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Nugent, J.

Publications and source records attributed to Nugent, J..

15 records
All sourcesNASA NTRSDOE OSTI
NASA NTRS1984

Results of a wind tunnel/flight test program to compare afterbody/nozzle pressures on a 1/12 scale model and an F-15 aircraft

In 1975 NASA Dryden Flight Research Facility received the No. 2 prototype F-15 aircraft from the USAF to conduct the F-15 Propulsion/Airframe Interactions Program. About the same time, NASA Langley Research Center acquired a 1/12 scale F-15 propulsion model, whose size made it suitable for detailed afterbody/nozzle static pressure distribution studies. Close coordination between Langley and Dryden assured identical orifice locations and nozzle geometries on the model and aircraft. This paper discusses the sequence of the test programs and how retesting the model after completion of the flight tests greatly increased the ability to match hardware and test conditions. The experience gained over the past decade from involvement in the program should prove valuable to any future programs attempting to match wind tunnel and flight test conditions and hardware.

Pendergraft, O. C., Jr.↗
NASA NTRS1983

Pressures measured in flight on the aft fuselage and external nozzle of a twin-jet fighter

Fuselage, boundary layer, and nozzle pressures were measured in flight for a twin jet fighter over a Mach number range from 0.60 to 2.00 at test altitudes of 6100, 10,700, and 13,700 meters for angles of attack ranging from 0 deg to 7 deg. Test data were analyzed to find the effects of the propulsion system geometry. The flight variables, and flow interference. The aft fuselage flow field was complex and showed the influence of the vertical tail, nacelle contour, and the wing. Changes in the boattail angle of either engine affected upper fuselage and lower fuselage pressure coefficients upstream of the nozzle. Boundary layer profiles at the forward and aft locations on the upper nacelles were relatively insensitive to Mach number and altitude. Boundary layer thickness decreased at both stations as angle of attack increased above 4 deg. Nozzle pressure coefficient was influenced by the vertical tail, horizontal tail boom, and nozzle interfairing; the last two tended to separate flow over the top of the nozzle from flow over the bottom of the nozzle. The left nozzle axial force coefficient was most affected by Mach number and left nozzle boattail angle. At Mach 0.90, the nozzle axial force coefficient was 0.0013.

Nugent, J.↗
NASA NTRS1982

Selected results of the F-15 propulsion interactions program

A better understanding of propulsion system/airframe flow interactions could aid in the reduction of aircraft drag. For this purpose, NASA and the United States Air Force have conducted a series of wind-tunnel and flight tests on the F-15 airplane. This paper presents a correlation of flight test data from tests conducted at the NASA Dryden Flight Research Facility of the Ames Research Center, with data obtained from wind-tunnel tests. Flights were made at stabilized Mach numbers around 0.6, 0.9, 1.2, and 1.5 with accelerations up to near Mach number 2. Wind-tunnel tests used a 7.5 percent-scale F-15 inlet/airframe model. Flight and wind-tunnel pressure coefficients showed good agreement in most cases. Correlation of interaction effects caused by changes in cowl angle, angle-of-attack, and Mach number are presented. For the afterbody region, the pressure coefficients on the nozzle surfaces were influenced by boattail angles and Mach number. Boundary-layer thickness decreased as angle of attack increased above 4 deg.

Webb, L. D.↗
NASA NTRS1981

Recent propulsion system flight tests at the NASA Dryden Flight Research Center

The article presents a summary of the propulsion system tests conducted on a number of aircraft at the NASA Dryden Flight Research Center. The tests included digital engine control systems, engine-inlet compatibility, inlet-airframe interactions, nozzle-boattail drag and advanced turboprop acoustics. Among the aircraft evaluated were the F-15, HiMAT, F-14, and the JetStar.

Burcham, F. W., Jr.↗
NASA NTRS1980

Flight-measured effects of boattail angle and Mach number on the nozzle afterbody flow of a twin-jet fighter

The paper presents the flight-measured nozzle afterbody surface pressures and engine exhaust nozzle pressure-area integrated axial force coefficients on a twin-jet fighter for varying boattail angles. The objective of the tests was to contribute to a full-scale flight data base applicable to the nozzle afterbody drag of advanced tactical fighter concepts. The data were acquired during the NASA F-15 Propulsion/Airframe Interactions Flight Research Program. Nozzle boattail angles from 7.7 deg to 18.1 deg were investigated. Results are presented for cruise angle of attack at Mach numbers from 0.6 to 2.0 at altitudes from 20,000 to 45,000 feet. The data show the nozle axial force coefficients to be a strong function of nozzle boattail angle and Mach number.

Plant, T. J.↗
NASA NTRS1979

Shock-swallowing air sensor

An air-data probe allows air to flow through it so that supersonic and hypersonic shock waves form behind pressure measuring orifices and tube instead of directly on them. Measured pressures are close to those in free-flowing air and are used to determine mach numbers of flying aircraft.

Nugent, J.↗
NASA NTRS1979

Reynolds number, scale and frequency content effects on F-15 inlet instantaneous distortion

An inlet instantaneous distortion study program sponsored by NASA was recently completed using an F-15 fighter aircraft. Peak distortion data from subscale inlet model wind tunnel tests are shown to be representative of full-scale flight test peak distortion. The effects on peak distortion are investigated for engine presence, Reynolds number, scale and frequency content. Data are presented which show that: (1) the effect of engine presence on total pressure recovery, peak distortion, and turbulence is small but favorable, (2) increasing the Reynolds number increases total pressure recovery, decreases peak distortion, and decreases turbulence, and (3) increasing the filter cutoff frequency increases the calculated values of both peak distortion and turbulence.

Stevens, C. H.↗
NASA NTRS1978

X-rays from Alpha Centauri

HEAO 1 observations of soft X-ray emission from a point source in the vicinity of Alpha Cen are reported. The source, designated H1437-61, is tentatively identified with Alpha Cen, and an X-ray luminosity comparable to that of the sun in an active state is estimated. A temperature of about 500,000 K and an emission integral of 5 x 10 to the 50th per cu cm are obtained. Coronal emission is suggested as the X-ray-producing mechanism.

Nugent, J.↗
NASA NTRS1978

Status of a nozzle-airframe study of a highly maneuverable fighter

NASA is sponsoring a research program that uses coordinated wind tunnel and flight tests to investigate nozzle-airframe flow interactions. The program objective is to compare transonic flight and wind tunnel measurements over a wide Reynolds number range. The paper discusses the progress of the program and the coordination of the wind tunnel and flight tests with regard to program elements, model-airplane differences, instrument locations, and test conditions. The real-time feedback techniques used to obtain steady flight conditions are presented. Available wind tunnel results are presented for the jet effects model showing the influence of the rear-end geometry and test variables on nozzle drag. Available flight results show the effect of the variable inlet ramp angle and angle of attack on fuselage pressures and upper surface boundary layers.

Nugent, J.↗
NASA NTRS1975

Exploratory wind tunnel tests of a shock-swallowing air data sensor at a Mach number of approximately 1.83

The test probe was designed to measure free-stream Mach number and could be incorporated into a conventional airspeed nose boom installation. Tests were conducted in the Langley 4-by 4-foot supersonic pressure tunnel with an approximate angle of attack test range of -5 deg to 15 deg and an approximate angle of sideslip test range of + or - 4 deg. The probe incorporated a variable exit area which permitted internal flow. The internal flow caused the bow shock to be swallowed. Mach number was determined with a small axially movable internal total pressure tube and a series of fixed internal static pressure orifices. Mach number error was at a minimum when the total pressure tube was close to the probe tip. For four of the five tips tested, the Mach number error derived by averaging two static pressures measured at horizontally opposed positions near the probe entrance were least sensitive to angle of attack changes. The same orifices were also used to derive parameters that gave indications of flow direction.

Nugent, J.↗
NASA NTRS1974

Flight-measured inlet pressure transients accompanying engine compressor surges on the F-111A airplane

Two-F-111A airplanes were subjected to conditions that caused engine compressor surges and accompanying duct hammershock pressure transients. Flight speed ranged from Mach 0.71 to Mach 2.23, and altitude varied from approximately 3200 meters to 14,500 meters. A wide range of compressor pressure ratios was covered. Stabilized free-stream, engine, and duct conditions were established before each compressor surge. Dynamic pressure instrumentation at the compressor face and in the duct recorded the pressure transients associated with the surges. Hammershock pressures were analyzed with respect to the stabilized conditions preceding the compressor surges. The hammershock transients caused large pressure rises at the compressor face and in the duct. Hammershock pressure ratios at the compressor face were not affected by free-stream Mach number or altitude but were functions of engine variables, such as compressor pressure ratio. The maximum hammershock pressure ratio of approximately 1.83 occurred at a compressor pressure ratio of approximately 21.7.

Nugent, J.↗
NASA NTRS1972

Flight-test results from two total temperature probes for air-data measurements up to 2014 K (3625 R)

An experimental temperature probe package containing a fluidic oscillator temperature probe and a shielded thermocouple temperature probe was tested during several X-15 flights. The X-15 flights provided greatly varying test conditions, including a wide range of rapidly changing total temperatures and Mach numbers which extended from subsonic to hypersonic speeds. Within restricted ranges of free-stream Mach number, free-stream unit weight flow, and local stagnation pressure, both probes yielded ramp outputs of temperature parallel to ramp inputs of free-stream total temperature. Within these ranges both probes were used to determine total temperature in the Mach 6 temperature environment. Because ambient temperature was known, both probes were used to estimate velocity and Mach number.

Nugent, J.↗