Apollo vehicle propulsion systems.
H-1, F-1, J-2 and RL-20 Apollo vehicle propulsion systems must have preflight design maturity due to man-rating requirement
PROPULSION SYSTEM PERFORMANCE↗
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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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RL10A-3-3 rocket engine Design report
Design features of RL-10 A-3 liquid propellant rocket engine
RL-10 A-3 ROCKET ENGINE↗
Development of LOX-hydrogen engines for the Saturn Apollo launch vehicles.
RL-10 and J-2 hydrogen oxygen engine design and development for Saturn Apollo launch vehicles, noting performance parameters from ground and flight tests
Burks, A. J.↗
FLOX/methane pump-fed engine study
FLOX and methane rocket propellant for use with modified RL-10-A-1 engine
Bailey, T. E.↗
FLOX/methane pump-fed engine systems.
Flox/methane pump fed engine, discussing RL-10-1-A engine tests and optimized 5000 lb thrust engine design
Brooke, A. W.↗
Low speed inducers for cryogenic upper stages
Briefing charts are presented, which were used in an oral presentation of the results and recommendations for the design and analysis of low speed hydrogen and oxygen inducers and their drive systems applicable to the space tug. A discussion of the design of the 15K and RL-10 inducers is included.
King, J. A.↗
Propulsion system tests on a full scale Centaur vehicle to investigate 3-burn mission capability of the D-lT configuration
Propulsion system tests were conducted on a full scale Centaur vehicle to investigate system capability of the proposed D-lT configuration for a three-burn mission. This particular mission profile requires that the engines be capable of restarting and firing for a final maneuver after a 5-1/2-hour coast to synchronous orbit. The thermal conditioning requirements of the engine and propellant feed system components for engine start under these conditions were investigated. Performance data were also obtained on the D-lT type computer controlled propellant tank pressurization system. The test results demonstrated that the RL-10 engines on the Centaur vehicle could be started and run reliably after being thermally conditioned to predicted engine start conditions for a one, two and three burn mission. Investigation of the thermal margins also indicated that engine starts could be accomplished at the maximum predicted component temperature conditions with prestart durations less than planned for flight.
Groesbeck, W. A.↗
Centaur feedline dynamics study using power spectral methods
Tests were conducted to determine the dynamic characteristics of the Centaur/RL-10 oxygen and hydrogen feedlines. The fundamental-mode resonant frequencies were determined by applying power spectral methods to noise-generated data from hot firings of the RL-10 engine. The effect of net positive suction pressure of the main feed pumps on resonant frequency characteristics was determined to be a straight-line relation. Power spectral methods were also used to determine the dynamic characteristics of the boost pumps.
Lorenzo, C. F.↗
The spectrum of cosmic electron with energies between 6 and 100 GeV
This experiment was carried out during three balloon flights which provided a total exposure of 3500 + or - 60 sq m sec sterad at an average depth of 4.8 g/sq cm The detector, in which the development of cascade showers in a 33.7 rl absorber was sampled by 10 scintillation counters and 216 Geiger-Muller tubes, was calibrated at the Cornell Electron Synchrotron, the separation of cosmic electrons from the nuclear background was confirmed by extensive analysis of data from the flights, from the calibration and from ground level exposure. The spectral intensity of primary cosmic ray electrons were found in particles/sq m sec sterad GeV. Similarly, the ground level spectrum of secondary cosmic ray electrons was also found. The steepness of the spectrum of cosmic electrons relative to that of nuclei implies one of the following conclusions: either the injection spectrum of electrons is steeper than that of nuclei, or the electron spectrum has been steepened by Compton/synchrotron losses in the energy range covered by the experiment.
Meegan, C. A.↗
Space transportation systems, 1980-2000
An overview of projected space transportation systems for the 1980-2000 period is presented. Methods for meeting the requirements in that area are documented through a description of the Space Shuttle and its current upper stages, reusable transfer vehicles, Shuttle uprating, fully reusable and single-stage transports, heavy lift vehicles, and advanced vehicles for orbital transfer and lunar landing. The structure of the liquid oxygen/liquid hydrogen boosters, cryogenic propellants, whose main propulsion system is a derivative of the RL-10 engine used in the Centaur launch, and mixed mode propulsion engine is also described, as is the flyback booster concept and such related issues as propellants and rocket engines, staging velocity, configuration shape, and whether it should be manned or unmanned. Especially noted are proposed propulsion systems going beyond the present Shuttle such as the nuclear and electric propulsion concepts, the resistojet, using liquid hydrogen, ion thruster, using argon, magnetoplasmadynamic thrusters, solar thermal rocket, and solar sail concept where the basic idea is to use a 'sail' of thin reflective material to generate a net force by reflection of solar radiation.
Cohan, C. J.↗
Orbit transfer vehicle engine study. Volume 1: Executive summary
The orbit transfer vehicle (OTV) engine study provided parametric performance, engine programmatic, and cost data on the complete propulative spectrum that is available for a variety of high energy, space maneuvering missions. Candidate OTV engines from the near term RL 10 engine to advanced high performance expander engines were examined.
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Transient solution for megajoule energy release in a lumped-parameter series RLC circuit
A method is developed for optimizing the energy release from a megajoule capacitive discharge in a series RLC circuit with an RL load. Both the resistance and inductance of the load are represented by effective values that characterize their behavior during the discharge. Using Kirchhoff's laws, equations utilizing the load impedance and the external circuit impedance are derived for determining the instantaneous load voltage and energy characteristics. A program (ERES) computes and displays the load characteristics and the circuit current. Use of the ERES program allows a designer to perturbate values of the circuit elements in order to produce the desired time distribution for the load energy input.
Barnes, G.↗
Ensuring the relocatability of programs in the operational system DOS YeS
Specific modifications in the Disk Operational System Unified Series to insure the relocatability of programs stored permanently in the core image library is described. A self-relocating method for loading programs into the working memory with re-editing all the programs recorded in the core image library is presented. The modified linkage editor can be included in a relocation dictionary containing data about each address constant at the assembly stage at the request of the programmer. The relocation dictionary increases the dimension of the RL-phase in comparison with the dimension of this same phase when edited by the standard method, making possible the creation of multiphase program complexes. Generation and use of the modified system using Assembly language is described. An example of the use of the system is given, and limitations of the use of the relocatable programs in the modified system are outlined.
Novoseltsev, S. K.↗
Advanced LOX/H sub 2 engine technologies for future OTVs
The optimization methods and trade studies which were used to define a 1980 state-of-the-art design of orbital transfer vehicles are summarized. The advanced thrust chamber, turbomachinery, and engine power cycle technologies recommended for development and verification during the 1980's are discussed. The state-of-the-art design serves as one of the reference points for evaluation of the advanced technology concepts. A second reference point is provided by the 1960 state-of-the-art RL-10 engine. Advanced heat energy extraction concepts in the combustor and injector are presented that will permit the high chamber pressures expected of the advanced expander cycle engine concept. Advanced turbomachinery concepts are selected that will help utilize efficiently the heat energy extracted in the thrust chamber and aid in uprating chamber pressures to values five times those realized with 1960 to 1980 technologies. The higher chamber pressure benefits are maximized with nozzle concepts that will package within a short retracted length and extend to the very high expansion area ratios required for specific impulse values significantly higher than those now realized by operational systems. Control system component concepts studied will allow utilization of the high turbine drive energy levels in a reliable repeatable manner that will reduce program risk and enhance overall space transportation system mission life cycle cost.
Glass, J.↗
Benefits and costs of low thrust propulsion systems
The results of costs/benefits analyses of three chemical propulsion systems that are candidates for transferring high density, low volume STS payloads from LEO to GEO are reported. Separate algorithms were developed for benefits and costs of primary propulsion systems (PPS) as functions of the required thrust levels. The life cycle costs of each system were computed based on the developmental, production, and deployment costs. A weighted criteria rating approach was taken for the benefits, with each benefit assigned a value commensurate to its relative worth to the overall system. Support costs were included in the costs modeling. Reference missions from NASA, commercial, and DoD catalog payloads were examined. The program was concluded reliable and flexible for evaluating benefits and costs of launch and orbit transfer for any catalog mission, with the most beneficial PPS being a dedicated low thrust configuration using the RL-10 system.
Robertson, R. I.↗
Shuttle/Centaur - More capability for the 1980's
Design features of the Centaur upper stage for the Shuttle are described, noting interfaces with the Orbiter and intended missions. The Shuttle will carry the Centaur stage into a 241 km eastward orbit, open the payload doors, and by the fourth orbit rotate the Centaur 45 deg so it points out of the bay. An integrated support system will limit the actual equipment added to the Orbiter to 122 kg. Separation from the Orbiter will be effected by a spring-loaded mechanism that will impart a 1/3 m/sec velocity to the Centaur, which carries its own LOX/LH2 fuel supply for two RL 10A-3-3A engines. The fuel is moved to the bottom of the tanks by auxiliary thrusters which propel the Centaur forward. Planned missions for the Shuttle-Centaur are boosting the ESA Solar Polar Mission and launching the Galileo probe in 1986, possibly followed by a Venus radar mapper mission in 1988.
Spurlock, O. F.↗
Characterization of CNRS Fizeau wedge laser tuner
A fringe detection and measurement system was constructed for use with the CNRS Fizeau wedge laser tuner, consisting of three circuit boards. The first board is a standard Reticon RC-100 B motherboard which is used to provide the timing, video processing, and housekeeping functions required by the Reticon RL-512 G photodiode array used in the system. The sampled and held video signal from the motherboard is processed by a second, custom fabricated circuit board which contains a high speed fringe detection and locating circuit. This board includes a dc level discriminator type fringe detector, a counter circuit to determine fringe center, a pulsed laser triggering circuit, and a control circuit to operate the shutter for the He-Ne reference laser beam. The fringe center information is supplied to the third board, a commercial single board computer, which governs the data collection process and interprets the results.
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Design study of an integrated aerobraking orbital transfer vehicle
An aerobraking orbital transfer vehicle (AOTV) concept, which has an aerobrake structure that is integrated with the propulsion stage, is discussed. The concept vehicle is to be assembled in space and is space-based. The advantages of aeroassist over an all propulsive vehicle are discussed and it is shown that the vehicle considered is very competitive with inflatable and deployable concepts from mass and performance aspects. The aerobrake geometry is an ellipsoidally blunted, raked-off, elliptical wide-angle cone with a toroidal skirt. Propellant tanks, engines, and subsystems are integrated into a closed, isogrid aerobrake structure which provides rigidity. The vehicle has two side-firing, gimbaled RL-10 type engines and carries 38,000 kg of useable propellant. The trajectory during aerobraking is determined from an adaptive guidance logic, and the heating is determined from engineering correlations as well as 3-D Navier-Stokes solutions. The AOTV is capable of placing 13,500 kg payload into geosynchronous Earth orbit (GEO) or carrying a LEO-GEO-LEO round-trip payload of 7100 kg. A two-stage version considered for lunar missions results in a lunar surface delivery capability of 18,000 kg or a round-trip capability of 6800 kg with 3860 kg delivery-only capability.
Scott, C. D.↗