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At least 1,045 records · Page 58
All sourcesNASA NTRSDOE OSTI
NASA NTRS1959

The Static Longitudinal Stability and Control Characteristics in the Presence of the Ground of a Model Having a Triangular Wing and Canard

A wind-tunnel investigation was made of the low-speed characteristics of a canard configuration having triangular wing and canard surfaces with an aspect ratio of 2. The exposed area of the canard was 6.9 percent of the total wing area. The canard hinge line was located at 0.35 of its mean aerodynamic chord and was 0.5 wing mean aerodynamic chord lengths forward of the wing apex. The ground effects, which made the lift more positive and the -Pitching moment more negative at a given angle of attack, were unaffected by the canard. The stability of the model at a constant canard hinge-moment coefficient decreased to 0 near a lift coefficient of 1.0. In addition, the maximum lift coefficient at which the canard could provide balance was decreased by ground effects to less than 1.0 if the moment center was as far forward as 0.21 of the wing mean aerodynamic chord. The relative magnitude of interference effects between the canard and the wing and body is presented.

Buell, Donald A.↗
NASA NTRS1959

The Effect of Moment of Area Rule Modifications on the Drag, Lift and Pitching Moment Characteristics of an Unswept Aspect Ratio 6 Wing and Body Combination

An experimental investigation was conducted to determine the effect of moment-of-area-rule modifications on the drag, lift, and pitching-moment characteristics of a wing-body combination with a relatively high aspect-ratio unswept wing. The basic configuration consisted of an aspect-ratio-6 wing with a sharp leading edge and a thickness ratio of 0.06 mounted on a cut-off Sears-Haack body. The model with full moment-of-area-rule modifications had four contoured pods mounted on the wing and indentations in the body to improve the longitudinal distributions of area and moments of area. Also investigated were modifications employing pods and indentations that were only half the size of the full modifications and modifications with partial body indentations. The models were tested at angles of attack from -2 deg to +12 deg at Mach numbers from 0.6 to 1.4. In general, the moment-of-area-rule modifications had a large effect on the drag characteristics of the models but only a small effect on their lift and pitching-moment characteristics. The modifications provided substantial reductions in the zero-lift drag at transonic and low supersonic speeds, but at subsonic speeds the drag was increased. Near Mach number 1.0, the model with full modification provided the greatest reduction in drag, but at the highest test Mach numbers the half modification gave the largest drag reduction. In general, the percent reductions of zero- lift drag obtained with the aspect-ratio-6 wing were as great or greater than those previously obtained with aspect-ratio-3 wings. The effect of the modifications on the drag due to lift was small except at Mach num- bers below 0.9 where the modified models had higher drag-rise factors. Above Mach number 0.9, the modified models had higher lift-drag ratios than the basic model. The modified models also had higher lift curve slopes and generally were slightly more stable than the basic configuration.

Dickey, Robert R.↗
NASA NTRS1959

Longitudinal Stability and Control Characteristics at Mach Numbers from 0.70 to 2.22 of a Triangular Wing Configuration Equipped with a Canard Control, a Trailing-Edge-Flap Control or a Cambered Forebody

Results of an investigation to determine the static longitudinal stability and control characteristics of an aspect-ratio-2 triangular wing and body configuration equipped with either a canard control, a trailing-edge-flap control, or a cambered forebody are presented without analysis for Mach numbers from 0.70 to 2.22. The canard surface had a triangular plan form and a ratio of exposed area to total wing area of 7.8 percent. The hinge line of the canard was in the extended wing chord plane, 0.83 wing mean aerodynamic chord ahead of the reference center of moments. The trailing-edge controls were constant-chord full-span flaps with exposed area equal to 10.7 percent of the total wing area. The cambered body was a modified Sears-Haack body with camber only ahead of the wing apex. Data are presented for various canard and flap deflections at angles of attack ranging from -6 deg to +18 deg.

Boyd, John W.↗
NASA NTRS2000

Equivalent Skin Analysis of Wing Structures Using Neural Networks

An efficient method of modeling trapezoidal built-up wing structures is developed by coupling. in an indirect way, an Equivalent Plate Analysis (EPA) with Neural Networks (NN). Being assumed to behave like a Mindlin-plate, the wing is solved using the Ritz method with Legendre polynomials employed as the trial functions. This analysis method can be made more efficient by avoiding most of the computational effort spent on calculating contributions to the stiffness and mass matrices from each spar and rib. This is accomplished by replacing the wing inner-structure with an "equivalent" material that combines to the skin and whose properties are simulated by neural networks. The constitutive matrix, which relates the stress vector to the strain vector, and the density of the equivalent material are obtained by enforcing mass and stiffness matrix equities with rec,ard to the EPA in a least-square sense. Neural networks for the material properties are trained in terms of the design variables of the wing structure. Examples show that the present method, which can be called an Equivalent Skin Analysis (ESA) of the wing structure, is more efficient than the EPA and still fairly good results can be obtained. The present ESA is very promising to be used at the early stages of wing structure design.

Liu, Youhua↗
NASA NTRS2000

In-Flight Wing Pressure Distributions for the NASA F/A-18A High Alpha Research Vehicle

Pressure distributions on the wings of the F/A-18A High Alpha Research Vehicle (HARV) were obtained using both flush-mounted pressure orifices and surface-mounted pressure tubing. During quasi-stabilized 1-g flight, data were gathered at ranges for angle of attack from 5 deg to 70 deg, for angle of sideslip from -12 deg to +12 deg, and for Mach from 0.23 to 0.64, at various engine settings, and with and without the leading edge extension fence installed. Angle of attack strongly influenced the wing pressure distribution, as demonstrated by a distinct flow separation pattern that occurred between the range from 15 deg to 30 deg. Influence by the leading edge extension fence was evident on the inboard wing pressure distribution, but little influence was seen on the outboard portion of the wing. Angle-of-sideslip influence on wing pressure distribution was strongest at low angle of attack. Influence of Mach number was observed in the regions of local supersonic flow, diminishing as angle of attack was increased. Engine throttle setting had little influence on the wing pressure distribution.

Davis, Mark C.↗
NASA NTRS1996

Navier-Stokes Simulation of the Canard-Wing-Body Longitudinal Dynamic Stability Characteristics

Many modern aircraft are canard-configured for aircraft control and improved aerodynamic performance. Canards can often enhance aircraft cruise performance, maneuverability and agility. For close-coupled canard configurations, the aerodynamic interaction between the canard and wing significantly changes the flow characteristics of the wing. In unsteady flow, such changes in the flow structure and performance of wings can be quite pronounced. Accurate modeling of the unsteady aerodynamics is essential for potential CFD design and analysis of such configurations. A time-accurate numerical simulation is performed to study the unsteady aerodynamic interaction between a canard and wing with emphasis on the effects of the canard on the configuration's dynamic response characteristics. The thin-layer Reynolds-averaged Navier-Stokes Equations with various turbulence models are used in this study. Computations are made on a generic, analytically-defined, close-coupled canard-wing-body configuration which has been the subject of numerous previously published experimental studies during the 1970's to mid-80's. More recently, a series of steady-flow simulations has been performed and published by the author. In the current study, the configuration is given prescribed ramp and oscillatory motions in order to predict characteristics such as the damping-in-pitch and oscillatory longitudinal stability parameters. The current computations are made at high-subsonic and transonic Mach numbers, moderate angles-of- attack from -4 to 20 degrees, and at various pitch rates and reduced frequencies. Comparisons of pressures and integrated force quantities (e.g. lift, drag, pitching moment and selected dynamic parameters) are made with other published computational results and available experimental data. Results showing the unsteady effects of the canard on surface pressures, integrated forces, canard-wing vortex interaction and vortex breakdown will be presented.

Tu, Eugene L.↗
NASA NTRS2002

Strain Gage Loads Calibration Testing of the Active Aeroelastic Wing F/A-18 Aircraft

This report describes strain-gage calibration loading through the application of known loads of the Active Aeroelastic Wing F/A-18 airplane. The primary goal of this test is to produce a database suitable for deriving load equations for left and right wing root and fold shear; bending moment; torque; and all eight wing control-surface hinge moments. A secondary goal is to produce a database of wing deflections measured by string potentiometers and the onboard flight deflection measurement system. Another goal is to produce strain-gage data through both the laboratory data acquisition system and the onboard aircraft data system as a check of the aircraft system. Thirty-two hydraulic jacks have applied loads through whiffletrees to 104 tension-compression load pads bonded to the lower wing surfaces. The load pads covered approximately 60 percent of the lower wing surface. A series of 72 load cases has been performed, including single-point, double-point, and distributed load cases. Applied loads have reached 70 percent of the flight limit load. Maximum wingtip deflection has reached nearly 16 in.

Lokos, William A.↗
NASA NTRS2003

Introduction to the Abrupt Wing Stall (AWS) Program

The Abrupt Wing Stall (AWS) Program has addressed the problem of uncommanded, transonic lateral motions, such as wing drop, with experimental, computational, and simulation tools. Background to the establishment of the AWS program is given as well as program objectives. In order to understand the fundamental flow mechanisms that caused the undesirable motions for a pre-production version of the F/A-18E, steady and unsteady flow field details were gathered from dedicated transonic wind-tunnel testing and computational studies. The AWS program has also adapted a free-to-roll (FTR) wind-tunnel testing technique traditionally used for low-speed studies of lateral dynamic stability to the transonic flow regime. This FTR capability was demonstrated first in a proof-of -concept study and then applied to an assessment of four different aircraft configurations. Figures of merit for static testing and for FTR testing have been evaluated for two configurations that demonstrated wing drop susceptibility during full-scale flight conditions (the pre-production F/A-18E and the AV-8B at the extremes of its flight envelope) and two configurations that do not exhibit wing drop (the F/A-18C and the F-16C). Design insights have been obtained from aerodynamic computational studies of the four aircraft configurations and from computations quantifying the impact of the various geometric wing differences between the F/A-18C and the F/A-18E wings. Finally, the AWS program provides guidance for assessing, in the simulator, the impact of experimentally determined lateral activity on flight characteristics before going to flight.

Hall, Robert M.↗
NASA NTRS2003

Accomplishments of the Abrupt Wing Stall (AWS) Program and Future Research Requirements

The Abrupt Wing Stall (AWS) Program has addressed the problem of uncommanded lateral motions, such as wing drop and wing rock, at transonic speeds. The genesis of this Program was the experience of the F/A-1 8E/F Program in the late 1990's, when wing drop was discovered in the heart of the maneuver envelope for the pre-production aircraft. While the F/A-1 8E/F problem was subsequently corrected by a leading-edge flap scheduling change and the addition of a porous door to the wing fold fairing, the AWS Program was initiated as a national response to the lack of technology readiness available at the time of the F/A-18E/F Development Program. The AWS Program objectives were to define causal factors for the F/A-18E/F experience, to gain insights into the flow physics associated with wing drop, and to develop methods and analytical tools so that future programs could identify this type of problem before going to flight test. The paper reviews, for the major goals of the AWS Program, the status of the technology before the

Hall, Robert M.↗
NASA NTRS2003

Introduction to the Abrupt Wing Stall (AWS) Program

The Abrupt Wing Stall (AWS) Program has addressed the problem of uncommanded, transonic lateral motions, such as wing drop, with experimental, computational, and simulation tools. Background to the establishment of the AWS program is given as well as program objectives. In order to understand the fundamental flow mechanisms that caused the undesirable motions for a pre-production version of the F/A-18E, steady and unsteady flow field details were gathered from dedicated transonic wind-tunnel testing and computational studies. The AWS program has also adapted a free-to- roll (FTR) wind-tunnel testing technique traditionally used for low-speed studies of lateral dynamic stability to the transonic flow regime. This FTR capability was demonstrated first in a proof-of-concept study and then applied to an assessment of four different aircraft configurations. Figures of merit for static testing and for FTR testing have been evaluated for two configurations that demonstrated wing drop susceptibility during full-scale flight conditions (the pre-production F/A-l8E and the AV-8B at the extremes of its flight envelope) and two configurations that do not exhibit wing drop (the F/A-l8C and the F-l6C). Design insights have been obtained from aerodynamic computational studies of the four aircraft configurations and from computations quantifying the impact of the various geometric wing differences between the F/A-18C and the F/A-l8E wings. Finally, the AWS program provides guidance for assessing, in the simulator, the impact of experimentally determined lateral activity on flight characteristics before going to flight.

Hall, Robert M.↗
NASA NTRS2003

Accomplishments of the Abrupt Wing Stall (AWS) Program and Future Research Requirements

The Abrupt Wing Stall (AWS) Program has addressed the problem of uncommanded lateral motions, such as wing drop and wing rock, at transonic speeds. The genesis of this Program was the experience of the F/A-18E/F Program in the late 199O's, when wing drop was discovered in the heart of the maneuver envelope for the pre-production aircraft. While the F/A-18E/F problem was subsequently corrected by a leading-edge flap scheduling change and the addition of a porous door to the wing fold fairing, the AWS Program was initiated as a national response to the lack of technology readiness available at the time of the F/A-18E/F Development Program. The AWS Program objectives were to define causal factors for the F/A-18E/F experience, to gain insights into the flow physics associated with wing drop, and to develop methods and analytical tools so that future programs could identify this type of problem before going to flight test. The paper reviews, for the major goals of the AWS Program, the status of the technology before the program began, the program objectives, accomplishments, and impacts. Lessons learned are presented for the benefit of future programs that must assess whether a vehicle will have uncommanded lateral motions before going to flight test. Finally, recommended future research needs are presented in light of the AWS Program experience.

Hall, Robert M.↗
NASA NTRS1947

Summary of Available Data Relating to Reynolds Number Effects on the Maximum Lift Coefficients of Swept-Back Wings

The available foreign and American data relating to Reynolds number effects on the maximum lift coefficients of swept-back wings are summarized and discussed. The data show that at low Reynolds numbers (below about 2.0 x 10(exp 6)) higher maximum lift coefficients were measured in most cases for moderately swept-back wings than for unswept wings of similar plan form; at high Reynolds numbers, however, increasing sweepback resulted in decreasing maximum lift coefficients. A smaller rate of increase of the maximum lift coefficient with Reynolds number was measured for the swept-back wings than for similar unswept wings in the critical range of Reynolds number. Increasing the Reynolds number resulted in decreases in the maximum lift coefficients of the two wings of approximately triangular plan form that were investigated.

Sweberg, Harold H.↗
NASA NTRS1960

Characteristics at Mach Number of 2.03 of a Series of Wings Having Various Spanwise Distributions of Thickness Ratio and Chord

A series of semispan wing models having various spanwise distributions of both thickness ratio and chord but having the same effective thickness ratio was tested in the Langley 4-by 4-foot supersonic pressure tunnel at Mach number 2.03 and Reynolds numbers from 1.9 x 10(exp 6) to 6.5 x 10(exp 6) complex wing forms with thickened roots, extended root chords, and higher volumes show appreciably lower zero-lift wave drag coefficients than the plain swept wings. A calculative technique for the determination of wave drag has been applied to one of the complex wings of the series and good agreement is shown with experimental results. The complex wing forms showed higher drags due to lift than the plain swept wings.

Robins, A. Warner↗
NASA NTRS1961

Heat-Transfer Measurements at a Mach Number of 4.95 on Two 60 deg Swept Delta Wings with Blunt Leading Edges and Dihedral Angles of 0 deg and 45 deg

An experimental investigation was conducted to evaluate the heat-transfer characteristics of two 60 deg swept delta wings with cylindrical leading edges of 0.25-inch radii and dihedral angles of 0 deg and 45 deg. The tests were conducted at a Mach number of 4.95 and a stagnation temperature of 400 F. The The test-section unit Reynolds number was varied from 1.95 x 10(exp 6) to 12.24 x 10(exp 6) per foot. The results of the investigation indicated that, in a plane normal to the leading edge, the laminar-flow heat-transfer distribution was in good agreement with two-dimensional blunt-body theory. The stagnation-line heat-transfer level could be predicted from two-dimensional blunt-body theory provided the stagnation-line heat-transfer coefficient was assumed to vary as the cosine of the effective sweep. A comparison of the heating rates to the 0 deg dihedral wing (planform sweep of 60 deg) and the 45 deg dihedral wing (planform sweep of 69.3 deg) with equal panel sweep and panel area indicated that the stagnation-line heat-transfer coefficient for the 45 deg dihedral wing could be as much as 40 percent less than the stagnation-line heat-transfer coefficient for the 0 deg dihedral wing at both equal angles of attack and equal lifts. The laminar-flow heat-transfer rate to both wings outside the vicinity of the stagnation line was essentially equal.

Stainback, P. Calvin↗
NASA NTRS1960

Effect of Body-Mounted Lateral Controls and Speed Brakes on the Aerodynamic Load Distribution over a 45 deg Swept Wing at Mach Numbers from 0.80 to 0.98

An investigation has been made in the Langley 16-foot transonic tunnel to determine the effect of body-mounted lateral controls and speed brakes on the aerodynamic load distribution over a swept wing. The lateral controls and speed brakes consisted of flat plates which rotated out of the side of the fuselage, were approximately perpendicular to the wing chord plane, and extended either above or below the chord plane. The wing had 45 deg sweep of the quarter-chord line, an aspect ratio of 3, a taper ratio of0.2, and 4-percent-thick airfoil section. Data were obtained at Mach numbers of 0.80, 0.94, and 0.98 fir angels of attack that usually ranged from about 0 deg to 21 deg. The results show that at the higher angles of attack a lower-surface body-mounted lateral control located along the wing trailing edge had higher effectiveness than a similar upper-surface control. Reduction in span from 0.3 to 0.2 of the wing semispan of an upper-surface body-mounted lateral control located along the wing trailing edge resulted in a less than proportiona1,change in control effectiveness.

West, F. E., Jr.↗
NASA NTRS1952

An Investigation of the Effects of a Vortex-Generator Configuration on the Aerodynamic Characteristics of a 1/4-Scale Model of the X-1 Airplane (10-Percent-Thick Wing)

An investigation of a vortex-generator configuration on the wings of a l/4-scale model of the X-1 airplane having a 10-percent-thick wing was conducted in the Langley 16-foot transonic tunnel. The effect of the vortex generators was determined by comparing the model aerodynamic characteristics, wing-pressure distributions, and wing-wake patterns for model configurations with and without vortex generators on the wings. Results are presented from tests at 0.1 increments in Mach number from about 0.69 to 0.99, at Reynolds numbers of about 4.1 x 10(exp 6) to 4.7 x 10(exp 6), and through an angle-of-attack range up to 1.5 deg at lower speeds and up to 5 deg at the highest speed. In general, little difference in the aerodynamic characteristics was observed, except at a Mach number of 0.90 where a rearward movement of the shock on the upper surface of the wing with the vortex generators installed resulted in less separation.

Runckel, Jack F.↗
NASA NTRS2005

Twist Model Development and Results From the Active Aeroelastic Wing F/A-18 Aircraft

Understanding the wing twist of the active aeroelastic wing F/A-18 aircraft is a fundamental research objective for the program and offers numerous benefits. In order to clearly understand the wing flexibility characteristics, a model was created to predict real-time wing twist. A reliable twist model allows the prediction of twist for flight simulation, provides insight into aircraft performance uncertainties, and assists with computational fluid dynamic and aeroelastic issues. The left wing of the aircraft was heavily instrumented during the first phase of the active aeroelastic wing program allowing deflection data collection. Traditional data processing steps were taken to reduce flight data, and twist predictions were made using linear regression techniques. The model predictions determined a consistent linear relationship between the measured twist and aircraft parameters, such as surface positions and aircraft state variables. Error in the original model was reduced in some cases by using a dynamic pressure-based assumption and by using neural networks. These techniques produced excellent predictions for flight between the standard test points and accounted for nonlinearities in the data. This report discusses data processing techniques and twist prediction validation, and provides illustrative and quantitative results.

Lizotte, Andrew↗
NASA NTRS2007

Modeling and Optimization for Morphing Wing Concept Generation

This report consists of two major parts: 1) the approach to develop morphing wing weight equations, and 2) the approach to size morphing aircraft. Combined, these techniques allow the morphing aircraft to be sized with estimates of the morphing wing weight that are more credible than estimates currently available; aircraft sizing results prior to this study incorporated morphing wing weight estimates based on general heuristics for fixed-wing flaps (a comparable "morphing" component) but, in general, these results were unsubstantiated. This report will show that the method of morphing wing weight prediction does, in fact, drive the aircraft sizing code to different results and that accurate morphing wing weight estimates are essential to credible aircraft sizing results.

Skillen, Michael D.↗