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Contract end item specifications required to define hardware end items for the expendable second stage on a reusable shuttle booster system are provided. The specifications are as follows: (1) expendable second stage, (2) space shuttle booster, and (3) space shuttle ground support system.
The specification for establishing the requirements for the system performance, design, development, and ground and flight operations of the expendable second stage on a reusable space shuttle booster system is presented. The basic specification is that the system shall be capable of placing payloads in excess of 100,000 pounds into earth orbit. In addition, the expendable second stage provides a multimission, economical, large capability system suitable for a variety of space missions in the 1980 time period.
Results are presented for an experimental space shuttle stage separation plume impingement program conducted in the NASA-Marshall Space Flight Center's impulse base flow facility (IBFF). Major objectives of the investigation were to: (1)determine the degree of dual engine exhaust plume simulation obtained using the equivalent engine; (2) determine the applicability of the analytical techniques; and (3) obtain data applicable for use in full-scale studies. The IBFF tests determined the orbiter rocket motor plume impingement loads, both pressure and heating, on a 3 percent General Dynamics B-15B booster configuration in a quiescent environment simulating a nominal staging altitude of 73.2 km (240,00 ft). The data included plume surveys of two 3 percent scale orbiter nozzles, and a 4.242 percent scaled equivalent nozzle - equivalent in the sense that it was designed to have the same nozzle-throat-to-area ratio as the two 3 percent nozzles and, within the tolerances assigned for machining the hardware, this was accomplished.
A conceptual model and an empirical application of the relationship between the manner of selecting observations and its effect on the precision of estimates from remote sensing are reported. This three stage sampling scheme considers flightlines, segments within flightlines, and units within these segments. The error of estimate is dependent on the number of observations in each of the stages.
The natural environment design requirements for the solar electric propulsion stage are presented. Environment criteria for the SEP stage will cover earth orbital operations out to geosynchronous altitudes and also interplanetary missions including comet and asteroid missions.
Discussion of the effects of capillarity-induced stresses on dislocations during early stage sintering. A special version of Hirth's (1963) theoretical calculation procedures modified to describe dislocation nucleation on planes meeting the sintering body's neck surface obliquely is shown to predict plastic flow at stress levels know to exist between micron size metal particles in the early stages of sintering.
The overall and blade-element performances are presented over the stable flow operating range from 50 to 100 percent of design speed. Stage peak efficiency of 0.834 was obtained at a weight flow of 26.4 kg/sec (58.3 lb/sec) and a pressure ratio of 1.581. The stall margin for the stage was 7.5 percent based on weight flow and pressure ratio at stall and peak efficiency conditions. The rotor minimum losses were approximately equal to design except in the blade vibration damper region. Stator minimum losses were less than design except in the tip and damper regions.
Performance was obtained for a 50-cm-diameter compressor designed for a high weight flow per unit annulus area of 208 (kg/sec)/sq m. Peak efficiency values of 0.83 and 0.79 were obtained for the rotor and stage, respectively. The stall margin for the stage was 23 percent, based on equivalent weight flow and total-pressure ratio at peak efficiency and stall.
Auger electron spectroscopy was used to examine the initial stages of oxidation of a polycrystalline copper - 19.6 a/o-aluminum alloy. The growth of the 55-eV aluminum oxide peak and the decay of the 59-, 62-, and 937-eV copper peaks were examined as functions of temperature, exposure, and pressure. Pressures ranged from 1x10 to the minus 7th power to 0.0005 torr of O2. Temperatures ranged from room temperature to 700 C. A completely aluminum oxide surface layer was obtained in all cases. Complete disappearance of the underlying 937-eV copper peak was obtained by heating at 700 C in O2 at 0.0005 torr for 1 hr. Temperature studies indicated that thermally activated diffusion was important to the oxidation studies. The initial stages of oxidation followed a logarithmic growth curve.
Tests were conducted on a high-tip-speed, low-loading transonic fan stage to determine the performance and inlet flow distortion tolerance of the design. The fan was designed for high efficiency at a moderate pressure ratio by designing the hub section to operate at minimum loss when the tip operates with an oblique shock. The design objective was an efficiency of 86 percent at a pressure ratio of 1.5, a specific flow (flow per unit annulus area) of 42 lb/sec-sq. ft (205.1 kgm/sec-m sq), and a tip speed of 1600 ft/sec (488.6 m/sec). During testing, a peak efficiency of 84 percent was achieved at design speed and design specific flow. At the design speed and pressure ratio, the flow was 4 percent greater than design, efficiency was 81 percent, and a stall margin of 24 percent was obtained. The stall line was improved with hub radial distortion but was reduced when the stage was tested with tip radial and circumferential flow distortions. Blade-to-blade values of static pressures were measured over the rotor blade tips.
The overall and blade-element performance of a low-solidity, low-pressure-ratio, low-tip-speed fan stage is presented over the stable operating range at rotative speeds from 90 to 120 percent of design speed. At design speed a stage peak efficiency of 0.836 was obtained at a weight flow of 30.27 kilograms per second and a pressure ratio of 1.111. The pressure ratio was less than design pressure ratio, and the design energy input into the rotor was not achieved. A mismatch of the rotor and stator blade elements resulted due to the lower than design pressure ratio of the rotor.
Program SEPSIM is a FORTRAN program which performs deployment, servicing, and retrieval missions to synchronous equatorial orbit using a space tug with a continuous low thrust upper stage known as a solar electric propulsion stage (SEPS). The SEPS ferries payloads back and forth between an intermediate orbit and synchronous orbit, and performs the necessary servicing maneuvers in synchronous orbit. The tug carries payloads between the orbiter and the intermediate orbit, deploys fully fueled SEPS vehicles, and retrieves exhausted SEPS vehicles when, and if, required. The program is presently contained in subroutine form in the Logistical On-orbit VEhicle Servicing (LOVES) Program, but can also be run independently with the addition of a simple driver program.
A two-stage fan was designed to reduce noise 20 db below current requirements. The first-stage rotor has a design tip speed of 365.8 m/sec and a hub/tip ratio of 0.4. The fan was designed to deliver a pressure ratio of 1.9 with an adiabatic efficiency of 85.3 percent at a specific inlet corrected flow of 209.2kg/sec/sq m. Noise reduction devices include acoustically treated casing walls, a flowpath exit acoustic splitter, a translating centerbody sonic inlet device, widely spaced blade rows, and the proper ratio of blades and vanes. Multiple-circular-arc rotor airfoils, resettable stators, split outer casings, and capability to go to close blade-row spacing are also included.
The features of a solar electric propulsion (SEP) stage are examined. An SEP stage (SEPS) contains all of the subsystems required to deliver separable payloads to the desired target and to support its own operations. These include attitude control, communications, data handling, and computer subsystems in addition to the solar array and thrust subsystems required for electric propulsion. SEPS can provide subsystem support to attached payloads for comet flyby and asteroid rendezvous. It can accommodate both planetary and geosynchronous missions by the use of mission-peculiar modules. In addition SEPS can assist the shuttle tug transportation system in earth-orbital applications. Aspects of mission analysis are discussed together with details of the electric propulsion system including solar array considerations.
Application of the SEP Stage (SEPS) for earth orbital missions in the Shuttle era has been investigated for the first time. In particular, the geosynchronous payload delivery/retrieval mission was evaluated using SEPS as the upper stage of the Shuttle/Tug system. Results show that SEPS can significantly increase the payload capability of the Shuttle/Tug, especially for the low-technology Tug. The missions can be accomplished using a modified planetary SEPS concept. Use of SEPS can significantly reduce the Shuttle/Tug launches to accomplish projected geosynchronous traffic requirements.
This paper summarizes the analyses and design studies performed for the Solar Electric Propulsion Stage (SEPS). Mission analysis results were used to establish general design requirements. A performance sensitivity analysis was then used to define an operational power level of 21 kW for accomplishing a broad range of interplanetary and earth-orbital missions. With this base, propulsion system and stage subsystem design analyses were performed to support concept and configuration studies. Design study areas include solar array, thrust subsystem, thermal control, structures, and astrionics. Results are presented primarily for areas not covered by more specific technical papers (i.e., thrust subsystem, solar array, design requirements and configuration studies).
The need for larger diameter fairings to encapsulate spacecraft designed to use Delta's increased lifting capabilities resulted in hammer-head design configurations that would have unaccepably degraded vehicle flight characteristics. This situation led to the development of the Straight Eight configuration. The towed mode arrangement for the second stage negated the requirement for the engine compartment structure, thus establishing the potential for an improved packaging of the hydraulic and pneumatic systems. Therefore, with cost reduction and reliability improvement objectives, a second stage propulsion system improvement program was initiated. The paper will describe these improvements and their system benefits.
The application of restartable solid motor stages to shuttle missions has been shown to provide a viable supplement to the shuttle program. Restartable solid motors in the 3000 pound class provide a small expendable transfer stage that reduces the demand on the shuttle for the lower energy missions. Shuttle operational requirements and preliminary performance data provided an input for defining design features required for restartable solid motor applications. These data provided a basis for a configuration definition that is compatible with shuttle operations. Mission by mission analysis showed the impact on a NASA supplied mission model. The results showed a 15% reduction in the number of shuttle flights required. In addition the amount of shuttle capability used to complete the mission objectives was significantly reduced. For example, in the 1979 missions there was a 62% reduction in shuttle capability used. The study also showed that the solid motor could provide a supplement to the TUG that would allow TUGS to be used in a recoverable rather than an expendable mode. The study shows a 71% reduction in the number of TUGs that would be expended.