Engineering topics

Campbell, R. L.

Publications and source records attributed to Campbell, R. L..

At least 19 records

NASA NTRS2010

Initial Multidisciplinary Design and Analysis Framework

Within the Supersonics (SUP) Project of the Fundamental Aeronautics Program (FAP), an initial multidisciplinary design & analysis framework has been developed. A set of low- and intermediate-fidelity discipline design and analysis codes were integrated within a multidisciplinary design and analysis framework and demonstrated on two challenging test cases. The first test case demonstrates an initial capability to design for low boom and performance. The second test case demonstrates rapid assessment of a well-characterized design. The current system has been shown to greatly increase the design and analysis speed and capability, and many future areas for development were identified. This work has established a state-of-the-art capability for immediate use by supersonic concept designers and systems analysts at NASA, while also providing a strong base to build upon for future releases as more multifidelity capabilities are developed and integrated.

Ozoroski, L. P.↗

NASA NTRS1995

Inverse Design of Nacelles Using Multi-Block Navier Stokes Codes

The objective of this work is to reshape a nacelle to achieve a specified nacelle pressure distribution. The nacelle may be either isolated or installed on an airplane. There are no restrictions on the attitude (toe, incidence, and roll) and position of the nacelle. The design algorithm is coupled to two different multi-block 3-D Navier Stokes flow solvers. The coupling between design and analysis is automated to the point where the design proceeds with minimal user input.

Naik, D. A.↗

NASA NTRS1992

Transonic airfoil and wing design using Navier-Stokes codes

An iterative design method has been implemented into 2D and 3D Navier-Stokes codes for the design of airfoils or wings with given target pressure distributions. The method begins with the analysis of an initial geometry, and obtains the analysis pressure distributions of that geometry. The differences between analysis pressures and target pressures are used to drive geometry changes through the use of a streamline curvature method. This paper describes the procedure that makes the iterative design method work for Navier-Stokes codes. Examples of 2D airfoil design, and 3D wing design are included. It is demonstrated that the method is highly effective for airfoil or wing design at flow conditions where no substantial separation occurs. Problems encountered in the airfoil design with shock induced flow separations are discussed.

Yu, N. J.↗

NASA NTRS1989

Optimum design of swept-forward high-aspect-ration graphite-epoxy wings

An analytical investigation of a swept-forward high-aspect-ratio graphite-epoxy transport wing is described. The objectives of this investigation are to illustrate an effective usage of the unique properties of composite materials by exploiting material tailoring and to demonstrate an integrated multidisciplinary approach for conducting this investigation.

Shuart, Mark J.↗

NASA NTRS1986

Design of a natural laminar flow wing for a transonic corporate transport

Two- and three-dimensional computational methods were used to design a wing for a transonic corporate transport that had significant runs of laminar flow on both upper and lower wing surfaces at the cruise condition. The airfoil was derived from a low-speed laminar-flow section. The contour was systematically modified based on results from a two-dimensional transonic code to give favorable pressure gradients to 50-percent chord on the upper surface and 65-percent chord on the lower surface. Three-dimensional transonic codes were used to determine the wing twist and to evaluate the aerodynamic characteristics of the complete configuration at various flight conditions. A leading-edge modification for improved stall characteristics was also designed. The theoretical pressure distributions for the final airfoil correlated well with results from wind-tunnel tests.

Campbell, R. L.↗

NASA NTRS1985

Computational wing design in support of an NLF variable sweep transition flight experiment

A natural laminar flow outer panel wing glove has been designed for a variable sweep fighter aircraft using state-of-the-art computational techniques. Testing of the design will yield wing pressure and boundary-layer data under actual flight conditions and environment. These data will be used to enhance the understanding of the interaction between crossflow and Tollmien-Schlichting disturbances on boundary-layer transition. The outer wing panel was contoured such that a wide range of favorable pressure gradients could be obtained on the wing upper surface. Extensive computations were performed to support the design effort which relied on two- and three-dimensional transonic design and analysis techniques. A detailed description of the design procedure that evolved during this study is presented. Results on intermediate designs at various stages in the design process demonstrate how the various physical and aerodynamic constraints were integrated into the design. Final results of the glove design analyzed as part of the complete aircraft configuration with a full-potential wing/body analysis code indicate that the aerodynamic design objectives were met.

Waggoner, E. G.↗

NASA NTRS1984

Aerodynamic design for improved manueverability by use of three-dimensional transonic theory

Improvements in transonic maneuver performance by the use of three-dimensional transonic theory and a transonic design procedure were examined. The FLO-27 code of Jameson and Caughey was used to design a new wing for a fighter configuration with lower drag at transonic maneuver conditions. The wing airfoil sections were altered to reduce the upper-surface shock strength by means of a design procedure which is based on the iterative application of the FLO-27 code. The plan form of the fighter configuration was fixed and had a leading edge sweep of 45 deg and an aspect ratio of 3.28. Wind-tunnel tests were conducted on this configuration at Mach numbers from 0.60 to 0.95 and angles of attack from -2 deg to 17 deg. The transonic maneuver performance of this configuration was evaluated by comparison with a wing designed by empirical methods and a wing designed primarily by two-dimensional transonic theory. The configuration designed by the use of FLO-27 had the same or lower drag than the empirical wing and, for some conditions, lower drag than the two-dimensional design. From some maneuver conditions, the drag of the two-dimensional design was somewhat lower.

Mann, M. J.↗

NASA NTRS1984

Calculated effects of varying Reynolds Number and dynamic pressure on flexible wings at transonic speeds

A computational method is described that includes the effects of static aeroelastic wing deflections in steady transonic aerodynamic calculations. This method, known as the Transonic Aero-elastic Program System (TAPS), interacts a 3D transonic computer code with boundary layer and a linear finite element structural analysis codes to calculate wing pressures and deflections. The nonlinear nature of the transonic flow makes it necessary to couple the aerodynamic and structures codes in an iterative manner. TAPS has been arranged in a modular fashion so that different aerodynamic or structures programs may be used with a minimum of coding changes required. Results obtained using two different aerodynamic codes in TAPS are given, and those results are correlated with experimental data.

Campbell, R. L.↗

NASA NTRS1983

Producibility of fibrous refractory composite insulation, FRCI 40-20

Fibrous Refractory Composite Insulation (FRCI) is a NASA-developed, second generation, reusable heat-shield material that comprises a mixture of aluminoborosilicate fibers, silica fibers, and silicon carbide. Under NASA contract, a program was conducted to demonstrate the capability for manufacturing FRCI 40-20 billets. A detailed fabrication procedure was written and validated by testing specimens from the first two billets. The material conformed to NASA requirements for density, tensile strength, modulus of rupture, thermal expansion, cristobalite content, and uniformity. Twenty-four billets were prepared to provide 20 deliverable articles. Production billets were checked for density, modulus of rupture, cristobalite content, and uniformity. Billet density ranged from 309.48 to 332.22 kg/cu m (19.32 to 20.74 lb/cu ft) and modulus of rupture from 4690 to 10,140 kPa (680 to 1470 psi). Cristobalite content was less than 1 percent. A Weibull analysis of modulus-of-rupture data indicated a 1.5 percent probability for failure below the specified strength of 4480 kPa (650 psi).

Strauss, E. L.↗

NASA NTRS1983

Design and true Reynolds number 2-D testing of an advanced technology airfoil

A NASA-industry program has been conducted to determine the accuracy of available 2-D airfoil analysis procedures over a wide range of Reynolds numbers. The program also served to develop and demonstrate effective wind tunnel model designs for use in a cryogenic environment. A Lockheed design, CRYO 12X, supercritical, shockfree airfoil was configured using a continuous curvature analytical definition of the ordinates. Test results show a very close ordinate tolerance was necessary to realize the intended pressure distribution. Correlation of test with Korn-Garabedian 2-D analysis pressure data were generally good. GRUMFOIL analysis with a sidewall correction gave a better correlation.

Reaser, J. S.↗

NASA NTRS1983

Aerodynamic design for improved maneuverability by the use of three-dimensional transonic theory

This study has examined the level of transonic maneuver performance that can be achieved by the use of three-dimensional transonic theory. The FLO-27 code of Jameson and Caughey (1977) was used to lower maneuver drag by the reduction of shock strength on the wing upper surface. A design procedure was used to provide a set of guidelines for the iterative application of FLO-27 during the wing design process. The fighter configuration utilized had a leading-edge sweep of 45 deg and an aspect ratio of 3.28. Tests were conducted at Mach numbers from 0.60 to 0.95 and angles of attack from -5 to 18 deg. The accurate prediction of wing shock systems required the inclusion of fuselage effects in the theory. The FLO-27 code gave a reasonable estimate of the experimental wing pressure distributions at transonic maneuver conditions in the region of the wing where the flow was attached. For Mach numbers between 0.60 and 0.95 and lift coefficients near 0.9, a wing designed by the use of FLO-27 showed substantially lower drag than an empirically developed, supercritical maneuver wing and a maneuver wing designed primarily by two-dimensional transonic theory. For some maneuver conditions, the drag of the two-dimensional design was somewhat lower.

Mann, M. J.↗

NASA NTRS1983

Experimental study of wing leading-edge devices for improved maneuver performance of a supercritical maneuvering fighter configuration

Wind tunnel tests were conducted to examine the use of wing leading-edge devices for improved subsonic and transonic maneuver performance. These devices were tested on a fighter configuration which utilized supercritical-wing technology. The configuration had a leading-edge sweep of 45 deg and an aspect ratio of 3.28. The tests were conducted at Mach numbers of 0.60 and 0.85 with angles of attack from -0.5 deg to 22 deg. At both Mach numbers, sharp leading-edge flaps produced vortices which greatly altered the flow pattern on the wing and resulted in substantial reductions in drag at high lift. Underwing or pylon-type vortex generators also reduced drag at high lift. The vortex generators worked better at a Mach number of 0.60. The vortex generators gave the best overall results with zero toe-in angle and when mounted on either the outboard part of the wing or at both an outboard location and halfway out the semispan. Both the flaps and the vortex generators had a minor effect on the pitching moment. Fluorescent minitufts were found to be useful for flow visualization at transonic maneuver conditions.

Mann, M. J.↗

NASA NTRS1982

Supercritical maneuvering fighter configuration. Wind-tunnel investigation at Mach numbers of 0.60 to 0.95

A wind tunnel investigation was conducted to study the application of supercritical technology to highly maneuverable combat aircraft. The configuration studied has a leading-edge sweep of 45 deg and an aspect ratio of 3.28. Two supercritical-wing shapes were tested at Mach numbers from 0.60 to 0.95 with angles of attack from -2 deg to 17 deg. On supercritical wing was designed to achieve a high level of transonic maneuver performance at a Mach number of 0.90; however, excessive flow separation developed on this wing at a Mach number of 0.85. A second supercritical wing was tested which had significantly reduced flow separation and improved drag characteristics at a Mach number of 0.85 and maintained the performance of the original wing at the higher Mach numbers. Leading-edge vortex generators did not improve the performance of the second wing; however, a sharp leading-edge flap produced sizable drag reductions at Mach numbers from 0.60 to 0.90.

Mann, M. J.↗

NASA NTRS1982

Effect of nacelles on aerodynamic characteristics of an executive-jet model with simulated, partial-chord, laminar-flow-control wing glove

Tests were conducted in the Langley High-Speed 7- by 10-Foot Tunnel using a 1/10-scale model of an executive jet to examine the effects of the nacelles on the wing pressures and model longitudinal aerodynamic characteristics. For the present investigation, each wing panel was modified with a simulated, partial-chord, laminar-flow-control glove. Horizontal-tail effects were also briefly examined. The tests covered a range of Mach numbers from 0.40 to 0.82 and lift coefficients from 0.20 to 0.55. Oil-flow photographs of the wing at selected conditions are included.

Campbell, R. L.↗

NASA NTRS1981

Computer analysis of flow perturbations generated by placement of choke bumps in a wind tunnel

An inviscid analytical study was conducted to determine the upstream flow perturbations caused by placing choke bumps in a wind tunnel. A computer program based on the stream-tube curvature method was used to calculate the resulting flow fields for a nominal free-stream Mach number range of 0.6 to 0.9. The choke bump geometry was also varied to investigate the effect of bump shape on the disturbance produced. Results from the study indicate that a region of significant variation from the free-stream conditions exists upstream of the throat of the tunnel. The extent of the disturbance region was, as a rule, dependent on Mach number and the geometry of the choke bump. In general, the upstream disturbance distance decreased for increasing nominal free-stream Mach number and for decreasing length-to-height ratio of the bump. A polynomial-curve choke bump usually produced less of a disturbance than did a circular-arc bump and going to an axisymmetric configuration (modeling choke bumps on all the tunnel walls) generally resulted in a lower disturbance than with the corresponding two dimensional case.

Campbell, R. L.↗

NASA NTRS1978

Full-scale testing of an Ogee tip rotor

Full scale tests were utilized to investigate the effect of the ogee tip on helicopter rotor acoustics, performance, and loads. Two facilities were used: the Langley whirl tower and a UH-1H helicopter. The text matrix for hover on the whirl tower involved thrust values from 0 to 44 480 N (10,000 lb) at several tip Mach numbers for both standard and Ogee rotors. The full scale testing on the UH-1H encompassed the major portion of the flight envelope for that aircraft. Both near field acoustic measurements and far field flyover data were obtained for both the ogee and standard rotors. Data analysis of the whirl tower test shows that the ogee tip does significantly diffuse the tip vortex while providing some improvement in hover performance at low and moderate thrust coefficients. Flight testing of both rotors indicates that the strong impulsive noise signature of the standard rotor can be reduced with the ogee rotor. Analysis of the spectra indicates a reduction in energy in the 250 Hz and 1000 Hz range for the ogee rotor. Forward flight performance was significantly improved with the ogee configuration for a large number of flight conditions. Further, rotor control loads were reduced through use of this advanced tip rotor.

Mantay, W. R.↗

NASA NTRS1977

Some results of the testing of a full-scale Ogee tip helicopter rotor; acoustics, loads, and performance

Full-scale tests were utilized to investigate the effect of the Ogee tip on helicopter rotor acoustics, performance, and loads. Two facilities were used for this study: the Langley whirl tower and a UH-1H helicopter. The test matrix for hover on the whirl tower involved thrust values from 0 to 44,480 N (10,000 lbs) at several tip Mach numbers for both standard and Ogee rotors. The full-scale testing on the UH-1H encompassed the major portion of the flight envelope for that aircraft. Both near-field acoustic measurements as well as far-field flyover data were obtained for both the Ogee and standard rotors. Data analysis of the whirl-tower test shows that the Ogee tip does significantly diffuse the tip vortex while providing some improvement in hover performance. Flight testing of both rotors indicates that the strong impulsive noise signature of the standard rotor can be reduced with the Ogee rotor. Forward flight performance was significantly improved with the Ogee configuration for a large number of flight conditions. Further, rotor control loads and vibrations were reduced through use of this advanced tip rotor.

Mantay, W. R.↗

NASA NTRS1976

Flight investigation of the response of a helicopter to the trailing vortex of a fixed-wing aircraft

A flight investigation was conducted to quantitatively determine the response of a medium-weight helicopter to the trailing-vortex system of a fixed-wing aircraft. Flight tests and analytical tools were both utilized in the investigation. The flight tests involved an extensively instrumented UH-1H helicopter and a C-54 aircraft. Penetrations of the vortex system by the UH-1H were made at the following nominal conditions: the C-54 flew at 5500 feet altitude at a nominal gross weight of 58,000 pounds and an indicated airspeed of 115 knots in a cruise configuration. The UH-1H, nominally 7200 pounds gross weight, flew at 60 knots indicated airspeed during the penetrations at separation distances of 6.64 nautical miles to 0.42 nautical mile between aircraft. In general, the data analyzed for the above tests indicated that no unsafe penetration occurred. Further, penetrating vehicle attitude changes and structural loads were nominal. In addition, the response of the helicopter did not change appreciably with decreased separation distance.

Mantay, W. R.↗