Optimum configuration of high-lift aeromaneuvering orbital transfer vehicles in viscous flow
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Davies, Carol B.↗
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The superfluid helium on-orbit transfer (SHOOT) flight experiment
The SHOOT flight demonstration is being undertaken to verify component and system level technology necessary to resupply large superfluid helium dewars in space. The baseline configuration uses two identical 210 liter dewars connected by a transfer line which contains a quick disconnect coupling. The helium will be transferred back and forth between the dewars under various conditions of flow rate, parasitic heat load, and temperature. An astronaut Extra-vehicular Activity (EVA) is also planned to manually demate and mate the coupling. A number of components necessary for the flight are being developed. These components are described here.
Dipirro, Michael J.↗
Cryogenic and thermal design for the Superfluid Helium On-Orbit Transfer (SHOOT) experiment
The analysis and trade-offs of the external thermal design of the two 200-liter dewars required in the SHOOT experiment to extend space mission life by superfluid helium replenishment are discussed. Also considered are the support electronics and the optimization and prediction of the performance of the dewar and cryostat assemblies. Particular attention is given to the ground-hold and standby performance of the dewars, along with the temperature of the helium bath during high-flow-rate helium transfers.
Lee, J. H.↗
Optimal reentry guidance for aeroassisted orbit transfer vehicles
A three-state model is presented for analyzing the problem of optimal changes in heading with minimum energy loss for a hypersonic gliding vehicle. A further model order reduction to a single state model is examined using singular perturbation theory. The optimal solution for the reduced problem defines an optimal altitude profile dependent on the current energy of the vehicle. A separate boundary-layer analysis is used to account for altitude and flight path angle dynamics, and to obtain lift and bank angle control solutions. By considering alternative approximations to solve the boundary-layer problem, three guidance laws are obtained, each having a feedback form. The guidance laws are evaluated for a hypothetical vehicle, and compared to an optimal solution obtained using a multiple shooting algorithm.
Calise, Anthony J.↗
Orbital transfer vehicle concept definition and system analysis study. Volume 10: Aerocapture for manned Mars missions
A manned expedition to Mars has been under consideration as a potential mission for the early 21st century. The necessarily large vehicle requirements have sparked interest in aerocapture as a means of reducing propellant usage. This volume summarizes the work performed to establish concepts and feasibility of such a mission which makes maximum use of aeroassist maneuvers.
Willcockson, W. H.↗
Orbital transfer vehicle concept definition and system analysis study, 1985. Volume 2: OTV concept definition and evaluation. Book 4: Operations
The benefits of the reusable Space Shuttle and the advent of the new Space Station hold promise for increasingly effective utilization of space by the scientific and commercial as well as military communities. A high energy reusable oribital transfer vehicle (OTV) represents an additional capability which also exhibits potential for enhancing space access by allowing more ambitious missions and at the same time reducing launch costs when compared to existing upper stages. This section, Vol. 2: Book 4, covers launch operations and flight operations. The launch operations section covers analyses of ground based and space based vehicles, launch site facilities, logistics requirements, propellant loading, space based maintenance and aft cargo carrier access options. The flight operations sections contain summary descriptions of ground based and space based OTV missions, operations and support requirements, and a discussion of fleet implications.
Mitchell, Jack C.↗
Orbital transfer vehicle concept definition and system analysis study, 1986. Volume 9: Study extension results
The purpose of this extension to the OTV Concept Definition and Systems Analysis Study was to improve the definition of the OTV Program that will be most beneficial to the nation in the 1995 to 2010 timeframe. The implications of the defined mission and defined launch vehicle are investigated. The key mission requirements identified for the Space Transportation Architecture Study (STAS) were established and reflect a need for early capability and more ambitious capability growth. The key technical objectives and related issues addressed are summarized. The analyses of selected areas including aerobrake design, proximity operations, and the balance of EVA and IVA operations used in the support of the OTV at the space-base were enhanced. Sensitivity studies were conducted to establish how the OTV program should be tailored to meet changing circumstances.
Kofal, Allen E.↗
Orbital transfer vehicle concept definition and system analysis study. Volume 2: OTV concept definition and evaluation. Book 3: Subsystem trade studies
The technical trade studies and analyses reported in this book represent the accumulated work of the technical staff for the contract period. The general disciplines covered are as follows: (1) Guidance, Navigation, and Control; (2) Avionics Hardware; (3) Aeroassist Technology; (4) Propulsion; (5) Structure and Materials; and (6) Thermal Control Technology. The objectives in each of these areas were to develop the latest data, information, and analyses in support of the vehicle design effort.
Dickman, Glen J.↗
Orbit transfer rocket engine technology program: Oxygen materials compatibility testing
Particle impact and frictional heating tests of metals in high pressure oxygen, are conducted in support of the design of an advanced rocket engine oxygen turbopump. Materials having a wide range of thermodynamic properties including heat of combustion and thermal diffusivity were compared in their resistance to ignition and sustained burning. Copper, nickel and their alloys were found superior to iron based and stainless steel alloys. Some materials became more difficult to ignite as oxygen pressure was increased from 7 to 21 MPa (1000 to 3000 psia).
Schoenman, Leonard↗
Unmanned, space-based, reusable orbital transfer vehicle, DARVES. Volume 2: Data and calculations
Results for trajectory analysis, fuel analysis, and analysis of aerobraking maneuvers are presented. Calculations are included for thickness and mass of the propellant tanks; determination of weight (avionics module); numerical integration program; hydrostatic pressure due to acceleration; and propellant tank pressurization.
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Performance of large area xenon ion thrusters for orbit transfer missions
Studies have indicated that xenon ion propulsion systems can enable the use of smaller Earth-launch vehicles for satellite placement which results in significant cost savings. These analyses have assumed the availability of advanced, high power ion thrusters operating at about 10 kW or higher. A program was initiated to explore the viability of operating 50 cm diameter ion thrusters at this power level. Operation with several discharge chamber and ion extraction grid set combinations has been demonstrated and data were obtained at power levels to 16 kW. Fifty cm diameter thrusters using state of the art 30 cm diameter grids or advanced technology 50 cm diameter grids allow discharge power and beam current densities commensurate with long life at power levels up to 10 kW. In addition, 50 cm diameter thrusters are shown to have the potential for growth in thrust and power levels beyond 10 KW.
Rawlin, Vincent K.↗
Integrated control and health monitoring capacitive displacement sensor development task. Orbit transfer rocket engine technology program
The development of a miniature multifunction turbomachinery shaft displacement sensor using state-of-the-art non-contract capacitive sensing technology is described. Axial displacement, radial displacement, and speed are sensed using a single probe within the envelope normally required for a single function. A survey of displacement sensing technology is summarized including inductive, capacitive, optical and ultrasonic techniques. The design and operation of an experimental triple function sensor is described. Test results are included showing calibration tests and simultaneous dynamic testing of multiple functions. Recommendations for design changes are made to improve low temperature performance, reliability, and for design of a flight type signal conditioning unit.
Collamore, Frank N.↗
GOES-I/M ascent maneuvers from transfer orbit to station
The Geostationary Operational Environmental Satellite (GOES)-I/M station acquisition sequence consists nominally of three in-plane/out-of-plane maneuvers at apogee on the line of relative nodes and a small in-plane maneuver at perigee. Existing software to determine maneuver attitude, ignition time, and burn duration required modification to optimize the out-of-plane parts and admit the noninertial, three-axis stabilized attitude. The Modified Multiple Impulse Station Acquisition Maneuver Planning Program (SENARIO2) was developed from its predecessor, SCENARIO, to optimize the out-of-plane components of the impulsive delta-V vectors. Additional new features include commputation of short term J sub 2 perturbations and output of all premaneuver and postmaneuver orbit elements, coarse maneuver attitudes, propellant usage, spacecraft antenna aspect angles, and ground station coverage. The output data are intended to be used in the launch window computation and by the maneuver targeting computation (General Maneuver (GMAN) Program) software. The maneuver targeting computation in GMAN was modified to admit the GOES-I/M maneuver attitude. Appropriate combinations of ignition time, burn duration, and attitude enable any reasonable target orbit to be achieved.
Abeyagunawardene, S.↗
Performance of Large Area Xenon Ion Thrusters for Orbit Transfer Missions
Studies have indicated that xenon ion propulsion systems can enable the use of smaller earth-launch vehicles for satellite placement which results in significant cost savings. These analyses have assumed the availability of advanced, high power ion thrusters-operating at about 10 kW or higher. A program was initiated to explore the viability of operating 50 cm diameter ion thrusters at this power level. Operation with several discharge chamber and ion extraction grid set combinations was demonstrated and data were obtained at power levels to 16 kW. Fifty cm diameter thrusters using state of the art 30 cm diameter grids or advanced technology 50 cm diameter grids allow discharge power and beam current densities commensurate with long life at power levels up to 10 kW. In addition, 50 cm diameter thrusters are shown to have potential for growth in thrust and power levels beyond 10 kW.
Rawlin, Vincent K.↗
Xenon ion propulsion for orbit transfer
The status of critical ion propulsion system elements is reviewed. Electron bombardment ion thrusters for primary propulsion have evolved to operate on xenon in the 5-10 kW power range. Thruster efficiencies of 0.7 and specific impulse values of 4000 s have been documented. The baseline thruster currently under development by NASA LeRC includes ring-cusp magnetic field plasma containment and dished two-grid ion optics. Based on past experience and demonstrated simplifications, power processors for these thrusters should have approximately 500 parts, a mass of 40 kg, and an efficiency near 0.94. Thrust vector control, via individual thruster gimbals, is a mature technology. High pressure, gaseous xenon propellant storage and control schemes, using flight qualified hardware, result in propellant tankage fractions between 0.1 and 0.2. In-space and ground integration testing has demonstrated that ion propulsion systems can be successfully integrated with their host spacecraft.
Rawlin, V. K.↗
Comparison of solution approaches for minimum-fuel, low-thrust, power-limited orbital transfers
An initial assessment of the feasibility of a function space gradient method for computing solutions to minimum-fuel power-limited transfers encompassing a wide range of thrust to weight ratios is conducted. Three transfers between coplanar ellipses are used as test cases. The gradient method performs best at the high end of the thrust to weight ratio range. At the lower end, there is reduced sensitivity of the fuel consumption to the control profiles. The minimum fuel consumption and the trajectory are computed quite accurately but the control profiles are in error. An approximate analytical solution, obtained by Edelbaum using the method of averaging, is discussed.
Mease, Kenneth D.↗
Impact of atmospheric density scale height on the performance of aeroassisted coplanar orbital transfer vehicles
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Naidu, D. S.↗
Orbital transfer vehicle oxygen turbopump technology. Volume 2: Nitrogen and ambient oxygen testing
The testing of a rocket engine oxygen turbopump using high pressure ambient temperature nitrogen and oxygen as the turbine drive gas in separate test series is discussed. The pumped fluid was liquid nitrogen or liquid oxygen. The turbopump (TPA) is designed to operate with 400 F oxygen turbine drive gas which will be demonstrated in a subsequent test series. Following bearing tests, the TPA was finish machined (impeller blading and inlet/outlet ports). Testing started on 15 February 1989 and was successfully concluded on 21 March 1989. Testing started using nitrogen to reduce the ignition hazard during initial TPA checkout. The Hydrostatic Bearing System requires a Bearing Pressurization System. Initial testing used a separate bearing supply to prevent a rubbing start. Two test series were successfully completed with the bearing assist supplied only by the pump second stage output which entailed a rubbing start until pump pressure builds up. The final test series used ambient oxygen drive and no external bearing assist. Total operating time was 2268 seconds. There were 14 starts without bearing assist and operating speeds up to 80,000 rpm were logged. Teardown examination showed some smearing of silverplated bearing surfaces but no exposure of the underlying monel material. There was no evidence of melting or oxidation due to the oxygen exposure. The articulating, self-centering hydrostatic bearing exhibited no bearing load or stability problems. The only anomaly was higher than predicted flow losses which were attributed to a faulty ring seal. The TPA will be refurbished prior to the 400 F oxygen test series but its condition is acceptable, as is, for continued operating. This was a highly successful test program.
Brannam, R. J.↗