SEARCH · Search NASA
Results for “swept wing”
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.
Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.
Relaxation methods for transonic flow about wing-cylinder combinations and lifting swept wings
The mixed elliptic-hyperbolic relaxation method for obtaining steady-state solutions to two-dimensional transonic potential equations is extended to the transonic small disturbance equation in three dimensions. In particular, transonic flow is considered both about nonlifting wing-cylinder combinations and over thin lifting wings with sweep and taper. The treatment is restricted to freestream Mach numbers less than one and to wings with subsonic trailing edges.
Analysis of unswept and swept wing chordwise pressure data from an oscillating NACA 0012 airfoil experiment. Volume 1: Technical Report
The unsteady chordwise force response on the airfoil surface was investigated and its sensitivity to the various system parameters was examined. A further examination of unsteady aerodynamic data on a tunnel spanning wing (both swept and unswept), obtained in a wind tunnel, was performed. The main body of this data analysis was carried out by analyzing the propagation speed of pressure disturbances along the chord and by studying the behavior of the unsteady part of the chordwise pressure distribution at various points of the airfoil pitching cycle. It was found that Mach number effects dominate the approach to and the inception of both static and dynamic stall. The stall angle decreases as the Mach number increases. However, sweep dominates the load behavior within the stall regime. Large phase differences between unswept and swept responses, that do not exist at low lift coefficient, appear once the stall boundary is penetrated. It was also found that reduced frequency is not a reliable indicator of the unsteady aerodynamic response in the high angle of attack regime.
Wing Load Distribution on a Swept Wing Airplane in Flight at Mach Numbers up to 1.11, and Comparison with Theory
Explore the source record for details and available documents.
Investigation at High Subsonic Speeds of Finned and Unfinned Bodies Mounted at Various Locations from the Wings of Unswept- and Swept-wing--fuselage Models, Including Measurements of Body Loads
Explore the source record for details and available documents.
Flight investigation and analysis of the wing deformations of a swept-wing bomber during push-pull maneuvers
Explore the source record for details and available documents.
Flight investigation and analysis of the wing deformations on a swept-wing bomber during rolling maneuvers
Explore the source record for details and available documents.
Effect of Geometric Fidelity on the Aerodynamics of a Swept Wing with Glaze Ice Accretion
Aerodynamic assessment of icing effects on swept wings is an important component of a larger effort to improve three-dimensional icing simulation capabilities. An understanding of ice-shape geometric fidelity on iced-wing aerodynamics and the associated flowfield features are needed to guide the development and validation of ice-accretion simulation tools. To this end, wind-tunnel testing was carried out for 8.9% and 13.3% scale semispan wing models based upon the Common Research Model airplane configuration. Various levels of geometric fidelity of an artificial ice shape representing a glaze-ice accretion on a swept wing were investigated. The highest fidelity artificial ice shape reproduced all of the three-dimensional features associated with the glaze ice accretion. The lowest fidelity artificial ice shapes were simple, spanwise-varying horn ice geometries intended to represent the maximum ice thickness on the wing upper surface. The results presented in this paper show that the addition of grit roughness to some lower-fidelity artificial ice shapes resulted in favorable lift and pitching moment comparisons to the wing with the highest fidelity artificial ice shape. In the range of 4.3 to 7.4 deg. angle of attack, surface oil flow visualization and pressure data show that the wing with the two lower fidelity simulations clearly demonstrated a leading edge vortex dominated flowfield, referred to as type I. For the wing with the high fidelity ice shape, the flowfield at lower angles of attack was characterized by streamwise-running, counter-rotating vortical flow referred to as type II. Between 6.4 and 7.4 deg. angle of attack, the effect of the type II flow structures was significantly altered and gave way to the type I leading edge vortex. This means that for angles of attack 7.4 deg. and higher, the wing with all three configurations exhibited the same type of flowfield. This helps to explain why there is reasonably good agreement in the lift and pitching moment coefficients among these configurations.
Exploratory wind tunnel investigation of the stability and control characteristics of a three-surface, forward-swept wing advanced turboprop model
The purpose of the present investigation was to parametrically study the stability and control characteristics of a forward-swept wing three-surface turboprop model through an extended angle of attack range, including the deep-stall region. As part of a joint research program between North Carolina State University and NASA Langley Research Center, a low-speed wind tunnel investigation was conducted with a three-surface, forward-swept wing, aft-mounted, twin-pusher propeller, model, representative of an advanced turboprop configuration. The tests were conducted in the NASA Langley 12-Foot Low-Speed Wind Tunnel. The model parameters varied in the test were horizontal tail location, canard size, sweep and location, and wing position. The model was equipped with air turbines, housed within the nacelles and driven by compressed air, to model turboprop power effects. A three-surface, forward-swept wing configuration that provided satisfactory static longitudinal and lateral/directional stability was identified. The three-surface configuration was found to have greater longitudinal control and increased center of gravity range relative to a conventional (two-surface) design. The test showed that power had a large favorable effect on stability and control about all three axis in the post-stall regime.
An Analysis of Flight-Test Measurements of the Wing Structural Deformations in Rough Air of a Large Flexible Swept-Wing Airplane
An analysis is made of wing deflection and streamwise twist measurements in rough-air flight of a large flexible swept-wing bomber. Random-process techniques are employed in analyzing the data in order to describe the magnitude and characteristics of the wing deflection and twist responses to rough air. Power spectra and frequency-response functions for the wing deflection and twist responses at several spanwise stations are presented. The frequency-response functions describe direct and absolute response characteristics to turbulence and provide a convenient basis for assessing analytic calculation techniques. The wing deformations in rough air are compared with the expected deformations for quasi-static loadings of the same magnitude, and the amplifications are determined. The results obtained indicate that generally the deflections are amplified by a small amount, while the streamwise twists are amplified by factors of the order of 2.0. The magnitudes of both the deflection velocities and the twist angles are shown to have significant effects on the local angles of attack at the various stations and provide the main source of aerodynamic loading, particularly at frequencies in the vicinity of the first wing-vibration mode.
Fluidic Fence Flow Control on A 30° Swept Wing in Compressible Freestreams
A 30° swept wing with a NACA 64(3)-618 airfoil cross-section was evaluated in four different freestream conditions with Mach numbers increasing from 0.05 to 0.7 using computational fluid dynamics simulations. Active flow control (AFC) through a streamwise row of vortex generating jets was applied at 70% span, blowing inboard to obstruct the development of spanwise flow and increase lift outboard of the control location. The effectiveness of the AFC was evaluated at α = 5° and 10° as freestream Mach number increased and the AFC mass flow coefficient was held constant. Increases to lift coefficient were reduced at higher Mach numbers, but the ratio of change in lift coefficient and the AFC momentum coefficient remained roughly constant. The fluidic fence caused decreased pressure coefficients on the outboard wing sections, but this region of decreased pressure coefficient moved aft on the chord with increasing Mach number. Mach 0.7 is higher than the critical Mach number of the wing and produced interesting interplay with the AFC, as the suction surface shock moved forward on the inboard wing section but moved aft on the outboard wing section.
Fluidic Fence Flow Control on A 30° Swept Wing in Compressible Freestreams
A 30° swept wing with a NACA 64(3)-618 airfoil cross-section was evaluated in four different freestream conditions with Mach numbers increasing from 0.05 to 0.7 using computational fluid dynamics simulations. Active flow control (AFC) through a streamwise row of vortex generating jets was applied at 70% span, blowing inboard to obstruct the development of spanwise flow and increase lift outboard of the control location. The effectiveness of the AFC was evaluated at α = 5° and 10° as freestream Mach number increased and the AFC mass flow coefficient was held constant. Increases to lift coefficient were reduced at higher Mach numbers, but the ratio of change in lift coefficient and the AFC momentum coefficient remained roughly constant. The fluidic fence caused decreased pressure coefficients on the outboard wing sections, but this region of decreased pressure coefficient moved aft on the chord with increasing Mach number. Mach 0.7 is higher than the critical Mach number of the wing and produced interesting interplay with the AFC, as the suction surface shock moved forward on the inboard wing section but moved aft on the outboard wing section.
The longitudinal stability of elastic swept wings at supersonic speed
The longitudinal stability characteristics of elastic swept wings of high aspect ratio experiencing bending and torsional deformations are calculated for supersonic speed by the application of linearized lifting-surface theory. A parabolic wing deflection curve is assumed and the analysis is simplified by a number of structural approximations. The method is thereby limited in application to wings of high aspect ratio for which the root effects are small. Expressions for the lift, pitching-moment, and span load distribution characteristics are derived in terms of the elastic properties of the wing; namely, the design stress, the modulus of elasticity, the shearing modulus, and the maximum design load factor. The analysis applies to wings with leading edges swept behind the Mach lines. In all cases, however, the trailing edge is sonic or supersonic. Application of the method of analysis to wings with leading edges swept ahead of the Mach lines is discussed.
Current Experimental Basis for Modeling Ice Accretions on Swept Wings
This work presents a review of the experimental basis for modeling ice accretions on swept wings. Experimental work related to ice accretion physics on swept wings conducted between 1954 and 2004 is reviewed. Proposed models or explanations of scallop formations are singled out and discussed. Special emphasis is placed on reviewing the work done to determine the basic macroscopic mechanisms of scallop formation. The role of feather growth and its connection to scallop growth is discussed. Conceptual steps in modeling scallop formations are presented. Research elements needed for modeling are discussed.
Longitudinal Aerodynamic Characteristics of a Wing-Body-Tail Model Having a Highly Tapered, Cambered 45 degree Swept Wing of Aspect Ratio 4 at Transonic Speeds
The longitudinal aerodynamic characteristics of a wing-body-horizontal-tail configuration designed for efficient performance at transonic speeds has been investigated at Mach numbers from 0.80 to 1.03 in the Langley 16-foot transonic tunnel. The effect of adding an outboard leading-edge chord-extension to the highly tapered 45 deg. swept wing was also obtained. The average Reynolds number for this investigation was 6.7 x 10(exp 6) based on the wing mean aerodynamic chord. The relatively low tail placement as well as the addition of a chord-extension achieved some alleviation of the pitchup tendencies of the wing-fuselage configuration. The maximum trimmed lift-drag ratio was 16.5 up to a Mach number of 0.9, with the moment center located at the quarter-chord point of the mean aerodynamic chord. For the untrimmed case, the maximum lift-drag ratio was approximately 19.5 up to a Mach number of 0.9.
Receptivity and Transition of Supersonic Boundary Layers Over Swept Wings
The receptivity, stability, and transition of three-dimensional supersonic boundary layers over (1) a swept cylinder, (2) a swept wing with a sharp leading edge, and (3) a swept wing with a blunt leading edge are numerically investigated for a free-stream Mach number of 3. These computations are compared to an earlier experimental and computational study performed by Archambaud et al.1 The steady flow fields with and without roughness elements are obtained by solving the full Navier-Stokes equations. The N-factors computed in this study at the transition onset locations reported in Ref. 1 for flow over the swept cylinder are approximately 16.5 for traveling crossflow disturbances and 9 for stationary disturbances. The N-factors for the traveling crossflow are high based on our past experiences. However, they are comparatively smaller than those reported by Archambaud et al., who found N-factor values in the range of 20 to 25 for traveling disturbances and 13 to 20 for stationary disturbances. Similarly, the N-factors computed in this study for the traveling and stationary disturbances for the flow over the sharp wing are approximately 7 and 2.5, respectively, and for the flow over the blunt wing are 6.5 and 4.8, respectively. Using the envelope method, Archambaud et al. obtained values of approximately 8.0 and 4.0 for the sharp wing case and 16.0 and 12.0 for the blunt wing case.
Exploratory investigation of the effects of vortex bursting on the high angle-of-attack lateral-directional stability characteristics of highly-swept wings
A recent low-speed wind-tunnel investigation of highly-swept wings has shown that vortex breakdown at high angles of attack can cause large destabilizing effects on static lateral-directional stability characteristics, and that the destabilizing effects of vortex breakdown can be greatly aggravated by model support strut interference effects. The present paper discusses these effects based on the results of static force tests of several highly-swept wing configurations for different wind-tunnel strut setup arrangements. Also included in the paper are photographs obtained during tuft-, smoke-, and helium-bubble flow visualization studies to indicate wing flow behavior patterns