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At least 739 records · Page 41
All sourcesNASA NTRSDOE OSTI
NASA NTRS1976

Wind tunnel investigation of Nacelle-Airframe interference at Mach numbers of 0.9 to 1.4-force data

Detailed interference force and pressure data were obtained on a representative wing-body-nacelle combination at Mach numbers of 0.9 to 1.4. The model consisted of a delta wing-body aerodynamic force model with four independently supported nacelles located beneath the wing-body combination. The model was mounted on a six-component force balance, and the left-hand wing was pressure-instrumented. Each of the two right-hand nacelles was mounted on a six-component force balance housed in the thickness of the nacelle, while each of the left-hand nacelles was pressure-instrumented. The primary variables examined included Mach number, angle of attack, nacelle position, and nacelle mass-flow ratio. Four different configurations were tested to identify various interference forces and pressures on each component; these included tests of the isolated nacelle, the isolated wing-body combination, the four nacelles as a unit, and the total wing-body-nacelle combination. Nacelle axial location, relative to both the wing-body combination and to each other, was the most important variable in determining the net interference among the components.

Bencze, D. P.↗
NASA NTRS1976

Wind tunnel investigation of Nacelle-Airframe interference at Mach numbers of 0.9 to 1.4-pressure data, volume 2

Detailed interference force and pressure data were obtained on a representative wing-body nacelle combination at Mach numbers of 0.9 to 1.4. The model consisted of a delta wing-body aerodynamic force model with four independently supported nacelles located beneath the wing-body combination. The primary variables examined included Mach number, angle of attack, nacelle position, and nacelle mass flow ratio. Four different configurations were tested to identify various interference forces and pressures on each component; these included tests of the isolated nacelle, the isolated wing-body combination, the four nacelles as a unit, and the total wing-body-nacelle combination. Nacelle axial location, relative to both the wing-body combination and to each other, was the most important variable in determining the net interference among the components. The overall interference effects were found to be essentially constant over the operating angle-of-attack range of the configuration, and nearly independent of nacelle mass flow ratio.

Bencze, D. P.↗
NASA NTRS1976

Wind tunnel investigation of nacelle-airframe interference at Mach numbers of 0.9 to 1.4 - pressure data, volume 1

Detailed interference force and pressure data were obtained on a representative wing-body nacelle combination at Mach numbers of 0.9 to 1.4. The model consisted of a delta wing-body aerodynamic force model with four independently supported nacelles located beneath the wing-body combination. The model was mounted on a six component force balance, and the left hand wing was pressure instrumented. Each of the two right hand nacelles was mounted on a six component force balance housed in the thickness of the nacelle, while each of the left hand nacelles was pressure instrumented. The primary variables examined included Mach number, angle of attack, nacelle position, and nacelle mass flow ratio. Nacelle axial location, relative to both the wing-body combination and to each other, was the most important variable in determining the net interference among the components.

Bencze, D. P.↗
NASA NTRS1977

An analytical model for predicting the aerodynamic performance of a turbine cascade with film cooling

Various analytical approaches to predicting the performance of film cooled turbine blades are reviewed. A two-dimensional cascade flow solution is developed for calculating the effects of the coolant injection on the total flow field. This solution is used with an available analytical performance predicting method to provide an improved method. Comparisons are made with experimental data and other analytical results.

Mcfarland, E. R.↗
NASA NTRS1978

Subsonic longitudinal and lateral-directional static aerodynamic characteristics for a model with swept back and swept forward wings

A general research fighter model was tested in the Langley 7- by 10-foot high speed tunnel at a Mach number of 0.3. With a conventional empennage, the model was tested with the wing in a 60 deg swept back configuration and in a 32 deg swept forward configuration. The 32 deg swept forward configuration was also tested with a strake. Very limited data was obtained with a wing in a 50 deg swept back configuration and a 7 deg swept forward configuration. The angle of attack range was from approximately -4 deg to 48 deg at sideslip angles of 0 deg, -5 deg, and 5 deg.

Huffman, J. K.↗
NASA NTRS1979

The effect of canard relative size and vertical location on the subsonic longitudinal and lateral-directional static aerodynamic characteristics for a model with a swept forward wing

A general research fighter model was tested in the Langley 7- by 10-foot high speed tunnel at a Mach number of 0.3. The model was tested with a 32 deg swept forward wing mounted in mid-, low-, and high-wing positions. For the mid-wing configuration, the model was tested with a 51.7 deg swept back canard mounted in mid-, low-, and high-canard positions. For the mid-wing mid-canard and the mid-wing high-canard configurations, canards of similar planform having two different areas were tested. The angle-of-attack range was from approximately -4 deg to 48 deg at sideslip angles of 0 deg, -5 deg, and 5 deg.

Huffman, J. K.↗
NASA NTRS1973

Effects of wing height on low-speed aerodynamic characteristics of a model having a 42 deg swept wing, a supercritical airfoil, double-slotted flaps, and a low tail

A low speed investigation was conducted in the Langley V/STOL tunnel to determine the static longitudinal lateral stability characteristics of a general research model with the wing in a high position and a low position on the fuselage. The model had a wing with a quarter chord sweep of 42 deg, an aspect ratio of 6.78, a supercritical airfoil, and a high lift system which consisted of a leading edge slat and a double slotted flap. Various slat and flap deflections represented clean, take off, and landing configurations. A 45 deg swept horizontal tail located slightly below the fuselage center line was investigated with both the low and high wing configurations.

Fournier, P. G.↗
NASA NTRS1986

A comparison of the acoustic and aerodynamic measurements of a model rotor tested in two anechoic wind tunnels

Two aeroacoustic facilities--the CEPRA 19 in France and the DNW in the Netherlands--are compared. The two facilities have unique acoustic characteristics that make them appropriate for acoustic testing of model-scale helicopter rotors. An identical pressure-instrumented model-scale rotor was tested in each facility and acoustic test results are compared with full-scale-rotor test results. Blade surface pressures measured in both tunnels were used to correlated nominal rotor operating conditions in each tunnel, and also used to assess the steadiness of the rotor in each tunnel's flow. In-the-flow rotor acoustic signatures at moderate forward speeds (35-50 m/sec) are presented for each facility and discussed in relation to the differences in tunnel geometries and aeroacoustic characteristics. Both reports are presented in appendices to this paper. ;.);

Boxwell, D. A.↗
NASA NTRS1986

A comparison of the acoustic and aerodynamic measurements of a model rotor tested in two anechoic wind tunnels

Two aeroacoustic facilities - the CEPRA 19 in France and the DNW in the Netherlands - are compared. The two facilities have unique acoustic characteristics that make them appropriate for acoustic testing of model-scale helicopter rotors. An identical pressure-instrumented model-scale rotor was tested in each facility and acoustic test results are compared with full-scale-rotor test results. Blade surface pressures measured in both tunnels were used to correlated nominal rotor operating conditions in each tunnel, and also used to assess the steadiness of the rotor in each tunnel's flow. In-the-flow rotor acoustic signatures at moderate forward speeds (35-50 m/sec) are presented for each facility and discussed in relation to the differences in tunnel geometries and aeroacoustic characteristics. Both reports are presented in appendices to this paper.

Boxwell, D. A.↗
NASA NTRS1987

Wall jet analysis for circulation control aerodynamics. Part 2: Zonal modeling concepts for wall jet/potential flow coupling

Work currently in progress to update an existing transonic circulation control airfoil analysis method is described. Existing methods suffer from two dificiencies: the inability to predict the shock structure of the underexpanded supersonic jets; and the insensitivity of the calculation to small changes in the Coanda surface geometry. A method developed for the analysis of jet exhaust plumes in supersonic flow is being modified for the case of the underexpanded wall jet. In the subsonic case, the same wall jet model was modified to include the calculation of the normal pressure gradient. This model is currently being coupled with the transonic circulation control airfoil analysis.

Dvorak, Frank A.↗
NASA NTRS1932

The aerodynamic characteristics of a model wing having a split flap deflected downward and moved to the rear

Tests were made on a model wing with three different sized split trailing-edged flaps, in the NACA 7 by 10 foot wind tunnel. The flaps were formed of the lower rear portion of the wing and were rotated downward about axes at their front edges. The lift, drag, and center of pressure were measured with the axis in its original position and also with it moved back in even steps to the trailing edge of the main wing, giving in effect an increase in area. The split flaps when deflected about their original axis locations gave slightly higher maximum lift coefficients than conventional trailing-edge flaps, and the lift coefficients were increased still further by moving the axes toward the rear. The highest value of C(sub L max), which was obtained with the largest flap hinged at 90 per cent of the chord from the leading edge, was 2.52 as compared with 1.27 for the basic wing.

Weick, Fred E↗