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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 703 records · Page 39
All sourcesNASA NTRSDOE OSTI
NASA NTRS1980

Investigation of advanced thrust vectoring exhaust systems for high speed propulsive lift

The paper presents the results of a wind tunnel investigation conducted at the NASA-Langley research center to determine thrust vectoring/induced lift characteristics of advanced exhaust nozzle concepts installed on a supersonic tactical airplane model. Specific test objectives include: (1) basic aerodynamics of a wing body configuration, (2) investigation of induced lift effects, (3) evaluation of static and forward speed performance, and (4) the effectiveness of a canard surface to trim thrust vectoring/induced lift forces and moments.

Hutchison, R. A.↗
NASA NTRS1980

Prediction of unsuppressed jet engine exhaust noise in flight from static data

In order to assess the impact of aircraft noise on the environment in the vicinity of an airport, it is essential that a methodology be developed for predicting in-flight exhaust noise from static data. Such a methodology is presented in this paper and is compared with experimental data for several unsuppressed turbojet engines. For each engine, static data over a range of jet velocities are compared with the predicted jet mixing noise and shock-cell noise. The static engine noise over and above the jet and shock noises is identified as 'excess' noise. The excess noise data are then empirically correlated to smooth the spectral and directivity relations and account for variations in test conditions. This excess noise is then projected to flight based on the assumption that the only effects of flight are a Doppler frequency shift and a level change. The effects of flight on jet mixing noise and shock noise are computed by published NASA methods.

Stone, J. R.↗
NASA NTRS1980

Model aerodynamic test results for two variable cycle engine coannular exhaust systems at simulated takeoff and cruise conditions

Wind tunnel tests were conducted to evaluate the aerodynamic performance of a coannular exhaust nozzle for a proposed variable stream control supersonic propulsion system. Tests were conducted with two simulated configurations differing primarily in the fan duct flowpaths: a short flap mechanism for fan stream control with an isentropic contoured flow splitter, and an iris fan nozzle with a conical flow splitter. Both designs feature a translating primary plug and an auxiliary inlet ejector. Tests were conducted at takeoff and simulated cruise conditions. Data were acquired at Mach numbers of 0, 0.36, 0.9, and 2.0 for a wide range of nozzle operating conditions. At simulated supersonic cruise, both configurations demonstrated good performance, comparable to levels assumed in earlier advanced supersonic propulsion studies. However, at subsonic cruise, both configurations exhibited performance that was 6 to 7.5 percent less than the study assumptions. At take off conditions, the iris configuration performance approached the assumed levels, while the short flap design was 4 to 6 percent less.

Nelson, D. P.↗
NASA NTRS1980

Effects of sidewall geometry on the installed performance of nonaxisymmetric convergent-divergent exhaust nozzles

The investigation was conducted at static conditions and over a Mach number range from 0.6 to 1.2. Angle of attack was held constant at 0 deg. High pressure air was used to simulate jet exhaust flow at ratios of jet total pressure to free-stream static pressure from 1 (jet off) to approximately 10. Sidewall cutback appears to be a viable way of reducing nozzle weight and cooling requirements without compromising installed performance.

Yetter, J. A.↗
NASA NTRS1981

Model aerodynamic test results for two variable cycle engine coannular exhaust systems at simulated takeoff and cruise conditions. Comprehensive data report. Volume 2: Tabulated aeroynamic data book 1

Tabulated data from wind tunnel tests conducted to evaluate the aerodynamic performance of an advanced coannular exhaust nozzle for a future supersonic propulsion system are presented. Tests were conducted with two test configurations: (1) a short flap mechanism for fan stream control with an isentropic contoured flow splitter, and (2) an iris fan nozzle with a conical flow splitter. Both designs feature a translating primary plug and an auxiliary inlet ejector. Tests were conducted at takeoff and simulated cruise conditions. Data were acquired at Mach numbers of 0, 0.36, 0.9, and 2.0 for a wide range of nozzle operating conditions. At simulated supersonic cruise, both configurations demonstrated good performance, comparable to levels assumed in earlier advanced supersonic propulsion studies. However, at subsonic cruise, both configurations exhibited performance that was 6 to 7.5 percent less than the study assumptions. At takeoff conditions, the iris configuration performance approached the assumed levels, while the short flap design was 4 to 6 percent less. Data are provided through test run 25.

Nelson, D. P.↗
NASA NTRS1981

Axisymmetric & non-axisymmetric exhaust jet induced-effects on a V/STOL vehicle design. Part 1: Data presentation

A 1/8 scale jet-effects model was tested in the NASA Ames 11 ft transonic tunnel at static conditions and over a range of Mach numbers from 0.4 to 1.4. The data presented show that significant differences in aeropropulsion performance can be expected by varying the exhaust nozzle type and its geometric parameters on a V/STOL underwing nacelle installation.

Schnell, W. C.↗
NASA NTRS1981

Mixing effectiveness test of an exhaust gas mixer in a high bypass turbofan at altitude

Thermal mixing effectiveness characteristics of an eighteen lobe, scalloped and unscalloped, partial, forced mixer were measured in a high-bypass turbofan engine. Data were also obtained without the mixer installed, i.e., free mixing. Tests were conducted at four combinations of simulated flight conditions from 0.3 to 0.8 Mach number and from 6,096 meters (20,000 ft) to 13,715 m (45,000 ft) altitude, mixing chamber lengths of L/D=0.52 and 0.65 were tested. For this range of test conditions and mixer configurations, the forced mixing effectiveness varied from 59 to 68 percent. Values of mixing effectiveness and total pressure loss were calculated from temperature and pressure data obtained at the mixer inlet and exhaust nozzle exit.

Cullom, R. R.↗
NASA NTRS1980

Effect of a semi-annular thermal acoustic shield on jet exhaust noise

The effect of a semi-annular acoustic shield on jet exhaust noise is investigated with the rationale that such a configuration would reduce or eliminate the multiple reflection mechanism. A limited range of flow conditions for one nozzle/shield configuration were studied at model scale. Noise measurements for a 10 cm conical nozzle with a semi-annular acoustical shield are presented in terms of lossless free field data at various angular locations with respect to the nozzle. Measurements were made on both the shielded and unshielded sides of the nozzle. Model scale overall sound pressure level directivity patterns and comparisons of model scale spectral data are provided. The results show that a semi-annular thermal acoustic shield consisting of a low velocity, high temperature gas stream partially surrounding a central jet exhibits lower noise levels than when the central jet is operated alone. The results are presented parametrically, showing the effects of various shield and central system velocities and temperatures.

Goodykoontz, J.↗
NASA NTRS1981

Mixing effectiveness test of an exhaust gas mixer in a high bypass turbofan at altitude

Thermal mixing effectiveness characteristics of an eighteen lobe, scalloped and unscalloped, partial, forced mixer were measured in a high-bypass turbofan engine. Data were also obtained without the mixer installed, i.e. free mixing. Tests were conducted at four combinations of simulated flight conditions from 0.3 to 0.8 Mach number and from 6,096 meters (20,000 ft) to 13,715 m (45,000 ft) altitude. Mixing chamber lengths of L/D = 0.52 and 0.65 were tested. For this range of test conditions and mixer configurations the forced mixing effectiveness varied from 59 to 68 percent. Values of mixing effectiveness and total pressure loss were calculated from temperature and pressure data obtained at the mixer inlet and exhaust nozzle exit.

Cullom, R. R.↗
NASA NTRS1983

Exhaust-Plume Impingement Characteristics

Model for rocket-exhaust plumes calculates effects of heat, pressure, and particle-impingement on nearby structures. Model has potential industrial applications in studies of impingement of multiphase mixtures inside coalgasification equipment and turbines.

Smith, S. D.↗
NASA NTRS1981

Prediction of subsonic aircraft flows with jet exhaust interactions

A numerical procedure to calculate the flow fields resulting from the viscous inviscid interactions that occur when a strong jet exhaust and aircraft flow field coupling exists was developed. The approach divides the interaction region into zones which are either predominantly viscous or inviscid. The flow in the inviscid zone, which surrounds most of the aircraft, is calculated using an existing potential flow code. The viscous flow zone, which encompasses the jet plume, is modeled using a parabolized Navier-Stokes code. The procedure features the coupling of the zonal solutions such that sufficient information is transferred between the zones to preserve the effects of the interactions. The zonal boundaries overlap and the boundary conditions are the information link between zones. An iteration scheme iterates the coupled analysis until convergence has been obtained.

Roberts, D. W.↗
NASA NTRS1982

Effect of vacuum exhaust pressure on the performance of MHD ducts at high B-field

The effect of area ratio variation on the performance of a supersonic Hall MHD duct is investigated. Results indicate that for a given combustion pressure there exists an area ratio below which the power generating region of the duct is shock free and the power output increases linearly with the square of the magnetic field. For area ratios greater than this, a shock forms in the power generating region which moves upstream with increasing magnetic field strength resulting in a less rapid raise in the power output. The shock can be moved downstream by either increasing the combustion pressure of decreasing the exhaust pressure. The influence of these effects upon duct performance is presented.

Smith, J. M.↗
NASA NTRS1982

RAXJET: A computer program for predicting transonic, axisymmetric flow over nozzle afterbodies with supersonic jet exhausts

A viscous-inviscid interaction method to calculate the subsonic and transonic flow over nozzle afterbodies with supersonic jet exhausts was developed. The method iteratively combines a relaxation solution of the full potential equation for the inviscid external flow, a shock capturing-shock fitting inviscid jet solution, an integral boundary layer solution, a control volume method for treating separated flows, and an overlaid mixing layer solution. A computer program called RAXJET which incorporates the method, illustrates the predictive capabilities of the method by comparison with experimental data is described, a user's guide to the computer program is provided. The method accurately predicts afterbody pressures, drag, and flow field properties for attached and separated flows for which no shock induced separation occurs.

Wilmoth, R. G.↗
NASA NTRS1982

Effects of installation of F101 DFE exhaust nozzles on the afterbody-nozzle characteristics of the F-14 airplane

A wind-tunnel investigation was conducted to determine the effects of F101 DFE (derivative fighter engine) nozzle axial positioning on the afterbody-nozzle longitudinal aerodynamic characteristics of the F-14 airplane. The model was tested in the Langley 16-Foot Transonic Tunnel at Mach numbers from 0.7 to 1.25 and angles of attack from about -2 to 6 degrees. Compressed air was used to simulate nozzle exhaust flow at jet total-pressure ratios from 1 (jet off) to about 8. The results of the investigation show that for subsonic Mach numbers the intermediate cruise nozzle position of the three positions tested resulted in the lowest drag.

Reubush, D. E.↗