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At least 181 records · Page 10

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NASA NTRS1990

New method of making advanced tube-bundle rocket thrust chambers

An improved method of fabrication rocket chambers for future space applications is described. Included are fabrication demonstrator and test chambers produced by this method. This concept offers the promise of improved cyclic life, reusability, and performance. The performance is improved because of the enhanced enthalpy extraction. The improved cyclic life, reusability, and reliability is improved because of the structural compliance inherent in the construction. The method of construction involves the forming of the combustion chamber by a tube-bundle of high conductivity copper or copper alloy tubes and the bonding of these tubes by a unique electroforming operation. Furthermore, the method of fabrication reduces chamber complexity by incorporating manifolds, and structural stiffeners while having the potential for thrust chamber cost and weight reduction.

Kazaroff, John M.↗

NASA NTRS1990

Test plan pressure fed thrust chamber technology

Aerojet is developing the technology for the design of a reliable, low cost, efficient, and lightweight LOX/RP-1 pressure fed engine. This technology program is a direct result of Aerojet's liquid rocket booster (LRB) study and previous NASA studies that identified liquid engines using high bulk density hydrocarbon fuels as very attractive for a space transportation system (STS). Previous large thrust LOX/RP-1 engine development programs were characterized by costly development problems due to combustion instability damage. The combustion stability solution was typically obtained through trial and error methods of minimizing instability damage by degrading engine performance. The approach to this program was to utilize existing and newly developed combustion analysis models and design methodology to create a thrust chamber design with features having the potential of producing reliable and efficient operation. This process resulted in an engine design with a unique high thrust-per-element OFO triplet injector utilizing a low cost modular approach. Cost efficient ablative materials are baselined for the injector face and chamber. Technology demonstration will be accomplished through a hot fire test program using appropriately sized subscale hardware. This subscale testing will provide a data base to supplement the current industry data bank and to anchor and validate the applied analysis models and design methodology. Once anchored and validated, these analysis models and design methodology can be applied with greatly increased confidence to design and characterize a large scale pressure fed LOX/RP-1 thrust chamber. The objective of this test program is to generate a data base that can be used to anchor and validate existing analysis models and design methodologies and to provide early concept demonstration of a low cost, efficient LOX/RP-1 thrust chamber. Test conditions and hardware instrumentation were defined to provide data sufficient to characterize combustion stability, performance, and thermal operation over a wide thrust chamber throttling range.

Dunn, Glenn↗

NASA NTRS1990

Cyclic hot firing results of tungsten-wire-reinforced, copper-lined thrust chambers

An advanced thrust liner material for potential long life reusable rocket engines is described. This liner material was produced with the intent of improving the reusable life of high pressure thrust chambers by strengthening the chamber in the hoop direction, thus avoiding the longitudinal cracking due to low cycle fatigue that is observed in conventional homogeneous copper chambers, but yet not reducing the high thermal conductivity that is essential when operating with high heat fluxes. The liner material produced was a tungsten wire reinforced copper composite. Incorporating this composite into two hydrogen-oxygen test rocket chambers was done so that its performance as a reusable liner material could be evaluated. Testing results showed that both chambers failed prematurely, but the crack sites were perpendicular to the normal direction of cracking indicating a degree of success in containing the tremendous thermal strain associated with high temperature rocket engines. The failures, in all cases, were associated with drilled instrumentation ports and no other damages or deformations were found elsewhere in the composite liners.

Kazaroff, John M.↗

NASA NTRS1992

A high-efficiency acoustic chamber and the anomalous sample rotation

A high efficiency acoustic chamber for the levitation and manipulation of liquid or molten samples in a microgravity environment has been developed. The chamber uses two acoustic drivers that are mounted at opposite corners of the chamber; by driving these at the same frequency, with 18-deg phase shifts, an increase in force of a factor of 3-4 is obtainable relative to the force of a single-driver system that is operated at the same power level. This enhancement is due to the increased coupling between the sound driver and the chamber. An anomalous rotation is noted to be associated with the chamber; this is found to be eliminated by a physically as-yet inexplicable empirical solution.

Wang, Taylor G.↗

NASA NTRS1931

Performance of a Compression-ignition Engine with a Precombustion Chamber Having High-Velocity Air Flow

Presented here are the results of performance tests made with a single-cylinder, four stroke cycle, compression-ignition engine. These tests were made on a precombustion chamber type of cylinder head designed to have air velocity and tangential air flow in both the chamber and cylinder. The performance was investigated for variable load and engine speed, type of fuel spray, valve opening pressure, injection period and, for the spherical chamber, position of the injection spray relative to the air flow. The pressure variations between the pear-shaped precombustion chamber and the cylinder for motoring and full load conditions were determined with a Farnboro electric indicator. The combustion chamber designs tested gave good mixing of a single compact fuel spray with the air, but did not control the ensuing combustion sufficiently. Relative to each other, the velocity of air flow was too high, the spray dispersion by injection too great, and the metering effect of the cylinder head passage insufficient. The correct relation of these factors is of the utmost importance for engine performance.

Spanogle, J A↗

NASA NTRS1938

Investigation of Ignition and Combustion Processes of Diesel Engines Operating with Turbulence and Air-storage Chambers

The flame photographs obtained with combustion-chamber models of engines operating respectively, with turbulence chamber and air-storage chambers or cells, provide an insight into the air and fuel movements that take place before and during combustion in the combustion chamber. The relation between air velocity, start of injection, and time of combustion was determined for the combustion process employing a turbulence chamber.

Petersen, Hans↗

NASA NTRS1994

Annular-Cross-Section CFE Chamber

Proposed continuous-flow-electrophoresis (CFE) chamber of annular cross section offers advantages over conventional CFE chamber, and wedge-cross-section chamber described in "Increasing Sensitivity in Continuous-Flow Electrophoresis" (MFS-26176). In comparison with wedge-shaped chamber, chamber of annular cross section virtually eliminates such wall effects as electro-osmosis and transverse gradients of velocity. Sensitivity enhanced by incorporating gradient maker and radial (collateral) flow.

Sharnez, Rizwan↗

NASA NTRS1993

Hot-gas-side heat transfer characteristics of subscale, plug-nozzle rocket calorimeter chamber

An experimental investigation was conducted to determine the hot-gas-side heat transfer characteristics for a liquid-hydrogen-cooled, subscale, plug-nozzle rocket test apparatus. This apparatus has been used since 1975 to evaluate rocket engine advanced cooling concepts and fabrication techniques, to screen candidate combustion chamber liner materials, and to provide data for model development. In order to obtain the data, a water-cooled calorimeter chamber having the same geometric configuration as the plug-nozzle test apparatus was tested. It also used the same two showerhead injector types that were used on the test apparatus: one having a Rigimesh faceplate and the other having a platelet faceplate. The tests were conducted using liquid oxygen and gaseous hydrogen as the propellants over a mixture ratio range of 5.8 to 6.3 at a nominal chamber pressure of 4.14 MPa abs (600 psia). The two injectors showed similar performance characteristics with the Rigimesh faceplate having a slightly higher average characteristic-exhaust-velocity efficiency of 96 percent versus 94.4 percent for the platelet faceplate. The throat heat flux was 54 MW/m(sup 2) (33 Btu/in.(sup 2)-sec) at the nominal operating condition, which was a chamber pressure of 4.14 MPa abs (600 psia), a hot-gas-side wall temperature of 730 K (1314 R), and a mixture ratio of 6.0. The chamber throat region correlation coefficient C(sub g) for a Nusselt number correlation of the form Nu =C(sub g)Re(sup 0.8)Pr(sup 0.3) averaged 0.023 for the Rigimesh faceplate and 0.026 for the platelet faceplate.

Quentmeyer, Richard J.↗

NASA NTRS1994

Prediction of engine performance and wall erosion due to film cooling for the 'fast track' ablative thrust chamber

Efforts have been made at the Propulsion Laboratory (MSFC) to design and develop new liquid rocket engines for small-class launch vehicles. Emphasis of the efforts is to reduce the engine development time with the use of conventional designs while meeting engine reliability criteria. Consequently, the engine cost should be reduced. A demonstrative ablative thrust chamber, called 'fast-track', has been built. To support the design of the 'fast-track' thrust chamber, predictions of the wall temperature and ablation erosion rate of the 'fast-track' thrust chamber have been performed using the computational fluid dynamics program REFLEQS (Reactive Flow Equation Solver). The analysis is intended to assess the amount of fuel to be used for film cooling so that the erosion rate of the chamber ablation does not exceed its allowable limit. In addition, the thrust chamber performance loss due to an increase of the film cooling is examined.

Trinh, Huu P.↗

NASA NTRS1995

Advanced Small Rocket Chambers. Option 3: 110 1Bf Ir-Re Rocket, Volume 1

This report describes the AJ10-221, a high performance Iridium-coated Rhenium (Ir-Re) 110 lbf (490N) welded rocket chamber with 286:1 area ratio nozzle. This engine was designed, built, and hot fired for over 6 hours on this program. It demonstrated an I(s) of 321.8 sec, which is 10 sec higher than conventional 110 lbf silicide coated Cb chambers now in use. The approach used in this portion of the program was to demonstrate the performance improvement that can be made by the elimination of fuel film cooling and the use of a high temperature (4000F) (2200C) iridium-coated rhenium (Ir-Re) rocket chamber. Detailed thermal, performance, mechanical, and dynamic design analyses of the full engine were conducted by Aerojet. Two Ir-Re chambers were built to the Aerojet design by Ultramet, using the chemical vapor deposition (CVD) process. Incorporation of a secondary mixing device or Boundary Layer Trip (BLT) within the combustion chamber (Aerojet Patents 4882904 and 4936091) results in improvement in flow uniformity, and a significant life and performance increase. The 110 lbf engine design was verified in bolt-up hardware tests at sea level and altitude. The effects of injector design on performance were studied. Two duplicate injectors were fabricated matching the preferred design and were demonstrated to be interchangeable in operation. One of these units was fabricated matching the preferred design and was demonstrated to be interchangeable in operation. One of these units was welded into a flight type thruster which was tested for an accumulated duration of 22,590 sec in 93 firings, one of which was a continuous burn of two hours. A design deficiency in the C-103 nozzle near the Re-Cb transition joint was discovered, studied and corrected design has been prepared. Workhardening studies have been conducted to investigate methods for increasing the low yield strength of the Re in the annealed conditions. An advanced 490N high performance engine has been demonstrated which, when proven to be capable of withstanding launch vibration, is ready for flight qualification.

Jassowski, Donald M.↗

NASA NTRS1995

Advanced small rocket chambers. Option 3: 110 1bf Ir-Re rocket, volume 2

This is the second part of a two-part report that describes the AJ10-221, a high performance iridium-coated rhenium (Ir-Re) 110 lbf (490N) welded rocket chamber with 286:1 area ratio nozzle. This engine was designed, built, and hot fired for over 6 hours on this program. It demonstrated an I(s) of 321.8 sec, which is 10 sec higher than conventional 110 lbf silicide coated Cb chambers now in use. The approach used in this portion of the program was to demonstrate the performance improvement that can be made by the elimination of fuel film cooling and the use of a high temperature (4000 F) (2200 C) iridium-coated rhenium (Ir-Re) rocket chamber. Detailed thermal, performance, mechanical, and dynamic design analyses of the full engine were conducted by Aerojet. Two Ir-Re chambers were built to the Aerojet design by Ultramet, using the chemical vapor deposition (CVD) process. Incorporation of a secondary mixing device or Boundary Layer Trip (BLT) within the combustion chamber (Aerojet Patents 4882904 and 4936091) results in improvement in flow uniformity, and a significant life and performance increase. The 110 lbf engine design was verified in bolt-up hardware tests at sea level and altitude. The effects of injector design on performance were studied. Two duplicate injectors were fabricated matching the preferred design and were demonstrated to be interchangeable in operation. One of these units were welded into a flight type thruster which was tested for an accumulated duration of 22,590 sec in 93 firings, one of which was a continuous burn of two hours. A design deficiency in the C-103 nozzle near the Re-Cb transition joint was discovered, studied and corrected design has been prepared. Workhardening studies have been conducted to investigate methods for increasing the low yield strength of the Re in the annealed conditions. An advanced 490N high performance engine has been demonstrated which, when proven to be capable of withstanding launch vibration, is ready for flight qualification.

Jassowski, Donald M.↗

NASA NTRS1996

Emulsion Chamber Technology Experiment (ECT)

The experimental objective of Emulsion Chamber Technology (ECT) was to develop space-borne emulsion chamber technology so that cosmic rays and nuclear interactions may subsequently be studied at extremely high energies with long exposures in space. A small emulsion chamber was built and flown on flight STS-62 of the Columbia in March 1994. Analysis of the several hundred layers of radiation-sensitive material has shown excellent post-flight condition and suitability for cosmic ray physics analysis at much longer exposures. Temperature control of the stack was 20 +/-1 C throughout the active control period and no significant deviations of temperature or pressure in the chamber were observed over the entire mission operations period. The unfortunate flight attitude of the orbiter (almost 90% Earth viewing) prevented any significant number of heavy particles (Z greater than or equal to 10) reaching the stack and the inverted flow of shower particles in the calorimeter has not allowed evaluation of absolute primary cosmic ray-detection efficiency nor of the practical time limits of useful exposure of these calorimeters in space to the level of detail originally planned. Nevertheless, analysis of the observed backgrounds and quality of the processed photographic and plastic materials after the flight show that productive exposures of emulsion chambers are feasible in low orbit for periods of up to one year or longer. The engineering approaches taken in the ECT program were proven effective and no major environmental obstacles to prolonged flight are evident.

Gregory, John C.↗

NASA NTRS1961

Combined Effect of Contraction Ratio and Chamber Pressure on the Performance of a Gaseous Hydrogen-Liquid-Oxygen Combustor for a Given Propellant Weight Flow and Oxidant-Fuel Ratio

The effect of contraction ratio and chamber pressure on the combustion performance of a gaseous-hydrogen-liquid-oxygen combustor was investigated analytically and experimentally. The experiment was conducted with a "two-dimensional" gaseous-hydrogen-liquid-oxygen engine of about 150-pound thrust. The contraction ratio was varied from 1.5 to 6 by changing the nozzle throat area. This variation resulted in a chamber pressure variation of about 25 to 120 pounds per square inch. The experimental results were corrected for heat transfer to the engine walls and momentum pressure losses. The experimental performance, as evaluated in terms of characteristic exhaust velocity, was 98 percent of theoretical at contraction ratios greater than 3 but decreased very rapidly at smaller contraction ratios. The heat-transfer rate increased with increasing contraction ratio and chamber pressure; it was about 1 Btu per square inch per second at a contraction ratio of 1.5 and increased to about 3 at a contraction ratio of 6. The combined effects of contraction-ratio and chamber-pressure changes on performance were investigated analytically with a mixing model and a vaporization model. The mixing model predicted very poor mixing at contraction ratios below 3 and almost perfect mixing at higher contraction ratios. The performance predicted by the vaporization model was very close to 100 percent for all contraction ratios. From these results, it was concluded that the performance was limited by poor mixing at low contraction ratios and chamber pressures.

Hersch, Martin↗

NASA NTRS1961

Theoretical Performance of Hydrogen-Oxygen Rocket Thrust Chambers

Data are presented for liquid-hydrogen-liquid-oxygen thrust chambers at chamber pressures from 15 to 1200 pounds per square inch absolute, area ratios to approximately 300, and percent fuel from about 8 to 34 for both equilibrium and frozen composition during expansion. Specific impulse in vacuum, specific impulse, combustion-chamber temperature, nozzle-exit temperature, characteristic velocity, and the ratio of chamber-to-nozzle-exit pressure are included. The data are presented in convenient graphical forms to allow quick calculation of theoretical nozzle performance with over- or underexpansion, flow separation, and introduction of the propellants at various initial conditions or heat loss from the combustion chamber.

Sievers, Gilbert K.↗

NASA NTRS1998

Fastrac Rocket Engine Combustion Chamber Acoustic Cavities

A three dimensional modal analysis was performed using finite fluid elements. The analysis shows four distinct modes of the Fastrac chamber plus cavities near the frequency of the chamber first tangential mode. The mode shapes illustrate the complexity of fluid oscillations in a three dimensional chamber and acoustic cavity. In addition, a first tangential forcing function was applied to the chamber with three different acoustic cavity fluid temperatures. It was observed that the acoustic cavity fluid temperature has a significant effect on the response of the chamber to first tangential mode oscillations.

Christensen, Eric↗

NASA NTRS1998

Experiment 8: Environmental Conditions in the ASTROCULTURE(trademark) Plant Chamber During the USML-2 Mission

Conducting plant research to assess the impact of microgravity on plant growth and development requires a plant chamber that has the capability to control other environmental parameters involved in plant growth and development. The environmental control in a space-based plant chamber must be equivalent to that available in such facilities used for terrestrial plant research. Additionally, plants are very sensitive to a number of atmospheric gaseous materials. Thus, the atmosphere of a plant chamber must be isolated from the space vehicle atmosphere, and the plant growth unit should have the capability to remove any such deleterious materials that may impact plant growth and development. The Wisconsin Center for Space Automation and Robotics (WCSAR), University of Wisconsin-Madison, has developed a totally enclosed controlled environment plant growth unit. The flight unit was used to support the ASTROCULTURE(TM) experiment conducted during the USML-2 mission. The experiment had two major objectives: 1) Provide further validation of the flight unit to control the experiment-defined environmental parameters in the plant chamber, and 2) support a plant experiment to assess the capability of potato plant material to produce tubers in microgravity. This paper describes the temperature, humidity, and carbon dioxide conditions of the plant chamber during the mission, from launch to landing. Another paper will present the plant response data.

Bula, R. J.↗

NASA NTRS1999

Robust Low Cost Aerospike/RLV Combustion Chamber by Advanced Vacuum Plasma Process

Next-generation, regeneratively cooled rocket engines will require materials that can withstand high temperatures while retaining high thermal conductivity. At the same time, fabrication techniques must be cost efficient so that engine components can be manufactured within the constraints of a shrinking NASA budget. In recent years, combustion chambers of equivalent size to the Aerospike chamber have been fabricated at NASA-Marshall Space Flight Center (MSFC) using innovative, relatively low-cost, vacuum-plasma-spray (VPS) techniques. Typically, such combustion chambers are made of the copper alloy NARloy-Z. However, current research and development conducted by NASA-Lewis Research Center (LeRC) has identified a Cu-8Cr-4Nb alloy which possesses excellent high-temperature strength, creep resistance, and low cycle fatigue behavior combined with exceptional thermal stability. In fact, researchers at NASA-LeRC have demonstrated that powder metallurgy (P/M) Cu-8Cr-4Nb exhibits better mechanical properties at 1,200 F than NARloy-Z does at 1,000 F. The objective of this program was to develop and demonstrate the technology to fabricate high-performance, robust, inexpensive combustion chambers for advanced propulsion systems (such as Lockheed-Martin's VentureStar and NASA's Reusable Launch Vehicle, RLV) using the low-cost, VPS process to deposit Cu-8Cr-4Nb with mechanical properties that match or exceed those of P/M Cu-8Cr-4Nb. In addition, oxidation resistant and thermal barrier coatings can be incorporated as an integral part of the hot wall of the liner during the VPS process. Tensile properties of Cu-8Cr-4Nb material produced by VPS are reviewed and compared to material produced previously by extrusion. VPS formed combustion chamber liners have also been prepared and will be reported on following scheduled hot firing tests at NASA-Lewis.

Holmes, Richard↗

NASA NTRS2000

Development of a Methodology for Conducting Hall Thruster EMI Tests in Metal Vacuum Chambers of Arbitrary Shape and Size

While the closed-drift Hall thruster (CDT) offers significant improvement in performance over conventional chemical rockets and other advanced propulsion systems such as the arcjet, its potential impact on spacecraft communication signals must be carefully assessed before widespread use of this device can take place. To this end, many of the potentially unique issues that are associated with these thrusters center on its plume plasma characteristics and the its interaction with electromagnetic waves. Although a great deal of experiments have been made in characterizing the electromagnetic interference (EMI) potential of these thrusters, the interpretation of the resulting data is difficult because most of these measurements have been made in vacuum chambers with metal walls which reflect radio waves emanating from the thruster. This project developed a means of assessing the impact of metal vacuum chambers of arbitrary size or shape on EMI experiments, thereby allowing for test results to be interpreted properly. Chamber calibration techniques were developed and initially tested at RIAME using their vacuum chamber. Calibration experiments were to have been made at Tank 5 of NASA GRC and the 6 m by 9 m vacuum chamber at the University of Michigan to test the new procedure, however the subcontract to RIAME was cancelled by NASA memorandum on Feb. 26. 1999.

Gallimore, Alec D.↗