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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 145 records · Page 8
All sourcesNASA NTRSDOE OSTI
NASA NTRS1971

High temperature insulation for reradiative TPS.

This paper presents the results of a program to evaluate thermal insulations for use under the radiative TPS of a shuttle orbiter vehicle. Evaluations were made for survival under the multimission requirements of shuttle acoustic/vibration loads and thermal profiles. During the first year of the program a total of 1100 specimen-cycles were completed using metal foil packaged refractory fiber insulations.

Hughes, T. A.↗
NASA NTRS1972

Mission analysis for design of a SEP multi-mission spacecraft.

Study of the multimission capability of a solar electric propulsion spacecraft. The study includes such factors as spacecraft constraints, payload variations with launch opportunity, propellant on-loading and off-loading requirements, and limitations in thrusting time, throttling range and steering. The study determined that a single spacecraft design, with P sub zero = 12 kW and I sub s = 3000 sec launched by a Titan IIIC or Titan IIID/Centaur can be used to conduct at least five diverse missions: Eros rendezvous, 1 AU out-of-ecliptic, Mercury orbiter, Jupiter flyby/probe, and Jupiter orbiter. Flight time and injection velocity vary as needed to achieve the desired payload capability. The mission analysis approach and results are presented.

Driver, J. M.↗
NASA NTRS1972

The solar electric propulsion stage concept for high energy missions.

Definition of multimission and engine performance requirements for candidate solar electric propulsion stage configurations, considering launch vehicle compatibility, electric propulsion integration, payload requirements, and the effects of environmental extremes. Electric propulsion power options include two solar array power levels (15/22 kW), up to twelve electric thrustors of 30 cm diameter and 2.7 kW each, five to eight power conditioning units, and a maximum mercury propellant capacity of 1530 kg. In performance, the stage with a dry weight of 700 to 900 kg can deliver a net mass of 756 kg into Saturn orbit, 329 kg into a tight Mercury orbit, and 334 kg within 0.1 AU of the sun. The stage can also deliver a round trip payload of 3350 kg to geosynchronous orbit and return from an intermediate elliptical orbit using the Shuttle/Tug. Thus, a versatile stage is developed which competes effectively in performance with existing integrated spacecraft and promises considerable savings in total program costs.

Guttman, C. H.↗
NASA NTRS1971

Expendable Second Stage Reusable Space Shuttle Booster: Preliminary System Specification - Volume 9

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.

Source record↗
NASA NTRS1973

The development and demonstration of hybrid programmable attitude control electronics

In the course of extended life attitude control system (ELACS) research sponsored by NASA a hybrid programable attitude control electronics (HYPACE) concept was developed and demonstrated. The wide variety of future planetary missions demanded a new control approach to accommodate the automatic fault tolerance and long the life requirements of such missions. HYPACE provides an adaptable, analog/digital design approach that permits preflight and in-flight accommodation of mission changes, component performance variations, and spacecraft changes, through programing. This enabled broad multimission flexibility of application in a cost effective manner. Previously, flight control computers have not been not flown on planetary missions because of weight and power problems. These problems were resolved in the design of HYPACE. The HYPACE design, which was demonstrated in breadboard form on a single-axis gas-bearing spacecraft simulation, uses a single control channel to perform the attitude control functions sequentially, thus significantly reducing the number of component parts over hard-wired designs.

Smith, L. S.↗
NASA NTRS1973

Processing of Rigidized Rei-mullite Insulative Composites

Systematic development and evaluation of ceramic fiber Mullite are summarized: (1) Major reductions in thermal protection system weight have been achieved by reducing the density and thermal conductivity of the insulation by 20 and 25 percent; (2) already adequate structural margins-of-safety have been greatly enhanced by increasing the tensile strength and strain-to-failure capabilities of the insulation by factors of 3 and 2; (3) cost effectiveness has been increased through the achievement of a high degree of uniformity and reproducibility of properties and through process simplification and binder modifications; and (4) maximization of multimission capability at surface temperatures of 1644 K has been achieved through firing cycle adjustments and the development of a material with high dimensional stability.

Gebhardt, J. J.↗