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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 289 records · Page 16
All sourcesNASA NTRSDOE OSTI
NASA NTRS1972

Acoustic characteristics of large-scale STOL model at forward speed

Wind-tunnel investigations of the acoustic characteristics of the externally blown jet flap (EBF) and augmentor wing STOL concepts are dicussed. The large-scale EBF model was equipped with a triple-slotted blown by four JT15D turbofan engines with circular, coannular exhaust nozzles. The large-scale augmentor wing model was equipped with an unlined augmentor blown by a slot primary nozzle. The effects of airspeed and angle of attack on the acoustics of the EBF were small. At a forward speed of 60 knots, the impingement noise of the landing flap was approximately 2 db lower than in the static tests. Angle of attack increased the impingement noise approximately 0.1 decibels per degree. Flap deflection had a greater effect on the acoustics of the augmentor wing than did airspeed. For a nozzle pressure ratio of 1.9, the peak perceived noise level of the landing flap was 3 to 5 PNdb higher than that of the takeoff flap. The total sound power was also significantly higher for landing indicating that turning in the augmentor generated acoustic energy. Airspeed produced a small aft shift in acoustic directivity with no significant change in the peak perceived noise levels or sound power levels.

Falarski, M. D.↗
NASA NTRS1973

Supersonic aerodynamic characteristics of the North American Rockwell ATP shuttle orbiter

A wind tunnel study to determine the supersonic aerodynamic characteristics of a 0.01925-scale model of the space shuttle orbiter configuration is reported. The model consisted of a low-finess-ratio body with a blended 50 swept delta wing forming an ogee planform and a center-line-mounted vertical tail. Tests were made at Mach numbers from 1.90 to 4.63, at angles of attack from -6 to 30, at angles of sideslip of 0 and 3, and at a Reynolds number, based on body length, of 5.3x 1 million.

Ware, G. M.↗
NASA NTRS1974

A simulator investigation of engine failure compensation for powered-lift STOL aircraft

A piloted simulator investigation of various engine failure compensation concepts for powered-lift STOL aircraft was carried out at the Ames Research Center. The purpose of this investigation was to determine the influence of engine failure compensation on recovery from an engine failure during the landing approach and on the precision of the STOL landing. The various concepts include: (1) cockpit warning lights to cue the pilot of an engine failure, (2) programmed thrust and roll trim compensation, (3) thrust command and (4) flight-path stabilization. The aircraft simulated was a 150 passenger four-engine, externally blown flap civil STOL transport having a 90 psf wing loading and a .56 thrust to weight ratio. Results of the simulation indicate that the combination of thrust command and flight-path stabilization offered the best engine-out landing performance in turbulence and did so over the entire range of altitudes for which engine failures occurred.

Nieuwenhuijse, A. W.↗
NASA NTRS1975

Method of representation of acoustic spectra and reflection corrections applied to externally blown flap noise

A computer method for obtaining a rational function representation of an acoustic spectrum and for correcting reflection effects is introduced. The functional representation provides a means of compact storage of data and the nucleus of the data analysis method. The method is applied to noise from a full-scale externally blown flap system with a quiet 6:1 bypass ratio turbofan engine and a three-flap wing section designed to simulate the take-off condition of a conceptual STOL aircraft.

Miles, J. H.↗
NASA NTRS1923

Speed limits of aircraft

This paper is restricted to the question of attainable speed limits and attacks the problem from different angles. Theoretical limits due to air resistance are presented along with design factors which may affect speed such as wing loads, wing areas, wing section shifting, landing speeds, drag-lift ratios, and power coefficients.

Everling, E↗
NASA NTRS2002

Evaluation of an Aircraft Concept With Over-Wing, Hydrogen-Fueled Engines for Reduced Noise and Emissions

This report describes the analytical modeling and evaluation of an unconventional commercial transport aircraft concept designed to address aircraft noise and emission issues. A strut-braced wing configuration with overwing, ultra-high bypass ratio, hydrogen fueled turbofan engines is considered. Estimated noise and emission characteristics are compared to a conventional configuration designed for the same mission and significant benefits are identified. The design challenges and technology issues which would have to be addressed to make the concept a viable alternative to current aircraft designs are discussed. This concept is one of the "Quiet Green Transport" aircraft concepts studied as part of NASA's Revolutionary Aerospace Systems Concepts (RASC) Program. The RASC Program seeks to develop revolutionary concepts that address strategic objectives of the NASA Enterprises, such as reducing aircraft noise and emissions, and to identify enabling advanced technology requirements for the concepts.

Guynn, Mark D.↗
NASA NTRS2021

Retrieval Algorithm for the Column CO2 Mixing Ratio from Pulsed Multi-Wavelength Lidar Measurements

The retrieval algorithm for the column mixing ratio of CO2 from the measurements of a pulsed multi-wavelength integrated path differential absorption (IPDA) lidar is described. The lidar samples the shape of the 1572.33 nm CO2 absorption line at 15 or 30 wavelengths. The algorithm uses a least-squares fit between the CO2 line shape computed from a layered 10 atmosphere model to that sampled by the lidar. In addition to the column average CO2 dry air mole fraction (XCO2), several other parameters are also solved simultaneously from the fit. These include the Doppler shift in the received laser signal wavelengths, the product of the surface reflectivity and atmospheric transmission and a linear trend in the lidar receiver’s spectral response. The algorithm can also be used to solve for the average water vapor mixing ratio, which causes a secondary absorption in the wings of the CO2 absorption line under high humidity conditions. The least-squares fit is linearized about the 15 expected XCO2 value which allows the use of a standard linear least-squares fitting method and software tools. The standard deviation of the retrieved XCO2 is obtained from covariance matrix of the fit. An averaging kernel is defined similarly to that used for passive trace-gas sounding. Examples are presented of using the algorithm to retrieve XCO2 from the measurements from NASA Goddard’s airborne CO2 Sounder lidar made at a constant altitude and during spiral-down maneuvers.

Xiaoli Sun↗
NASA NTRSDate not supplied

Design Under Uncertainty with Design-Dependent Uncertain Variables

Uncertainty quantification (UQ) can provide a more robust understanding of a system, leading to better informed decisions earlier in the design process. The additional information that UQ provides can be leveraged during a design optimization process known as design under uncertainty that, when incorporated with multidisciplinary design and optimization, can become computationally infeasible due to the large number of responses required for meaningful results. Previous work addressed reducing the computational expense in design under uncertainty by incorporating analytic derivatives throughout polynomial chaos expansion. Although this allows design under uncertainty to be feasible for more systems, some multidisciplinary systems have design-dependent uncertain variables. This paper details an implementation of design dependent uncertain variables in a manner than preserves derivatives required for efficient gradient-based optimization throughout the process. Two analytic examples of design-dependent uncertain variables are given: the first transforms a uniform uncertain variable with one design variable and the second transforms a normal uncertain variable with two design variables. The polynomial chaos expansion (PCE) results are comparable to both the Monte Carlo (MC) results and the analytic results for the two examples. A case study that maximizes the lift-to-drag ratio with a design-dependence between the wing leading edge sweep angle and uncertain parameter percentage of laminar flow is compared to a MC and alternative optimization formulations. This paper demonstrates that design-dependent uncertain variables are valid and hold throughout PCE.

Joanna N Schmidt↗
NASA NTRSDate not supplied

Design Under Uncertainty with Design-Dependent Uncertain Variables

Uncertainty quantification (UQ) can provide a more robust understanding of a system, leading to better informed decisions earlier in the design process. The additional information that UQ provides can be leveraged during a design optimization process known as design under uncertainty that, when incorporated with multidisciplinary design and optimization, can become computationally infeasible due to the large number of responses required for meaningful results. Previous work addressed reducing the computational expense in design under uncertainty by incorporating analytic derivatives throughout polynomial chaos expansion. Although this allows design under uncertainty to be feasible for more systems, some multidisciplinary systems have design-dependent uncertain variables. This paper details an implementation of design dependent uncertain variables in a manner than preserves derivatives required for efficient gradient-based optimization throughout the process. Two analytic examples of design-dependent uncertain variables are given: the first transforms a uniform uncertain variable with one design variable and the second transforms a normal uncertain variable with two design variables. The polynomial chaos expansion (PCE) results are comparable to both the Monte Carlo (MC) results and the analytic results for the two examples. A case study that maximizes the lift-to-drag ratio with a design-dependence between the wing leading edge sweep angle and uncertain parameter percentage of laminar flow is compared to a MC and alternative optimization formulations. This paper demonstrates that design-dependent uncertain variables are valid and hold throughout PCE.

robust design↗
NASA NTRS1973

Heating and flow-field studies on a straight-wing hypersonic reentry vehicle at angles of attack from 20 to 80 deg with simulation of real-gas trends

Data are presented from a series of phase-change heat transfer and flow visualization tests at Mach 7.4, 8, and 10.3 in air, Mach 19.5 in nitrogen, Mach 20.3 in helium, and Mach 6 in tetrafluoromethane (CF4) on the windward surface of a straight wing hypersonic reentry configuration for angles of attack from 20 deg to 80 deg. The results indicate that: (1) for hypersonic stream Mach numbers, the flow field over the straight-wing configuration is essentially independent of Mach number, (2) transition Reynolds number decreases with increasing angle of attack, (3) at some critical angle of attack, the wing-shock standoff distance is greatly increased and the stagnation line moves downstream from the wing leading edge, (4) value of the critical angle of attack is very sensitive to the flow shock density ratio or effective gamma, and (5) at angles of attack above the critical value for all gases, the nondimensional level of heat transfer to the wing is higher for the higher shock density ratio flows.

Hunt, J. L.↗
NASA NTRS1960

Aerodynamic Performance and Static Stability at Mach Number 3.3 of an Aircraft Configuration Employing Three Triangular Wing Panels and a Body Equal Length

An aircraft configuration, previously conceived as a means to achieve favorable aerodynamic stability characteristics., high lift-drag ratio, and low heating rates at high supersonic speeds., was modified in an attempt to increase further the lift-drag ratio without adversely affecting the other desirable characteristics. The original configuration consisted of three identical triangular wing panels symmetrically disposed about an ogive-cylinder body equal in length to the root chord of the panels. This configuration was modified by altering the angular disposition of the wing panels, by reducing the area of the panel forming the vertical fin, and by reshaping the body to produce interference lift. Six-component force and moment tests of the modified configuration at combined angles of attack and sideslip were made at a Mach number of 3.3 and a Reynolds number of 5.46 million. A maximum lift-drag ratio of 6.65 (excluding base drag) was measured at a lift coefficient of 0.100 and an angle of attack of 3.60. The lift-drag ratio remained greater than 3 up to lift coefficient of 0.35. Performance estimates, which predicted a maximum lift-drag ratio for the modified configuration 27 percent greater than that of the original configuration, agreed well with experiment. The modified configuration exhibited favorable static stability characteristics within the test range. Longitudinal and directional centers of pressure were slightly aft of the respective centroids of projected plan-form and side area.

James, Carlton S.↗
NASA NTRS1941

Characteristics of NACA 4400R Series Rectangular and Tapered Airfoils, Including the Effect of Split Flaps

At the request of the Bureau of Aeronautics, Navy Department, tests were made in the variable-density wind tunnel of a tapered wing of 3-10-18 plan form and based on the NACA 4400R series sections. The wing was also tested with 0.2 chord spit flaps, deflected 60 deg span ratios of 0.3, 0.5, 0.7 and 1.0 respectively. In order to get data from which to calculate the characteristics of the flapped wing, the investigation was extended to include tests of the four rectangular airfoils of the NACA 4400R series (4409R, 4412R, 4415R, and 4418R) with full-span 0.2 chord, trailing edge split flaps deflected 60 deg.

Greenberg, Harry↗
NASA NTRS2013

High Bypass Ratio Jet Noise Reduction and Installation Effects Including Shielding Effectiveness

An experimental investigation was performed to study the propulsion airframe aeroacoustic installation effects of a separate flow jet nozzle with a Hybrid Wing Body aircraft configuration where the engine is installed above the wing. Prior understanding of the jet noise shielding effectiveness was extended to a bypass ratio ten application as a function of nozzle configuration, chevron type, axial spacing, and installation effects from additional airframe components. Chevron types included fan chevrons that are uniform circumferentially around the fan nozzle and T-fan type chevrons that are asymmetrical circumferentially. In isolated testing without a pylon, uniform chevrons compared to T-fan chevrons showed slightly more low frequency reduction offset by more high frequency increase. Phased array localization shows that at this bypass ratio chevrons still move peak jet noise source locations upstream but not to nearly the extent, as a function of frequency, as for lower bypass ratio jets. For baseline nozzles without chevrons, the basic pylon effect has been greatly reduced compared to that seen for lower bypass ratio jets. Compared to Tfan chevrons without a pylon, the combination with a standard pylon results in more high frequency noise increase and an overall higher noise level. Shielded by an airframe surface 2.17 fan diameters from nozzle to airframe trailing edge, the T-fan chevron nozzle can produce reductions in jet noise of as much as 8 dB at high frequencies and upstream angles. Noise reduction from shielding decreases with decreasing frequency and with increasing angle from the jet inlet. Beyond an angle of 130 degrees there is almost no noise reduction from shielding. Increasing chevron immersion more than what is already an aggressive design is not advantageous for noise reduction. The addition of airframe control surfaces, including vertical stabilizers and elevon deflection, showed only a small overall impact. Based on the test results, the best overall nozzle configuration design was selected for application to the N2A Hybrid Wing Body concept that will be the subject of the NASA Langley 14 by 22 Foot Subsonic Tunnel high fidelity aeroacoustic characterization experiment. The best overall nozzle selected includes T-fan type chevrons, uniform chevrons on the core nozzle, and no additional pylon of the type that created a strong acoustic effect at lower bypass ratios. The T-fan chevrons are oriented azimuthally away from the ground observer locations. This best overall nozzle compared to the baseline nozzle was assessed, at equal thrust, to produce sufficient installed noise reduction of the jet noise component to enable the N2A HWB to meet NASA s noise goal of 42 dB cumulative below Stage 4.

Thomas, Russell H.↗
NASA NTRS1980

Performance improvement of delta wings at subsonic speeds due to vortex flaps

Subsonic wind tunnel tests were conducted to determine performance improvements possible from the use of leading edge vortex flaps (LEVF) on delta wings. Various flap sizes and deflection angles were examined and lift-to-drag ratio improvements of up to 40% were found at moderate angles of attack on 60 deg and 75 deg swept wings. The LEVF is concluded to be effective in moving the wing's leading edge vortex onto the flaps, tilting the vortex-induced force vector forward to produce a thrust or reduce the wing's drag while maintaining attached flow and lift on the wing's upper surface.

Marchman, J. F., III↗
NASA NTRS1960

Performance, Stability, and Control Investigation at Mach Numbers from 0.4 to 0.9 of a Model of the "Swallow" with Outer Wing Panels Swept 25 degree with and without Power Simulation

An investigation of the performance, stability, and control characteristics of a variable-sweep arrow-wing model (the "Swallow") with the outer wing panels swept 25 deg has been conducted in the Langley 16-foot transonic tunnel. The wing was uncambered and untwisted and had RAE 102 airfoil sections with a thickness-to-chord ratio of 0.14 normal to the leading edge. Four outboard engines located above and below the wing provided propulsive thrust, and, by deflecting in the pitch direction and rotating in the lateral plane, also produced control forces. A pair of swept lateral fins and a single vertical fin were mounted on each engine nacelle to provide aerodynamic stability and control. Jets-off data were obtained with flow-through nacelles, stimulating the effects of inlet flow; jet thrust and hot-jet interference effects were obtained with faired-nose nacelles housing hydrogen peroxide gas generators. Six-component force and moment data were obtained through a Mach number range of 0.40 to 0.90 at angles of attack and angles of sideslip from 0 deg to 15 deg. Longitudinal, directional, and lateral control were obtained by deflecting the nacelle-fin combinations as elevators, rudders, and ailerons at several fixed angles for each control.

Runckel, Jack F.↗
NASA NTRS1955

Effect of Ground Interference on the Aerodynamic and Flow Characteristics of a 42 Degree Sweptback Wing at Reynolds Numbers up to 6.8 x 10(6)

Report presents the results of an investigation of the effects of ground interference on the aerodynamic characteristics of a 42 degree sweptback wing at distances 0.68 and 0.92 of the mean aerodynamic chord from the simulated ground to the 0.25-chord point of the mean aerodynamic chord. Survey data behind the wing, both with and without the simulated ground, are presented in the form of contour charts of downwash, sidewash, and dynamic-pressure ratio at longitudinal stations of 2.0 and 2.8 mean aerodynamic chords behind the wing.

Furlong, G Chester↗
NASA NTRS1976

Comparison of supercritical and conventional wing flutter characteristics

A wind-tunnel study was undertaken to directly compare the measured flutter boundaries of two dynamically similar aeroelastic models which had the same planform, maximum thickness-to-chord ratio, and as nearly identical stiffness and mass distributions as possible, with one wing having a supercritical airfoil and the other a conventional airfoil. The considerations and problems associated with flutter testing supercritical wing models at or near design lift coefficients are discussed, and the measured transonic boundaries of the two wings are compared with boundaries calculated with a subsonic lifting surface theory.

Farmer, M. G.↗
NASA NTRS1934

The effects of full-span and partial-span split flaps on the aerodynamic characteristics of a tapered wing

The investigation was made to determine the effects of full-span and of partial-span split flaps on the aerodynamic characteristics of a tapered wing. Aerodynamic force tests were made in the N.A.C.A. 7 by 10 foot wind tunnel on a highly tapered Clark Y wing equipped with various split flaps. Two sizes of tapered-chord flaps were tested as full-span flaps, and a narrow tapered-chord flap was tested as a partial-span flap by cutting off portions first from the tip and then from the center. The investigation showed that with full-span split flaps the lift and drag characteristics of the tapered wing up to the stall are similar to those of a rectangular wing with flaps of comparable size, but that the stall of the tapered wing with full-span flaps occurs at progressively lower angles of attack with increasing flap deflection up to that for maximum lift. For partial-span tapered split flaps on a tapered wing it was found that the maximum lift is greater, and the lift-drag ratio at maximum lift is less, when the partial-span flap is located at the center of the wing than when it is located at the tip portion.

Wenzinger, Carl J↗