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  • 1
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    Advanced tube-bundle rocket thrust chambers (1992)
    In:  Other Sources
    Details
    Publication Date: 2011-08-24
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: Journal of Propulsion and Power (ISSN 0748-4658); 8; 4, Ju; 786-791
    Format: text
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    NASA TECHNICAL REPORTS
  • 2
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    Experimental performance of a high-area-ratio rocket nozzle at high combustion chamber pressure (1996)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An experimental investigation was conducted to determine the thrust coefficient of a high-area-ratio rocket nozzle at combustion chamber pressures of 12.4 to 16.5 MPa (1800 to 2400 psia). A nozzle with a modified Rao contour and an expansion area ratio of 1025:1 was tested with hydrogen and oxygen at altitude conditions. The same nozzle, truncated to an area ratio of 440:1, was also tested. Values of thrust coefficient are presented along with characteristic exhaust velocity efficiencies, nozzle wall temperatures, and overall thruster specific impulse.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA-TP-3576 , NAS 1.60:3576 , E-9849
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 3
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    Advanced tube-bundle rocket thrust chamber (1990)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: An advanced rocket thrust chamber for future space application is described along with an improved method of fabrication. Potential benefits of the concept are improved cyclic life, reusability, and performance. Performance improvements are anticipated because of the enhanced heat transfer into the coolant which will enable higher chamber pressure in expander cycle engines. Cyclic life, reusability and reliability improvements are anticipated because of the enhanced structural compliance inherent in the construction. The method of construction involves the forming of the combustion chamber with a tube-bundle of high conductivity copper or copper alloy tubes, and the bonding of these tubes by an electroforming operation. Further, the method of fabrication reduces chamber complexity by incorporating manifolds, jackets, and structural stiffeners while having the potential for thrust chamber cost and weight reduction.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: AIAA PAPER 90-2726
    Format: text
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    NASA TECHNICAL REPORTS
  • 4
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    Hot fire test results of subscale tubular combustion chambers (1992)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Advanced, subscale, tubular combustion chambers were built and test fired with hydrogen-oxygen propellants to assess the increase in fatigue life that can be obtained with this type of construction. Two chambers were tested: one ran for 637 cycles without failing, compared to a predicted life of 200 cycles for a comparable smooth-wall milled-channel liner configuration. The other chamber failed at 256 cycles, compared to a predicted life of 118 cycles for a comparable smooth-wall milled-channel liner configuration. Posttest metallographic analysis determined that the strain-relieving design (structural compliance) of the tubular configuration was the cause of this increase in life.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-TP-3222 , E-6849 , NAS 1.60:3222
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 5
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    Experimental thrust performance of a high-area-ratio rocket nozzle (1987)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An experimental investigation was conducted to determine the thrust performance attainable from high-area-ratio rocket nozzles. A modified Rao-contoured nozzle with an expansion area of 1030 was test fired with hydrogen-oxygen propellants at altitude conditions. The nozzle was also tested as a truncated nozzle, at an expansion area ratio of 428. Thrust coefficient and thrust coefficient efficiency values are presented for each configuration at various propellant mixture ratios (oxygen/fuel). Several procedural techniques were developed permitting improved measurement of nozzle performance. The more significant of these were correcting the thrust for the aneroid effects, determining the effective chamber pressure, and referencing differential pressure transducers to a vacuum reference tank.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-TP-2720 , E-3236-1 , NAS 1.60:2720
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 6
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    Experimental evaluation of heat transfer on a 1030:1 area ratio rocket nozzle (1987)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A 1030:1 carbon steel, heat-sink nozzle was tested. The test conditions included a nominal chamber pressure of 2413 kN/sq m and a mixture ratio range of 2.78 to 5.49. The propellants were gaseous oxygen and gaseous hydrogen. Outer wall temperature measurements were used to calculate the inner wall temperature and the heat flux and heat rate to the nozzle at specified axial locations. The experimental heat fluxes were compared to those predicted by the Two-Dimensional Kinetics (TDK) computer model analysis program. When laminar boundary layer flow was assumed in the analysis, the predicted values were within 15% of the experimental values for the area ratios of 20 to 975. However, when turbulent boundary layer conditions were assumed, the predicted values were approximately 120% higher than the experimental values. A study was performed to determine if the conditions within the nozzle could sustain a laminar boundary layer. Using the flow properties predicted by TDK, the momentum-thickness Reynolds number was calculated, and the point of transition to turbulent flow was predicted. The predicted transition point was within 0.5 inches of the nozzle throat. Calculations of the acceleration parameter were then made to determine if the flow conditions could produce relaminarization of the boundary layer. It was determined that if the boundary layer flow was inclined to transition to turbulent, the acceleration conditions within the nozzle would tend to suppress turbulence and keep the flow laminar-like.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-TP-2726 , E-3558 , NAS 1.60:2726 , AIAA PAPER 87-2070
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 7
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    Comparison of theoretical and experimental thrust performance of a 1030:1 area ratio rocket nozzle at a chamber pressure of 2413 kN/m2 (350 psia) (1987)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The joint Army. Navy, NASA. Air Force (JANNAF) rocket engine peformnace prediction procedure is based on the use of various reference computer programs. One of the reference programs for nozzle analysis is the Two-Dimensional Kinetics (TDK) Program. The purpose of this report is to calibrate the JANNAF procedure incorporated into the December l984 version of the TDK program for the high-area-ratio rocket engine regime. The calibration was accomplished by modeling the performance of a 1030:1 rocket nozzle tested at NASA Lewis Research Center. A detailed description of the experimental test conditions and TDK input parameters is given. The results show that the computer code predicts delivered vacuum specific impulse to within 0.12 to 1.9 percent of the experimental data. Vacuum thrust coefficient predictions were within + or - 1.3 percent of experimental results. Predictions of wall static pressure were within approximately + or - 5 percent of the measured values. An experimental value for inviscid thrust was obtained for the nozzle extension between area ratios of 427.5 and 1030 by using an integration of the measured wall static pressures. Subtracting the measured thrust gain produced by the nozzle between area ratios of 427.5 and 1030 from the inviscid thrust gain yielded experimental drag decrements of 10.85 and 27.00 N (2.44 and 6.07 lb) for mixture ratios of 3.04 and 4.29, respectively. These values correspond to 0.45 and 1.11 percent of the total vacuum thrust. At a mixture ratio of 4.29, the TDK predicted drag decrement was 16.59 N (3.73 lb), or 0.71 percent of the predicted total vacuum thrust.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-TP-2725 , E-3523 , NAS 1.60:2725 , AIAA PAPER 87-2069
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 8
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    Hot fire fatigue testing results for the compliant combustion chamber (1992)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A hydrogen-oxygen subscale rocket combustion chamber was designed incorporating an advanced design concept to reduce strain and increase life. The design permits unrestrained thermal expansion of a circumferential direction and, thereby, provides structural compliance during the thermal cycling of hot-fire testing. The chamber was built and test fired at a chamber pressure of 4137 kN/sq m (600 psia) and a hydrogen-oxygen mixture ratio of 6.0. Compared with a conventional milled-channel configuration, the new structurally compliant chamber had a 134 or 287 percent increase in fatigue life, depending on the life predicted for the conventional configuration.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-TP-3223 , E-6850 , NAS 1.60:3223
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 9
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    Experimental evaluation of heat transfer on a 1030:1 area ratio rocket nozzle (1987)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: A 1030:1 carbon steel, heat-sink nozzle was tested. The test conditions included a nominal chamber pressure of 2413 kN/sq m and a mixture ratio range of 2.78 to 5.49. The propellants were gaseous oxygen and gaseous hydrogen. Outer wall temperature measurements were used to calculate the inner wall temperature and the heat flux and heat rate to the nozzle at specified axial locations. The experimental heat fluxes were compared to those predicted by the Two-Dimensional Kinetics (TDK) computer model analysis program. When laminar boundary layer flow was assumed in the analysis, the predicted values were within 15 percent of the experimental values for the area ratios of 20 to 975. However, when turbulent boundary layer conditions were assumed, the predicted values were approximately 120 percent higher than the experimental values. A study was performed to determine if the conditions within the nozzle could sustain a laminar boundary layer. Using the flow properties predicted by TDK, the momentum-thickness Reynolds number was calculated, and the point of transition to turbulent flow was predicted. The predicted transition point was within 0.5 inches of the nozzle throat. Calculations of the acceleration parameter were then made to determine if the flow conditions could produce relaminarization of the boundary layer. It was determined that if the boundary layer flow was inclined to transition to turbulent, the acceleration conditions within the nozzle would tend to suppress turbulence and keep the flow laminar-like.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: AIAA PAPER 87-2070
    Format: text
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    NASA TECHNICAL REPORTS
  • 10
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    Comparison of theoretical and experimental thrust performance of a 1030:1 area ratio rocket nozzle at a chamber pressure of 2413 kN/sq m (350 psia) (1987)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: The Joint Army, Navy, NASA, Air Force (JANNAF) rocket-engine performance-prediction procedure is based on the use of various reference computer programs. One of the reference programs for nozzle analysis is the Two-Dimensional Kinetics (TDK) Program. The purpose of this report is to calibrate the JANNAF procedure that has been incorporated into the December 1984 version of the TDK program for the high-area-ratio rocket-engine regime. The calibration was accomplished by modeling the performance of a 1030:1 rocket nozzle tested at NASA Lewis. A detailed description of the test conditions and TDK input parameters is given. The reuslts indicate that the computer code predicts delivered vacuum specific impulse to within 0.12 to 1.9 percent of the experimental data. Vacuum thrust coefficient predictions were within + or - 1.3 percent of experimental results. Predictions of wall static pressure were within approximately + or - 5 percent of the measured values.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: AIAA PAPER 87-2069
    Format: text
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    NASA TECHNICAL REPORTS
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