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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 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.↗
NASA NTRS2007

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

Understanding the wing twist of the active aeroelastic wing (AAW) 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. This technique 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 M.↗
NASA NTRS1993

Lift augmentation for highly swept wing aircraft

A pair of spaced slots, disposed on each side of an aircraft centerline and spaced well inboard of the wing leading edges, are provided in the wing upper surfaces and directed tangentially spanwise toward thin sharp leading wing edges of a highly swept, delta wing aircraft. The slots are individually connected through separate plenum chambers to separate compressed air tanks and serve, collectively, as a system for providing aircraft lift augmentation. A compressed air supply is tapped from the aircraft turbojet power plant. Suitable valves, under the control of the aircraft pilot, serve to selective provide jet blowing from the individual slots to provide spanwise sheets of jet air closely adjacent to the upper surfaces and across the aircraft wing span to thereby create artificial vortices whose suction generate additional lift on the aircraft. When desired, or found necessary, unequal or one-side wing blowing is employed to generate rolling moments for augmented lateral control. Trailing flaps are provided that may be deflected differentially, individually, or in unison, as needed for assistance in take-off or landing of the aircraft.

Rao, Dhanvada M.↗
NASA NTRS2008

Fiber Optic Wing Shape Sensing on NASA's Ikhana UAV

This document discusses the development of fiber optic wing shape sensing on NASA's Ikhana vehicle. The Dryden Flight Research Center's Aerostructures Branch initiated fiber-optic instrumentation development efforts in the mid-1990s. Motivated by a failure to control wing dihedral resulting in a mishap with the Helios aircraft, new wing displacement techniques were developed. Research objectives for Ikhana included validating fiber optic sensor measurements and real-time wing shape sensing predictions; the validation of fiber optic mathematical models and design tools; assessing technical viability and, if applicable, developing methodology and approaches to incorporate wing shape measurements within the vehicle flight control system; and, developing and flight validating approaches to perform active wing shape control using conventional control surfaces and active material concepts.

Richards, Lance↗
NASA NTRS2009

Hybrid Wing Body Configuration System Studies

The objective of this study was to develop a hybrid wing body (HWB) sizing and analysis capability, apply that capability to estimate the fuel burn potential for an HWB concept, and identify associated technology requirements. An advanced tube with wings concept was also developed for comparison purposes. NASA s Flight Optimization System (FLOPS) conceptual aircraft sizing and synthesis software was modified to enable the sizing and analysis of HWB concepts. The noncircular pressurized centerbody of the HWB concept was modeled, and several options were created for defining the outboard wing sections. Weight and drag estimation routines were modified to accommodate the unique aspects of an HWB configuration. The resulting capability was then utilized to model a proprietary Boeing blended wing body (BWB) concept for comparison purposes. FLOPS predicted approximately a 15 percent greater drag, mainly caused by differences in compressibility drag estimation, and approximately a 5 percent greater takeoff gross weight, mainly caused by the additional fuel required, as compared with the Boeing data. Next, a 777-like reference vehicle was modeled in FLOPS and calibrated to published Boeing performance data; the same mission definition was used to size an HWB in FLOPS. Advanced airframe and propulsion technology assumptions were applied to the HWB to develop an estimate for potential fuel burn savings from such a concept. The same technology assumptions, where applicable, were then applied to an advanced tube-with-wings concept. The HWB concept had a 39 percent lower block fuel burn than the reference vehicle and a 12 percent lower block fuel burn than the advanced tube-with-wings configuration. However, this fuel burn advantage is partially derived from assuming the high-risk technology of embedded engines with boundary-layer-ingesting inlets. The HWB concept does have the potential for significantly reduced noise as a result of the shielding advantages that are inherent with an over-body engine installation.

Nickol, Craig L.↗
NASA NTRS1934

Wing-Nacelle-Propeller Tests - Comparative Tests of Liquid-Cooled and Air-Cooled Engine Nacelles

This report gives the results of measurements of the lift, drag, and propeller characteristics of several wing and nacelle combinations with a tractor propeller. The nacelles were so located that the propeller was about 31% of the wing chord directly ahead of the leading edge of the wing, a position which earlier tests (NASA Report No. 415) had shown to be efficient. The nacelles were scale models of an NACA cowled nacelle for a radial air-cooled engine, a circular nacelle with the V-type engine located inside and the radiator for the cooling liquid located inside and the radiator for the type, and a nacelle shape simulating the housing which would be used for an extension shaft if the engine were located entirely within the wing. The propeller used in all cases was a 4-foot model of Navy No. 4412 adjustable metal propeller. The results of the tests indicate that, at the angles of attack corresponding to high speeds of flight, there is no marked advantage of one type of nacelle over the others as far as low drag is concerned, since the drag added by any of the nacelles in the particular location ahead of the wing is very small. The completely cowled nacelle for a radial air-cooled engine appears to have the highest drag, the liquid-cooled engine appears to have the highest drag, the liquid-cooled engine nacelle with external radiator slightly less drag. The liquid-cooled engine nacelle with radiator in the cowling hood has about half the drag of the cowled radial air-cooled engine nacelle. The extension-shaft housing shows practically no increase in drag over that of the wing alone. A large part of the drag of the liquid-cooled engine nacelle appears to be due to the external radiator. The maximum propulsive efficiency for a given propeller pitch setting is about 2% higher for the liquid-cooled engine nacelle with the radiator in the cowling hood than that for the other cowling arrangements.

Wood, Donald H.↗
NASA NTRS2010

Slot Nozzle Effects for Reduced Sonic Boom on a Generic Supersonic Wing Section

NASA has conducted research programs to reduce or eliminate the operational restrictions of supersonic aircraft over populated areas. Restrictions are due to the disturbance from the sonic boom, caused by the coalescence of shock waves formed off the aircraft. Results from two-dimensional computational fluid dynamic (CFD) analyses (performed on a baseline Mach 2.0 nozzle in a simulated Mach 2.2 flow) indicate that over-expanded and under-expanded operation of the nozzle has an effect on the N-wave boom signature. Analyses demonstrate the feasibility of reducing the magnitude of the sonic boom N-wave by controlling the nozzle plume interaction with the nozzle boat tail shock structure. This work was extended to study the impact of integrating a high aspect ratio exhaust nozzle or long slot nozzle on the trailing edge of a supersonic wing. The nozzle is operated in a highly under-expanded condition, creating a large exhaust plume and a shock at the trailing edge of the wing. This shock interacts with and suppresses the expansion wave caused by the wing, a major contributor to the sonic boom signature. The goal was to reduce the near field pressures caused by the expansion using a slot nozzle located at the wing trailing edge. Results from CFD analysis on a simulated wing cross-section and a slot nozzle indicate potential reductions in sonic boom signature compared to a baseline wing with no propulsion or trailing edge exhaust. Future studies could investigate if this effect could be useful on a supersonic aircraft for main propulsion, auxiliary propulsion, or flow control.

Castner, Raymond S.↗
NASA NTRS2010

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.

Balakumar, Ponnampalam↗