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  • 1
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    The Superorbital Expansion Tube concept, experiment and analysis (1995)
    In:  CASI
    Details
    Publication Date: 2004-12-03
    Description: In response to the need for ground testing facilities for super orbital re-entry research, a small scale facility has been set up at the University of Queensland to demonstrate the superorbital expansion tube concept. This unique device is a free piston driven, triple diaphragm, impulse shock facility which uses the enthalpy multiplication mechanism of the unsteady expansion process and the addition of a secondary shock driver to further heat the driver gas. The pilot facility has been operated to produce quasi-steady test flows in air with shock velocities in excess of 13 km/s and with a usable test flow duration of the order of 15 micro sec. an experimental condition produced in the facility with total enthalpy of 108 MJ/kg and a total pressure of 335 MPa is reported. A simple analytical flow model which accounts for non-ideal rupture of the light tertiary diaphragm and the resulting entropy increase in the test gas is discussed. It is shown that equilibrium calculations more accurately model the unsteady expansion process than calculations assuming frozen chemistry. This is because the high enthalpy flows produced in the facility can only be achieved if the chemical energy stored in the test flow during shock heating of the test gas is partially returned to the flow during the process of unsteady expansion. Measurements of heat transfer rates to a flat plate demonstrate the usability of test flow for aerothermodynamic testing and comparison of these rates with empirical calculations confirms the usable accuracy of the flow model.
    Keywords: RESEARCH AND SUPPORT FACILITIES (AIR)
    Type: Shock Tunnel Studies of Scramjet Phenomena 1993; p 97-105
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  • 2
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    Particle impact and optical property analysis of the surfaces of Surveyor 3 materials (1972)
    In:  CASI
    Details
    Publication Date: 2005-11-30
    Description: As part of the investigation of Surveyor 3 materials, a study was conducted to determine the effect of the lunar environment on some of the painted and unpainted exterior surfaces. Examination of the camera parts and tube sections was conducted using three techniques: (1) optical and scanning electron microscopy, (2) energy dispersive X-ray probe analysis, and (3) spectral reflectance measurements.
    Keywords: SPACE SCIENCES
    Type: Analysis of Surveyor 3 Mater. and Phot. Returned by Apollo 12; p 76-88
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  • 3
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    Pressure scaling effects in a scramjet combustion chamber (1987)
    In:  Other Sources
    Details
    Publication Date: 2011-08-19
    Description: The test results obtained for a model scramjet over a range of pressure levels corresponding to different flight altitudes involve enthalpies that vary from the ignition limit, at the low temperature end, to temperatures where the dissociation of combustion products severely limits heat release. The minimum temperature is noted to be highly pressure-sensitive; above the ignition limit, the amount of heat release increased markedly with pressure and with combustion chamber length. A FEM computer code has been used to model the mixing and combustion processes.
    Keywords: AIRCRAFT PROPULSION AND POWER
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  • 4
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    Hydrogen scramjet with sidewall injection - Shock tunnel simulations (1987)
    In:  Other Sources
    Details
    Publication Date: 2011-08-19
    Description: A free-piston shock tunnel has been used to obtain test data on a scramjet combustion chamber with sidewall injection. The results obtained indicate that combustion was strongly influenced by a region of fuel whose temperature was held below its ignition temperature by wall-cooling effects; this increased the fraction of unburned fuel and resulted in a significant loss of specific impulse. Aerodynamic heating would keep the walls above hydrogen ignition temperature in an actual scramjet powerplant, however. Maximum specific impulse was obtained with a combination of parallel and transverse injection in a long combustion chamber, followed by a dual stage expansion.
    Keywords: AIRCRAFT PROPULSION AND POWER
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  • 5
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    Preliminary calibration of a generic scramjet combustor (1991)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The results of a preliminary investigation of the combustion of hydrogen fuel at hypersonic flow conditions are provided. The tests were performed in a generic, constant-area combustor model with test gas supplied by a free-piston-driven reflected-shock tunnel. Static pressure measurements along the combustor wall indicated that burning did occur for combustor inlet conditions of P(static) approximately equal to 19kPa, T(static) approximately equal to 1080 K, and U approximately equal to 3630 m/s with a fuel equivalence ratio approximately equal to 0.9. These inlet conditions were obtained by operating the tunnel with stagnation enthalpy approximately equal to 8.1 MJ/kg, stagnation pressure approximately equal to 52 MPa, and a contoured nozzle with a nominal exit Mach number of 5.5.
    Keywords: AIRCRAFT PROPULSION AND POWER
    Type: NASA-CR-187539 , NAS 1.26:187539 , ICASE-16 , AD-A234873
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  • 6
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    Shock tunnel studies of scramjet phenomena, supplement 5 (1990)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A series of reports are presented on SCRAMjet studies, shock tunnel studies, and expansion tube studies. The SCRAMjet studies include: (1) Investigation of a Supersonic Combustion Layer; (2) Wall Injected SCRAMjet Experiments; (3) Supersonic Combustion with Transvers, Circular, Wall Jets; (4) Dissociated Test Gas Effects on SCRAMjet Combustors; (5) Use of Silane as a Fuel Additive for Hypersonic Thrust Production, (6) Pressure-length Correlations in Supersonic Combustion; (7) Hot Hydrogen Injection Technique for Shock Tunnels; (8) Heat Release - Wave Interaction Phenomena in Hypersonic Flows; (9) A Study of the Wave Drag in Hypersonic SCRAMjets; (10) Parametric Study of Thrust Production in the Two Dimensional SCRAMjet; (11) The Design of a Mass Spectrometer for use in Hypersonic Impulse Facilities; and (12) Development of a Skin Friction Gauge for use in an Impulse Facility. The shock tunnel studies include: (1) Hypervelocity flow in Axisymmetric Nozzles; (2) Shock Tunnel Development; and (3) Real Gas Efects in Hypervelocity Flows over an Inclined Cone. The expansion tube studies include: (1) Investigation of Flow Characteristics in TQ Expansion Tube; and (2) Disturbances in the Driver Gas of a Shock Tube.
    Keywords: FLUID MECHANICS AND HEAT TRANSFER
    Type: NASA-CR-182096-SUPPL-5 , NAS 1.26:182096-SUPPL-5
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  • 7
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    Scramjet sidewall burning: Preliminary shock tunnel results (1985)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Experiments performed with a two dimensional model scramjet with particular emphasis on the effect of fuel injection from a wall are reported. Air low with a nominal Mach number of 3.5 and varied enthalpies was produced. It was found that neither hydrogen injection angle nor combustor divergence angle had any appreciable effect on thrust values while increased combustor length appeared to increase thrust levels. Specific impulse was observed to peak when hydrogen was injected at an equivalence ratio of about 2. Lowering the Mach number of the injected hydrogen at low equivalence ratios, less than 4, appeared to benefit specific impulse while hydrogen Mach number had little effect at higher equivalence ratios. When a 1:1 mixture by volume of nitrogen and oxygen is used instead of air as a test gas, it is found that hydrogen combustion is enhanced but only at high enthalpies.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-176420 , NAS 1.26:176420 , REPT-12/85
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  • 8
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    Further shock tunnel studies of scramjet phenomena (1986)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Scramjet phenomena were studied using the shock tunnel T3 at the Australian National University. Simple two dimensional models were used with a combination of wall and central injectors. Silane as an additive to hydrogen fuel was studied over a range of temperatures and pressures to evaluate its effect as an ignition aid. The film cooling effect of surface injected hydrogen was measured over a wide range of equivalence. Heat transfer measurements without injection were repeated to confirm previous indications of heating rates lower than simple flat plate predictions for laminar boundary layers in equilibrium flow. The previous results were reproduced and the discrepancies are discussed in terms of the model geometry and departures of the flow from equilibrium. In the thrust producing mode, attempts were made to increase specific impulse with wall injection. Some preliminary tests were also performed on shock induced ignition, to investigate the possibility in flight of injecting fuel upstream of the combustion chamber, where it could mix but not burn.
    Keywords: AERODYNAMICS
    Type: NASA-CR-179937 , NAS 1.26:179937 , RR-10-86
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  • 9
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    Shock tunnel studies of scramjet phenomena, supplement 6 (1993)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Reports by the staff of the University of Queensland on various research studies related to the advancement of scramjet technology are presented. These reports document the tests conducted in the reflected shock tunnel T4 and supporting research facilities that have been used to study the injection, mixing, and combustion of hydrogen fuel in generic scramjets at flow conditions typical of hypersonic flight. In addition, topics include the development of instrumentation and measurement technology, such as combustor wall shear and stream composition in pulse facilities, and numerical studies and analyses of the scramjet combustor process and the test facility operation.
    Keywords: FLUID MECHANICS AND HEAT TRANSFER
    Type: NASA-CR-191428 , NAS 1.26:191428
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  • 10
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    An investigation of a wall-injected scramjet using a shock tunnel (1992)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: Experimental results are presented for several wall-injected, constant-area scramjet experiments, performed in the T4 free-piston-driven shock tunnel. Stagnation enthalpies range from 5.7 to 10.1 MJ/kg, corresponding to flight Mach numbers of Mach 9.7 to Mach 13. Hydrogen fuel is injected from the rear of a step into a supersonic air stream, with combustion-chamber Mach numbers of 3.6 and 4.4, and static pressures of 40 kPa. The experimental results are compared to numerical predictions obtained from a two-dimensional, parabolic Navier-Stokes code, which modeled turbulence with a compressibility modified k-epsilon model, and used multi-reaction, finite-rate chemistry. The numerical and experimental results show reasonable agreement and indicate that the scramjet flows studied are predominantly mixing limited.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 92-3965
    Format: text
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