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NASA NTRS2019

Independent Effects of Reynolds Number and Mach Number on the Aerodynamics of an Iced Swept Wing

Aerodynamic assessment of icing effects on swept wings is an important component of a larger effort to improve three-dimensional (3D) icing simulation capabilities. An understanding of ice-shape geometric fidelity and Reynolds and Mach number effects on the iced-wing aerodynamics is needed to guide the development and validation of ice-accretion simulation tools. To this end, wind tunnel testing was carried out for a 13.3-percent-scale semispan wing based upon the Common Research Model airplane configuration. The wind tunnel testing was conducted at the Office National d’Etudes et de Recherches Aérospatiales (ONERA) F1 pressurized wind tunnel with Reynolds numbers of 1.6 x 10(exp 6) to 11.9 x 10(exp 6 ) and Mach numbers of 0.09 to 0.34. Five different configurations were investigated using fully 3D, high-fidelity artificial ice shapes that maintain nearly all of the 3D ice-accretion features documented in prior icing wind tunnel tests. These large, leading-edge ice shapes were nominally based upon airplane holding in icing conditions scenarios. For three of these configurations, lower fidelity simulations were also built and tested. The results presented in this paper show that while Reynolds and Mach number effects are important for quantifying the clean-wing performance, there is very little to no effect for an iced wing with 3D, high-fidelity artificial ice shapes or 3D smooth ice shapes with grit roughness. These conclusions are consistent with the large volume of past research on iced airfoils. However, some differences were also noted for the associated stalling angle of the iced swept wing and for various lower fidelity versions of the leading-edge ice accretion. More research is planned to further investigate the key features of ice-accretion geometry that must be simulated in lower fidelity versions in order to capture the essential aerodynamics.

Aircraft icing↗
NASA NTRS1981

Transonic swept wings studied by the lifting-line theory

Transonic swept wings are analyzed as a lifting-line problem under a small-disturbance approximation. Basic concepts and principal results of the asymptotic theory are discussed. The study focuses on straight oblique wings and V-shaped swept wings, of which the local centerline curvature can be equated to zero. The three-dimensional (3-D) perturbation of the nonlinear component flow admits a similarity flow structure but requires that all wing sections are generated from a single airfoil profile; the reduced 2-D problems in this case are solved only once for all span stations. Examples of solutions involving high subcritical and slightly supercritical component flows are demonstrated and compared with surface pressure data from 3-D computer codes based on the full-potential equation (FLO 22). Except in the neighborhood of leading edges, where the small-disturbance assumption breaks down, and in the vicinities of wing tips and the symmetry plane, where neither the theory nor the 3-D codes may claim full validity, reasonable agreement is consistently found. The explicit results from the upwash analysis, along with the similarity flow structure, provides a rational approach to the control of 3-D effects in transonic aerodynamic design studies.

Cheng, H. K.↗
NASA NTRS1982

On the track of practical forward-swept wings

Structural laminates which comprise wing-cover skins for forward swept winged aircraft are examined. The laminates are themselves composed of lamina arranged in a symmetrical and unbalanced fashion. The fibers are oriented so that no fiber has a counterpart in the same ply which is at an exact anti-angle to itself. The laminate orientation creates a wash-out in a forward swept wing and alleviates aeroelastic loading. Further discussion is devoted to center-of-pressure movement, flutter behavior, aeroelasticity and aeroelastic divergence, and wind tunnel testing of aerodynamically tailored wings. It is found that rotating the laminate to increase the divergence dynamic pressure decreases strain under aerodynamic loading. Flight tests with three models are reported, and it is concluded that divergence can be avoided by the use of an efficient composite structure.

Hertz, T. J.↗
NASA NTRS1980

Dynamic response of a forward-swept-wing model at angles of attack up to 15 deg at a Mach number of 0.8

Root mean square (rms) bending moments for a dynamically scaled, aeroelastic wing of a proposed forward swept wing, flight demonstrator airplane are presented for angles of attack up to 15 deg at a Mach number of 0.8 The 0.6 size semispan model had a leading edge forward sweep of 44 deg and was constructed of composite material. In addition to broad band responses, individual rms responses and total damping ratios are presented for the first two natural modes. The results show that the rms response increases with angle of attack and has a peak value at an angle of attack near 13 deg. In general, the response was characteristic of buffeting and similar to results often observed for aft swept wings. At an angle of attack near 13 deg, however, the response had characteristics associated with approaching a dynamic instability, although no instability was observed over the range of parameters investigated.

Doggett, R. V., Jr.↗
NASA NTRS1985

A forward-swept wing configuration designed for high maneuverability by use of three-dimensional transonic theory

Supercritical technology has been applied to the design of a forward-swept-wing fighter configuration, and an assessment has been made of the relative performance of forward versus aft sweep. The wing and canard for this forward-swept wing configuration were designed for transonic maneuver by the use of a transonic computational analysis method and a transonic design procedure. The computational method calculates the transonic flow over a canard-wing-fuselage combination so that the strong transonic induced-flow effects of the canard on the wing are taken into account. A model of this configuration was constructed and was tested in the Langley 16-Foot Transonic Tunnel. The transonic theory gave a reasonably good estimate of the wing pressure distributions at transonic maneuver conditions. Comparison of this configuration with an equivalent aft-swept wing configuration showed that, at a Mach number of 0.9 and a lift coefficient of 0.9, the two configurations have essentially the same drag. This forward-swept wing configuration was also found to have very good maneuver performance relative to the Rockwell International HiMAT highly-maneuverable aircraft configuration.

Mann, M. J.↗
NASA NTRS1992

Forebody vortex control for suppressing wing rock on a highly-swept wing configuration

Free-to-roll tests were conducted in a wind tunnel with a configuration that consisted of a highly-slender forebody and a 78 deg swept delta wing. A limit cycle oscillation was observed for angles of attack between 22 and 30 deg. In general, the main flow phenomena responsible for the wing-body-tail wing rock are the interactions between the forebody and the wing vortices. Various blowing techniques were evaluated as means of wing rock suppression. Blowing tangentially aft from leeward side nozzles near the forebody tip can damp the roll motion at low blowing rates and stop it completely at higher blowing rates. At the high rates, significant vortex asymmetries are created, causing the model to stop at a non-zero roll angle. Forward blowing and alternating right/left pulsed blowing appear to be more efficient techniques for suppressing wing rock. The oscillations can be damped almost completely at lower blowing coefficients, and, apparently, no major vortex asymmetries are induced. Good agreement is observed between this study and previous water tunnel tests on the same configuration.

Suarez, Carlos J.↗
NASA NTRS2017

Development of Experimental Icing Simulation Capability for Full-Scale Swept Wings: Hybrid Design Process, Years 1 and 2

This report presents the key results from the first two years of a program to develop experimental icing simulation capabilities for full-scale swept wings. This investigation was undertaken as a part of a larger collaborative research effort on ice accretion and aerodynamics for large-scale swept wings. Ice accretion and the resulting aerodynamic effect on large-scale swept wings presents a significant airplane design and certification challenge to air frame manufacturers, certification authorities, and research organizations alike. While the effect of ice accretion on straight wings has been studied in detail for many years, the available data on swept-wing icing are much more limited, especially for larger scales.

Aircraft icing↗
NASA NTRS1989

Crossflow-vortex instability and transition on a 45 deg swept wing

A crossflow vortex experiment on a 45 deg swept wing is currently being conducted in the Arizona State University Unsteady Wind Tunnel. The experimental apparatus is designed to produce crossflow-dominated transition by simulating infinite swept wing flow using contoured end liners in a closed throat wind tunnel. Stationary fixed-wavelength crossflow vortices are observed at several chord Reynolds numbers. The vortex wavelength which is fixed for a given Reynolds number varies with Reynolds number approximately as predicted by linear stability theory, but with the predicted wavelengths about 30 percent larger than the observed wavelengths. Travelling waves are observed both in the frequency range predicted by linear stability theory and at higher frequencies. These higher frequency waves may be harmonics of the primary crossflow waves generated by a nonlinear parametric resonance phenomena.

Dagenhart, J. Ray↗
NASA NTRS1991

Numerical simulation of swept-wing flows

The transition process characteristics of flows over swept wings were computationally modelled. The crossflow instability and crossflow/T-S wave interaction are analyzed through the numerical solution of the full three dimensional Navier-Stokes equations including unsteadiness, curvature, and sweep. The leading-edge region of a swept wing is considered in a three-dimensional spatial simulation with random disturbances as the initial conditions.

Reed, Helen L.↗
NASA NTRS1988

Current flight test experience related to structural divergence of forward-swept wings

Flight testing the X-29A forward-swept wing aircraft has required development of new flight test techniques to accomplish subcritical extrapolations to the actual structural divergence dynamic pressure of the aircraft. This paper provides current experience related to applying these techniques to analysis of flight data from the forward-swept wing in order to assess the applicability of these techniques to flight test data. The measurements required, maneuvers flown, and flight test conditions are described. Supporting analytical predictions for the techniques are described and the results using flight data are compared to these predictions. Use of the results during envelope expansion and the resulting modifications to the techniques are discussed. Some of the analysis challenges that occurred are addressed and some preliminary conclusions and recommendations are made relative to the usefulness of these techniques in the flight test environment.

Schuster, Lawrence S.↗
NASA NTRS1988

Current flight test experience related to structural divergence of forward-swept wings

Flight testing the X-29A forward-swept wing aircraft has required development of new flight test techniques to accomplish subcritical extrapolations to the actual structural divergence dynamic pressure of the aircraft. This paper provides current experience related to applying these techniques to analysis of flight data from the forward-swept wing in order to assess the applicability of these techniques to flight test data. The measurements required, manuevers flown, and flight test conditions are described. Supporting analytical predictions for the techniques are described and the results using flight data are compared to these predictions. Use of the results during envelope expansion and the resulting modifications to the techniques are discussed. Some of the analysis challenges that occurred are addressed and some preliminary conclusions and recommendations are made relative to the usefulness of these techniques in the flight test environment.

Schuster, Lawrence S.↗
NASA NTRS2022

Experimental and Computational Icing Simulation for Large Swept Wings

The purpose of this final report is to summarize the key results, findings and contributions of a large, multi-organization, multi-year effort focused on swept wing icing and aerodynamics. This research effort was jointly sponsored by NASA, FAA and ONERA and was supported by Boeing and the Universities of Illinois, Virginia, and Washington. The overall goal of this research was to improve the experimental and computational simulation capability for icing on large swept wings typical of commercial transports. This research included both ice accretion and aerodynamic studies using the NASA Common Research Model (CRM) as the reference geometry. For this work, a 65% scaled version—CRM65—was used as the full-scale baseline airplane geometry. This program marks the first non-proprietary research to generate and document full-scale swept wing ice accretions in an icing tunnel and then conduct semispan wing aerodynamic testing on scaled ice accretions at high Reynolds number (i.e., up to 11.9×106). During this work, several new experimental and computational techniques were developed or documented and new knowledge about swept wing icing aerodynamics was identified. All are documented in this report.

Aircraft icing↗
NASA NTRS1960

An Experimental Investigation of Boundary-Layer Control for Drag Reduction of a Swept-Wing Section at Low Speed and High Reynolds Numbers

An investigation of laminar boundary-layer control by suction for purposes of drag reduction at low speed and high Reynolds numbers has been conducted in the Ames 12-Foot Pressure Wind Tunnel. The model was a 72.96-inch-chord wing panel, swept back 30 deg., which was installed between end plates to approximate a wing of infinite span. The airfoil section employed was a modified NACA 66-012 in the streamwise direction. Tests were limited to controlling the flow over only the upper surface of the model. Seventeen individually controllable suction chambers were provided below the surface to induce flow through 93 spanwise slots in the surface between the 0.0052- and 0.97-chord stations. Tests were made at angles of attack of 0 deg., +/- 1.0 deg., +/- 1.5 deg., and -2.0 deg. for Reynolds numbers from approximately 1.5 x 10(exp 6) to 4.0 x 10(exp 6) per foot. In general, essentially full-chord laminar flow was obtained for all conditions with small suction quantities. Minimum profile-drag coefficients of about 0.0005 to 0.0006 were obtained for the slotted surface at maximum values of the Reynolds number; these values include the Power required to induce suction as an equivalent drag.

Gault, Donald E.↗
NASA NTRS1987

A swept wing panel in a low speed flexible walled test section

The testing of two-dimensional airfoil sections in adaptive wall tunnels is relatively widespread and has become routine at all speeds up to transonic. In contrast, the experience with the three-dimensional testing of swept panels in adaptive wall test sections is very limited, except for some activity in the 1940's at NPL, London. The current interest in testing swept wing panels led to the work covered by this report, which describes the design of an adaptive-wall swept-wing test section for a low speed wind tunnel and gives test results for a wing panel swept at 40 deg. The test section has rigid flat sidewalls supporting the panel, and features flexible top and bottom wall with ribs swept at the same angle as the wing. When streamlined, the walls form waves swept at the same angle as the wing. The C sub L (-) curve for the swept wing, determined from its pressure distributions taken with the walls streamlined, compare well with reference data which was taken on the same model, unswept, in a test section deep enough to avoid wall interference.

Goodyer, M. J.↗
NASA NTRS1989

Effect of milling machine roughness and wing dihedral on the supersonic aerodynamic characteristics of a highly swept wing

An experimental investigation was conducted to assess the effect of surface finish on the longitudinal and lateral aerodynamic characteristics of a highly-swept wing at supersonic speeds. A study of the effects of wing dihedral was also made. Included in the tests were four wing models: three models having 22.5 degrees of outboard dihedral, identical except for surface finish, and a zero-dihedral, smooth model of the same planform for reference. Of the three dihedral models, two were taken directly from the milling machine without smoothing: one having a maximum scallop height of 0.002 inches and the other a maximum scallop height of 0.005 inches. The third dihedral model was handfinished to a smooth surface. Tests were conducted in Test Section 1 of the Unitary Plan Wind Tunnel at NASA-Langley over a range of Mach numbers from 1.8 to 2.8, a range of angle of attack from -5 to 8 degrees, and at a Reynolds numbers per foot of 2 x 10(6). Selected data were also taken at a Reynolds number per foot of 6 x 10(6). Drag coefficient increases, with corresponding lift-drag ratio decreases were the primary aerodynamic effects attributed to increased surface roughness due to milling machine grooves. These drag and lift-drag ratio increments due to roughness increased as Reynolds number increased.

Darden, Christine M.↗
NASA NTRS1980

Experimental and computational study of transonic flow about swept wings

An experimental investigation of NACA 0010 and 10% circular arc wing models, swept at 45 deg, spanning a channel, and at zero angle of attack is described. Measurements include chordwise and spanwise surface pressure distributions and oil-flow patterns for a range of transonic Mach numbers and Reynolds numbers. Calculations using a new three-dimensional Navier-Stokes code and a two-equation turbulence model are included for the circular-arc wing flow. Reasonable agreement between measurements and computations is obtained.

Bertelrud, A.↗