level,” “average,” “nominal,” and“maximum.” Thrust levels vary during a launch and were sometimes pre-sented as a range of values or as a percentage of “rated thrust.”Frequently, there was no indication of which definition of thrust wasbeing used.

This chapter uses the following abbreviations for propellants: LH2 =liquid hydrogen, LOX = liquid oxygen, N2H2 = hydrazine, N2O4 = nitro-gen tetroxide, RJ-1 = liquid hydrocarbon, and RP-1 = kerosene.

LAUNCH SYSTEMS 23

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Expendable Launch Vehicles

From 1979 through 1988, NASA attempted seventy-four launcheswith a 94.6-percent success rate using the expendable Atlas E/F, Atlas-Centaur, Delta, or all-solid-fueled Scout vehicle—all vehicles that hadbeen used during NASA’s second decade. During this time, the agencycontinued to built Deltas and maintained its capability to build Scouts andAtlases on demand. It did not emphasize ELV development but ratherfocused on Space Shuttle development and the start of STS operationalstatus. However, the adoption of the mixed fleet strategy returned someattention to ELV development

The following section summarizes ELV activities during the decadefrom 1979 through 1988. Figure 2–7 and Table 2–13 present the successrate of each launch vehicle.

1979

NASA conducted nine launches during 1979, all successful. These usedthe Scout, the Atlas E/F, the Atlas-Centaur, and the Delta. Of the nine launch-es, three launched NASA scientific and application payloads, and six sup-ported other U.S. government and nongovernment reimbursing customers.6

A Scout vehicle launched the NASA Stratospheric Aerosol and GasExperiment (SAGE), a NASA magnetic satellite (Magsat), and a reim-bursable United Kingdom scientific satellite (UK-6/Ariel). An Atlas-Centaur launched a FltSatCom DOD communications satellite and aNASA scientific satellite (HEAO-3). Three launches used the Delta: onedomestic communications satellite for Western Union, another for RCA,and an experimental satellite, called SCATHA, for DOD. A weather satel-lite was launched on an Atlas F by the Air Force for NASA and theNational Oceanic and Atmospheric Administration (NOAA).


6Aeronautics and Space Report of the President, 1979 (Washington, DC:U.S. Government Printing Office (GPO), 1980), p. 39.

Figure 2–7. Expendable Launch Vehicle Success Rate

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1980

Seven ELV launches took place in 1980: three on Deltas, three onAtlas-Centaurs, and one on an Atlas F. Of the seven, one was for NASA;the other six were reimbursable launches for other U.S. government,international, and domestic commercial customers that paid NASA forthe launch and launch support costs.7

A Delta launched the Solar Maximum Mission, the single NASAmission, with the goal of observing solar flares and other active Sun phe-nomena and measuring total radiative output of the Sun over a six-monthperiod. A Delta also launched GOES 4 (Geostationary OperationalEnvironmental Satellite) for NOAA. The third Delta launch, for SatelliteBusiness Systems (SBS), provided integrated, all-digital, interference-free transmission of telephone, computer, electronic mail, and videocon-ferencing to clients.

An Atlas-Centaur launched FltSatCom 3 and 4 for the Navy andDOD. An Atlas-Centaur also launched Intelsat V F-2. This was the first ina series of nine satellites launched by NASA for Intelsat and was the firstthree-axis stabilized Intelsat satellite. An Atlas F launched NOAA-B, thethird in a series of Sun-synchronous operational environmental monitor-ing satellites launched by NASA for NOAA. A booster failed to place thissatellite in proper orbit, causing mission failure.

1981

During 1981, NASA launched missions on eleven ELVs: one on aScout, five using Deltas (two with dual payloads), four on Atlas-Centaurs,and one using an Atlas F. All but two were reimbursable launches forother agencies or commercial customers, and all were successful.8

A Scout vehicle launched the DOD navigation satellite, NOVA 1. Infive launches, the Delta, NASA’s most-used launch vehicle, deployedseven satellites. Two of these launches placed NASA’s scientific Explorersatellites into orbit: Dynamics Explorer 1 and 2 on one Delta and theSolar Mesosphere Explorer (along with Uosat for the University ofSurrey, England) on the other. The other three Delta launches had payingcustomers, including the GOES 5 weather satellite for NOAA and twocommunications satellites, one for SBS and one for RCA.

An Atlas-Centaur, which was the largest ELV being used by NASA,launched four missions: Comstar D-4, a domestic communications satel-lite for Comsat; two Intelsat V communications satellites for Intelsat; andthe last in the current series of FltSatCom communications satellites forDOD. An Atlas F launched the NOAA 7 weather satellite for NOAA.

LAUNCH SYSTEMS 25

7Aeronautics and Space Report of the President, 1980 (Washington, DC:GPO, 1981).


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In addition, ELVs continued to provide backup support to STS cus-tomers during the early development and transition phase of the STS system.

1982

NASA launched nine missions on nine ELVs in 1982, using sevenDeltas and two Atlas-Centaurs. Of the nine, eight were reimbursablelaunches for other agencies or commercial customers, and one was aNASA applications mission.9

The Delta supported six commercial and international communicationsmissions for which NASA was fully reimbursed: RCA’s Satcom 4 and 5,Western Union’s Westar 4 and 5, India’s Insat 1A, and Canada’s Telesat G(Anik D-1). In addition, a Delta launched Landsat 4 for NASA. The Landsatand Telesat launches used improved, more powerful Deltas. An Aerojetengine and a tank with a larger diameter increased the Delta weight-carry-ing capability into geostationary-transfer orbit by 140 kilograms. An Atlas-Centaur launched two communications satellites for the Intelsat.

1983

During 1983, NASA launched eleven satellites on eleven ELVs, usingeight Deltas, one Atlas E, one Atlas-Centaur, and one Scout. A Deltalaunch vehicle carried the European Space Agency’s EXOSAT x-rayobservatory to a highly elliptical polar orbit. Other 1983 payloadslaunched into orbit on NASA ELVs were the NASA-Netherlands InfraredAstronomy Satellite (IRAS), NOAA 8 and GOES 6 for NOAA, Hilat forthe Air Force, Intelsat VF-6 for Intelsat, Galaxy 1 and 2 for HughesCommunications, Telstar 3A for AT&T, and Satcom 1R and 2R for RCA;all except IRAS were reimbursable.10

The increased commercial use of NASA’s launch fleet and launch ser-vices conformed to President Reagan’s policy statement on May 16,1983, in which he announced that the U.S. government would facilitatethe commercial operation of the ELV program.

1984

During 1984, NASA’s ELVs provided launch support to seven satel-lite missions using four Deltas, one Scout, one Atlas-Centaur, and oneAtlas E. During this period, the Delta vehicle completed its forty-thirdconsecutive successful launch with the launching of the NATO-IIID satel-lite in November 1984. In addition, a Delta successfully launched Landsat5 for NOAA in March (Landsat program management had transferred to


9Aeronautics and Space Report of the President, 1982 (Washington, DC:GPO, 1983), p. 19.

10Aeronautics and Space Report of the President, 1983 (Washington, DC:GPO, 1984), p. 17.

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NOAA in 1983); AMPTE, a joint American, British, and German spacephysics mission involving three satellites, in August; and Galaxy-C inSeptember. Other payloads launched during 1984 by NASA ELVs includ-ed a Navy navigation satellite by a Scout, an Intelsat communicationssatellite by an Atlas-Centaur, and a NOAA weather satellite by an Atlas Fvehicle. The launch of the Intelsat satellite experienced an anomaly in thelaunch vehicle that resulted in mission failure. All missions, except theNASA scientific satellite AMPTE, were reimbursable launches for otherU.S. government, international, and domestic commercial missions thatpaid NASA for launch and launch support.11

In accordance with President Reagan’s policy directive to encouragecommercialization of the launch vehicle program, Delta, Atlas-Centaur,and Scout ELVs were under active consideration during this time by com-mercial operators for use by private industry. NASA and TranspaceCarriers, Inc. (TCI), signed an interim agreement for exclusive rights tomarket the Delta vehicle, and negotiations took place with GeneralDynamics on the Atlas-Centaur. A Commerce Business Daily announce-ment, published August 8, 1984, solicited interest for the private use ofthe Scout launch vehicle. Ten companies expressed interest in assuming atotal or partial takeover of this vehicle system.

Also in August 1984, President Reagan approved a National SpaceStrategy intended to implement the 1983 National Space Policy. Thisstrategy called for the United States to encourage and facilitate commer-cial ELV operations and minimize government regulation of these opera-tions. It also mandated that the U.S. national security sector pursue animproved assured launch capability to satisfy the need for a launch sys-tem that complemented the STS as a hedge against “unforeseen technicaland operational problems” and to use in case of crisis situations. Toaccomplish this, the national security sector should “pursue the use of alimited number of ELVs.”12

1985

In 1985, NASA’s ELVs continued to provide launch support duringthe transition of payloads to the Space Shuttle. Five launches took placeusing ELVs. Two of these were DOD satellites launched on Scouts—onefrom the Western Space and Missile Center and the other from theWallops Flight Facility. Atlas-Centaurs launched the remaining three mis-sions for Intelsat on a reimbursable basis.13

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11Aeronautics and Space Report of the President, 1984 (Washington, DC:GPO, 1985), p. 23

12White House Fact Sheet, “National Space Strategy,” August 15, 1984.13Aeronautics and Space Report of the President, 1985 (Washington, DC:

GPO, 1986).

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1986

In 1986, NASA’s ELVs launched five space application missions forNOAA and DOD. A Scout launched the Polar Beacon Experiments andAuroral Research satellite (Polar Bear) from Vandenberg Air Force Base; anAtlas-Centaur launched a FltSatCom satellite in December; an Atlas Elaunched a NOAA satellite; and two Delta vehicles were used—one tolaunch a NOAA GOES satellite and the other to launch a DOD mission. Oneof the Delta vehicles failed during launch and was destroyed before boostingthe GOES satellite into transfer orbit. An investigation concluded that thefailure was caused by an electrical short in the vehicle wiring. Wiring modi-fications were incorporated into all remaining Delta vehicles. In September,the second Delta vehicle successfully launched a DOD mission.14

Partly as a result of the Challenger accident, NASA initiated studies in1986 on the need to establish a Mixed Fleet Transportation System, consist-ing of the Space Shuttle and existing or new ELVs. This policy replaced theearlier stated intention to make the Shuttle NASA’s sole launch vehicle.

1987

In 1987, NASA launched four spacecraft missions using ELVs. Threeof these missions were successful: a Delta launch of GOES 7 for NOAAinto geostationary orbit in February; a Delta launch of Palapa B-2, a com-munications satellite for the Indonesian government, in March; and aScout launch of a Navy Transit satellite in September. In March, an Atlas-Centaur launch attempt of FltSatCom 6, a Navy communications satellite,failed when lightning in the vicinity of the vehicle caused the engines tomalfunction. The range safety officer destroyed the vehicle approximate-ly fifty-one seconds after launch.15

1988

The ELV program had a perfect launch record in 1988 with six success-ful launches. In February, a Delta ELV lifted a classified DOD payload intoorbit. This launch marked the final east coast Delta launch by a NASA launchteam. A NASA-Air Force agreement, effective July 1, officially transferredcustody of Delta Launch Complex 17 at Cape Canaveral Air Force Station tothe Air Force. Over a twenty-eight-year period, NASA had launched 143Deltas from the two Complex 17 pads. A similar transaction transferredaccountability for Atlas/Centaur Launch Complex 36 to the Air Force.16





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Also in 1988, a Scout launched San Marcos DL from the SanMarco launch facility in the Indian Ocean, a NASA-Italian scientificmission, during March. Its goal was to explore the relationshipbetween solar activity and meteorological phenomena by studying thedynamic processes that occur in the troposphere, stratosphere, andthermosphere. In April, another Scout deployed the SOOS-3, a Navynavigation satellite. In June, a third Scout carried the NOVA-II, thethird in a series of improved Navy Transit navigation satellites, intospace. The final Scout launch of the year deployed a fourth SOOS mis-sion in August. In September, an Atlas E launched NOAA H, aNational Weather Service meteorological satellite funded by NOAA,into Sun-synchronous orbit. This satellite payload included on-boardsearch-and-rescue instruments.

In addition to arranging for the purchase of launch services fromthe commercial sector, NASA took steps to divest itself of an adjunctELV capability and by making NASA-owned ELV property and ser-vices available to the private sector. During 1988, NASA finalized abarter agreement with General Dynamics that gave the company own-ership of NASA’s Atlas-Centaur flight and nonflight assets. Inexchange, General Dynamics agreed to provide the agency with twoAtlas-Centaur launches at no charge. An agreement was signed for thefirst launch service—supporting the FltSatCom F-8 Navy mission.NASA and General Dynamics also completed a letter contract for asecond launch service to support the NASA-DOD Combined Releaseand Radiation Effects Satellite (CRRES) mission. In addition, NASAtransferred its Delta vehicle program to the U.S. Air Force. Finally,enabling agreements were completed to allow ELV companies to nego-tiate directly with the appropriate NASA installation. During 1988,NASA Headquarters signed enabling agreements with McDonnellDouglas, Martin Marietta, and LTV Corporation. The Kennedy SpaceCenter and General Dynamics signed a subagreement in March toallow General Dynamics to take over maintenance and operations forLaunch Complex 36.

ELV Characteristics

The Atlas Family

The basic Atlas launch vehicle was a one-and-a-half stage stainlesssteel design built by the Space Systems Division of General Dynamics. Itwas designed as an intercontinental ballistic missile (ICBM) and was con-sidered an Air Force vehicle. However, the Atlas launch vehicle was alsoused successfully in civilian space missions dating from NASA’s earlydays. The Atlas launched all three of the unmanned lunar exploration pro-grams (Ranger, Lunar Orbiter, and Surveyor). Atlas vehicles alsolaunched the Mariner probes to Mars, Venus, and Mercury and thePioneer probes to Jupiter, Saturn, and Venus.

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NASA used two families of Atlas vehicles during the 1979–1988period: the Atlas E/F series and the Atlas-Centaur series. The Atlas E/Flaunched seven satellites during this time, six of them successful (Table2–14). The Atlas E/F space booster was a refurbished ICBM. It burnedkerosene (RP-1) and liquid oxygen in its three main engines, twoRocketdyne MA-3 booster engines, and one sustainer engine. The AtlasE/F also used two small vernier engines located at the base of the RP-1tank for added stability during flight (Table 2–15). The Atlas E/F wasdesigned to deliver payloads directly intolow-Earth orbit without the use of an upperstage.

The Atlas-Centaur (Figure 2–8) was thenation’s first high-energy launch vehicle pro-pelled by liquid hydrogen and liquid oxygen.Developed and launched under the directionof the Lewis Research Center, it becameoperational in 1966 with the launch ofSurveyor 1, the first U.S. spacecraft to soft-land on the Moon’s surface. Beginning in1979, the Centaur stage was used only incombination with the Atlas booster, but it hadbeen successfully used earlier in combinationwith the Titan III booster to launch payloadsinto interplanetary trajectories, sending twoHelios spacecraft toward the Sun and twoViking spacecraft toward Mars.17 From 1979through 1988, the Atlas-Centaur launched 18satellites with only two failures (Table 2–16).

The Centaur stage for the Atlas boosterwas upgraded in 1973 and incorporated anintegrated electronic system controlled by adigital computer. This flight-proven “astrion-ics” system checked itself and all other sys-tems prior to and during the launch phase;during flight, it controlled all events after theliftoff. This system was located on the equipment module on the forwardend of the Centaur stage. The 16,000-word capacity computer replacedthe original 4,800-word capacity computer and enabled it to take overmany of the functions previously handled by separate mechanical andelectrical systems. The new Centaur system handled navigation, guidancetasks, control pressurization, propellant management, telemetry formatsand transmission, and initiation of vehicle events (Table 2–17).


Figure 2–8. Atlas-CentaurLaunch Vehicle

17For details, see Linda Neuman Ezell, NASA Historical Data Book, VolumeIII: Programs and Projects, 1969–1978 (Washington, DC: NASA SP-4012,1988).

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The Delta Family

NASA has used the Delta launch vehicle since the agency’s inception.In 1959, NASA’s Goddard Space Flight Center awarded a contract toDouglas Aircraft Company (later McDonnell Douglas) to produce andintegrate twelve launch vehicles. The Delta, using components from theAir Force’s Thor intermediate range ballistic missile (IRBM) programand the Navy’s Vanguard launch program, was available eighteen monthslater. The Delta has evolved since that time to meet the increasingdemands of its payloads and has been the most widely used launch vehi-cle in the U.S. space program, with thirty-five launches from 1979through 1988 and thirty-four of them successful (Table 2–18).

The Delta configurations of the late 1970s and early 1980s were des-ignated the 3900 series. Figure 2–9 illustrates the 3914, and Figure 2–10shows the 3920 with the Payload Assist Module (PAM) upper stage. The3900 series resembled the earlier 2900 series (Table 2–19), except for thereplacement of the Castor II solid strap-on motors with nine larger andmore powerful Castor IV solid motors (Tables 2–20 and 2–21).

The RS-27 engine, manufactured by the Rocketdyne Division ofRockwell International, powered the first stage of the Delta. It was a single-start power plant, gimbal-mounted and operated on a combination of liquidoxygen and kerosene (RP-1). The thrust chamber was regeneratively

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Figure 2–9. Delta 3914

Figure 2–10. Delta 3920/PAM-D

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cooled, with the fuel circulating through 292 tubes that comprised theinner wall of the chamber.

The following four-digit code designated the type of Delta launchvehicle:• 1st digit designated the type of strap-on engines:

2 = Castor II, extended long tank Thor with RS-27 mainengine

3 = Castor IV, extended long tank Thor with RS-27 mainengine

• 2nd digit designated the number of strap-on engines• 3rd digit designated the type of second stage and manufacturer:

1 = ninety-six-inch manufactured by TRW (TR-201)2 = ninety-six-inch stretched tank manufactured by Aerojet

(AJ10-118K)• 4th digit designated the type of third stage:

0 = no third stage3 = TE-364-34 = TE-364-4

For example, a model desig-nation of 3914 indicated the use ofCastor IV strap-on engines,extended long tank with an RS-27main engine; nine strap-ons; aninety-six-inch second stage man-ufactured by TRW; and a TE-364-4 third stage engine. A PAMdesignation appended to the lastdigit indicated the use of aMcDonnell-Douglas PAM.

Scout Launch Vehicle

The standard Scout launchvehicle (Scout is an acronym forSolid Controlled Orbital UtilityTest) was a solid propellant four-stage booster system. It was theworld’s first all-solid propellantlaunch vehicle and was one ofNASA’s most reliable launch vehi-cles. The Scout was the smallest ofthe basic launch vehicles used byNASA and was used for orbit,probe, and reentry Earth missions(Figure 2–11).


Figure 2–11. Scout-D Launch Vehicle(Used in 1979)

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The first Scout launch took place in 1960. Since that time, forty-sixNASA Scout launches have taken place, including fourteen between 1979and 1988, when every launch was successful (Table 2–22). In addition toNASA payloads, Scout clients included DOD, the European SpaceResearch Organization, and several European governments. The Scoutwas used for both orbital and suborbital missions and has participated inresearch in navigation, astronomy, communications, meteorology, geo-desy, meteoroids, reentry materials, biology, and Earth and atmosphericsensing. It was the only U.S. ELV launched from three launch sites:Wallops on the Atlantic Ocean, Vandenberg on the Pacific Ocean, and theSan Marco platform in the Indian Ocean. It could also inject satellites intoa wider range of orbital inclinations than any other launch vehicle.

Unlike NASA’s larger ELVs, the Scout was assembled and the pay-load integrated and checked out in the horizontal position. The vehiclewas raised to the vertical orientation prior to launch. The propulsionmotors were arranged in tandem with transition sections between thestages to tie the structure together and to provide space for instrumenta-tion. A standard fifth stage was available for highly elliptical and solarorbit missions.

Scout’s first-stage motor was based on an earlier version of theNavy’s Polaris missile motor; the second-stage motor was developedfrom the Army’s Sergeant surface-to-surface missile; and the third- andfourth-stage motors were adapted by NASA’s Langley Research Centerfrom the Navy’s Vanguard missile. The fourth-stage motor used on the G model could carry almost four times as much payload to low-Earthorbit as the original model in 1960—that is, 225 kilograms versus fifty-nine kilograms (Table 2–23).

Vought Corporation, a subsidiary of LTV Corporation, was the primecontractor for the Scout launch vehicle. The Langley Research Centermanaged the Scout program.

Space Shuttle

The reusable, multipurpose Space Shuttle was designed to replace theELVs that NASA used to deliver commercial, scientific, and applicationsspacecraft into Earth’s orbit. Because of its unique design, the SpaceShuttle served as a launch vehicle, a platform for scientific laboratories,an orbiting service center for other satellites, and a return carrier for pre-viously orbited spacecraft. Beginning with its inaugural flight in 1981 andthrough 1988, NASA flew twenty-seven Shuttle missions (Table 2–24).This section focuses on the Shuttle’s use as a launch vehicle. Chapter 3discusses its use as a platform for scientific laboratories and servicingfunctions.

The Space Shuttle system consisted of four primary elements: anorbiter spacecraft, two solid rocket boosters (SRBs), an external tank tohouse fuel and an oxidizer, and three main engines. RockwellInternational built the orbiter and the main engines; Thiokol Corporation

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produced the SRB motors; and the external tank was built by MartinMarietta Corporation. The Johnson Space Center directed the orbiter andintegration contracts, while the Marshall Space Flight Center managedthe SRB, external tank, and main engine contracts.

The Shuttle could transport up to 29,500 kilograms of cargo into near-Earth orbit (185.2 to 1,111.2 kilometers). This payload was carried in a bayabout four and a half meters in diameter and eighteen meters long. Majorsystem requirements were that the orbiter and the two SRBs be reusableand that the orbiter have a maximum 160-hour turnaround time after land-ing from the previous mission. The orbiter vehicle carried personnel andpayloads to orbit, provided a space base for performing their assigned tasks,and returned personnel and payloads to Earth. The orbiter provided a hab-itable environment for the crew and passengers, including scientists andengineers. Additional orbiter characteristics are addressed in Chapter 3.

The Shuttle was launched in an upright position, with thrust provid-ed by the three main engines and the two SRBs. After about two minutes,at an altitude of about forty-four kilometers, the two boosters were spentand were separated from the orbiter. They fell into the ocean at predeter-mined points and were recovered for reuse.

The main engines continued firing for about eight minutes, cutting offat about 109 kilometers altitude just before the spacecraft was insertedinto orbit. The external tank was separated, and it followed a ballistic tra-jectory back into a remote area of the ocean but was not recovered.

Two smaller liquid rocket engines made up the orbital maneuveringsystem (OMS). The OMS injected the orbiter into orbit, performedmaneuvers while in orbit, and slowed the vehicle for reentry. After reen-try, the unpowered orbiter glided to Earth and landed on a runway.

The Shuttle used two launch sites: the Kennedy Space Center inFlorida and Vandenberg Air Force Base in California. Under optimumconditions, the orbiter landed at the site from which it was launched.However, as shown in the tables in Chapter 3 that describe the individualShuttle missions, weather conditions frequently forced the Shuttle to landat Edwards Air Force Base in California, even though it had beenlaunched from Kennedy.

Main Propulsion System

The main propulsion system (MPS) consisted of three Space Shuttlemain engines (SSMEs), three SSME controllers, the external tank, theorbiter MPS propellant management subsystem and helium subsystem,four ascent thrust vector control units, and six SSME hydraulic servo-actu-ators. The MPS, assisted by the two SRBs during the initial phases of theascent trajectory, provided the velocity increment from liftoff to a prede-termined velocity increment before orbit insertion. The Shuttle jettisonedthe two SRBs after their fuel had been expended, but the MPS continuedto thrust until the predetermined velocity was achieved. At that time, mainengine cutoff (MECO) was initiated, the external tank was jettisoned, and


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the OMS was ignited to provide the final velocity increment for orbitalinsertion. The magnitude of the velocity increment supplied by the OMSdepended on payload weight, mission trajectory, and system limitations.

Along with the start of the OMS thrusting maneuver (which settled theMPS propellants), the remaining liquid oxygen propellant in the orbiterfeed system and SSMEs was dumped through the nozzles of the engines.At the same time, the remaining liquid hydrogen propellant in the orbiterfeed system and SSMEs was dumped overboard through the hydrogen filland drain valves for six seconds. Then the hydrogen inboard fill and drainvalve closed, and the hydrogen recirculation valve opened, continuing thedump. The hydrogen flowed through the engine hydrogen bleed valves tothe orbiter hydrogen MPS line between the inboard and outboard hydro-gen fill and drain valves, and the remaining hydrogen was dumped throughthe outboard fill and drain valve for approximately 120 seconds.

During on-orbit operations, the flight crew vacuum made the MPSinert by opening the liquid oxygen and liquid hydrogen fill and drainvalves, which allowed the remaining propellants to be vented to space.Before entry into the Earth’s atmosphere, the flight crew repressurized theMPS propellant lines with helium to prevent contaminants from beingdrawn into the lines during entry and to maintain internal positive pres-sure. MPS helium also purged the spacecraft’s aft fuselage. The last activ-ity involving the MPS occurred at the end of the landing rollout. At thattime, the helium remaining in on-board helium storage tanks was releasedinto the MPS to provide an inert atmosphere for safety.

Main Engine

The SSME represented a major advance in propulsion technology.Each engine had an operating life of seven and a half hours and fifty-fivestarts and the ability to throttle a thrust level that extended over a widerange (65 percent to 109 percent of rated power level). The SSME was thefirst large, liquid-fuel rocket engine designed to be reusable.

A cluster of three SSMEs housed in the orbiter’s aft fuselage provid-ed the main propulsion for the orbiter. Ignited on the ground prior tolaunch, the cluster of liquid hydrogen–liquid oxygen engines operated inparallel with the SRBs during the initial ascent. After the boosters sepa-rated, the main engines continued to operate. The nominal operating timewas approximately eight and a half minutes. The SSMEs developed thrustby using high-energy propellants in a staged combustion cycle. The pro-pellants were partially combusted in dual preburners to produce high-pressure hot gas to drive the turbopumps. Combustion was completed inthe main combustion chamber. The cycle ensured maximum performancebecause it eliminated parasitic losses. The various thrust levels providedfor high thrust during liftoff and the initial ascent phase but allowed thrustto be reduced to limit acceleration to three g’s during the final ascentphase. The engines were gimbaled to provide pitch, yaw, and roll controlduring the orbiter boost phase.

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Key components of each engine included four turbopumps (two low-and two high-pressure), two preburners, the main injector, the main com-bustion chamber, the nozzle, and the hot-gas manifold. The manifold wasthe structural backbone of the engine. It supported the two preburners, thehigh-pressure pumps, the main injector, the pneumatic control assembly,and the main combustion chamber with the nozzle. Table 2–25 summa-rizes SSME characteristics.

The SSME was the first rocket engine to use a built-in electronic dig-ital controller. The controller accepted commands from the orbiter forengine start, shutdown, and change in throttle setting and also monitoredengine operation. In the event of a failure, the controller automaticallycorrected the problem or shut down the engine safely.

Main Engine Margin Improvement Program. Improvements to theSSMEs for increased margin and durability began with a formal Phase IIprogram in 1983. Phase II focused on turbomachinery to extend the timebetween high-pressure fuel turbopump (HPFT) overhauls by reducing theoperating temperature in the HPFT and by incorporating margin improve-ments to the HPFT rotor dynamics (whirl), turbine blade, and HPFT bear-ings. Phase II certification was completed in 1985, and all the changeswere incorporated into the SSMEs for the STS-26 mission.

In addition to the Phase II improvements, NASA made additionalchanges to the SSME to further extend the engine’s margin and durability.The main changes were to the high-pressure turbomachinery, main combus-tion chamber, hydraulic actuators, and high-pressure turbine discharge tem-perature sensors. Changes were also made in the controller software toimprove engine control. Minor high-pressure turbomachinery design changesresulted in margin improvements to the turbine blades, thereby extending theoperating life of the turbopumps. These changes included applying surfacetexture to important parts of the fuel turbine blades to improve the materialproperties in the pressure of hydrogen and incorporating a damper into thehigh-pressure oxidizer turbine blades to reduce vibration.

Plating a welded outlet manifold with nickel increased the main com-bustion chamber’s life. Margin improvements were also made to fivehydraulic actuators to preclude a loss in redundancy on the launch pad.Improvements in quality were incorporated into the servo-component coildesign, along with modifications to increase margin. To address a tem-perature sensor in-flight anomaly, the sensor was redesigned and exten-sively tested without problems.

To certify the improvements to the SSMEs and demonstrate their reli-ability through margin (or limit) testing, NASA initiated a ground test pro-gram in December 1986. Its primary purposes were to certify theimprovements and demonstrate the engine’s reliability and operating mar-gin. From December 1986 to December 1987, 151 tests and 52,363 secondsof operation (equivalent to 100 Shuttle missions) were performed. Thesehot-fire ground tests were performed at the single-engine test stands at theStennis Space Center in Mississippi and at the Rockwell InternationalRocketdyne Division’s Santa Susana Field Laboratory in California.


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NASA also conducted checkout and acceptance tests of the threemain engines for the STS-26 mission. Those tests, also at Stennis, beganin August 1987, and all three STS-26 engines were delivered to theKennedy Space Center by January 1988.

Along with hardware improvements, NASA conducted several majorreviews of requirements and procedures. These reviews addressed suchtopics as possible failure modes and effects, as well as the associated crit-ical items list. Another review involved having a launch/abort reassess-ment team examine all launch-commit criteria, engine redlines, andsoftware logic. NASA also performed a design certification review. Table2–26 lists these improvements, as well as events that occurred earlier inthe development of the SSME.

A related effort involved Marshall Space Flight Center engineerswho, working with their counterparts at Kennedy, accomplished a com-prehensive launch operations and maintenance review. This ensured thatengine processing activities at the launch site were consistent with the lat-est operational requirements.

External Tank

The external tank contained the propellants (liquid hydrogen and liq-uid oxygen) for the SSMEs and supplied them under pressure to the threemain engines in the orbiter during liftoff and ascent. Just prior to orbitalinsertion, the main engines cut off, and the external tank separated fromthe orbiter, descended through a ballistic trajectory over a predesignatedarea, broke up, and impacted in a remote ocean area. The tank was notrecovered.

The largest and heaviest (when loaded) element of the Space Shuttle,the external tank had three major components: a forward liquid oxygentank; an unpressurized intertank, which contained most of the electricalcomponents; and an aft liquid hydrogen tank. Beginning with the STS-6mission, NASA used a lightweight external tank (LWT). For each kilogram of weight reduced from the original external tank, the cargo-carrying capability of the Space Shuttle spacecraft increased one kilo-gram. The weight reduction was accomplished by eliminating portions ofstringers (structural stiffeners running the length of the hydrogen tank),using fewer stiffener rings, and by modifying major frames in the hydro-gen tank. Also, significant portions of the tank were milled differently toreduce thickness, and the weight of the external tank’s aft SRB attach-ments was reduced by using a stronger, yet lighter and less expensive,titanium alloy. Earlier, the use of the LWT reduced the total weight bydeleting the antigeyser line. The line paralleled the oxygen feed line andprovided a circulation path for liquid oxygen to reduce the accumulationof gaseous oxygen in the feed line while the oxygen tank was being filledbefore launch. After NASA assessed propellant loading data from groundtests and the first four Space Shuttle missions, engineers removed theantigeyser line for STS-5 and subsequent missions. The total length and

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diameter of the external tank remained unchanged (Figure 2–12). Table2–27 summarizes the external tank characteristics, and Table 2–28 pre-sents a chronology of external development.

As well as containing and delivering the propellant, the external tankserved as the structural backbone of the Space Shuttle during launch oper-ations. The external tank consisted of two primary tanks: a large hydro-gen tank and a smaller oxygen tank, joined by an intertank to form onelarge propellant-storage container. Superlight ablator (SLA-561) andfoam insulation sprayed on the forward part of the oxygen tank, the inter-tank, and the sides of the hydrogen tank protected the outer surfaces. Theinsulation reduced ice or frost formation during launch preparation, pro-tecting the orbiter from free-falling ice during flight. This insulation alsominimized heat leaks into the tank, avoided excessive boiling of the liq-uid propellants, and prevented liquification and solidification of the airnext to the tank.

The external tank attached to the orbiter at one forward attachmentpoint and two aft points. In the aft attachment area, umbilicals carried flu-ids, gases, electrical signals, and electrical power between the tank andthe orbiter. Electrical signals and controls between the orbiter and the twoSRBs also were routed through those umbilicals.

Liquid Oxygen Tank. The liquid oxygen tank was an aluminummonocoque structure composed of a fusion-welded assembly of pre-formed, chem-milled gores, panels, machined fittings, and ring chords. Itoperated in a pressure range of 1,035 to 1,138 mmHg. The tank containedantislosh and antivortex provisions to minimize liquid residuals and dampfluid motion. The tank fed into a 0.43-meter-diameter feedline that sentthe liquid oxygen through the intertank, then outside the external tank tothe aft righthand external tank/orbiter disconnect umbilical. The feedlinepermitted liquid oxygen to flow at approximately 1,268 kilograms per


Figure 2–12. External Tank

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second, with the SSMEs operating at 104 percent of rated thrust, or per-mitted a maximum flow of 71,979 liters per minute. The liquid oxygentank’s double-wedge nose cone reduced drag and heating, contained thevehicle’s ascent air data system, and served as a lightning rod.

Intertank. The intertank was not a tank in itself but provided amechanical connection between the liquid oxygen and liquid hydrogentanks. The primary functions of the intertank were to provide structuralcontinuity to the propellant tanks, to serve as a protective compartment tohouse instruments, and to receive and distribute thrust loads from theSRBs. The intertank was a steel/aluminum semimonocoque cylindricalstructure with flanges on each end for joining the liquid oxygen and liq-uid hydrogen tanks. It housed external tank instrumentation componentsand provided an umbilical plate that interfaced with the ground facilityarm for purging the gas supply, hazardous gas detection, and hydrogengas boiloff during ground operations. It consisted of mechanically joinedskin, stringers, and machined panels of aluminum alloy. The intertankwas vented during flight. It contained the forward SRB-external tankattach thrust beam and fittings that distributed the SRB loads to the liquidoxygen and liquid hydrogen tanks.

Liquid Hydrogen Tank. The liquid hydrogen tank was an aluminumsemimonocoque structure of fusion-welded barrel sections, five majorring frames, and forward and aft ellipsoidal domes. Its operating pressurewas 1,759 mmHg. The tank contained an antivortex baffle and siphon out-let to transmit the liquid hydrogen from the tank through a 0.43-meter lineto the left aft umbilical. The liquid hydrogen feedline flow rate was 211.4 kilograms per second, with the SSMEs at 104 percent of ratedthrust, or a maximum flow of 184,420 liters per minute. At the forwardend of the liquid hydrogen tank was the external tank/orbiter forwardattachment pod strut, and at its aft end were the two external tank/orbiteraft attachment ball fittings as well as the aft SRB-external tank stabiliz-ing strut attachments.

External Tank Thermal Protection System. The external tank ther-mal protection system consisted of sprayed-on foam insulation and pre-molded ablator materials. The system also included the use of phenolicthermal insulators to preclude air liquefaction. Thermal isolators wererequired for liquid hydrogen tank attachments to preclude the liquefactionof air-exposed metallic attachments and to reduce heat flow into the liq-uid hydrogen. The thermal protection system weighed 2,192 kilograms.

External Tank Hardware. The external hardware, externaltank/orbiter attachment fittings, umbilical fittings, and electrical andrange safety system weighed 4,136.4 kilograms.

Each propellant tank had a vent and relief valve at its forward end.This dual-function valve could be opened by ground support equipmentfor the vent function during prelaunch and could open during flight whenthe ullage (empty space) pressure of the liquid hydrogen tank reached1,966 mmHg or the ullage pressure of the liquid oxygen tank reached1,293 mmHg.

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The liquid oxygen tank contained a separate, pyrotechnically operat-ed, propulsive tumble vent valve at its forward end. At separation, the liq-uid oxygen tumble vent valve was opened, providing impulse to assist inthe separation maneuver and more positive control of the entry aerody-namics of the external tank.

There were eight propellant-depletion sensors, four each for fuel andoxidizer. The fuel-depletion sensors were located in the bottom of the fueltank. The oxidizer sensors were mounted in the orbiter liquid oxygenfeedline manifold downstream of the feedline disconnect. During SSMEthrusting, the orbiter general purpose computers constantly computed theinstantaneous mass of the vehicle because of the usage of the propellants.Normally, MECO was based on a predetermined velocity; however, if anytwo of the fuel or oxidizer sensors sensed a dry condition, the engineswould be shut down.

The locations of the liquid oxygen sensors allowed the maximumamount of oxidizer to be consumed in the engines, while allowing suffi-cient time to shut down the engines before the oxidizer pumps ran dry. Inaddition, 500 kilograms of liquid hydrogen were loaded over and abovethat required by the six-to-one oxidizer/fuel engine mixture ratio. Thisassured that MECO from the depletion sensors was fuel rich; oxidizer-rich engine shutdowns could cause burning and severe erosion of enginecomponents.

Four pressure transducers located at the top of the liquid oxygen andliquid hydrogen tanks monitored the ullage pressures. Each of the two aftexternal tank umbilical plates mated with a corresponding plate on theorbiter. The plates helped maintain alignment among the umbilicals.Physical strength at the umbilical plates was provided by bolting corre-sponding umbilical plates together. When the orbiter general purposecomputers commanded external tank separation, the bolts were severedby pyrotechnic devices.

The external tank had five propellant umbilical valves that interfacedwith orbiter umbilicals—two for the liquid oxygen tank and three for theliquid hydrogen tank. One of the liquid oxygen tank umbilical valves wasfor liquid oxygen, the other for gaseous oxygen. The liquid hydrogen tankumbilical had two valves for liquid and one for gas. The intermediate-diameter liquid hydrogen umbilical was a recirculation umbilical usedonly during the liquid hydrogen chill-down sequence during prelaunch.

The external tank also had two electrical umbilicals that carried elec-trical power from the orbiter to the tank and the two SRBs and providedinformation from the SRBs and external tank to the orbiter. A swing-arm-mounted cap to the fixed service structure covered the oxygen tank venton top of the external tank during countdown and was retracted about twominutes before liftoff. The cap siphoned off oxygen vapor that threatenedto form large ice on the external tank, thus protecting the orbiter’s ther-mal protection system during launch.

External Tank Range Safety System. A range safety system, moni-tored by the flight crew, provided for dispersing tank propellants if nec-


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essary. It included a battery power source, a receiver/decoder, antennas,and ordnance.

Post-Challenger Modification. Prior to the launch of STS-26, NASAmodified the external tank by strengthening the hydrogen pressurizationline. In addition, freezer wrap was added to the hydrogen line. This per-mitted the visual detection of a hydrogen fire (Table 2–28).

Solid Rocket Boosters

The two SRBs provided the main thrust to lift the Space Shuttle offthe pad and up to an altitude of about forty-four and a half kilometers. Inaddition, the two SRBs carried the entire weight of the external tank andorbiter and transmitted the weight load through their structure to themobile launcher platform. The SRBs were ignited after the three SSMEs’thrust level was verified. The two SRBs provided 71.4 percent of thethrust at liftoff and during first-stage ascent. Seventy-five seconds afterSRB separation, SRB apogee occurred at an altitude of approximatelysixty-five kilometers. SRB impact occurred in the ocean approximately226 kilometers downrange, to be recovered and returned for refurbish-ment and reuse.

The primary elements of each booster were the motor (includingcase, propellant, igniter, and nozzle), structure, separation systems, oper-ational flight instrumentation, recovery avionics, pyrotechnics, decelera-tion system, thrust vector control system, and range safety destructsystem (Figure 2–13). Each booster attached to the external tank at theSRB’s aft frame with two lateral sway braces and a diagonal attachment.The forward end of each SRB joined the external tank at the forward end

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Figure 2–13. Solid Rocket Booster

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of the SRB’s forward skirt. On the launch pad, each booster also con-nected to the mobile launcher platform at the aft skirt with four bolts andnuts that were severed by small explosives at liftoff.

The SRBs were used as matched pairs. Each consisted of four solidrocket motor (SRM) segments. The pairs were matched by loading eachof the four motor segments in pairs from the same batches of propellantingredients to minimize any thrust imbalance. The exhaust nozzle in theaft segment of each motor, in conjunction with the orbiter engines, steeredthe Space Shuttle during the powered phase of launch. The segmented-casing design assured maximum flexibility in fabrication and ease oftransportation and handling. Each segment was shipped to the launch siteon a heavy-duty rail car with a specially built cover.

The propellant mixture in each SRB motor consisted of an ammoni-um perchlorate (oxidizer, 69.6 percent by weight), aluminum (fuel,16 percent), iron oxide (a catalyst, 0.4 percent), a polymer (a binder thatheld the mixture together, 12.04 percent), and an epoxy curing agent(1.96 percent). The propellant was an eleven-point star-shaped perfora-tion in the forward motor segment and a double-truncated-cone perfora-tion in each of the aft segments and aft closure. This configurationprovided high thrust at ignition and then reduced the thrust by approxi-mately one-third fifty seconds after liftoff to prevent overstressing thevehicle during maximum dynamic pressure.

The cone-shaped aft skirt supported the four aft separation motors.The aft section contained avionics, a thrust vector control system that con-sisted of two auxiliary power units and hydraulic pumps, hydraulic sys-tems, and a nozzle extension jettison system. The forward section of eachbooster contained avionics, a sequencer, forward separation motors, a nosecone separation system, drogue and main parachutes, a recovery beacon, arecovery light, a parachute camera on selected flights, and a range safetysystem. Each SRB incorporated a range safety system that included a bat-tery power source, a receiver-decoder, antennas, and ordnance.

Each SRB had two integrated electronic assemblies, one forward andone aft. After burnout, the forward assembly initiated the release of thenose cap and frustum and turned on the recovery aids. The aft assembly,mounted in the external tank-SRB attach ring, connected with the forwardassembly and the orbiter avionics systems for SRB ignition commandsand nozzle thrust vector control. Each integrated electronic assembly hada multiplexer-demultiplexer, which sent or received more than one mes-sage, signal, or unit of information on a single communications channel.

Eight booster separation motors (four in the nose frustum and four inthe aft skirt) of each SRB thrust for 1.02 seconds at SRB separation fromthe external tank. SRB separation from the external tank was electricallyinitiated. Each solid rocket separation motor was 0.8 meter long and 32.5 centimeters in diameter (Table 2–29).

Location aids were provided for each SRB, frustum-drogue chutes,and main parachutes. These included a transmitter, antenna, strobe/con-verter, battery, and saltwater switch electronics. The recovery crew


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retrieved the SRBs, frustum/drogue chutes, and main parachutes. Thenozzles were plugged, the solid rocket motors were dewatered, and thecrew towed the SRBs back to the launch site. Each booster was removedfrom the water, and its components disassembled and washed with freshand de-ionized water to limit saltwater corrosion. The motor segments,igniter, and nozzle were shipped back to Thiokol for refurbishment. TheSRB nose caps and nozzle extensions were not recovered.

Testing and production of the SRB were well under way in 1979. Thebooster performed well until the Challenger accident revealed flaws thathad very likely existed for several missions but had resulted in little reme-dial action. The 1986 Challenger accident forced major modifications tothe SRB and SRM.

Post-Challenger Modifications. On June 13, 1986, President Reagandirected NASA to implement, as soon as possible, the recommendationsof the Presidential Commission on the Space Shuttle ChallengerAccident. During the downtime following the Challenger accident,NASA analyzed critical structural elements of the SRB, primarilyfocused in areas where anomalies had been noted during postflightinspection of recovered hardware.

Anomalies had been noted in the attach ring where the SRBs joinedthe external tank. Some of the fasteners showed distress where the ringattached to the SRB motor case. Tests attributed this to the high loadsencountered during water impact. To correct the situation and ensurehigher strength margins during ascent, the attach ring was redesigned toencircle the motor case completely (360 degrees). Previously, the attachring formed a “C” and encircled the motor case 270 degrees.

In addition, NASA performed special structural tests on the aft skirt.During this test program, an anomaly occurred in a critical weld betweenthe hold-down post and skin of the skirt. A redesign added reinforcementbrackets and fittings in the aft ring of the skirt. These modifications addedapproximately 200 kilograms to the weight of each SRB.

Solid Rocket Motor Redesign. The Presidential Commission deter-mined that the cause of the loss of the Challenger was “a failure in thejoint between the two lower segments of the right solid rocket motor. Thespecific failure was the destruction of the seals that are intended to pre-vent hot gases from leaking through the joint during the propellant burnof the rocket motor.”18

Consequently, NASA developed a plan for a redesigned solid rocketmotor (RSRM). Safety in flight was the primary objective of the SRMredesign. Minimizing schedule impact by using existing hardware, to theextent practical, without compromising safety was another objective.

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18Report at a Glance, report to the President by the Presidential Commissionon the Space Shuttle Challenger Accident, Chapter IV, “The Cause of theAccident,” Finding (no pg. number).

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NASA established a joint redesign team with participants from theMarshall Space Flight Center, other NASA centers, Morton Thiokol, andoutside NASA. The team developed an “SRM Redesign Project Plan” toformalize the methodology for SRM redesign and requalification. Theplan provided an overview of the organizational responsibilities and rela-tionships; the design objectives, criteria, and process; the verificationapproach and process; and a master schedule. Figure 2–14 shows theSRM Project Schedule as of August 1986. The companion “Developmentand Verification Plan” defined the test program and analyses required toverify the redesign and unchanged components of the SRM. The SRMwas carefully and extensively redesigned. The RSRM received intensescrutiny and was subjected to a thorough certification process to verifythat it worked properly and to qualify the motor for human spaceflight.

NASA assessed all aspects of the existing SRM and required designchanges in the field joint, case-to-nozzle joint, nozzle, factory joint, pro-pellant grain shape, ignition system, and ground support equipment. Thepropellant, liner, and castable inhibitor formulations did not requirechanges. Design criteria were established for each component to ensure asafe design with an adequate margin of safety. These criteria focused onloads, environments, performance, redundancy, margins of safety, andverification philosophy.

The team converted the criteria into specific design requirements dur-ing the Preliminary Requirements Reviews held in July and August 1986.NASA assessed the design developed from these requirements at thePreliminary Design Review held in September 1986 and baselined inOctober 1986. NASA approved the final design at the Critical Design


Figure 2–14. Solid Rocket Motor Redesign Schedule

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Review held in October 1987. Manufacture of the RSRM test hardwareand the first flight hardware began prior to the Preliminary DesignReview and continued in parallel with the hardware certification program.The Design Certification Review considered the analyses and test resultsversus the program and design requirements to certify that the RSRM wasready to fly.

Specific Modifications. The SRM field-joint metal parts, internalcase insulation, and seals were redesigned, and a weather protection sys-tem was added. The major change in the motor case was the new tangcapture feature to provide a positive metal-to-metal interference fitaround the circumference of the tang and clevis ends of the mating seg-ments. The interference fit limited the deflection between the tang andclevis O-ring sealing surfaces caused by motor pressure and structuralloads. The joints were designed so that the seals would not leak undertwice the expected structural deflection and rate.

The new design, with the tang capture feature, the interference fit,and the use of custom shims between the outer surface of the tang andinner surface of the outer clevis leg, controlled the O-ring sealing gapdimension. The sealing gap and the O-ring seals were designed so that apositive compression (squeeze) was always on the O-rings. The minimumand maximum squeeze requirements included the effects of temperature,O-ring resiliency and compression set, and pressure. The redesignincreased the clevis O-ring groove dimension so that the O-ring neverfilled more than 90 percent of the O-ring groove, and pressure actuationwas enhanced.

The new field-joint design also included a new O-ring in the capturefeature and an additional leak check port to ensure that the primary O-ringwas positioned in the proper sealing direction at ignition. This new orthird O-ring also served as a thermal barrier in case the sealed insulationwas breached. The field-joint internal case insulation was modified to besealed with a pressure-actuated flap called a j-seal, rather than with puttyas in the STS 51-L (Challenger) configuration.

The redesign added longer field-joint-case mating pins, with a recon-figured retainer band, to improve the shear strength of the pins andincrease the metal parts’ joint margin of safety. The joint safety margins,both thermal and structural, were demonstrated over the full ranges ofambient temperature, storage compression, grease effect, assembly stress-es, and other environments. The redesign incorporated external heaterswith integral weather seals to maintain the joint and O-ring temperatureat a minimum of 23.9 degrees Celsius. The weather seal also preventedwater intrusion into the joint.

Original Versus Redesigned SRM Case-to-Nozzle Joint. The SRMcase-to-nozzle joint, which experienced several instances of O-ring ero-sion in flight, was redesigned to satisfy the same requirements imposedon the case field joint. Similar to the field joint, case-to-nozzle joint mod-ifications were made in the metal parts, internal insulation, and O-rings.The redesign added radial bolts with Stato-O-Seals to minimize the joint

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sealing gap opening. The internal insulation was modified to be sealedadhesively, and a third O-ring was included. The third O-ring served as adam or wiper in front of the primary O-ring to prevent the polysulfideadhesive from being extruded in the primary O-ring groove. It also servedas a thermal barrier in case the polysulfide adhesive was breached. Thepolysulfide adhesive replaced the putty used in the STS 51-L joint. Also,the redesign added an another leak check port to reduce the amount oftrapped air in the joint during the nozzle installation process and to aid inthe leak check procedure.

Nozzle. Redesigned internal joints of the nozzle metal parts incorpo-rated redundant and verifiable O-rings at each joint. The modified nozzlesteel fixed housing part permitted the incorporation of the 100 radial boltsthat attached the fixed housing to the case’s aft dome. The new nozzlenose inlet, cowl/boot, and aft exit cone assemblies used improved bond-ing techniques. Increasing the thickness of the aluminum nose inlet hous-ing and improving the bonding process eliminated the distortion of thenose inlet assembly’s metal-part-to-ablative-parts bond line. The changedtape-wrap angle of the carbon cloth fabric in the areas of the nose inletand throat assembly parts improved the ablative insulation erosion toler-ance. Some of these ply-angle changes had been in progress prior to STS51-L. Additional structural support with increased thickness and contourchanges to the cowl and outer boot ring increased their margins of safety.In addition, the outer boot ring ply configuration was altered.

Factory Joint. The redesign incorporated minor modifications in thecase factory joints by increasing the insulation thickness and layup toincrease the margin of safety on the internal insulation. Longer pins werealso added, along with a reconfigured retainer band and new weather sealto improve factory joint performance and increase the margin of safety. Inaddition, the redesign changed the O-ring and O-ring groove size to beconsistent with the field joint.

Propellant. The motor propellant forward transition region wasrecontoured to reduce the stress fields between the star and cylindricalportions of the propellant grain.

Ignition System. The redesign incorporated several minor modifica-tions into the ignition system. The aft end of the igniter steel case, whichcontained the igniter nozzle insert, was thickened to eliminate a localizedweakness. The igniter internal case insulation was tapered to improve themanufacturing process. Finally, although vacuum putty was still used atthe joint of the igniter and case forward dome, it eliminated asbestos asone of its constituents.

Ground Support Equipment. Redesigned ground support equipment(1) minimized the case distortion during handling at the launch site,(2) improved the segment tang and clevis joint measurement system formore accurate reading of case diameters to facilitate stacking, (3) mini-mized the risk of O-ring damage during joint mating, and (4) improvedleak testing of the igniter, case, and nozzle field joints. A ground supportequipment assembly aid guided the segment tang into the clevis and


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rounded the two parts with each other. Other ground support equipmentmodifications included transportation monitoring equipment and the lift-ing beam.

Testing. Tests of the redesigned motor were carried out in a horizon-tal attitude, providing a more accurate simulation of actual conditions ofthe field joint that failed during the STS 51-L mission. In conjunction withthe horizontal attitude for the RSRM full-scale testing, NASA incorporat-ed externally applied loads. Morton Thiokol constructed a second hori-zontal test stand for certification of the redesigned SRM. The contractorused this new stand to simulate environmental stresses, loads, and tem-peratures experienced during an actual Space Shuttle launch and ascent.The new test stand also provided redundancy for the original stand.

The testing program included five full-scale firings of the RSRMprior to STS-26 to verify the RSRM performance. These included twodevelopment motor tests, two qualification motor tests, and a productionverification motor test. The production verification motor test in August1988 intentionally introduced severe artificial flaws into the test motor tomake sure that the redundant safety features implemented during theredesign effort worked as planned. Laboratory and component tests wereused to determine component properties and characteristics. Subscalemotor tests simulated gas dynamics and thermal conditions for compo-nents and subsystem design. Simulator tests, consisting of motors usingfull-size flight-type segments, verified joint design under full flight loads,pressure, and temperature.

Full-scale tests verified analytical models and determined hardwareassembly characteristics; joint deflection characteristics; joint perfor-mance under short duration, hot-gas tests, including joint flaws and flightloads; and redesigned hardware structural characteristics. Table 2–30 liststhe events involved in the redesign of the SRB and SRM as well as earli-er events in their development.19

Upper Stages

The upper stages boost payloads from the Space Shuttle’s parkingorbit or low-Earth orbit to geostationary-transfer orbit or geosynchronousorbit. They are also used on ELV missions to boost payloads from anearly stage of the orbit maneuver into geostationary-transfer orbit or geo-synchronous orbit. The development of the upper stages used by NASAbegan prior to 1979 and continued throughout the 1980s (Table 2–31).

The upper stages could be grouped into three categories, according totheir weight delivery capacity:• Low capacity: 453- to 1,360-kilogram capacity to geosynchronous

orbit

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19See Ezell, NASA Historical Data Book, Volume III, for earlier events inSRB development.

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• Medium capacity: 1,360- to 3,175-kilogram capacity to geosynchro-nous orbit

• High capacity: 3,175- to 5,443-kilogram capacity to geosynchronousorbit

Inertial Upper Stages

DOD designed and developed the Inertial Upper Stage (IUS) medium-capacity system for integration with both the Space Shuttle andTitan launch vehicle. It was used to deliver spacecraft into a wide rangeof Earth orbits beyond the Space Shuttle’s capability. When used with theShuttle, the solid-propellant IUS and its payload were deployed from theorbiter in low-Earth orbit. The IUS was then ignited to boost its payloadto a higher energy orbit. NASA used a two-stage configuration of the IUSprimarily to achieve geosynchronous orbit and a three-stage version forplanetary orbits.

The IUS was 5.18 meters long and 2.8 meters in diameter andweighed approximately 14,772 kilograms. It consisted of an aft skirt, anaft stage SRM with 9,707 kilograms of solid propellant generating202,828.8 newtons of thrust, an interstage, a forward stage SRM with2,727.3 kilograms of propellant generating 82,288 newtons of thrust andusing an extendible exit cone, and an equipment support section. Theequipment support section contained the avionics that provided guidance,navigation, telemetry, command and data management, reaction control,and electrical power. All mission-critical components of the avionics sys-tem and thrust vector actuators, reaction control thrusters, motor igniter,and pyrotechnic stage separation equipment were redundant to ensurebetter than 98-percent reliability (Figure 2–15).


Figure 2–15. Inertial Upper Stage

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The spacecraft was attached to the IUS at a maximum of eight attach-ment points. These points provided substantial load-carrying capabilitywhile minimizing thermal transfer. Several IUS interface connectors pro-vided power and data transmission to the spacecraft. Access to these con-nectors could be provided on the spacecraft side of the interface plane orthrough the access door on the IUS equipment bay.

The IUS provided a multilayer insulation blanket of aluminizedKapton with polyester net spacers and an aluminized beta cloth outerlayer across the IUS and spacecraft interface. All IUS thermal blanketsvented toward and into the IUS cavity. All gases within the IUS cavityvented to the orbiter payload bay. There was no gas flow between thespacecraft and the IUS. The thermal blankets were grounded to the IUSstructure to prevent electrostatic charge buildup.

Beginning with STS-26, the IUS incorporated a number of advancedfeatures. It had the first completely redundant avionics system developedfor an uncrewed space vehicle. This system could correct in-flight fea-tures within milliseconds. Other advanced features included a carboncomposite nozzle throat that made possible the high-temperature, long-duration firing of the IUS motor and a redundant computer system inwhich the second computer could take over functions from the primarycomputer, if necessary.

Payload Assist Module

The Payload Assist Module (PAM), which was originally called theSpinning Stage Upper Stage, was developed by McDonnell Douglas at itsown expense for launching smaller spacecraft to geostationary-transferorbit. It was designed as a higher altitude booster of satellites deployed innear-Earth orbit but operationally destined for higher altitudes. The PAM-D could launch satellites weighing up to 1,247 kilograms. It wasoriginally configured for satellites that used the Delta ELV but was usedon both ELVs and the Space Shuttle. The PAM-DII (used on STS 61-Band STS 61-C) could launch satellites weighing up to 1,882 kilograms. Athird PAM, the PAM-A, had been intended for satellites weighing up to1,995 kilograms and was configured for missions using the Atlas-Centaur.NASA halted its development in 1982, pending definition of spacecraftneeds. Commercial users acquired the PAM-D and PAM-DII directlyfrom the manufacturer.

The PAM consisted of a deployable (expendable) stage and reusableairborne support equipment. The deployable stage consisted of a spin-stabilized SRM, a payload attach fitting to mate with the unmannedspacecraft, and the necessary timing, sequencing, power, and controlassemblies.

The PAM’s airborne support equipment consisted of the reusable hard-ware elements required to mount, support, control, monitor, protect, andoperate the PAM’s expendable hardware and untended spacecraft fromliftoff to deployment from the Space Shuttle or ELV. It also provided these

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functions for the safing and return of the stage and spacecraft in case of anaborted mission. The airborne support equipment was designed to be asself-contained as possible. The major airborne support equipment elementsincluded the cradle for structural mounting and support, the spin table anddrive system, the avionics system to control and monitor the airborne sup-port equipment and the PAM vehicle, and the thermal control system.

The PAM stages were supported through the spin table at the base ofthe motor and through restraints at the PAF. The forward restraints wereretracted before deployment. The sunshield of the PAM-D and DII pro-vided thermal protection of the PAM/untended spacecraft when the SpaceShuttle orbiter payload bay doors were open on orbit.

Transfer Orbit Stage

The development of the Transfer Orbit Stage (TOS) began in April1983 when NASA signed a Space System Development Agreement withOrbital Sciences Corporation (OSC) to develop a new upper stage. Underthe agreement, OSC provided technical direction, systems engineering,mission integration, and program management of the design, production,and testing of the TOS. NASA, with participation by the Johnson andKennedy Space Centers, provided technical assistance during TOS devel-opment and agreed to provide technical monitoring and advice duringTOS development and operations to assure its acceptability for use withmajor national launch systems, including the STS and Titan vehicles.NASA also established a TOS Program Office at the Marshall SpaceFlight Center. OSC provided all funding for the development and manu-facturing of TOS (Figure 2–16).

In June 1985, Marshall awarded a 16-month contract to OSC for alaser initial navigation system (LINS) developed for the TOS. Marshallwould use the LINS for guidance system research, testing, and other pur-poses related to the TOS program.

Production of the TOS began in mid-1986. It was scheduled to be used on theAdvanced Communications TechnologySatellite (ACTS) and the PlanetaryObserver series of scientific explorationspacecraft, beginning with the MarsObserver mission in the early 1990s.

The TOS could place 2,490 to6,080 kilograms payloads into geosta-tionary-transfer orbit from the STS andup to 5,227 kilograms from the TitanIII and IV and could also deliver space-craft to planetary and other high-ener-gy trajectories. The TOS allowedsmaller satellites to be placed into geo-stationary-transfer orbit in groups of


Figure 2–16. Transfer Orbit Stage

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two or three. Two payloads of the Atlas class (1,136 kilograms) or threepayloads of the Delta class (636 kilograms) could be launched on a sin-gle TOS mission. Besides delivery of commercial communications satel-lites, its primary market, the TOS would be used for NASA and DODmissions.

The TOS system consisted of flight vehicle hardware and softwareand associated airborne and ground support equipment required forbuildup. Table 3–32 lists its characteristics. Performance capabilities ofthe TOS included:• Earth escape transfer capability• Geosynchronous transfer orbit capability• Orbit inclination change capability• Low-altitude transfer capability• Intermediate transfer orbit capability• De-orbit maneuver• Satellite repair and retrieval

Apogee and Maneuvering System

The liquid bipropellant Apogee and Maneuvering System (AMS) wasdesigned to be used both with and independently of the TOS. The AMSwould boost the spacecraft into a circular orbit and allow on-orbit maneu-vering. Martin Marietta Denver Aerospace worked to develop the AMSwith Rockwell International’s Rocketdyne Division, providing the AMSRS-51 bipropellant rocket engine, and Honeywell, Inc., supplied theTOS/AMS LINS avionics system.

When it became operational, the TOS/AMS combination woulddeliver up to approximately 2,950 kilograms into geosynchronous orbitfrom the orbiter’s parking orbit into final geosynchronous orbit. TheTOS/AMS would have a delivery capability 30 percent greater than theIUS and would reduce stage and STS user costs. The main propulsion,reaction control, avionics, and airborne support equipment systems wouldbe essentially the same as those used on the TOS. In particular, the avion-ics would be based on a redundant, fault-tolerant LINS.

Operating alone, the AMS would be able to place communicationssatellites weighing up to approximately 2,500 kilograms into geostation-ary-transfer orbit after deployment in the standard Space Shuttle parkingorbit. Other missions would include low-orbit maneuvering between theShuttle and the planned space station, delivery of payloads to Sun-synchronous and polar orbits, and military on-demand maneuvering capa-bility. The AMS was planned to be available for launch in early 1989 andwould provide an alternative to the PAM-DII.

The avionics, reaction control system, and airborne support equip-ment designs of the AMS would use most of the standard TOS compo-nents. Main propulsion would be provided by the 2,650-pound thrustRocketdyne RS-51 engine. This engine was restartable and operable overextended periods. A low-thrust engine option that provided 400 pounds ofthrust would also be available for the AMS.

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Centaur Upper Stage

NASA studied and began production in the early 1980s of a modifiedCentaur upper stage for use with the STS for planetary and heavier geo-synchronous mission applications. The proposed modifications wouldincrease the size of the propellant tanks to add about 50 percent more pro-pellant capacity and make the stage compatible with the Space Shuttle.This wide-body version would use the same propulsion system and about85 percent of the existing Centaur’s avionics systems. Contracts werenegotiated with General Dynamics, Honeywell, Pratt & Whitney, andTeledyne for the design, development, and procurement of Centaur upperstages for the Galileo and International Solar Polar missions that werescheduled for 1986.

However, following the Challenger accident, NASA determined thateven with modifications, the Centaur could not comply with necessarysafety requirements for use on the Shuttle. The Centaur upper stage ini-tiative was then dropped.

Advanced Programs

Advanced programs focused on future space transportation activi-ties, including improving space transportation operations through theintroduction of more advanced technologies and processes, and on ser-vicing and protecting U.S. space assets. The following sectionsdescribe NASA’s major advanced program initiatives. Several of theefforts progressed from advanced program status to operational statusduring this decade.

Orbital Transfer Vehicle

NASA’s Advanced Planning/Programs Division of the Office ofSpace Transportation identified the need for an Orbital Transfer Vehicle(OTV) in the early 1980s, when it became obvious that a way was need-ed to transport payloads from the Space Shuttle’s low-Earth orbit to ahigher orbit and to retrieve and return payloads to the Shuttle or futurespace station. The Marshall Space Flight Center was designated as thelead center for the development effort, and the Lewis Research Center ledthe propulsion system studies. An untended OTV was proposed for a firstflight in the early 1990s.

NASA believed that the use of aerobraking was necessary to make theOTV affordable. Studies beginning in 1981 conducted at Marshall by def-inition phase contractors Boeing Aerospace Company and GeneralElectric Reentry Systems determined that aerodynamic braking was anefficient fuel-saving technique for the OTV, perhaps doubling payloadcapacity. This technique would use the Earth’s atmosphere as a brakingmechanism for return trips, possibly supplemented by the use of a ballute,an inflatable drag device. When the transfer vehicle passed through the


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atmosphere, the friction of the air against the vehicle would provideenough drag to slow the vehicle. Otherwise, a rocket engine firing wouldbe required to brake the vehicle. Aeroassist braking would save one burn,and the extra fuel could be used to transport a larger payload to a highorbit. The aeroassisted braking could result in about a twofold increase inthe amount of payload that could be ferried to high altitudes.

Boeing’s studies emphasized low lifting-body designs—“low lift-to-drag ratio”—designs with a relatively low capability of lift to enable themto fly, but ones that weigh less. General Electric Reentry Systems focusedon moderate lift-to-drag ratio designs—relatively moderate lift capabilityand somewhat heavier weight.

In 1981, NASA designated the Lewis Research Center the lead cen-ter for OTV propulsion technology. This program supported technologyfor three advanced engine concepts that were developed by AerojetTechSystems, Pratt & Whitney, and Rocketdyne to satisfy a NASA-supplied set of goals. The proposed engines would be used to transferloads—both personnel and cargo—between low-Earth orbit and geosyn-chronous orbit, and beyond. In addition, because OTVs would facerequirements ranging from high-acceleration round-trip transfers forresupply to very low-acceleration one-way transfers of large, flexiblestructures, NASA investigated variable thrust propulsion systems, whichwould provide high performance over a broad throttling range.

In 1983, NASA chose the same three contractors to begin a programleading to the design, development, test, and engineering of the OTV.These contracts expired in 1986. NASA sponsored another competitiveprocurement to continue the OTV propulsion program. Funding wasreduced, and only Rocketdyne and Aerojet continued the advancedengine technology development. Component testing began in 1988, andfurther investigations into aerobraking continued into the 1990s.

The OTV would be used primarily to place NASA, DOD, and com-mercial satellites and space platforms into geosynchronous orbit. TheOTV could also deliver large payloads into other orbits and boost plane-tary exploration spacecraft into high-velocity orbits approaching theirmission trajectory. The vehicle was expected to use liquid oxygen–liquidhydrogen propellants.

The OTV’s reusable design provided for twenty flights before it hadto be refurbished or replaced. Because of its reusability, the OTV wouldsignificantly reduce payload transportation costs.

At the same time, that Lewis was leading propulsion studies,Marshall initiated studies in 1984 to define OTV concepts and choseBoeing Aerospace and Martin Marietta to conduct the conceptual studies.The studies examined the possibilities of both a space-based and anEarth-based OTV. Both would initially be uncrewed upper stages. Theultimate goal, however, was to develop a crewed vehicle capable of fer-rying a crew capsule to geosynchronous orbit. The vehicle would thenreturn the crew and capsule for other missions. The development of acrew capsule for the OTV was planned for the 1990s.

LAUNCH SYSTEMS 53

*DB Chap 2 (11-58) 1/17/02 2:51 PM Page 53

The Space Shuttle would carry the Earth-based OTV into space. Itwould be launched from the Shuttle’s payload bay or from an aft cargocarrier attached to the aft end of the Shuttle’s external tank. The OTVwould transfer payloads from a low orbit to a higher one. It would alsoretrieve payloads in high orbits and return them to the Shuttle. The OTVwould then return to Earth in the Shuttle’s payload bay. The OTV wouldseparate from the Shuttle’s external tank at about the same time that thepayload was deployed from the orbiter’s cargo bay. The two componentswould then join together and begin to travel to a higher orbit. This Earth-based OTV offered the advantage of performing vehicle maintenance andrefueling on the ground with the help of gravity, ground facilities, andworkers who do not have to wear spacesuits.

A space-based OTV would be based at the future space station. Itwould move payloads into higher orbit from the space station and thenreturn to its home there. It would be refueled and maintained at the spacestation. Studies showed cost savings for space-based OTVs. This type ofOTV could be assembled in orbit rather than on the ground so it could belarger than a ground-based unit and capable of carrying more payload.Initial studies of an OTV that would be based at the space station werecompleted in 1985.

A single-stage OTV could boost payloads of up to 7,272 kilograms tohigh-Earth or geosynchronous orbit. A multistage OTV could provide upto 36,363 kilograms to lunar orbit with 6,818.2 kilograms returned tolow-Earth orbit. After completing its delivery or servicing mission, theOTV would use its rocket engines to start a descent. Skimming throughthe thin upper atmosphere (above sixty kilometers), the OTV’s aerobrakewould slow the OTV without consuming extra propellant. Then, becauseof orbital dynamics, the OTV would navigate back to a low-Earth orbit.When the OTV reached the desired orbital altitude, its rocket engineswould again fire, circularizing its orbit until it was retrieved by the SpaceShuttle or an orbital maneuvering vehicle (OMV) dispatched from thespace station.

NASA Administrator James M. Beggs stated in June 1985 that theOTV would complement the proposed OMV. The OTV would transportpayloads from low-Earth orbit to destinations much higher than the OMVcould reach. The majority of the payloads transported by the OTV wouldbe delivered to geostationary orbit. Beggs envisioned that most OTVswould be based at the space station, where they would be maintained,fueled, and joined to payloads. In time, the OTV would also be used totransport people to geostationary orbit.

Orbital Maneuvering Vehicle

The OMV (Figure 2–17) was designed to aid satellite servicing andretrieval. This uncrewed vehicle could be characterized as a “space tug,”which would move satellites and other orbiting objects from place to


*DB Chap 2 (11-58) 1/17/02 2:51 PM Page 54

place above the Earth. A reusable,remotely operated unmanned propulsivevehicle to increase the range of the STS,the OMV was designed to be used pri-marily for spacecraft delivery, retrieval,boost, deboost, and close proximity visu-al observation beyond the operatingrange of the Space Shuttle. The vehiclewould extend the reach of the Shuttle upto approximately 2,400 kilometers.

Concept definition studies were com-pleted in 1983, and development begantoward a flight demonstration of the abil-ity to refuel propellant tanks of an orbit-ing satellite. In 1984, an in-flightdemonstration of hydrazine fuel transfertook place successfully on STS 41-G.System definition studies were complet-ed in 1985, and in June 1986, TRW wasselected by NASA for negotiations lead-ing to the award of a contract to develop

the OMV. The Preliminary Requirements Review took place in 1987, andthe Preliminary Design Review was held in 1988, with the MarshallSpace Flight Center managing the effort.

NASA planned for the OMV to be available for its first mission in1993, when it would be remotely controlled from Earth. In the early yearsof use, NASA envisioned that the OMV would be deployed from theSpace Shuttle for each short-duration mission and returned to Earth forservicing. Later, the vehicle would be left parked in orbit for extendedperiods, for use with both the Shuttle and the space station. However, theOMV was the victim of budget cuts, and the contract with TRW was can-celed in June 1990.

Tethered Satellite System

The Tethered Satellite System (TSS) program was a cooperativeeffort between the government of Italy and NASA to provide the capabil-ity to perform science in areas of space outside the reach of the SpaceShuttle. The TSS would enable scientists to conduct experiments in theupper atmosphere and ionosphere while tethered to the Space Shuttle asits operating base. The system consisted of a satellite anchored to theSpace Shuttle by a tether up to 100 kilometers long. (Tethers are long,superstrong tow lines joining orbiting objects together.)

The advanced development stage of the program was completed in1983, and management for the TSS moved to the Space Transportation

LAUNCH SYSTEMS 55

Figure 2–17. Orbital Maneuvering Vehicle

*DB Chap 2 (11-58) 1/17/02 2:51 PM Page 55

and Capability Development Division. In 1984, a study and laboratoryprogram was initiated to define and evaluate several applications of teth-ers in space. Possible applications included power generation, orbit rais-ing in the absence of propellants, artificial gravity, and space vehicleconstellations. In 1986, the Critical Design and Manufacturing Reviewswere conducted on the satellite and the deployer. In 1988, manufactureand qualification of the flight subsystems continued. The twelve-meterdeployer boom, reel motor, and on-board computer were all qualified anddelivered. Also, manufacture of the deployer structure was initiated, andthe tether control mechanisms were functionally tested. A test programwas completed for the satellite structural and engineering models. Theflight satellite structure was due for delivery in early 1989. The develop-ment of the scientific instruments continued, with delivery of flight satel-lite instruments scheduled for early 1989. The first TSS mission wasscheduled for 1991.

Advanced Launch System

The Advanced Launch System, a joint NASA-DOD effort, was a sys-tems definition and technology advanced development program aimed atdefining a new family of launchers for use after 2000, including a newheavy-lift vehicle. President Reagan signed a report to Congress inJanuary 1988 that officially created the program. Within this DOD-funded program, NASA managed the liquid engine system and advanceddevelopment efforts.

Next Manned Launch Vehicle

In 1988, attention was focused on examining various next-generationmanned launch vehicle concepts. Three possible directions were consid-ered: Space Shuttle evolution, a personnel launch system, and anadvanced manned launch system. The evolution concept referred to theoption of improving the current Shuttle design through the incorporationof upgraded technologies and capabilities. The personnel launch systemwould be a people carrier and have no capability to launch payloads intospace. The advanced manned launch system represented an innovativecrewed transportation system. Preliminary studies on all three possibili-ties progressed during 1988.

Shuttle-C

Shuttle-C (cargo) was a concept for a large, uncrewed launch vehiclethat would make maximum use of existing Space Shuttle systems with acargo canister in place of the orbiter. This proposed cargo-carrying launchvehicle would be able to lift 45,454.5 to 68,181.8 kilograms to low-Earthorbit. This payload capacity is two to three times greater than the SpaceShuttle payload capability.


*DB Chap 2 (11-58) 1/17/02 2:51 PM Page 56

In October 1987, NASA selected three contractors to perform thefirst of a two-phase systems definition study for Shuttle-C. The effortsfocused on vehicle configuration details, including the cargo element’slength and diameter, the number of liquid-fueled main engines, and anoperations concept evaluation that included ground and flight supportsystems. A major purpose of the study was to determine whether Shuttle-C would be cost effective in supporting the space station. Using Shuttle-C could free the Space Shuttle for STS-unique missions, such as solarsystem exploration, astronomy, life sciences, space station crew rotation,and logistics and materials processing experiments. Shuttle-C also wouldbe used to launch planetary missions and serve as a test bed for newShuttle boosters.

The results of the Shuttle-C efforts were to be coordinated with otherongoing advanced launch systems studies to enable a joint steering group,composed of DOD and NASA senior managers. The purpose of the steer-ing group was to formulate a national heavy-lift vehicle strategy that bestaccommodated both near-term requirements and longer term objectivesfor reducing space transportation operational costs.

Advanced Upper Stages

Advanced missions in the future would require even greater capabil-ities to move from low- to high-Earth orbit and beyond. During 1988,activity in the advanced upper stages area focused on the space transfervehicle (STV) and the possibility of upgrading the existing Centaur upperstage. The STV concept involved a cryogenic hydrogen-oxygen vehiclethat could transport payloads weighing from 909.1 to 8,636 kilogramsfrom low-Earth orbit to geosynchronous orbit or the lunar surface, as wellas for unmanned planetary missions. The STV concept could potentiallylead to a vehicle capable of supporting human exploration missions to theMoon or Mars.

Advanced Solid Rocket Motor

The Advanced Solid Rocket Motor (ASRM) was an STS improve-ment intended to replace the RSRM that was used on STS-26. The ASRMwould be based on a better design than the former rocket motor, containmore reliable safety margins, and use automated manufacturing tech-niques. The ASRM would also enhance Space Shuttle performance byoffering a potential increase of payload mass to orbit from 5454.5 kilo-grams to 9090.9 kilograms for the Shuttle. In addition, a new study on liq-uid rocket boosters was conducted that examined the feasibility ofreplacing SRMs with liquid engines.

In March 1988, NASA submitted the “Space Shuttle Advanced SolidRocket Motor Acquisition Plan” to Congress. This plan reviewed pro-curement strategy for the ASRM and discussed implementation plansand schedules. Facilities in Mississippi would be used for production and

LAUNCH SYSTEMS 57

*DB Chap 2 (11-58) 1/17/02 2:51 PM Page 57

testing of the new rocket motor. In August 1988, NASA issued an requestfor proposals to design, develop, test, and evaluate the ASRM. Contractaward was anticipated for early 1989, and the first flight using the newmotor was targeted for 1994.


*DB Chap 2 (11-58) 1/17/02 2:51 PM Page 58

LAUNCH SYSTEMS 59

Tabl

e 2–

1. A

ppro

pria

ted

Bud

get

by L

aunc

h Ve

hicl

e an

d L

aunc

h-R

elat

ed C

ompo

nent

(in

tho

usan

ds o

f do

llar

s)V

ehic

le/Y

ear

1979

Supp

. App

r.19

80Su

pp. A

ppr.

1981

1982

1983

Atla

s E

/F—

a—

——

Atla

s C

enta

urb

c5,

600

——

Del

tab

d47

,900

30,4

0042

,800

Scou

tb

e90

080

0—

Spac

e Sh

uttle

Mai

n E

ngin

e(S

SME

) D

esig

n,D

evel

opm

ent,

Test

,and

Eva

luat

ion

(DD

TE

)b

fg

h14

5,70

012

7,00

026

2,00

0SS

ME

Pro

duct

ion

bf

gh

121,

500

105,

000

—So

lid R

ocke

t Boo

ster

(SR

B)

bi

jh

14,0

0017

,000

kE

xter

nal T

ank

bl

mh

48,0

0025

,000

nST

S U

pper

Sta

ges

(ST

S O

pera

tions

Cap

abili

tyD

evel

opm

ent)

b—

o—

29,0

0075

,000

p

qU

pper

Sta

ges

Ope

ratio

ns(S

TS

Ope

ratio

ns)

b—

o—

30,9

0040

,000

p—

Orb

ital M

aneu

veri

ng V

ehic

lePr

ogra

m d

id n

ot b

egin

unt

il 19

83 w

hen

it w

as in

corp

orat

ed in

to A

dvan

ced

Prog

ram

s.r

Teth

ered

Sat

ellit

e Sy

stem

Prog

ram

did

not

beg

in u

ntil

1982

whe

n it

was

r

rin

corp

orat

ed in

to A

dvan

ced

Prog

ram

s.A

dvan

ced

Prog

ram

sb

—s

—t

8,80

011

,900

*DB Chap 2 Tables (59-104) 1/17/02 2:50 PM Page 59


Tabl

e 2–

1 co

ntin

ued

Veh

icle

/Yea

r19

8419

85Su

pp. A

ppr.

1986

1987

1988

Atla

s E

/Fu

v—

vv

—A

tlas

Cen

taur

uv

—v

v—

Del

tau

v—

vv

28,0

00Sc

out

uv

—v

v—

SSM

Ew

x—

yz

aaSR

B (

Prop

ulsi

on S

yste

m)

wx

—y

zaa

Solid

Roc

ket B

oost

er(F

light

Har

dwar

e)w

xy

zE

xter

nal T

ank

(Pro

puls

ion

Syst

em)

wx

—y

zz

Ext

erna

l Tan

k(F

light

Har

dwar

e)w

yz

aaU

pper

Sta

ges

143,

200

bb92

,400

40,0

0012

2,00

020

2,10

015

9,70

0O

rbita

l Man

euve

ring

Veh

icle

rr

—10

,000

45,0

0055

,000

Teth

ered

Sat

ellit

e Sy

stem

3,30

018

,200

—10

,000

10,6

007,

300

Adv

ance

d Pr

ogra

ms

15,0

0020

,500

—21

,000

16,6

0030

,900

aU

ndis

trib

uted

. Onl

y to

tal 1

980

R&

D a

ppro

pria

tion

spec

ifie

d:$4

,091

,086

,000

. (A

utho

riza

tion

for A

tlas

F =

$2,

000,

000.

)b

Und

istr

ibut

ed. T

otal

197

9 R

&D

app

ropr

iatio

n =

$3,

477,

200,

000.

cU

ndis

trib

uted

. Onl

y to

tal 1

980

R&

D a

ppro

pria

tion

spec

ifie

d:$4

,091

,086

,000

. (A

utho

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tion

for A

tlas

Cen

taur

= $

18,3

00,0

00.)

dU

ndis

trib

uted

. Onl

y to

tal 1

980

R&

D a

ppro

pria

tion

spec

ifie

d:$4

,091

,086

,000

. (A

utho

riza

tion

for

Del

ta =

$43

,100

,000

.)e

Und

istr

ibut

ed. O

nly

tota

l 198

0 R

&D

app

ropr

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n sp

ecif

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$4,0

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00. (

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300,

000

fSu

pple

men

tal a

ppro

pria

tion

spec

ifie

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r ov

eral

l R&

D a

ctiv

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= $

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000,

000.

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utho

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for

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E =

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,000

,000

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Und

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nly

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&D

app

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$4,0

91,0

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00. (

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r SS

ME

DD

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0,60

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hU

ndis

trib

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al S

pace

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l app

ropr

iatio

n =

$28

5,00

0,00

0. N

o sp

ecif

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huttl

e el

emen

ts li

sted

.


LAUNCH SYSTEMS 61

Tabl

e 2–

1 co

ntin

ued

iSu

pple

men

tal a

ppro

pria

tion

spec

ifie

d fo

r ov

eral

l R&

D a

ctiv

ities

= $

185,

000,

000.

(A

utho

riza

tion

for

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= $

36,7

00,0

00.)

jU

ndis

trib

uted

. Onl

y to

tal 1

980

R&

D a

ppro

pria

tion

spec

ifie

d:$4

,091

,086

,000

. (A

utho

riza

tion

for

SRB

= $

57,5

00,0

00.)

kN

o bu

dget

item

list

ed. S

uppo

rtin

g co

mm

ittee

doc

umen

tatio

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clud

es S

RB

in S

pace

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ry w

ith n

o am

ount

spe

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otal

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$1,

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spec

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r ov

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l R&

D a

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= $

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000,

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utho

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for

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tal 1

980

R&

D a

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utho

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dget

item

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he S

huttl

e Pr

oduc

tion

and

Cap

abili

ty D

evel

opm

ent /

Prop

ulsi

on S

yste

ms

“pro

vide

s fo

r th

e pr

oduc

tion

of th

e Sp

ace

Shut

tle’s

mai

n en

gine

s an

d th

e de

velo

pmen

t of

the

capa

bilit

y to

sup

port

ope

ratio

nal r

equi

rem

ents

est

ablis

hed

for

the

mai

n en

gine

,sol

id r

ocke

t boo

ster

,and

ext

erna

l tan

k.”

Con

gres

sion

al d

ocum

ents

als

o st

ate

that

the

cate

gory

incl

udes

con

tinui

ng “

capa

bilit

y de

velo

pmen

t tas

ks f

or th

e or

bite

r,m

ain

engi

ne,e

xter

nal t

ank,

and

SRB

,. .

.”an

d “t

he d

evel

opm

ent o

f th

e fi

lam

ent w

ound

cas

e(F

WC

) SR

B.”

Som

e la

unch

sys

tem

–rel

ated

app

ropr

iate

d fu

ndin

g is

incl

uded

in th

e Fl

ight

Har

dwar

e/Sh

uttle

Ope

ratio

ns c

ateg

ory

(als

o in

SFC

&D

C)

undi

stri

bute

d,in

clud

-ed

in S

huttl

e O

pera

tions

app

ropr

iatio

n =

$1,

520,

600,

000.

NA

SA B

udge

t Est

imat

e do

cum

ents

indi

cate

est

imat

ed a

mou

nt o

f $3

36,2

00,0

00 f

or e

xter

nal t

ank

and

$353

,200

,000

for

SR

B u

nder

Flig

ht H

ardw

are/

Shut

tle O

pera

tions

cat

egor

y.



Tabl

e 2–

1 co

ntin

ued

xPr

oduc

tion

and

resi

dual

dev

elop

men

t tas

ks f

or th

e or

bite

r,SS

ME

,ext

erna

l tan

k,an

d SR

B f

all u

nder

Spa

ce P

rodu

ctio

n an

d O

pera

tiona

l Cap

abili

ty,P

ropu

lsio

n Sy

stem

s.SR

Bs

and

exte

rnal

tank

pro

cure

men

t (pr

oduc

tion)

fal

ls u

nder

Spa

ce T

rans

port

atio

n O

pera

tions

,Flig

ht H

ardw

are.

No

brea

kdow

n is

pro

vide

d fo

r in

divi

dual

Spa

ce S

huttl

epr

opul

sion

com

pone

nts.

The

198

5 ap

prop

riat

ion

for

Prop

ulsi

on S

yste

ms

= $

599,

000,

000;

Flig

ht H

ardw

are

appr

opri

atio

n =

$75

8,00

0,00

0.y

No

brea

kdow

n fo

r in

divi

dual

Spa

ce S

huttl

e pr

opul

sion

com

pone

nts.

The

198

6 ap

prop

riat

ion

for

Prop

ulsi

on S

yste

ms

= $

454,

000,

000;

no

Flig

ht H

ardw

are

appr

opri

atio

nin

198

6.z

No

brea

kdow

n fo

r in

divi

dual

Spa

ce S

huttl

e pr

opul

sion

com

pone

nts.

The

198

7 ap

prop

riat

ion

for

Prop

ulsi

on S

yste

ms

= $

338,

400,

000;

app

ropr

iatio

n fo

r Fl

ight

Har

dwar

e=

$64

6,20

0,00

0.aa

No

brea

kdow

n fo

r in

divi

dual

Spa

ce S

huttl

e pr

opul

sion

com

pone

nts.

The

198

8 ap

prop

riat

ion

for

Prop

ulsi

on S

yste

ms

= $

249,

300,

000;

app

ropr

iatio

n fo

r Fl

ight

Har

dwar

e=

$92

3,10

0,00

0.bb

Incl

udes

fun

ding

for

mod

ific

atio

n of

the

Cen

taur

for

use

in th

e Sh

uttle

.So

urce

:N

ASA

Chr

onol

ogic

al H

isto

ry F

isca

l Yea

rs 1

979–

1983

Bud

get

Subm

issi

ons.


LAUNCH SYSTEMS 63

Tabl

e 2–

2. A

tlas

E/F

Fun

ding

His

tory

(in

tho

usan

ds o

f do

llar

s)Y

ear

(Fis

cal)

Subm

issi

onA

utho

riza

tion

App

ropr

iati

onP

rogr

amm

ed

(Act

ual)

1980

2,00

0a2,

000

b1,

200

1981

No

dire

ct f

unds

aut

hori

zed

or a

ppro

pria

ted;

no

prop

osed

use

of A

tlas

E/F

aft

er 1

980

by N

ASA

—19

82N

o bu

dget

line

item

—19

83R

eim

burs

able

onl

y—

1984

N

o bu

dget

line

item

for

spe

cifi

c E

LVs

c—

1985

dT

here

wer

e no

dir

ect a

ppro

pria

ted

fund

req

uire

men

ts f

or th

e—

1986

eE

LV p

rogr

am. D

OD

and

NO

AA

con

tinue

d to

use

the

Del

ta,

—19

87 f

Scou

t,A

tlas,

and

Atla

s—C

enta

ur E

LVs

on a

ful

ly r

eim

burs

able

bas

is.

—19

88A

tlas

E/F

not

in u

se b

y N

ASA

—a

Atla

s F

only

.b

Und

istr

ibut

ed. I

nclu

ded

in 1

980

R&

D a

ppro

pria

tion

of $

4,09

1,08

6,00

0.c

No

budg

et s

ubm

issi

on,a

utho

riza

tion,

or a

ppro

pria

tion

for

spec

ific

ELV

s. T

otal

und

istr

ibut

ed E

LV s

ubm

issi

on =

$50

,000

,000

; aut

hori

zatio

n =

$50

,000

,000

; and

app

ropr

i-at

ion

= $

50,0

00,0

00. E

LV a

ppro

pria

tion

rem

oved

fro

m R

&D

and

pla

ced

in S

pace

Flig

ht,C

ontr

ol &

Dat

a C

omm

unic

atio

ns (

SFC

&D

C)

(Off

ice

of S

pace

Tra

nspo

rtat

ion

Syst

ems)

app

ropr

iatio

n.d

No

budg

et li

ne it

em f

or E

LVs.

Sup

port

for

ELV

s pa

id f

or a

s pa

rt o

f Sp

ace

Tra

nspo

rtat

ion

Ope

ratio

ns P

rogr

am. B

udge

t dat

a fo

r Sp

ace

Tra

nspo

rtat

ion

Ope

ratio

ns P

rogr

amfo

und

in C

hapt

er 3

bud

get t

able

s.

eN

o bu

dget

line

item

for

ELV

s. S

uppo

rt f

or E

LVs

paid

for

as

part

of

Spac

e T

rans

port

atio

n O

pera

tions

Pro

gram

. Bud

get d

ata

for

Spac

e T

rans

port

atio

n O

pera

tions

Pro

gram

foun

d in

Cha

pter

3 b

udge

t tab

les.

fIn

clud

ed in

Flig

ht H

ardw

are

cate

gory

. Bud

get d

ata

for

Flig

ht H

ardw

are

foun

d in

Cha

pter

3 b

udge

t tab

les.



Tabl

e 2–

3. A

tlas

-Cen

taur

Fun

ding

His

tory

(in

tho

usan

ds o

f do

llar

s)Y

ear

(Fis

cal)

Subm

issi

onA

utho

riza

tion

App

ropr

iati

onP

rogr

amm

ed

(Act

ual)

1979

21,5

00a

b17

,320

1980

18,3

0018

,300

c18

,000

1981

5,60

05,

600

5,60

05,

600

d19

82R

eim

burs

able

onl

y—

1983

Rei

mbu

rsab

le o

nly

—19

84N

o bu

dget

line

item

for

spe

cifi

c E

LVs

e—

1985

fT

here

wer

e no

dir

ect a

ppro

pria

ted

fund

req

uire

men

ts f

or th

e,—

1986

gE

LV p

rogr

am. D

OD

and

NO

AA

con

tinue

d to

use

the

Del

ta—

1987

hSc

out,

Atla

s,an

d A

tlas-

Cen

taur

ELV

s on

a f

ully

rei

mbu

rsab

le b

asis

.—

1988

Atla

s-C

enta

ur n

ot in

use

by

NA

SA—

aN

ot d

istr

ibut

ed b

y ve

hicl

e—to

tal 1

979

ELV

aut

hori

zatio

n =

$74

,000

,000

.b

Not

dis

trib

uted

by

vehi

cle—

1979

R&

D a

ppro

pria

tion

= $

3,47

7,20

0,00

0.c

Und

istr

ibut

ed. I

nclu

ded

in R

&D

app

ropr

iatio

n of

$4,

091,

086,

000.

dB

ased

on

antic

ipat

ed c

lose

out o

f th

e N

ASA

pro

gram

by

the

end

of 1

981.

eN

o bu

dget

sub

mis

sion

,aut

hori

zatio

n,or

app

ropr

iatio

n fo

r sp

ecif

ic E

LVs.

Tot

al u

ndis

trib

uted

ELV

sub

mis

sion

= $

50,0

00,0

00; a

utho

riza

tion

= $

50,0

00,0

00; a

nd a

ppro

pri-

atio

n =

$50

,000

,000

. ELV

app

ropr

iatio

n re

mov

ed f

rom

R&

D a

nd p

lace

d in

Spa

ce F

light

,Con

trol

& D

ata

Com

mun

icat

ions

(SF

C&

DC

) (O

ffic

e of

Spa

ce T

rans

port

atio

nSy

stem

s) a

ppro

pria

tion.

fN

o bu

dget

line

item

for

ELV

s. S

uppo

rt f

or E

LVs

paid

for

as

part

of

Spac

e T

rans

port

atio

n O

pera

tions

Pro

gram

. Bud

get d

ata

for

Spac

e T

rans

port

atio

n O

pera

tions

Pro

gram

foun

d in

Cha

pter

3 b

udge

t tab

les.

gN

o bu

dget

line

item

for

ELV

s. S

uppo

rt f

or E

LVs

paid

for

as

part

of

Spac

e T

rans

port

atio

n O

pera

tions

Pro

gram

. Bud

get d

ata

for

Spac

e T

rans

port

atio

n O

pera

tions

Pro

gram

foun

d in

Cha

pter

3 b

udge

t tab

les.

hIn

clud

ed in

Flig

ht H

ardw

are

cate

gory

. Bud

get d

ata

for

Flig

ht H

ardw

are

foun

d in

Cha

pter

3 b

udge

t tab

les.


LAUNCH SYSTEMS 65

Tabl

e 2–

4. D

elta

Fun

ding

His

tory

(in

tho

usan

ds o

f do

llar

s)Y

ear

(Fis

cal)

Subm

issi

onA

utho

riza

tion

App

ropr

iati

onP

rogr

amm

ed

(Act

ual)

1979

38,6

00a

b45

,680

1980

43,1

0043

,100

c43

,100

1981

47,9

0047

,900

47,9

0047

,900

1982

30,4

0030

,400

30,4

0030

,400

1983

42,8

0042

,800

42,8

0083

,000

1984

No

budg

et li

ne it

em f

or s

peci

fic

ELV

s d

50,0

00e

1985

fN

o bu

dget

line

item

for

ELV

s—

1986

gN

o bu

dget

line

item

for

ELV

s—

1987

hN

o bu

dget

line

item

for

ELV

s—

1988

28,0

00i

60,0

00i

28,0

00i

28,0

00i

aN

ot d

istr

ibut

ed b

y ve

hicl

e—to

tal 1

979

ELV

aut

hori

zatio

n =

$74

,000

,000

.b

Not

dis

trib

uted

by

vehi

cle—

1979

R&

D a

ppro

pria

tion

= $

3,47

7,20

0,00

0.c

Und

istr

ibut

ed. I

nclu

ded

in R

&D

app

ropr

iatio

n of

$4,

091,

086,

000.

dN

o bu

dget

sub

mis

sion

,aut

hori

zatio

n,or

app

ropr

iatio

n fo

r sp

ecif

ic E

LVs.

Tot

al u

ndis

trib

uted

ELV

sub

mis

sion

= $

50,0

00,0

00; a

utho

riza

tion

= $

50,0

00,0

00; a

nd a

ppro

pri-

atio

n =

$50

,000

,000

. ELV

app

ropr

iatio

n re

mov

ed f

rom

R&

D a

nd p

lace

d in

SFC

&D

C (

Off

ice

of S

pace

Tra

nspo

rtat

ion

Syst

ems)

app

ropr

iatio

n. C

ongr

essi

onal

sup

port

ing

docu

men

tatio

n in

dica

tes

that

$50

,000

,000

is f

or “

cont

inue

d pr

ocur

emen

t of

the

Del

ta E

LVs

in F

Y 1

984.

”e

NA

SA b

udge

t sum

mar

y da

ta d

o no

t spe

cifi

cally

indi

cate

that

pro

gram

med

am

ount

was

for

the

Del

ta. H

owev

er,t

he n

arra

tive

that

acc

ompa

nies

con

gres

sion

al c

omm

ittee

repo

rts

desc

ribe

s pr

ogra

ms

that

use

the

Del

ta a

s th

e la

unch

veh

icle

.f

No

budg

et li

ne it

em f

or E

LVs.

Sup

port

for

ELV

s pa

id f

or a

s pa

rt o

f Sp

ace

Tra

nspo

rtat

ion

Ope

ratio

ns P

rogr

am. I

t was

ant

icip

ated

that

the

NA

SA E

LV p

rogr

am w

ould

be

com

plet

ely

fund

ed o

n a

reim

burs

able

bas

is in

198

5.g

No

budg

et li

ne it

em f

or E

LVs.

Sup

port

for

ELV

s pa

id f

or a

s pa

rt o

f Sp

ace

Tra

nspo

rtat

ion

Ope

ratio

ns P

rogr

am. I

t was

ant

icip

ated

that

the

NA

SA E

LV p

rogr

am w

ould

be

com

plet

ely

fund

ed o

n a

reim

burs

able

bas

is in

198

6.h

Incl

uded

in F

light

Har

dwar

e ca

tego

ry. I

t was

ant

icip

ated

that

the

NA

SA E

LV p

rogr

am w

ould

be

com

plet

ely

fund

ed o

n a

reim

burs

able

bas

is in

198

7.i

Veh

icle

not

spe

cifi

ed in

bud

get f

igur

es b

ut in

dica

ted

in s

uppo

rtin

g co

ngre

ssio

nal c

omm

ittee

doc

umen

tatio

n,w

hich

spe

cifi

es tw

o D

elta

II

vehi

cles

for

199

0 an

d 19

91la

unch

es.


Tabl

e 2–

5. S

cout

Fun

ding

His

tory

(in

tho

usan

ds o

f do

llar

s)Y

ear

(Fis

cal)

Subm

issi

onA

utho

riza

tion

App

ropr

iati

onP

rogr

amm

ed

(Act

ual)

1979

16,4

00a

b10

,600

1980

7,30

07,

300

c5,

100

1981

2,20

02,

200

900

d90

019

8280

080

080

080

019

83N

o bu

dget

line

item

(Sc

out n

ot in

use

by

NA

SA a

fter

198

2)—

aN

ot d

istr

ibut

ed b

y ve

hicl

e—to

tal 1

979

ELV

aut

hori

zatio

n =

$74

,000

,000

.b

Not

dis

trib

uted

by

vehi

cle—

1979

R&

D a

ppro

pria

tion

= $

3,47

7,20

0,00

0.c

Und

istr

ibut

ed. I

nclu

ded

in R

&D

app

ropr

iatio

n of

$4,

091,

086,

000.

dB

asic

app

ropr

iatio

n of

$2,

200,

000.

Eff

ect o

f G

ener

al P

rovi

sion

,Sec

. 412

(Pu

blic

Law

96–

526)

,red

uced

fun

ding

leve

l to

$900

,000

.e

No

budg

et s

ubm

issi

on,a

utho

riza

tion,

or a

ppro

pria

tion

for

spec

ific

ELV

s. T

otal

und

istr

ibut

ed E

LV s

ubm

issi

on =

$50

,000

,000

; aut

hori

zatio

n =

$50

,000

,000

; and

app

ropr

i-at

ion

= $

50,0

00,0

00. E

LV a

ppro

pria

tion

rem

oved

fro

m R

&D

and

pla

ced

in S

pace

Flig

ht,C

ontr

ol &

Dat

a C

omm

unic

atio

ns (

SFC

&D

C)

(Off

ice

of S

pace

Tra

nspo

rtat

ion

Syst

ems)

app

ropr

iatio

n.f

No

budg

et li

ne it

em f

or E

LVs.

Sup

port

for

ELV

s pa

id f

or a

s pa

rt o

f Sp

ace

Tra

nspo

rtat

ion

Ope

ratio

ns P

rogr

am. I

t was

ant

icip

ated

that

the

NA

SA E

LV p

rogr

am w

ould

be

com

plet

ely

fund

ed o

n a

reim

burs

able

bas

is in

198

5.g

No

budg

et li

ne it

em f

or E

LVs.

Sup

port

for

ELV

s pa

id f

or a

s pa

rt o

f Sp

ace

Tra

nspo

rtat

ion

Ope

ratio

ns P

rogr

am. I

t was

ant

icip

ated

that

the

NA

SA E

LV p

rogr

am w

ould

be

com

plet

ely

fund

ed o

n a

reim

burs

able

bas

is in

198

6.h

Incl

uded

in F

light

Har

dwar

e ca

tego

ry. I

t was

ant

icip

ated

that

the

NA

SA E

LV p

rogr

am w

ould

be

com

plet

ely

fund

ed o

n a

reim

burs

able

bas

is in

198

7.



Tabl

e 2–

6. S

pace

Shu

ttle

Mai

n E

ngin

e F

undi

ng H

isto

ry (

in t

hous

ands

of

doll

ars)

Yea

r (F

isca

l)Su

bmis

sion

Aut

hori

zati

onA

ppro

pria

tion

Pro

gram

med

(A

ctua

l)19

79 D

DT

&E

176,

700

176,

700

a17

2,70

0Pr

oduc

tion

18,0

00b

a26

4,50

0Su

ppl.

App

ropr

iatio

nc

48,0

00d

e19

80 D

DT

&E

140,

600

140,

600

f14

0,60

0Pr

oduc

tion

109,

900

109,

900

f12

3,60

0Su

ppl.

App

ropr

iatio

n g

1981

DD

T&

E14

5,70

014

5,70

014

5,70

013

4,00

0Pr

oduc

tion

121,

500

121,

500

121,

500

779,

000

1982

DD

T&

E12

7,00

012

7,00

012

7,00

0h

Prod

uctio

n10

5,00

010

5,00

010

5,00

016

3,30

019

8326

2,00

026

2,00

026

2,00

035

5,70

019

84i

ii

418,

100

1985

ji

i41

9,00

019

86k

kk

394,

400

1987

ll

l43

2,70

019

88m

mm

395,

900

aN

ot d

istr

ibut

ed b

y el

emen

t/veh

icle

—19

79 R

&D

app

ropr

iatio

n =

$3,

477,

200,

000.

bN

o SS

ME

Pro

duct

ion

cate

gory

bro

ken

out.

Tota

l Pro

duct

ion

amou

nt =

$45

8,00

0,00

0.c

No

brea

kout

of

Supp

lem

enta

l App

ropr

iatio

n su

bmis

sion

; inc

lude

d in

gen

eral

R&

D s

uppl

emen

tal a

ppro

pria

tion

subm

issi

on.

dB

reak

dow

n of

sup

plem

enta

l aut

hori

zatio

n no

t pro

vide

d in

bud

get r

eque

st o

r pu

blic

law

. Bre

akdo

wn

prov

ided

in s

uppo

rtin

g do

cum

enta

tion

for

auth

oriz

atio

n on

ly.

eSu

pple

men

tal a

ppro

pria

tion

spec

ifie

d fo

r ov

eral

l R&

D a

ctiv

ities

= $

185,

000,

000.

fU

ndis

trib

uted

. Inc

lude

d in

R&

D a

ppro

pria

tion

of $

4,09

1,08

6,00

0.g

Supp

lem

enta

l app

ropr

iatio

n fo

r Sp

ace

Shut

tle in

res

pons

e to

am

ende

d N

ASA

bud

get s

ubm

issi

on o

f $3

00,0

00,0

00. N

o au

thor

izat

ion

activ

ity. S

uppl

emen

tal a

ppro

pria

tion

of $

285,

000,

000

appr

oved

with

no

dist

ribu

tion

to in

divi

dual

com

pone

nts.

hPr

ogra

mm

ed a

mou

nt f

or S

SME

DD

T&

E in

198

2 no

t ind

icat

ed.

LAUNCH SYSTEMS 67



Tabl

e 2–

6 co

ntin

ued

iN

o sp

ecif

ic a

utho

riza

tion

for

SSM

E,e

xter

nal t

ank,

or S

RB

. Acc

ordi

ng to

con

gres

sion

al r

epor

ts,t

he S

pace

Tra

nspo

rtat

ion

and

Cap

abili

ty D

evel

opm

ent p

rogr

am s

uppo

rt-

ed th

e pr

oduc

tion

of th

e SS

ME

,SR

B,a

nd e

xter

nal t

ank,

in a

dditi

on to

pro

vidi

ng f

or c

ritic

al s

pare

s (a

s w

ell a

s ot

her

item

s). T

he to

tal a

utho

riza

tion

for

this

cat

egor

y =

$2,0

09,4

00,0

00. C

ateg

ory

also

incl

uded

con

tinui

ng “

capa

bilit

y de

velo

pmen

t tas

ks f

or th

e or

bite

r,m

ain

engi

ne,e

xter

nal t

ank,

and

SR

B,.

. .”

and

“the

dev

elop

men

t of

the

fila

men

t wou

nd c

ase

(FW

C)

SRB

.”A

ppro

pria

tion

for

Spac

e Sh

uttle

act

iviti

es o

f $1

,545

,000

,000

mov

ed f

rom

R&

D to

SFC

&D

C. S

ome

Spac

e Sh

uttle

fun

ding

was

incl

uded

in th

e Fl

ight

Har

dwar

e ca

tego

ry:s

ubm

issi

on =

$84

8,40

0,00

0; a

utho

riza

tion

undi

stri

bute

d,in

clud

ed in

Shu

ttle

Ope

ratio

ns a

utho

riza

tion

= $

1,49

5,60

0,00

0; a

ndap

prop

riat

ion

(mov

ed to

SFC

&D

C)

undi

stri

bute

d,in

clud

ed in

Shu

ttle

Ope

ratio

ns a

ppro

pria

tion

= $

1,52

0,60

0,00

0. A

mou

nt o

f $4

27,4

00,0

00 r

emai

ned

in R

&D

for

oth

erac

tiviti

es.

jSS

ME

pro

duct

ion

and

resi

dual

dev

elop

men

t tas

ks f

or th

e or

bite

r,SS

ME

,ext

erna

l tan

k,an

d SR

B f

ell u

nder

Spa

ce P

rodu

ctio

n an

d O

pera

tiona

l Cap

abili

ty,P

ropu

lsio

nSy

stem

s. S

RB

s an

d ex

tern

al ta

nk p

rocu

rem

ent (

prod

uctio

n) f

ell u

nder

Spa

ce T

rans

port

atio

n O

pera

tions

,Flig

ht H

ardw

are.

No

brea

kdow

n fo

r in

divi

dual

Spa

ce S

huttl

epr

opul

sion

com

pone

nts.

The

198

5 am

ount

s fo

r Pr

opul

sion

Sys

tem

s w

ere:

subm

issi

on =

$59

9,00

0,00

0; a

utho

riza

tion

= $

599,

000,

000;

and

app

ropr

iatio

n =

$59

9,00

0,00

0.Fl

ight

Har

dwar

e su

bmis

sion

= $

758,

000,

000;

aut

hori

zatio

n =

$75

8,00

0,00

0; a

nd a

ppro

pria

tion

= $

758,

000,

000.

kN

o br

eakd

own

for

indi

vidu

al S

pace

Shu

ttle

prop

ulsi

on c

ompo

nent

s. T

he 1

986

amou

nts

for

Prop

ulsi

on S

yste

ms

wer

e:su

bmis

sion

= $

454,

000,

000;

aut

hori

zatio

n =

$454

,000

,000

; and

app

ropr

iatio

n =

$45

4,00

0,00

0. N

o Fl

ight

Har

dwar

e bu

dget

cat

egor

y in

198

6.l

No

brea

kdow

n fo

r in

divi

dual

Spa

ce S

huttl

e pr

opul

sion

com

pone

nts.

The

198

7 am

ount

s fo

r Pr

opul

sion

Sys

tem

s w

ere:

subm

issi

on =

$33

8,40

0,00

0; a

utho

riza

tion

=$3

38,4

00,0

00; a

nd a

ppro

pria

tion

= $

338,

400,

000.

Flig

ht H

ardw

are

subm

issi

on =

$64

6,20

0,00

0; a

utho

riza

tion

= $

879,

100,

000;

and

app

ropr

iatio

n =

$64

6,20

0,00

0.m

No

brea

kdow

n fo

r in

divi

dual

Spa

ce S

huttl

e pr

opul

sion

com

pone

nts.

The

198

8 am

ount

s fo

r Pr

opul

sion

Sys

tem

s w

ere:

subm

issi

on =

$55

2,10

0,00

0; a

utho

riza

tion

=$5

52,1

00,0

00; a

nd a

ppro

pria

tion

= $

249,

300,

000.

Flig

ht H

ardw

are

subm

issi

on =

$92

3,10

0,00

0; a

utho

riza

tion

= $

923,

100,

000;

and

app

ropr

iatio

n =

$92

3,10

0,00

0.


Tabl

e 2–

7. S

olid

Roc

ket

Boo

ster

s F

undi

ng H

isto

ry (

in t

hous

ands

of

doll

ars)

Yea

r (F

isca

l)Su

bmis

sion

Aut

hori

zati

onA

ppro

pria

tion

Pro

gram

med

(A

ctua

l)19

7963

,500

63,5

00a

115,

400

Supp

l. A

ppro

pria

tion

b36

,700

cd

1980

57,5

0057

,500

e65

,200

Supp

l. A

ppro

pria

tion

f—

——

1981

14,0

0014

,000

14,0

0050

,500

1982

Pro

puls

ion

Syst

ems

g17

,000

17,0

0017

,000

22,0

00Fl

ight

Har

dwar

e g

156,

200

1983

Pro

puls

ion

Syst

ems

gh

hh

102,

300

Flig

ht H

ardw

are

g30

9,20

019

84 P

ropu

lsio

n Sy

stem

si

ii

140,

500

Flig

ht H

ardw

are

ii

i34

1,20

019

85 P

ropu

lsio

n Sy

stem

sj

jj

105,

100

Flig

ht H

ardw

are

jj

j29

8,60

019

86 P

ropu

lsio

n Sy

stem

sk

kk

328,

500

Flig

ht H

ardw

are

kk

k33

5,00

019

87 P

ropu

lsio

n Sy

stem

sl

ll

322,

100

Flig

ht H

ardw

are

ll

l14

4,30

019

88 P

ropu

lsio

n Sy

stem

sm

mm

161,

200

Flig

ht H

ardw

are

mm

m20

0,50

0a

Not

dis

trib

uted

by

elem

ent/v

ehic

le—

1979

R&

D a

ppro

pria

tion

= $

3,47

7,20

0,00

0.b

No

brea

kout

of

Supp

lem

enta

l App

ropr

iatio

n su

bmis

sion

; inc

lude

d in

gen

eral

R&

D r

eque

st o

f $1

85,0

00,0

00.

cB

reak

dow

n of

sup

plem

enta

l aut

hori

zatio

n no

t pro

vide

d in

bud

get r

eque

st o

r pu

blic

law

. Bre

akdo

wn

prov

ided

in s

uppo

rtin

g do

cum

enta

tion

for

auth

oriz

atio

n on

ly.

dSu

pple

men

tal A

ppro

pria

tion

for

gene

ral R

&D

act

iviti

es.

eU

ndis

trib

uted

. Inc

lude

d in

198

0 R

&D

app

ropr

iatio

n of

$4,

091,

086,

000.

LAUNCH SYSTEMS 69



Tabl

e 2–

7 co

ntin

ued

fSu

pple

men

tal a

ppro

pria

tion

for

Spac

e Sh

uttle

in r

espo

nse

to a

men

ded

NA

SA b

udge

t sub

mis

sion

of

$300

,000

,000

. No

auth

oriz

atio

n ac

tivity

. Sup

plem

enta

l app

ropr

iatio

nof

$28

5,00

0,00

0 ap

prov

ed w

ith n

o di

stri

butio

n to

indi

vidu

al c

ompo

nent

s.g

Prop

ulsi

on S

yste

ms

and

Flig

ht H

ardw

are

budg

et c

ateg

orie

s w

ere

not u

sed

in N

ASA

’s b

udge

t pri

or to

198

4. H

owev

er,p

rogr

amm

ed a

mou

nts

used

thes

e ca

tego

ries

to b

eco

nsis

tent

with

cat

egor

ies

used

in e

stim

ates

for

the

futu

re y

ears

.h

No

budg

et it

em li

sted

. Sup

port

ing

com

mitt

ee d

ocum

enta

tion

incl

uded

SR

B in

Spa

ce S

huttl

e Pr

oduc

tion

cate

gory

with

no

amou

nt s

peci

fied

. Tot

al P

rodu

ctio

n am

ount

:su

bmis

sion

= $

1,58

5,50

0,00

0; a

utho

riza

tion

= $

1,67

0,50

0,00

0; a

nd a

ppro

pria

tion

= $

1,63

6,60

0,00

0.i

No

spec

ific

aut

hori

zatio

n fo

r SS

ME

,ext

erna

l tan

k,or

SR

B. A

ccor

ding

to c

ongr

essi

onal

rep

orts

,the

Spa

ce T

rans

port

atio

n an

d C

apab

ility

Dev

elop

men

t pro

gram

sup

port

-ed

the

prod

uctio

n of

the

SSM

E,S

RB

,and

ext

erna

l tan

k,in

add

ition

to p

rovi

ding

for

cri

tical

spa

res

(as

wel

l as

othe

r ite

ms)

. The

tota

l aut

hori

zatio

n fo

r th

is c

ateg

ory

=$2

,009

,400

,000

. Cat

egor

y al

so in

clud

ed c

ontin

uing

“ca

pabi

lity

deve

lopm

ent t

asks

for

the

orbi

ter,

mai

n en

gine

,ext

erna

l tan

k,an

d S

RB

,. .

.”an

d “t

he d

evel

opm

ent o

f th

efi

lam

ent w

ound

cas

e (F

WC

) SR

B.”

App

ropr

iatio

n m

oved

fro

m R

&D

to S

FC&

DC

= $

1,54

5,00

0,00

0. S

ome

Spac

e Sh

uttle

fun

ding

was

incl

uded

in th

e Fl

ight

Har

dwar

eca

tego

ry (

see

abov

e fo

r de

fini

tion)

:sub

mis

sion

= $

848,

400,

000;

aut

hori

zatio

n un

dist

ribu

ted,

incl

uded

in S

huttl

e O

pera

tions

aut

hori

zatio

n of

$1,

495,

600,

000;

and

app

ro-

pria

tion

(mov

ed to

SFC

&D

C)

undi

stri

bute

d,in

clud

ed in

Shu

ttle

Ope

ratio

ns a

ppro

pria

tion

= $

1,52

0,60

0,00

0. A

ppro

pria

tion

mov

ed f

rom

R&

D to

SFC

&D

C =

$1,5

45,0

00,0

00.

jSS

ME

pro

duct

ion

and

resi

dual

dev

elop

men

t tas

ks f

or th

e or

bite

r,SS

ME

,ext

erna

l tan

k,an

d SR

B f

ell u

nder

Spa

ce P

rodu

ctio

n an

d O

pera

tiona

l Cap

abili

ty,P

ropu

lsio

nSy

stem

s. S

RB

and

ext

erna

l tan

k pr

ocur

emen

t (pr

oduc

tion)

fel

l und

er S

pace

Tra

nspo

rtat

ion

Ope

ratio

ns,F

light

Har

dwar

e. N

o br

eakd

own

for

indi

vidu

al S

pace

Shu

ttle

prop

ulsi

on c

ompo

nent

s. T

he 1

985

amou

nt f

or P

ropu

lsio

n Sy

stem

s w

as:s

ubm

issi

on =

$59

9,00

0,00

0; a

utho

riza

tion

= $

599,

000,

000;

and

app

ropr

iatio

n =

$59

9,00

0,00

0.A

utho

riza

tion

for

subm

issi

on =

$75

8,00

0,00

0. P

rocu

rem

ent o

f ex

tern

al ta

nk,s

olid

roc

ket m

otor

,and

SR

B h

ardw

are

incl

uded

in S

pace

Tra

nspo

rtat

ion

Ope

ratio

nsPr

ogra

m,F

light

Har

dwar

e am

ount

of

$758

,000

,000

; app

ropr

iatio

n =

$75

8,00

0,00

0. S

SME

pro

duct

ion

and

resi

dual

dev

elop

men

t tas

ks f

or th

e or

bite

r,SS

ME

,ext

erna

lta

nk,a

nd S

RB

fal

l und

er S

pace

Pro

duct

ion

and

Ope

ratio

nal C

apab

ility

,Pro

puls

ion

Syst

ems.

SR

B a

nd e

xter

nal t

ank

proc

urem

ent (

prod

uctio

n) f

ell u

nder

Spa

ceT

rans

port

atio

n O

pera

tions

,Flig

ht H

ardw

are.

Flig

ht H

ardw

are

subm

issi

on =

$75

8,00

0,00

0; a

utho

riza

tion

= $

758,

000,

000;

and

app

ropr

iatio

n =

$75

8,00

0,00

0.k

No

brea

kdow

n fo

r in

divi

dual

Spa

ce S

huttl

e pr

opul

sion

com

pone

nts.

The

198

6 am

ount

s fo

r Pr

opul

sion

Sys

tem

s w

ere:

subm

issi

on =

$45

4,00

0,00

0; a

utho

riza

tion

=$4

54,0

00,0

00; a

nd a

ppro

pria

tion

= $

454,

000,

000.

No

Flig

ht H

ardw

are

budg

et c

ateg

ory

in 1

986.

lN

o br

eakd

own

for

indi

vidu

al S

pace

Shu

ttle

prop

ulsi

on c

ompo

nent

s. T

he 1

987

amou

nts

for

Prop

ulsi

on S

yste

ms

wer

e:su

bmis

sion

= $

338,

400,

000;

aut

hori

zatio

n =

$338

,400

,000

; and

app

ropr

iatio

n =

$33

8,40

0,00

0. F

light

Har

dwar

e su

bmis

sion

= $

646,

200,

000;

aut

hori

zatio

n =

$87

9,10

0,00

0; a

nd a

ppro

pria

tion

= $

646,

200,

000.

mN

o br

eakd

own

for

indi

vidu

al S

pace

Shu

ttle

prop

ulsi

on c

ompo

nent

s. T

he 1

988

amou

nts

for

Prop

ulsi

on S

yste

ms

wer

e:su

bmis

sion

= $

552,

100,

000;

aut

hori

zatio

n =

$552

,100

,000

; and

app

ropr

iatio

n =

$24

9,30

0,00

0. F

unds

del

eted

fro

m P

ropu

lsio

n Sy

stem

s; $

302,

800,

000

appr

opri

ated

mov

ed to

Lau

nch

and

Mis

sion

Sup

port

cat

egor

y.Fl

ight

Har

dwar

e su

bmis

sion

= $

923,

100,

000;

aut

hori

zatio

n =

$92

3,10

0,00

0; a

nd a

ppro

pria

tion

= $

923,

100,

000.


Tabl

e 2–

8. E

xter

nal T

ank

Fun

ding

His

tory

(in

tho

usan

ds o

f do

llar

s)Y

ear

(Fis

cal)

Subm

issi

onA

utho

riza

tion

App

ropr

iati

onP

rogr

amm

ed

(Act

ual)

1979

80,5

0080

,500

a10

4,80

0Su

ppl.

App

ropr

iatio

nb

27,1

00c

1980

68,4

0068

,400

d79

,400

Supp

l. A

ppro

pria

tion

e—

——

1981

48,0

0048

,000

48,0

0063

,500

1982

Pro

puls

ion

Syst

ems

f25

,000

25,0

0025

,000

45,7

00Fl

ight

Har

dwar

e f

176,

200

1983

Pro

puls

ion

Syst

ems

fg

gg

97,6

00Fl

ight

Har

dwar

e f

gg

g26

9,40

019

84 P

ropu

lsio

n Sy

stem

sh

hh

74,4

00Fl

ight

Har

dwar

eh

hh

242,

700

1985

Pro

puls

ion

Syst

ems

ii

i60

,500

Flig

ht H

ardw

are

ii

i26

7,00

019

86 P

ropu

lsio

n Sy

stem

sj

jj

63,2

00Fl

ight

Har

dwar

ej

jj

285,

100

1987

Pro

puls

ion

Syst

ems

kk

k51

,700

Flig

ht H

ardw

are

kk

k25

1,40

019

88 P

ropu

lsio

n Sy

stem

sl

ll

36,0

00Fl

ight

Har

dwar

el

ll

286,

600

aU

ndis

trib

uted

. No

amou

nt s

peci

fied

for

ext

erna

l tan

k ap

prop

riat

ion.

Inc

lude

d in

tota

l R&

D a

ppro

pria

tion

of $

3,47

7,20

0,00

0.b

No

brea

kout

of

supp

lem

enta

l app

ropr

iatio

n su

bmis

sion

; inc

lude

d in

R&

D s

ubm

issi

on o

f $1

85,0

00,0

00.

cSu

pple

men

tal A

ppro

pria

tion

of $

185,

000,

000

spec

ifie

d fo

r ge

nera

l R&

D a

ctiv

ities

.d

Und

istr

ibut

ed. I

nclu

ded

in R

&D

app

ropr

iatio

n of

$4,

091,

096,

000.

eSu

pple

men

tal a

ppro

pria

tion

for

Spac

e Sh

uttle

in r

espo

nse

to a

men

ded

NA

SA b

udge

t sub

mis

sion

of

$300

,000

,000

. No

auth

oriz

atio

n ac

tivity

. Sup

plem

enta

l app

ropr

iatio

nof

$28

5,00

0,00

0 ap

prov

ed w

ith n

o di

stri

butio

n to

indi

vidu

al c

ompo

nent

s.

LAUNCH SYSTEMS 71



Tabl

e 2–

8 co

ntin

ued

fPr

opul

sion

Sys

tem

s an

d Fl

ight

Har

dwar

e bu

dget

cat

egor

ies

wer

e no

t use

d by

NA

SA p

rior

to F

Y 1

984.

How

ever

,pro

gram

med

am

ount

s us

ed th

ese

cate

gori

es in

FY

198

2an

d FY

198

3 to

be

cons

iste

nt w

ith c

ateg

orie

s us

ed in

est

imat

es f

or f

utur

e ye

ars.

gN

o bu

dget

item

list

ed. S

uppo

rtin

g co

mm

ittee

doc

umen

tatio

n in

clud

ed e

xter

nal t

ank

in S

pace

Shu

ttle

Prod

uctio

n w

ith n

o am

ount

spe

cifi

ed. T

otal

Pro

duct

ion

amou

nt:s

ub-

mis

sion

= $

1,58

5,50

0,00

0; a

utho

riza

tion

= $

1,67

0,50

0,00

0; a

nd a

ppro

pria

tion

= $

1,63

6,50

0,00

0.h

No

spec

ific

aut

hori

zatio

n fo

r SS

ME

,ext

erna

l tan

k,or

SR

B. A

ccor

ding

to c

ongr

essi

onal

rep

orts

,the

Spa

ce T

rans

port

atio

n an

d C

apab

ility

Dev

elop

men

t pro

gram

sup

port

-ed

the

prod

uctio

n of

the

SSM

E,S

RB

,and

ext

erna

l tan

k,in

add

ition

to p

rovi

ding

for

cri

tical

spa

res

(as

wel

l as

othe

r ite

ms)

. The

tota

l aut

hori

zatio

n fo

r th

is c

ateg

ory

=$2

,009

,400

,000

. Som

e Sp

ace

Shut

tle f

undi

ng w

as in

clud

ed in

the

Flig

ht H

ardw

are

cate

gory

(se

e ab

ove

for

defi

nitio

n):s

ubm

issi

on =

$84

8,40

0,00

0; a

utho

riza

tion

undi

s-tr

ibut

ed,i

nclu

ded

in S

huttl

e O

pera

tions

aut

hori

zatio

n =

$1,

495,

600,

000;

and

app

ropr

iatio

n (m

oved

to S

FC&

DC

) un

dist

ribu

ted,

incl

uded

in S

huttl

e O

pera

tions

app

ropr

ia-

tion

= $

1,52

0,60

0,00

0.i

SSM

E p

rodu

ctio

n an

d re

sidu

al d

evel

opm

ent t

asks

for

the

orbi

ter,

SSM

E,e

xter

nal t

ank,

and

SRB

fel

l und

er S

pace

Pro

duct

ion

and

Ope

ratio

nal C

apab

ility

,Pro

puls

ion

Syst

ems.

SR

B a

nd e

xter

nal t

ank

proc

urem

ent (

prod

uctio

n) f

ell u

nder

Spa

ce T

rans

port

atio

n O

pera

tions

,Flig

ht H

ardw

are.

No

brea

kdow

n fo

r in

divi

dual

Spa

ce S

huttl

epr

opul

sion

com

pone

nts.

The

198

5 am

ount

s fo

r Pr

opul

sion

Sys

tem

s w

ere:

subm

issi

on =

$59

9,00

0,00

0; a

utho

riza

tion

= $

599,

000,

000;

and

app

ropr

iatio

n =

$59

9,00

0,00

0.A

utho

riza

tion

for

subm

issi

on =

$75

8,00

0,00

0. P

rocu

rem

ent o

f ex

tern

al ta

nk,s

olid

roc

ket m

otor

,and

SR

B h

ardw

are

incl

uded

in S

pace

Tra

nspo

rtat

ion

Ope

ratio

nsPr

ogra

m,F

light

Har

dwar

e am

ount

of

$758

,000

,000

; app

ropr

iatio

n =

$75

8,00

0,00

0. S

SME

pro

duct

ion

and

resi

dual

dev

elop

men

t tas

ks f

or th

e or

bite

r,SS

ME

,ext

erna

lta

nk,a

nd S

RB

fel

l und

er S

pace

Pro

duct

ion

and

Ope

ratio

nal C

apab

ility

,Pro

puls

ion

Syst

ems.

SR

B a

nd e

xter

nal t

ank

proc

urem

ent (

prod

uctio

n) f

ell u

nder

Spa

ceT

rans

port

atio

n O

pera

tions

,Flig

ht H

ardw

are.

Flig

ht H

ardw

are

subm

issi

on =

$75

8,00

0,00

0; a

utho

riza

tion

= $

758,

000,

000;

and

app

ropr

iatio

n =

$75

8,00

0,00

0.j

No

brea

kdow

n fo

r in

divi

dual

Spa

ce S

huttl

e pr

opul

sion

com

pone

nts.

The

198

6 am

ount

s fo

r Pr

opul

sion

Sys

tem

s w

ere:

subm

issi

on =

$45

4,00

0,00

0; a

utho

riza

tion

=$4

54,0

00,0

00; a

nd a

ppro

pria

tion

= $

454,

000,

000.

No

Flig

ht H

ardw

are

budg

et c

ateg

ory

in 1

986.

kN

o br

eakd

own

for

indi

vidu

al S

pace

Shu

ttle

prop

ulsi

on c

ompo

nent

s. T

he 1

987

amou

nts

for

Prop

ulsi

on S

yste

ms

wer

e:su

bmis

sion

= $

338,

400,

000;

aut

hori

zatio

n =

$338

,400

,000

; and

app

ropr

iatio

n =

$33

8,40

0,00

0. F

light

Har

dwar

e su

bmis

sion

= $

646,

200,

000;

aut

hori

zatio

n =

$87

9,10

0,00

0; a

ppro

pria

tion

= $

646,

200,

000.

lN

o br

eakd

own

for

indi

vidu

al S

pace

Shu

ttle

prop

ulsi

on c

ompo

nent

s. T

he 1

988

amou

nts

for

Prop

ulsi

on S

yste

ms

wer

e:su

bmis

sion

= $

552,

100,

000;

aut

hori

zatio

n =

$552

,100

,000

; and

app

ropr

iatio

n =

$24

9,30

0,00

0. F

unds

del

eted

fro

m P

ropu

lsio

n Sy

stem

s; $

302,

800,

000

mov

ed to

Lau

nch

and

Mis

sion

Sup

port

cat

egor

y. F

light

Har

dwar

e su

bmis

sion

= $

923,

100,

000;

aut

hori

zatio

n =

$92

3,10

0,00

0; a

nd a

ppro

pria

tion

= $

923,

100,

000.


Tabl

e 2–

9. U

pper

Sta

ges

Fun

ding

His

tory

(in

tho

usan

ds o

f do

llar

s)Y

ear

(Fis

cal)

Subm

issi

onA

utho

riza

tion

App

ropr

iati

onP

rogr

amm

ed

(Act

ual)

1979

ST

S U

pper

Sta

ges

aa

a19

,300

Upp

er S

tage

Ope

ratio

ns6,

300

1980

ST

S U

pper

Sta

ges

bb

b18

,300

Upp

er S

tage

Ope

ratio

nsb

bb

18,7

0019

81 S

TS

Upp

er S

tage

sc

cc

38,3

00U

pper

Sta

ge O

pera

tions

cc

c30

,900

1982

d—

——

106,

700

1983

e—

——

167,

000

1984

f14

3,20

014

3,20

014

3,20

014

3,20

019

85 g

92,4

0092

,400

92,4

0013

7,40

0Su

ppl.

App

ropr

iatio

n40

,000

h19

8612

2,00

012

2,00

012

2,00

012

2,00

019

8720

2,10

0i

200,

100

j20

2,10

0k

156,

100

1988

159,

700

159,

700

159,

700

aN

o sp

ecif

ic f

undi

ng. S

ubm

issi

on f

or S

pace

Tra

nspo

rtat

ion

Syst

em O

pera

tions

Cap

abili

ty D

evel

opm

ent =

$11

0,50

0,00

0; a

utho

riza

tion

for

Spac

e T

rans

port

atio

n Sy

stem

Ope

ratio

ns C

apab

ility

Dev

elop

men

t by

Sena

te c

omm

ittee

= $

110,

500,

000

(no

fina

l aut

hori

zatio

n); a

nd a

ppro

pria

tion

for

Spac

e T

rans

port

atio

n Sy

stem

Ope

ratio

nsC

apab

ility

Dev

elop

men

t was

und

istr

ibut

ed. T

otal

R&

D a

ppro

pria

tion

= $

3,47

7,20

0,00

0. S

ubm

issi

on f

or S

pace

Tra

nspo

rtat

ion

Syst

em O

pera

tions

= $

33,4

00,0

00.

Aut

hori

zatio

n fo

r Sp

ace

Tra

nspo

rtat

ion

Syst

em O

pera

tions

by

Sena

te c

omm

ittee

= $

33,4

00,0

00 (

no f

inal

aut

hori

zatio

n). A

ppro

pria

tion

undi

stri

bute

d.b

No

spec

ific

fun

ding

for

Upp

er S

tage

s.c

Upp

er S

tage

s w

ere

incl

uded

in th

e Sp

ace

Flig

ht O

pera

tions

Spa

ce T

rans

port

atio

n Sy

stem

s O

pera

tiona

l Cap

abili

ty b

udge

t lin

e ite

m. H

ouse

Com

mitt

ee d

ocum

enta

tion

indi

cate

d th

at N

ASA

sub

mis

sion

,as

wel

l as

cong

ress

iona

l aut

hori

zatio

n,fo

r up

per

stag

e ac

tiviti

es w

as $

29,0

00,0

00; P

ublic

Law

96–

526

repo

rt r

efer

red

to b

oth

STS

Upp

er S

tage

s an

d ST

S O

pera

tions

Upp

er S

tage

s:ap

prop

riat

ion

= $

29,0

00,0

00 f

or S

TS

Upp

er S

tage

s an

d $3

0,90

0,00

0 fo

r ST

S O

pera

tions

Upp

er S

tage

s.d

No

NA

SA s

ubm

issi

on,f

inal

aut

hori

zatio

n,or

app

ropr

iatio

n fo

r U

pper

Sta

ges

indi

cate

d.e

Upp

er S

tage

s in

clud

ed in

Spa

ce F

light

Ope

ratio

ns,b

ut n

o am

ount

spe

cifi

ed. T

otal

Spa

ce F

light

Ope

ratio

ns:s

ubm

issi

on =

$1,

707,

000,

000;

aut

hori

zatio

n =

$1,6

99,0

00,0

00; a

nd a

ppro

pria

tion

= $

1,79

6,00

0,00

0.

LAUNCH SYSTEMS 73



Tabl

e 2–

9 co

ntin

ued

fIn

clud

ed m

odif

icat

ion

of th

e C

enta

ur f

or u

se in

the

Shut

tle.

gIn

clud

ed d

evel

opm

ent o

f T

rans

fer

Orb

it St

age

for

use

in la

unch

ing

the

Mar

s ge

osci

ence

/clim

atol

ogy

orbi

ter

in 1

990.

Als

o in

clud

ed jo

int d

evel

opm

ent p

rogr

am b

etw

een

NA

SA a

nd D

OD

for

use

of

the

Cen

taur

as

an S

TS

uppe

r st

age.

Pro

cure

men

t wou

ld b

e in

itiat

ed in

FY

198

5 fo

r tw

o C

enta

ur G

veh

icle

s to

sup

port

the

Ven

us R

adar

Map

per

mis

sion

pla

nned

for

198

8 an

d th

e T

DR

S-E

mis

sion

.h

Supp

lem

enta

l App

ropr

iatio

n ad

ded

$40,

000,

000

to in

itial

app

ropr

iatio

n fo

r U

pper

Sta

ges

for

tota

l of

$132

,400

,000

iA

men

ded

budg

et s

ubm

issi

on in

crea

sed

amou

nt f

rom

$85

,100

,000

to $

202,

100,

000.

jFi

gure

ref

lect

s au

thor

izat

ion

act,

whi

ch w

as v

etoe

d.k

Figu

re r

efle

cts

App

ropr

iatio

n C

onfe

renc

e C

omm

ittee

act

ion,

whi

ch w

as s

ubse

quen

tly in

clud

ed in

the

Om

nibu

s A

ppro

pria

tion

Act

of

1987

(Pu

blic

Law

99–

591)

.


Tabl

e 2–

10. O

rbit

al M

aneu

veri

ng V

ehic

le F

undi

ng H

isto

ry (

in t

hous

ands

of

doll

ars)

Yea

r (F

isca

l)Su

bmis

sion

Aut

hori

zati

onA

ppro

pria

tion

Pro

gram

med

(A

ctua

l)19

83In

clud

ed in

Adv

ance

d Pr

ogra

ms

1984

Incl

uded

in A

dvan

ced

Prog

ram

s19

85In

clud

ed in

Adv

ance

d Pr

ogra

ms

1986

25,0

0013

,000

10,0

005,

000

1987

45,0

00a

50,0

00b

45,0

00c

45,0

0019

8880

,000

75,0

0055

,000

aR

efle

cts

revi

sed

budg

et s

ubm

issi

on,w

hich

dec

reas

ed a

mou

nt f

rom

$70

,000

,000

to $

45,0

00,0

00b

Figu

re r

efle

cts

auth

oriz

atio

n ac

t,w

hich

was

vet

oed.

cFi

gure

ref

lect

s A

ppro

pria

tion

Con

fere

nce

Com

mitt

ee a

ctio

n,w

hich

was

sub

sequ

ently

incl

uded

in th

e O

mni

bus

App

ropr

iatio

n A

ct o

f 19

87 (

Publ

ic L

aw 9

9–59

1).

LAUNCH SYSTEMS 75



Tabl

e 2–

11. T

ethe

red

Sate

llit

e Sy

stem

Fun

ding

His

tory

(in

tho

usan

ds o

f do

llar

s)Y

ear

(Fis

cal)

Subm

issi

onA

utho

riza

tion

App

ropr

iati

onP

rogr

amm

ed

(Act

ual)

1979

Incl

uded

in A

dvan

ced

Prog

ram

s19

80In

clud

ed in

Adv

ance

d Pr

ogra

ms

1981

Incl

uded

in A

dvan

ced

Prog

ram

s19

82In

clud

ed in

Adv

ance

d Pr

ogra

ms

1983

Incl

uded

in A

dvan

ced

Prog

ram

s19

843,

300

3,30

03,

300

3,30

019

8518

,200

18,2

0018

,200

15,8

0019

8621

,000

14,0

0021

,000

15,0

0019

8710

,600

a11

,600

b10

,600

c10

,600

1988

7,30

07,

300

7,30

0a

Figu

re r

efle

cts

$1,0

00,0

00 r

educ

tion

from

initi

al b

udge

t sub

mis

sion

.b

Figu

re r

efle

cts

auth

oriz

atio

n ac

t,w

hich

was

vet

oed.

cFi

gure

ref

lect

s A

ppro

pria

tion

Con

fere

nce

Com

mitt

ee a

ctio

n,w

hich

was

sub

sequ

ently

incl

uded

in th

e O

mni

bus

App

ropr

iatio

n A

ct o

f 19

87 (

Publ

ic L

aw 9

9–59

1).


LAUNCH SYSTEMS 77

Tabl

e 2–

12. A

dvan

ced

Pro

gram

s/P

lann

ing

Fun

ding

His

tory

(in

tho

usan

ds o

f do

llar

s)Y

ear

(Fis

cal)

Subm

issi

onA

utho

riza

tion

App

ropr

iati

onP

rogr

amm

ed

(Act

ual)

1979

5,00

0a

b7,

000

1980

13,0

00c

d13

,000

1981

8,80

013

,800

ef

11,8

0019

828,

800

12,8

00g

8,80

09,

700

1983

11,9

00h

11,9

0012

,600

1984

15,0

0025

,000

i15

,000

21,5

0019

8514

,500

14,5

0020

,500

20,5

0019

8621

,000

21,0

0021

,000

19,4

0019

8716

,600

16,6

00j

16,6

00k

33,6

0019

8824

,900

24,9

0030

,900

aU

ndis

trib

uted

. Inc

lude

d in

Spa

ce F

light

Ope

ratio

ns P

rogr

am a

utho

riza

tion

of $

315,

900,

000.

bU

ndis

trib

uted

. Inc

lude

d in

R&

D a

ppro

pria

tion

of $

3,47

7,20

0,00

0.c

Und

istr

ibut

ed. I

nclu

ded

in S

pace

Flig

ht O

pera

tions

Pro

gram

aut

hori

zatio

n of

$46

3,30

0,00

0.d

Und

istr

ibut

ed. I

nclu

ded

in R

&D

app

ropr

iatio

n of

$4,

091,

086,

000.

eIn

crea

sed

auth

oriz

atio

n re

com

men

ded

by H

ouse

Com

mitt

ee to

sup

port

enh

ance

d Ph

ase

B d

efin

ition

stu

dies

and

tech

nica

l dev

elop

men

t for

the

pow

er e

xten

sion

pac

kage

(PE

P) a

nd th

e 25

-kilo

wat

t (kW

) po

wer

mod

ule.

fU

ndis

trib

uted

. Inc

lude

d in

R&

D a

ppro

pria

tion

of $

4,39

6,20

0,00

0 (m

odif

ied

by G

ener

al P

rovi

sion

,Sec

. 412

,to

$4,3

40,7

88).

gH

ouse

rec

omm

ende

d ad

ditio

nal a

utho

riza

tion

of $

5,00

0,00

0 fo

r PE

P,25

-kW

pow

er m

odul

e,sp

ace

plat

form

s,sp

ace

oper

atio

ns d

efin

ition

stu

dies

,and

adv

ance

d te

chni

cal

deve

lopm

ent.

Con

fere

nce

Com

mitt

ee r

educ

ed a

dditi

onal

aut

hori

zatio

n to

$12

,800

,000

.h

Und

istr

ibut

ed. I

nclu

ded

in S

pace

Flig

ht O

pera

tions

aut

hori

zatio

n of

$1,

699,

000,

000.

iH

ouse

aut

hori

zed

addi

tiona

l $10

,000

,000

for

spa

ce s

tatio

n st

udie

s an

d sp

ace

plat

form

. Sen

ate

auth

oriz

ed a

dditi

onal

$5,

000,

000

for

spac

e st

atio

n st

udie

s. C

onfe

renc

eC

omm

ittee

aut

hori

zed

addi

tiona

l $10

,000

,000

for

spa

ce s

tatio

n st

udie

s.j

Figu

re r

efle

cts

auth

oriz

atio

n ac

t,w

hich

was

vet

oed.

kFi

gure

ref

lect

s A

ppro

pria

tion

Con

fere

nce

Com

mitt

ee a

ctio

n,w

hich

was

sub

sequ

ently

incl

uded

in th

e O

mni

bus

App

ropr

iatio

n A

ct o

f 19

87 (

Publ

ic L

aw 9

9–59

1).



Tabl

e 2–

13. E

LVSu

cces

s R

ate

by Y

ear

and

Lau

nch

Vehi

cle

for

NA

SAL

aunc

hes

Yea

r (F

isca

l)A

tlas

-Cen

taur

Atl

as E

/FD

elta

Scou

tTo

tal

1979

2/2

1/1

3/3

3/3

9/9

1980

3/3

0/1

1/1

—1/

119

814/

41/

15/

51/

111

/11

1982

2/2

—7/

7—

9/9

1983

1/1

1/1

8/8

1/1

11/1

119

840/

11/

14/

41/

16/

719

853/

3—

—2/

25/

519

861/

11/

11/

21/

14/

519

870/

1—

2/2

1/1

3/4

1988

—1/

11/

14/

46/

6To

tal

16/1

8 (8

8.9%

)6/

7 (8

5.7%

)34

/35

(97.

1%)

14/1

4 (1

00%

)70

/74

(94.

6%)


LAUNCH SYSTEMS 79

Tabl

e 2–

14. N

ASA

Atl

as E

/F V

ehic

le L

aunc

hes

Atl

as-E

/F V

ehic

leD

ate

Mis

sion

Atl

as S

ucce

ssfu

l aA

tlas

FJu

ne 2

7,19

79N

OA

A-6

Yes

Atla

s F

May

29,

1980

NO

AA

-BN

o. L

aunc

h ve

hicl

e m

alfu

nctio

ned;

faile

d to

pla

ce s

atel

lite

into

pro

per

orbi

t.A

tlas

FJu

ne 2

3,19

81N

OA

A-7

Yes

Atla

s E

Mar

ch 2

8,19

83N

OA

A-8

Yes

Atla

s E

Dec

. 12,

1984

NO

AA

-9Y

esA

tlas

ESe

pt. 1

7,19

86N

OA

A-1

0Y

esA

tlas

ESe

pt. 2

4,19

88N

OA

A-1

1Y

esa

One

fai

lure

out

of

seve

n at

tem

pts

(85.

7% s

ucce

ss r

ate)

.



Tabl

e 2–

15. A

tlas

E/F

Cha

ract

eris

tics

1-1/

2 St

ages

(B

oost

er &

Sus

tain

er)

Apo

gee

Kic

k M

otor

Fai

ring

Len

gth

21.3

met

ers

(m)

—7.

0 m

Ove

rall

Len

gth

Up

to 2

8.3

m in

clud

ing

fair

ing

Dia

met

er3.

05 m

2.1

mG

ross

Wei

ght (

Lif

toff

)12

1,00

0 ki

logr

ams

(kg)

47.7

kg

(wei

ght o

f m

otor

)73

5 kg

asse

mbl

y ca

se a

fter

depl

etio

n of

fue

l)Fu

el W

eigh

t11

2,90

0 kg

666

kgE

ngin

e Ty

pe/N

ame

MA

-3 s

yste

m c

onsi

stin

g of

TE

-M-3

64-1

5L

R 8

9-N

A-5

boo

ster

,L

R 1

05-N

A-5

sus

tain

er,

LR

101

-NA

-7 v

erni

er e

ngin

esN

umbe

r of

Eng

ines

2 bo

oste

r en

gine

s,1

1 su

stai

ner

engi

ne,&

2 ve

rnie

r en

gine

s (V

E)

Prop

ella

ntL

OX

& R

J-1-

1So

lidB

urn

Tim

e (A

vg.)

120-

sec

boos

ter,

309-

sec.

sus

tain

er45

sec

.L

ifto

ff T

hrus

t1,

743,

000

new

tons

Avg

. Thr

ust p

er E

ngin

e1,

470,

000

new

tons

(bo

oste

rs);

650,

800

new

tons

(267

,000

new

tons

(su

stai

ner)

;3,

000

new

tons

(ea

ch V

E)

Max

. Pay

load

2,09

0 kg

in 1

85-k

m o

rbit

from

pol

ar la

unch

with

dua

l TE

-364

4 e

ngin

es; 1

,500

kg

in 1

85-k

m o

rbit

from

po

lar

laun

ch w

ith s

ingl

e T

E 3

74-4

eng

ine


LAUNCH SYSTEMS 81

Tabl

e 2–

15 c

onti

nued

1-1/

2 St

ages

(B

oost

er &

Sus

tain

er)

Apo

gee

Kic

k M

otor

Fai

ring

Prim

e C

ontr

acto

rG

ener

al D

ynam

ics

Con

trac

tors

Roc

ketd

yne

Thi

okol

How

Util

ized

To la

unch

met

eoro

logi

cal s

atel

lites

Rem

arks

The

Atla

s E

/F s

erie

s w

as o

rigi

nally

dep

loye

d as

IC

BM

s. B

y th

e la

te 1

970s

,the

rem

aini

ng A

tlas

E/F

s w

ere

conv

erte

d fo

r sp

ace

laun

ch. D

urin

g 19

79–1

988,

they

wer

e us

ed o

nly

to la

unch

met

eoro

logi

cal

sate

llite

s. O

n pa

rtic

ular

mis

sion

s,th

e fa

irin

gs w

ere

leng

then

ed to

7.4

m to

acc

omm

odat

e ad

ditio

nal

equi

pmen

t—fo

r in

stan

ce,s

earc

h-an

d-re

scue

equ

ipm

ent o

n N

OA

A m

issi

ons.



Tabl

e 2–

16. N

ASA

Atl

as-C

enta

ur V

ehic

le L

aunc

hes

Atl

as-C

enta

ur V

ehic

leD

ate

Mis

sion

Atl

as-C

enta

urSe

rial

Num

ber

Succ

essf

ul a

AC

-47

May

4,1

979

FltS

atC

om 2

Yes

AC

-53

Sept

. 20,

1979

HE

AO

3Y

esA

C-4

9Ja

n. 1

7,19

80Fl

tSat

Com

3Y

esA

C-5

2O

ct. 3

0,19

80Fl

tSat

Com

4Y

esA

C-5

4D

ec. 6

,198

0In

tels

at V

-A F

-2Y

esA

C-4

2Fe

b. 2

1,19

81C

omst

ar 4

Yes

AC

-56

May

23,

1981

Inte

lsat

V-B

F-1

Yes

AC

-59

Aug

. 6,1

981

FltS

atC

om 5

Yes

AC

-55

Dec

. 15,

1981

Inte

lsat

V F

-3Y

esA

C-5

8M

ar. 4

,198

2In

tels

at V

-D F

-4Y

esA

C-6

0Se

pt. 2

8,19

82In

tels

at V

-E F

-5Y

esA

C-6

1M

ay 1

9,19

83In

tels

at V

-F F

-6Y

esA

C-6

2Ju

ne 9

,198

4In

tels

at V

-G F

-9N

o. V

ehic

le f

aile

d to

pla

cesa

telli

te in

use

ful o

rbit.

AC

-36

Mar

. 22,

1985

Inte

lsat

V-A

F-1

0Y

esA

C-6

4Ju

ne 2

9,19

85In

tels

at V

-A F

-11

Yes

AC

-65

Sept

. 28,

1985

Inte

lsat

V-A

F-1

2Y

esA

C-6

6D

ec. 4

,198

6Fl

tSat

Com

7Y

esA

C-6

7M

ar. 2

6,19

87Fl

tSat

Com

6N

o. T

elem

etry

lost

sho

rtly

aft

er la

unch

;de

stru

ct s

igna

l sen

t at 7

0.7

seco

nds

into

flig

ht. A

n el

ectr

ical

tran

sien

t,ca

used

by

light

ning

str

ike

on la

unch

vehi

cle,

was

mos

t pro

babl

e ca

use

of lo

ss.

aTw

o fa

ilure

s ou

t of

eigh

teen

atte

mpt

s (8

8.9%

suc

cess

rat

e).


Tabl

e 2–

17. A

tlas

-Cen

taur

Cha

ract

eris

tics

Atl

as B

oost

er &

Sust

aine

r SL

V-3

DC

enta

ur S

tage

D-1

AL

engt

h21

.1 m

eter

s (m

)9.

1 m

with

out f

airi

ng; 1

8.6

m (

with

pay

load

fai

ring

)O

vera

ll L

engt

h40

.8 m

incl

udin

g fa

irin

gD

iam

eter

3.05

m3.

05 m

Eng

ine

Type

/Nam

eM

A-5

sys

tem

con

sist

ing

of 2

boo

ster

sR

L-1

01

sust

aine

r,an

d 2

vern

ier

engi

nes

Prim

e C

ontr

acto

rG

ener

al D

ynam

ics

Con

trac

tors

Roc

ketd

yne

Prat

t & W

hitn

ey A

ircr

aft

Num

ber

of E

ngin

es5

(2 b

oost

er e

ngin

es,1

thru

st s

usta

iner

eng

ine,

2 th

rust

eng

ines

and

14

smal

l hyd

roge

n pe

roxi

de th

rust

ers

2 ve

rnie

r en

gine

s)L

ifto

ff T

hrus

t (A

vg.)

1,93

1,00

0 ne

wto

ns (

at s

ea le

vel)

usi

ng tw

o13

3,44

0 ne

wto

ns (

vacu

um)

usin

g tw

o

828,

088-

new

ton-

thru

st b

oost

er e

ngin

es,

67,0

00-n

ewto

n-th

rust

RL

-10

engi

nes

and

one

267,

000-

new

ton-

thru

st s

usta

iner

eng

ine,

14 s

mal

l hyd

roge

n pe

roxi

de th

rust

ers

and

two

vern

ier

engi

nes

deve

lopi

ng3,

006

new

tons

eac

hB

urn

Tim

e17

4-se

c. b

oost

er,2

26-s

ec. s

usta

iner

450

sec.

Prop

ella

ntL

OX

as

the

oxid

izer

and

RP-

1L

OX

and

LH

2

Max

. Pay

load

6,10

0 ki

logr

ams

(kg)

in 1

85-k

m o

rbit;

2,3

60 k

g in

geo

sync

hron

ous

tran

sfer

orb

it; 9

00 k

g to

Ven

us o

r M

ars

Lau

nch

Wei

ght

128,

934

kg17

,676

kg

How

Util

ized

Prim

arily

to la

unch

com

mun

icat

ions

sat

ellit

esR

emar

ksU

nlik

e ea

rlie

r Atla

s-C

enta

ur c

ombi

natio

ns,t

he S

LV-3

D a

nd la

ter

mod

els

wer

e in

tegr

ated

ele

ctro

nica

lly w

ith th

e C

enta

ur D

-1A

upp

er s

tage

. The

Int

elsa

t V-A

F-1

0,In

tels

at V

-A F

-11,

Inte

lsat

V-A

F-1

2,Fl

tSat

Com

5,a

nd

FltS

atC

om 6

mis

sion

s us

ed th

e A

tlas

G c

onfi

gura

tion.

The

Atla

s st

age

on th

e “G

”co

nfig

urat

ion

is 2

.06

m lo

nger

th

an th

e SL

V-3

D,a

nd it

s en

gine

pro

vide

d 33

,600

new

tons

mor

e th

rust

than

the

SLV

-3D

eng

ines

.

LAUNCH SYSTEMS 83



Tabl

e 2–

18. C

hron

olog

y of

Del

ta V

ehic

le L

aunc

hes

Del

ta V

ehic

le T

ype

Dat

eM

issi

onD

elta

Suc

cess

ful a

2914

/148

Jan.

30,

1979

SCA

TH

AY

es29

14/1

49A

ug. 9

,197

9W

esta

r-C

Yes

3914

/150

Dec

. 6,1

979

RC

A-C

Yes

3910

/151

Feb.

14,

1980

Sola

r M

ax M

issi

onY

es39

14/1

52Se

pt. 9

,198

0G

OE

S 4

Yes

3910

-PA

M/1

53N

ov. 1

5,19

80SB

S-A

(fi

rst u

se o

f PA

M)

Yes

3914

/154

May

22,

1981

GO

ES

5Y

es39

13/1

55

Aug

. 3,1

981

Dyn

amic

Exp

lore

r D

E-A

/BY

es39

10-P

AM

/156

Sept

. 24,

1981

SBS-

BY

es23

10/1

57 b

Oct

. 6,1

981

SME

/Uos

atY

es39

10-P

AM

/158

Nov

. 20,

1981

RC

A-D

Yes

3910

-PA

M/1

59Ja

n. 1

5,19

82R

CA

-CY

es39

10-P

AM

/160

Feb.

25,

1982

Wes

tar

IVY

es32

910-

PAM

/161

Apr

. 10,

1982

Insa

t-1A

Yes

3915

/162

June

8,1

982

Wes

tar-

VY

es39

20/1

63Ju

ly 1

6,19

82L

ands

at-D

Yes

3920

-PA

M/1

64A

ug. 2

6,19

82(T

eles

at-F

) Ani

k-D

-1Y

es39

24/1

65O

ct. 2

7,19

82R

CA

-EY

es39

10/1

66Ja

n. 2

5,19

83IR

AS/

PIX

II

Yes

3924

/167

Apr

. 11,

1983

RC

A-F

Yes

3914

/168

Apr

. 28,

1983

GO

ES

FY

es39

14/1

69M

ay 2

6,19

83E

XO

SAT

Yes

3920

-PA

M/1

70Ju

ne 2

8,19

83G

alax

y-A

Yes

3920

-PA

M/1

71

July

28,

1983

Tels

tar-

3AY

es


LAUNCH SYSTEMS 85

Tabl

e 2–

18 c

onti

nued

Del

ta V

ehic

le T

ype

Dat

eM

issi

onD

elta

Suc

cess

ful a

3924

/172

Sept

. 8,1

983

Satc

om-I

IR (

RC

A-G

)Y

es39

20-P

AM

/173

Sept

. 22,

1983

Gal

axy-

BY

es39

20/1

74M

ar. 1

,198

4L

ands

at-D

/Uos

atY

es39

24/1

75A

ug. 1

6,19

84A

MPT

EY

es39

20-P

AM

/176

Sept

. 21,

1984

Gal

axy-

CY

es39

14/1

77N

ov. 1

3,19

84N

AT

O-I

IID

Yes

3914

/178

May

3,1

986

GO

ES

GN

o. V

ehic

le f

aile

d.39

20/1

80Se

pt. 5

,198

6SD

IY

es39

24/1

79Fe

b. 2

6,19

87G

OE

S H

Yes

3920

-PA

M/1

82M

ar. 2

0,19

87Pa

lapa

-B2P

Yes

3910

/181

Feb.

8,1

988

SDI

Yes

aO

ne f

ailu

re o

ut o

f th

irty

-fiv

e at

tem

pts

(97.

1% s

ucce

ss r

ate)

.b

Thr

ee s

trap

-on

engi

nes.



Tabl

e 2–

19. D

elta

291

4 C

hara

cter

isti

csSt

rap-

onSt

age

ISt

age

IISt

age

III

Len

gth

21.3

m6.

4 m

1.4

mO

vera

ll L

engt

h35

.5 m

incl

udin

g sp

acec

raft

shr

oud

Dia

met

erO

vera

ll ba

sic

diam

eter

of

2.4

mE

ngin

e Ty

pe/N

ame

TX

-354

-5 C

asto

r II

RS-

27 e

xten

ded

long

8-fo

ot-d

iam

eter

TE

-364

-4ta

nk T

hor

TR

-201

No.

of

Eng

ines

91

mai

n an

d 2

vern

ier

11

Thr

ust (

per

Eng

ine)

(A

vg.)

233,

856

new

tons

911,

840

new

tons

45,8

00 n

ewto

ns66

,586

new

tons

Lif

toff

Thr

ust

1,76

5,31

5 ne

wto

ns (

incl

udes

6 o

f 9

stra

p-on

s,w

hich

are

igni

ted

at li

ftof

f)B

urn

Tim

e37

sec

.20

9 se

c.33

5 se

c.44

sec

.Pr

opel

lant

Solid

RP-

1/L

OX

N2 O

4&

aer

ozin

e-50

Solid

Fuel

Wei

ght

18,0

84 k

g ea

ch s

trap

-on

80,2

64 k

g4,

593

kg1,

039

kgL

ifto

ff W

eigh

t40

,320

kg

84,3

30 k

g6,

125

kg1,

120

kgPr

ime

Con

trac

tor

McD

onne

ll D

ougl

asC

ontr

acto

rsT

hiok

olR

ocke

tdyn

eT

RW

Thi

okol

How

Util

ized

Med

ium

-wei

ght p

aylo

ads

Rem

arks

Onl

y th

ree

Del

tas

in th

e 29

00 s

erie

s w

ere

used

bet

wee

n 19

79 a

nd 1

988.

Tw

o w

ere

2914

s an

d on

e w

as a

23

10,w

hich

had

onl

y th

ree

stra

p-on

mot

ors

and

two

stag

es.


LAUNCH SYSTEMS 87

Tabl

e 2–

20. D

elta

391

0/39

14 C

hara

cter

isti

csSt

rap-

onSt

age

ISt

age

IISt

age

III

Ove

rall

Len

gth

35.5

m in

clud

ing

spac

ecra

ft s

hrou

dL

engt

h 11

.3 m

21.3

m6.

0 m

a1.

8 m

aD

iam

eter

Ove

rall

2.4

m m

ax.

Eng

ine

Type

/Nam

eC

asto

r IV

/TX

-526

-2R

S-27

mod

ifie

d lo

ng ta

nkT

R-2

01T

hiok

ol T

E-3

64-3

or

Tho

r bo

oste

rT

E-3

64-4

No.

of

Eng

ines

91

mai

n an

d 2

vern

ier

11

Bur

n T

ime

(Avg

.)57

sec

.22

4 se

c.32

0 se

c.44

sec

.Sp

ecif

ic I

mpu

lse

(Avg

.)22

9.9

sec.

262.

4 se

c.31

9 se

c.28

3 se

c.T

hrus

t 37

7,16

5 ne

wto

ns91

1,88

7 ne

wto

ns43

,815

new

tons

TE

364

-3 e

ngin

e:42

,169

new

tons

TE

364

-4 e

ngin

e:66

,586

new

tons

Prop

ella

ntSo

lid T

P-H

-803

8L

OX

and

RP-

1 (h

ydra

zine

) or

LO

XN

2 O4

and

Solid

and

RJ-

1 (l

iqui

d hy

droc

arbo

n) b

aero

zine

-50

Fuel

Wei

ght

9,37

3 kg

80,2

64 k

g4,

593

kg1,

039

kgG

ross

Wei

ght

10,8

40 k

g ea

ch85

,076

kg

6,11

5 kg

1,15

8 kg

Prim

e C

ontr

acto

rM

cDon

nell

Dou

glas

Con

trac

tors

Thi

okol

Roc

ketd

yne

TR

WT

hiok

olH

ow U

tiliz

edM

ediu

m-w

eigh

t pay

load

sR

emar

ksW

ith th

e ex

cept

ions

not

ed b

elow

in n

otes

aan

d b,

the

3910

was

iden

tical

to th

e 39

14 b

ut h

ad o

nly

two

stag

es.

For

laun

ches

fro

m th

e E

aste

rn S

pace

and

Mis

sile

Cen

ter,

six

stra

p-on

mot

ors

wer

e ig

nite

d at

lift

off

and

jetti

sone

d ap

prox

imat

ely

nine

sec

onds

aft

er ig

nitio

n of

the

seco

nd s

et o

f th

ree

stra

p-on

mot

ors.

The

rem

aini

ngth

ree

mot

ors

wer

e je

ttiso

ned

at a

ppro

xim

atel

y 12

6 se

cond

s af

ter

lifto

ff. F

or th

e W

este

rn S

pace

and

Mis

sile

Cen

ter,

the

six

grou

nd-i

gnite

d m

otor

s w

ere

jetti

sone

d at

a la

ter

time

for

rang

e sa

fety

con

side

ratio

ns.

aT

he le

ngth

of

the

seco

nd s

tage

on

the

3910

equ

aled

the

sum

of

the

leng

ths

of th

e se

cond

and

thir

d st

ages

on

the

3914

. The

leng

ths

of th

e in

divi

dual

sta

ges

did

not i

nclu

deth

e le

ngth

of

the

spac

ecra

ft s

hrou

d.b

The

391

0 us

ed L

OX

and

RP-

1 pr

opel

lant

; the

391

4 us

ed L

OX

and

RJ-

1 pr

opel

lant

.



Tabl

e 2–

21. D

elta

392

0/39

24 C

hara

cter

isti

cs

Stra

p-on

Stag

e I

Stag

e II

Stag

e II

IL

engt

h21

.3 m

6m1.

8mO

vera

ll L

engt

h35

.5 m

incl

udin

g sp

acec

raft

shr

oud

Dia

met

erO

vera

ll 2.

4 m

max

.E

ngin

e Ty

pe/N

ame

Cas

tor

IV T

X-5

26-2

sol

id b

oost

ers

RS-

27 m

odif

ied

long

tank

Impr

oved

Tra

nsta

geT

E-3

64-4

Tho

r bo

oste

rIn

ject

or P

rogr

amN

o. o

f E

ngin

es9

1 m

ain

& 2

ver

nier

11

Spec

ific

Im

puls

e (A

vg.)

229.

9 se

c.26

2.4

sec.

319

sec.

283.

6 se

c.T

hrus

t (A

vg.)

377,

165

new

tons

911,

007

new

tons

44

,000

new

tons

66,5

86 n

ewto

nsB

urn

Tim

e57

sec

.22

4 se

c.32

0 se

c.44

sec

.Pr

opel

lant

Solid

TP-

H-8

038

RP-

1 an

d L

OX

Aer

ozin

e-50

and

So

lidN

2 O4

oxid

e

Fuel

Wei

ght

9,37

3 kg

79,3

80 k

g4,

593

kg1,

039

kgM

ax P

aylo

ad3,

045

kg in

185

-km

orb

it w

ith d

ue e

ast l

aunc

h; 1

,275

kg

in g

eosy

nchr

onou

s tr

ansf

er o

rbit

with

due

eas

t lau

nch;

2,

135

kg in

cir

cula

r Su

n-sy

nchr

onou

s or

bit w

ith p

olar

laun

ch; 2

,180

kg

in 1

85-k

m o

rbit

with

pol

ar la

unch

Gro

ss W

eigh

t10

,840

kg

85,0

76 k

g6,

920

kg1,

122

kgPr

ime

Con

trac

tor

McD

onne

ll D

ougl

asC

ontr

acto

rsT

hiok

olR

ocke

tdyn

eA

eroj

etT

hiok

olH

ow U

tiliz

edM

id-s

ize

com

mun

icat

ion

and

met

eoro

logi

cal s

atel

lite

Rem

arks

The

leng

th o

f th

e se

cond

sta

ge o

f th

e 39

20 e

qual

led

the

com

bine

d le

ngth

s of

the

seco

nd a

nd th

ird

stag

es o

f th

e 39

24. L

engt

hs o

f in

divi

dual

sta

ges

did

not i

nclu

de le

ngth

of

the

spac

ecra

ft s

hrou

d.


LAUNCH SYSTEMS 89

Tabl

e 2–

22. N

ASA

Sco

ut L

aunc

hes

Scou

t Veh

icle

Dat

eM

issi

onSc

out

Succ

essf

ulS-

202

Feb.

18,

1979

SAG

EY

esS-

198

June

2,1

979

UK

-6Y

esS-

203

Oct

. 30,

1979

Mag

sat

Yes

S-19

2M

ay 1

4,19

81N

OV

A-I

Yes

S-20

5Ju

ne 2

7,19

83H

ilat

Yes

S-20

8O

ct. 1

1,19

84N

OV

A-I

IIY

esS-

209

Aug

. 2,1

985

SOO

S-1

Yes

S-20

7D

ec. 1

2,19

85A

FIT

VY

esS-

199

Nov

. 13,

1986

AF

Pola

r B

EA

RY

esS-

204

Sept

. 16,

1987

SOO

S-2

Yes

S-20

6M

ar. 2

5,19

88Sa

n M

arco

-DL

Yes

S-21

1A

pr. 2

5,19

88SO

OS-

3Y

esS-

213

June

15,

1988

NO

VA

-II

Yes

S-21

4A

ug. 2

5,19

88SO

OS-

4Y

esA

ll of

atte

mpt

ed la

unch

es w

ere

succ

essf

ul.



Tabl

e 2–

23. S

cout

Cha

ract

eris

tics

(G

-1)

Fir

st S

tage

Seco

nd S

tage

Thi

rd S

tage

Fou

rth

Stag

eL

engt

h9.

94 m

6.56

m3.

28 m

1.97

mO

vera

ll L

engt

h22

.86

m in

clud

ing

tran

sitio

n an

d pa

yloa

d se

ctio

nsW

eigh

t14

,255

kg

4,42

4 kg

1,39

5 kg

302

kgD

iam

eter

1.01

m m

ax.

Eng

ine

Type

/ Nam

eA

lgol

III

AC

asto

r II

AA

ntar

es I

IIA

aA

ltair

III

A/ S

tar

31T

hrus

t (A

vg.)

481,

000

new

tons

281,

000

new

tons

83,1

00 n

ewto

ns25

,593

new

tons

Fuel

Solid

Solid

Solid

Solid

Fuel

Wei

ght

12,6

84 k

g3,

762

kg1,

286

kg27

5 kg

Lau

nch

Wei

ght

14,2

15 k

g4,

433

kg1,

394

kg30

1 kg

Bur

n T

ime

(Avg

.)90

sec

.46

sec

.48

.4 s

ec.

30 s

ec.

Payl

oad

Cap

acity

227.

2 kg

pay

load

to a

480

-km

Ear

th o

rbit

Prim

e C

ontr

acto

rV

ough

t Cor

p. (

LTV

Cor

p.)

Con

trac

tors

Uni

ted

Tech

nolo

gies

Thi

okol

Thi

okol

Thi

okol

How

Util

ized

Smal

ler

payl

oads

Rem

arks

An

optio

nal f

ifth

sta

ge u

sed

the

Alc

yone

IA

eng

ine,

with

a th

rust

of

appr

oxim

atel

y 26

,230

new

tons

,a b

urn

time

of 8

.42

sec.

,and

a to

tal w

eigh

t of

98.2

kg.

aM

issi

ons

prio

r to

Mag

sat (

SAG

E a

nd U

K-6

) us

ed th

e A

ntar

es I

I th

ird

stag

e en

gine

.


LAUNCH SYSTEMS 91

Tabl

e 2–

24. S

TS-

Lau

nche

d M

issi

ons

Veh

icle

Mis

sion

Dep

loye

d P

aylo

adD

ate

Col

umbi

aST

S-1

Firs

t tes

t flig

ht,n

o de

ploy

able

pay

load

Apr

. 12–

14,1

981

Col

umbi

aST

S-2

Seco

nd te

st f

light

,no

depl

oyab

le p

aylo

adN

ov. 1

2–14

,198

1C

olum

bia

STS-

3T

hird

test

flig

ht,n

o de

ploy

able

pay

load

Mar

. 22–

30,1

982

Col

umbi

aST

S-4

Four

th a

nd f

inal

test

flig

ht; D

OD

pay

load

82-

1Ju

ne 2

7–Ju

ly 4

,198

2C

olum

bia

STS-

5SB

S-C

/PA

M-D

,Ani

k C

-3/P

AM

-D (

Tele

sat-

E)

(Can

ada)

Nov

. 11–

16,1

982

Cha

llen

ger

STS-

6T

rack

ing

and

Dat

a R

elay

Sat

ellit

e (T

DR

S)-1

/IU

SA

pr. 4

–9,1

983

Cha

llen

ger

STS-

7Te

lesa

t 7 (

Ani

k C

-2)/

PAM

-D (

Can

ada)

/PA

M-D

,Pal

apa

B-1

June

18–

24,1

983

(Ind

ones

ia)/

PAM

-DC

hall

enge

rST

S-8

INSA

T-1B

/PA

M-D

(In

dia)

Aug

. 30–

Sept

. 5,1

983

Col

umbi

aST

S-9

Spac

elab

-1 (

no s

atel

lites

dep

loye

d)N

ov. 2

8–D

ec. 8

,198

3C

hall

enge

rST

S 41

-BPa

lapa

-B2/

PAM

-D (

Indo

nesi

a),W

esta

r V

I/PA

M-D

Feb.

3–1

1,19

84C

hall

enge

rST

S 41

-CL

ong

Dur

atio

n E

xpos

ure

Faci

lity

(LD

EF-

1)A

pr. 6

–13,

1984

Dis

cove

ryST

S 41

-DSy

ncom

IV

-2 (

Lea

sat 2

)/U

niqu

e U

pper

Sta

ge*,

Tels

tar

3-C

/PA

M-D

,SB

S-D

/PA

M-D

Aug

. 30–

Sept

. 5,1

984

Cha

llen

ger

STS

41-G

Ear

th R

adia

tion

Bud

get S

atel

lite

(ER

BS)

Oct

. 5–1

3,19

84D

isco

very

STS

51-A

Sync

om I

V-1

(L

easa

t 1)/

Uni

que

Upp

er S

tage

*,A

nik

Nov

. 8–1

6,19

84(T

eles

at-H

)/PA

M-D

Dis

cove

ryST

S 51

-CD

OD

cla

ssif

ied

payl

oad/

IUS

Jan.

24–

27,1

984

Dis

cove

ryST

S 51

-DA

nik

C-1

(Te

lesa

t-I)

/PA

M-D

,Syn

com

IV

(L

easa

t 3)/

Uni

que

Apr

. 12–

19,1

985

Upp

er S

tage

*C

hall

enge

rST

S 51

-BSp

acel

ab 3

,NU

SAT,

GL

OM

R (

faile

d to

dep

loy)

Apr

. 29–

May

6,1

985

Dis

cove

ryST

S 51

-GM

orel

os-A

/PA

M-D

(M

exic

o),A

rabs

at-A

/PA

M-D

,Ju

ne 1

7–24

,198

5Te

lsta

r 3-

D/P

AM

-D,S

part

an-1

/MPE

SS



Tabl

e 2–

24 c

onti

nued

Veh

icle

Mis

sion

Dep

loye

d P

aylo

adD

ate

Cha

llen

ger

STS

51-F

Spac

elab

2 (

no s

atel

lites

dep

loye

d)Ju

ly 2

9–A

ug. 6

,198

5D

isco

very

STS

51-I

ASC

-1/P

AM

-D,A

ussa

t-1/

PAM

-D (

Aus

tral

ia),

Sync

om I

VA

ug. 2

7–Se

pt. 3

,198

5(L

easa

t-4)

/Uni

que

Upp

er S

tage

*A

tlan

tis

STS

51-J

DO

D M

issi

onO

ct. 3

–7,1

985

Cha

llen

ger

STS

61-A

GL

OM

R G

AS

(DO

D c

lass

ifie

d m

issi

on)

Oct

. 30–

Nov

. 6,1

985

Atl

anti

sST

S 61

-BM

orel

os-B

/PA

M-D

(M

exic

o),A

ussa

t-2/

PAM

-D (

Aus

tral

ia),

Nov

. 26–

Dec

. 3,1

985

Satc

om K

u-2/

PAM

-DII

(R

CA

)C

olum

bia

STS

61-C

Satc

om K

u-1/

PAM

-DII

(R

CA

)Ja

n. 1

2–18

,198

6C

hall

enge

rST

S 51

-LT

DR

S-B

/IU

S an

d Sp

arta

n 20

3 (c

arri

ed b

ut n

ot d

eplo

yed

Jan.

28,

1986

beca

use

of th

e de

stru

ctio

n of

Cha

llen

ger)

D

isco

very

STS-

26T

DR

S-3/

IUS

Sept

. 29–

Oct

. 3,1

988

Atl

anti

sST

S-27

DO

D p

aylo

adD

ec. 2

–6,1

988

* U

niqu

e U

pper

Sta

ge—

Min

utem

an m

issi

le th

ird

stag

e us

ed a

s a

solid

pro

pella

nt p

erig

ee m

otor

.


LAUNCH SYSTEMS 93

Table 2–25. Space Shuttle Main Engine CharacteristicsNumber of Engines Three on each ShuttleThrust 2,000,000 newtons eachOperating Life 7.5 hours and 55 startsRange of Thrust Level 65%–109% of rated power levelPropellant LOX/LH2

Nominal Burn Time 522 sec.Prime Contractor Rockwell International



Tabl

e 2–

26. M

ain

Eng

ine

Dev

elop

men

t an

d Se

lect

ed E

vent

sD

ate

Eve

ntJu

ne 1

980

Surp

asse

d or

igin

al g

oal o

f ac

hiev

ing

80,0

00 s

econ

ds o

f en

gine

test

tim

e be

fore

the

firs

t orb

ital f

light

.Fe

b. 2

0,19

81Fl

ight

rea

dine

ss f

irin

g (2

0-se

cond

fir

ing

of a

ll th

ree

SSM

Es)

,Col

umbi

a(O

V-1

02)

at K

enne

dy S

pace

Cen

ter.

Feb.

28,

1982

Com

plet

ed m

ain

prop

ulsi

on te

st p

rogr

am,N

atio

nal S

pace

Tec

hnol

ogy

Lab

orat

orie

s (N

STL

),M

issi

ssip

pi.

1983

Phas

e II

pro

gram

beg

an f

or im

prov

emen

ts to

SSM

Es

for

incr

ease

d m

argi

n an

d du

rabi

lity.

Dec

. 198

3C

ompl

eted

cer

tific

atio

n of

mai

n en

gine

s at

109

per

cent

of

pres

ent r

ated

pow

er le

vel t

o fu

ll po

wer

leve

l. C

ertif

icat

ion

proc

ess

incl

uded

400

test

s of

mor

e th

an 4

0,00

0 se

cond

s of

sta

tic f

irin

g op

erat

ion.

June

26,

1984

Lau

nch

of S

TS

41-D

pos

tpon

ed in

defi

nite

ly b

ecau

se o

f sh

utdo

wn

of S

SME

s 3

and

2 at

T-4

sec

onds

cau

sed

by s

low

ope

ning

SS

ME

3 m

ain

fuel

val

ve. S

SME

1 n

ever

rec

eive

d a

star

t com

man

d.A

ug. 3

0,19

84ST

S 41

-D c

ondu

cted

suc

cess

fully

.Ju

ly 1

2,19

85ST

S 51

-F la

unch

scr

ubbe

d at

T-3

sec

onds

and

shu

tdow

n of

SSM

Es

beca

use

of lo

ss o

f re

dund

ancy

(ch

anne

l A)

on S

SME

2

cham

ber

cool

ant v

alve

.Ju

ly 2

9,19

85ST

S 51

-F c

ondu

cted

suc

cess

fully

.Ju

ly 1

6,19

8625

0-se

cond

test

con

duct

ed s

ucce

ssfu

lly a

t NST

L. T

he te

st w

as th

e fi

rst i

n a

seri

es to

ver

ify

a m

odif

icat

ion

desi

gned

to e

xten

d th

e op

erat

iona

l ser

vice

life

of

turb

ine

blad

es o

n th

e en

gine

’s h

igh-

pres

sure

oxi

dize

r tu

rbop

ump.

A

ug. 1

3,19

86N

ASA

ann

ounc

ed s

elec

tion

of P

ratt

& W

hitn

ey f

or a

ltern

ate

turb

opum

p de

velo

pmen

t con

trac

t,w

hich

wou

ld p

rovi

de e

xten

ded

life

capa

bilit

y an

d en

hanc

e sa

fety

mar

gins

.D

ec. 1

986

Gro

und

test

pro

gram

initi

ated

.D

ec. 1

986-

Dec

. 198

715

1 te

sts

and

52,3

63 s

econ

ds o

f op

erat

ion

(equ

ival

ent t

o 10

0 Sh

uttle

mis

sion

s) w

ere

perf

orm

ed a

t NST

L (

Mis

siss

ippi

) an

d R

ockw

ell I

nter

natio

nal’s

Roc

ketd

yne

Div

isio

n (C

alif

orni

a).

Aug

. 198

7–Ja

n. 1

988

Acc

epta

nce

test

s at

Ste

nnis

Spa

ce C

ente

r (f

orm

erly

NST

L).

Sept

. 198

7B

egin

ning

of

acce

ptan

ce te

stin

g of

mai

n en

gine

s to

be

used

on

STS-

26 a

t NST

L. A

num

ber

of im

prov

emen

ts w

ere

mad

e on

the

engi

nes

as a

res

ult o

f an

ext

ensi

ve,o

ngoi

ng te

st p

rogr

am.

Jan.

6,1

988

Eng

ine

2016

arr

ived

at K

enne

dy.

Jan.

10,

1988

Eng

ine

2106

inst

alle

d in

num

ber

one

posi

tion

on D

isco

very

.Ja

n. 1

5,19

88E

ngin

e 20

22 a

rriv

ed a

t Ken

nedy

.Ja

n. 2

1,19

88E

ngin

e 20

28 a

rriv

ed a

t Ken

nedy

.Ja

n. 2

4,19

88E

ngin

e 20

22 in

stal

led

in n

umbe

r-tw

o po

sitio

n an

d E

ngin

e 20

28 in

stal

led

in n

umbe

r-th

ree

posi

tion

on D

isco

very

.A

ug. 1

0,19

88C

ondu

cted

a 2

2-se

cond

flig

ht r

eadi

ness

firi

ng o

f D

isco

very

’s m

ain

engi

ne. V

erifi

ed th

at th

e en

tire

Shut

tle s

yste

m w

as r

eady

for

flig

ht


Table 2–27. Space Shuttle External Tank CharacteristicsPropellants LH2, LOXLength 46.8 mDiameter 8.4 mWeight of Propellant 700,000 kgGross Liftoff Weight 750,980 kgInert Weight of Lightweight Tank 30, 096 kgLiquid Oxygen Max. Weight 617,774 kgLiquid Oxygen Tank Volume 542,583 litersLiquid Oxygen Tank Diameter 8.4 mLiquid Oxygen Tank Length 15 mLiquid Oxygen Tank Weight 5,454.5 kg emptyLiquid Hydrogen Max. Weight 103, 257 kgLiquid Hydrogen Tank Diameter 8.4 mLiquid Hydrogen Tank Length 29.46 mLiquid Hydrogen Tank Volume 1,458,228 litersLiquid Hydrogen Tank Weight (Empty) 13,181.8 kgIntertank Length 6.9 mIntertank Diameter 8.4 mIntertank Weight 5,500 kgPrime Contractor Martin Marietta Aerospace

LAUNCH SYSTEMS 95



Tabl

e 2–

28. E

xter

nal T

ank

Dev

elop

men

t an

d Se

lect

ed E

vent

s*D

ate

Eve

ntM

ar. 1

9,19

79Fi

rst e

xter

nal t

ank

leav

es M

arsh

all S

pace

Flig

ht C

ente

r fo

r K

enne

dy S

pace

Cen

ter.

June

25,

1979

Firs

t ext

erna

l tan

k re

ady

for

flig

ht.

Feb.

28,

1980

Succ

essf

ul c

ompl

etio

n of

ful

l dur

atio

n te

st o

f M

PTA

-098

.Ju

ne 3

0,19

80N

ASA

aw

ards

con

trac

t for

ext

erna

l tan

k to

Mar

tin M

arie

tta C

orp.

Oct

. 8,1

980

“All

Syst

ems

Test

”co

nduc

ted.

Nov

. 3,1

980

Firs

t ext

erna

l tan

k m

ated

to S

RB

s fo

r ST

S-1.

Nov

. 11,

1980

Ext

erna

l tan

k an

d SR

Bs

mat

ed to

orb

iter

for

STS-

1.D

ec. 2

,198

0A

ssem

bly

of f

irst

ligh

twei

ght t

ank

begi

ns.

Jan.

17,

1981

Stat

ic f

irin

g at

NST

L. E

xter

nal t

ank

test

with

out a

nti-

geys

er li

ne to

ver

ify

feas

ibili

ty o

f ev

entu

ally

rem

ovin

g

it fr

om la

ter

exte

rnal

tank

ver

sion

s.Ja

n. 2

2 an

d 24

,198

1E

xter

nal t

ank

liqui

d hy

drog

en lo

ad o

f C

olum

bia

at K

enne

dy.

Apr

. 12,

1981

Firs

t tan

k fl

own

succ

essf

ully

.N

ov. 1

2,19

81Se

cond

tank

flo

wn

succ

essf

ully

.O

ct. 1

981

Thi

rd ta

nk d

eliv

ered

to K

enne

dy.

1981

Maj

or w

eldi

ng a

nd s

truc

tura

l ass

embl

y co

mpl

eted

on

the

firs

t pro

duct

ion

vers

ion

of a

ligh

twei

ght t

ank.

Apr

. 4,1

983

Firs

t lig

htw

eigh

t tan

k (L

WT

R 1

) fl

own

on S

TS-

6 m

issi

on; d

esig

n ch

ange

s re

duce

d w

eigh

t of

exte

rnal

tank

by

4,0

00 k

g,pe

rmitt

ing

heav

ier

payl

oad.

Aug

. 1,1

988

Wet

Cou

ntdo

wn

Dem

onst

ratio

n Te

st h

eld;

ext

erna

l tan

k lo

aded

with

liqu

id o

xyge

n an

d liq

uid

hydr

ogen

.*

Rel

ativ

ely

few

eve

nts

wer

e as

soci

ated

with

the

deve

lopm

ent o

f th

e ex

tern

al ta

nk,a

nd th

ere

wer

e no

eve

nts

over

a f

ive-

year

per

iod

from

Apr

il 19

83 to

Aug

ust 1

988.

The

exte

rnal

tank

per

form

ed s

ucce

ssfu

lly o

n th

e ST

S m

issi

ons

duri

ng th

is p

erio

d an

d re

quir

ed li

ttle

atte

ntio

n.


LAUNCH SYSTEMS 97

Tabl

e 2–

29. S

pace

Shu

ttle

Sol

id R

ocke

t B

oost

er C

hara

cter

isti

csL

engt

h45

.5 m

Dia

met

er3.

7 m

Out

side

Dia

met

er o

f N

ozzl

e an

d T

hrus

t12

.4 f

eet

Vec

tor

Con

trol

Sys

tem

Wei

ght a

t Lau

nch

(Eac

h)58

9,68

0 kg

Pr

opel

lant

Wei

ght (

Eac

h)

500,

000

kgIn

ert W

eigh

t (E

ach)

87,2

73 k

gPr

opel

lant

Mix

ture

Am

mon

ium

per

chlo

rate

,alu

min

um,i

ron

oxid

e,a

poly

mer

,an

epox

y cu

ring

age

ntT

hrus

t (Se

a L

evel

) of

Eac

h B

oost

er in

Vac

uum

14,4

09,7

40 n

ewto

ns a

t lau

nch

Sepa

ratio

n M

otor

sFo

ur m

otor

s in

the

nose

fru

stum

and

fou

r m

otor

s in

the

aft s

kirt

Len

gth

0.8

mD

iam

eter

32.5

cm

Thr

ust o

f Se

para

tion

Mot

ors

98,0

78 n

ewto

ns e

ach

Iner

t Wei

ght

8737

3.6

kgB

urn

Tim

e (N

omin

al)

123

sec.

Prim

e C

ontr

acto

rsSR

B m

otor

s:M

orto

n T

hiok

ol C

orp.

SRB

ass

embl

y,ch

ecko

ut,a

nd r

efur

bish

men

t for

all

non–

solid

roc

ket m

otor

com

pone

nts

and

for

SRB

inte

grat

ion:

Boo

ster

Pro

duct

ion

Co.

NA

SA L

ead

Cen

ter

Mar

shal

l Spa

ce F

light

Cen

ter

Rem

arks

Stru

ctur

al m

odif

icat

ions

fol

low

ing

the

Cha

lleng

er a

ccid

ent a

dded

app

roxi

mat

ely

204

kg to

th

e w

eigh

t of

each

SR

B.



Tabl

e 2–

30. C

hron

olog

y of

Sel

ecte

d So

lid

Roc

ket

Boo

ster

Dev

elop

men

t E

vent

sD

ate

Eve

ntJa

n. 3

0,19

79B

egan

orb

iter/

exte

rnal

tank

/SR

B b

urno

ut m

ated

ver

tical

gro

und

vibr

atio

n te

st a

t Mar

shal

l Spa

ce F

light

Cen

ter.

Feb.

17,

1979

Four

th S

RB

fir

ing

at T

hiok

ol,U

tah.

June

15,

1979

Firs

t SR

B q

ualif

icat

ion

firi

ng,T

hiok

ol,1

22 s

econ

ds; n

ozzl

e ex

tens

ion

seve

red

at e

nd o

f ru

n as

in a

ctua

l mis

sion

; ful

l cy

cle

gim

bal.

July

23,

1979

Ent

erpr

ise

(OV

-101

),ex

tern

al ta

nk,a

nd S

RB

s tr

ansp

orte

d on

mob

ile la

unch

er p

latf

orm

fro

m L

aunc

h C

ompl

ex 3

9-A

to

Veh

icle

Ass

embl

y B

uild

ing

at K

enne

dy S

pace

Cen

ter.

Aug

. 197

9Se

cond

SR

B q

ualif

icat

ion

firi

ng,T

hiok

ol.

Feb.

14,

1980

Fina

l qua

lific

atio

n fi

ring

SR

B,T

hiok

ol.

Aug

. 4,1

980

Col

umbi

am

ated

with

SR

Bs

and

exte

rnal

tank

for

ST

S-2.

Nov

. 3,1

980

Ext

erna

l tan

k m

ated

to S

RB

s in

Veh

icle

Ass

embl

y B

uild

ing,

Ken

nedy

,for

ST

S-1.

Nov

. 5,1

980

Ext

erna

l tan

k m

ated

to S

RB

s at

Ken

nedy

.N

ov. 2

6,19

80M

atin

g of

Col

umbi

a(O

V-1

02)

to e

xter

nal t

ank

and

SRB

s in

Veh

icle

Ass

embl

y B

uild

ing

for

STS-

1,K

enne

dy.

Apr

. 20,

1981

SRB

sta

ckin

g be

gan

on m

obile

laun

cher

pla

tfor

m f

or S

TS-

2,K

enne

dy.

July

30,

1981

Star

t mat

ing

of e

xter

nal t

ank

to S

RB

s on

mob

ile la

unch

er p

latf

orm

for

ST

S-2,

Ken

nedy

.A

pr. 1

2,19

81,

STS-

1 an

d ST

S-2

flig

hts

veri

fied

reu

sabi

lity

of S

RB

s; s

ome

rede

sign

of

aft s

kirt

s in

dica

ted.

Nov

. 12,

1981

Sept

. 9,1

981

Col

umbi

am

ated

with

SR

Bs

and

exte

rnal

tank

in p

repa

ratio

n fo

r ST

S-5.

Nov

. 23,

1981

Star

t SR

B s

tack

ing

on m

obile

laun

cher

pla

tfor

m f

or S

TS-

3,K

enne

dy.

Dec

. 19,

1981

Star

t mat

ing

of e

xter

nal t

ank

to S

RB

s on

mob

ile la

unch

er p

latf

orm

for

ST

S-3,

Ken

nedy

.A

pr. 1

6,19

82C

ompl

ete

mat

ing

of S

RB

s an

d ex

tern

al ta

nk f

or S

TS-

4 in

Veh

icle

Ass

embl

y B

uild

ing,

Ken

nedy

.A

pr. 4

,198

3N

ew li

ghtw

eigh

t SR

B c

ase

firs

t flo

wn

on S

TS-

6.A

ug. 3

0,19

83Fi

rst h

igh-

perf

orm

ance

sol

id-f

uele

d ro

cket

mot

or f

low

n on

ST

S-8.

Aug

. 2,1

984

Dis

cove

ry(O

V-1

03)

tran

spor

ted

from

Orb

iter

Proc

essi

ng F

acili

ty to

Veh

icle

Ass

embl

y B

uild

ing

for

rem

ate

with

or

igin

al 4

1-D

SR

Bs

and

exte

rnal

tank

,Ken

nedy

.D

ec. 1

9,19

85ST

S 61

-C,s

even

th f

light

of

Col

umbi

a(O

V-1

02),

laun

ch s

crub

bed

at T

-13

seco

nds

beca

use

of r

ight

hand

SR

B

auxi

liary

pow

er u

nit t

urbi

ne s

yste

m B

ove

rspe

ed.


LAUNCH SYSTEMS 99

Tabl

e 2–

30 c

onti

nued

Dat

eE

vent

Jan.

6,1

986

STS

61-C

laun

ch s

crub

bed

at T

-31

seco

nds

beca

use

of a

laun

ch f

acili

ty li

quid

oxy

gen

repl

enis

h va

lve

prob

lem

.Ja

n. 7

,198

6ST

S 61

-C la

unch

scr

ubbe

d at

T-9

sec

onds

bec

ause

of

adve

rse

wea

ther

con

ditio

ns.

Jan.

12,

1986

STS

61-C

laun

ch c

ondu

cted

suc

cess

fully

.Ja

n. 2

8,19

86ST

S 51

-L la

unch

ed. D

estr

uctio

n of

Cha

llen

ger

and

all c

rew

abo

ard.

Mar

. 25,

1986

Form

atio

n of

Sol

id R

ocke

t Mot

or R

edes

ign

Team

to r

equa

lify

the

SRB

mot

or.

July

–Aug

. 198

6Pr

elim

inar

y R

equi

rem

ents

Rev

iew

s he

ld.

Aug

. 22,

1986

NA

SA a

nnou

nced

the

begi

nnin

g of

a s

erie

s of

test

s de

sign

ed to

ver

ify

the

igni

tion

pres

sure

dyn

amic

s of

the

Spac

e Sh

uttle

sol

id r

ocke

t mot

or (

SRM

) fi

eld

join

t. T

he s

erie

s w

as c

ondu

cted

ove

r th

e ne

xt y

ear

at T

hiok

ol’s

fac

ility

and

at

Mar

shal

l.Se

pt. 5

,198

6St

udy

cont

ract

s aw

arde

d to

fiv

e ae

rosp

ace

firm

s fo

r co

ncep

tual

des

igns

of

an a

ltern

ativ

e or

Blo

ck I

I Sp

ace

Shut

tle S

RM

.Se

pt. 1

986

Prel

imin

ary

Des

ign

Rev

iew

hel

d to

ass

ess

desi

gn r

equi

rem

ents

.O

ct. 2

,198

6N

ASA

ann

ounc

ed th

e de

cisi

on to

test

-fir

e th

e re

desi

gned

SR

M in

a h

oriz

onta

l atti

tude

to b

est s

imul

ate

the

criti

cal

cond

ition

s on

the

fiel

d jo

int t

hat f

aile

d du

ring

the

51-L

mis

sion

.O

ct. 9

,198

6T

rans

fer

of A

tlan

tis

(OV

-104

) m

ated

,min

us S

SME

s,fr

om th

e V

ehic

le A

ssem

bly

Bui

ldin

g to

Lau

nch

Com

plex

39-

B

for

wea

ther

pro

tect

ion

fit c

heck

s,pa

yloa

d ba

y op

erat

ions

,SR

B f

light

rea

dine

ss te

st,t

erm

inal

cou

ntdo

wn

dem

onst

ratio

n te

st,a

nd e

mer

genc

y eg

ress

sim

ulat

ion,

Ken

nedy

.O

ct. 1

6,19

86N

ASA

ann

ounc

ed it

wou

ld p

roce

ed w

ith c

onst

ruct

ing

a se

cond

hor

izon

tal t

est s

tand

for

red

esig

n an

d ce

rtif

icat

ion

of

the

Spac

e Sh

uttle

SR

M a

t the

Thi

okol

fac

ility

. T

he n

ew te

st s

tand

was

des

igne

d to

sim

ulat

e,m

ore

clos

ely

than

the

exis

ting

SRM

sta

nd,t

he s

tres

ses

on th

e SR

M d

urin

g an

act

ual S

huttl

e la

unch

and

asc

ent.

Oct

. 198

6D

esig

n re

quir

emen

ts b

asel

ined

.19

87Pr

imar

y de

sign

cha

nges

mad

e to

the

SRM

fie

ld jo

ints

,noz

zle-

to-c

ase

join

ts,c

ase

insu

latio

n,an

d se

als.

Jan.

198

7R

esul

ts o

f st

udie

s re

latin

g to

inno

vativ

e ch

ange

to th

e ex

istin

g (p

re-C

hall

enge

r) S

RM

join

t des

ign

and

desi

gn o

f ne

w

conc

epts

for

impr

oved

SR

M p

erfo

rman

ce w

ere

repo

rted

to C

ongr

ess.

Mar

. 198

7SR

M A

cqui

sitio

n St

rate

gy a

nd P

lan

subm

itted

to C

ongr

ess.

Pla

n in

dica

ted

that

NA

SA p

ropo

sed

to in

itiat

e Ph

ase

B

(Def

initi

on)

stud

ies

for

an A

dvan

ced

Solid

Roc

ket M

otor

(A

SRM

); S

RM

red

esig

n te

am e

valu

ated

des

ign

alte

rnat

ives

th

at w

ould

min

imiz

e th

e re

desi

gn ti

me

but e

nsur

e ad

equa

te s

afet

y m

argi

ns. T

he te

am c

ondu

cted

ana

lyse

s an

d te

sts

of

the

rede

sign

bas

elin

e.



Tabl

e 2–

30 c

onti

nued

Dat

eE

vent

May

22,

1987

Firs

t in

a se

ries

of

test

fir

ings

con

duct

ed a

t Thi

okol

’s f

acili

ty. O

bjec

tives

of

the

test

,cal

led

the

Noz

zle

Join

t E

nvir

onm

ent S

imul

ator

test

,inc

lude

d di

vers

ifie

d m

otor

sys

tem

ope

ratio

ns,s

uch

as e

valu

atin

g an

d ch

arac

teri

zing

the

SRM

noz

zle-

to-c

ase

join

t,ob

tain

ing

info

rmat

ion

on th

e jo

int d

efle

ctio

n da

ta,a

nd v

alid

atin

g th

e te

st a

rtic

le in

its

orig

inal

des

ign.

May

27,

1987

An

engi

neer

ing

test

mot

or (

ET

M)

was

test

-fir

ed a

t Thi

okol

’s f

acili

ty in

Uta

h as

par

t of

the

Shut

tle m

otor

red

esig

n pr

ogra

m. T

he e

xten

sive

ly in

stru

men

ted

ET

M-1

A w

as s

ucce

ssfu

lly f

ired

for

120

sec

onds

,a f

ull-

dura

tion

test

.A

ug. 2

9,19

87Fi

rst f

ull-

dura

tion

test

fir

ing

of th

e re

desi

gned

SR

M a

t Thi

okol

. Des

igna

ted

DM

-8,t

he 2

-min

ute

test

eva

luat

ed th

e pe

rfor

man

ce o

f th

e m

ajor

fea

ture

s of

the

rede

sign

ed m

otor

and

com

plet

ed s

ever

al te

sts

of c

ase

and

nozz

le-t

o-ca

se

join

ts w

ith in

tent

iona

lly f

law

ed in

sula

tion

and

O-r

ings

.A

ug. 1

987

Five

sol

id p

ropu

lsio

n co

ntra

ctor

s w

ere

awar

ded

cont

ract

s fo

r 9-

mon

th p

relim

inar

y de

sign

and

def

initi

on s

tudi

es o

f bo

th a

mon

olith

ic a

nd s

egm

ente

d A

SRM

that

wou

ld p

erm

it pe

rfor

man

ce in

crea

ses

of u

p to

5,4

43 k

g of

pay

load

.O

ct. 1

987

Cri

tical

Des

ign

Rev

iew

—fi

nal d

esig

n w

as a

ppro

ved.

Dec

. 19,

1987

Seco

nd f

ull-

dura

tion

test

fir

ing

of th

e re

desi

gned

Spa

ce S

huttl

e SR

M a

t Thi

okol

.M

ar. 1

,198

8R

edes

igne

d SR

M s

egm

ents

beg

an a

rriv

ing

at K

enne

dy.

Mar

. 28,

1988

Beg

an s

tack

ing

of D

isco

very

’s S

RM

seg

men

ts b

egin

ning

with

left

aft

boo

ster

.A

pr. 1

988

Full-

dura

tion

test

fir

ing

of r

edes

igne

d so

lid r

ocke

t mot

or (

RSR

M)

at T

hiok

ol’s

fac

ility

in U

tah.

May

5,1

988

Beg

an s

tack

ing

left

hand

boo

ster

seg

men

ts.

May

28,

1988

Com

plet

e st

acki

ng o

f D

isco

very

’s S

RB

s.Ju

ne 1

988

Full-

dura

tion

test

fir

ing

of R

SRM

at T

hiok

ol’s

fac

ility

in U

tah.

June

10,

1988

SRB

s an

d ex

tern

al ta

nk a

re m

ated

for

ST

S-26

; int

erfa

ce te

st b

etw

een

boos

ters

and

ext

erna

l tan

k co

nduc

ted

to v

erif

y co

nnec

tion.

July

198

8So

lid P

ropu

lsio

n In

tegr

ity P

rogr

am c

ondu

cted

mos

t hig

hly

inst

rum

ente

d SR

M n

ozzl

e te

st u

p to

that

tim

e.A

ug. 1

988

Full-

dura

tion

test

fir

ing

of R

SRM

. For

this

test

,pro

duct

ion

veri

fica

tion

mot

or-1

was

ext

ensi

vely

fla

wed

to

dem

onst

rate

the

fail-

safe

cha

ract

eris

tics

of th

e re

desi

gn.

Aug

. 198

8R

FP is

sued

for

des

ign,

deve

lopm

ent,

test

,and

eva

luat

ion

of a

Spa

ce S

huttl

e A

SRM

to r

epla

ce th

e cu

rren

t RSR

M in

th

e m

id-1

990s

. Con

trac

t aw

ard

was

ant

icip

ated

for

the

spri

ng o

f 19

89.

Sept

. 29,

1988

STS-

26 r

etur

ns S

huttl

e to

ope

ratio

nal s

tatu

s.


LAUNCH SYSTEMS 101

Tabl

e 2–

31. U

pper

Sta

ge D

evel

opm

ent

Dat

eE

vent

1979

DO

D’s

det

aile

d de

sign

of

the

two-

stag

e In

ertia

l Upp

er S

tage

(IU

S) c

onfi

gura

tion

com

plet

ed.

1979

Det

aile

d de

sign

of

the

NA

SA th

ree-

stag

e IU

S co

nfig

urat

ion

initi

ated

.19

79D

esig

ns c

ompl

eted

and

qua

lific

atio

n pr

ogra

m in

itiat

ed f

or th

e PA

M.

1979

Mos

t PA

M f

light

har

dwar

e m

anuf

actu

red

and

read

y fo

r as

sem

bly.

1979

NA

SA o

rder

ed P

AM

-As

for

Com

sat’s

Int

elsa

t V c

omm

unic

atio

ns s

atel

lite

mis

sion

s.N

ov. 1

5,19

80Fi

rst f

light

of

PAM

-D o

n th

e D

elta

laun

ched

the

SBS

1 sp

acec

raft

.Se

pt. 1

981

SBS

2 us

ed P

AM

-D o

n D

elta

veh

icle

.N

ov. 1

981

RC

A-S

atco

m 3

-R la

unch

ed u

sing

PA

M-D

.19

82PA

M-A

qua

lific

atio

n an

d pr

oduc

tion

halte

d,pe

ndin

g de

fini

tion

of s

pace

craf

t nee

ds a

nd la

unch

sch

edul

es.

1982

PAM

-D c

ompl

eted

qua

lific

atio

n an

d ve

rifi

catio

n te

sts.

1982

PAM

-D f

lew

six

com

mer

cial

flig

hts

as th

ird

stag

e of

Del

ta E

LV.

May

198

2T

rans

fer

Orb

it St

age

(TO

S) c

once

ptua

l stu

dies

initi

ated

.O

ct. 3

0,19

82Fi

rst I

US

flow

n on

DO

D m

issi

on.

Dec

. 198

2N

ASA

/Orb

ital S

cien

ces

Cor

p. T

OS

Mem

oran

dum

of

Und

erst

andi

ng.

Apr

. 4,1

983

Firs

t IU

S la

unch

ed f

rom

Spa

ce S

huttl

e on

ST

S-6,

carr

ying

the

Tra

ckin

g an

d D

ata

Rel

ay S

atel

lite

(TD

RS)

. Sec

ond

stag

e fa

iled

to p

lace

sat

ellit

e in

fin

al g

eosy

nchr

onou

s or

bit.

Add

ition

al m

aneu

vers

pla

ced

TD

RS

1 in

its

requ

ired

fu

nctio

nal o

rbit.

NA

SA-A

ir F

orce

team

det

erm

ined

the

IUS

prob

lem

was

in th

e gi

mba

l mec

hani

sm o

f th

e se

cond

sta

ge.

1983

PAM

-D la

unch

ed n

ine

com

mun

icat

ions

sat

ellit

es,t

hree

fro

m th

e Sp

ace

Shut

tle’s

car

go b

ay a

nd s

ix f

rom

ELV

s.A

pr. 1

983

NA

SA a

nd O

rbita

l Sci

ence

s si

gned

a jo

int a

gree

men

t for

com

mer

cial

dev

elop

men

t of

the

TO

S.M

ay 1

983

TO

S de

sign

stu

dies

initi

ated

.O

ct. 1

983

TO

S fu

ll-sc

ale

deve

lopm

ent i

nitia

ted.

Feb.

3,1

984

PAM

fai

led

to b

oost

Wes

tar

6 an

d Pa

lapa

B-2

to p

rope

r or

bit o

n ST

S 41

-B m

issi

on.

May

198

4T

OS

Prel

imin

ary

Des

ign

Rev

iew

.Ju

ne 1

984

Las

er in

itial

nav

igat

ion

syst

em d

evel

opm

ent b

egun

by

Hon

eyw

ell f

or u

se in

upp

er s

tage

s.D

ec. 1

984

Orb

ital S

cien

ces

and

Mar

tin M

arie

tta s

ign

a de

velo

pmen

t con

trac

t for

TO

S/A

poge

e M

aneu

veri

ng S

tage

.



Tabl

e 2–

31 c

onti

nued

Dat

eE

vent

Mar

. 198

5T

OS

Cri

tical

Des

ign

Rev

iew

.Ju

ne 1

985

Con

trac

t aw

arde

d to

Orb

ital S

cien

ces

for

the

lase

r in

itial

nav

igat

ion

syst

em.

Aug

. 198

5T

OS

Prod

uctio

n R

eadi

ness

Rev

iew

; fac

tory

rol

lout

of

firs

t TO

S up

per

stag

e.N

ov. 2

6,19

85PA

M D

II u

sed

on S

TS

61-B

.Ja

n. 1

2,19

86PA

M D

II u

sed

on S

TS

61-C

.Fe

b. 1

986

Boe

ing

Aer

ospa

ce s

elec

ted

to p

rovi

de u

pper

sta

ge f

or T

DR

S-E

and

-F.

Mar

. 198

6M

ars

Obs

erve

r T

OS

cont

ract

sel

ectio

n.Ju

ne 1

9,19

86Te

rmin

atio

n of

Cen

taur

upp

er s

tage

dev

elop

men

t.N

ov. 2

6,19

86N

ASA

ann

ounc

ed th

e se

lect

ion

of th

e IU

S as

the

base

line

optio

n fo

r th

ree

plan

etar

y m

issi

ons:

Gal

ileo,

Mag

ella

n,an

d U

lyss

es.

Nov

. 26,

1986

The

TO

S w

as s

elec

ted

to p

lace

the

Mar

s O

bser

ver

spac

ecra

ft in

to th

e pr

oper

inte

rpla

neta

ry tr

ajec

tory

.19

89M

artin

Mar

ietta

and

Boe

ing

chos

en to

con

duct

stu

dies

on

spac

e tr

ansf

er c

once

pts,

succ

esso

r to

the

Orb

ital

Tra

nsfe

r V

ehic

le.

July

14,

1993

Adv

ance

d C

omm

unic

atio

ns T

echn

olog

y Sa

telli

te (

AC

TS)

laun

ched

fro

m S

huttl

e w

ith T

OS

for

tran

sfer

to h

ighe

r or

bit.


LAUNCH SYSTEMS 103

Table 2–32. Transfer Orbit Stage CharacteristicsLength 3.3 mWeight With Full Propellant Load 10,886 kgAirborne Support Equipment Weight 1,450 kgPayload to Geotransfer Orbit 6,080 kg from ShuttlePayload to Planetary and High-Energy Orbits 5,227 kg from Titan III and IVPropulsion System Orbis 21 solid rocket motor

and attitude control systemCapacity 1,360 kg to 3,175 kg capacity
