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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 37 records · Page 2
All sourcesNASA NTRSDOE OSTI
NASA NTRS1975

Development, fabrication and calibration of a porous surface microphone in an aerofoil

The development of a porous surface microphone in an airfoil intended to measure acoustic signals in a turbulent airflow and to minimize the flow noise is described. The sensor because of its airfoil operates over a wide range of yaw angles and flow velocities without excessive flow noise. The acoustic properties of the porous materials used in the airfoil sensor and their effects on the frequency response of the sensor were analyzed and tested. An accurate airfoil was selected, having a smaller thickness-to-chord ratio and an airfoil sensor was designed. The sensor was calibrated acoustically and its flow noise evaluated in the quiet BBN wind tunnel at flow velocities up to 70 m/sec. Results are presented.

Noiseux, D. U.↗
NASA NTRS1976

Effects of thickness on the aerodynamic characteristics of an initial low-speed family of airfoils for general aviation applications

Wind tunnel tests were conducted to determine the effects of airfoil thickness-ratio on the low speed aerodynamic characteristics of an initial family of airfoils. The results were compared with theoretical predictions obtained from a subsonic viscous method. The tests were conducted over a Mach number range from 0.10 to 0.28. Chord Reynolds numbers varied from about 2.0 x 1 million to 9.0 x 1 million.

Mcghee, R. J.↗
NASA NTRS1995

Grid Sensitivity and Aerodynamic Optimization of Generic Airfoils

An algorithm is developed to obtain the grid sensitivity with respect to design parameters for aerodynamic optimization. The procedure is advocating a novel (geometrical) parameterization using spline functions such as NURBS (Non-Uniform Rational B- Splines) for defining the airfoil geometry. An interactive algebraic grid generation technique is employed to generate C-type grids around airfoils. The grid sensitivity of the domain with respect to geometric design parameters has been obtained by direct differentiation of the grid equations. A hybrid approach is proposed for more geometrically complex configurations such as a wing or fuselage. The aerodynamic sensitivity coefficients are obtained by direct differentiation of the compressible two-dimensional thin-layer Navier-Stokes equations. An optimization package has been introduced into the algorithm in order to optimize the airfoil surface. Results demonstrate a substantially improved design due to maximized lift/drag ratio of the airfoil.

Sadrehaghighi, Ideen↗
NASA NTRS1976

Aeroelastic stability and control of an oblique wing - Wind tunnel experiments

Results are presented for wind tunnel tests of an elastic wing model to verify the theoretical predictions for the aeroelastic instability of an oblique wing. The model wing has an elliptic planform of 10 to 1 axis ratio and a symmetrical airfoil section of 7-1/2% thickness/chord ratio. The wing is of wood and as may be seen in the photographs presented, slack wires are used to limit the amplitude of unstable motions. The fuselage is mounted on bearings permitting freedom of roll, but provision is made to clamp the fuselage for some of the tests. It is found that freedom in roll increases the dynamic pressure at which aeroelastic instability first appears. With the model free in roll, the effectiveness of the ailerons in maintaining trim is not noticeably affected by passage through the speed at which the wing would become unstable if clamped.

Jones, R. T.↗
NASA NTRS1975

Pressure data from a 64A010 airfoil at transonic speeds in heavy gas media of ratio of specific heats from 1.67 to 1.12

A NACA 64A010 pressure-instrumented airfoil was tested at transonic speeds over a range of angle of attack from -1 to 12 degrees at various Reynolds numbers ranging from 2 to 6 million in air, argon, Freon 12, and a mixture of argon and Freon 12 having a ratio of specific heats corresponding to air. Good agreement of results is obtained for conditions where compressibility is not significant and for the air and comparable argon-Freon 12 mixture. Comparison of heavy gas results with air, when adjusted for transonic similarity, show improved, but less than desired agreement.

Gross, A. R.↗
NASA NTRS1979

NASA low- and medium-speed airfoil development

The status of NASA low and medium speed airfoil research is discussed. Effects of airfoil thickness-chord ratios varying from 9 percent to 21 percent on the section characteristics for a design lift coefficient of 0.40 are presented for the initial low speed family of airfoils. Also, modifications to the 17-percent low-speed airfoil to reduce the pitching-moment coefficient and to the 21-percent low speed airfoil results are shown for two new medium speed airfoils with thickness ratios of 13 percent and 17 percent and design-lift coefficients of 0.30. Applications of NASA-developed airfoils to general aviation aircraft are summarized.

Mcghee, R. J.↗
NASA NTRS1991

Generalized multi-point inverse airfoil design

In a rather general sense, inverse airfoil design can be taken to mean the problem of specifying a desired set of airfoil characteristics, such as the airfoil maximum thickness ratio, pitching moment, part of the velocity distribution or boundary-layer development, etc., then from this information determine the corresponding airfoil shape. This paper presents a method which approaches the design problem from this perspective. In particular, the airfoil is divided into segments along which, together with the design conditions, either the velocity distribution or boundary-layer development may be prescribed. In addition to these local desired distributions, single parameters like the airfoil thickness can be specified. The problem of finding the airfoil shape is determined by coupling an incompressible, inviscid, inverse airfoil design method with a direct integral boundary-layer analysis method and solving the resulting nonlinear equations via a multidimensional Newton iteration technique. The approach is fast and easily allows for interactive design. It is also flexible and could be adapted to solving compressible, inverse airfoil design problems.

Selig, Michael S.↗
NASA NTRS1986

Transonic airfoil analysis and design in nonuniform flow

A nonuniform transonic airfoil code is developed for applications in analysis, inverse design and direct optimization involving an airfoil immersed in propfan slipstream. Problems concerning the numerical stability, convergence, divergence and solution oscillations are discussed. The code is validated by comparing with some known results in incompressible flow. A parametric investigation indicates that the airfoil lift-drag ratio can be increased by decreasing the thickness ratio. A better performance can be achieved if the airfoil is located below the slipstream center. Airfoil characteristics designed by the inverse method and a direct optimization are compared. The airfoil designed with the method of direct optimization exhibits better characteristics and achieves a gain of 22 percent in lift-drag ratio with a reduction of 4 percent in thickness.

Chang, J. F.↗
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 NTRS1994

Wing design for a civil tiltrotor transport aircraft

The goal of this research is the proper tailoring of the civil tiltrotor's composite wing-box structure leading to a minimum-weight wing design. With focus on the structural design, the wing's aerodynamic shape and the rotor-pylon system are held fixed. The initial design requirement on drag reduction set the airfoil maximum thickness-to-chord ratio to 18 percent. The airfoil section is the scaled down version of the 23 percent-thick airfoil used in V-22's wing. With the project goal in mind, the research activities began with an investigation of the structural dynamic and aeroelastic characteristics of the tiltrotor configuration, and the identification of proper procedures to analyze and account for these characteristics in the wing design. This investigation led to a collection of more than thirty technical papers on the subject, some of which have been referenced here. The review of literature on the tiltrotor revealed the complexity of the system in terms of wing-rotor-pylon interactions. The aeroelastic instability or whirl flutter stemming from wing-rotor-pylon interactions is found to be the most critical mode of instability demanding careful consideration in the preliminary wing design. The placement of wing fundamental natural frequencies in bending and torsion relative to each other and relative to the rotor 1/rev frequencies is found to have a strong influence on the whirl flutter. The frequency placement guide based on a Bell Helicopter Textron study is used in the formulation of frequency constraints. The analysis and design studies are based on two different finite-element computer codes: (1) MSC/NASATRAN and (2) WIDOWAC. These programs are used in parallel with the motivation to eventually, upon necessary modifications and validation, use the simpler WIDOWAC code in the structural tailoring of the tiltrotor wing. Several test cases were studied for the preliminary comparison of the two codes. The results obtained so far indicate a good overall agreement between the two codes.

Rais-Rohani, Masoud↗
NASA NTRS1981

Numerical optimization of circulation control airfoils

A numerical procedure for optimizing circulation control airfoils, which consists of the coupling of an optimization scheme with a viscous potential flow analysis for blowing jet, is presented. The desired airfoil is defined by a combination of three baseline shapes (cambered ellipse, and cambered ellipse with drooped and spiralled trailing edges). The coefficients of these shapes are used as design variables in the optimization process. Under the constraints of lift augmentation and lift-to-drag ratios, the optimal airfoils are found to lie between those of cambered ellipse and the drooped trailing edge, towards the latter as the angle of attack increases. Results agree qualitatively with available experimental data.

Tai, T. C.↗
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 NTRS1954

Impingement of water droplets on wedges and double-wedge airfoils at supersonic speeds

An analytical solution has been obtained for the equations of motion of water droplets impinging on a wedge in a two-dimensional supersonic flow field with a shock wave attached to the wedge. The closed-form solution yields analytical expressions for the equation of the droplet trajectory, the local rate of impingement and the impingement velocity at any point on the wedge surface, and the total rate of impingement. The analytical expressions are utilized to determine the impingement on the forward surfaces of diamond airfoils in supersonic flow fields with attached shock waves. The results presented include the following conditions: droplet diameters from 2 to 100 microns, pressure altitudes from sea level to 30,000 feet, free-stream static temperatures from 420 degrees r, free stream Mach numbers from 1.1 to 2.0, semiapex angles for the wedge from 1.14 degrees to 7.97 degrees, thickness-to-chord ratios for the diamond airfoil from 0.02 to 0.14, chord lengths from 1 to 20 feet, and angles of attack from zero to the inverse tangent of the airfoil thickness-to-chord ratio.

Serafini, John S↗
NASA NTRS1953

Impingement of water droplets on wedges and diamond airfoils at supersonic speeds

An analytical solution has been obtained for the equations of motion of water droplets impinging on a wedge in a two-dimensional supersonic flow field with a shock wave attached to the wedge. The closed-form solution yields analytical expressions for the equation of the droplet trajectory, the local rate of impingement and the impingement velocity at any point on the wedge surface, and the total rate of impingement. The analytical expressions are utilized to determine the impingement on the forward surfaces of diamond airfoils in supersonic flow fields with attached shock waves. The results presented include the following conditions: droplet diameters from 2 to 100 microns, pressure altitudes from sea level to 30,000 feet, free-stream static temperatures from 420 degrees to 460 degrees R. Also, free-stream Mach numbers from 1.1 to 2.0, semi-apex angles for the wedge from 1.14 degrees to 7.97 degrees, thickness-to-chord ratios for the diamond airfoil from 0.02 to 0.14, chord lengths from 1 to 20 feet, and angles of attack from zero to the inverse tangent of the airfoil thickness-to-chord ratio.

SURFACE CONDITIONS - WING SECTIONS↗
NASA NTRS2016

NASA Phase2 Unlimited Rights Final Report Research and Technology for Aerospace Propulsion Systems (RTAPS) Task Order: ERA Advanced Core Compressor Technology Program

As part of the NASA Integrated Systems Research Program (ISRP), Environmentally Responsible Aviation (ERA) Project, NASA GRC and GE Aviation have partnered to investigate the technology barriers associated with improved fuel economy of large gas turbine engines. Crucial to improving fuel economy is the increase in pressure ratio (PR) of the core high pressure compressor (HPC). A key enabler to increasing HPC pressure ratios are new technology front stage airfoil designs that have higher efficiency, higher pressure ratios and improvements in front stage part-power operability. The purpose of this work is to design and test two high pressure ratio, high efficiency front block designs. The baseline design, hereafter referred to as Build 1, is similar to what would be found in a high efficiency 30:1 PR class HPC. A redesigned front block with improved blading will also be tested and hereafter be referred to as Build 2. The goal of Build 2 is to increase stage aero loading (increase stage pressure ratio) beyond Build 1 while maintaining equivalent efficiency and stall margin. The detailed inter-stage and traverse data from the two builds will be used to improve understanding of the complex flow physics at high-speed and part-speed conditions that impact performance and operability. By addressing, understanding and solving the challenges associated with aerodynamic losses and stage matching better design strategies can be incorporated in future core compressors with higher efficiency, thereby enabling improvement in mission fuel burn while still meeting operability requirements.

Axial Compressors↗
NASA NTRS1947

Preliminary Tests in the Supersonic Sphere

This report presents preliminary data obtained in the Langley supersonic sphere. The supersonic sphere is essentially a whirling mechanism enclosed in a steel shell which can be filled with either air or Freon gas. The test models for two-dimensional study are of propeller form having the same plan form and diameter but varying only in the airfoil shape and thickness ratio. Torque coefficients for the 16-006, 65-110, and the 15 percent thick ellipse models are presented, as well as pressure distributions on a circular-arc supersonic airfoil section having a maximum thickness of 10 percent chord at the 1/3-chord position. Torque coefficients were measured in both Freon and air on the 15 percent thick ellipse, and the data obtained in air and Freon are found to be in close agreement. The torque coefficients for the three previously mentioned models showed large differences in magnitude at tip Mach numbers above 1, the model with the thickest airfoil section having the largest torque coefficient. Pressure distribution on the previously mentioned circular-arc airfoil section are presented at Mach numbers of 0.69, 1.26, and 1.42. At Mach numbers of 1.26 and 1.42 the test section is in the mixed flow region where both subsonic and supersonic speeds occur on the airfoil. No adequate theory has been developed for this condition of mixed flow, but the experimental data have been compared with values of pressure based on Ackeret's theory. The experimental data obtained at a Mach number of 1.26 on the rear portion of the airfoil section agree fairly well with the values calculated by Ackeret's theory. At a Mach number of 1.42 a larger percentage of the airfoil is in supersonic flow, and the experimental data for the entire airfoil agree fairly well with the values obtained using Ackeret's theory.

Baker, John E.↗
NASA NTRS1978

Wind tunnel tests on a tail-less swept wing span-distributed cargo aircraft configuration

The configuration consisted of a 30 deg -swept, untapered, untwisted wing utilizing a low-moment cambered airfoil of 20 percent streamwise thickness designed for low wave drag at M = 0.6, C sub L = 0.4. The tests covered a range of Mach numbers 0.3 to 0.725 and chord Reynolds number 1,100,000 to 2,040,000, angles of attack up to model buffet and sideslip angles + or - 4 deg. Configuration build up, wing pod filleting, airfoil modification and trailing edge control deflection effects were briefly investigated. Three wing tip vertical tail designs were also tested. Wing body filleting and a simple airfoil modification both produced increments to maximum lift/drag ratio. Addition of pods eliminated pitch instability of the basic wing. While the magnitude of these benefits probably was Reynolds number sensitive, they underline the potential for improving the aerodynamics of the present configuration. The cruise parameter (product of Mach number and lift/drag ratio) attained a maximum close to the airfoil design point. The configuration was found to be positively stable with normal control effectiveness about all three axes in the Mach number and C sub L range of interest.

Rao, D. M.↗