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Compton, W. B., III

Publications and source records attributed to Compton, W. B., III.

10 records
All sourcesNASA NTRSDOE OSTI
NASA NTRS1984

An experimental investigation of nacelle-pylon installation on an unswept wing at subsonic and transonic speeds

A wind tunnel investigation was conducted to determine the aerodynamic interference associated with the installation of a long duct, flow-through nacelle on a straight unswept untapered supercritical wing. Experimental data was obtained for the verification of computational prediction techniques. The model was tested in the 16-Foot Transonic Tunnel at Mach numbers from 0.20 to 0.875 and at angles of attack from about 0 deg to 5 deg. The results of the investigation show that strong viscous and compressibility effects are present at the transonic Mach numbers. Numerical comparisons show that linear theory is adequate for subsonic Mach number flow prediction, but is inadequate for prediction of the extreme flow conditions that exist at the transonic Mach numbers.

Carlson, J. R.↗
NASA NTRS1983

Three-dimensional Euler solutions for long-duct nacelles

A three-dimensional Euler-equation computational technique has been developed to solve for the transonic flow past flow-through nacelles. The technique employs an approximately-factored alternating-direction implicit numerical algorithm and a radiation treatment of the outflow boundary. Studies are presented which show that the radiation treatment gives better numerical convergence than the condition of specifying the pressure at the outflow boundary. Calculations made with the technique are presented for a long-duct turbofan engine nacelle at a Mach number of 0.80 and angles of attack of 0 deg and 4 deg. Good agreement is shown between the computational results and wind-tunnel data. Problem areas are identified and recommendations are made for further numerical studies.

Compton, W. B., III↗
NASA NTRS1976

Jet exhaust and support interference effects on the transonic aerodynamic characteristics of a fighter model with two widely spaced engines

Jet exhaust, nozzle installation, and model support interference effects on the longitudinal aerodynamic characteristics of a twin-engine fighter model were determined. Realistic jet exhaust nozzle configurations and a reference configuration with a simulated vertical-tail support were tested. Free-stream Mach number was varied from 0.6 to 1.2, and model angle of attack from 0 deg to 9 deg. The jet exhaust affected drag more than it affected lift and pitching moment. The largest effects occurred at a Mach number of 0.9 and for the afterburning mode of exhaust nozzle operation. The combined differences between the aerodynamic characteristics of the realistic and reference configurations (which were due to afterbody and nozzle contours, jet operation, and simulated reference support interference) were considerably different from those for the jet interference alone.

Compton, W. B., III↗
NASA NTRS1975

Effects of jet exhaust gas properties on exhaust simulation and afterbody drag

The effect of varying the jet exhaust's ratio of specific heats, gas constant, and temperature on airplane afterbody drag was investigated. Jet exhaust simulation parameters were evaluated also. Subsonic and transonic tests were made using a single nacelle model with afterbodies having boattail angles of 10 deg and 20 deg. Besides air, three other jet exhaust gases were investigated. The ratios of specific heats, gas constants, and total temperatures of the four exhaust gases ranged from 1.40 to 1.26, 287 to 376 J/kg-K, and 300 to 1013 K, respectively. For steep boattail angles, and transonic speeds and typical turbojet pressure ratios, the current data indicate that the use of air to simulate a dry turbojet exhaust can result in an overprediction of afterbody drag as high as 17 percent of the dry turbojet value.

Compton, W. B., III↗
NASA NTRS1975

Contribution of the National Aeronautics and Space Administration Langley Research Center

As part of a special international effort, three nozzles were designed and tested on single nacelle models in wind tunnels of several nations belonging to the North Atlantic Treaty Organization. All three of these nozzles were investigated in the Langley 16-foot transonic wind tunnel at the National Aeronautics and Space Administration's Langley Research Center. Langley Research Center also contributed theoretical calculations of the jet plume boundary and afterbody pressures. The calculations were obtained using an iterative solution which combined the inviscid Douglas Neumann method for the external flow with the method of characteristics for the flow in the jet plume. For the investigation, the nozzles were mounted on a single nacelle model 15.24 centimeters in diameter and 162.56 centimeters long. Tests were made at free stream Mach number from 0.4 to 1.2, and at Reynolds numbers per meter from 7.38 million to 13.78 million depending on the Mach number. Four types of data were recorded: afterbody pressure data, afterbody force data, model boundary layer data, and tunnel wall pressure data. The ratio of jet total pressure to free stream static pressure ranged up to 8.5. A description of the wind tunnel, model, and test procedure is included.

Compton, W. B., III↗
NASA NTRS1975

An experimental study of jet exhaust simulation

Afterbody drag predictions for jet aircraft are usually made experimentally with the jet exhaust flow simulated. The physical gas properties of the fluid used for the model jet exhaust can affect the accuracy of simulation of the airplane's jet exhaust plume. The effect of the accuracy of this simulation on afterbody drag was investigated by wind-tunnel tests with single engine model. In addition to unheated air as the exhaust gas, the decomposition products of three different concentrations of hydrogen peroxide were utilized. The air jet simulation consistently resulted in higher boattail drag than hydrogen peroxide simulation. The differences in drag for the various exhaust gases are attributed to different plume shapes and entrainment properties of the gases. The largest differences in drag due to exhaust gas properties were obtained for the combination of high transonic Mach numbers and high boattail angles. For these conditions, the current data indicate that the use of air to simulate a nonafterburning turbojet exhaust can result in an increase in afterbody amounting to 20 percent of the nonafterburning turbojet value.

Compton, W. B., III↗
NASA NTRS1974

An Experimental Study of Jet Exhaust Simulation

An investigation was conducted to determine the effect of varying the jet exhaust ratio of specific heats, gas constants, and temperatures on jet interference on afterbody drag. Jet exhaust simulation parameters were also evaluated. In addition to air, three other exhaust gases, each with a different value of each of the gas parameters, were tested. The range of the ratios of specific heats, gas constants, and total temperatures of the four gases are identified. Tests were made using a single nacelle model with afterbodies having boattail angles of 10 degrees and 20 degrees with sonic and Mach two jet exits. Wind tunnel tests were conducted through a Mach number range of 0.60 to 1.20 and a Reynolds number per meter from 10.06 million to 14.05 million.

Compton, W. B., III↗
NASA NTRS1972

Jet effects on the drag of conical afterbodies at supersonic speeds

A parametric investigation has been conducted to determine the jet effects on the boattail drag of nozzles with truncated conical afterbodies. The boattail drag for nozzle configurations with boattail angles of 3 deg, 5 deg, and 10 deg and ratios of boattail length to maximum diameter of 1.0, 0.8, and 0.6 was compared for the jet-off condition and for a wide range of jet pressure ratios. A nozzle configuration with a circular-arc boattail was tested also. The tests were run at Mach numbers of 1.83 and 2.20 with the model at an angle of attack of 0 deg.

Compton, W. B., III↗