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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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An advanced lightweight insulation system suitable for use in long duration, low pressure planetary orbiter-type motor applications was developed. Experiments included the screening of various filler and binder materials with optimization studies combining the best of each. Small scale test motor data were used to judge the degree of success.
A program to apply holography for nondestructive testing of the X-248 rocket motor was undertaken. The objective was to establish the capability of holography in detecting known unbonding between liner and propellant. Holography was performed employing stressing techniques: (1) acoustical, (2) thermal, (3) radiative, and (4) static loading. Radiative stressing was successful in locating a large area of liner/propellant unbond. The results were correlated with destructive testing. Theoretical analysis provided an understanding of motor case holography in conjunction with radiative stressing.
The development of magnetic bearings for use in direct current electric motors is discussed. The characteristics of the magnets used in the construction of the bearings are described. A magnetic bearing using steel armoring on permanent magnets was selected for performance tests. The specifications of the motor are presented. The test equipment used in the evaluation is described.
Four iron-less armature torque motors, four Hall device position sensor assemblies, and two test fixtures were fabricated. The design approach utilized samarium cobalt permanent magnets, a large airgap, and a three-phase winding in a stationary ironless armature. Hall devices were employed to sense rotor position. An ironless armature torque motor having an outer diameter of 4.25 inches was developed to produce a torque constant of 65 ounce-inches per ampere with a resistance of 20.5 ohms. The total weight, including structural elements, was 1.58 pounds. Test results indicated that all specifications were met except for generated voltage waveform. It is recommended that investigations be made concerning the generated voltage waveform to determine if it may be improved.
Fluidic sounding rocket motor ignition has been found to be feasible using a system without stored energy and with the complete absence of electrical energy and wiring. The fluidic ignitor is based on a two component aerodynamic resonance heating device called the pneumatic match. Temperatures in excess of 600 C were generated in closed resonance tubes which were excited by a free air jet from a simple convergent nozzle. Using nitrocellulose interface material, ignition of boron potassium nitrate was accomplished with air supply pressures as low as 45 psi. This paper describes an analytical and experimental program which established a fluidic rocket motor ignition system concept incorporating a pneumatic match with a simple hand pump as the only energy source.
The evolution of the space shuttle booster system is reviewed from its initial concepts based on liquid-propellant reusable boosters to the final selection of recoverable, solid-fuel rocket motors. The rationale associated with each of the several major decisions in the evolution process is discussed. It is shown that the external tank orbiter configuration emerging from the latest studies takes maximum advantage of the solid rocket motor development experience and promises to be the optimum configuration for fulfilling the paramount shuttle program requirements of minimum total development risk within acceptable costs.
Motor is proposed to convert radiation from sun directly into mechanical energy. Motor utilizes thermal expansion of liquid, heated by sun, as driving force. Unlike most thermally powered systems, it does not require that liquid be converted into vapor.
The fabrication procedures are described for filament-wound rocket motor cases, approximately 26.25 inches long by 25.50 inches diameter, utilizing graphite fibers. The process utilized prepreg tape which consists of Fortafil 4-R fibers in the E-759 epoxy resin matrix. This fabrication effect demonstrated an ability to fabricate high quality graphite/epoxy rocket motor cases in the 26.25 inch by 25.50 inch size range.
A review of typical surveillance and monitoring practices followed during the flight phases of representative solid-propellant upper stages and apogee motors was conducted to evaluate the need for improved flight diagnostic instrumentation on future spacecraft. The capabilities of the flight instrumentation package were limited to the detection of whether or not the solid motor was the cause of failure and to the identification of probable primary failure modes. Conceptual designs of self-contained flight instrumentation packages capable of meeting these reqirements were generated and their performance, typical cost, and unit characteristics determined. Comparisons of a continuous real time and a thresholded hybrid design were made on the basis of performance, mass, power, cost, and expected life. The results of this analysis substantiated the feasibility of a self-contained independent flight instrumentation module as well as the existence of performance margins by which to exploit growth option applications.
A miniature solid propellant rocket motor has been developed to be used in a program to determine those parameters which must be duplicated in a cold gas flow to produce aerodynamic effects on an experimental model similar to those produced by hot, particle-laden exhaust plumes. Phenomena encountered during the testing of the miniature solid propellant motors included erosive propellant burning caused by high flow velocities parallel to the propellant surface, regressive propellant burning as a result of exposed propellant edges, the deposition of aluminum oxide on the nozzle surfaces sufficient to cause aerodynamic nozzle throat geometry changes, and thermal erosion of the nozzle throat at high chamber pressures. A series of tests was conducted to establish the stability of the rocket chamber pressure and the repeatibility of test conditions. Data are presented which define the tests selected to represent the final test matrix. Qualitative observations are also presented concerning the phenomena experienced based on the results of a large number or rocket tests not directly applicable to the final test matrix.
Brushless dc motors were intensively developed and tested over several years before qualification as the prime movers for Apollo Spacecraft life support blowers, and for circulating oxygen in the lunar portable life support system. Knowledge gained through prototype development and critical testing has significantly influenced the technology employed, broadened markets and applications, and reduced the cost of present day motors.
The material presented on the motor activity under weightless conditions (brief and long) leads to the conclusion that it is not significantly disrupted, if those being examined are secured at the workplaces. Some discoordination of movement, moderately expressed disruption of the precision of reproduction of assigned muscular forces, etc., were observed. Motor disorders decrease significantly in proportion to the length of stay under weightless conditions. This apparently takes place, as a consequence of formation of a new functional system, adequate to the conditions of weightlessness. Tests on intact and labyrinthectomized animals have demonstrated that signaling from the inner ear receptors is superfluous in weightlessness, since it promotes the onset of disruptions in the combined work of the position analyzers.
Experiments are described in which it was attempted to define the state of HCl when rocket motors using ammonium perchlorate and powdered aluminum filler in the fuel system are fired under various relative humidity conditions. A small, variable load rocket motor was developed so that controlled quantities of exhaust products could be produced in an environmental chamber. A technique is described for distinguishing the difference between HCl existing as a gas or existing as an acid aerosol.
The Solid Rocket Motor (SRM) for the Space Shuttle Solid Rocket Booster (SRB) is being developed by Thiokol Corporation, Wasatch Division under the cognizance of Marshall Space Flight Center (MSFC), Contract NAS8-30490. Solid Rocket Motor requirements are summarized, and the overall SRM configuration and performance characteristics are presented in this paper. Design details of the major SRM components are reviewed and the approach taken in each component area is identified and related to previous industry experience. The overall scope of the SRM Project is presented. Activities on the project since authorization to proceed in June 1974 are reviewed. A summary schedule is presented which relates key SRM Project activities to the overall Shuttle Program.
The tailoff thrust and impulse imbalance between pairs of solid rocket motors is of particular interest for the Space Shuttle Vehicle because of the potential control problems that exist with this asymmetric configuration. Although a similar arrangement of solid rocket motors was utilized for the Titan Program, they produced less than one-half the thrust level of the Space Shuttle at web action time, and the overall vehicle was symmetric. Since the Titan Program does provide the most applicable actual test data, 23 flight pairs were analyzed to determine the actual tailoff thrust and impulse imbalance experienced. The results were scaled up using the predicted web action time thrust and tailoff time to arrive at values for the Space Shuttle. These values were then statistically treated to obtain a prediction of the maximum imbalance one could expect to experience during the Shuttle Program.
A FORTRAN program for induction motor analysis is described. The analysis includes calculations of torque-speed characteristics, efficiency, losses, magnetic flux densities, weights, and various electrical parameters. The program is limited to three-phase Y-connected, squirrel-cage motors. Detailed instructions for using the program are given. The analysis equations are documented, and the sources of the equations are referenced. The appendixes include a FORTRAN symbol list, a complete explanation of input requirements, and a list of error messages.
A three dimensional finite element solution of the acoustic vibration problem in a solid propellant rocket motor is presented. The solution yields the natural circular frequencies of vibration and the corresponding acoustic pressure mode shapes, considering the coupled response of the propellant grain to the acoustic oscillations occurring in the motor cavity. The near incompressibility of the solid propellant is taken into account in the formulation. A relatively simple example problem is solved in order to illustrate the applicability of the analysis and the developed computer code.
As part of a continuing study of the environmental effects of solid rocket motor (SRM) operations in the troposphere, a numerical model was used to simulate the afterburning processes occurring in solid rocket motor plumes and to predict the quantities of potentially harmful chemical species which are created. The calculations include the effects of finite-rate chemistry and turbulent mixing. It is found that the amount of NO produced is much less than the amount of HCl present in the plume, that chlorine will appear predominantly in the form of HCl although some molecular chlorine is present, and that combustion is complete as is evident from the predominance of carbon dioxide over carbon monoxide.