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Morgan, R. G.

Publications and source records attributed to Morgan, R. G..

29 records · Page 2
All sourcesNASA NTRSDOE OSTI
NASA NTRS1987

Scramjet testing in impulse facilities

The use of impulse facilities for gas dynamic experimentation on scramjets at high stagnation enthalpies is discussed. It is seen that, although such facilities will produce adequate stagnation enthalpies, it is necessary to compromise somewhat on the requirements for simulating test section densities, in order to allow realistic operating pressure levels. The shock tunnel and the expansion tube are briefly described, before focusing on operation of the reflected shock tunnel. It is shown that test times and flow starting times are such as to allow testing with models of reasonable size, provided that large regions of flow recirculation do not exist. Some examples of experimental scramjet studies, conducted in a reflected shock tunnel, are presented, demonstrating that the reflected shock tunnel is suitable for basic studies of scramjet flow processes. Provided that the effective enhancement of fuel calorific value by free stream 'freezing' of oxygen in the shock tunnel nozzle expansion does not prove to be an insuperable obstacle, the reflected shock tunnel may be expected to provide realistic simulation of scramjet flows up to speeds approaching earth orbital velocity.

Stalker, R. J.↗
NASA NTRS1987

Hydrogen scramjet with sidewall injection - Shock tunnel simulations

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.

Morgan, R. G.↗
NASA NTRS1987

Pressure scaling effects in a scramjet combustion chamber

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.

Morgan, R. G.↗
NASA NTRS1987

High Reynolds number heat transfer to the cold walls of a model scramjet

The results of simultaneous heat transfer and pressure measurements at the walls of three different configurations of a model scramjet are presented. The heat transfer results are compared with results empirically predicted from the pressure measurements. It is shown that the measured heat transfer rate is comparable with, or lower than, that predicted for a laminar boundary layer. A mathematical model is proposed for the film-cooling effect observed when a hydrogen fuel is injected along a wall of the scramjet. In this mathematical model, the heat transfer rate is shown to be insensitive to the velocity profile in the insulating layer of fuel. The model suggests that the cooling layer is turbulent and that 90 percent of the fuel is mixed with the air.

Paull, A.↗
NASA NTRS1986

Further shock tunnel studies of scramjet phenomena

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.

Morgan, R. G.↗
NASA NTRS1986

Hydrogen scramjet with side wall injection

Preliminary results are presented from an experimental evaluation of a hydrogen-burning wall-injection scramjet engine, performed in the free-piston shock tunnel at the Australian National University. The advantages of scramjet propulsion for high-Mach-number high-altitude flight are reviewed, and the need for an alternative to injection struts is indicated. Pressure profiles and heat-transfer measurements for constant-area and diverging ducts are presented graphically and characterized in detail. The ability of the injected flow to shield the chamber wall from the heat of the freestream flow is demonstrated. The extrapolation of the shock-tunnel results to flight conditions, however, is made difficult by the fact that a significant amount of fuel passes through a quenched zone without burning, seriously degrading overall performance.

Morgan, R. G.↗
NASA NTRS1985

Shock tunnel measurements of heat transfer in a model scramjet

The results of heat transfer measurements to the walls of a two dimensional scramjet combustion chamber in a shock tunnel are presented. Thin film heat transfer gauges on a ceramic glass substrate were used. The range of experimental conditions covered produced boundary layers ranging from laminar to transitional, as was independently checked by flow visualization. Empirical flat plate correlations, corrected for local pressure disturbances were used to make a comparison with the experimental results. In the fully laminar regime the heating rates were found to give approximate agreement with the empirical estimates. In the nonlaminar tests the heating rate is found to be well below the fully turbulent levels. It is not known at present if this is a transition effect, or if the pressure corrected flat plate turbulent correlations do not apply to the configuration used.

Morgan, R. G.↗
NASA NTRS1985

Scramjet sidewall burning: Preliminary shock tunnel results

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.

Morgan, R. G.↗
NASA NTRS1972

Particle impact and optical property analysis of the surfaces of Surveyor 3 materials

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.

Anderson, D. L.↗