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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 361 records · Page 20
All sourcesNASA NTRSDOE OSTI
NASA NTRS1939

A comparison of several tapered wings designed to avoid tip stalling

Optimum proportions of tapered wings were investigated by a method that involved a comparison of wings designed to be aerodynamically equal. The conditions of aerodynamic equality were equality in stalling speed, in induced drag at a low speed, and in the total drag at cruising speed. After the wings were adjusted to aerodynamic equivalence, the weights of the wings were calculated as a convenient method of indicating the optimum wing. The aerodynamic characteristics were calculated from wing theory and test data for the airfoil sections. Various combinations of washout, camber increase in the airfoil sections from the center to the tips, and sharp leading edges at the center were used to bring about the desired equivalence of maximum lift and center-stalling characteristics. In the calculation of the weights of the wings, a simple type of spar structure was assumed that permitted an integration across the span to determine the web and the flange weights. The covering and the remaining weight were taken in proportion to the wing area. The total weights showed the wings with camber and washout to have the lowest weights and indicated the minimum for wings with a taper ratio between 1/2 and 1/3.

Anderson, Raymond F↗
NASA NTRS1957

Systematic two-dimensional cascade tests of NACA 65-series compressor blades at low speeds

A two-dimensional low-speed porous-wall cascade tunnel investigation has been conducted to establish the performance of the NACA 65-series compressor blade sections over the useful range of inlet angle, solidity, and section camber. Design points for optimum high-speed operation are presented. The loading limitation is determined for some conditions. Trends of section operating range with increasing section camber are determined for the four inlet angles tested.

COMPRESSORS - AXIAL-FLOW↗
NASA NTRS1925

Aerodynamic Characteristics of Airfoils at High Speeds

This report deals with an experimental investigation of the aerodynamical characteristics of airfoils at high speeds. Lift, drag, and center of pressure measurements were made on six airfoils of the type used by the air service in propeller design, at speeds ranging from 550 to 1,000 feet per second. The results show a definite limit to the speed at which airfoils may efficiently be used to produce lift, the lift coefficient decreasing and the drag coefficient increasing as the speed approaches the speed of sound. The change in lift coefficient is large for thick airfoil sections (camber ratio 0.14 to 0.20) and for high angles of attack. The change is not marked for thin sections (camber ratio 0.10) at low angles of attack, for the speed range employed. At high speeds the center of pressure moves back toward the trailing edge of the airfoil as the speed increases. The results indicate that the use of tip speeds approaching the speed of sound for propellers of customary design involves a serious loss in efficiency.

Briggs, L J↗
NASA NTRS1934

Wind-tunnel tests on combinations of a wing with fixed auxiliary airfoils having various chords and profiles

This report presents the results of wind tunnel tests on various auxiliary airfoils having three different airfoil sections and several different chord lengths in combination with a Clark y model wing in a sufficient number of relative positions to determine the optimum with regard to certain criterions of aerodynamic performance. The airfoil sections included a symmetrical profile, one of medium camber, and a highly cambered one. The chord sizes of the auxiliary airfoils ranged from 7.5 to 25 percent of the chord of the main wing, and the span was equal to that of the main wing.

Weick, Fred E↗
NASA NTRS1951

Systematic Two-Dimensional Cascade Tests of NACA 65-Series Compressor Blades at Low Speeds

The performance of NACA 65-series compressor blade section in cascade has been investigated systematically in a low-speed cascade tunnel. Porous test-section side walls and for high-pressure-rise conditions, porous flexible end walls were employed to establish conditions closely simulating two-dimensional flow. Blade sections of design lift coefficients from 0 to 2.7 were tested over the usable angle-of-attack range for various combinations of inlet-flow angle. A sufficient number of combinations were tested to permit interpolation and extrapolation of the data to all conditions within the usual range of application. The results of this investigation indicate a continuous variation of blade-section performance as the major cascade parameters, blade camber, inlet angle, and solidity were varied over the test range. Summary curves of the results have been prepared to enable compressor designers to select the proper blade camber and angle of attack when the compressor velocity diagram and desired solidity have been determined.

Herrig, L. Joseph↗
NASA NTRS1945

Tests of Four Full-scale Propellers to Determine the Effect of Trailing-edge Extensions on Propeller Aerodynamic Characteristics

Propellers with trailing-edge extensions were studied to determine aerodynamic characteristics. Trailing-edge extension increased power absorbed by propeller with little loss in efficiency. Power coefficient for maximum efficiency was greater for 20% camber type extension than for 20% straight type extension over range of advance ratio of 1.0 to 2.5 although camber type was less efficient. Efficiency was about the same for cruising and high-speed at a high power coefficient for propeller with extension.

PROPELLERS - TRAILING EDGE, EXTENDED↗
NASA NTRS1956

Low-speed cascade investigation of loaded leading-edge compressor blades

Six percent thick NACA 63-series compressor-blade sections having a loaded leading-edge A4K6 mean line have been investigated systematically in a two-dimensional porous-wall cascade over a range of Reynolds numbers from 160,000 to 385,000. Blades cambered to have isolated-airfoil lift coefficients of 0.6, 1.2, 1.8, and 2.4 were tested over the usable angle-of-attack range at inlet-air angles of 30 degrees, 45 degrees, and 60 degrees and solidities of 1.0 and 1.5. A comparison with data of NACA RM L51G31, shows that the angle-of-attack operating range is 2 degrees to 4 degrees less than the range for the uniformly loaded section; however, the wake losses near design angle of attack are slightly lower than those for the uniformly loaded section. Except for highly cambered blades at high inlet angles, the 63-(C s oA4K6)06 compressor-blade sections are capable of more efficient operation for moderate-speed subsonic compressors at design angle of attack than are the 65-(C s oa10)10 or the 65-(c s oA2I8b)10 compressor-blade sections. In contrast to the other sections, the loaded leading-edge sections are capable of operating efficiently at the lower Reynolds numbers.

Emery, James C↗
NASA NTRS1994

Experimental and theoretical study of aerodynamic characteristics of some lifting bodies at angles of attack from -10 degrees to 53 degrees at Mach numbers from 2.30 to 4.62

Lifting bodies are of interest for possible use as space transportation vehicles because they have the volume required for significant payloads and the aerodynamic capability to negotiate the transition from high angles of attack to lower angles of attack (for cruise flight) and thus safely reenter the atmosphere and perform conventional horizontal landings. Results are presented for an experimental and theoretical study of the aerodynamic characteristics at supersonic speeds for a series of lifting bodies with 75 deg delta planforms, rounded noses, and various upper and lower surface cambers. The camber shapes varied in thickness and in maximum thickness location, and hence in body volume. The experimental results were obtained in the Langley Unitary Plan Wind Tunnel for both the longitudinal and the lateral aerodynamic characteristics. Selected experimental results are compared with calculated results obtained through the use of the Hypersonic Arbitrary-Body Aerodynamic Computer Program.

Spearman, M. Leroy↗
NASA NTRS1994

An initial investigation into methods of computing transonic aerodynamic sensitivity coefficients

The primary accomplishments of the project are as follows: (1) Using the transonic small perturbation equation as a flowfield model, the project demonstrated that the quasi-analytical method could be used to obtain aerodynamic sensitivity coefficients for airfoils at subsonic, transonic, and supersonic conditions for design variables such as Mach number, airfoil thickness, maximum camber, angle of attack, and location of maximum camber. It was established that the quasi-analytical approach was an accurate method for obtaining aerodynamic sensitivity derivatives for airfoils at transonic conditions and usually more efficient than the finite difference approach. (2) The usage of symbolic manipulation software to determine the appropriate expressions and computer coding associated with the quasi-analytical method for sensitivity derivatives was investigated. Using the three dimensional fully conservative full potential flowfield model, it was determined that symbolic manipulation along with a chain rule approach was extremely useful in developing a combined flowfield and quasi-analytical sensitivity derivative code capable of considering a large number of realistic design variables. (3) Using the three dimensional fully conservative full potential flowfield model, the quasi-analytical method was applied to swept wings (i.e. three dimensional) at transonic flow conditions. (4) The incremental iterative technique has been applied to the three dimensional transonic nonlinear small perturbation flowfield formulation, an equivalent plate deflection model, and the associated aerodynamic and structural discipline sensitivity equations; and coupled aeroelastic results for an aspect ratio three wing in transonic flow have been obtained.

Carlson, Leland A.↗
NASA NTRS1995

Wing-section optimization for supersonic viscous flow

To improve the shape of a supersonic wing, an automated method that also includes higher fidelity to the flow physics is desirable. With this impetus, an aerodynamic optimization methodology incorporating thin-layer Navier-Stokes equations and sensitivity analysis had been previously developed. Prior to embarking upon the wind design task, the present investigation concentrated on testing the feasibility of the methodology, and the identification of adequate problem formulations, by defining two-dimensional, cost-effective test cases. Starting with two distinctly different initial airfoils, two independent shape optimizations resulted in shapes with similar features: slightly cambered, parabolic profiles with sharp leading- and trailing-edges. Secondly, the normal section to the subsonic portion of the leading edge, which had a high normal angle-of-attack, was considered. The optimization resulted in a shape with twist and camber which eliminated the adverse pressure gradient, hence, exploiting the leading-edge thrust. The wing section shapes obtained in all the test cases had the features predicted by previous studies. Therefore, it was concluded that the flowfield analyses and sensitivity coefficients were computed and fed to the present gradient-based optimizer correctly. Also, as a result of the present two-dimensional study, suggestions were made for the problem formulations which should contribute to an effective wing shape optimization.

Item, Cem C.↗
NASA NTRS1995

Tire footprint studies

This presentation covers the results of tire footprint studies conducted in the Landing Gear Development Facility of the USAF Wright Laboratory at the Wright Patterson Air force Base, OH. Tire footprint studies are essential in understanding tire wear mechanisms and computing tire tread wear rates. The power input into the tread is the driving force for tread wear. Variables needed for power input calculations include the footprint pressure and slip velocity distributions. Studies were performed on the effects of power input distributions due to vertical load, camber, yaw, inflation pressure, and tire construction. For the present study, two tire constructions, one radial and the other bias, were selected. These tires were for the F-16 Block 30 fighter aircraft, both of which were previously worn. The present study was limited to steady straight roll with a 14,000 lb vertical load, a 310 psi inflation pressure, and zero yaw and camber. All tests were conducted on the Tire Force Machine (TFM) with a specialized sensor plate with embedded pressure sensors (X, Y, and Z) and slip sensors (X and Y). All tests were conducted for a table speed of 1 in/s. Tests on the TFM show that the power intensity distributions and total power for both tire constructions are quite similar for straight roll. Later on, tests were also conducted on a modified dynamometer which was overlaid with a grit wear surface. The tire speed was maintained at 40 miles per hour and yaw was set to four degrees. Dynamometer tests showed that radial tires have more tread wear than the bias tire; however, in the field, radial tires have longer life.

Chawla, Mangal↗
NASA NTRS1998

Evaluation of 3D Inverse Code Using Rotor 67 as Test Case

A design modification of Rotor 67 is carried out with a full 3D inverse method. The blade camber surface is modified to produce a prescribed pressure loading distribution, with the blade tangential thickness distribution and the blade stacking line at midchord kept the same as the original Rotor 67 design. Because of the inviscid-flow assumption used in the current version of the method, Rotor 67 geometry is modified for use at a design point different from the original design value. A parametric study with the prescribed pressure loading distribution yields the following results. In the subsonic section, smooth pressure loading shapes generally produce blades with well-behaved blade surface pressure distributions. In the supersonic section, the study shows that the strength and position of the passage shock correlate with the characteristics of the blade pressure loading shape. In general, "smooth" prescribed blade pressure loading distributions generate blade designs with reverse cambers which have the effect of weakening the passage shock.

Dang, T.↗
NASA NTRS1999

A PTC Optimization and Control Surface Interference Study HSR Airframe Technical Review

This report considers the effect of canard and horizontal tail vertical position on the aerodynamic characteristics of the PTC configuration without nacelles and diverters. This analysis is followed by three optimization studies using canard and tail incidence as design variables in the first problem followed by an optimization run with canard and tail incidence and wing camber design variables and finally an optimization run with canard incidence and wing camber. The first problem was run at fixed lift while the other two problems were run at fixed angle of attack. The final investigation reported here will show data from a component buildup study using the PTC configuration. This final study will show the aerodynamic interference between the canard, wing and horizontal tail.

Hicks, Raymond↗
NASA NTRS1999

Supersonic Aerodynamic Design Improvements of an Arrow-Wing HSCT Configuration Using Nonlinear Point Design Methods

This paper is a discussion of the supersonic nonlinear point design optimization efforts at McDonnell Douglas Aerospace under the High-Speed Research (HSR) program. The baseline for these optimization efforts has been the M2.4-7A configuration which represents an arrow-wing technology for the High-Speed Civil Transport (HSCT). Optimization work on this configuration began in early 1994 and continued into 1996. Initial work focused on optimization of the wing camber and twist on a wing/body configuration and reductions of 3.5 drag counts (Euler) were realized. The next phase of the optimization effort included fuselage camber along with the wing and a drag reduction of 5.0 counts was achieved. Including the effects of the nacelles and diverters into the optimization problem became the next focus where a reduction of 6.6 counts (Euler W/B/N/D) was eventually realized. The final two phases of the effort included a large set of constraints designed to make the final optimized configuration more realistic and they were successful albeit with a loss of performance.

Unger, Eric R.↗
NASA NTRS2005

Theoretical and Experimental Unsteady Aerodynamics Compared for a Linear Oscillating Cascade With a Supersonic Leading-Edge Locus

Experimental data were obtained to help validate analytical and computational fluid dynamics (CFD) codes used to compute unsteady cascade aerodynamics in a supersonicaxial- flow regime. Results from two analytical codes and one CFD code were compared with experimental data. One analytical code did not account for airfoil thickness or camber; another, using piston theory (piston code), accounted for thickness and camber upstream of the first shockwave/airfoil impingement locations. The Euler CFD code accounted fully for airfoil shape.

Ramsey, John K.↗
NASA NTRS2009

SWIFT Code Assessment for Two Similar Transonic Compressors

One goal of the NASA Fundamental Aeronautics Program is the assessment of computational fluid dynamic (CFD) codes used for the design and analysis of many aerospace systems. This paper describes the assessment of the SWIFT turbomachinery analysis code for two similar transonic compressors, NASA rotor 37 and stage 35. The two rotors have identical blade profiles on the front, transonic half of the blade but rotor 37 has more camber aft of the shock. Thus the two rotors have the same shock structure and choking flow but rotor 37 produces a higher pressure ratio. The two compressors and experimental data are described here briefly. Rotor 37 was also used for test cases organized by ASME, IGTI, and AGARD in 1994-1998. Most of the participating codes over predicted pressure and temperature ratios, and failed to predict certain features of the downstream flowfield. Since then the AUSM+ upwind scheme and the k- turbulence model have been added to SWIFT. In this work the new capabilities were assessed for the two compressors. Comparisons were made with overall performance maps and spanwise profiles of several aerodynamic parameters. The results for rotor 37 were in much better agreement with the experimental data than the original blind test case results although there were still some discrepancies. The results for stage 35 were in very good agreement with the data. The results for rotor 37 were very sensitive to turbulence model parameters but the results for stage 35 were not. Comparison of the rotor solutions showed that the main difference between the two rotors was not blade camber as expected, but shock/boundary layer interaction on the casing.

Chima, Rodrick V.↗
NASA NTRS2013

Elastically Shaped Wing Optimization and Aircraft Concept for Improved Cruise Efficiency

This paper presents the findings of a study conducted tn 2010 by the NASA Innovation Fund Award project entitled "Elastically Shaped Future Air Vehicle Concept". The study presents three themes in support of meeting national and global aviation challenges of reducing fuel burn for present and future aviation systems. The first theme addresses the drag reduction goal through innovative vehicle configurations via non-planar wing optimization. Two wing candidate concepts have been identified from the wing optimization: a drooped wing shape and an inflected wing shape. The drooped wing shape is a truly biologically inspired wing concept that mimics a seagull wing and could achieve about 5% to 6% drag reduction, which is aerodynamically significant. From a practical perspective, this concept would require new radical changes to the current aircraft development capabilities for new vehicles with futuristic-looking wings such as this concept. The inflected wing concepts could achieve between 3% to 4% drag reduction. While the drag reduction benefit may be less, the inflected-wing concept could have a near-term impact since this concept could be developed within the current aircraft development capabilities. The second theme addresses the drag reduction goal through a new concept of elastic wing shaping control. By aeroelastically tailoring the wing shape with active control to maintain optimal aerodynamics, a significant drag reduction benefit could be realized. A significant reduction in fuel burn for long-range cruise from elastic wing shaping control could be realized. To realize the potential of the elastic wing shaping control concept, the third theme emerges that addresses the drag reduction goal through a new aerodynamic control effector called a variable camber continuous trailing edge flap. Conventional aerodynamic control surfaces are discrete independent surfaces that cause geometric discontinuities at the trailing edge region. These discontinuities promote vorticities which result in drag rises as well as noise sources. The variable camber trailing edge flap concept could provide a substantial drag reduction benefit over a conventional discrete flap system. Aerodynamic simulations show a drag reduction of over 50% could be achieved with the flap concept over a conventional discrete flap system.

Nguyen, Nhan↗
NASA NTRS2014

Computational Investigation of Novel Tip Leakage Mitigation Methods for High Pressure Turbine Blades

This paper presents preliminary findings on a possible approach to reducing tip leakage losses. In this paper a computational study was conducted on the Energy Efficient Engine (EEE) High Pressure Turbine (HPT) rotor tip geometry using the commercial numerical solver ANSYS FLUENT. The flow solver was validated against aerodynamic data acquired in the NASA Transonic Turbine Blade Cascade facility. The scope of the ongoing study is to computationally investigate how the tip leakage and overall blade losses are affected by (1) injection from the tip near the pressure side, (2) injection from the tip surface at the camber line, and (3) injection from the tip surface into the tip separation bubble. The objective is to identify the locations on the tip surface at which to place appropriately configured blowing keeping in mind the film cooling application of tip blowing holes. The validation was conducted at Reynolds numbers of 85,000, 343,000, and 685,000 and at engine realistic flow conditions. The coolant injection simulations were conducted at a Reynolds number of 343,000 based on blade chord and inlet velocity and utilized the SST turbulence model in FLUENT. The key parameters examined are the number of jets, jet angle and jet location. A coolant to inlet pressure ratio of 1.0 was studied for angles of +30 deg, -30 deg, and 90 deg to the local free stream on the tip. For the 3 hole configuration, 3 holes spaced 3 hole diameters apart with length to diameter ratio of 1.5 were used. A simulation including 11 holes along the entire mean camber line is also presented (30 deg toward suction side). In addition, the effect of a single hole is also compared to a flat tip with no injection. The results provide insight into tip flow control methods and can be used to guide further investigation into tip flow control. As noted in past research it is concluded that reducing leakage flow is not necessarily synonymous with reducing losses due to leakage.

Ibrahim, Mounir↗