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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 127 records · Page 7
All sourcesNASA NTRSDOE OSTI
NASA NTRS1987

High angle of attack hypersonic aerodynamics

A new aerodynamics force model is presented which is based on modified Newtonian theory and empirical correlations. The algebraic model was developed for complete vehicles from take off to orbital speeds and for large angles of attack. Predictions are compared to results for a wind tunnel model at a Mach number of 20, and the full scale Shuttle Orbiter for Mach numbers from 0.25 to 20 for angles of attack from 0 to 50 deg. The maximum shuttle orbiter lift/drag at Mach 10 and 20 is 1.85 at 20-deg angle-of-attack. Aerodynamic force predictions are made for a transatmospheric vehicle, which is a derivative of the Shuttle Orbiter, for Mach numbers from 4 to 27 at angles of attack from 5 to 40 deg. Predicted aerodynamic force data indicate that lift/drag ratios of 5.2 at Mach number 10 and 3.6 at Mach number 26 are obtainable. Changes in force coefficients with changes in: nose angle, sweep angle, and (volume exp 2/3)/planform area are quantified for Mach numbers of 10 and 26. Lift/drag ratios increase with decreasing nose angle and (volume exp 2/3)/planform area and increasing wing sweep angle. Lift/drag ratios are independent of these variables for angles of attack in excess of 20 deg at Mach 10 and 30 deg at Mach 26.

Harloff, Gary J.↗
NASA NTRS1972

Status of Langley studies of aerodynamics and interference effects of tandem launch vehicles

Some of the aerodynamic and interference effects measured from wind-tunnel tests of three fully reusable space shuttle configurations at transonic and low-supersonic speeds are presented, and the status is given of an assessment of the analytical approach to date. The three configurations tested were a low-fineness-ratio booster combined in tandem with a delta-wing orbiter, designated the "low-fineness-ratio configuration,' and a phase B shuttle configuration consisting of the delta-wing orbiter and high-fineness-ratio booster, designated the "stage-arrangement configuration,' arranged in tandem and in parallel. The force measurements obtained at a Mach number of 1.5 for the low-fineness-ratio configuration were representative of the transonic and low-supersonic speed ranges where the interference effects are predominant and were used for comparison with values calculated at the same Mach number. This comparison is the basis for a preliminary assessment of the initial application of the analytical approach.

Scallion, W. I.↗
NASA NTRS1978

Static tests of the J97 powered, external augmenter V/STOL wind tunnel model

Results of the static testing (zero forward speed) of the J97-powered, external augmentor, large scale, V/STOL model are discussed. With a ground clearance of 7.5 feet, believed to have put the model essentially out of ground effect, a gross thrust augmentation ratio of 1.60 at nozzle pressure ratio (NPR) = 3.0 was measured for the fuselage augmentor. A similar figure was apparent for the wing augmentor. An overall ratio of model thrust to bare engine thrust of 1.52 was determined at NPR = 3.0. The structural integrity of the model was well demonstrated and duct pressure losses were small.

Garland, D. B.↗
NASA NTRS1989

Some aerodynamic characteristics of the scissor wing configuration

A scissor wing configuration, consisting of four adjustable wing surfaces, is compared with a comparable fixed wing baseline configuration. Wave drag, induced drag, viscous drag, thrust required, and gust loading are calculated for both configurations. The scissor wing is shown to have lower zero lift wave drag and higher total lift to drag ratios than the baseline. It is demonstrated that the scissor configurations' sweep can be programmed to keep the static margin fixed. Thrust required for both the fixed static margin case and a constant sweep angle case are presented with the scissor configuration requiring lower thrust levels. The gust loading ratio of the scissor wing to the baseline is also shown to be significantly less than 1.0 for sweep angles greater than 20 degrees.

Selberg, Bruce P.↗
NASA NTRS1994

Thin tailored composite wing for civil tiltrotor

The tiltrotor aircraft is a flight vehicle which combines the efficient low speed (i.e., take-off, landing, and hover) characteristics of a helicopter with the efficient cruise speed of a turboprop airplane. A well-known example of such vehicle is the Bell-Boeing V-22 Osprey. The high cruise speed and range constraints placed on the civil tiltrotor require a relatively thin wing to increase the drag-divergence Mach number which translates into lower compressibility drag. It is required to reduce the wing maximum thickness-to-chord ratio t/c from 23% (i.e., V-22 wing) to 18%. While a reduction in wing thickness results in improved aerodynamic efficiency, it has an adverse effect on the wing structure and it tends to reduce structural stiffness. If ignored, the reduction in wing stiffness leads to susceptibility to aeroelastic and dynamic instabilities which may consequently cause a catastrophic failure. By taking advantage of the directional stiffness characteristics of composite materials the wing structure may be tailored to have the necessary stiffness, at a lower thickness, while keeping the weight low. The goal of this study is to design a wing structure for minimum weight subject to structural, dynamic and aeroelastic constraints. The structural constraints are in terms of strength and buckling allowables. The dynamic constraints are in terms of wing natural frequencies in vertical and horizontal bending and torsion. The aeroelastic constraints are in terms of frequency placement of the wing structure relative to those of the rotor system. The wing-rotor-pylon aeroelastic and dynamic interactions are limited in this design study by holding the cruise speed, rotor-pylon system, and wing geometric attributes fixed. To assure that the wing-rotor stability margins are maintained a more rigorous analysis based on a detailed model of the rotor system will need to ensue following the design study. The skin-stringer-rib type architecture is used for the wing-box structure. The design variables include upper and lower skin ply thicknesses and orientation angles, spar and rib web thicknesses and cap areas, and stringer cross-sectional areas. These design variables will allow the maximum tailoring of the structure to meet the design requirements most efficiently. Initial dynamic analysis has been conducted using MSC/NASTRAN to determine the baseline wing's frequencies and mode shapes. For the design study we intend to use the finite-element based code called WIDOWAC (Wing Design Optimization With Aeroeastic Constraints) that was developed at NASA Langley in early 1970's for airplane wing structural analysis and preliminary design. Currently, the focus is on modification and validation of this code which will be used for the civil tiltrotor design efforts.

Rais-Rohani, Masoud↗
NASA NTRS2006

A Note about Self-Induced Velocity Generated by a Lifting-Line Wing or Rotor Blade

This report presents an elementary analysis of the induced velocity created by a field of vortices that reside in the wake of a rotor blade. Progress achieved by other researchers in the last 70 years is briefly reviewed. The present work is presented in four stages of complexity that carry a lifting-line representation of a fixed wing into a single-blade rotor. The analysis leads to the conclusion that the lifting rotor's spiraling vortex wake structure has very high induced power when compared to the ideal wing. For an advanced ratio of one-half, induced power is on the order of 10 times that of the wing when the comparison is made at wingspan equal to rotor diameter and wing and rotor having equal lift.

Harris, Franklin D.↗
NASA NTRS1961

An Investigation of the Influence of Body Size and Indentation Asymmetry of the Effectiveness of Body Indentation in Combination with a Cambered Wing

An investigation has been made of a 450 sweptback cambered wing in combination with an unindented body and a body symmetrically indented with respect to its axes designed for a Mach number of 1.2. The ratio of body frontal area to wing planform area was 0.08 for these wing-body combinations. In order to determine the influence of body size on the effectiveness of indentation, the test data have been compared with previously obtained data for similar configurations having a ratio of body frontal area to wing planform area of 0.04. Also, in order to investigate the relative effectiveness of indentation asymmetry, a specially indented body designed to account for the wing camber and also designed for a Mach number of 1.2 has been included in these tests. The investigation was conducted in the Langley 8-Foot Tunnels Branch at Mach numbers from 0.80 to 1.43 and a Reynolds number of approximately 1.85 x 10(exp 6), based on a mean aerodynamic chord length of 5.955 inches. The data indicate that the configurations with larger ratio of body frontal area to wing planform area had smaller reductions in zero-lift wave drag associated with body indentation than the configurations with smaller ratio of body frontal area to wing planform area. The 0.08-area-ratio configurations also had correspondingly smaller increases in the values of maximum lift-drag ratio than the 0.04-area-ratio configurations. The consideration of wing camber in the body indentation design resulted in a 35.5-percent reduction in zero-lift wave drag, compared with a 21.5-percent reduction associated with the symmetrical indentation, but had a negligible effect on the values of maximum lift-drag ratio.

Patterson, James C., Jr.↗
NASA NTRS1974

Comparisons of two-dimensional shock-expansion theory with experimental aerodynamic data for delta-planform wings at high supersonic speeds

An investigation has been conducted to explore the potential for optimizing airfoil shape at high supersonic speeds by utilizing the two-dimensional shock-expansion method. Theoretical and experimental force and moment coefficients are compared for four delta-planform semispan wings having a leading-edge sweep angle of 65 deg and incorporating modified diamond airfoils with a thickness-chord ratio of 0.06. The wings differ only in airfoil maximum-thickness position and camber. The experimental data are obtained at Mach numbers of 3.95 and 4.63 and at a Reynolds number of 9.84 million per meter. A relatively simple method is developed for predicting, in terms of lift-drag ratio, the optimum modified diamond airfoil at high supersonic and hypersonic speeds.

Jernell, L. S.↗
NASA NTRS1972

Analysis of the effect of engine characteristics on the external aerodynamics of STOL wing propulsion systems

The effects of engine presssure ratio, engine size, and engine location on the pressure distribution, lift coefficient, and flow field of a STOL wing propulsion system are presented. The flow variables of the engines are included in the two-dimensional potential flow analysis by considering the effects of mass flow coefficient at the engine inlet and thrust coefficient at the engine exit. A functional relation between these coefficients and engine pressure ratio is given. The results of this study indicate that the effect of engine pressure ratio on the external aerodynamics is a function of engine location. For engines located on the bottom of the wing, the highest pressure ratio engine resulted in the highest lift coefficient. For engines located on the top of the wing, the lowest pressure ratio engine resulted in the highest lift coefficient.

Albers, J. A.↗
NASA NTRS1974

Geometry considerations for jet noise shielding with CTOL engine-over-the-wing concept

Jet noise shielding benefits for CTOL engine-over-the-wing installations were obtained with various model-scale circular nozzles and wing chord geometries. Chord-to-nozzle diameter ratios were varied from 3 to 20, while ratios of nozzle height above the wing to the diameter were varied from near zero to 3. Spectral noise data were obtained with jet velocities from 640 to 1110 ft/sec. Characteristics of low frequency noise sources are discussed. Jet-noise shielding is correlated in terms of acoustic and geometric parameters. Implications of extending the model-scale data to full-scale are discussed.

Vonglahn, U.↗
NASA NTRS1974

Geometry considerations for jet noise shielding with CTOL engine-over-the-wing concept

Jet noise shielding benefits for CTOL engine-over-the-wing installations were obtained with various model-scale circular nozzles and wing chord geometries. Chord-to-nozzle diameter ratios were varied from 3 to 20, while ratios of nozzle height above the wing to the diameter were varied from near zero to 3. Spectral noise data were obtained with jet velocities from 640 to 1110 ft/sec. Characteristics of low frequency noise sources are discussed. Jet-noise shielding is correlated in terms of acoustic and geometric parameters. Implications of extending the model-scale data to full-scale are discussed.

Von Glahn, U.↗
NASA NTRS1977

Wing design by numerical optimization

A study was conducted to assess the feasibility of performing computerized wing design by numerical optimization. The design program combined a full potential, inviscid aerodynamics code with a conjugate gradient optimization algorithm. Three design problems were selected to demonstrate the design technique. The first involved modifying the upper surface of the inboard 50% of a swept wing to reduce the shock drag subject to a constraint on wing volume. The second involved modifying the entire upper surface of the same swept wing (except the tip section) to increase the lift-drag ratio subject to constraints on wing volume and lift coefficient. The final problem involved modifying the inboard 50% of a low-speed wing to achieve good stall progression. Results from the three cases indicate that the technique is sufficiently accurate to permit substantial improvement in the design objectives.

Hicks, R. M.↗
NASA NTRS1933

Wind-tunnel tests of a Clark Y wing with a narrow auxiliary airfoil in different positions

Aerodynamic force tests were made on a combination of a Clark Y wing and a narrow auxiliary airfoil to find the best location of the auxiliary airfoil with respect to the main wing. The auxiliary was a highly cambered airfoil of medium thickness having a chord 14.5 per cent that of the main wing. It was tested in 141 different positions ahead of, above, and behind the nose portion of the main wing, the range of the test points being extended until the best aerodynamic conditions were covered. A range of positions was found in which the combination of main wing and auxiliary gave substantially greater aerodynamic efficiency and higher maximum lift coefficients (based on total area) than the main Clark Y wing alone. In the optimum position tested, considering both the maximum lift and the speed-range ratio, the combination of main wing and auxiliary gave an increase in the maximum lift coefficient of 32 per cent together with an increase in the ratio of 21 per cent of the respective values for the main Clark Y wing alone.

Weick, Fred E↗
DOE OSTI2022

Tidal Turbine Test, Downeast Turbines, July 12, 2021

Downeast Turbines tested a tidal turbine prototype with novel rotor/channel system and lateral effluent discharge apparatus (LEDA), during five days of testing in the flume at Alden Lab. Three days of testing (July 12-14, 2021) focused on turbine power metrics of torque and rpm, which were low, and then two days of follow-up testing (July 27-28, 2021) focused on LEDA performance metrics of pressure differential and rates of volumetric flow, with encouraging results. Next step is to to characterize, and even optimize, configurations of the LEDA, using 3D-CFD as a helpful tool, to refine its shape and explore its limits of performance as a means of effluent discharge that augments performance of an instream turbine. An improved configuration of the LEDA will be re-combined with the rotor/channel system of the turbine prototype, and ultimately, the rotor will be re-sized (enlarged), to better match the LEDA's performance capabilities in drawing through a rate of volumetric flow. This submission is Downeast Turbines' Post Access Report for the test event. It includes the files described here (next below), and several reference links. "Downeast TEAMER-Post-Access-Report...docx" is a document file containing the report. "Appendices A, B, and C" are included in this file, and so are "Figures #1-7." "Appendix D - Test Data Workbooks.zip" is an archive file containing all post access data (raw data tables, calculating tables, and graphs), presented in fourteen Excel workbooks as described in the report. "Appendix E - Post Access Figures.zip" is an archive file containing "Figures #8-52," of the report.

16 TIDAL AND WAVE POWER↗
NASA NTRS2014

Noise and Fuel Burn Reduction Potential of an Innovative Subsonic Transport Configuration

A study is presented for the noise and fuel burn reduction potential of an innovative double deck concept aircraft with two three-shaft direct-drive turbofan engines. The engines are mounted from the fuselage so that the engine inlet is over the main wing. It is shown that such an aircraft can achieve a cumulative Effective Perceived Noise Level (EPNL) about 28 dB below the current aircraft noise regulations of Stage 4. The combination of high bypass ratio engines and advanced wing design with laminar flow control technologies provide fuel burn reduction and low noise levels simultaneously. For example, the fuselage mounted engine position provides more than 4 EPNLdB of noise reduction by shielding the inlet radiated noise. To identify the potential effect of noise reduction technologies on this concept, parametric studies are presented to reveal the system level benefits of various emerging noise reduction concepts, for both engine and airframe noise reduction. These concepts are discussed both individually to show their respective incremental noise reduction potential and collectively to assess their aggregate effects on the total noise. Through these concepts approximately about 8 dB of additional noise reduction is possible, bringing the cumulative noise level of this aircraft to 36 EPNLdB below Stage 4, if the entire suite of noise reduction technologies would mature to practical application. In a final step, an estimate is made for this same aircraft concept but with higher bypass ratio, geared, turbofan engines. With this geared turbofan propulsion system, the noise is estimated to reach as low as 40-42 dB below Stage 4 with a fuel burn reduction of 43-47% below the 2005 best-in-class aircraft baseline. While just short of the NASA N+2 goals of 42 dB and 50% fuel burn reduction, for a 2025 in service timeframe, this assessment shows that this innovative concept warrants refined study. Furthermore, this design appears to be a viable potential future passenger aircraft, not only in meeting the regulatory requirements, but also in competing with aircraft of different advanced designs within this N+2 timeframe and goal framework.

Guo, Yueping↗
NASA NTRS2010

Propulsion Airframe Aeroacoustic Integration Effects for a Hybrid Wing Body Aircraft Configuration

An extensive experimental investigation was performed to study the propulsion airframe aeroacoustic effects of a high bypass ratio engine for a hybrid wing body aircraft configuration where the engine is installed above the wing. The objective was to provide an understanding of the jet noise shielding effectiveness as a function of engine gas condition and location as well as nozzle configuration. A 4.7% scale nozzle of a bypass ratio seven engine was run at characteristic cycle points under static and forward flight conditions. The effect of the pylon and its orientation on jet noise was also studied as a function of bypass ratio and cycle condition. The addition of a pylon yielded significant spectral changes lowering jet noise by up to 4dB at high polar angles and increasing it by 2 to 3dB at forward angles. In order to assess jet noise shielding, a planform representation of the airframe model, also at 4.7% scale was traversed relative to the jet nozzle from downstream to several diameters upstream of the wing trailing edge. Installations at two fan diameters upstream of the wing trailing edge provided only limited shielding in the forward arc at high frequencies for both the axisymmetric and a conventional round nozzle with pylon. This was consistent with phased array measurements suggesting that the high frequency sources are predominantly located near the nozzle exit and, consequently, are amenable to shielding. The mid to low frequencies sources were observed further downstream and shielding was insignificant. Chevrons were designed and used to impact the distribution of sources with the more aggressive design showing a significant upstream migration of the sources in the mid frequency range. Furthermore, the chevrons reduced the low frequency source levels and the typical high frequency increase due to the application of chevron nozzles was successfully shielded. The pylon was further modified with a technology that injects air through the shelf of the pylon which was effective in reducing low frequency noise and moving jet noise sources closer to the nozzle exit. In general, shielding effectiveness varied as a function of cycle condition with the cutback condition producing higher shielding compared to sideline power. The configuration with a more strongly immersed chevron and a pylon oriented opposite to the microphones produced the largest reduction in jet noise. In addition to the jet noise source, the shielding of a broadband point noise source was documented with up to 20 dB of noise reduction at directivity angles directly under the shielding surface.

Czech, Michael J.↗