Design of a flight director/configuration management system for piloted STOL approaches

NASA Technical Reports Server (NTRS)

Hoh, R. H.; Klein, R. H.; Johnson, W. A.

1973-01-01

The design and characteristics of a flight director for V/STOL aircraft are discussed. A configuration management system for piloted STOL approaches is described. The individual components of the overall system designed to reduce pilot workload to an acceptable level during curved, decelerating, and descending STOL approaches are defined. The application of the system to augmentor wing aircraft is analyzed. System performance checks and piloted evaluations were conducted on a flight simulator and the results are summarized.

Low-speed wind tunnel investigation of a semispan STOL jet transport wing body with an upper surface blown jet flap

NASA Technical Reports Server (NTRS)

Phelps, A. E., III; Letko, W.; Henderson, R. L.

1973-01-01

An investigation of the static longitudinal aerodynamic characteristics of a semispan STOL jet transport wing-body with an upper-surface blown jet flap for lift augmentation was conducted in a low-speed wind tunnel having a 12-ft octagonal test section. The semispan swept wing had an aspect ratio of 3.92 (7.84 for the full span) and had two simulated turbofan engines mounted ahead of and above the wing in a siamese pod equipped with an exhaust deflector. The purpose of the deflector was to spread the engine exhaust into a jet sheet attached to the upper surface of the wing so that it would turn downward over the flap and provide lift augmentation. The wing also had optional boundary-layer control provided by air blowing through a thin slot over a full-span plain trailing-edge flap.

Aerodynamic characteristics of a propulsive wing-canard concept at STOL speeds

NASA Technical Reports Server (NTRS)

Stewart, V. R.

1985-01-01

A full span model of a wing/canard concept representing a fighter configuration has been tested at STOL conditions in the NASA Langley 4 x 7 meter tunnel. The results of this test are presented, and comparisons are made to previous data of the same configuration tested as a semispan model. The potential of the propulsive wing/canard to develop very high lift coefficients was investigated with several nozzle spans (nozzle aspect ratios). Although longitudinal trim was not accomplished with the blowing distributions and configurations tested, the propulsive wing/canard appears to offer an approach to managing the large negative pitching moments associated with trailing edge flap blowing. Also presented are data showing the effects of large flap deflections and relative wing/canard positions. Presented in the appendix to the report are limited lateral-directional and ground effects data, as well as wing downwash measurements.

Approach path control for powered-lift STOL aircraft

NASA Technical Reports Server (NTRS)

Clymer, D. J.; Flora, C. C.

1973-01-01

A flight control system concept is defined for approach flightpath control of an augmentor wing (or similar) powered-lift STOL configuration. The proposed STOL control concept produces aircraft transient and steady-state control responses that are familiar to pilots of conventional jet transports, and has potential for good handling qualities ratings in all approach and landing phases. The effects of trailing-edge rate limits, real-engine dynamics, and atmospheric turbulence are considered in the study. A general discussion of STOL handling qualities problems and piloting techniques is included.

An economic assessment of STOL aircraft potential including terminal area environmental considerations, volume 1

NASA Technical Reports Server (NTRS)

Solomon, H. L.; Sokolsky, S.

1974-01-01

The results of an economic and environmental study of short haul airline systems using short takeoff and landing (STOL) aircraft are presented. The STOL system characteristics were optimized for maximum patronage at a specified return on investment, while maintaining noise impact compatibility with the terminal area. Supporting studies of aircraft air pollution and hub airport congestion relief were also performed. The STOL concept specified for this study was an Augmentor Wing turbofan aircraft having a field length capability of 2,000 ft. and an effective perceived noise level of 95 EPNdB at 500 ft. sideline distance. An economic and environmental assessment of the defined STOL system and a summary of the methodology, STOL system characteristics and arena characteristics are provided.

Study of aircraft in intraurban transportation systems, volume 3

NASA Technical Reports Server (NTRS)

Stout, E. G.; Kesling, P. H.; Matteson, D. E.; Sherwood, D. E.; Tuck, W. R., Jr.; Vaughn, L. A.

1971-01-01

An investigation of three aircraft concepts, deflected slipstream STOL, helicopter VTOL, and fixed wing STOL, is presented. An attempt was made to determine the best concept for the intraurban transportation system. Desirability of the concept was based on ease of maintenance, development timing, reliability, operating costs, and the noise produced. Indications are that the deflected slipstream STOL is best suited for intraurban transportation. Tables and graphs are included.

Fluctuating pressures in flow fields of jets

NASA Technical Reports Server (NTRS)

Schroeder, J. C.; Haviland, J. K.

1976-01-01

The powered lift configurations under present development for STOL aircraft are the externally blown flap (EBF), involving direct jet impingement on the aircraft flaps, and the upper surface blown (USB), where the jet flow is attached on the upper surface of the wing and directed downwards. Towards the goal of developing scaling laws to predict unsteady loads imposed on the structural components of these STOL aircraft from small model tests, the near field fluctuating pressure behavior for the simplified cases of a round free cold jet and the same jet impinging on a flat plate was investigated. Examples are given of coherences, phase lags (giving convection velocities), and overall fluctuating pressure levels measured. The fluctuating pressure levels measured on the flat plate are compared to surface fluctuating pressure levels measured on full-scale powered-lift configuration models.

Assessment at full scale of exhaust nozzle-to-wing size on STOL-OTW acoustic characteristics

NASA Technical Reports Server (NTRS)

Von Glahn, U.; Groesbeck, D.

1979-01-01

On the basis of static zero/acoustic data obtained at model scale, the effect of exhaust nozzle size on flyover noise is evaluated at full scale for different STOL-OTW nozzle configurations. Three types of nozzles are evaluated: a circular/deflector nozzle mounted above the wing, a slot/deflector nozzle mounted on the wing, and a slot nozzle mounted on the wing. The nozzle exhaust plane location, measured from the wing leading edge was varied from 10 to 46 percent of the wing chord (flaps retracted). Flap angles of 20 deg (takeoff) and 60 deg (approach) are included in the study. Initially, perceived noise levels (PNL) are calculated as a function of flyover distance at 152 m altitude. From these plots static EPNL values, defined as flyover relative noise levels, then are obtained as functions of nozzle size for equal aerodynamic performance (lift and thrust). On the basis of these calculations, the acoustic benefits attributable to nozzle size relative to a given wing chord size are assessed.

Assessment at full scale of exhaust nozzle to wing size on STOL-OTW acoustic characteristics

NASA Technical Reports Server (NTRS)

Vonglahn, U.; Grosbeck, D.

1979-01-01

On the basis of static aero/acoustic data obtained at model scale, the effect of exhaust nozzle size on flyover noise is evaluated at full scale for different STOL-OTW nozzle configurations. Three types of nozzles are evaluated: a circular/deflector nozzle mounted above the wing; a slot/deflector nozzle mounted on the wing; and a slot nozzle mounted on the wing. The nozzle exhaust plane location, measured from the wing leading edge, was varied from 10 to 46 percent of the wing chord (flaps retracted). Flap angles of 20 deg (takeoff) and 60 deg (approach) are included in the study. Initially, perceived noise levels (PNL) are calculated as a function flyover distance at 152m altitude. From these plots, static EPNL values (defined as flyover relative noise levels), are obtained as functions of nozzle size for equal aerodynamic performance (lift and thrust). The acoustic benefits attributable to nozzle size relative to a given wing chord size are assessed.

Correlation of laser velocimeter measurements over a wing with results of two prediction techniques. [in the Langley V/STOL tunnel

NASA Technical Reports Server (NTRS)

Hoad, D. R.; Meyers, J. F.; Young, W. H., Jr.; Hepner, T. P.

1978-01-01

The flow field at the center line of an unswept wing with an aspect ratio of eight was determined using a two dimensional viscous flow prediction technique for the flow field calculation, and a three dimensional potential flow panel method to evaluate the degree of two dimensionality achieved at the wing center line. The analysis was made to provide an acceptable reference for comparison with velocity measurements obtained from a fringe type laser velocimeter optics systems operating in the backscatter mode in the Langley V/STOL tunnel. Good agreement between laser velocimeter measurements and theoretical results indicate that both methods provide a true representation of the velocity field about the wing at angles of attack of 0.6 and 4.75 deg.

Design integration and noise studies for jet STOL aircraft. Volume 1: Program summary

NASA Technical Reports Server (NTRS)

Okeefe, V. O.; Kelley, G. S.

1972-01-01

This program was undertaken to develop, through analysis, design, experimental static testing, wind tunnel testing, and design integration studies, an augmentor wing jet flap configuration for a jet STOL transport aircraft having maximum propulsion and aerodynamic performance with minimum noise generation. The program had three basic elements: (1) static testing of a scale wing section to demonstrate augmentor performance and noise characteristics; (2) two-dimensional wind tunnel testing to determine flight speed effects on performance; and (3) system design and evaluation which integrated the augmentor information obtained into a complete system and ensured that the design was compatible with the requirements for a large STOL transport having a 500-ft sideline noise of 95 PNdB or less. This objective has been achieved.

Evaluation of Flying Qualities and Guidance Displays for an Advanced Tilt-Wing STOL Transport Aircraft in Final Approach and Landing

NASA Technical Reports Server (NTRS)

Frost, Chad R.; Franklin, James A.; Hardy, Gordon H.

2002-01-01

A piloted simulation was performed on the Vertical Motion Simulator at NASA Ames Research Center to evaluate flying qualities of a tilt-wing Short Take-Off and Landing (STOL) transport aircraft during final approach and landing. The experiment was conducted to assess the design s handling qualities, and to evaluate the use of flightpath-centered guidance for the precision approach and landing tasks required to perform STOL operations in instrument meteorological conditions, turbulence, and wind. Pilots rated the handling qualities to be satisfactory for all operations evaluated except those encountering extreme crosswinds and severe windshear; even in these difficult meteorological conditions, adequate handling qualities were maintained. The advanced flight control laws and guidance displays provided consistent performance and precision landings.

Assessment at full scale of nozzle/wing geometry effects on OTW aeroacoustic characteristics. [Over The Wing STOL engine configurations

NASA Technical Reports Server (NTRS)

Groesbeck, D.; Von Glahn, U.

1979-01-01

The effects on acoustic characteristics of nozzle type and location on a wing for STOL engine over-the-wing configurations are assessed at full scale on the basis of model-scale data. Three types of nozzle configurations are evaluated: a circular nozzle with external deflector mounted above the wing, a slot nozzle with external deflector mounted on the wing and a slot nozzle mounted on the wing. Nozzle exhaust plane locations with respect to the wing leading edge are varied from 10 to 46 percent chord (flaps retracted) with flap angles of 20 deg (take-off attitude) and 60 deg (approach attitude). Perceived noise levels (PNL) are calculated as a function of flyover distance at 152 m altitude. From these plots, static EPNL values, defined as flyover relative noise levels, are calculated and plotted as a function of lift and thrust ratios. From such plots the acoustic benefits attributable to variations in nozzle/deflector/wing geometry at full scale are assessed for equal aerodynamic performance.

Flyover noise characteristics of a tilt-wing V/STOL aircraft (XC-142A)

NASA Technical Reports Server (NTRS)

Pegg, R. J.; Henderson, H. R.; Hilton, D. A.

1974-01-01

A field noise measurement investigation was conducted during the flight testing of an XC-142A tilt-wing V/STOL aircraft to define its external noise characteristics. Measured time histories of overall sound pressure level show that noise levels are higher at lower airspeeds and decrease with increased speed up to approximately 160 knots. The primary noise sources were the four high-speed, main propellers. Flyover-noise time histories calculated by existing techniques for propeller noise prediction are in reasonable agreement with the experimental data.

Identification of Spey engine dynamics in the augmentor wing jet STOL research aircraft from flight data

NASA Technical Reports Server (NTRS)

Dehoff, R. L.; Reed, W. B.; Trankle, T. L.

1977-01-01

The development and validation of a spey engine model is described. An analysis of the dynamical interactions involved in the propulsion unit is presented. The model was reduced to contain only significant effects, and was used, in conjunction with flight data obtained from an augmentor wing jet STOL research aircraft, to develop initial estimates of parameters in the system. The theoretical background employed in estimating the parameters is outlined. The software package developed for processing the flight data is described. Results are summarized.

Development of an integrated configuration management/flight director system for piloted STOL approaches

NASA Technical Reports Server (NTRS)

Hoh, R. H.; Klein, R. H.; Johnson, W. A.

1977-01-01

A system analysis method for the development of an integrated configuration management/flight director system for IFR STOL approaches is presented. Curved descending decelerating approach trajectories are considered. Considerable emphasis is placed on satisfying the pilot centered requirements (acceptable workload) as well as the usual guidance and control requirements (acceptable performance). The Augmentor Wing Jet STOL Research Aircraft was utilized to allow illustration by example, and to validate the analysis procedure via manned simulation.

Study of short haul high-density V/STOL transportation systems, volume 1

NASA Technical Reports Server (NTRS)

Solomon, H. L.

1972-01-01

The relative advantages of STOL aircraft concepts were examined by simulating the operations of a short haul high-density intercity STOL system set in two arenas, the California corridor and the Chicago-Detroit-Cleveland triangle, during the 1980 time period. The study was constrained to the use of three aircraft concepts designated as the deflected slipstream turboprop, externally blown flap, and augmentor wing turbofan configurations. The projected demographic, economic, travel demand, and travel characteristics of the representative arenas were identified. The STOL airline operating scenarios were then formulated and through the use of the aerospace modal split simulation program, the traveler modal choices involving alternative STOL concepts were estimated in the context of the total transportation environment for 1980. System combinations that presented the best potential for economic return and traveler acceptance were then identified for each STOL concept.

Interior and exterior fuselage noise measured on NASA's C-8a augmentor wing jet-STOL research aircraft

NASA Technical Reports Server (NTRS)

Shovlin, M. D.

1977-01-01

Interior and exterior fuselage noise levels were measured on NASA's C-8A Augmentor Wing Jet-STOL Research Aircraft in order to provide design information for the Quiet Short-Haul Research Aircraft (QSRA), which will use a modified C-8A fuselage. The noise field was mapped by 11 microphones located internally and externally in three areas: mid-fuselage, aft fuselage, and on the flight deck. Noise levels were recorded at four power settings varying from takeoff to flight idle and were plotted in one-third octave band spectra. The overall sound pressure levels of the external noise field were compared to previous tests and found to correlate well with engine primary thrust levels. Fuselage values were 145 + or - 3 dB over the aircraft's normal STOL operating range.

A flight-test evaluation of a go-around control system for a twin engine powered-lift STOL airplane

NASA Technical Reports Server (NTRS)

Watson, D. M.; Hardy, G. H.

1983-01-01

An automatic go-around control system was evaluated on the Augmentor Wing Jet Short Takeoff and Landing (STOL) Research Airplane (AWJSRA) as part of a study of an automatic landing system for a powered-lift STOL airplane. The results of the evaluation indicate that the go-around control system can successfully transition the airplane to a climb configuration from any initiation point during the glide-slope track or the flare maneuver prior to touchdown.

Evaluation of Four Advanced Nozzle Concepts for Short Takeoff and Landing Performance

NASA Technical Reports Server (NTRS)

Quinto, P. Frank; Kemmerly, Guy T.; Paulson, John W., Jr.

1993-01-01

Four advanced nozzle concepts were tested on a canard-wing fighter in the Langley 14- by 22-Foot Subsonic Tunnel. The four vectoring-nozzle concepts were as follows: (1) an axisymmetric nozzle (AXI); (2) an asymmetric, load balanced exhaust nozzle (ALBEN); (3) a low aspect ratio, single expansion ramp nozzle (LASERN); and (4) a high aspect ratio, single expansion ramp nozzle (HASERN). The investigation was conducted to determine the most suitable nozzle concept for short takeoff and landing (STOL) performance. The criterion for the best STOL performance was a takeoff ground roll of less than 1000 ft. At approach, the criteria were high lift and sufficient drag to maintain a glide slope of -3 to -6 deg with enough pitching-moment control from the canards. The test was performed at a dynamic pressure of 45 lb/sq ft and an angle-of-attack range of 0 to 20 deg. The nozzle pressure ratio was varied from 1.0 to 4.3 at both dry power and after burning nozzle configurations with nozzle vectoring to 60 deg. In addition, the model was tested in and out of ground effects. The ALBEN concept was the best of the four nozzle concepts tested for STOL performance.

STOL aircraft transient ground effects. Part 1: Fundamental analytical study

NASA Technical Reports Server (NTRS)

Goldhammer, M. I.; Crowder, J. P.; Smyth, D. N.

1975-01-01

The first phases of a fundamental analytical study of STOL ground effects were presented. Ground effects were studied in two dimensions to establish the importance of nonlinear effects, to examine transient aspects of ascent and descent near the ground, and to study the modelling of the jet impingement on the ground. Powered lift system effects were treated using the jet-flap analogy. The status of a three-dimensional jet-wing ground effect method was presented. It was shown, for two-dimensional unblown airfoils, that the transient effects are small and are primarily due to airfoil/freestream/ground orientation rather than to unsteady effects. The three-dimensional study showed phenomena similar to the two-dimensional results. For unblown wings, the wing/freestream/ground orientation effects were shown to be of the same order of magnitude as for unblown airfoils. This may be used to study the nonplanar, nonlinear, jet-wing ground effect.

Flight-test of the glide-slope track and flare-control laws for an automatic landing system for a powered-lift STOL airplane

NASA Technical Reports Server (NTRS)

Watson, D. M.; Hardy, G. H.; Warner, D. N., Jr.

1983-01-01

An automatic landing system was developed for the Augmentor Wing Jet STOL Research Airplane to establish the feasibility and examine the operating characteristics of a powered-lift STOL transport flying a steep, microwave landing system (MLS) glide slope to automatically land on a STOL port. The flight test results address the longitudinal aspects of automatic powered lift STOL airplane operation including glide slope tracking on the backside of the power curve, flare, and touchdown. Three different autoland control laws were evaluated to demonstrate the tradeoff between control complexity and the resulting performance. The flight test and simulation methodology used in developing conventional jet transport systems was applied to the powered-lift STOL airplane. The results obtained suggest that an automatic landing system for a powered-lift STOL airplane operating into an MLS-equipped STOL port is feasible. However, the airplane must be provided with a means of rapidly regulation lift to satisfactorily provide the glide slope tracking and control of touchdown sink rate needed for automatic landings.

Flow visualization study of close-coupled canard wing and strake wing configuration

NASA Technical Reports Server (NTRS)

Miner, D. D.; Gloss, B. B.

1975-01-01

The Langley 1/8-scale V/STOL model tunnel was used to qualitatively determine the flow fields associated with semi-span close coupled canard wing and strake wing models. Small helium filled bubbles were injected upstream of the models to make the flow visible. Photographs were taken over the angle-of-attack ranges of -10 deg to 40 deg.

Low speed wind tunnel test of a propulsive wing/canard concept in the STOL configuration. Volume 2: Test data

NASA Technical Reports Server (NTRS)

Stewart, V. R.

1987-01-01

A propulsive wind/canard model was tested at STOL operating conditions in the NASA Langley Research Center 4 x 7 meter wind tunnel. Longitudinal and lateral/directional aerodynamic characteristics were measured for various flap deflections, angles of attack and sideslip, and blowing coefficients. Testing was conducted for several model heights to determine ground proximity effects on the aerodynamic characteristics. Flow field surveys of local flow angles and velocities were performed behind both the canard and the wing. This is volume 2 of a 2 volume report. All of the test data in three appendices are presented. Appendix A presented tabulated six component force and moment data, Appendix B presents tabulated wing pressure coefficients, and Appendix C presents the flow field data.

Use of the flight simulator in the design of a STOL research aircraft.

NASA Technical Reports Server (NTRS)

Spitzer, R. E.; Rumsey, P. C.; Quigley, H. C.

1972-01-01

Piloted simulator tests on the NASA-Ames Flight Simulator for Advanced Aircraft motion base played a major role in guiding the design of the Modified C-8A 'Buffalo' augmentor wing jet flap STOL research airplane. Design results are presented for the flight control systems, lateral-directional SAS, hydraulic systems, and engine and thrust vector controls. Emphasis is given to lateral control characteristics on STOL landing approach, engine-out control and recovery techniques in the powered-lift regime, and operational flight procedures which affected airplane design.

V/STOL aircraft and method

DOEpatents

Owens, Phillip R.

1997-01-01

Aircraft apparatus and method capable of V/STOL (vertical, short takeoff and landing) in addition to conventional flight. For V/STOL operation, induced lift is provided by blowing air over the upper surface of each wing through a duct installed near the leading edge. Intake air is supplied to the blowing fan through a duct installed near the trailing edge, thus providing suction as well as blowing. Two fans in series are required. The engine provides power not only to the propeller but also to a transmission which provides power to the pulleys driving the belt-driven fans.

V/STOL tilt rotor aircraft study: Wind tunnel tests of a full scale hingeless prop/rotor designed for the Boeing Model 222 tilt rotor aircraft

NASA Technical Reports Server (NTRS)

Magee, J. P.; Alexander, H. R.

1973-01-01

The rotor system designed for the Boeing Model 222 tilt rotor aircraft is a soft-in-plane hingeless rotor design, 26 feet in diameter. This rotor has completed two test programs in the NASA Ames 40' X 80' wind tunnel. The first test was a windmilling rotor test on two dynamic wing test stands. The rotor was tested up to an advance ratio equivalence of 400 knots. The second test used the NASA powered propeller test rig and data were obtained in hover, transition and low speed cruise flight. Test data were obtained in the areas of wing-rotor dynamics, rotor loads, stability and control, feedback controls, and performance to meet the test objectives. These data are presented.

The development of an augmentor wing jet STOL research aircraft (modified C-8A). Volume 2: Analysis of contractor's flight test

NASA Technical Reports Server (NTRS)

Skavdahl, H.; Patterson, D. H.

1972-01-01

The initial flight test phase of the modified C-8A airplane was conducted. The primary objective of the testing was to establish the basic airworthiness of the research vehicle. This included verification of the structural design and evaluation of the aircraft's systems. Only a minimum amount of performance testing was scheduled; this has been used to provide a preliminary indication of the airplane's performance and flight characteristics for future flight planning. The testing included flutter and loads investigations up to the maximum design speed. The operational characteristics of all systems were assessed including hydraulics, environmental control system, air ducts, the vectoring conical nozzles, and the stability augmentation system (SAS). Approaches to stall were made at three primary flap settings: up, 30 deg and 65 deg, but full stalls were not scheduled. Minimum control speeds and maneuver margins were checked. All takeoffs and landings were conventional, and STOL performance was not scheduled during this phase of the evaluation.

V/STOL aircraft and method

DOEpatents

Owens, P.R.

1997-11-18

Aircraft apparatus and method capable of V/STOL (vertical, short takeoff and landing) in addition to conventional flight are disclosed. For V/STOL operation, induced lift is provided by blowing air over the upper surface of each wing through a duct installed near the leading edge. Intake air is supplied to the blowing fan through a duct installed near the trailing edge, thus providing suction as well as blowing. Two fans in series are required. The engine provides power not only to the propeller but also to a transmission which provides power to the pulleys driving the belt-driven fans. 10 figs.

Noise tests of a model engine-over-the-wing STOL configuration using a multijet nozzle with deflector

NASA Technical Reports Server (NTRS)

Olsen, W. A.; Friedman, R.

1973-01-01

Noise data were obtained with a small scale model stationary STOL configuration that used an eight lobe mixer nozzle with deflector mounted above a 32-cm-chord wing section. The factors varied to determine their effect upon the noise were wing flap angle, nozzle shape, nozzle location, deflector configuration, and jet velocity. The noise from the mixer nozzle model was compared to the noise from a model using a circular nozzle of the same area. The mixer nozzle model was quieter at the low to middle frequencies, while the circular nozzle was quieter at high frequencies. The perceived noise level (PNL) was calculated for an aircraft 10 times larger than the model. The PNL at 500 feet for the mixer nozzle turned out to be within 1 db of the PNL for the circular nozzle. For some configurations at highly directional broadband noise, which could be eliminated by changes in nozzle and/or deflector location, occurred below the wing.

Quiet Cruise Efficient Short Take-off and Landing Subsonic Transport System

NASA Technical Reports Server (NTRS)

Kawai, Ron

2008-01-01

This NASA funded study conceived a revolutionary airplane concept to enable future traffic growth by using regional air space. This requires a very quiet airplane with STOL capability. Starting with a Blended Wing Body that is cruise efficient with inherent low noise characteristics from forward noise shielding and void of aft downward noise reflections, integration of embedded distributed propulsion enables incorporation of the revolutionary concept for jet noise shielding. Embedded distributed propulsion also enables incorporation of a fan bleed internally blown flap for quiet powered lift. The powered lift provides STOL capability for operation at regional airports with rapid take-off and descent to further reduce flyover noise. This study focused on configuring the total engine noise shielding STOL concept with a BWB airplane using the Boeing Phantom Works WingMOD multidisciplinary optimization code to define a planform that is pitch controllable. The configuration was then sized and mission data developed to enable NASA to assess the flyover and sideline noise. The foundational technologies needed are identified including military dual use benefits.

High angle-of-attack aerodynamics of a strake-canard-wing V/STOL fighter configuration

NASA Technical Reports Server (NTRS)

Durston, D. A.; Schreiner, J. A.

1983-01-01

High angle-of-attack aerodynamic data are analyzed for a strake-canard-wing V/STOL fighter configuration. The configuration represents a twin-engine supersonic V/STOL fighter aircraft which uses four longitudinal thrust-augmenting ejectors to provide vertical lift. The data were obtained in tests of a 9.39 percent scale model of the configuration in the NASA Ames 12-Foot Pressure Wind Tunnel, at a Mach number of 0.2. Trimmed aerodynamic characteristics, longitudinal control power, longitudinal and lateral/directional stability, and effects of alternate strake and canard configurations are analyzed. The configuration could not be trimmed (power-off) above 12 deg angle of attack because of the limited pitch control power and the high degree of longitudinal instability (28 percent) at this Mach number. Aerodynamic center location was found to be controllable by varying strake size and canard location without significantly affecting lift and drag. These configuration variations had relatively little effect on the lateral/directional stability up to 10 deg angle of attack.

Fundamental limitations on V/STOL terminal guidance due to aircraft characteristics

NASA Technical Reports Server (NTRS)

Wolkovitch, J.; Lamont, C. W.; Lochtie, D. W.

1971-01-01

A review is given of limitations on approach flight paths of V/STOL aircraft, including limits on descent angle due to maximum drag/lift ratio. A method of calculating maximum drag/lift ratio of tilt-wing and deflected slipstream aircraft is presented. Derivatives and transfer functions for the CL-84 tilt-wing and X-22A tilt-duct aircraft are presented. For the unaugmented CL-84 in steep descents the transfer function relating descent angle to thrust contains a right-half plane zero. Using optimal control theory, it is shown that this zero causes a serious degradation in the accuracy with which steep flight paths can be followed in the presence of gusts.

Thrust reverser design studies for an over-the-wing STOL transport

NASA Technical Reports Server (NTRS)

Ammer, R. C.; Sowers, H. D.

1977-01-01

Aerodynamic and acoustics analytical studies were conducted to evaluate three thrust reverser designs for potential use on commercial over-the-wing STOL transports. The concepts were: (1) integral D nozzle/target reverser, (2) integral D nozzle/top arc cascade reverser, and (3) post exit target reverser integral with wing. Aerodynamic flowpaths and kinematic arrangements for each concept were established to provide a 50% thrust reversal capability. Analytical aircraft stopping distance/noise trade studies conducted concurrently with flow path design showed that these high efficiency reverser concepts are employed at substantially reduced power settings to meet noise goals of 100 PNdB on a 152.4 m sideline and still meet 609.6 m landing runway length requirements. From an overall installation standpoint, only the integral D nozzle/target reverser concept was found to penalize nacelle cruise performance; for this concept a larger nacelle diameter was required to match engine cycle effective area demand in reverse thrust.

Effect of crossflow velocity on VTOL lift fan blade passing frequency noise generation

NASA Technical Reports Server (NTRS)

Stimpert, D. L.

1973-01-01

Analysis of noise measurements taken during tests of a remote lift fan wing installation, a V/STOL model transport with both lift and lift/cruise fans, and XV5B research aircraft flight tests has indicated a definite increase in pure tone sound pressure level due to crossflow over the face of the life fans. The fan-in-wing and V/STOL model transport tests were conducted in the NASA Ames 40 ft. by 80 ft. wing tunnel and the XV5B flight tests at Moffett Field. Increases up to 10 db were observed for the lift fan installation tested at crossflow to fan tip velocity ratios up to 0.25. Cruise fan noise levels were found to be unaffected by the external flow. The noise level increase was shown to be related to an increase in fan distortion levels.

Wind Tunnel Testing of Powered Lift, All-Wing STOL Model

NASA Technical Reports Server (NTRS)

Collins, Scott W.; Westra, Bryan W.; Lin, John C.; Jones, Gregory S.; Zeune, Cal H.

2008-01-01

Short take-off and landing (STOL) systems can offer significant capabilities to warfighters and, for civil operators thriving on maximizing efficiencies they can improve airspace use while containing noise within airport environments. In order to provide data for next generation systems, a wind tunnel test of an all-wing cruise efficient, short take-off and landing (CE STOL) configuration was conducted in the National Aeronautics and Space Administration (NASA) Langley Research Center (LaRC) 14- by 22-foot Subsonic Wind Tunnel. The test s purpose was to mature the aerodynamic aspects of an integrated powered lift system within an advanced mobility configuration capable of CE STOL. The full-span model made use of steady flap blowing and a lifting centerbody to achieve high lift coefficients. The test occurred during April through June of 2007 and included objectives for advancing the state-of-the-art of powered lift testing through gathering force and moment data, on-body pressure data, and off-body flow field measurements during automatically controlled blowing conditions. Data were obtained for variations in model configuration, angles of attack and sideslip, blowing coefficient, and height above ground. The database produced by this effort is being used to advance design techniques and computational tools for developing systems with integrated powered lift technologies.

STOL Characteristics of a Propeller-Driven, Aspect-Ratio-10, Straight-Wing Airplane with Boundary-Layer Control Flaps, as Estimated from Large-Scale Wind-Tunnel Tests

NASA Technical Reports Server (NTRS)

Weiberg, James A; Holzhauser, Curt A.

1961-01-01

A study is presented of the improvements in take-off and landing distances possible with a conventional propeller-driven transport-type airplane when the available lift is increased by propeller slipstream effects and by very effective trailing-edge flaps and ailerons. This study is based on wind-tunnel tests of a 45-foot span, powered model, with BLC on the trailing-edge flaps and controls. The data were applied to an assumed airplane with four propellers and a wing loading of 50 pounds per square foot. Also included is an examination of the stability and control problems that may result in the landing and take-off speed range of such a vehicle. The results indicated that the landing and take-off distances could be more than halved by the use of highly effective flaps in combination with large amounts of engine power to augment lift (STOL). At the lowest speeds considered (about 50 knots), adequate longitudinal stability was obtained but the lateral and directional stability were unsatisfactory. At these low speeds, the conventional aerodynamic control surfaces may not be able to cope with the forces and moments produced by symmetric, as well as asymmetric, engine operation. This problem was alleviated by BLC applied to the control surfaces.

Tactical STOL moment balance through innovative configuration technology

NASA Technical Reports Server (NTRS)

Eckard, G. J.; Sutton, R. C.; Poth, G. E.

1981-01-01

Innovative and conventional thrust vectoring moment balance mechanisms, as applied to advanced tactical fighters, are examined. The innovative mechanisms include thrust line translation, life line translation, and auxiliary power control; the conventional mechanisms under investigation are horizontal tails, canards, and variable sweep wings. These mechanisms are tested for their ability to provide negative static margins for landing approach or relocation of the vectored thrust line nearer the aircraft's center of gravity. The net pitching moment due to wing, flaps, and vectored thrust lift would then be small, making possible beneficial trim forces from small trimming devices. These innovative mechanisms are, however, possibly heavy and must be evaluated on their complexity, reliability, maintainability, and STOL capabilities. Several candidate fighter configurations are compared and evaluated.

Rotary-wing aircraft systems for the short-haul market

NASA Technical Reports Server (NTRS)

Magee, J. P.; Clark, R. D.; Giulianetti, D.

1975-01-01

This paper describes preliminary designs of tilt-rotor and tandem-rotor helicopter V/STOL aircraft for the 1958 short-haul market. These designs include a tilt-rotor aircraft designed for STOL-only operation. The baseline designs are presented with technological and cost data. The impact of noise and ride qualities on aircraft size and cost, and on passenger acceptance are discussed. The results of the study are compared against competitive alternatives in air transportation.

Effect of rotor wake on aerodynamic characteristics of a 1/6 scale model of the rotor systems research aircraft. [in the Langley V/STOL tunnel

NASA Technical Reports Server (NTRS)

Mineck, R. E.

1977-01-01

Tests were conducted in the Langley V/STOL tunnel to determine the effect of the main-rotor wake on the aerodynamic characteristics of the rotor systems research aircraft. A 1/6-scale model with a 4-blade articulated rotor was used to determine the effect of the rotor wake for the compound configuration. Data were obtained over a range of angles of attack, angles of sideslip, auxiliary engine thrusts, rotor collective pitch angles, and rotor tip-path plane angles for several main-rotor advance ratios. Separate results are presented for the forces and moments on the airframe, the wing, and the tail. An analysis of the test data indicates significant changes in the aerodynamic characteristics. The rotor wake increases the longitudinal static stability, the effective dihedral, and the lateral static stability of the airframe. The rotor induces a downwash on the wing. This downwash decreases the wing lift and increases the drag. The asymmetrical rotor wake induces a differential lift across the wing and a subsequent rolling moment. These rotor induced effects on the wing become smaller with increasing forward speed.

Evaluation of pressure and thermal data from a wind tunnel test of a large-scale, powered, STOL fighter model

NASA Technical Reports Server (NTRS)

Howell, G. A.; Crosthwait, E. L.; Witte, M. C.

1981-01-01

A STOL fighter model employing the vectored-engine-over wing concept was tested at low speeds in the NASA/Ames 40 by 80-foot wind tunnel. The model, approximately 0.75 scale of an operational fighter, was powered by two General Electric J-97 turbojet engines. Limited pressure and thermal instrumentation were provided to measure power effects (chordwise and spanwise blowing) and control-surface-deflection effects. An indepth study of the pressure and temperature data revealed many flow field features - the foremost being wing and canard leading-edge vortices. These vortices delineated regions of attached and separated flow, and their movements were often keys to an understanding of flow field changes caused by power and control-surface variations. Chordwise blowing increased wing lift and caused a modest aft shift in the center of pressure. The induced effects of chordwise blowing extended forward to the canard and significantly increased the canard lift when the surface was stalled. Spanwise blowing effectively enhanced the wing leading-edge vortex, thereby increasing lift and causing a forward shift in the center of pressure.

A look at V/STOL for business aircraft.

NASA Technical Reports Server (NTRS)

Feistel, T. W.; Stewart, E. C.; Gerdes, R. M.; Smith, K. R., Jr.

1972-01-01

Attempt to ascertain the economic viability of the V/STOL capability for business aircraft and the manner in which this viability depends on the aircraft concept. A cost-benefit analysis is presented which indicates that a VTOL business aircraft would be more viable economically than a contemporary turbine-powered business aircraft. The combination of traveler's time value and trip distance for which each aircraft dominates is shown. The significance of disk loading in V/STOL concept application is discussed, and preliminary design configuration studies for three different business-aircraft-sized V/STOLs, using three concepts covering a range of disk loading, are presented as examples. Finally, a discussion of operational aspects of interest to future users of V/STOL business aircraft is presented which centers around the requirements for routine IFR terminal-area operations.

Study of aerodynamic technology for single-cruise-engine V/STOL fighter/attack aircraft

NASA Technical Reports Server (NTRS)

Hess, J. R.; Bear, R. L.

1982-01-01

A viable, single engine, supersonic V/STOL fighter/attack aircraft concept was defined. This vectored thrust, canard wing configuration utilizes an advanced technology separated flow engine with fan stream burning. The aerodynamic characteristics of this configuration were estimated and performance evaluated. Significant aerodynamic and aerodynamic propulsion interaction uncertainties requiring additional investigation were identified. A wind tunnel model concept and test program to resolve these uncertainties and validate the aerodynamic prediction methods were defined.

Static, noise, and transition tests of a combined-surface-blowing V/STOL lift/propulsion system

NASA Technical Reports Server (NTRS)

Schoen, A. H.; Kolesar, C. E.; Schaeffer, E. G.

1977-01-01

Efficient thrust vectoring and high levels of circulatory lift were obtained in tests of a half model V/STOL airplane by using a type of externally blown jet flap in which the jet exhaust from wing-mounted cruise fans is directed over both upper and lower surfaces of a flapped wing. Approximately 90% thrust recovery with 87 deg of thrust vectoring was achieved under static conditions using 89 deg of trailing edge flap deflection. The approximately 10% loss appears to be associated primarily with pressure losses due to the flap brackets or slot entries. The jet induced lift was shown to be 55% of the theoretical value for a fullspan jet-flapped wing, even though only 27.5% of the wingspan was immersed in the jet. Steady rate of descent capability in excess of 1,000 feet per minute is predicted. The possibility of significant aerodynamic-noise cancelling when blowing over both surfaces at high velocities is indicated.

Static noise tests on augmentor wing jet STOL research aircraft (C8A Buffalo)

NASA Technical Reports Server (NTRS)

Marrs, C. C.; Harkonen, D. L.; Okeefe, J. V.

1974-01-01

Results are presented for full scale ground static acoustic tests of over-area conical nozzles and a lobe nozzle installed on the Augmentor Wing Jet STOL Research Aircraft, a modified C8A Buffalo. The noise levels and spectrums of the test nozzles are compared against those of the standard conical nozzle now in use on the aircraft. Acoustic evaluations at 152 m (500 ft), 304 m (1000 ft), and 1216 m (4000 ft) are made at various engine power settings with the emphasis on approach and takeoff power. Appendix A contains the test log and propulsion calculations. Appendix B gives the original test plan, which was closely adhered to during the test. Appendix C describes the acoustic data recording and reduction systems, with calibration details.

Fixed base simulator study of an externally blown flap STOL transport airplane during approach and landing

NASA Technical Reports Server (NTRS)

Grantham, W. D.; Nguyen, L. T.; Patton, J. M., Jr.; Deal, P. L.; Champine, R. A.; Carter, C. R.

1972-01-01

A fixed-base simulator study was conducted to determine the flight characteristics of a representative STOL transport having a high wing and equipped with an external-flow jet flap in combination with four high-bypass-ratio fan-jet engines during the approach and landing. Real-time digital simulation techniques were used. The computer was programed with equations of motion for six degrees of freedom and the aerodynamic inputs were based on measured wind-tunnel data. A visual display of a STOL airport was provided for simulation of the flare and touchdown characteristics. The primary piloting task was an instrument approach to a breakout at a 200-ft ceiling with a visual landing.

A manual control theory analysis of vertical situation displays for STOL aircraft

NASA Technical Reports Server (NTRS)

Baron, S.; Levison, W. H.

1973-01-01

Pilot-vehicle-display systems theory is applied to the analysis of proposed vertical situation displays for manual control in approach-to-landing of a STOL aircraft. The effects of display variables on pilot workload and on total closed-loop system performance was calculated using an optimal-control model for the human operator. The steep approach of an augmentor wing jet STOL aircraft was analyzed. Both random turbulence and mean-wind shears were considered. Linearized perturbation equations were used to describe longitudinal and lateral dynamics of the aircraft. The basic display configuration was one that abstracted the essential status information (including glide-slope and localizer errors) of an EADI display. Proposed flight director displays for both longitudinal and lateral control were also investigated.

Experimental measurements of motion cue effects on STOL approach tasks

NASA Technical Reports Server (NTRS)

Ringland, R. F.; Stapleford, R. L.

1972-01-01

An experimental program to investigate the effects of motion cues on STOL approach is presented. The simulator used was the Six-Degrees-of-Freedom Motion Simulator (S.01) at Ames Research Center of NASA which has ?2.7 m travel longitudinally and laterally and ?2.5 m travel vertically. Three major experiments, characterized as tracking tasks, were conducted under fixed and moving base conditions: (1) A simulated IFR approach of the Augmentor Wing Jet STOL Research Aircraft (AWJSRA), (2) a simulated VFR task with the same aircraft, and (3) a single-axis task having only linear acceleration as the motion cue. Tracking performance was measured in terms of the variances of several motion variables, pilot vehicle describing functions, and pilot commentary.

The XFV-12A Thrust-Augmented Wing (TAW) prototype aircraft

NASA Technical Reports Server (NTRS)

Murphy, R.; Lewis, E. L.

1979-01-01

The XFV-12A, a unique V/STOL technology prototype aircraft being developed for the Navy, is described. The innovative design features a thrust augmented wing and a canard ejector. Structural, functional, and control test performances are discussed. Static tether test results are also discussed. Assessment of test results are given along with projections for future modification areas.

The development of an augmentor wing jet STOL research airplane (modified C-8A). Volume 1: Summary

NASA Technical Reports Server (NTRS)

Ashleman, R. H.; Kavdahl, H.

1972-01-01

A project to develop an experimental aircraft for use as an inflight demonstrator of the augmentor wing, short takeoff concept is discussed. The required modifications were made on a de Havilland C-8A aircraft. The modifications to the aircraft are explained and the performance of the modified aircraft is reported.

The FM-007: An advanced jet commuter for HUB to spoke transportation

NASA Technical Reports Server (NTRS)

Blouke, Peter Scott; Engel, George Bryan; Fordham, Kari Suzanne; Layne, Steven James; Moore, Joel David; Shaver, Frederick Martin; Thornton, Douglas Hershal, Jr.

1991-01-01

Due to the increasing need for new commuter aircraft, the FM-007 is proposed, a technologically advanced jet propelled short takeoff and landing (STOL) airplane. The proposed commuter is designed for hub to spoke air travel. In order to reduce drag, natural laminar flow technology is integrated into the design using the natural laminar flow airfoil section for the wing. A three lifting surface configuration provides for more efficient cruise flight. This unique design includes a small forward wing (canard), a rear mounted high aspect ratio main wing, and a small horizontal stabilizer high atop the vertical tail. These three surfaces act together to reduce drag by minimizing the downward force the horizontal stabilizer has to account for due to the nose down pitching moment. Commuter aircraft must also incorporate passenger comfort. This is achieved by providing a spacious pressurized cabin with a large galley and reduced cabin noise due to incorporation of noise reduction gear. A basic oval design is adopted, as opposed to a circular design in order to allow for the seating of five passengers abreast. To get STOL capability, an over the wing blown flap is used using a Rolls Royce Tay series engine.

Recent progress in VSTOL technology

NASA Technical Reports Server (NTRS)

Roberts, L.; Deckert, W. R.

1982-01-01

Progress in vertical and short takeoff and landing (V/STOL) aircraft technology, in particular, during the 1970 to 1980 period at Ames Research Center is discussed. Although only two kinds of V/STOL aircraft (the helicopter and the British direct lift Harrier) have achieved operational maturity, understanding of the technology has vastly improved during this 10 year period. To pursue an aggressive R and D program at a reasonable cost, it was decided to conduct extensive large scale testing in wind tunnel and flight simulation facilities, to develop low cost research aircraft using modified airframes or engines, and to involve other agencies and industry contractors in joint technical and funding arrangements. The STOL investigations include exploring STOL performance using the rotating cylinder flap concept, the augmentor wing, upon initiation of the Quiet Short Haul Research Aircraft program, the upper surface blown flap concept. The VTOL investigations were conducted using a tilt rotor aircraft, resulting in the XV-15 tilt rotor research aircraft. Direct jet lift is now being considered for application to future supersonic fighter aircraft.

Aerodynamic characteristics of a six-jet V/STOL configuration with four swing-out lift jets in the transition speed range

NASA Technical Reports Server (NTRS)

Carter, A. W.

1970-01-01

A wind-tunnel investigation has been made of the longitudinal aerodynamic characteristics and jet-interference effects of a model of a jet V/STOL variable-sweep fighter airplane that employs four direct-lift engines which swing out from the fuselage and two lift-cruise engines located in the rear part of the fuselage. Data were obtained with two wing areas for various forward speeds and power conditions in the transition speed range. The data are presented without analysis or discussion.

A progress report on the development of an augmentor wing jet STOL research aircraft.

NASA Technical Reports Server (NTRS)

Quigley, H. C.; Sinclair, S. R. M.; Nark, T. C., Jr.; O'Keefe, J. V.

1971-01-01

The development of the aircraft has progressed to the point where the design of the modifications to the de Havilland C-8A Buffalo is complete and the engines are being tested. The predicted performance shows that the aircraft will be able to take off and land in less than 1500 ft. Simulation studies indicate that the handling qualities of the aircraft, with stability augmentation, will be acceptable for STOL research missions. Special techniques were required, however, for flight path control and transition from cruise to landing configuration .

Application of Nonlinear Systems Inverses to Automatic Flight Control Design: System Concepts and Flight Evaluations

NASA Technical Reports Server (NTRS)

Meyer, G.; Cicolani, L.

1981-01-01

A practical method for the design of automatic flight control systems for aircraft with complex characteristics and operational requirements, such as the powered lift STOL and V/STOL configurations, is presented. The method is effective for a large class of dynamic systems requiring multi-axis control which have highly coupled nonlinearities, redundant controls, and complex multidimensional operational envelopes. It exploits the concept of inverse dynamic systems, and an algorithm for the construction of inverse is given. A hierarchic structure for the total control logic with inverses is presented. The method is illustrated with an application to the Augmentor Wing Jet STOL Research Aircraft equipped with a digital flight control system. Results of flight evaluation of the control concept on this aircraft are presented.

Comparison of model and flight test data for an augmented jet flap STOL research aircraft

NASA Technical Reports Server (NTRS)

Cook, W. L.; Whittley, D. C.

1975-01-01

Aerodynamic design data for the Augmented Jet Flap STOL Research Aircraft or commonly known as the Augmentor-Wing Jet-STOL Research Aircraft was based on results of tests carried out on a large scale research model in the NASA Ames 40- by 80-Foot Wind Tunnel. Since the model differs in some respects from the aircraft, precise correlation between tunnel and flight test is not expected, however the major areas of confidence derived from the wind tunnel tests are delineated, and for the most part, tunnel results compare favorably with flight experience. In some areas the model tests were known to be nonrepresentative so that a degree of uncertainty remained: these areas of greater uncertainty are identified, and discussed in the light of subsequent flight tests.

Laser velocimeter survey about a NACA 0012 wing at low angles of attack

NASA Technical Reports Server (NTRS)

Hoad, D. R.; Meyers, J. F.; Young, W. H., Jr.; Hepner, T. E.

1978-01-01

An investigation was conducted in the Langley V/STOL tunnel with a laser velocimeter to obtain measurements of airflow velocities about a wing at low angles of attack. The applicability of the laser velocimeter technique for this purpose in the V/STOL tunnel was demonstrated in this investigation with measurement precision bias calculated at -1.33 percent to 0.91 percent and a random uncertainty calculated at + or - 0.47 percent. Free stream measurements were obtained with this device and compared with velocity calculations from pitot static probe data taken near the laser velocimeter measurement location. The two measurements were in agreement to within 1 percent. Velocity measurement results about the centerline at 0.6 degrees angle of attack were typically those expected. At 4.75 degrees, the velocity measurements indicated that a short laminar separation bubble existed near the leading edge with an oscillating shear layer.

Effect of twist and camber on the low-speed aerodynamic characteristics of a powered close-coupled wing-canard configuration

NASA Technical Reports Server (NTRS)

Paulson, J. W., Jr.; Thomas, J. L.

1978-01-01

A series of wind-tunnel tests were conducted in a V/STOL tunnel to determine the low-speed longitudinal aerodynamic characteristics of a powered close-coupled wing/canard fighter configuration. The data was obtained for a high angle-of-attack maneuvering configuration and a takeoff and landing configuration. The data presented in tabulated form are intended for reference purposes.

NASA aeronautics

NASA Technical Reports Server (NTRS)

Anderton, D. A.

1982-01-01

Aeronautical research programs are discussed in relation to research methods and the status of the programs. The energy efficient aircraft, STOL aircraft and general aviation aircraft are considered. Aerodynamic concepts, rotary wing aircraft, aircraft safety, noise reduction, and aircraft configurations are among the topics included.

Flight evaluation of configuration management system concepts during transition to the landing approach for a powered-lift STOL aircraft

NASA Technical Reports Server (NTRS)

Franklin, J. A.; Innis, R. C.

1980-01-01

Flight experiments were conducted to evaluate two control concepts for configuration management during the transition to landing approach for a powered-lift STOL aircraft. NASA Ames' augmentor wing research aircraft was used in the program. Transitions from nominal level-flight configurations at terminal area pattern speeds were conducted along straight and curved descending flightpaths. Stabilization and command augmentation for attitude and airspeed control were used in conjunction with a three-cue flight director that presented commands for pitch, roll, and throttle controls. A prototype microwave system provided landing guidance. Results of these flight experiments indicate that these configuration management concepts permit the successful performance of transitions and approaches along curved paths by powered-lift STOL aircraft. Flight director guidance was essential to accomplish the task.

V/STOL Aerodynamics

DTIC Science & Technology

1974-10-01

jet exhaust, m (ft) Ro radius of engine exhaust, m (ft) 1. INTRODUCTION free deg S wing area, m2 (ft2) t time, see T Thrust, N (lb) u...dimensional potential flow method to lift prediction for a wing with internally blown flaps is described. INTRODUCTION The objectives of this paper are...twofold. The first is to provide an introduction to this session on research into the aerodynamics of powered high lift systems. This will be

Propulsion simulation test technique for V/STOL configurations

NASA Technical Reports Server (NTRS)

Bailey, R. O.; Smith, S. C.; Bustie, J. B.

1983-01-01

Ames Research Center is developing the technology for turbine-powered jet engine simulators so that airframe/propulsion system interactions on V/STOL fighter aircraft and other highly integrated configurations can be studied. This paper describes the status of the compact multimission aircraft propulsion simulator (CMAPS) technology. Three CMAPS units have accumulated a total of 340 hr during approximately 1-1/2 yr of static and wind-tunnel testing. A wind-tunnel test of a twin-engine CMAPS-equipped close-coupled canard-wing V/STOL model configuration with nonaxisymmetric nozzles was recently completed. During this test approximately 140 total hours were logged on two CMAPS units, indicating that the rotating machinery is reliable and that the CMAPS and associated control system provide a usable test tool. However, additional development is required to correct a drive manifold O-ring problem that limits the engine-pressure-ratio (EPR) to approximately 3.5.

Design criteria for flightpath and airspeed control for the approach and landing of STOL aircraft

NASA Technical Reports Server (NTRS)

Franklin, J. A.; Innis, R. C.; Hardy, G. H.; Stephenson, J. D.

1982-01-01

A flight research program was conducted to assess requirements for flightpath and airspeed control for glide-slope tracking during a precision approach and for flare control, particularly as applied to powered-lift, short takeoff and landing (STOL) aircraft. Ames Research Center's Augmentor Wing Research Aircraft was used to fly approaches on a 7.5 deg glide slope to landings on a 30 X 518 m (100 X 1700 ft) STOL runway. The dominant aircraft response characteristics determined were flightpath overshoot, flightpath-airspeed coupling, and initial flightpath response time. The significant contribution to control of the landing flare using pitch attitude was the short-term flightpath response. The limiting condition for initial flightpath response time for flare control with thrust was also identified. It is possible to define flying-qualities design criteria for glide-slope and flare control based on the aforementioned response characteristics.

Effects of Blowing Spanwise from the Tips of Low-Aspect Ratio Wings of Varying Taper Ratio, with Application to Improving STOL Capability of Fighter Aircraft.

DTIC Science & Technology

1983-02-01

aspect ratio is relatively small. Brooks (ref. 1) worked with rectangular fins of 0.62 and 1.24 aspect ratio in a water medium and showed very large ...airflow rates. Lloyd (ref. 3) worked with an aspect ratio 2.0 rectangular wing using a very wide range of jet momentum coefficient; his results were in...D-A1i35 688 EFFECTS OF BLOWING SPANWISE FROM THE TIPS OF LOW ASPECT in, RATIO WINGS OF VA .(U) NIELSEN ENGINEERING AND RESEARCH INC MOUNTAIN VIEW CA

Flap noise measurements for STOL configurations using external upper surface blowing

NASA Technical Reports Server (NTRS)

Dorsch, R. G.; Reshotko, M.; Olsen, W. A.

1972-01-01

Screening tests of upper surface blowing on externally blown flaps configurations were conducted. Noise and turning effectiveness data were obtained with small-scale, engine-over-the-wing models. One large model was tested to determine scale effects. Nozzle types included circular, slot, D-shaped, and multilobed. Tests were made with and without flow attachment devices. For STOL applications the particular multilobed mixer and the D-shaped nozzles tested were found to offer little or no noise advantage over the round convergent nozzle. High aspect ratio slot nozzles provided the quietest configurations. In general, upper surface blowing was quieter than lower surface blowing for equivalent EBF models.

Pressure distribution on a 1- by 3-meter semispan wing at sweep angles from 0 deg to 40 deg in subsonic flow

NASA Technical Reports Server (NTRS)

Yip, L. P.; Shubert, G. L.

1976-01-01

A 1- by 3-meter semispan wing of taper ratio 1.0 with NACA 0012 airfoil section contours was tested in the Langley V/STOL tunnel to measure the pressure distribution at five sweep angles, 0 deg, 10 deg, 20 deg, 30 deg, and 40 deg, through an angle-of-attack range from -6 deg to 20 deg. The pressure data are presented as plots of pressure coefficients at each static-pressure tap location on the wing. Flow visualization wing-tuft photographs are also presented for a wing of 40 deg sweep. A comparison between theory and experiment using two inviscid theories and a viscous theory shows good agreement for pressure distributions, normal forces, and pitching moments for the wing at 0 deg sweep.

Fan and wing force data from wind tunnel investigation of a 0.38 meter (15 inch) diameter VTOL model lift fan installed in a two dimensional wing

NASA Technical Reports Server (NTRS)

Yuska, J. A.; Diedrich, J. H.

1972-01-01

Test data are presented for a 38-cm (15-in.) diameter, 1.28 pressure ratio model VTOL lift fan installed in a two-dimensional wing and tested in a 2.74-by 4.58-meter (9-by 15-ft)V/STOL wind tunnel. Tests were run with and without exit louvers over a wide range of crossflow velocities and wing angle of attack. Tests were also performed with annular-inlet vanes, inlet bell-mouth surface disconuities, and fences to induce fan windmilling. Data are presented on the axial force of the fan assembly and overall wing forces and moments as measured on force balances for various static and crossflow test conditions. Midspan wing surface pressure coefficient data are also given.

Wind-tunnel investigation of a large-scale VTOL aircraft model with wing root and wing thrust augmentors. [Ames 40 by 80 foot wind tunnel

NASA Technical Reports Server (NTRS)

Aoyagi, K.; Aiken, T. N.

1979-01-01

Tests were conducted in the Ames 40 by 80 foot wind tunnel to determine the aerodynamic characteristics of a large-scale V/STOL aircraft model with thrust augmentors. The model had a double-delta wing of aspect ratio 1.65 with augmentors located in the wing root and the wing trailing edge. The supply air for the augmentor primary nozzles was provided by the YJ-97 turbojet engine. The airflow was apportioned approximately 74 percent to the wing root augmentor and 24 percent to wing augmentor. Results were obtained at several trailing-edge flap deflections with the nozzle jet-momentum coefficients ranging from 0 to 7.9. Three-component longitudinal data are presented with the agumentor operating with and without the horizontal tail. A limited amount of six component data are also presented.

V/STOL tilt rotor aircraft study. Volume 6: Preliminary design of a composite wing for tilt rotor research aircraft

NASA Technical Reports Server (NTRS)

Soule, V. A.; Badri-Nath, Y.

1973-01-01

The results of a study of the use of composite materials in the wing of a tilt rotor aircraft are presented. An all-metal tilt rotor aircraft was first defined to provide a basis for comparing composite with metal structure. A configuration study was then done in which the wing of the metal aircraft was replaced with composite wings of varying chord and thickness ratio. The results of this study defined the design and performance benefits obtainable with composite materials. Based on these results the aircraft was resized with a composite wing to extend the weight savings to other parts of the aircraft. A wing design was then selected for detailed structural analysis. A development plan including costs and schedules to develop this wing and incorporate it into a proposed flight research tilt rotor vehicle has been devised.

Civil applications of high-speed rotorcraft and powered-lift aircraft configurations

NASA Technical Reports Server (NTRS)

Albers, James A.; Zuk, John

1987-01-01

Advanced subsonic vertical and short takeoff and landing (V/STOL) aircraft configurations offer new transportation options for civil applications. Described is a range of vehicles from low-disk to high-disk loading aircraft, including high-speed rotorcraft, V/STOL aircraft, and short takeoff and landing (STOL) aircraft. The status and advantages of the various configurations are described. Some of these show promise for relieving congestion in high population-density regions and providing transportation opportunities for low population-density regions.

V/STOL aircraft configurations and opportunities in the Pacific Basin

NASA Technical Reports Server (NTRS)

Albers, James A.; Zuk, John

1987-01-01

Advanced aircraft configurations offer new transportation options for the Pacific Basin. Described is a range of vehicles from low-disk to high-disk loading aircraft, including high-speed rotorcraft, subsonic vertical and short takeoff and landing (V/STOL) aircraft, and subsonic short takeoff and landing (STOL) aircraft. The status and advantages of the various configurations are described. Some of these show promise for satisfying many of the transportation requirements of the Pacific Basin; as such, they could revolutionize short-haul transportation in that region.

Effect of spanwise blowing on leading-edge vortex bursting of a highly swept aspect ratio 1.18 delta wing

NASA Technical Reports Server (NTRS)

Scantling, W. L.; Gloss, B. B.

1974-01-01

An investigation was conducted in the Langley 1/8-scale V/STOL model tunnel on a semispan delta wing with a leading-edge sweep of 74 deg, to determine the effectiveness of various locations of upper surface and reflection plane blowing on leading-edge vortex bursting. Constant area nozzles were located on the wing upper surface along a ray swept 79 deg, which was beneath the leading-edge vortex core. The bursting and reformation of the leading-edge vortex was viewed by injecting helium into the vortex core, and employing a schlieren system.

An engineering optimization method with application to STOL-aircraft approach and landing trajectories

NASA Technical Reports Server (NTRS)

Jacob, H. G.

1972-01-01

An optimization method has been developed that computes the optimal open loop inputs for a dynamical system by observing only its output. The method reduces to static optimization by expressing the inputs as series of functions with parameters to be optimized. Since the method is not concerned with the details of the dynamical system to be optimized, it works for both linear and nonlinear systems. The method and the application to optimizing longitudinal landing paths for a STOL aircraft with an augmented wing are discussed. Noise, fuel, time, and path deviation minimizations are considered with and without angle of attack, acceleration excursion, flight path, endpoint, and other constraints.

Design of a powered elevator control system. [powered elevator system for modified C-8A aircraft for STOL operation

NASA Technical Reports Server (NTRS)

Glende, W. L. B.

1974-01-01

The design, fabrication and flight testing of a powered elevator system for the Augmentor Wing Jet STOL Research Aircraft (AWJSRA or Mod C-8A) are discussed. The system replaces a manual spring tab elevator control system that was unsatisfactory in the STOL flight regime. Pitch control in the AWJSRA is by means of a single elevator control surface. The elevator is used for both maneuver and trim control as the stabilizer is fixed. A fully powered, irreversible flight control system powered by dual hydraulic sources was designed. The existing control columns and single mechanical cable system of the AWJSRA have been retained as has been the basic elevator surface, except that the elevator spring tab is modified into a geared balance tab. The control surface is directly actuated by a dual tandem moving body actuator. Control signals are transmitted from the elevator aft quadrant to the actuator by a linkage system that includes a limited authority series servo actuator.

Assessment at full scale of nozzle/wing geometry effects on OTW aero-acoustic characteristics. [short takeoff aircraft noise

NASA Technical Reports Server (NTRS)

Groesbeck, D.; Vonglahn, U.

1979-01-01

The effects on acoustic characteristics of nozzle type and location on a wing for STOL engine over-the-wing configurations are assessed at full scale on the basis of model-scale data. Three types of nozzle configurations are evaluated: a circular nozzle with external deflector mounted above the wing, a slot nozzle with external deflector mounted on the wing and a slot nozzle mounted on the wing. Nozzle exhaust plane locations with respect to the wing leading edge are varied from 10 to 46 percent chord (flaps retracted) with flap angles of 20 (takeoff altitude) and 60 (approach attitude). Perceived noise levels (PNL) are calculated as a function of flyover distance at 152 m altitude. From these plots, static EPNL values, defined as flyover relative noise levels, are calculated and plotted as a function of lift and thrust ratios. From such plots the acoustic benefits attributable to variations in nozzle/deflector/wing geometry at full scale are assessed for equal aerodynamic performance.

Vector thrust induced lift effects for several ejector exhaust locations on a V/STOL wind tunnel model at forward speed

NASA Technical Reports Server (NTRS)

Sharon, A. D.

1975-01-01

The results and analysis of aerodynamic force data obtained from a small scale model of a V/STOL research vehicle in a low speed wind tunnel are presented. The analysis of the data includes the evaluation of aerodynamic-propulsive lift performance when operating twin ejector nozzles with thrust deflected. Three different types of thrust deflector systems were examined: 90 deg downward deflected nozzle, 90 deg slotted nozzle with boundary layer control, and an externally blown flap configuration. Several nozzle locations were tested, including over and underwing positions. The interference lift of the nacelle and model due to jet exhaust thrust is compared and results show that 90 deg turned nozzles located over the wing (near the trailing edge) produce the largest interference lift increment for an untrimmed aircraft, and that the slotted nozzle located under the wing near the trailing edge (in conjunction with a BLC flap) gives a comparable interference lift in the trimmed condition. The externally blown flap nozzle produced the least interference lift and significantly less total lift due to jet thrust effects.

Application of variable structure system theory to aircraft flight control. [AV-8A and the Augmentor Wing Jet STOL Research Aircraft

NASA Technical Reports Server (NTRS)

Calise, A. J.; Kadushin, I.; Kramer, F.

1981-01-01

The current status of research on the application of variable structure system (VSS) theory to design aircraft flight control systems is summarized. Two aircraft types are currently being investigated: the Augmentor Wing Jet STOL Research Aircraft (AWJSRA), and AV-8A Harrier. The AWJSRA design considers automatic control of longitudinal dynamics during the landing phase. The main task for the AWJSRA is to design an automatic landing system that captures and tracks a localizer beam. The control task for the AV-8A is to track velocity commands in a hovering flight configuration. Much effort was devoted to developing computer programs that are needed to carry out VSS design in a multivariable frame work, and in becoming familiar with the dynamics and control problems associated with the aircraft types under investigation. Numerous VSS design schemes were explored, particularly for the AWJSRA. The approaches that appear best suited for these aircraft types are presented. Examples are given of the numerical results currently being generated.

Summary of Lift and Lift/Cruise Fan Powered Lift Concept Technology

NASA Technical Reports Server (NTRS)

Cook, Woodrow L.

1993-01-01

A summary is presented of some of the lift and lift/cruise fan technology including fan performance, fan stall, ground effects, ingestion and thrust loss, design tradeoffs and integration, control effectiveness and several other areas related to vertical short takeoff and landing (V/STOL) aircraft conceptual design. The various subjects addressed, while not necessarily pertinent to specific short takeoff/vertical landing (STOVL) supersonic designs being considered, are of interest to the general field of lift and lift/cruise fan aircraft designs and may be of importance in the future. The various wind tunnel and static tests reviewed are: (1) the Doak VZ-4 ducted fan, (2) the 0.57 scale model of the Bell X-22 ducted fan aircraft, (3) the Avrocar, (4) the General Electric lift/cruise fan, (5) the vertical short takeoff and landing (V/STOL) lift engine configurations related to ingestion and consequent thrust loss, (6) the XV-5 and other fan-in-wing stall consideration, (7) hybrid configurations such as lift fan and lift/cruise fan or engines, and (8) the various conceptual design studies by air-frame contractors. Other design integration problems related to small and large V/STOL transport aircraft are summarized including lessons learned during more recent conceptual design studies related to a small executive V/STOL transport aircraft.

Development of circulation control technology for powered-lift STOL aircraft

NASA Technical Reports Server (NTRS)

Englar, Robert J.

1987-01-01

The flow entraining capabilities of the Circulation Control Wing high lift system were employed to provide an even stronger STOL potential when synergistically combined with upper surface mounted engines. The resulting configurations generate very high supercirculation lift in addition to a vertical component of the pneumatically deflected engine thrust. A series of small scale wind tunnel tests and full scale static thrust deflection tests are discussed which provide a sufficient data base performance. These tests results show thrust deflections of greater than 90 deg produced pneumatically by nonmoving aerodynamic surfaces, and the ability to maintain constant high lift while varying the propulsive force from high thrust recovery required for short takeoff to high drag generation required for short low speed landings.

Simulation of decelerating landing approaches on an externally blown flap STOL transport airplane

NASA Technical Reports Server (NTRS)

Grantham, W. D.; Nguyen, L. T.; Deal, P. L.

1974-01-01

A fixed-base simulator program was conducted to define the problems and methods for solution associated with performing decelerating landing approaches on a representative STOL transport having a high wing and equipped with an external-flow jet flap in combination with four high-bypass-ratio fan-jet engines. Real-time digital simulation techniques were used. The computer was programed with equations of motion for six degrees of freedom and the aerodynamic inputs were based on measured wind-tunnel data. The pilot's task was to capture the localizer and the glide slope and to maintain them as closely as possible while decelerating from an initial airspeed of 140 knots to a final airspeed of 75 knots, while under IFR conditions.

Initial piloted simulation study of geared flap control for tilt-wing V/STOL aircraft

NASA Technical Reports Server (NTRS)

Guerrero, Lourdes M.; Corliss, Lloyd D.

1991-01-01

A simulation study of a representative tilt wing transport aircraft was conducted in 1990 on the Ames Vertical Motion Simulator. This simulation is in response to renewed interest in the tilt wing concept for use in future military and civil applications. For past tilt wing concepts, pitch control in hover and low-speed flight has required a tail rotor or reaction jets at the tail. Use of mono cyclic propellers or a geared flap have also been proposed as alternate methods for providing pitch control at low speed. The geared flap is a subject of this current study. This report describes the geared flap concept, the tilt wing aircraft, the simulation model, the simulation facility and experiment setup, the pilots' evaluation tasks and procedures, and the results obtained from the simulation experiment. The pilot evaluations and comments are also documented in the report appendix.

XC-142 Tilt Wing; 0.6 Scale Model in the 40x80 Foot Wind Tunnel at NASA Ames Research Center.

NASA Image and Video Library

1964-01-22

3/4 front right side only with Tim Wills on right and Charles Greco, mechanic. Large flaps on Variable height struts. XC-142 was a tri-service tiltwing experimental aircraft designed to investigate the operational suitability of vertical/short takeoff and landing (V/STOL) transports.

Brake control system modification, augmentor Wing Jet STOL Research Airplane (AWJSRA)

NASA Technical Reports Server (NTRS)

Amberg, R. L.; Arline, J. A.; Jenny, R. W.

1974-01-01

The braking system for a short takeoff aircraft is discussed and the deficiencies are described. The installation of a Boeing 727 aircraft brake system was made to correct the deficiencies. Tests of the modified system were conducted using an analog computer/hardware simulator. Actual performance tests were conducted and the characteristics of the system were satisfactory.

The reduction of takeoff ground roll by the application of a nose gear jump strut

NASA Technical Reports Server (NTRS)

Eppel, Joseph C.; Maisel, Martin D.; Mcclain, J. Greer; Luce, W.

1994-01-01

A series of flight tests were conducted to evaluate the reduction of takeoff ground roll distance obtainable from a rapid extension of the nose gear strut. The NASA Quiet Short-haul Research Aircraft (QSRA) used for this investigation is a transport-size short take off and landing (STOL) research vehicle with a slightly swept wing that employs the upper surface blowing (USB) concept to attain the high lift levels required for its low-speed, short-field performance. Minor modifications to the conventional nose gear assembly and the addition of a high-pressure pneumatic system and a control system provided the extendable nose gear, or jump strut, capability. The limited flight test program explored the effects of thrust-to-weight ratio, wing loading, storage tank initial pressure, and control valve open time duration on the ground roll distance. The data show that a reduction of takeoff ground roll on the order of 10 percent was achieved with the use of the jump strut, as predicted. Takeoff performance with the jump strut was also found to be essentially independent of the pneumatic supply pressure and was only slightly affected by control valve open time within the range of the parameters examined.

Understanding the Potential of Aeroelastic Couplings to Stabilize Ground and Air Resonance in a Soft-Inplane Tiltrotor

NASA Technical Reports Server (NTRS)

Howard, Anna K. T.

1999-01-01

The tiltrotor offers the best mix of hovering and cruise flight of any of the current V/STOL configurations. One possible improvement on the tiltrotors of today designs would be using a soft-inplane hingeless hub. The advantages to a soft-inplane hingeless hub range from reduced weight and maintenance to reduced vibration and loads. However, soft-inplane rotor systems are inherently in danger of the aeromechanical instabilities of ground and air resonance. Furthermore tiltrotors can be subject to whirl flutter. At least in part because of the potential for air and ground resonance in a soft-inplane rotor, the Bell XV-15, the Bell-Boeing V-22 Osprey, and the new Bell Augusta 609 have stiff-inplane, gimballed rotors which do not experience these instabilities. In order to design soft-inplane V/STOL aircraft that do not experience ground or air resonance, it is important to be able to predict these instabilities accurately. Much of the research studying the stability of tiltrotors has been focused on the understanding and prediction of whirl flutter. As this instability is increasingly well understood, air and ground resonance for a tiltrotor need to be investigated. Once we understand the problems of air and ground resonance in a tiltrotor, we must look for solutions to these instabilities. Other researchers have found composite or kinematic couplings in the blades of a helicopter helpful for ground and air resonance stability. Tiltrotor research has shown composite couplings in the wing to be helpful for whirl flutter. Therefore, this project will undertake to model ground and air resonance of a soft-inplane hingeless tiltrotor to understand the mechanisms involved and to evaluate whether aeroelastic couplings in the wing or kinematic couplings in the blades would aid in stabilizing these instabilities in a tiltrotor.

Motion-base simulator study of control of an externally blown flap STOL transport aircraft after failure of an outboard engine during landing approach

NASA Technical Reports Server (NTRS)

Middleton, D. B.; Hurt, G. J., Jr.; Bergeron, H. P.; Patton, J. M., Jr.; Deal, P. L.; Champine, R. A.

1975-01-01

A moving-base simulator investigation of the problems of recovery and landing of a STOL aircraft after failure of an outboard engine during final approach was made. The approaches were made at 75 knots along a 6 deg glide slope. The engine was failed at low altitude and the option to go around was not allowed. The aircraft was simulated with each of three control systems, and it had four high-bypass-ratio fan-jet engines exhausting against large triple-slotted wing flaps to produce additional lift. A virtual-image out-the-window television display of a simulated STOL airport was operating during part of the investigation. Also, a simple heads-up flight director display superimposed on the airport landing scene was used by the pilots to make some of the recoveries following an engine failure. The results of the study indicated that the variation in visual cues and/or motion cues had little effect on the outcome of a recovery, but they did have some effect on the pilot's response and control patterns.

Investigation of surface fluctuating pressures on a 1/4 scale YC-14 upper surface blown flap model

NASA Technical Reports Server (NTRS)

Pappa, R. S.

1979-01-01

Fluctuating pressures were measured at 30 positions on the surface of a 1/4-scale YC-14 wing and fuselage model during an outdoor static testing program. These data were obtained as part of a NASA program to study the fluctuating loads imposed on STOL aircraft configurations and to further the understanding of the scaling laws of unsteady surface pressure fields. Fluctuating pressure data were recorded at several discrete engine thrust settings for each of 16 configurations of the model. These data were reduced using the technique of random data analysis to obtain auto-and cross-spectral density functions and coherence functions for frequencies from 0 to 10 kHz, and cross-correlation functions for time delays from 0 to 10.24 ms. Results of this program provide the following items of particular interest: (1) Good collapse of normalized PSD functions on the USB flap was found using a technique applied by Lilley and Hodgson to data from a laboratory wall-jet apparatus. (2) Results indicate that the fluctuating pressure loading on surfaces washed by the jet exhaust flow was dominated by hydrodynamic pressure variations, loading on surface well outside the flow region dominated by acoustic pressure variations, and loading near the flow boundaries from a mixture of the two.

Noise reduction tests of large-scale-model externally blown flap using trailing-edge blowing and partial flap slot covering. [jet aircraft noise reduction

NASA Technical Reports Server (NTRS)

Mckinzie, D. J., Jr.; Burns, R. J.; Wagner, J. M.

1976-01-01

Noise data were obtained with a large-scale cold-flow model of a two-flap, under-the-wing, externally blown flap proposed for use on future STOL aircraft. The noise suppression effectiveness of locating a slot conical nozzle at the trailing edge of the second flap and of applying partial covers to the slots between the wing and flaps was evaluated. Overall-sound-pressure-level reductions of 5 db occurred below the wing in the flyover plane. Existing models of several noise sources were applied to the test results. The resulting analytical relation compares favorably with the test data. The noise source mechanisms were analyzed and are discussed.

Flight investigation of a four-dimensional terminal area guidance system for STOL aircraft

NASA Technical Reports Server (NTRS)

Neuman, F.; Hardy, G. H.

1981-01-01

A series of flight tests and fast-time simulations were conducted, using the augmentor wing jet STOL research aircraft and the STOLAND 4D-RNAV system to add to the growing data base of 4D-RNAV system performance capabilities. To obtain statistically meaningful data a limited amount of flight data were supplemented by a statistically significant amount of data obtained from fast-time simulation. The results of these tests are reported. Included are comparisons of the 4D-RNAV estimated winds with actual winds encountered in flight, as well as data on along-track navigation and guidance errors, and time-of-arrival errors at the final approach waypoint. In addition, a slight improvement of the STOLAND 4D-RNAV system is proposed and demonstrated, using the fast-time simulation.

Phase 2 and 3 wind tunnel tests of the J-97 powered, external augmentor V/STOL model. [at Ames 40 by 80 wind tunnel

NASA Technical Reports Server (NTRS)

Garland, D. B.; Harris, J. L.

1980-01-01

Static and forward speed tests were made in a 40 multiplied by 80 foot wind tunnel of a large-scale, ejector-powered V/STOL aircraft model. Modifications were made to the model following earlier tests primarily to improve longitudinal acceleration capability during transition from hovering to wingborne flight. A rearward deflection of the fuselage augmentor thrust vector was shown to be beneficial in this regard. Other augmentor modifications were tested, notably the removal of both endplates, which improved acceleration performance at the higher transition speeds. The model tests again demonstrated minimal interference of the fuselage augmentor on aerodynamic lift. A flapped canard surface also showed negligible influence on the performance of the wing and of the fuselage augmentor.

Effects of deflected thrust on the longitudinal aerodynamic characteristics of a close-coupled wing-canard configuration. [in the Langley V/STOL tunnel

NASA Technical Reports Server (NTRS)

Yip, L. P.; Paulson, J. W., Jr.

1977-01-01

The effects of power on the longitudinal aerodynamic characteristics of a close-coupled wing-canard fighter configuration with partial-span rectangular nozzles at the trailing edge of the wing were investigated. Data were obtained on a basic wing-strake configuration for nozzle and flap deflections from 0 deg to 30 deg and for nominal thrust coefficients from 0 to 0.30. The model was tested over an angle-of-attack range from -2 deg to 40 deg at Mach numbers of 0.15 and 0.18. Results show substantial improvements in lift-curve slope, in maximum lift, and in drag-due-to-lift efficiency when the canard and strakes have been added to the basic wing-fuselage (wing-alone) configuration. Addition of power increased both lift-curve slope and maximum lift, improved longitudinal stability, and reduced drag due to lift on both the wing-canard and wing-canard-strake configurations. These beneficial effects are primarily derived from boundary-layer control due to moderate thrust coefficients which delay flow separation on the nozzle and inboard portion of the wing flaps.

Application of the aerodynamic energy concept to flutter suppression and gust alleviation by use of active controls

NASA Technical Reports Server (NTRS)

Nissim, E.; Caspi, A.; Lottati, I.

1976-01-01

The effects of active controls on flutter suppression and gust alleviation of the Arava twin turboprop STOL transport and the Westwind twinjet business transport are investigated. The active control surfaces are introduced in pairs which include, in any chosen wing strip, a 20-percent chord leading-edge control and a 20-percent chord trailing-edge control. Each control surface is driven by a combined linear-rotational sensor system located on the activated strip. The control law is based on the concept of aerodynamic energy and utilizes previously optimized control law parameters based on two-dimensional aerodynamic theory. The best locations of the activated system along the span of the wing are determined for bending-moment alleviation, reduction in fuselage accelerations, and flutter suppression. The effectiveness of the activated system over a wide range of maximum control deflections is also determined. Two control laws are investigated. The first control law utilizes both rigid-body and elastic contributions of the motion. The second control law employs primarily the elastic contribution of the wing and leads to large increases in the activated control effectiveness as compared with the basic control law. The results indicate that flutter speed can be significantly increased (over 70 percent increase) and that the bending moment due to gust loading can be almost totally eliminated by a control system of about 10 to 20 percent span with reasonable control-surface rotations.

Phase 2 and 3 wind tunnel tests of the J-97 powered, external augmentor V/STOL model. [conducted in Ames 40- by 80-foot wind tunnel

NASA Technical Reports Server (NTRS)

Garland, D. B.

1980-01-01

Modifications were made to the model to improve longitudinal acceleration capability during transition from hovering to wing borne flight. A rearward deflection of the fuselage augmentor thrust vector is shown to be beneficial in this regard. Other agmentor modifications were tested, notably the removal of both endplates, which improved acceleration performance at the higher transition speeds. The model tests again demonstrated minimal interference of the fuselage augmentor on aerodynamic lift. A flapped canard surface also shows negligible influence on the performance of the wing and of the fuselage augmentor.

Configuration management and automatic control of an augmentor wing aircraft with vectored thrust

NASA Technical Reports Server (NTRS)

Cicolani, L. S.; Sridhar, B.; Meyer, G.

1979-01-01

An advanced structure for automatic flight control logic for powered-lift aircraft operating in terminal areas is under investigation at Ames Research Center. This structure is based on acceleration control; acceleration commands are constructed as the sum of acceleration on the reference trajectory and a corrective feedback acceleration to regulate path tracking errors. The central element of the structure, termed a Trimmap, uses a model of the aircraft aerodynamic and engine forces to calculate the control settings required to generate the acceleration commands. This report describes the design criteria for the Trimmap and derives a Trimmap for Ames experimental augmentor wing jet STOL research aircraft.

Flight directors for STOl aircraft

NASA Technical Reports Server (NTRS)

Rabin, U. H.

1983-01-01

Flight director logic for flight path and airspeed control of a powered-lift STOL aircraft in the approach, transition, and landing configurations are developed. The methods for flight director design are investigated. The first method is based on the Optimal Control Model (OCM) of the pilot. The second method, proposed here, uses a fixed dynamic model of the pilot in a state space formulation similar to that of the OCM, and includes a pilot work-load metric. Several design examples are presented with various aircraft, sensor, and control configurations. These examples show the strong impact of throttle effectiveness on the performance and pilot work-load associated with manual control of powered-lift aircraft during approach. Improved performed and reduced pilot work-load can be achieved by using direct-lift-control to increase throttle effectiveness.

Aircraft aerodynamic prediction method for V/STOL transition including flow separation

NASA Technical Reports Server (NTRS)

Gilmer, B. R.; Miner, G. A.; Bristow, D. R.

1983-01-01

A numerical procedure was developed for the aerodynamic force and moment analysis of V/STOL aircraft operating in the transition regime between hover and conventional forward flight. The trajectories, cross sectional area variations, and mass entrainment rates of the jets are calculated by the Adler-Baron Jet-in-Crossflow Program. The inviscid effects of the interaction between the jets and airframe on the aerodynamic properties are determined by use of the MCAIR 3-D Subsonic properties are determined by use of the MCAIR 3-D Subsonic Potential Flow Program, a surface panel method. In addition, the MCAIR 3-D Geometry influence Coefficient Program is used to calculate a matrix of partial derivatives that represent the rate of change of the inviscid aerodynamic properties with respect to arbitrary changes in the effective wing shape.

Robustness of linear quadratic state feedback designs in the presence of system uncertainty. [application to Augmentor Wing Jet STOL Research Aircraft flare control autopilot design

NASA Technical Reports Server (NTRS)

Patel, R. V.; Toda, M.; Sridhar, B.

1977-01-01

The paper deals with the problem of expressing the robustness (stability) property of a linear quadratic state feedback (LQSF) design quantitatively in terms of bounds on the perturbations (modeling errors or parameter variations) in the system matrices so that the closed-loop system remains stable. Nonlinear time-varying and linear time-invariant perturbations are considered. The only computation required in obtaining a measure of the robustness of an LQSF design is to determine the eigenvalues of two symmetric matrices determined when solving the algebraic Riccati equation corresponding to the LQSF design problem. Results are applied to a complex dynamic system consisting of the flare control of a STOL aircraft. The design of the flare control is formulated as an LQSF tracking problem.

Helicopters: The Platform of Choice for Defense and Development

DTIC Science & Technology

2013-04-25

FID effort by facilitating the IDAD tasks of balanced development, security, neutralization, and mobilization. In order to prevent instability...security, neutralization, and mobilization. In order to prevent instability inside a Host Nation (HN), it is important to build cohesive IDAD and...derived from the H -53 (arguably the oldest, biggest, and most cost-prohibitive platform of this type) compared to STOL-capable 4-9 passenger fixed wing

The lateral/directional stability characteristics of a four-propeller tilt-wing V/STOL model in low-speed steep descent. M.S. Thesis - Princeton Univ., N.J.

NASA Technical Reports Server (NTRS)

Dicarlo, D. J.

1971-01-01

Lateral-directional dynamic stability derivatives are presented for a O.1-scale model of the XC-142A tilt-wing transport. The tests involved various descending flight conditions achieved at constant speed and wing incidence by varying the vehicle angle of attack. The propeller blade angle and the speed were also changed in the steepest descent case. The experimental data were analyzed assuming that the dynamic motions of the vehicle may be described by linearized equations, with the lateral-directional characteristics of the full-scale aircraft also presented and discussed. Results from this experimental investigation indicated that the full-scale aircraft would have a stable lateral-directional motion in level flight, with the dynamic motion becoming less stable as the descent angle was increased.

Summary of low-speed longitudinal aerodynamics of two powered close-coupled wing-canard fighter configurations. [conducted in Langley C/STOL tunnel

NASA Technical Reports Server (NTRS)

Paulson, J. W., Jr.; Thomas, J. L.

1979-01-01

Investigations of the low speed longitudinal characteristics of two powered close coupled wing-canard fighter configurations are discussed. Data obtained at angles of attack from -2 deg to 42 deg, Mach numbers from 0.12 to 0.20, nozzle and flap deflections from 0 deg to 40 deg, and thrust coefficients from 0 to 2.0, to represent both high angle of attack subsonic maneuvering characteristics and conventional takeoff and landing characteristics are examined. Data obtained with the nozzles deflected either 60 deg or 90 deg and the flaps deflected 60 deg to represent vertical or short takeoff and landing characteristics are discussed.

Longitudinal handling qualities during approach and landing of a powered lift STOL aircraft

NASA Technical Reports Server (NTRS)

Franklin, J. A.; Innis, R. C.

1972-01-01

Longitudinal handling qualities evaluations were conducted on the Ames Research Center Flight Simulator for Advanced Aircraft (FSAA) for the approach and landing tasks of a powered lift STOL research aircraft. The test vehicle was a C-8A aircraft modified with a new wing incorporating internal blowing over an augmentor flap. The investigation included: (1) use of various flight path and airspeed control techniques for the basic vehicle; (2) assessment of stability and command augmentation schemes for pitch attitude and airspeed control; (3) determination of the influence of longitudinal and vertical force coupling for the power control; (4) determination of the influence of pitch axis coupling with the thrust vector control; and (5) evaluations of the contribution of stability and command augmentation to recovery from a single engine failure. Results are presented in the form of pilot ratings and commentary substantiated by landing approach time histories.

Whirl Flutter Studies for a SSTOL Transport Demonstrator

NASA Technical Reports Server (NTRS)

Acree, C. W., Jr.; Hoffman, Krishna

2004-01-01

A proposed new class of aircraft - the Advanced Theater Transport (ATT) will combine strategic range and high payload with 'Super-STOL' (short take-off and landing) capability. It is also proposed to modify a YC-15 into a technology demonstrator with a 20-deg tilt wing; four, eight-bladed propellers; cross-shafted gearboxes and V-22 engines. These constitute a unique combination of design features that potentially affect performance, loads and whirl-mode stability (whirl flutter). NASA Ames Research Center is working with Boeing and Hamilton Sundstrand on technology challenges presented by the concept; the purpose of NASA involvement is to establish requirements for the demonstrator and for early design guidance, with emphasis on whirl flutter. CAMRAD II is being used to study the effects of various design features on whirl flutter, with special attention to areas where such features differ from existing aircraft, notably tiltrotors. Although the stability margins appear to be more than adequate, the concept requires significantly different analytical methods, principally including far more blade modes, than typically used for tiltrotors.

Flap survey test of a combined surface blowing model: Flow measurements at static flow conditions

NASA Technical Reports Server (NTRS)

Fukushima, T.

1978-01-01

The Combined Surface Blowing (CSB) V/STOL lift/propulsion system consists of a blown flap system which deflects the exhaust from a turbojet engine over a system of flaps deployed at the trailing edge of the wing. Flow measurements consisting of velocity measurements using split film probes and total measure surveys using a miniature Kiel probe were made at control stations along the flap systems at two spanwise stations, the centerline of the nozzle and 60 percent of the nozzle span outboard of the centerline. Surface pressure measurements were made in the wing cove and the upper surface of the first flap element. The test showed a significant flow separation in the wing cove. The extent of the separation is so large that the flow into the first flap takes place only at the leading edge of the flap. The velocity profile measurements indicate that large spanwise (3 dimensional) flow may exist.

Correlation of low speed wind tunnel and flight test data for V/STOL aircraft

NASA Technical Reports Server (NTRS)

Cook, W. L.; Hickey, D. H.

1975-01-01

The XV-5B fan-in-wing aircraft and the Y0V-10 RCF rotating cylinder flap aircraft were subjected to wind tunnel tests. These tests were conducted specifically to provide for correlation between wind tunnel and inflight aerodynamics and noise test data. Correlation between aerodynamic and noise data are presented and testing techniques that are related to the accuracy of the data, or that might affect the correlations, are discussed.

Longitudinal aerodynamic characteristics of a low-wing lift-fan transport including hover characteristics in and out of ground effect

NASA Technical Reports Server (NTRS)

Hoad, D. R.; Gentry, G. L., Jr.

1977-01-01

The longitudinal aerodynamic characteristics of a six-fan, tip-driven (remote) lift-fan VTOL transport through transition were determined by an investigation conducted in the Langley V/STOL tunnel. Tests were also made with the large midspan lift-fan pods and lift-cruise fans removed to determine their their influence on the stability and control of the configuration. Data were obtained for a range of model height above ground.

A comparison of flight and simulation data for three automatic landing system control laws for the Augmentor wing jet STOL research airplane

NASA Technical Reports Server (NTRS)

Feinreich, B.; Gevaert, G.

1980-01-01

Automatic flare and decrab control laws for conventional takeoff and landing aircraft were adapted to the unique requirements of the powered lift short takeoff and landing airplane. Three longitudinal autoland control laws were developed. Direct lift and direct drag control were used in the longitudinal axis. A fast time simulation was used for the control law synthesis, with emphasis on stochastic performance prediction and evaluation. Good correlation with flight test results was obtained.

Application of Powered High Lift Systems to STOL Aircraft Design.

DTIC Science & Technology

1979-09-01

century by da Vinci, an English - man named Sir George Cayley first attempted to integrate the features of the helicopter and the airplane. In 1843 his...horizontal flight Jun 1959 WEIGHT: 6500 LBS ENGINES: (1) SNECMA ATAR 101 E.V. Turbojet of 8155 lbs thrust LAYOUT: See Fig. 21 COMMENTS: SNECMA was engaged...34 ATAR VOLANT" test vehicle fitted with an annular wing to permit transition to horizontal flight. The aircraft was controlled from a tilting ejec- tion

Fuel-Conservation Guidance System for Powered-Lift Aircraft

NASA Technical Reports Server (NTRS)

Erzberger, Heinz; McLean, John D.

1981-01-01

A technique is described for the design of fuel-conservative guidance systems and is applied to a system that was flight tested on board NASA's sugmentor wing jet STOL research aircraft. An important operational feature of the system is its ability to rapidly synthesize fuel-efficient trajectories for a large set of initial aircraft positions, altitudes, and headings. This feature allows the aircraft to be flown efficiently under conditions of changing winds and air traffic control vectors. Rapid synthesis of fuel-efficient trajectories is accomplished in the airborne computer by fast-time trajectory integration using a simplified dynamic performance model of the aircraft. This technique also ensures optimum flap deployment and, for powered-lift STOL aircraft, optimum transition to low-speed flight. Also included in the design is accurate prediction of touchdown time for use in four-dimensional guidance applications. Flight test results have demonstrated that the automatically synthesized trajectories produce significant fuel savings relative to manually flown conventional approaches.

Simulation test results for lift/cruise fan research and technology aircraft

NASA Technical Reports Server (NTRS)

Bland, M. P.; Konsewicz, R. K.

1976-01-01

A flight simulation program was conducted on the flight simulator for advanced aircraft (FSAA). The flight simulation was a part of a contracted effort to provide a lift/cruise fan V/STOL aircraft mathematical model for flight simulation. The simulated aircraft is a configuration of the Lift/Cruise Fan V/STOL research technology aircraft (RTA). The aircraft was powered by three gas generators driving three fans. One lift fan was installed in the nose of the aircraft, and two lift/cruise fans at the wing root. The thrust of these fans was modulated to provide pitch and roll control, and vectored to provide yaw, side force control, and longitudinal translation. Two versions of the RTA were defined. One was powered by the GE J97/LF460 propulsion system which was gas-coupled for power transfer between fans for control. The other version was powered by DDA XT701 gas generators driving 62 inch variable pitch fans. The flight control system in both versions of the RTA was the same.

Wind tunnel and ground static investigation of a large scale model of a lift/cruise fan V/STOL aircraft

NASA Technical Reports Server (NTRS)

1976-01-01

An investigation was conducted in a 40 foot by 80 foot wind tunnel to determine the aerodynamic/propulsion characteristics of a large scale powered model of a lift/cruise fan V/STOL aircraft. The model was equipped with three 36 inch diameter turbotip X376B fans powered by three T58 gas generators. The lift fan was located forward of the cockpit area and the two lift/cruise fans were located on top of the wing adjacent to the fuselage. The three fans with associated thrust vectoring systems were used to provide vertical, and short, takeoff and landing capability. For conventional cruise mode operation, only the lift/cruise fans were utilized. The data that were obtained include lift, drag, longitudinal and lateral-directional stability characteristics, and control effectiveness. Data were obtained up to speeds of 120 knots at one model height of 20 feet for the conventional aerodynamic lift configuration and at several thrust vector angles for the powered lift configuration.

V/STOL tilt rotor aircraft study. Volume 10: Performance and stability test of A 1-14.622 Froude scaled Boeing Vertol Model 222 tilt rotor aircraft (Phase 1)

NASA Technical Reports Server (NTRS)

Mchugh, F. J.; Eason, W.; Alexander, H. R.; Mutter, H.

1973-01-01

Wind tunnel test data obtained from a 1/4.622 Froude scale Boeing Model 222 with a full span, two prop, tilt rotor, powered model in the Boeing V/STOL wind tunnel are reported. Data were taken in transition and cruise flight conditions and include performance, stability and control and blade loads information. The effects of the rotors, tail surfaces and airframe on the performance and stability are isolated as are the effects of the airframe on the rotors.

Unsteady loads due to propulsive lift configurations. Part C: Development of experimental techniques for investigation of unsteady pressures behind a cold model jet

NASA Technical Reports Server (NTRS)

Haviland, J. K.; Schroeder, J. C.

1978-01-01

As part of an overall study of the scaling laws for the fluctuating pressures induced on the wings and flaps of STOL aircraft by jet engine impingement, an experimental investigation was made of the near field fluctuating pressures behind a cold circular jet, both when it was free and when it was impinging on a flat plate. Miniature static pressure probes were developed for measurements in the free jet and on the flat plate which were connected by plastic tubing to 1/8 inch microphones and acted as pressure transducers. Using a digital correlator together with an FFT program on the CDC 6400 computer, spectral densities, relative amplitudes, phase lags, and coherences were also obtained for the signals from pairs of these probes, and were used to calibrate these probes directly against microphones. This system of instrumentation was employed to obtain single point rms and third octave surveys of the static pressures in the free jet and on the surface of the plate.

A lift-cancellation technique in linearized supersonic-wing theory

NASA Technical Reports Server (NTRS)

Mirels, Harold

1951-01-01

A lift-cancellation technique is presented for determining load distributions on thin wings at supersonic speeds. The loading on a wing having a prescribed plan form is expressed as the loading of a known related wing (such as a two-dimensional or triangular wing) minus the loading of an appropriate cancellation wing. The lift-cancellation technique can be used to find the loading on a large variety of wings. Applications to swept wings having curvilinear plan forms and to wings having reentrant side edges are indicated.

Low Speed Wind Tunnel Tests on a One-Seventh Scale Model of the H.126 Jet Flap Aircraft

NASA Technical Reports Server (NTRS)

Laub, G. H.

1975-01-01

Low speed wind tunnel tests were performed on a one-seventh scale model of the British H.126 jet flap research aircraft over a range of jet momentum coefficients. The primary objective was to compare model aerodynamic characteristics with those of the aircraft, with the intent to provide preliminary data needed towards establishing small-to-full scale correlating techniques on jet flap V/STOL aircraft configurations. Lift and drag coefficients from the model and aircraft tests were found to be in reasonable agreement. The pitching moment coefficient and trim condition correlation was poor. A secondary objective was to evaluate a modified thrust nozzle having thrust reversal capability. The results showed there was a considerable loss of lift in the reverse thrust operational mode because of increased nozzle-wing flow interference. A comparison between the model simulated H.126 wing jet efflux and the model uniform pressure distribution wing jet efflux indicated no more than 5% loss in weight flow rate.

Analytical study of STOL Aircraft in ground effect. Part 1: Nonplanar, nonlinear wing/jet lifting surface method

NASA Technical Reports Server (NTRS)

Shollenberger, C. A.; Smyth, D. N.

1978-01-01

A nonlinear, nonplanar three dimensional jet flap analysis, applicable to the ground effect problem, is presented. Lifting surface methodology is developed for a wing with arbitrary planform operating in an inviscid and incompressible fluid. The classical, infintely thin jet flap model is employed to simulate power induced effects. An iterative solution procedure is applied within the analysis to successively approximate the jet shape until a converged solution is obtained which closely satisfies jet and wing boundary conditions. Solution characteristics of the method are discussed and example results are presented for unpowered, basic powered and complex powered configurations. Comparisons between predictions of the present method and experimental measurements indicate that the improvement of the jet with the ground plane is important in the analyses of powered lift systems operating in ground proximity. Further development of the method is suggested in the areas of improved solution convergence, more realistic modeling of jet impingement and calculation efficiency enhancements.

Determination of mean camber surfaces for wings having uniform chordwise loading and arbitrary spanwise loading in subsonic flow

NASA Technical Reports Server (NTRS)

Katzoff, S; Faison, M Frances; Dubose, Hugh C

1954-01-01

The field of a uniformly loaded wing in subsonic flow is discussed in terms of the acceleration potential. It is shown that, for the design of such wings, the slope of the mean camber surface at any point can be determined by a line integration around the wing boundary. By an additional line integration around the wing boundary, this method is extended to include the case where the local section lift coefficient varies with spanwise location (the chordwise loading at every section still remaining uniform). For the uniformly loaded wing of polygonal plan form, the integrations necessary to determine the local slope of the surface and the further integration of the slopes to determine the ordinate can be done analytically. An outline of these integrations and the resulting formulas are included. Calculated results are given for a sweptback wing with uniform chordwise loading and a highly tapered spanwise loading, a uniformly loaded delta wing, a uniformly loaded sweptback wing, and the same sweptback wing with uniform chordwise loading but elliptical span load distribution.

Prediction of circulation control performance characteristics for Super STOL and STOL applications

NASA Astrophysics Data System (ADS)

Naqvi, Messam Abbas

The rapid air travel growth during the last three decades, has resulted in runway congestion at major airports. The current airports infrastructure will not be able to support the rapid growth trends expected in the next decade. Changes or upgrades in infrastructure alone would not be able to satisfy the growth requirements, and new airplane concepts such as the NASA proposed Super Short Takeoff and Landing and Extremely Short Takeoff & Landing (ESTOL) are being vigorously pursued. Aircraft noise pollution during Takeoff & Landing is another serious concern and efforts are aimed to reduce the airframe noise produced by Conventional High Lift Devices during Takeoff & Landing. Circulation control technology has the prospect of being a good alternative to resolve both the aforesaid issues. Circulation control airfoils are not only capable of producing very high values of lift (Cl values in excess of 8.0) at zero degree angle of attack, but also eliminate the noise generated by the conventional high lift devices and their associated weight penalty as well as their complex operation and storage. This will ensure not only satisfying the small takeoff and landing distances, but minimal acoustic signature in accordance with FAA requirements. The Circulation Control relies on the tendency of an emanating wall jet to independently control the circulation and lift on an airfoil. Unlike, conventional airfoil where rear stagnation point is located at the sharp trailing edge, circulation control airfoils possess a round trailing edge, therefore the rear stagnation point is free to move. The location of rear stagnation point is controlled by the blown jet momentum. This provides a secondary control in the form of jet momentum with which the lift generated can be controlled rather the only available control of incidence (angle of attack) in case of conventional airfoils. The use of Circulation control despite its promising potential has been limited only to research applications due to the lack of a simple prediction capability. This research effort was focused on the creation of a rapid prediction capability of Circulation Control Aerodynamic Characteristics which could help designers with rapid performance estimates for design space exploration. A morphological matrix was created with the available set of options which could be chosen to create this prediction capability starting with purely analytical physics based modeling to high fidelity CFD codes. Based on the available constraints, and desired accuracy meta-models have been created around the two dimensional circulation control performance results computed using Navier Stokes Equations (Computational Fluid Dynamics). DSS2, a two dimensional RANS code written by Professor Lakshmi Sankar was utilized for circulation control airfoil characteristics. The CFD code was first applied to the NCCR 1510-7607N airfoil to validate the model with available experimental results. It was then applied to compute the results of a fractional factorial design of experiments array. Metamodels were formulated using the neural networks to the results obtained from the Design of Experiments. Additional validation runs were performed to validate the model predictions. Metamodels are not only capable of rapid performance prediction, but also help generate the relation trends of response matrices with control variables and capture the complex interactions between control variables. Quantitative as well as qualitative assessments of results were performed by computation of aerodynamic forces & moments and flow field visualizations. Wing characteristics in three dimensions were obtained by integration over the whole wing using Prandtl's Wing Theory. The baseline Super STOL configuration [3] was then analyzed with the application of circulation control technology. The desired values of lift and drag to achieve the target values of Takeoff & Landing performance were compared with the optimal configurations obtained by the model. The same optimal configurations were then subjected to Super STOL cruise conditions to perform a trade off analysis between Takeoff and Cruise Performance. Supercritical airfoils modified for circulation control were also thoroughly analyzed for Takeoff and Cruise performance and may constitute a viable option for Super STOL & STOL Designs. The prediction capability produced by this research effort can be integrated with the current conceptual aircraft modeling & simulation framework. The prediction tool is applicable within the selected ranges of each variable, but methodology and formulation scheme adopted can be applied to any other design space exploration.

A preliminary design proposal for a maritime patrol strike aircraft: MPS-2000 Condor

NASA Technical Reports Server (NTRS)

1994-01-01

The four member graduate design team assembled to submit a proposal for the 1993/1994 RFP at the University of Kansas has designed a four seat, variable swept wing, twin turbofan aircraft with STOL capabilities. The aircraft is named the MPS-2000 Condor and is capable of carrying air-to-surface or air-to-air weapon systems along with attack and surveillance radar and IRF systems. The aircraft has a cruise range of 800 nautical miles, a loiter of 4 hours, and a dash speed of 500 kts.

V/STOL tilt rotor aircraft study mathematical model for a real time simulation of a tilt rotor aircraft (Boeing Vertol Model 222), volume 8

NASA Technical Reports Server (NTRS)

Rosenstein, H.; Mcveigh, M. A.; Mollenkof, P. A.

1973-01-01

A mathematical model for a real time simulation of a tilt rotor aircraft was developed. The mathematical model is used for evaluating aircraft performance and handling qualities. The model is based on an eleven degree of freedom total force representation. The rotor is treated as a point source of forces and moments with appropriate response time lags and actuator dynamics. The aerodynamics of the wing, tail, rotors, landing gear, and fuselage are included.

Flow visualization studies of VTOL aircraft models during Hover in ground effect

NASA Technical Reports Server (NTRS)

Mourtos, Nikos J.; Couillaud, Stephane; Carter, Dale; Hange, Craig; Wardwell, Doug; Margason, Richard J.

1995-01-01

A flow visualization study of several configurations of a jet-powered vertical takeoff and landing (VTOL) aircraft model during hover in ground effect was conducted. A surface oil flow technique was used to observe the flow patterns on the lower surfaces of the model. There were significant configuration effects. Wing height with respect to fuselage, the presence of an engine inlet duct beside the fuselage, and nozzle pressure ratio are seen to have strong effects on the surface flow angles on the lower surface of the wing. This test was part of a program to improve the methods for predicting the hot gas ingestion (HGI) for jet-powered vertical/short takeoff and landing (V/STOL) aircraft. The tests were performed at the Jet Calibration and Hover Test (JCAHT) Facility at Ames Research Center.

Development and analysis of a STOL supersonic cruise fighter concept

NASA Technical Reports Server (NTRS)

Dollyhigh, S. M.; Foss, W. E., Jr.; Morris, S. J., Jr.; Walkley, K. B.; Swanson, E. E.; Robins, A. W.

1984-01-01

The application of advanced and emerging technologies to a fighter aircraft concept is described. The twin-boom fighter (TBF-1) relies on a two dimensional vectoring/reversing nozzle to provide STOL performance while also achieving efficient long range supersonic cruise. A key feature is that the propulsion package is placed so that the nozzle hinge line is near the aircraft center-of-gravity to allow large vector angles and, thus, provide large values of direct lift while minimizing the moments to be trimmed. The configurations name is derived from the long twin booms extending aft of the engine to the twin vertical tails which have a single horizontal tail mounted atop and between them. Technologies utilized were an advanced engine (1985 state-of-the-art), superplastic formed/diffusion bonded titanium structure, advanced controls/avionics/displays, supersonic wing design, and conformal weapons carriage. The integration of advanced technologies into this concept indicate that large gains in takeoff and landing performance, maneuver, acceleration, supersonic cruise speed, and range can be acieved relative to current fighter concepts.

Experimental aerodynamic characteristics of two V/STOL fighter/attack aircraft configurations at Mach numbers from 0.4 to 1.4

NASA Technical Reports Server (NTRS)

Nelms, W. P.; Durston, D. A.; Lummus, J. R.

1980-01-01

A wind tunnel test was conducted to measure the aerodynamic characteristics of two horizontal attitude takeoff and landing V/STOL fighter/attack aircraft concepts. In one concept, a jet diffuser ejector was used for the vertical lift system; the other used a remote augmentation lift system (RALS). Wind tunnel tests to investigate the aerodynamic uncertainties and to establish a data base for these types of concepts were conducted over a Mach number range from 0.2 to 2.0. The present report covers tests, conducted in the 11 foot transonic wind tunnel, for Mach numbers from 0.4 to 1.4. Detailed effects of varying the angle of attack (up to 27 deg), angle of sideslip (-4 deg to +8 deg), Mach number, Reynolds number, and configuration buildup were investigated. In addition, the effects of wing trailing edge flap deflections, canard incidence, and vertical tail deflections were explored. Variable canard longitudinal location and different shapes of the inboard nacelle body strakes were also investigated.

Robustness of linear quadratic state feedback designs in the presence of system uncertainty. [applied to STOL autopilot design

NASA Technical Reports Server (NTRS)

Patel, R. V.; Toda, M.; Sridhar, B.

1977-01-01

In connection with difficulties concerning an accurate mathematical representation of a linear quadratic state feedback (LQSF) system, it is often necessary to investigate the robustness (stability) of an LQSF design in the presence of system uncertainty and obtain some quantitative measure of the perturbations which such a design can tolerate. A study is conducted concerning the problem of expressing the robustness property of an LQSF design quantitatively in terms of bounds on the perturbations (modeling errors or parameter variations) in the system matrices. Bounds are obtained for the general case of nonlinear, time-varying perturbations. It is pointed out that most of the presented results are readily applicable to practical situations for which a designer has estimates of the bounds on the system parameter perturbations. Relations are provided which help the designer to select appropriate weighting matrices in the quadratic performance index to attain a robust design. The developed results are employed in the design of an autopilot logic for the flare maneuver of the Augmentor Wing Jet STOL Research Aircraft.

Flutter suppression and gust alleviation using active controls

NASA Technical Reports Server (NTRS)

Nissim, E.

1974-01-01

The effects of active controls on the suppression of flutter and gust alleviation of two different types of subsonic aircraft (the Arava, twin turboprop STOL transport, and the Westwind twin-jet business transport) are investigated. The active controls are introduced in pairs which include, in any chosen wing strip, a leading-edge (LE) control and a trailing-edge (TE) control. Each control surface is allowed to be driven by a combined linear-rotational sensor system, located on the activated strip. The control law, which translates the sensor signals into control surface rotations, is based on the concept of aerodynamic energy. The results indicate the extreme effectiveness of the active systems in controlling flutter. A single system spanning 10% of the wing semispan made the Arava flutter-free, and a similar active system, for the Westwind aircraft, yielded a reduction of 75% in the maximum bending moment of the wing and a reduction of 90% in the acceleration of the cg of the aircraft. Results for simultaneous activation of several LE - TE systems are presented. Further work needed to bring the investigation to completion is also discussed.

The effect of wing dihedral and section suction distribution on vortex bursting

NASA Technical Reports Server (NTRS)

Washburn, K. E.; Gloss, B. B.

1975-01-01

Eleven semi-span wing models were tested in the 1/8-scale model of the Langley V/STOL tunnel to qualitatively study vortex bursting. Flow visualization was achieved by using helium filled soap bubbles introduced upstream of the model. The angle of attack range was from 0 deg to 45 deg. The results show that the vortex is unstable, that is, the bursting point location is not fixed at a given angle of attack but moves within certain bounds. Upstream of the trailing edge, the bursting point location has a range of two inches; downstream, the range is about six inches. Anhedral and dihedral appear to have an insignificant effect on the vortex and its bursting point location. Altering the section suction distribution by improving the triangularity generally increases the angle of attack at which vortex bursting occurs at the trailing edge.

Wing loading in 15 species of North American owls

Treesearch

David H. Johnson

1997-01-01

Information on wing morphology is important in understanding the mechanics and energetics of flight and in aspects related to reversed sexual size dimorphism in owls. I summarized wing span, wing area, wing loading, root box, and aspect ratio calculations from the available literature and from 113 owls examined in this study. Wing loading estimates for 15 species...

Evaluation of advanced lift concepts and potential fuel conservation for short-haul aircraft

NASA Technical Reports Server (NTRS)

Sweet, H. S.; Renshaw, J. H.; Bowden, M. K.

1975-01-01

The effect of different field lengths, cruise requirements, noise level, and engine cycle characteristics on minimizing fuel consumption and minimizing operating cost at high fuel prices were evaluated for some advanced short-haul aircraft. The conceptual aircraft were designed for 148 passengers using the upper surface-internally blown jet flap, the augmentor wing, and the mechanical flap lift systems. Advanced conceptual STOL engines were evaluated as well as a near-term turbofan and turboprop engine. Emphasis was given to designs meeting noise levels equivalent to 95-100 EPNdB at 152 m (500 ft) sideline.

Investigations of simulated aircraft flight through thunderstorm outflows

NASA Technical Reports Server (NTRS)

Frost, W.; Crosby, B.

1978-01-01

The effects of wind shear on aircraft flying through thunderstorm gust fronts were investigated. A computer program was developed to solve the two dimensional, nonlinear equations of aircraft motion, including wind shear. The procedure described and documented accounts for spatial and temporal variations of the aircraft within the flow regime. Analysis of flight paths and control inputs necessary to maintain specified trajectories for aircraft having characteristics of DC-8, B-747, augmentor wing STOL, and DHC-6 aircraft was recorded. From the analysis an attempt was made to find criteria for reduction of the hazards associated with landing through thunderstorm gust fronts.

Externally-blown-flap noise

NASA Technical Reports Server (NTRS)

Dorsch, R. G.; Kreim, W. J.; Olsen, W. A.

1972-01-01

Noise data were obtained with a large externally blown flap model. A fan-jet engine exhaust was simulated by a 1/2-scale bypass nozzle supplied by pressurized air. The nozzle was pylon mounted on a wing section having a double-slotted flap for lift augmentation. Noise radiation patterns and spectra were obtained for nozzle exhaust velocities between 400 and 1150 ft/sec. The blown flap noise data are in good agreement with previous small model results extrapolated to test conditions by Strouhal scaling. The results indicate that blown flap noise must be suppressed to meet STOL aircraft noise goals.

Jet noise suppressor nozzle development for augmentor wing jet STOL research aircraft (C-8A Buffalo)

NASA Technical Reports Server (NTRS)

Harkonen, D. L.; Marks, C. C.; Okeefe, J. V.

1974-01-01

Noise and performance test results are presented for a full-scale advanced design rectangular array lobe jet suppressor nozzle (plain wall and corrugated). Flight design and installation considerations are also discussed. Noise data are presented in terms of peak PNLT (perceived noise level, tone corrected) suppression relative to the existing airplane and one-third octave-band spectra. Nozzle performance is presented in terms of velocity coefficient. Estimates of the hot thrust available during emergency (engine out) with the suppressor nozzle installed are compared with the current thrust levels produced by the round convergent nozzles.

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

NASA Technical Reports Server (NTRS)

Lokos, William A.; Olney, Candida D.; Chen, Tony; Crawford, Natalie D.; Stauf, Rick; Reichenbach, Eric Y.; Bessette, Denis (Technical Monitor)

2002-01-01

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.

Lift-fan aircraft: Lessons learned-the pilot's perspective

NASA Technical Reports Server (NTRS)

Gerdes, Ronald M.

1993-01-01

This paper is written from an engineering test pilot's point of view. Its purpose is to present lift-fan 'lessons learned' from the perspective of first-hand experience accumulated during the period 1962 through 1988 while flight testing vertical/short take-off and landing (V/STOL) experimental aircraft and evaluating piloted engineering simulations of promising V/STOL concepts. Specifically, the scope of the discussions to follow is primarily based upon a critical review of the writer's personal accounts of 30 hours of XV-5A/B and 2 hours of X-14A flight testing as well as a limited simulator evaluation of the Grumman Design 755 lift-fan aircraft. Opinions of other test pilots who flew these aircraft and the aircraft simulator are also included and supplement the writer's comments. Furthermore, the lessons learned are presented from the perspective of the writer's flying experience: 10,000 hours in 100 fixed- and rotary-wing aircraft including 330 hours in 5 experimental V/STOL research aircraft. The paper is organized to present to the reader a clear picture of lift-fan lessons learned from three distinct points of view in order to facilitate application of the lesson principles to future designs. Lessons learned are first discussed with respect to case histories of specific flight and simulator investigations. These principles are then organized and restated with respect to four selected design criteria categories in Appendix I. Lastly, Appendix Il is a discussion of the design of a hypothetical supersonic short take-off vertical landing (STOVL) fighter/attack aircraft.

Design optimization of high-speed proprotor aircraft

NASA Technical Reports Server (NTRS)

Schleicher, David R.; Phillips, James D.; Carbajal, Kevin B.

1993-01-01

NASA's high-speed rotorcraft (HSRC) studies have the objective of investigating technology for vehicles that have both low downwash velocities and forward flight speed capability of up to 450 knots. This paper investigates a tilt rotor, a tilt wing, and a folding tilt rotor designed for a civil transport mission. Baseline aircraft models using current technology are developed for each configuration using a vertical/short takeoff and landing (V/STOL) aircraft design synthesis computer program to generate converged vehicle designs. Sensitivity studies and numerical optimization are used to illustrate each configuration's key design tradeoffs and constraints. Minimization of the gross takeoff weight is used as the optimization objective function. Several advanced technologies are chosen, and their relative impact on future configurational development is discussed. Finally, the impact of maximum cruise speed on vehicle figures of merit (gross weight, productivity, and direct operating cost) is analyzed. The three most important conclusions from the study are payload ratios for these aircraft will be commensurate with current fixed-wing commuter aircraft; future tilt rotors and tilt wings will be significantly lighter, more productive, and cheaper than competing folding tilt rotors; and the most promising technologies are an advanced-technology proprotor for both tilt rotor and tilt wing and advanced structural materials for the folding tilt rotor.

Simulation studies of STOL airplane operations in metropolitan downtown and airport air traffic control environments

NASA Technical Reports Server (NTRS)

Sawyer, R. H.; Mclaughlin, M. D.

1974-01-01

The operating problems and equipment requirements for STOL airplanes in terminal area operations in simulated air traffic control (ATC) environments were studied. These studies consisted of Instrument Flight Rules (IFR) arrivals and departures in the New York area to and from a downtown STOL port, STOL runways at John F. Kennedy International Airport, or STOL runways at a hypothetical international airport. The studies were accomplished in real time by using a STOL airplane flight simulator. An experimental powered lift STOL airplane and two in-service airplanes having high aerodynamic lift (i.e., STOL) capability were used in the simulations.

Effect of wing loading, aspect ratio, and span loading of flight performances

NASA Technical Reports Server (NTRS)

Gothert, B

1940-01-01

An investigation is made of the possible improvements in maximum, cruising, and climbing speeds attainable through increase in the wing loading. The decrease in wing area was considered for the two cases of constant aspect ratio and constant span loading. For a definite flight condition, an investigation is made to determine what loss in flight performance must be sustained if, for given reasons, certain wing loadings are not to be exceeded. With the aid of these general investigations, the trend with respect to wing loading is indicated and the requirements to be imposed on the landing aids are discussed

An investigation of tip planform influence on the aerodynamic load characteristics of semispan, upswept wing and wing-tip

NASA Technical Reports Server (NTRS)

Vanaken, Johannes M.

1986-01-01

A semi-span wing, equipped with an interchangeable tip, which was varied in planform and size was examined. Total wing aerodynamic loading was obtained from the wind tunnel scale system. The wing tip was mounted on a separate six-component strain gauge balance, which provided the aerodynamic loads on the tip. The tests were accomplished in the NASA Ames 7X10-Foot Wind Tunnel at a Mach number of 0.178. The aerodynamic load characteristics of the wing and of the tip were presented with the tip at several incidence angles relative to the wing inboard section.

Effects of winglets on a first-generation jet transport wing. 7: Sideslip effects on winglet loads and selected wing loads at subsonic speeds for a full-span model

NASA Technical Reports Server (NTRS)

Meyer, Robert R., Jr.; Covell, Peter F.

1986-01-01

The effect of sideslip on winglet loads and selected wing loads was investigated at high and low subsonic Mach numbers. The investigation was conducted in two separate wind tunnel facilities, using two slightly different 0.035-scale full-span models. Results are presented which indicate that, in general, winglet loads as a result of sideslip are analogous to wing loads caused by angle of attack. The center-of-pressure locations on the winglets are somewhat different than might be expected for an analogous wing. The spanwise center of pressure for a winglet tends to be more inboard than for a wing. The most notable chordwise location is a forward center-of-pressure location on the winglet at high sideslip angles. The noted differences between a winglet and an analogous wing are the result of the influence of the wing on the winglet.

Structural analysis and testing of a carbon-composite wing using fiber Bragg gratings

NASA Astrophysics Data System (ADS)

Nicolas, Matthew James

The objective of this study was to determine the deflected wing shape and the out-of-plane loads of a large-scale carbon-composite wing of an ultralight aerial vehicle using Fiber Bragg Grating (FBG) technology. The composite wing was instrumented with an optical fiber on its top and bottom surfaces positioned over the main spar, resulting in approximately 780 strain sensors bonded to the wings. The strain data from the FBGs was compared to that obtained from four conventional strain gages, and was used to obtain the out-of-plane loads as well as the wing shape at various load levels using NASA-developed real-time load and displacement algorithms. The composite wing measured 5.5 meters and was fabricated from laminated carbon uniaxial and biaxial prepreg fabric with varying laminate ply patterns and wall thickness dimensions. A three-tier whiffletree system was used to load the wing in a manner consistent with an in-flight loading condition.

Loads Model Development and Analysis for the F/A-18 Active Aeroelastic Wing Airplane

NASA Technical Reports Server (NTRS)

Allen, Michael J.; Lizotte, Andrew M.; Dibley, Ryan P.; Clarke, Robert

2005-01-01

The Active Aeroelastic Wing airplane was successfully flight-tested in March 2005. During phase 1 of the two-phase program, an onboard excitation system provided independent control surface movements that were used to develop a loads model for the wing structure and wing control surfaces. The resulting loads model, which was used to develop the control laws for phase 2, is described. The loads model was developed from flight data through the use of a multiple linear regression technique. The loads model input consisted of aircraft states and control surface positions, in addition to nonlinear inputs that were calculated from flight-measured parameters. The loads model output for each wing consisted of wing-root bending moment and torque, wing-fold bending moment and torque, inboard and outboard leading-edge flap hinge moment, trailing-edge flap hinge moment, and aileron hinge moment. The development of the Active Aeroelastic Wing loads model is described, and the ability of the model to predict loads during phase 2 research maneuvers is demonstrated. Results show a good match to phase 2 flight data for all loads except inboard and outboard leading-edge flap hinge moments at certain flight conditions. The average load prediction errors for all loads at all flight conditions are 9.1 percent for maximum stick-deflection rolls, 4.4 percent for 5-g windup turns, and 7.7 percent for 4-g rolling pullouts.

Flight investigation of the use of a nose gear jump strut to reduce takeoff ground roll distance of STOL aircraft

NASA Technical Reports Server (NTRS)

Eppel, Joseph C.; Hardy, Gordon; Martin, James L.

1994-01-01

A series of flight tests was conducted to evaluate the reduction of takeoff ground roll distance obtainable from a rapid extension of the nose gear strut. The NASA Quiet Short-haul Research Aircraft (QSRA) used for this investigation is a transport-size short takeoff and landing (STOL) research vehicle with a slightly swept wing that employs the upper surface blowing (USB) concept to attain the high lift levels required for its low speed, short-field performance. Minor modifications to the conventional nose gear assembly and the addition of a high pressure pneumatic system and a control system provided the extendible nose gear, or 'jump strut,' capability. The limited flight test program explored the effects of thrust-to-weight ratio, storage tank initial pressure, and control valve open time duration on the ground roll distance. The data show that the predicted reduction of takeoff ground roll on the order of 10 percent was achieved with the use of the jump strut. Takeoff performance with the jump strut was also found to be essentially independent of the pneumatic supply pressure and was only slightly affected by control valve open time within the range of the parameters examined.

An analysis of available data on effects of wing-fuselage-tail and wing-nacelle interference on the distribution of the air load among components of airplanes

NASA Technical Reports Server (NTRS)

Wollner, Bertram C

1949-01-01

Available information on the effects of wing-fuselage-tail and wing-nacelle interference on the distribution of the air load among components of airplanes is analyzed. The effects of wing and nacelle incidence, horizontal andvertical position of wing and nacelle, fuselage shape, wing section and filleting are considered. Where sufficient data were unavailable to determine the distribution of the air load, the change in lift caused by interference between wing and fuselage was found. This increment is affected to the greatest extent by vertical wing position.

Loads calibrations of strain gage bridges on the DAST project Aeroelastic Research Wing (ARW-1)

NASA Technical Reports Server (NTRS)

Eckstrom, C. V.

1980-01-01

The details of and results from the procedure used to calibrate strain gage bridges for measurement of wing structural loads for the DAST project ARW-1 wing are presented. Results are in the form of loads equations and comparison of computed loads vs. actual loads for two simulated flight loading conditions.

Flight Speeds among Bird Species: Allometric and Phylogenetic Effects

PubMed Central

Alerstam, Thomas; Rosén, Mikael; Bäckman, Johan; Ericson, Per G. P; Hellgren, Olof

2007-01-01

Flight speed is expected to increase with mass and wing loading among flying animals and aircraft for fundamental aerodynamic reasons. Assuming geometrical and dynamical similarity, cruising flight speed is predicted to vary as (body mass)1/6 and (wing loading)1/2 among bird species. To test these scaling rules and the general importance of mass and wing loading for bird flight speeds, we used tracking radar to measure flapping flight speeds of individuals or flocks of migrating birds visually identified to species as well as their altitude and winds at the altitudes where the birds were flying. Equivalent airspeeds (airspeeds corrected to sea level air density, U e) of 138 species, ranging 0.01–10 kg in mass, were analysed in relation to biometry and phylogeny. Scaling exponents in relation to mass and wing loading were significantly smaller than predicted (about 0.12 and 0.32, respectively, with similar results for analyses based on species and independent phylogenetic contrasts). These low scaling exponents may be the result of evolutionary restrictions on bird flight-speed range, counteracting too slow flight speeds among species with low wing loading and too fast speeds among species with high wing loading. This compression of speed range is partly attained through geometric differences, with aspect ratio showing a positive relationship with body mass and wing loading, but additional factors are required to fully explain the small scaling exponent of U e in relation to wing loading. Furthermore, mass and wing loading accounted for only a limited proportion of the variation in U e. Phylogeny was a powerful factor, in combination with wing loading, to account for the variation in U e. These results demonstrate that functional flight adaptations and constraints associated with different evolutionary lineages have an important influence on cruising flapping flight speed that goes beyond the general aerodynamic scaling effects of mass and wing loading. PMID:17645390

Fuel-conservative guidance system for powered-lift aircraft

NASA Technical Reports Server (NTRS)

Erzberger, H.; Mclean, J. D.

1979-01-01

A concept for automatic terminal area guidance, comprising two modes of operation, was developed and evaluated in flight tests. In the predictive mode, fuel efficient approach trajectories are synthesized in fast time. In the tracking mode, the synthesized trajectories are reconstructed and tracked automatically. An energy rate performance model derived from the lift, drag, and propulsion system characteristics of the aircraft is used in the synthesis algorithm. The method optimizes the trajectory for the initial aircraft position and wind and temperature profiles encountered during each landing approach. The design theory and the results of simulations and flight tests using the Augmentor Wing Jet STOL Research Aircraft are described.

Study of aerodynamic technology for VSTOL fighter attack aircraft

NASA Technical Reports Server (NTRS)

Burhans, W., Jr.; Crafta, V. J., Jr.; Dannenhoffer, N.; Dellamura, F. A.; Krepski, R. E.

1978-01-01

Vertical short takeoff aircraft capability, supersonic dash capability, and transonic agility were investigated for the development of Fighter/attack aircraft to be accommodated on ships smaller than present aircraft carriers. Topics covered include: (1) description of viable V/STOL fighter/attack configuration (a high wing, close-coupled canard, twin-engine, control configured aircraft) which meets or exceeds specified levels of vehicle performance; (2) estimates of vehicle aerodynamic characteristics and the methodology utilized to generate them; (3) description of propulsion system characteristics and vehicle mass properties; (4) identification of areas of aerodynamic uncertainty; and (5) a test program to investigate the areas of aerodynamic uncertainty in the conventional flight mode.

Static noise tests on modified augmentor wing jet STOL research aircraft

NASA Technical Reports Server (NTRS)

Cook, G. R.; Lilley, B. F.

1981-01-01

Noise measurements were made to determine if recent modifications made to the bifurcated jetpipe to increase engine thrust had at the same time reduced the noise level. The noise field was measured by a 6-microphone array positioned on a 30.5m (100 ft) sideline between 90 and 150 degrees from the left engine inlet. Noise levels were recorded at three flap angles over a range of engine thrust settings from flight idle to emergency power and plotted in one-third octave band spectra. Little attenuation was observed at maximum power, but significant attenuation was achieved at approach and cruise power levels.

Further Studies of Aerodynamic Loads at Spin Entry

DTIC Science & Technology

1977-06-30

the model and the loads on the model for a research coafiguration having certain essential features of a modern fighter-bomber aircraft, since no such...percent at the tip. The model is geometrically similar to that described in reference 3 but approximately 2.4 times larger. The NASA model is designed to...measurements turned out to be extremely time consuming. Even though the V/STOL tunnel staff granted three additional days over the 10 originally scheduled for

78 FR 31851 - Harmonization of Airworthiness Standards-Gust and Maneuver Load Requirements

Federal Register 2010, 2011, 2012, 2013, 2014

2013-05-28

... airplanes equipped with wing-mounted engines; revise the engine torque loads criteria; add an engine failure... equipped with wing-mounted engines. Following an accident in which an airplane shed a large wing- mounted...-93-137, November 15, 1993). This recommendation was specifically aimed at gust loads on wing-mounted...

Forward velocity effects on fan noise and the suppression characteristics of advanced inlets as measured in the NASA-Ames 40 by 80 foot wind tunnel

NASA Technical Reports Server (NTRS)

Moore, M. T.

1980-01-01

Forward velocity effects on the forward radiated fan noise and on the suppression characteristics of three advanced inlets relative to a baseline cylindrical inlet were measured in the NASA Ames Research Center 40 x 80 foot Wind Tunnel. A modified JT15D turbofan engine in a quiet nacelle was the source of fan noise; the advanced inlets were a Conventional Takeoff/Landing (CTOL) hybrid inlet, a Short Takeoff/Landing (STOL) hybrid inlet, and a treated deflector inlet. Also measured were the static to flight effects on the fan noise of canting the baseline inlet 4 deg downward to simulate typical wing mounted turbofan engines. The CTOL hybrid inlet suppressed the high tip speed fan noise as much as 18 PNdB on a 61 m (200 ft) sideline scaled to a CF6 size engine while the STOL hybrid inlet suppressed the low tip speed fan noise as much as 13 PNdB on a 61 m (200 ft) sideline scaled to a OCSEE size engine. The deflector inlet suppressed the high tip speed fan noise as much as 13 PNdB at 61 m (200 ft) overhead scaled to a CF6 size engine. No significant changes in fan noise suppression for the CTOL and STOL hybrid inlets occurred for forward velocity changes above 21 m/s (68 ft/s) or for angle of attack changes up to 15 deg. However, changes in both forward velocity and angle of attack changed the deflector inlet noise unpredictably due to the asymmetry of the inlet flow field into the fan.

Effectiveness evaluation of STOL transport operations

NASA Technical Reports Server (NTRS)

Hitt, E. F.; Bruckner, J. M. H.; Drago, V. J.; Brown, R. A.; Rea, F. G.; Porter, R. F.

1973-01-01

A short-takeoff and landing (STOL) systems simulation model has been developed and implemented in a computer code (known as STOL OPS) which permits evaluation of the operation of a STOL aircraft and its avionics in a commercial airline operating environment. STOL OPS concentrated on the avionics functions of navigation, guidance, control, communication, hazard aviodance, and systems management. External world factors influencing the operation of the STOL aircraft include each airport and its geometry, air traffic at each airport, air traffic control equipment and procedures, weather (including winds and visibility), and the flight path between each airport served by the route. The development of the STOL OPS program provides NASA a set of computer programs which can be used for detailed analysis of a STOL aircraft and its avionics and permit establishment of system requirements as a function of airline mission performance goals.

A Flight Investigation of the STOL Characteristics of an Augmented Jet Flap STOL Research Aircraft

NASA Technical Reports Server (NTRS)

Quigley, H. C.; Innis, R. C.; Grossmith, S.

1974-01-01

The flight test program objectives are: (1) To determine the in-flight aerodynamic, performance, and handling qualities of a jet STOL aircraft incorporating the augmented jet flap concept; (2) to compare the results obtained in flight with characteristics predicted from wind tunnel and simulator test results; (3) to contribute to the development of criteria for design and operation of jet STOL transport aircraft; and (4) to provide a jet STOL transport aircraft for STOL systems research and development. Results obtained during the first 8 months of proof-of-concept flight testing of the aircraft in STOL configurations are reported. Included are a brief description of the aircraft, fan-jet engines, and systems; a discussion of the aerodynamic, stability and control, and STOL performance; and pilot opinion of the handling qualities and operational characteristics.

Test stand performance of a convertible engine for advanced V/STOL and rotorcraft propulsion

NASA Technical Reports Server (NTRS)

Mcardle, Jack G.

1987-01-01

A variable inlet guide vane (VIGV) convertible engine that could be used to power future high-speed V/STOL and rotorcraft was tested on an outdoor stand. The engine ran stably and smoothly in the turbofan, turboshaft, and dual (combined fan and shaft) power modes. In the turbofan mode with the VIGV open, fuel consumption was comparable to that of a conventional turbofan engine. In the turboshaft mode with the VIGV closed, fuel consumption was higher than that of present turboshaft engines because power was wasted in churning fan-tip air flow. In dynamic performance tests with a specially built digital engine control and using a waterbrake dynamometer for shaft load, the engine responded effectively to large steps in thrust command and shaft torque.

Test stand performance of a convertible engine for advanced V/STOL and rotorcraft propulsion

NASA Technical Reports Server (NTRS)

Mcardle, Jack G.

1988-01-01

A variable inlet guide vane (VIGV) convertible engine that could be used to power future high-speed V/STOL and rotorcraft was tested on an outdoor stand. The engine ran stably and smoothly in the turbofan, turboshaft, and dual (combined fan and shaft) power modes. In the turbofan mode with the VIGV open, fuel consumption was comparable to that of a conventional turbofan engine. In the turboshaft mode with the VIGV closed, fuel consumption was higher than that of present turboshaft engines because power was wasted in churning fan-tip air flow. In dynamic performance tests with a specially built digital engine control and using a waterbrake dynamometer for shaft load, the engine responded effectively to large steps in thrust command and shaft torque.

Evaluation of Load Analysis Methods for NASAs GIII Adaptive Compliant Trailing Edge Project

NASA Technical Reports Server (NTRS)

Cruz, Josue; Miller, Eric J.

2016-01-01

The Air Force Research Laboratory (AFRL), NASA Armstrong Flight Research Center (AFRC), and FlexSys Inc. (Ann Arbor, Michigan) have collaborated to flight test the Adaptive Compliant Trailing Edge (ACTE) flaps. These flaps were installed on a Gulfstream Aerospace Corporation (GAC) GIII aircraft and tested at AFRC at various deflection angles over a range of flight conditions. External aerodynamic and inertial load analyses were conducted with the intention to ensure that the change in wing loads due to the deployed ACTE flap did not overload the existing baseline GIII wing box structure. The objective of this paper was to substantiate the analysis tools used for predicting wing loads at AFRC. Computational fluid dynamics (CFD) models and distributed mass inertial models were developed for predicting the loads on the wing. The analysis tools included TRANAIR (full potential) and CMARC (panel) models. Aerodynamic pressure data from the analysis codes were validated against static pressure port data collected in-flight. Combined results from the CFD predictions and the inertial load analysis were used to predict the normal force, bending moment, and torque loads on the wing. Wing loads obtained from calibrated strain gages installed on the wing were used for substantiation of the load prediction tools. The load predictions exhibited good agreement compared to the flight load results obtained from calibrated strain gage measurements.

Method for calculating the aerodynamic loading on an oscillating finite wing in subsonic and sonic flow

NASA Technical Reports Server (NTRS)

Runyan, Harry L; Woolston, Donald S

1957-01-01

A method is presented for calculating the loading on a finite wing oscillating in subsonic or sonic flow. The method is applicable to any plan form and may be used for determining the loading on deformed wings. The procedure is approximate and requires numerical integration over the wing surface.

V/STOL flight simulation

NASA Technical Reports Server (NTRS)

1979-01-01

The requirements for a new research aircraft to provide in-flight V/STOL simulation were reviewed. The required capabilities were based on known limitations of ground based simulation and past/current experience with V/STOL inflight simulation. Results indicate that V/STOL inflight simulation capability is needed to aid in the design and development of high performance V/STOL aircraft. Although a new research V/STOL aircraft is preferred, an interim solution can be provided by use of the X-22A, the CH-47B, or the 4AV-8B aircraft modified for control/display flight research.

Accelerated development and flight evaluation of active controls concepts for subsonic transport aircraft. Volume 1: Load alleviation/extended span development and flight tests

NASA Technical Reports Server (NTRS)

Johnston, J. F.

1979-01-01

Active wing load alleviation to extend the wing span by 5.8 percent, giving a 3 percent reduction in cruise drag is covered. The active wing load alleviation used symmetric motions of the outboard ailerons for maneuver load control (MLC) and elastic mode suppression (EMS), and stabilizer motions for gust load alleviation (GLA). Slow maneuvers verified the MLC, and open and closed-loop flight frequency response tests verified the aircraft dynamic response to symmetric aileron and stabilizer drives as well as the active system performance. Flight tests in turbulence verified the effectiveness of the active controls in reducing gust-induced wing loads. It is concluded that active wing load alleviation/extended span is proven in the L-1011 and is ready for application to airline service; it is a very practical way to obtain the increased efficiency of a higher aspect ratio wing with minimum structural impact.

A review of technologies applicable to low-speed flight of high-performance aircraft investigated in the Langley 14- x 22-foot subsonic tunnel

NASA Technical Reports Server (NTRS)

Paulson, John W., Jr.; Quinto, P. Frank; Banks, Daniel W.; Kemmerly, Guy T.; Gatlin, Gregory M.

1988-01-01

An extensive research program has been underway at the NASA Langley Research Center to define and develop the technologies required for low-speed flight of high-performance aircraft. This 10-year program has placed emphasis on both short takeoff and landing (STOL) and short takeoff and vertical landing (STOVL) operations rather than on regular up and away flight. A series of NASA in-house as well as joint projects have studied various technologies including high lift, vectored thrust, thrust-induced lift, reversed thrust, an alternate method of providing trim and control, and ground effects. These technologies have been investigated on a number of configurations ranging from industry designs for advanced fighter aircraft to generic wing-canard research models. Test conditions have ranged from hover (or static) through transition to wing-borne flight at angles of attack from -5 to 40 deg at representative thrust coefficients.

Strain Gage Loads Calibration Testing with Airbag Support for the Gulfstream III SubsoniC Research Aircraft Testbed (SCRAT)

NASA Technical Reports Server (NTRS)

Lokos, William; Miller, Eric; Hudson, Larry; Holguin, Andrew; Neufeld, David; Haraguchi, Ronnie

2015-01-01

This paper describes the design and conduct of the strain gage load calibration ground test of the SubsoniC Research Aircraft Testbed, Gulfstream III aircraft, and the subsequent data analysis and its results. The goal of this effort was to create and validate multi-gage load equations for shear force, bending moment, and torque for two wing measurement stations. For some of the testing the aircraft was supported by three air bags in order to isolate the wing structure from extraneous load inputs through the main landing gear. Thirty-two strain gage bridges were installed on the left wing. Hydraulic loads were applied to the wing lower surface through a total of 16 load zones. Some dead weight load cases were applied to the upper wing surface using shot bags. Maximum applied loads reached 54,000 pounds.

Investigation of the Effect of Tip Tanks on the Wing Loading of a Republic F-84 Airplane in the Ames 40- by 80-foot Wind Tunnel

NASA Technical Reports Server (NTRS)

Hunton, Lynn W.; Dew, Joseph K.; Salisbury, Ralph D.

1949-01-01

Wind-tunnel tests at low Mach number of a Republic F-84C airplane were conducted to determine by pressure-distribution measurements the air loads on wing-tip tanks and the change in wing load distribution due to the presence of tip tanks. Measurements of the aeroelastic twist of the wing were also obtained. Results are presented in the form of loading coefficient, center-of- pressure location, pitching-moment coefficient, aerodynamic-center location, and aeroelastic twist. The investigation revealed that the redistributions in loading brought about by either the tip tanks or elastic deformation of the wing were relatively small when compared with the chnnges in loading normally associated with the deflection of an aileron.

Comparison of wing-span averaging effects on lift, rolling moment, and bending moment for two span load distributions and for two turbulence representations

NASA Technical Reports Server (NTRS)

Lichtenstein, J. H.

1978-01-01

An analytical method of computing the averaging effect of wing-span size on the loading of a wing induced by random turbulence was adapted for use on a digital electronic computer. The turbulence input was assumed to have a Dryden power spectral density. The computations were made for lift, rolling moment, and bending moment for two span load distributions, rectangular and elliptic. Data are presented to show the wing-span averaging effect for wing-span ratios encompassing current airplane sizes. The rectangular wing-span loading showed a slightly greater averaging effect than did the elliptic loading. In the frequency range most bothersome to airplane passengers, the wing-span averaging effect can reduce the normal lift load, and thus the acceleration, by about 7 percent for a typical medium-sized transport. Some calculations were made to evaluate the effect of using a Von Karman turbulence representation. These results showed that using the Von Karman representation generally resulted in a span averaging effect about 3 percent larger.

Minimization theory of induced drag subject to constraint conditions

NASA Technical Reports Server (NTRS)

Deyoung, J.

1979-01-01

Exact analytical solutions in terms of induced drag influence coefficients can be attained which define the spanwise loading with minimized induced drag, subject to specified constraint conditions, for any nonplanar wing shape or number of lift plus wing bending moment about a given wing span station. Example applications of the theory are made to a biplane, a wing in ground effect, a cruciform wing, a V-wing, a planar-wing winglet, and linked wingtips in formation flying. For minimal induced drag, the spanwise loading, relative to elliptic, is outboard for the biplane and is inboard for the wing in ground effect and for the planar-wing winglet. A spinoff of the triplane solution provides mathematically exact equations for downwash and sidewash about a planar vorticity sheet having an arbitrary loading distribution.

Advanced prediction technique for the low speed aerodynamics of V/STOL aircraft. Volume 1: Technical discussion

NASA Technical Reports Server (NTRS)

Beatty, T. D.; Worthey, M. K.

1984-01-01

The V/STOL Aircraft Propulsive Effects (VAPE) computerized prediction method is evaluated. The program analyzes viscous effects, various jet, inlet, and Short TakeOff and Landing (STOL) models, and examines the aerodynamic configurations of V/STOL aircraft.

Strain Gage Loads Calibration Testing with Airbag Support for the Gulfstream III SubsoniC Research Aircraft Testbed (SCRAT)

NASA Technical Reports Server (NTRS)

Lokos, William A.; Miller, Eric J.; Hudson, Larry D.; Holguin, Andrew C.; Neufeld, David C.; Haraguchi, Ronnie

2015-01-01

This paper describes the design and conduct of the strain-gage load calibration ground test of the SubsoniC Research Aircraft Testbed, Gulfstream III aircraft, and the subsequent data analysis and results. The goal of this effort was to create and validate multi-gage load equations for shear force, bending moment, and torque for two wing measurement stations. For some of the testing the aircraft was supported by three airbags in order to isolate the wing structure from extraneous load inputs through the main landing gear. Thirty-two strain gage bridges were installed on the left wing. Hydraulic loads were applied to the wing lower surface through a total of 16 load zones. Some dead-weight load cases were applied to the upper wing surface using shot bags. Maximum applied loads reached 54,000 lb. Twenty-six load cases were applied with the aircraft resting on its landing gear, and 16 load cases were performed with the aircraft supported by the nose gear and three airbags around the center of gravity. Maximum wing tip deflection reached 17 inches. An assortment of 2, 3, 4, and 5 strain-gage load equations were derived and evaluated against independent check cases. The better load equations had root mean square errors less than 1 percent. Test techniques and lessons learned are discussed.

Nonplanar wing load-line and slender wing theory

NASA Technical Reports Server (NTRS)

Deyoung, J.

1977-01-01

Nonplanar load line, slender wing, elliptic wing, and infinite aspect ratio limit loading theories are developed. These are quasi two dimensional theories but satisfy wing boundary conditions at all points along the nonplanar spanwise extent of the wing. These methods are applicable for generalized configurations such as the laterally nonplanar wing, multiple nonplanar wings, or wing with multiple winglets of arbitrary shape. Two dimensional theory infers simplicity which is practical when analyzing complicated configurations. The lateral spanwise distribution of angle of attack can be that due to winglet or control surface deflection, wing twist, or induced angles due to multiwings, multiwinglets, ground, walls, jet or fuselage. In quasi two dimensional theory the induced angles due to these extra conditions are likewise determined for two dimensional flow. Equations are developed for the normal to surface induced velocity due to a nonplanar trailing vorticity distribution. Application examples are made using these methods.

STOL Traffic environment and operational procedures

NASA Technical Reports Server (NTRS)

Schlundt, R. W.; Dewolf, R. W.; Ausrotas, R. A.; Curry, R. E.; Demaio, D.; Keene, D. W.; Speyer, J. L.; Weinreich, M.; Zeldin, S.

1972-01-01

The expected traffic environment for an intercity STOL transportation system is examined, and operational procedures are discussed in order to identify problem areas which impact STOL avionics requirements. Factors considered include: traffic densities, STOL/CTOL/VTOL traffic mix, the expect ATC environment, aircraft noise models and community noise models and community noise impact, flight paths for noise abatement, wind considerations affecting landing, approach and landing considerations, STOLport site selection, runway capacity, and STOL operations at jetports, suburban airports, and separate STOLports.

Study of quiet turbofan STOL aircraft for short-haul transportation. Volume 4: Markets

NASA Technical Reports Server (NTRS)

1973-01-01

A marketing study to determine the acceptance and utilization of a STOL aircraft short-haul air transportation system was conducted. The relationship between STOL characteristics and the economic and social viability of STOL as a short-haul reliever system was examined. A study flow chart was prepared to show the city pair and traffic split analysis. The national demand for STOL aircraft, as well as the foreign and military markets, were analyzed.

Reduction of structural loads using maneuver load control on the Advanced Fighter Technology Integration (AFTI)/F-111 mission adaptive wing

NASA Technical Reports Server (NTRS)

Thornton, Stephen V.

1993-01-01

A transonic fighter-bomber aircraft, having a swept supercritical wing with smooth variable-camber flaps was fitted with a maneuver load control (MLC) system that implements a technique to reduce the inboard bending moments in the wing by shifting the spanwise load distribution inboard as load factor increases. The technique modifies the spanwise camber distribution by automatically commanding flap position as a function of flap position, true airspeed, Mach number, dynamic pressure, normal acceleration, and wing sweep position. Flight test structural loads data were obtained for loads in both the wing box and the wing root. Data from uniformly deflected flaps were compared with data from flaps in the MLC configuration where the outboard segment of three flap segments was deflected downward less than the two inboard segments. The changes in the shear loads in the forward wing spar and at the roots of the stabilators also are presented. The camber control system automatically reconfigures the flaps through varied flight conditions. Configurations having both moderate and full trailing-edge flap deflection were tested. Flight test data were collected at Mach numbers of 0.6, 0.7, 0.8, and 0.9 and dynamic pressures of 300, 450, 600, and 800 lb/sq ft. The Reynolds numbers for these flight conditions ranged from 26 x 10(exp 6) to 54 x 10(exp 6) at the mean aerodynamic chord. Load factor increases of up to 1.0 g achieved with no increase in wing root bending moment with the MLC flap configuration.

Aerodynamic comparison of a butterfly-like flapping wing-body model and a revolving-wing model

NASA Astrophysics Data System (ADS)

Suzuki, Kosuke; Yoshino, Masato

2017-06-01

The aerodynamic performance of flapping- and revolving-wing models is investigated by numerical simulations based on an immersed boundary-lattice Boltzmann method. As wing models, we use (i) a butterfly-like model with a body and flapping-rectangular wings and (ii) a revolving-wing model with the same wings as the flapping case. Firstly, we calculate aerodynamic performance factors such as the lift force, the power, and the power loading of the two models for Reynolds numbers in the range of 50-1000. For the flapping-wing model, the power loading is maximal for the maximum angle of attack of 90°, a flapping amplitude of roughly 45°, and a phase shift between the flapping angle and the angle of attack of roughly 90°. For the revolving-wing model, the power loading peaks for an angle of attack of roughly 45°. In addition, we examine the ground effect on the aerodynamic performance of the revolving-wing model. Secondly, we compare the aerodynamic performance of the flapping- and revolving-wing models at their respective maximal power loadings. It is found that the revolving-wing model is more efficient than the flapping-wing model both when the body of the latter is fixed and where it can move freely. Finally, we discuss the relative agilities of the flapping- and revolving-wing models.

Strain Gage Load Calibration of the Wing Interface Fittings for the Adaptive Compliant Trailing Edge Flap Flight Test

NASA Technical Reports Server (NTRS)

Miller, Eric J.; Holguin, Andrew C.; Cruz, Josue; Lokos, William A.

2014-01-01

This is the presentation to follow conference paper of the same name. The adaptive compliant trailing edge (ACTE) flap experiment safety of flight requires that the flap to wing interface loads be sensed and monitored in real time to ensure that the wing structural load limits are not exceeded. This paper discusses the strain gage load calibration testing and load equation derivation methodology for the ACTE interface fittings. Both the left and right wing flap interfaces will be monitored and each contains four uniquely designed and instrumented flap interface fittings. The interface hardware design and instrumentation layout are discussed. Twenty one applied test load cases were developed using the predicted in-flight loads for the ACTE experiment.

Recent Loads Calibration Experience With a Delta Wing Airplane

NASA Technical Reports Server (NTRS)

Jenkins, Jerald M.; Kuhl, Albert E.

1977-01-01

Aircraft which are designed for supersonic and hypersonic flight are evolving with delta wing configurations. An integral part of the evolution of all new aircraft is the flight test phase. Included in the flight test phase is an effort to identify and evaluate the loads environment of the aircraft. The most effective way of examining the loads environment is to utilize calibrated strain gages to provide load magnitudes. Using strain gage data to accomplish this has turned out to be anything but a straightforward task. The delta wing configuration has turned out to be a very difficult type of wing structure to calibrate. Elevated structural temperatures result in thermal effects which contaminate strain gage data being used to deduce flight loads. The concept of thermally calibrating a strain gage system is an approach to solving this problem. This paper will address how these problems were approached on a program directed toward measuring loads on the wing of a large, flexible supersonic aircraft. Structural configurations typical of high-speed delta wing aircraft will be examined. The temperature environment will be examined to see how it induces thermal stresses which subsequently cause errors in loads equations used to deduce the flight loads.

Fuel-conservative guidance system for powered-lift aircraft

NASA Technical Reports Server (NTRS)

Erzberger, H.; Mclean, J. D.

1979-01-01

A concept for automatic terminal-area guidance, comprising two modes of operation, has been developed and evaluated in flight tests. In the first or predictive mode, fuel-efficient approach trajectories are synthesized in fast time. In the second or tracking mode, the synthesized trajectories are reconstructed and tracked automatically. An energy rate performance model derived from the lift, drag, and propulsion-system characteristics of the aircraft is used in the synthesis algorithm. The method optimizes the trajectory for the initial aircraft position and wind and temperature profiles encountered during each landing approach. The paper describes the design theory and discusses the results of simulations and flight tests using the Augmentor Wing Jet STOL Research Aircraft.

V/STOL tilt rotor study. Volume 5: A mathematical model for real time flight simulation of the Bell model 301 tilt rotor research aircraft

NASA Technical Reports Server (NTRS)

Harendra, P. B.; Joglekar, M. J.; Gaffey, T. M.; Marr, R. L.

1973-01-01

A mathematical model for real-time flight simulation of a tilt rotor research aircraft was developed. The mathematical model was used to support the aircraft design, pilot training, and proof-of-concept aspects of the development program. The structure of the mathematical model is indicated by a block diagram. The mathematical model differs from that for a conventional fixed wing aircraft principally in the added requirement to represent the dynamics and aerodynamics of the rotors, the interaction of the rotor wake with the airframe, and the rotor control and drive systems. The constraints imposed on the mathematical model are defined.

Top-mounted inlet system feasibility for transonic-supersonic fighter aircraft. [V/STOL aircraft

NASA Technical Reports Server (NTRS)

Williams, T. L.; Hunt, B. L.; Smeltzer, D. B.; Nelms, W. P.

1981-01-01

The more salient findings are presented of recent top inlet performance evaluations aimed at assessing the feasibility of top-mounted inlet systems for transonic-supersonic fighter aircraft applications. Top inlet flow field and engine-inlet performance test data show the influence of key aircraft configuration variables-inlet longitudinal position, wing leading-edge extension planform area, canopy-dorsal integration, and variable incidence canards-on top inlet performance over the Mach range of 0.6 to 2.0. Top inlet performance data are compared with those or more conventional inlet/airframe integrations in an effort to assess the viability of top-mounted inlet systems relative to conventional inlet installations.

Implementation of an optimum profile guidance system on STOLAND

NASA Technical Reports Server (NTRS)

Flanagan, P. F.

1978-01-01

The implementation on the STOLAND airborne digital computer of an optimum profile guidance system for the augmentor wing jet STOL research aircraft is described. Major tasks were to implement the guidance and control logic to airborne computer software and to integrate the module with the existing STOLAND navigation, display, and autopilot routines. The optimum profile guidance system comprises an algorithm for synthesizing mimimum fuel trajectories for a wide range of starting positions in the terminal area and a control law for flying the aircraft automatically along the trajectory. The avionics software developed is described along with a FORTRAN program that was constructed to reflect the modular nature and algorthms implemented in the avionics software.

Analysis of Effect of Rolling Pull-Outs on Wing and Aileron Loads of a Fighter Airplane

NASA Technical Reports Server (NTRS)

Pearson, Henry A.; Aiken, William S.

1946-01-01

An analysis was made to determine the effect of rolling pull-out maneuvers on the wing and aileron loads of a typical fighter airplane, the P-47B. The results obtained indicate that higher loads are imposed upon wings and ailerons because of the rolling pull-out maneuver, than would be obtained by application of the loading requirements to which the airplane was designed. An increase of 102 lb or 15 percent of wing weight would be required if the wing were designed for rolling pull-out maneuver. It was also determined that the requirements by which the aileron was originally designed were inadequate.

77 FR 45515 - Airworthiness Directives; The Boeing Company Airplanes

Federal Register 2010, 2011, 2012, 2013, 2014

2012-08-01

... could result in the wing structure not supporting the limit load condition, which could lead to loss of... the limit load condition, which could lead to loss of the structural integrity of the wing. Relevant... could result in the wing structure not supporting the limit load condition, which could lead to loss of...

Seakeeping considerations in the employment of V/STOL on Naval ships

NASA Technical Reports Server (NTRS)

Olson, S. R.

1977-01-01

Compatibility of Naval ships as V/STOL support platforms and the ship motions that V/STOL aircraft must endure are discussed. A methodology which evaluates the impact of motion criteria such as the maximum ship motion allowable during V/STOL landing/launch is presented. Emphasis is given to design alternatives that reduce ship motion.

Terminal-area STOL operating systems experiments program

NASA Technical Reports Server (NTRS)

Smith, D. W.; Watson, D.; Christensen, J. V.

1973-01-01

Information which will aid in the choice by the U.S. Government and industry of system concepts, design criteria, operating procedures for STOL aircraft and STOL ports, STOL landing guidance systems, air traffic control systems, and airborne avionics and flight control systems. Ames has developed a terminal-area STOL operating systems experiments program which is a part of the joint DOT/NASA effort is discussed. The Ames operating systems experiments program, its objectives, the program approach, the program schedule, typical experiments, the research facilities to be used, and the program status are described.

Summary of the recent short-haul systems studies

NASA Technical Reports Server (NTRS)

Savin, R. C.; Galloway, T. L.; Wilcox, D. E.; Kenyon, G. C.; Ardema, M. D.; Waters, M. H.

1975-01-01

The results of several NASA sponsored high density short haul air transportation systems studies are reported as well as analyzed. Included are the total STOL systems analysis approach, a companion STOL composites study conducted in conjunction with STOL systems studies, a STOL economic assessment study, an evaluation of STOL aircraft with and without externally blown flaps, an alternative STOL systems for the San Francisco Bay Area, and the quiet, clean experimental engine studies. Assumptions and results of these studies are summarized, their differences, analyzed, and the results compared with those in-house analyses performed by the Systems Studies Division of the NASA-Ames Research Center. Pertinent conclusions are developed and the more significant technology needs for the evaluation of a viable short haul transportation system are identified.

Design of a Large Span-Distributed Load Flying-Wing Cargo Airplane

NASA Technical Reports Server (NTRS)

Jernell, L. S.; Quartero, C. B.

1977-01-01

The design and operation of very large, long-range, subsonic cargo aircraft are considered. A design concept which distributes the payload along the wingspan to counterbalance the aerodynamic loads, with a resultant decrease in the in-flight wing bending moments and shear forces, is described. The decreased loading of the wing structure, coupled with the very thick wing housing the cargo, results in a relatively low overall structural weight in comparison to that of conventional aircraft.

An experimental comparative study of the efficiency of twisted and flat flapping wings during hovering flight.

PubMed

Phan, Hoang Vu; Truong, Quang Tri; Park, Hoon Cheol

2017-04-19

This work presents a parametric study to find a proper wing configuration for achieving economical flight using unsteady blade element theory, which is based on the 3D kinematics of a flapping wing. Power loading was first considered as a performance parameter for the study. The power loadings at each wing section along the wingspan were obtained for various geometric angles of attack (AoAs) by calculating the ratios of the vertical forces generated and the power consumed by that particular wing section. The results revealed that the power loading of a negatively twisted wing could be higher than the power loading that a flat wing can have; the power loading of the negatively twisted wing was approximately 5.9% higher. Given the relatively low average geometric AoA (α A,root   ≈  44° and α A,tip   ≈  25°), the vertical force produced by the twisted wing for the highest power loading was approximately 24.4% less than that produced by the twisted wing for the strongest vertical force. Therefore, for a given wing geometry and flapping amplitude, a flapping-wing micro air vehicle required a 13.5% increase in flapping frequency to generate the same strongest cycle-average vertical force while saving about 24.3% power. However, when force 3 /power 2 and force 2 /power ratios were considered as performance indices, the twisted wings for the highest force 3 /power 2 (α A,root   ≈  43° and α A,tip   ≈  30°) and force 2 /power (α A,root   ≈  43° and α A,tip   ≈  36°) required only 6.5% and 4% increases in flapping frequency and consumed 26.2% and 25.3% less power, respectively. Thus, it is preferable to use a flapping wing operating at a high frequency using the geometric AoAs for the highest power loading, force 3 /power 2 ratio, and force 2 /power ratio over a flapping wing operating at a low frequency using a high geometric AoA with the strongest vertical force. Additionally, by considering both aerodynamic and inertial forces, this study obtained average geometric AoAs in the range of 30° to 40°, which are similar to those of a typical hovering insect's wings. Therefore, the operation of an aerodynamically uneconomical, high AoA in a hovering insect's wings during flight is explainable.

Active Dihedral Control System for a Torsionally Flexible Wing

NASA Technical Reports Server (NTRS)

Morgan, Walter R. (Inventor); Kendall, Greg T. (Inventor); Lisoski, Derek L. (Inventor); Griecci, John A. (Inventor)

2017-01-01

A span-loaded, highly flexible flying wing, having horizontal control surfaces mounted aft of the wing on extended beams to form local pitch-control devices. Each of five spanwise wing segments of the wing has one or more motors and photovoltaic arrays, and produces its own lift independent of the other wing segments, to minimize inter-segment loads. Wing dihedral is controlled by separately controlling the local pitch-control devices consisting of a control surface on a boom, such that inboard and outboard wing segment pitch changes relative to each other, and thus relative inboard and outboard lift is varied.

Active Dihedral Control System for a Torisionally Flexible Wing

NASA Technical Reports Server (NTRS)

Kendall, Greg T. (Inventor); Lisoski, Derek L. (Inventor); Morgan, Walter R. (Inventor); Griecci, John A. (Inventor)

2015-01-01

A span-loaded, highly flexible flying wing, having horizontal control surfaces mounted aft of the wing on extended beams to form local pitch-control devices. Each of five spanwise wing segments of the wing has one or more motors and photovoltaic arrays, and produces its own lift independent of the other wing segments, to minimize inter-segment loads. Wing dihedral is controlled by separately controlling the local pitch-control devices consisting of a control surface on a boom, such that inboard and outboard wing segment pitch changes relative to each other, and thus relative inboard and outboard lift is varied.

Aircraft control system

NASA Technical Reports Server (NTRS)

Kendall, Greg T. (Inventor); Morgan, Walter R. (Inventor)

2010-01-01

A span-loaded, highly flexible flying wing, having horizontal control surfaces mounted aft of the wing on extended beams to form local pitch-control devices. Each of five spanwise wing segments of the wing has one or more motors and photovoltaic arrays, and produces its own lift independent of the other wing segments, to minimize inter-segment loads. Wing dihedral is controlled by separately controlling the local pitch-control devices consisting of a control surface on a boom, such that inboard and outboard wing segment pitch changes relative to each other, and thus relative inboard and outboard lift is varied.

MODEL TESTS AND 3D ELASTIC FINITE ELEMENT ANALYSIS FOR STEEL PIPE PILES WITH WINGS IN STALLED IN SOIL CEMENT COLUMN

NASA Astrophysics Data System (ADS)

Tamai, Toshiyuki; Teramoto, Shuntarou; Kimura, Makoto

Steel pipe piles with wings installed in soil cement column is a composite foundation of pile consisting of soil improvement with cement and steel pipe with wings. This type of pile shows higher vertical bearing capacity when compared to steel pipe piles that are installed without soil cement. It is thought the wings contribute to higher bearing capacity of this type of piles. The wings are also thought to play the role of structural unification of pile foundations and load transfer. In this study, model test and 3D elastic finite element analysis was carried out in order to elucidate the effect of wings on the structural unification of pile foundation and the load transfer mechanism. Firstly, the model test was carried out in order to grasp the influence of pile with and without wings, the shape of wings of the pile and the unconfined compression strength of the soil cement on the structural unification of the pile foundation. The numerical analysis of the model test was then carried out on the intermediate part of the pile foundation with wings and mathematical model developed. Finally load tran sfer mechanism was checked for the entire length of the pile through this mathematical model and the load sharing ratio of the wings and stress distribution occurring in the soil cement clarified. In addition, the effect of the wing interval on the structural unification of the pile foundation and load transfer was also checked and clarified.

The calculation of downwash behind supersonic wings with an application to triangular plan forms

NASA Technical Reports Server (NTRS)

Lomax, Harvard; Sluder, Loma; Heaslet, Max A

1950-01-01

A method is developed consistent with the assumptions of small perturbation theory which provides a means of determining the downwash behind a wing in supersonic flow for a known load distribution. The analysis is based upon the use of supersonic doublets which are distributed over the plan form and wake of the wing in a manner determined from the wing loading. The equivalence in subsonic and supersonic flow of the downwash at infinity corresponding to a given load distribution is proved.

A study of altitude-constrained supersonic cruise transport concepts

NASA Technical Reports Server (NTRS)

Tice, David C.; Martin, Glenn L.

1992-01-01

The effect of restricting maximum cruise altitude on the mission performance of two supersonic transport concepts across a selection of cruise Mach numbers is studied. Results indicate that a trapezoidal wing concept can be competitive with an arrow wing depending on the altitude and Mach number constraints imposed. The higher wing loading of trapezoidal wing configurations gives them an appreciably lower average cruise altitude than the lower wing loading of the arrow wing configurations, and this advantage increases as the maximum allowable cruise altitude is reduced.

Evaluation of the Hinge Moment and Normal Force Aerodynamic Loads from a Seamless Adaptive Compliant Trailing Edge Flap in Flight

NASA Technical Reports Server (NTRS)

Miller, Eric J.; Cruz, Josue; Lung, Shun-Fat; Kota, Sridhar; Ervin, Gregory; Lu, Kerr-Jia; Flick, Pete

2016-01-01

A seamless adaptive compliant trailing edge (ACTE) flap was demonstrated in flight on a Gulfstream III aircraft at the NASA Armstrong Flight Research Center. The trailing edge flap was deflected between minus 2 deg up and plus 30 deg down in flight. The safety-of-flight parameters for the ACTE flap experiment require that flap-to-wing interface loads be sensed and monitored in real time to ensure that the structural load limits of the wing are not exceeded. The attachment fittings connecting the flap to the aircraft wing rear spar were instrumented with strain gages and calibrated using known loads for measuring hinge moment and normal force loads in flight. The safety-of-flight parameters for the ACTE flap experiment require that flap-to-wing interface loads be sensed and monitored in real time to ensure that the structural load limits of the wing are not exceeded. The attachment fittings connecting the flap to the aircraft wing rear spar were instrumented with strain gages and calibrated using known loads for measuring hinge moment and normal force loads in flight. The interface hardware instrumentation layout and load calibration are discussed. Twenty-one applied calibration test load cases were developed for each individual fitting. The 2-sigma residual errors for the hinge moment was calculated to be 2.4 percent, and for normal force was calculated to be 7.3 percent. The hinge moment and normal force generated by the ACTE flap with a hinge point located at 26-percent wing chord were measured during steady state and symmetric pitch maneuvers. The loads predicted from analysis were compared to the loads observed in flight. The hinge moment loads showed good agreement with the flight loads while the normal force loads calculated from analysis were over-predicted by approximately 20 percent. Normal force and hinge moment loads calculated from the pressure sensors located on the ACTE showed good agreement with the loads calculated from the installed strain gages.

Modeling Aircraft Wing Loads from Flight Data Using Neural Networks

NASA Technical Reports Server (NTRS)

Allen, Michael J.; Dibley, Ryan P.

2003-01-01

Neural networks were used to model wing bending-moment loads, torsion loads, and control surface hinge-moments of the Active Aeroelastic Wing (AAW) aircraft. Accurate loads models are required for the development of control laws designed to increase roll performance through wing twist while not exceeding load limits. Inputs to the model include aircraft rates, accelerations, and control surface positions. Neural networks were chosen to model aircraft loads because they can account for uncharacterized nonlinear effects while retaining the capability to generalize. The accuracy of the neural network models was improved by first developing linear loads models to use as starting points for network training. Neural networks were then trained with flight data for rolls, loaded reversals, wind-up-turns, and individual control surface doublets for load excitation. Generalization was improved by using gain weighting and early stopping. Results are presented for neural network loads models of four wing loads and four control surface hinge moments at Mach 0.90 and an altitude of 15,000 ft. An average model prediction error reduction of 18.6 percent was calculated for the neural network models when compared to the linear models. This paper documents the input data conditioning, input parameter selection, structure, training, and validation of the neural network models.

Conceptual design study of 1985 commercial tilt rotor transports. Volume 3: STOL design summary

NASA Technical Reports Server (NTRS)

Sambell, K. W.

1976-01-01

A conceptual design study is presented of 1,985 commercial tilt rotor STOL transports for a NASA 200 n. mi. (370 km) STOL Mission. A 100-passenger STOL Variant (Bell D313) of the Phase I VTOL Tilt Rotor Aircraft is defined. Aircraft characteristics are given; with the aircraft redesigned to meet 2,000-foot (610 m) field criteria, with emphasis on low fuel consumption and low direct operating cost. The 100-passenger STOL Tilt Rotor Aircraft was analyzed for performance, weights, economics, handling qualities, noise footprint and aeroelastic stability.

14 CFR 23.343 - Design fuel loads.

Code of Federal Regulations, 2012 CFR

2012-01-01

... zero fuel to the selected maximum fuel load. (b) If fuel is carried in the wings, the maximum allowable weight of the airplane without any fuel in the wing tank(s) must be established as “maximum zero wing... part and— (1) The structure must be designed to withstand a condition of zero fuel in the wing at limit...

14 CFR 23.343 - Design fuel loads.

Code of Federal Regulations, 2014 CFR

2014-01-01

... zero fuel to the selected maximum fuel load. (b) If fuel is carried in the wings, the maximum allowable weight of the airplane without any fuel in the wing tank(s) must be established as “maximum zero wing... part and— (1) The structure must be designed to withstand a condition of zero fuel in the wing at limit...

14 CFR 23.343 - Design fuel loads.

Code of Federal Regulations, 2011 CFR

2011-01-01

... zero fuel to the selected maximum fuel load. (b) If fuel is carried in the wings, the maximum allowable weight of the airplane without any fuel in the wing tank(s) must be established as “maximum zero wing... part and— (1) The structure must be designed to withstand a condition of zero fuel in the wing at limit...

14 CFR 23.343 - Design fuel loads.

Code of Federal Regulations, 2013 CFR

2013-01-01

... zero fuel to the selected maximum fuel load. (b) If fuel is carried in the wings, the maximum allowable weight of the airplane without any fuel in the wing tank(s) must be established as “maximum zero wing... part and— (1) The structure must be designed to withstand a condition of zero fuel in the wing at limit...

Aircraft control system

NASA Technical Reports Server (NTRS)

Kendall, Greg T. (Inventor); Lisoski, Derek L. (Inventor)

2007-01-01

A solar rechargeable, long-duration, span-loaded flying wing, having no fuselage or rudder. Having a two-hundred foot wingspan that mounts photovoltaic cells on most all of the wing's top surface, the aircraft uses only differential thrust of its eight propellers to turn, pitch and yaw. The wing is configured to deform under flight loads to position the propellers such that the control can be achieved. Each of five segments of the wing has one or more motors and photovoltaic arrays, and produces its own lift independent of the other segments, to avoid loading them. Five two-sided photovoltaic arrays, in all, are mounted on the wing, and receive photovoltaic energy both incident on top of the wing, and which is incident also from below, through a bottom, transparent surface.

Effect of leading-edge load constraints on the design and performance of supersonic wings

NASA Technical Reports Server (NTRS)

Darden, C. M.

1985-01-01

A theoretical and experimental investigation was conducted to assess the effect of leading-edge load constraints on supersonic wing design and performance. In the effort to delay flow separation and the formation of leading-edge vortices, two constrained, linear-theory optimization approaches were used to limit the loadings on the leading edge of a variable-sweep planform design. Experimental force and moment tests were made on two constrained camber wings, a flat uncambered wing, and an optimum design with no constraints. Results indicate that vortex strength and separation regions were mildest on the severely and moderately constrained wings.

Effect of an alternate winglet on the pressure and spanwise load distributions of a first generation jet transport wing

NASA Technical Reports Server (NTRS)

Montoya, L. C.; Flechner, S. G.; Jacobs, P. F.

1978-01-01

Pressure and spanwise load distributions on a first-generation jet transport semispan model at subsonic speeds are presented. The wind tunnel data were measured for the wing with and without an alternate winglet. The results show that the winglet affected outboard wing pressure distributions and increased the spanwise loads near the tip.

Deformed Shape Calculation of a Full-Scale Wing Using Fiber Optic Strain Data from a Ground Loads Test

NASA Technical Reports Server (NTRS)

Jutte, Christine V.; Ko, William L.; Stephens, Craig A.; Bakalyar, John A.; Richards, W. Lance

2011-01-01

A ground loads test of a full-scale wing (175-ft span) was conducted using a fiber optic strain-sensing system to obtain distributed surface strain data. These data were input into previously developed deformed shape equations to calculate the wing s bending and twist deformation. A photogrammetry system measured actual shape deformation. The wing deflections reached 100 percent of the positive design limit load (equivalent to 3 g) and 97 percent of the negative design limit load (equivalent to -1 g). The calculated wing bending results were in excellent agreement with the actual bending; tip deflections were within +/- 2.7 in. (out of 155-in. max deflection) for 91 percent of the load steps. Experimental testing revealed valuable opportunities for improving the deformed shape equations robustness to real world (not perfect) strain data, which previous analytical testing did not detect. These improvements, which include filtering methods developed in this work, minimize errors due to numerical anomalies discovered in the remaining 9 percent of the load steps. As a result, all load steps attained +/- 2.7 in. accuracy. Wing twist results were very sensitive to errors in bending and require further development. A sensitivity analysis and recommendations for fiber implementation practices, along with, effective filtering methods are included

Static investigation of the circulation control wing/upper surface blowing concept applied to the quiet short haul research aircraft

NASA Technical Reports Server (NTRS)

Eppel, J. C.; Shovlin, M. D.; Jaynes, D. N.; Englar, R. J.; Nichols, J. H., Jr.

1982-01-01

Full scale static investigations were conducted on the Quiet Short Haul Research Aircraft (QSRA) to determine the thrust deflecting capabilities of the circulation control wing/upper surface blowing (CCW/USB) concept. This scheme, which combines favorable characteristics of both the A-6/CCW and QSRA, employs the flow entrainment properties of CCW to pneumatically deflect engine thrust in lieu of the mechanical USB flap system. Results show that the no moving parts blown system produced static thrust deflections in the range of 40 deg to 97 deg (depending on thrust level) with a CCW pressure of 208,900 Pa (30.3 psig). In addition, the ability to vary horizontal forces from thrust to drag while maintaining a constant vertical (or lift) value was demonstrated by varying the blowing pressure. The versatility of the CCW/USB system, if applied to a STOL aircraft, was confirmed, where rapid conversion from a high drag approach mode to a thrust recovering waveoff or takeoff configuration could be achieved by nearly instantaneous blowing pressure variation.

Study of quiet turbofan STOL aircraft for short-haul transportation. Volume 2: Aircraft

NASA Technical Reports Server (NTRS)

1973-01-01

A study of the quiet turbofan STOL aircraft for short haul transportation was conducted. The objectives of the study were as follows: (1) to determine the relationships between STOL characteristics and economic and social viability of short haul air transportation, (2) to identify critical technology problems involving introduction of STOL short haul systems, (3) to define representative aircraft configurations, characteristics, and costs, and (4) to identify high payoff technology areas to improve STOL systems. The analyses of the aircraft designs which were generated to fulfill the objectives are summarized. The baseline aircraft characteristics are documented and significant trade studies are presented.

Viper cabin-fuselage structural design concept with engine installation and wing structural design

NASA Technical Reports Server (NTRS)

Marchesseault, B.; Carr, D.; Mccorkle, T.; Stevens, C.; Turner, D.

1993-01-01

This report describes the process and considerations in designing the cabin, nose, drive shaft, and wing assemblies for the 'Viper' concept aircraft. Interfaces of these assemblies, as well as interfaces with the sections of the aircraft aft of the cabin, are also discussed. The results of the design process are included. The goal of this project is to provide a structural design which complies with FAR 23 requirements regarding occupant safety, emergency landing loads, and maneuvering loads. The design must also address the interfaces of the various systems in the cabin, nose, and wing, including the drive shaft, venting, vacuum, electrical, fuel, and control systems. Interfaces between the cabin assembly and the wing carrythrough and empennage assemblies were required, as well. In the design of the wing assemblies, consistency with the existing cabin design was required. The major areas considered in this report are materials and construction, loading, maintenance, environmental considerations, wing assembly fatigue, and weight. The first three areas are developed separately for the nose, cabin, drive shaft, and wing assemblies, while the last three are discussed for the entire design. For each assembly, loading calculations were performed to determine the proper sizing of major load carrying components. Table 1.0 lists the resulting margins of safety for these key components, along with the types of the loads involved, and the page number upon which they are discussed.

Installation effects on performance of multiple model V/STOL lift fans

NASA Technical Reports Server (NTRS)

Diedrich, J. H.; Clough, N.; Lieblein, S.

1972-01-01

An experimental program was performed in which the individual performance of multiple VTOL model lift fans was measured. The model tested consisted of three 5.5 in. diameter tip-turbine driven model VTOL lift fans mounted chordwise in a two-dimensional wing to simulate a pod-type array. The performance data provided significant insight into possible thrust variations and losses caused by the presence of cover doors, adjacent fuselage panels, and adjacent fans. The effect of a partial loss of drive air supply (simulated gas generator failure) on fan performance was also investigated. The results of the tests demonstrated that lift fan installation variables and hardware can have a significant effect on the thrust of the individual fans.

An evaluation of proposed acoustic treatments for the NASA LaRC 4 x 7 meter wind tunnel

NASA Technical Reports Server (NTRS)

Abrahamson, A. L.

1985-01-01

The NASA LaRC 4 x 7 Meter Wind Tunnel is an existing facility specially designed for powered low speed (V/STOL) testing of large scale fixed wing and rotorcraft models. The enhancement of the facility for scale model acoustic testing is examined. The results are critically reviewed and comparisons are drawn with a similar wind tunnel (the DNW Facility in the Netherlands). Discrepancies observed in the comparison stimulated a theoretical investigation using the acoustic finite element ADAM System, of the ways in which noise propagating around the tunnel circuit radiates into the open test section. The reasons for the discrepancies noted above are clarified and assists in the selection of acoustic treatment options for the facility.

Shock Location Dominated Transonic Flight Loads on the Active Aeroelastic Wing

NASA Technical Reports Server (NTRS)

Lokos, William A.; Lizotte, Andrew; Lindsley, Ned J.; Stauf, Rick

2005-01-01

During several Active Aeroelastic Wing research flights, the shadow of the over-wing shock could be observed because of natural lighting conditions. As the plane accelerated, the shock location moved aft, and as the shadow passed the aileron and trailing-edge flap hinge lines, their associated hinge moments were substantially affected. The observation of the dominant effect of shock location on aft control surface hinge moments led to this investigation. This report investigates the effect of over-wing shock location on wing loads through flight-measured data and analytical predictions. Wing-root and wing-fold bending moment and torque and leading- and trailing-edge hinge moments have been measured in flight using calibrated strain gages. These same loads have been predicted using a computational fluid dynamics code called the Euler Navier-Stokes Three Dimensional Aeroelastic Code. The computational fluid dynamics study was based on the elastically deformed shape estimated by a twist model, which in turn was derived from in-flight-measured wing deflections provided by a flight deflection measurement system. During level transonic flight, the shock location dominated the wing trailing-edge control surface hinge moments. The computational fluid dynamics analysis based on the shape provided by the flight deflection measurement system produced very similar results and substantially correlated with the measured loads data.

Load distribution on a closed-coupled wing canard at transonic speeds

NASA Technical Reports Server (NTRS)

Gloss, B. B.; Washburn, K. E.

1977-01-01

A wind tunnel test where load distributions were obtained at transonic speeds on both the canard and wing surfaces of a closely coupled wing canard configuration is reported. Detailed component and configuration arrangement studies to provide insight into the various aerodynamic interference effects for the leading edge vortex flow conditions encountered are included. Data indicate that increasing the Mach number from 0.70 to 0.95 caused the wing leading edge vortex to burst over the wing when the wing was in the presence of the high canard.

Wing Weight Optimization Under Aeroelastic Loads Subject to Stress Constraints

NASA Technical Reports Server (NTRS)

Kapania, Rakesh K.; Issac, J.; Macmurdy, D.; Guruswamy, Guru P.

1997-01-01

A minimum weight optimization of the wing under aeroelastic loads subject to stress constraints is carried out. The loads for the optimization are based on aeroelastic trim. The design variables are the thickness of the wing skins and planform variables. The composite plate structural model incorporates first-order shear deformation theory, the wing deflections are expressed using Chebyshev polynomials and a Rayleigh-Ritz procedure is adopted for the structural formulation. The aerodynamic pressures provided by the aerodynamic code at a discrete number of grid points is represented as a bilinear distribution on the composite plate code to solve for the deflections and stresses in the wing. The lifting-surface aerodynamic code FAST is presently being used to generate the pressure distribution over the wing. The envisioned ENSAERO/Plate is an aeroelastic analysis code which combines ENSAERO version 3.0 (for analysis of wing-body configurations) with the composite plate code.

A Structural Weight Estimation Program (SWEEP) for Aircraft. Volume 11 - Flexible Airloads Stand-Alone Program

DTIC Science & Technology

1974-06-01

stiffness, lb-in. I Integer used to designate wing strip number 2 I Airplanw pitching moment of inertia, slug ft 2 I Airplane yawing moment of inertia...slug ft J Integer used to designated wing-loading distribution, i.e., J-l, loading due to angle of attack J=2> loading due to flap deflection J-3...moment at intersection of load reference line and body interface station (for vertical tail), in.-lb Integer used to designate type of wing airload

Lifting-surface theory for calculating the loading induced on a wing by a flap

NASA Technical Reports Server (NTRS)

Johnson, W. A.

1972-01-01

A method is described for using lifting-surface theory to obtain the pressure distribution on a wing with a trailing-edge flap or control surface. The loading has a logarithmic singularity at the flap edges, which may be determined directly by the method of matched asymptotic expansions. Expressions are given for the singular flap loading for various flap hinge line and side edge geometries, both for steady and unsteady flap deflection. The regular part of the flap loading must be obtained by inverting the lifting-surface-theory integral equation relating the pressure and the downwash on the wing: procedures are described to accomplish this for a general wing and flap geometry. The method is applied to several example wings, and the results are compared with experimental data. Theory and test correlate well.

Loads calibrations of strain gage bridges on the DAST project Aeroelastic Research Wing (ARW-2)

NASA Technical Reports Server (NTRS)

Eckstrom, C. V.

1986-01-01

Results from and details of the procedure used to calibrate strain gage bridges for measurements of wing structural loads, shear (V), bending moment (M), and torque (T), at three semispan stations on both the left and right semispans of the ARW-2 wing are presented. The ARW-2 wing has a reference area of 35 square feet, a span of 19 feet, an aspect ratio of 10.3, a midchord line sweepback angle of 25 degrees, and a taper ratio of 0.4. The ARW-2 wing was fabricated using aluminum spars and ribs covered with a fiberglass/honeycomb sandwich skin material. All strain gage bridges are mounted along with an estimate of their accuracy by means of a comparison of computed loads versus actual loads for three simulated flight conditions.

Deflection-Based Structural Loads Estimation From the Active Aeroelastic Wing F/A-18 Aircraft

NASA Technical Reports Server (NTRS)

Lizotte, Andrew M.; Lokos, William A.

2005-01-01

Traditional techniques in structural load measurement entail the correlation of a known load with strain-gage output from the individual components of a structure or machine. The use of strain gages has proved successful and is considered the standard approach for load measurement. However, remotely measuring aerodynamic loads using deflection measurement systems to determine aeroelastic deformation as a substitute to strain gages may yield lower testing costs while improving aircraft performance through reduced instrumentation weight. This technique was examined using a reliable strain and structural deformation measurement system. The objective of this study was to explore the utility of a deflection-based load estimation, using the active aeroelastic wing F/A-18 aircraft. Calibration data from ground tests performed on the aircraft were used to derive left wing-root and wing-fold bending-moment and torque load equations based on strain gages, however, for this study, point deflections were used to derive deflection-based load equations. Comparisons between the strain-gage and deflection-based methods are presented. Flight data from the phase-1 active aeroelastic wing flight program were used to validate the deflection-based load estimation method. Flight validation revealed a strong bending-moment correlation and slightly weaker torque correlation. Development of current techniques, and future studies are discussed.

Recommendations for ground effects research for V/STOL and STOL aircraft and associated equipment for large scale testing

NASA Technical Reports Server (NTRS)

Kuhn, Richard E.

1986-01-01

The current understanding of the effects of ground proximity on V/STOL and STOL aircraft is reviewd. Areas covered include (1) single jet suckdown in hover, (2) fountain effects on multijet configurations, (3) STOL ground effects including the effect of the ground vortex flow field, (4) downwash at the tail, and (5) hot gas ingestion in both hover and STOL operation. The equipment needed for large scale testing to extend the state of the art is reviewed and developments in three areas are recommended as follows: (1) improve methods for simulating the engine exhaust and inlet flows; (2) develop a model support system that can simulate realistic rates of climb and descent as well as steady height operation; and (3) develop a blowing BLC ground board as an alternative to a moving belt ground board to properly simulate the flow on the ground.

Rolling moments in a trailing vortex flow field

NASA Technical Reports Server (NTRS)

Mcmillan, O. J.; Schwind, R. G.; Nielsen, J. N.; Dillenius, M. F. E.

1977-01-01

Pressure distributions are presented which were measured on a wing in close proximity to a tip vortex of known structure generated by a larger, upstream semispan wing. Overall loads calculated by integration of these pressures are checked by independent measurements made with an identical model mounted on a force balance. Several conventional methods of wing analysis are used to predict the loads on the following wing. Strip theory is shown to give uniformly poor results for loading distribution, although predictions of overall lift and rolling moment are sometimes acceptable. Good results are obtained for overall coefficients and loading distribution by using linearized pressures in vortex-lattice theory in conjunction with a rectilinear vortex. The equivalent relation from reverse-flow theory that can be used to give economic predictions for overall loads is presented.

Application of winglets and/or wing tip extensions with active load control on the Boeing 747

NASA Technical Reports Server (NTRS)

Allison, R. L.; Perkin, B. R.; Schoenman, R. L.

1978-01-01

The application of wing tip modifications and active control technology to the Boeing 747 airplane for the purpose of improving fuel efficiency is considered. Wing tip extensions, wing tip winglets, and the use of the outboard ailerons for active wing load alleviation are described. Modest performance improvements are indicated. A costs versus benefits approach is taken to decide which, if any, of the concepts warrant further development and flight test leading to possible incorporation into production airplanes.

Effects of wing flexibility and variable air lift upon wing bending moment during landing impacts of a small seaplane

NASA Technical Reports Server (NTRS)

Merten, Kenneth F; Beck, Edgar B

1951-01-01

A smooth-water-landing investigation was conducted with a small seaplane to obtain experimental wing-bending-moment time histories together with time histories of the various parameters necessary for the prediction of wing bending moments during hydrodynamic forcing functions. The experimental results were compared with calculated results which include inertia-load effects and the effects of air-load variation during impact. The responses of the fundamental mode were calculated with the use of the measured hydrodynamic forcing functions. From these responses, the wing bending moments due to the hydrodynamic load were calculated according to the procedure given in R.M. No. 2221. The comparison of the time histories of the experimental and calculated wing bending moments showed good agreement both in phase relationship of the oscillations and in numerical values.

Vortex wake alleviation studies with a variable twist wing

NASA Technical Reports Server (NTRS)

Holbrook, G. T.; Dunham, D. M.; Greene, G. C.

1985-01-01

Vortex wake alleviation studies were conducted in a wind tunnel and a water towing tank using a multisegmented wing model which provided controlled and measured variations in span load. Fourteen model configurations are tested at a Reynolds number of one million and a lift coefficient of 0.6 in the Langley 4- by 7-Meter Tunnel and the Hydronautics Ship Model Basin water tank at Hydronautics, Inc., Laurel, Md. Detailed measurements of span load and wake velocities at one semispan downstream correlate well with each other, with inviscid predictions of span load and wake roll up, and with peak trailing-wing rolling moments measured in the far wake. Average trailing-wing rolling moments are found to be an unreliable indicator of vortex wake intensity because vortex meander does not scale between test facilities and free-air conditions. A tapered-span-load configuration, which exhibits little or no drag penalty, is shown to offer significant downstream wake alleviation to a small trailing wing. The greater downstream wake alleviation achieved with the addition of spoilers to a flapped-wing configuration is shown to result directly from the high incremental drag and turbulence associated with the spoilers and not from the span load alteration they cause.

Correlation of Structural Analysis and Test Results for the McDonnell Douglas Stitched/RFI All-Composite Wing Stub Box

NASA Technical Reports Server (NTRS)

Wang, John T.; Jegley, Dawn C.; Bush, Harold G.; Hinrichs, Stephen C.

1996-01-01

The analytical and experimental results of an all-composite wing stub box are presented in this report. The wing stub box, which is representative of an inboard portion of a commercial transport high-aspect-ratio wing, was fabricated from stitched graphite-epoxy material with a Resin Film Infusion manufacturing process. The wing stub box was designed and constructed by the McDonnell Douglas Aerospace Company as part of the NASA Advanced Composites Technology program. The test article contained metallic load-introduction structures on the inboard and outboard ends of the graphite-epoxy wing stub box. The root end of the inboard load introduction structure was attached to a vertical reaction structure, and an upward load was applied to the outermost tip of the outboard load introduction structure to induce bending of the wing stub box. A finite element model was created in which the center portion of the wing-stub-box upper cover panel was modeled with a refined mesh. The refined mesh was required to represent properly the geometrically nonlinear structural behavior of the upper cover panel and to predict accurately the strains in the stringer webs of the stiffened upper cover panel. The analytical and experimental results for deflections and strains are in good agreement.

A candidate V/STOL research aircraft design concept using an S-3A aircraft and 2 Pegasus 11 engines

NASA Technical Reports Server (NTRS)

Lampkin, B. A.

1980-01-01

A candidate V/STOL research aircraft concept which uses an S-3A airframe and two Pegasus 11 engines was studied to identify a feasible V/STOL national flight facility that could be obtained at the lowest possible cost for the demonstration of V/STOL technology, inflight simulation, and flight research. The rationale for choosing the configuration, a description of the configuration, and the capability of a fully developed aircraft are discussed.

Design and fabrication of graphite-epoxy bolted wing skin splice specimens

NASA Technical Reports Server (NTRS)

Johnson, R. W.; Mccarty, J. E.

1977-01-01

Graphite-epoxy bolted joint specimens were designed and fabricated. These specimens were to be representative of a side-of-body wing skin splice with a 20-year life expectancy in a commercial transport environment. Preliminary tests were performed to determine design values of bearing and net tension stresses. Based upon the information developed, a three-fastener-wide representative wing skin splice was designed for a load of 2627 KN/m (15,000 lbf/in.). One joint specimen was fabricated and tested at NASA. The wing skin splice failed at 106 percent of design ultimate load. This joint design achieved all static load objectives. Fabrication of six specimens, together with their loading fixtures, was completed, and the specimens were delivered to NASA-LRC.

A study of canard-wing interference using experimental pressure data at transonic speeds

NASA Technical Reports Server (NTRS)

Gloss, B. B.; Washburn, K. E.

1979-01-01

The canard had an exposed area of 28.0 percent of the wing reference area and was located in the chord plane of the wing or in a position 18.5 percent of the wing mean geometric chord above or below the wing chord plane. The canard leading edge sweep was 51.7 deg and the wing leading-edge sweep was 60 deg. The results indicated that the direct canard downwash effects on the wing loading are limited to the forward half of the wing directly behind the canard. The wing leading-edge vortex is located farther forward for the wing in the presence of the canard than for the wing-alone configuration. The wake, from the canard located below the wing chord plane, physically interacts with the wing inboard surface and produces a substantial loss of wing lift. For the Mach number 0.70 case, the presence of the wing increased the loading on the canard for the higher angles of attack. However, at Mach numbers of 0.95 and 1.20, the presence of the wing had the unexpected result of unloading the canard.

L∞-Optimal feedforward gust load alleviation design for a large blended wing body airliner

NASA Astrophysics Data System (ADS)

Wildschek, A.; Haniš, T.; Stroscher, F.

2013-12-01

The potential advantages of Blended Wing Body (BWB) aircraft in terms of fuel efficiency are opposed by technical challenges such as the alleviation of gust loads. Due to the low wing, loading gusts, generally, have a more severe impact on BWB aircraft than on conventional aircraft. This paper presents the design and optimization of a Gust Load Alleviation System (GLAS) for a large BWB airliner. Numerical simulations are performed with an aeroelastic model of the aircraft including GLAS in order to compute time series of modal displacements for deriving equivalent static load cases which are used for the resizing of the aircraft structure.

Flight Test of the F/A-18 Active Aeroelastic Wing Airplane

NASA Technical Reports Server (NTRS)

Clarke, Robert; Allen, Michael J.; Dibley, Ryan P.; Gera, Joseph; Hodgkinson, John

2005-01-01

Successful flight-testing of the Active Aeroelastic Wing airplane was completed in March 2005. This program, which started in 1996, was a joint activity sponsored by NASA, Air Force Research Laboratory, and industry contractors. The test program contained two flight test phases conducted in early 2003 and early 2005. During the first phase of flight test, aerodynamic models and load models of the wing control surfaces and wing structure were developed. Design teams built new research control laws for the Active Aeroelastic Wing airplane using these flight-validated models; and throughout the final phase of flight test, these new control laws were demonstrated. The control laws were designed to optimize strategies for moving the wing control surfaces to maximize roll rates in the transonic and supersonic flight regimes. Control surface hinge moments and wing loads were constrained to remain within hydraulic and load limits. This paper describes briefly the flight control system architecture as well as the design approach used by Active Aeroelastic Wing project engineers to develop flight control system gains. Additionally, this paper presents flight test techniques and comparison between flight test results and predictions.

Measured and predicted structural behavior of the HiMAT tailored composite wing

NASA Technical Reports Server (NTRS)

Nelson, Lawrence H.

1987-01-01

A series of load tests was conducted on the HiMAT tailored composite wing. Coupon tests were also run on a series of unbalanced laminates, including the ply configuration of the wing, the purpose of which was to compare the measured and predicted behavior of unbalanced laminates, including - in the case of the wing - a comparison between the behavior of the full scale structure and coupon tests. Both linear and nonlinear finite element (NASTRAN) analyses were carried out on the wing. Both linear and nonlinear point-stress analyses were performed on the coupons. All test articles were instrumented with strain gages, and wing deflections measured. The leading and trailing edges were found to have no effect on the response of the wing to applied loads. A decrease in the stiffness of the wing box was evident over the 27-test program. The measured load-strain behavior of the wing was found to be linear, in contrast to coupon tests of the same laminate, which were nonlinear. A linear NASTRAN analysis of the wing generally correlated more favorably with measurements than did a nonlinear analysis. An examination of the predicted deflections in the wing root region revealed an anomalous behavior of the structural model that cannot be explained. Both hysteresis and creep appear to be less significant in the wing tests than in the corresponding laminate coupon tests.

Aeroelasticity of Axially Loaded Aerodynamic Structures for Truss-Braced Wing Aircraft

NASA Technical Reports Server (NTRS)

Nguyen, Nhan; Ting, Eric; Lebofsky, Sonia

2015-01-01

This paper presents an aeroelastic finite-element formulation for axially loaded aerodynamic structures. The presence of axial loading causes the bending and torsional sitffnesses to change. For aircraft with axially loaded structures such as the truss-braced wing aircraft, the aeroelastic behaviors of such structures are nonlinear and depend on the aerodynamic loading exerted on these structures. Under axial strain, a tensile force is created which can influence the stiffness of the overall aircraft structure. This tension stiffening is a geometric nonlinear effect that needs to be captured in aeroelastic analyses to better understand the behaviors of these types of aircraft structures. A frequency analysis of a rotating blade structure is performed to demonstrate the analytical method. A flutter analysis of a truss-braced wing aircraft is performed to analyze the effect of geometric nonlinear effect of tension stiffening on the flutter speed. The results show that the geometric nonlinear tension stiffening effect can have a significant impact on the flutter speed prediction. In general, increased wing loading results in an increase in the flutter speed. The study illustrates the importance of accounting for the geometric nonlinear tension stiffening effect in analyzing the truss-braced wing aircraft.

Low speed wind tunnel test of ground proximity and deck edge effects on a lift cruise fan V/STOL configuration, volume 1

NASA Technical Reports Server (NTRS)

Stewart, V. R.

1979-01-01

The characteristics were determined of a lift cruise fan V/STOL multi-mission configuration in the near proximity to the edge of a small flat surface representation of a ship deck. Tests were conducted at both static and forward speed test conditions. The model (0.12 scale) tested was a four fan configuration with modifications to represent a three fan configuration. Analysis of data showed that the deck edge effects were in general less critical in terms of differences from free air than a full deck (in ground effect) configuration. The one exception to this was when the aft edge of the deck was located under the center of gravity. This condition, representative of an approach from the rear, showed a significant lift loss. Induced moments were generally small compared to the single axis control power requirements, but will likely add to the pilot work load.

Future V/STOL Aircraft For The Pacific Basin

NASA Technical Reports Server (NTRS)

Albers, James A.; Zuk, John

1992-01-01

Report describes geography and transportation needs of Asian Pacific region, and describes aircraft configurations suitable for region and compares performances. Examines applications of high-speed rotorcraft, vertical/short-takeoff-and-landing (V/STOL) aircraft, and short-takeoff-and landing (STOL) aircraft. Configurations benefit commerce, tourism, and development of resources.

Effects on sleep of noise from two proposed STOL aircraft

NASA Technical Reports Server (NTRS)

Lukas, J. S.; Peeler, D. J.; Davis, J. E.

1975-01-01

Responses, both overt behavior and those measured by electroencephalograph, to noise by eight male subjects were studied for sixteen consecutive nights. Test stimuli were: (1) The simulated sideline noise of a short takeoff and landing aircraft with blown flaps; (2) the simulated sideline noise of a STOL aircraft of turbofan design; (3) the simulated takeoff noise of the blown flap STOL aircraft; and (4) a four second burst of simulated pink noise. Responses to each noise were tested at three noise intensities selected to represent levels expected indoors from operational aircraft. The results indicate that the blown flap STOL aircraft noise resulted in 8 to 10 percent fewer sleep disturbance responses than did the turbofan STOL aircraft when noises of comparable intensities from similar maneuvers were used.

Flight through thunderstorm outflows. [aircraft landing

NASA Technical Reports Server (NTRS)

Frost, W.; Crosby, B.; Camp, D. W.

1978-01-01

Computer simulation of aircraft landing through thunderstorm gust fronts is carried out. The two-dimensional, nonlinear equations or aircraft motion containing all wind shear terms are solved numerically. The gust front spatial wind field inputs are provided in the form of tabulated experimental data which are coupled with a computer table lookup routine to provide the required wind components and shear at any given position within an approximate 500 m by 1 km vertical plane. The aircraft is considered to enter the wind field at a specified position under trimmed conditions. Both fixed control and automatic control landings are simulated. Flight paths, as well as control inputs necessary to maintain specified trajectories, are presented and discussed for aircraft having characteristics of a DC-8, B-747, augmentor-wing STOL, and a DHC-6.

Robustness in linear quadratic feedback design with application to an aircraft control problem

NASA Technical Reports Server (NTRS)

Patel, R. V.; Sridhar, B.; Toda, M.

1977-01-01

Some new results concerning robustness and asymptotic properties of error bounds of a linear quadratic feedback design are applied to an aircraft control problem. An autopilot for the flare control of the Augmentor Wing Jet STOL Research Aircraft (AWJSRA) is designed based on Linear Quadratic (LQ) theory and the results developed in this paper. The variation of the error bounds to changes in the weighting matrices in the LQ design is studied by computer simulations, and appropriate weighting matrices are chosen to obtain a reasonable error bound for variations in the system matrix and at the same time meet the practical constraints for the flare maneuver of the AWJSRA. Results from the computer simulation of a satisfactory autopilot design for the flare control of the AWJSRA are presented.

Wing Torsional Stiffness Tests of the Active Aeroelastic Wing F/A-18 Airplane

NASA Technical Reports Server (NTRS)

Lokos, William A.; Olney, Candida D.; Crawford, Natalie D.; Stauf, Rick; Reichenbach, Eric Y.

2002-01-01

The left wing of the Active Aeroelastic Wing (AAW) F/A-18 airplane has been ground-load-tested to quantify its torsional stiffness. The test has been performed at the NASA Dryden Flight Research Center in November 1996, and again in April 2001 after a wing skin modification was performed. The primary objectives of these tests were to characterize the wing behavior before the first flight, and provide a before-and-after measurement of the torsional stiffness. Two streamwise load couples have been applied. The wing skin modification is shown to have more torsional flexibility than the original configuration has. Additionally, structural hysteresis is shown to be reduced by the skin modification. Data comparisons show good repeatability between the tests.

Effects of flaps on buffet characteristics and wind-rock onset of an F-8C airplane at subsonic and transonic speeds

NASA Technical Reports Server (NTRS)

Monaghan, R. C.; Friend, E. L.

1973-01-01

Wind-up-turn maneuvers were performed to establish the values of airplane normal force coefficient for buffet onset, wing-rock onset, and buffet loads with various combinations of leading- and trailing-edge flap deflections. Data were gathered at both subsonic and transonic speeds covering a range from Mach 0.64 to Mach 0.92. Buffet onset and buffet loads were obtained from wingtip acceleration and wing-root bending-moment data, and wing-rock onset was obtained from airplane roll rate data. Buffet onset, wing-rock onset, and buffet loads were similarly affected by the various combinations of leading- and training-edge flaps. Subsonically, the 12 deg leading-edge-flap and trailing-edge-flap combination was most effective in delaying buffet onset, wing-rock onset, and equivalent values of buffet loads to a higher value of airplane normal force coefficient. This was the maximum flap deflection investigated. Transonically, however, the optimum leading-edge flap position was generally less than 12 deg.

Utilization of Optimization for Design of Morphing Wing Structures for Enhanced Flight

NASA Astrophysics Data System (ADS)

Detrick, Matthew Scott

Conventional aircraft control surfaces constrain maneuverability. This work is a comprehensive study that looks at both smart material and conventional actuation methods to achieve wing twist to potentially improve flight capability using minimal actuation energy while allowing minimal wing deformation under aerodynamic loading. A continuous wing is used in order to reduce drag while allowing the aircraft to more closely approximate the wing deformation used by birds while loitering. The morphing wing for this work consists of a skin supported by an underlying truss structure whose goal is to achieve a given roll moment using less actuation energy than conventional control surfaces. A structural optimization code has been written in order to achieve minimal wing deformation under aerodynamic loading while allowing wing twist under actuation. The multi-objective cost function for the optimization consists of terms that ensure small deformation under aerodynamic loading, small change in airfoil shape during wing twist, a linear variation of wing twist along the length of the wing, small deviation from the desired wing twist, minimal number of truss members, minimal wing weight, and minimal actuation energy. Hydraulic cylinders and a two member linkage driven by a DC motor are tested separately to provide actuation. Since the goal of the current work is simply to provide a roll moment, only one actuator is implemented along the wing span. Optimization is also used to find the best location within the truss structure for the actuator. The active structure produced by optimization is then compared to simulated and experimental results from other researchers as well as characteristics of conventional aircraft.

Optimal Topology of Aircraft Rib and Spar Structures under Aeroelastic Loads

NASA Technical Reports Server (NTRS)

Stanford, Bret K.; Dunning, Peter D.

2014-01-01

Several topology optimization problems are conducted within the ribs and spars of a wing box. It is desired to locate the best position of lightening holes, truss/cross-bracing, etc. A variety of aeroelastic metrics are isolated for each of these problems: elastic wing compliance under trim loads and taxi loads, stress distribution, and crushing loads. Aileron effectiveness under a constant roll rate is considered, as are dynamic metrics: natural vibration frequency and flutter. This approach helps uncover the relationship between topology and aeroelasticity in subsonic transport wings, and can therefore aid in understanding the complex aircraft design process which must eventually consider all these metrics and load cases simultaneously.

Dynamic Structural Fault Detection and Identification

NASA Technical Reports Server (NTRS)

Smith, Timothy; Reichenbach, Eric; Urnes, James M.

2009-01-01

Aircraft structures are designed to guarantee safety of flight in some required operational envelope. When the aircraft becomes structurally impaired, safety of flight may not be guaranteed within that previously safe operational envelope. In this case the safe operational envelope must be redefined in-flight and a means to prevent excursion from this new envelope must be implemented. A specific structural failure mode that may result in a reduced safe operating envelope, the exceedance of which could lead to catastrophic structural failure of the aircraft, will be addressed. The goal of the DFEAP program is the detection of this failure mode coupled with flight controls adaptation to limit critical loads in the damaged aircraft structure. The DFEAP program is working with an F/A-18 aircraft model. The composite wing skins are bonded to metallic spars in the wing substructure. Over time, it is possible that this bonding can deteriorate due to fatigue. In this case, the ability of the wing spar to transfer loading between the wing skins is reduced. This failure mode can translate to a reduced allowable compressive strain on the wing skin and could lead to catastrophic wing buckling if load limiting of the wing structure is not applied. The DFEAP program will make use of a simplified wing strain model for the healthy aircraft. The outputs of this model will be compared in real-time to onboard strain measurements at several locations on the aircraft wing. A damage condition is declared at a given location when the strain measurements differ sufficiently from the strain model. Parameter identification of the damaged structure wing strain parameters will be employed to provide load limiting control adaptation for the aircraft. This paper will discuss the simplified strain models used in the implementation and their interaction with the strain sensor measurements. Also discussed will be the damage detection and identification schemes employed and the means by which the damaged aircraft parameters will be used to provide load limiting that keeps the aircraft within the safe operational envelope.

Analysis of high aspect ratio jet flap wings of arbitrary geometry.

NASA Technical Reports Server (NTRS)

Lissaman, P. B. S.

1973-01-01

Paper presents a design technique for rapidly computing lift, induced drag, and spanwise loading of unswept jet flap wings of arbitrary thickness, chord, twist, blowing, and jet angle, including discontinuities. Linear theory is used, extending Spence's method for elliptically loaded jet flap wings. Curves for uniformly blown rectangular wings are presented for direct performance estimation. Arbitrary planforms require a simple computer program. Method of reducing wing to equivalent stretched, twisted, unblown planform for hand calculation is also given. Results correlate with limited existing data, and show lifting line theory is reasonable down to aspect ratios of 5.

Application of a transonic potential flow code to the static aeroelastic analysis of three-dimensional wings

NASA Technical Reports Server (NTRS)

Whitlow, W., Jr.; Bennett, R. M.

1982-01-01

Since the aerodynamic theory is nonlinear, the method requires the coupling of two iterative processes - an aerodynamic analysis and a structural analysis. A full potential analysis code, FLO22, is combined with a linear structural analysis to yield aerodynamic load distributions on and deflections of elastic wings. This method was used to analyze an aeroelastically-scaled wind tunnel model of a proposed executive-jet transport wing and an aeroelastic research wing. The results are compared with the corresponding rigid-wing analyses, and some effects of elasticity on the aerodynamic loading are noted.

Theoretical antisymmetric span loading for wings of arbitrary plan form at subsonic speeds

NASA Technical Reports Server (NTRS)

Deyoung, John

1951-01-01

A simplified lifting-surface theory that includes effects of compressibility and spanwise variation of section lift-curve slope is used to provide charts with which antisymmetric loading due to arbitrary antisymmetric angle of attack can be found for wings having symmetric plan forms with a constant spanwise sweep angle of the quarter-chord line. Consideration is given to the flexible wing in roll. Aerodynamic characteristics due to rolling, deflected ailerons, and sideslip of wings with dihedral are considered. Solutions are presented for straight-tapered wings for a range of swept plan forms.

Conceptual design study of a 1985 commercial STOL tilt rotor transport

NASA Technical Reports Server (NTRS)

Widdison, C. A.; Magee, J. P.; Alexander, H. R.

1974-01-01

Results of conceptual engineering design studies of a STOL tilt rotor commercial aircraft for the 1985 time frame are presented. The details of aircraft size, performance, flying qualities, noise, and cost are included. The savings in terms of fuel economy resulting from STOL operations compared with VTOL vehicles are determined.

Flight testing a V/STOL aircraft to identify a full-envelope aerodynamic model

NASA Technical Reports Server (NTRS)

Mcnally, B. David; Bach, Ralph E., Jr.

1988-01-01

Flight-test techniques are being used to generate a data base for identification of a full-envelope aerodynamic model of a V/STOL fighter aircraft, the YAV-8B Harrier. The flight envelope to be modeled includes hover, transition to conventional flight and back to hover, STOL operation, and normal cruise. Standard V/STOL procedures such as vertical takeoff and landings, and short takeoff and landings are used to gather data in the powered-lift flight regime. Long (3 to 5 min) maneuvers which include a variety of input types are used to obtain large-amplitude control and response excitations. The aircraft is under continuous radar tracking; a laser tracker is used for V/STOL operations near the ground. Tracking data are used with state-estimation techniques to check data consistency and to derive unmeasured variables, for example, angular accelerations. A propulsion model of the YAV-8B's engine and reaction control system is used to isolate aerodynamic forces and moments for model identification. Representative V/STOL flight data are presented. The processing of a typical short takeoff and slow landing maneuver is illustrated.

Knowledge acquisition and representation for the Systems Test and Operations Language (STOL) Intelligent Tutoring System (ITS)

NASA Technical Reports Server (NTRS)

Seamster, Thomas L.; Eike, David R.; Ames, Troy J.

1990-01-01

This presentation concentrates on knowledge acquisition and its application to the development of an expert module and a user interface for an Intelligent Tutoring System (ITS). The Systems Test and Operations Language (STOL) ITS is being developed to assist NASA control center personnel in learning a command and control language as it is used in mission operations rooms. The objective of the tutor is to impart knowledge and skills that will permit the trainee to solve command and control problems in the same way that the STOL expert solves those problems. The STOL ITS will achieve this object by representing the solution space in such a way that the trainee can visualize the intermediate steps, and by having the expert module production rules parallel the STOL expert's knowledge structures.

Approximate calculation of multispar cantilever and semicantilever wings with parallel ribs under direct and indirect loading

NASA Technical Reports Server (NTRS)

Sanger, Eugen

1932-01-01

A method is presented for approximate static calculation, which is based on the customary assumption of rigid ribs, while taking into account the systematic errors in the calculation results due to this arbitrary assumption. The procedure is given in greater detail for semicantilever and cantilever wings with polygonal spar plan form and for wings under direct loading only. The last example illustrates the advantages of the use of influence lines for such wing structures and their practical interpretation.

77 FR 45518 - Airworthiness Directives; The Boeing Company Airplanes

Federal Register 2010, 2011, 2012, 2013, 2014

2012-08-01

... structure not supporting the limit load condition, which could lead to loss of structural integrity of the... wing structure not supporting the limit load condition, which could lead to loss of the structural... wing structure not supporting the limit load condition, which could lead to loss of structural...

Minimum trim drag design for interfering lifting surfaces using vortex-lattice methodology

NASA Technical Reports Server (NTRS)

Lamar, J. E.

1976-01-01

A new method has been developed by which the mean camber surface can be determined for trimmed noncoplanar planforms with minimum vortex drag under subsonic conditions. The method uses a vortex lattice and overcomes previous difficulties with chord loading specification; it uses a Trefftz plane analysis to determine the optimum span loading for minimum drag, then solves for the mean camber surface of the wing which will provide the required loading. Pitching-moment or root-bending-moment constraints can be employed as well at the design lift coefficient. Sensitivity studies of vortex-lattice arrangement have been made with this method and are presented. Comparisons with other theories show generally good agreement. The versatility of the method is demonstrated by applying it to (1) isolated wings, (2) wing-canard configurations, (3) a tandem wing, and (4) a wing-winglet configuration.

Velocity and rolling-moment measurements in the wake of a swept-wing model in the 40 by 80 foot wind tunnel

NASA Technical Reports Server (NTRS)

Rossow, V. J.; Corsiglia, V. R.; Schwind, R. G.; Frick, J. K. D.; Lemmer, O. J.

1975-01-01

Measurements were made in the wake of a swept wing model to study the structure of lift generated vortex wakes shed by conventional span loadings and by several span loadings designed to reduce wake velocities. Variations in the span loading on the swept wing generator were obtained by deflecting seven flap segments on each side by amounts determined by vortex lattice theory to approximate the desired span loadings. The resulting wakes were probed with a three component, hot wire probe to measure velocity, and with a wing to measure the rolling moment that would be induced on a following aircraft. The experimental techniques are described herein, and the measured velocity and rolling moments are presented, along with some comparisons with the applicable theories.

Strain-gage bridge calibration and flight loads measurements on a low-aspect-ratio thin wing

NASA Technical Reports Server (NTRS)

Peele, E. L.; Eckstrom, C. V.

1975-01-01

Strain-gage bridges were used to make in-flight measurements of bending moment, shear, and torque loads on a low-aspect-ratio, thin, swept wing having a full depth honeycomb sandwich type structure. Standard regression analysis techniques were employed in the calibration of the strain bridges. Comparison of the measured loads with theoretical loads are included.

Transonic Aerodynamic Loading Characteristics of a Wing-Body-Tail Combination Having a 52.5 deg. Sweptback Wing of Aspect Ratio 3 With Conical Wing Camber and Body Indentation for a Design Mach Number of Square Root of 2

NASA Technical Reports Server (NTRS)

Cassetti, Marlowe D.; Re, Richard J.; Igoe, William B.

1961-01-01

An investigation has been made of the effects of conical wing camber and body indentation according to the supersonic area rule on the aerodynamic wing loading characteristics of a wing-body-tail configuration at transonic speeds. The wing aspect ratio was 3, taper ratio was 0.1, and quarter-chord-line sweepback was 52.5 deg. with 3-percent-thick airfoil sections. The tests were conducted in the Langley 16-foot transonic tunnel at Mach numbers from 0.80 to 1.05 and at angles of attack from 0 deg. to 14 deg., with Reynolds numbers based on mean aerodynamic chord varying from 7 x 10(exp 6) to 8 x 10(exp 6). Conical camber delayed wing-tip stall and reduced the severity of the accompanying longitudinal instability but did not appreciably affect the spanwise load distribution at angles of attack below tip stall. Body indentation reduced the transonic chordwise center-of-pressure travel from about 8 percent to 5 percent of the mean aerodynamic chord.

Study of aerodynamic technology for single-cruise engine V/STOL fighter/attack aircraft

NASA Technical Reports Server (NTRS)

Driggers, H. H.; Powers, S. A.; Roush, R. T.

1982-01-01

A conceptual design analysis is performed on a single engine V/STOL supersonic fighter/attack concept powered by a series flow tandem fan propulsion system. Forward and aft mounted fans have independent flow paths for V/STOL operation and series flow in high speed flight. Mission, combat and V/STOL performance is calculated. Detailed aerodynamic estimates are made and aerodynamic uncertainties associated with the configuration and estimation methods identified. A wind tunnel research program is developed to resolve principal uncertainties and establish a data base for the baseline configuration and parametric variations.

Gyroscopic sensing in the wings of the hawkmoth Manduca sexta: the role of sensor location and directional sensitivity.

PubMed

Hinson, Brian T; Morgansen, Kristi A

2015-10-06

The wings of the hawkmoth Manduca sexta are lined with mechanoreceptors called campaniform sensilla that encode wing deformations. During flight, the wings deform in response to a variety of stimuli, including inertial-elastic loads due to the wing flapping motion, aerodynamic loads, and exogenous inertial loads transmitted by disturbances. Because the wings are actuated, flexible structures, the strain-sensitive campaniform sensilla are capable of detecting inertial rotations and accelerations, allowing the wings to serve not only as a primary actuator, but also as a gyroscopic sensor for flight control. We study the gyroscopic sensing of the hawkmoth wings from a control theoretic perspective. Through the development of a low-order model of flexible wing flapping dynamics, and the use of nonlinear observability analysis, we show that the rotational acceleration inherent in wing flapping enables the wings to serve as gyroscopic sensors. We compute a measure of sensor fitness as a function of sensor location and directional sensitivity by using the simulation-based empirical observability Gramian. Our results indicate that gyroscopic information is encoded primarily through shear strain due to wing twisting, where inertial rotations cause detectable changes in pronation and supination timing and magnitude. We solve an observability-based optimal sensor placement problem to find the optimal configuration of strain sensor locations and directional sensitivities for detecting inertial rotations. The optimal sensor configuration shows parallels to the campaniform sensilla found on hawkmoth wings, with clusters of sensors near the wing root and wing tip. The optimal spatial distribution of strain directional sensitivity provides a hypothesis for how heterogeneity of campaniform sensilla may be distributed.

Flexible Wing Model for Structural Sizing and Multidisciplinary Design Optimization of a Strut-Braced Wing

NASA Technical Reports Server (NTRS)

Gern, Frank H.; Naghshineh, Amir H.; Sulaeman, Erwin; Kapania, Rakesh K.; Haftka, Raphael T.

2000-01-01

This paper describes a structural and aeroelastic model for wing sizing and weight calculation of a strut-braced wing. The wing weight is calculated using a newly developed structural weight analysis module considering the special nature of strut-braced wings. A specially developed aeroelastic model enables one to consider wing flexibility and spanload redistribution during in-flight maneuvers. The structural model uses a hexagonal wing-box featuring skin panels, stringers, and spar caps, whereas the aerodynamics part employs a linearized transonic vortex lattice method. Thus, the wing weight may be calculated from the rigid or flexible wing spanload. The calculations reveal the significant influence of the strut on the bending material weight of the wing. The use of a strut enables one to design a wing with thin airfoils without weight penalty. The strut also influences wing spanload and deformations. Weight savings are not only possible by calculation and iterative resizing of the wing structure according to the actual design loads. Moreover, as an advantage over the cantilever wing, employment of the strut twist moment for further load alleviation leads to increased savings in structural weight.

KC-135 wing and winglet flight pressure distributions, loads, and wing deflection results with some wind tunnel comparisons

NASA Technical Reports Server (NTRS)

Montoya, L. C.; Jacobs, P.; Flechner, S.; Sims, R.

1982-01-01

A full-scale winglet flight test on a KC-135 airplane with an upper winglet was conducted. Data were taken at Mach numbers from 0.70 to 0.82 at altitudes from 34,000 feet to 39,000 feet at stabilized flight conditions for wing/winglet configurations of basic wing tip, 15/-4 deg, 15/-2 deg, and 0/-4 deg winglet cant/incidence. An analysis of selected pressure distribution and data showed that with the basic wing tip, the flight and wind tunnel wing pressure distribution data showed good agreement. With winglets installed, the effects on the wing pressure distribution were mainly near the tip. Also, the flight and wind tunnel winglet pressure distributions had some significant differences primarily due to the oilcanning in flight. However, in general, the agreement was good. For the winglet cant and incidence configuration presented, the incidence had the largest effect on the winglet pressure distributions. The incremental flight wing deflection data showed that the semispan wind tunnel model did a reasonable job of simulating the aeroelastic effects at the wing tip. The flight loads data showed good agreement with predictions at the design point and also substantiated the predicted structural penalty (load increase) of the 15 deg cant/-2 deg incidence winglet configuration.

Derivation of charts for determining the horizontal tail load variation with any elevator motion

NASA Technical Reports Server (NTRS)

Pearson, Henry A

1943-01-01

The equations relating the wing and tail loads are derived for a unit elevator displacement. These equations are then converted into a nondimensional form and charts are given by which the wing- and tail-load-increment variation may be determined under dynamic conditions for any type of elevator motion and for various degrees of airplane stability. In order to illustrate the use of the charts, several examples are included in which the wing and tail loads are evaluated for a number of types of elevator motion. Methods are given for determining the necessary derivatives from results of wind-tunnel tests when such tests are available.

An exploratory investigation of the STOL landing maneuver

NASA Technical Reports Server (NTRS)

Whyte, P. H.

1979-01-01

The parameters influencing the STOL landing are identified and their effect on the ease and quality of the flare maneuver is discussed. Data from actual landings, supported by pilot commentary and pilot opinion rating, are analyzed. Hypotheses concerning the prediction of STOL handling qualities in the flare are proposed, and suggestions for future research are presented.

V/STOL and STOL ground effects and testing techniques

NASA Technical Reports Server (NTRS)

Kuhn, R. E.

1987-01-01

The ground effects associated with V/STOL operation were examined and an effort was made to develop the equipment and testing techniques needed for that understanding. Primary emphasis was on future experimental programs in the 40 x 80 and the 80 x 120 foot test sections and in the outdoor static test stand associated with these facilities. The commonly used experimental techniques are reviewed and data obtained by various techniques are compared with each other and with available estimating methods. These reviews and comparisons provide insight into the limitations of past studies and the testing techniques used and identify areas where additional work is needed. The understanding of the flow mechanics involved in hovering and in transition in and out of ground effect is discussed. The basic flow fields associated with hovering, transition and STOL operation of jet powered V/STOL aircraft are depicted.

Deflection-Based Aircraft Structural Loads Estimation with Comparison to Flight

NASA Technical Reports Server (NTRS)

Lizotte, Andrew M.; Lokos, William A.

2005-01-01

Traditional techniques in structural load measurement entail the correlation of a known load with strain-gage output from the individual components of a structure or machine. The use of strain gages has proved successful and is considered the standard approach for load measurement. However, remotely measuring aerodynamic loads using deflection measurement systems to determine aeroelastic deformation as a substitute to strain gages may yield lower testing costs while improving aircraft performance through reduced instrumentation weight. With a reliable strain and structural deformation measurement system this technique was examined. The objective of this study was to explore the utility of a deflection-based load estimation, using the active aeroelastic wing F/A-18 aircraft. Calibration data from ground tests performed on the aircraft were used to derive left wing-root and wing-fold bending-moment and torque load equations based on strain gages, however, for this study, point deflections were used to derive deflection-based load equations. Comparisons between the strain-gage and deflection-based methods are presented. Flight data from the phase-1 active aeroelastic wing flight program were used to validate the deflection-based load estimation method. Flight validation revealed a strong bending-moment correlation and slightly weaker torque correlation. Development of current techniques, and future studies are discussed.

Damage Arresting Composites for Shaped Vehicles - Phase II Final Report

NASA Technical Reports Server (NTRS)

Velicki, Alex; Yovanof, Nicolette; Baraja, Jaime; Linton, Kim; Li, Victor; Hawley, Arthur; Thrash, Patrick; DeCoux, Steve; Pickell, Robert

2011-01-01

This report describes the development of a novel structural concept, Pultruded Rod Stitched Efficient Unitized Structure (PRSEUS), that addresses the demanding fuselage loading requirements for the Hybrid Wing or Blended Wing Body (BWB) airplane configuration. In addition to the analytical studies, a three specimen test program was also completed to assess the concept under axial tension loading, axial compression loading, and internal pressure loading.

Theoretical study of aerodynamic characteristics of wings having vortex flow

NASA Technical Reports Server (NTRS)

Reddy, C. S.

1979-01-01

The aerodynamic characteristics of slender wings having separation induced vortex flows are investigated by employing three different computer codes--free vortex sheet, quasi vortex lattice, and suction analogy methods. Their capabilities and limitations are examined, and modifications are discussed. Flat wings of different configurations: arrow, delta, and diamond shapes, as well as cambered delta wings, are studied. The effect of notch ratio on the load distributions and the longitudinal characteristics of a family of arrow and diamond wings is explored. The sectional lift coefficients and the accumulated span loadings are determined for an arrow wing and are seen to be unusual in comparison with the attached flow results. The theoretically predicted results are compared with the existing experimental values.

Real-time monitoring system of composite aircraft wings utilizing Fibre Bragg Grating sensor

NASA Astrophysics Data System (ADS)

Vorathin, E.; Hafizi, Z. M.; Che Ghani, S. A.; Lim, K. S.

2016-10-01

Embedment of Fibre Bragg Grating (FBG) sensor in composite aircraft wings leads to the advancement of structural condition monitoring. The monitored aircraft wings have the capability to give real-time response under critical loading circumstances. The main objective of this paper is to develop a real-time FBG monitoring system for composite aircraft wings to view real-time changes when the structure undergoes some static loadings and dynamic impact. The implementation of matched edge filter FBG interrogation system to convert wavelength variations to strain readings shows that the structure is able to response instantly in real-time when undergoing few loadings and dynamic impact. This smart monitoring system is capable of updating the changes instantly in real-time and shows the weight induced on the composite aircraft wings instantly without any error. It also has a good agreement with acoustic emission (AE) sensor in the dynamic test.

A Summary of Numerous Strain-Gage Load Calibrations on Aircraft Wings and Tails in a Technological Format

NASA Technical Reports Server (NTRS)

Jenkins, Jerald M.; DeAngelis, V. Michael

1997-01-01

Fifteen aircraft structures that were calibrated for flight loads using strain gages are examined. The primary purpose of this paper is to document important examples of load calibrations on airplanes during the past four decades. The emphasis is placed on studying the physical procedures of calibrating strain-gaged structures and all the supporting analyses and computational techniques that have been used. The results and experiences obtained from actual data from 14 structures (on 13 airplanes and 1 laboratory test structure) are presented. This group of structures includes fins, tails, and wings with a wide variety of aspect ratios. Straight- wing, swept-wing, and delta-wing configurations are studied. Some of the structures have skin-dominant construction; others are spar-dominant. Anisotropic materials, heat shields, corrugated components, nonorthogonal primary structures, and truss-type structures are particular characteristics that are included.

An Investigation of Wing and Aileron Loads Due to Deflected Inboard and Outboard Ailerons on a 4-Percent-Thick 30 deg Sweptback Wing at Transonic Speeds

NASA Technical Reports Server (NTRS)

Whitcomb, Charles F.; Critzos, Chris C.; Brown, Philippa F.

1961-01-01

An investigation has been conducted in the Langley 16-foot transonic tunnel to determine the changes in wing loading characteristics due to deflections of a plain faired flap-type inboard aileron, a plain faired flap-type outboard aileron, and a slab-sided thickened trailing edge outboard aileron. The test wing was 4 percent thick and had 30 sweep of the quarter chord, an aspect ratio of 3.0, a taper ratio of 0.2, and NACA 65A004 airfoil sections. The loading characteristics of the deflected ailerons were also investigated. The model was a sting-mounted wing-body combination, and pressure measurements over one wing panel (exposed area) and the ailerons were obtained for angles of attack from 0 to 20 at deflections up to +/- 15 deg for Mach numbers between 0.80 and 1.03. The test Reynolds number based on the wing mean aerodynamic chord was about 7.4 x 10(exp 6). The results of the investigation indicated that positive deflection of the plain faired flap-type inboard aileron caused significant added loading over the wing sections outboard of the aileron at all Mach numbers for model angles of attack from 0 deg or 4 deg up to 12 deg. Positive deflection of the two outboard ailerons (plain faired and slab sided with thickened trailing edge) caused significant added loading over the wing sections inboard of the ailerons for different model angle-of-attack ranges at the several test Mach numbers. The loading shapes over the ailerons were irregular and would be difficult to predict from theoretical considerations in the transonic speed range. The longitudinal and lateral center-of-pressure locations for the ailerons varied only slightly with increasing angle of attack and/or Mach number. Generally, the negative slopes of the variations of aileron hinge-moment coefficient with aileron deflection for all three ailerons varied similarly with Mach number at the test angles of attack.

Advanced prediction technique for the low speed aerodynamics of V/STOL aircraft. Volume 2: User's manual

NASA Technical Reports Server (NTRS)

Beatty, T. D.; Worthey, M. K.

1984-01-01

A computerized prediction method known as the Vought V/STOL Aircraft Propulsive Effects computer program (VAPE) for propulsive induced forces and moments in transition and Short TakeOff and Landing (STOL) flight is improved and evaluated. The VAPE program is capable of evaluating: (1) effects of relative wind about an aircraft, (2) effects of propulsive lift jet entrainment, vorticity and flow blockage, (3) effects of engine inlet flow on the aircraft flow field, (4) engine inlet forces and moments including inlet separation, (5) ground effects in the STOL region of flight, and (6) viscous effects on lifting surfaces.

Learjet Model 55 Wing Analysis with Landing Loads

NASA Technical Reports Server (NTRS)

Boroughs, R. R.

1985-01-01

The NASTRAN analysis was used to determine the impact of new landing loads on the Learjet Model 55 wing. These new landing loads were the result of a performance improvement effort to increase the landing weight of the aircraft to 18,000 lbs. from 17,000 lbs. and extend the life of the tires and brakes by incorporating larger tires and heavy duty brakes. Landing loads for the original 17,000 lb. airplane landing configuration were applied to the full airplane NASTRAN model. The analytical results were correlated with the strain gage data from the original landing load static tests. The landing loads for the 18,000 lb. airplane were applied to the full airplane NASTRAN model, and a comparison was made with the original Model 55 data. The results of this comparison enable Learjet to determine the difference in stress distribution in the wing due to these two different sets of landing loads.

Strain Gage Load Calibration of the Wing Interface Fittings for the Adaptive Compliant Trailing Edge Flap Flight Test

NASA Technical Reports Server (NTRS)

Miller, Eric J.; Holguin, Andrew C.; Cruz, Josue; Lokos, William A.

2014-01-01

The safety-of-flight parameters for the Adaptive Compliant Trailing Edge (ACTE) flap experiment require that flap-to-wing interface loads be sensed and monitored in real time to ensure that the structural load limits of the wing are not exceeded. This paper discusses the strain gage load calibration testing and load equation derivation methodology for the ACTE interface fittings. Both the left and right wing flap interfaces were monitored; each contained four uniquely designed and instrumented flap interface fittings. The interface hardware design and instrumentation layout are discussed. Twenty-one applied test load cases were developed using the predicted in-flight loads. Pre-test predictions of strain gage responses were produced using finite element method models of the interface fittings. Predicted and measured test strains are presented. A load testing rig and three hydraulic jacks were used to apply combinations of shear, bending, and axial loads to the interface fittings. Hardware deflections under load were measured using photogrammetry and transducers. Due to deflections in the interface fitting hardware and test rig, finite element model techniques were used to calculate the reaction loads throughout the applied load range, taking into account the elastically-deformed geometry. The primary load equations were selected based on multiple calibration metrics. An independent set of validation cases was used to validate each derived equation. The 2-sigma residual errors for the shear loads were less than eight percent of the full-scale calibration load; the 2-sigma residual errors for the bending moment loads were less than three percent of the full-scale calibration load. The derived load equations for shear, bending, and axial loads are presented, with the calculated errors for both the calibration cases and the independent validation load cases.

Flow effects with cross-blown lifting jets of V/STOL aircraft and their reactions on aerodynamical forces and moments of the airframe

NASA Technical Reports Server (NTRS)

Viehweger, G.

1977-01-01

Systematic basic studies on the close and distant effects of cross blown single and twin lifting jets were performed with the aid of a principle model. The different effects are described in detail. The number of the experimental parameters is reduced to the most essential ones: (1) the angle of attack, (2) the flight and the jet velocities as well as the jet diameter, (3) the distance between the twin jets, (4) the location of the wing relative to the jets and the fuselage, and (5) the ground distance. The results of systematic pressure distribution measurements on the fuselage surface are studied, especially in the close vicinity of the jet exits. From these results, functions on the influence of the parameters are deduced.

Forecasting the demand potential for STOL air transportation

NASA Technical Reports Server (NTRS)

Fan, S.; Horonjeff, R.; Kanafani, A.; Mogharabi, A.

1973-01-01

A process for predicting the potential demand for STOL aircraft was investigated to provide a conceptual framework, and an analytical methodology for estimating the STOL air transportation market. It was found that: (1) schedule frequency has the strongest effect on the traveler's choice among available routes, (2) work related business constitutes approximately 50% of total travel volume, and (3) air travel demand follows economic trends.

Static Structural Analysis of a Variable Span Morphing Wing for Unmanned Aerial Vehicle

NASA Astrophysics Data System (ADS)

Bashir, M.; Rajendran, P.

2018-05-01

While the primary reason to develop an adaptive wing is the aerodynamic benefits, the primary hindrance is the structural and vibrational considerations due to the unsteady nature of the airflow during the flight. Hence this study forms an important part of the morphable wing technology. In this paper, the design of a moderate aspect ratio variable span wing will be performed. The morphing wing is modeled structurally to observe the effect of spanwise load distribution on the wing structure. For the structural design and analysis of the unmanned aerial vehicle (UAV) under this study, commercial software Solidworks and Ansys/Static Structural/Modal are used. The static structural analyses of the wing are performed under different load conditions. The results of these analyses show that the designed structure is safe within the flight envelope. It is observed that the wing-root bending moment increases drastically due to an increase in the wingspan. Thus, the bending moment along the wingspan of the morphing wing is much larger than that of the conventional wing which results in an increase in the deflection of the free-end. The maximum stress for the un-extended wing configuration increases for the extended wing configuration.

Bone laminarity in the avian forelimb skeleton and its relationship to flight mode: testing functional interpretations.

PubMed

Simons, Erin L R; O'connor, Patrick M

2012-03-01

Wing bone histology in three species of birds was characterized in order to test hypotheses related to the relationship between skeletal microstructure and inferred wing loading during flight. Data on the degree of laminarity (the proportion of circular vascular canals) and the occurrence of secondary osteons were obtained from three species that utilize different primary flight modes: the Double-crested cormorant, a continuous flapper; the Brown pelican, a static soarer; and the Laysan albatross, a dynamic soarer. Laminarity indices were calculated for four quadrants for each of the three main wing elements. Ulnae and carpometacarpi were predicted to exhibit quadrant specific patterns of laminarity due to hypothesized differences in locally applied loads related to the attachment of flight feathers. However, few differences among the quadrants were identified. No significant differences were identified among the three elements, which is notable as different bones are likely experiencing different loading conditions. These results do not support the concept of bone functional adaptation in the primary structure of the wing elements. Significant differences in laminarity were found among the three primary flight modes. The dynamic soaring birds exhibited significantly lower laminarity than the flapping and static soaring birds. These results support the proposed hypothesis that laminarity is an adaptation for resisting torsional loading. This may be explained by overall wing shape: whereas dynamic soaring birds have long slender wings, flappers and static soaring birds have broader wings with a larger wing chord that would necessarily impart a higher torsional moment on the feather-bearing bones. Copyright © 2012 Wiley Periodicals, Inc.

A vortex-lattice method for the mean camber shapes of trimmed noncoplanar planforms with minimum vortex drag

NASA Technical Reports Server (NTRS)

Lamar, J. E.

1976-01-01

A new subsonic method has been developed by which the mean camber surface can be determined for trimmed noncoplanar planforms with minimum vortex drag. This method uses a vortex lattice and overcomes previous difficulties with chord loading specification. A Trefftz plane analysis is utilized to determine the optimum span loading for minimum drag, then solved for the mean camber surface of the wing, which provides the required loading. Sensitivity studies, comparisons with other theories, and applications to configurations which include a tandem wing and a wing winglet combination have been made and are presented.

Testing and Analysis of a Composite Non-Cylindrical Aircraft Fuselage Structure. Part 1; Ultimate Design Loads

NASA Technical Reports Server (NTRS)

Przekop, Adam; Jegley, Dawn C.; Lovejoy, Andrew E.; Rouse, Marshall; Wu, Hsi-Yung T.

2016-01-01

The Environmentally Responsible Aviation Project aimed to develop aircraft technologies enabling significant fuel burn and community noise reductions. Small incremental changes to the conventional metallic alloy-based 'tube and wing' configuration were not sufficient to achieve the desired metrics. One airframe concept identified by the project as having the potential to dramatically improve aircraft performance was a composite-based hybrid wing body configuration. Such a concept, however, presented inherent challenges stemming from, among other factors, the necessity to transfer wing loads through the entire center fuselage section which accommodates a pressurized cabin confined by flat or nearly flat panels. This paper discusses finite element analysis and testing of a large-scale hybrid wing body center section structure developed and constructed to demonstrate that the Pultruded Rod Stitched Efficient Unitized Structure concept can meet these challenging demands of the next generation airframes. Part I of the paper considers the five most critical load conditions, which are internal pressure only and positive and negative g-loads with and without internal pressure. Analysis results are compared with measurements acquired during testing. Performance of the test article is found to be closely aligned with predictions and, consequently, able to support the hybrid wing body design loads in pristine and barely visible impact damage conditions.

BMI Sandwich Wing Box Analysis and Test

NASA Technical Reports Server (NTRS)

Palm, Tod; Mahler, Mary; Shah, Chandu; Rouse, Marshall; Bush, Harold; Wu, Chauncey; Small, William J.

2000-01-01

A composite sandwich single bay wing box test article was developed by Northrop Grumman and tested recently at NASA Langley Research Center. The objectives for the wing box development effort were to provide a demonstration article for manufacturing scale up of structural concepts related to a high speed transport wing, and to validate the structural performance of the design. The box concept consisted of highly loaded composite sandwich wing skins, with moderately loaded composite sandwich spars. The dimensions of the box were chosen to represent a single bay of the main wing box, with a spar spacing of 30 inches, height of 20 inches constant depth, and length of 64 inches. The bismaleimide facesheet laminates and titanium honeycomb core chosen for this task are high temperature materials able to sustain a 300F service temperature. The completed test article is shown in Figure 1. The tests at NASA Langley demonstrated the structures ability to sustain axial tension and compression loads in excess of 20,000 lb/in, and to maintain integrity in the thermal environment. Test procedures, analysis failure predictions, and test results are presented.

Calculation of Wing Bending Moments and Tail Loads Resulting from the Jettison of Wing Tips During a Symmetrical Pull-Up

NASA Technical Reports Server (NTRS)

Boshar, John

1947-01-01

A preliminary analytical investigation was made to determine the feasibility of the basic idea of controlled failure points as safety valves for the primary airplane structure. The present analysis considers the possibilities of the breakable wing tip which, in failing as a weak link, would relieve the bending moments on the wing structure. The analysis was carried out by computing the time histories of the wing and stabilizer angle of attack in a 10g pull-up for an XF8F airplane with tips fixed and comparing the results with those for the same maneuver, that is, elevator motion but with tips jettisoned at 8g. The calculations indicate that the increased stability accompanying the loss of the wing tips reduces the bending moment an additional amount above that which would be expected from the initial loss in lift and the inboard shift in load. The vortex shed when the tips are lost may induce a transient load requiring that the tail be made stronger than otherwise.

Load distribution on a close-coupled wing canard at transonic speeds

NASA Technical Reports Server (NTRS)

Gloss, B. B.; Washburn, K. E.

1977-01-01

This paper reports on a wind-tunnel test where load distributions were obtained at transonic speeds on both the canard and wing surfaces of a closely-coupled wing-canard configuration. The investigation included detailed component and configuration arrangement studies to provide insight into the various aerodynamic interference effects for the leading-edge vortex flow conditions encountered. Data indicate that increasing the Mach number from 0.70 to 0.95 caused the wing leading-edge vortex to burst over the wing when the wing was in the presence of the high canard. For some of the outboard span locations, the leading-edge vortex reattachment streamline intersects the wing trailing edge inboard of these span locations, thus, the Kutta condition was not satisfied. In general, the effect of adding a canard was to reduce the lift inboard and somewhat increase the lift outboard similar to the trends that would have been expected had the flow been attached.

Prediction of Aerodynamic Loading

DTIC Science & Technology

1977-02-01

d~collage et a atterrissage court ou vertical (V/STOL), trois m~thodes thgoriques ont 6t4 d6crites concernant l𔄀valuation de la charge agrodynamique...Aeronautique, No. 44, Mars-Avril 1955, pp. 3-8. 2. Legendret Robert, ŕcoulement subsonique transveral i un secteur angulaire plan," Comptes Renduo, Vol. 2...d’entrde nbtevu dans ces conditions est propos~se ; e~ie est lbas~e coefficient de frottersent Sur des lois a.-sez empirique-. 6tahlieos ý partir de

V/STOL tilt rotor aircraft study. Volume 7: Tilt rotor flight control program feedback studies

NASA Technical Reports Server (NTRS)

Alexander, H. R.; Eason, W.; Gillmore, K.; Morris, J.; Spittle, R.

1973-01-01

An exploratory study has been made of the use of feedback control in tilt rotor aircraft. This has included the use of swashplate cyclic and collective controls and direct lift control. Various sensor and feedback systems are evaluated in relation to blade loads alleviation, improvement in flying qualities, and modal suppression. Recommendations are made regarding additional analytical and wind tunnel investigations and development of feedback systems in the full scale flight vehicle. Estimated costs and schedules are given.

Fundamental Characterization of Spanwise Loading and Trailed Wake Vortices

DTIC Science & Technology

2016-07-01

the close interaction of the tip vortex with a following blade . Such vortex interactions are fundamental determinants of rotor performance, loads, and...wing loading distribution differs from a typical loading on a hovering rotor blade in that the maximum bound circulation occurs at the blade root...and not close to the tip; this is similar to a very highly twisted rotor blade , like a tilt-rotor, in hover. The wing-vortex interaction alters the

Assessment of aerodynamic performance of V/STOL and STOVL fighter aircraft

NASA Technical Reports Server (NTRS)

Nelms, W. P.

1984-01-01

The aerodynamic performance of V/STOL and STOVL fighter/attack aircraft was assessed. Aerodynamic and propulsion/airframe integration activities are described and small and large scale research programs are considered. Uncertainties affecting aerodynamic performance that are associated with special configuration features resulting from the V/STOL requirement are addressed. Example uncertainties relate to minimum drag, wave drag, high angle of attack characteristics, and power induced effects.

Study of quiet turbofan STOL aircraft for short-haul transportation. Volume 5: Economics

NASA Technical Reports Server (NTRS)

1973-01-01

The economic aspects of the STOL aircraft for short-haul air transportation are discussed. The study emphasized the potential market, the preferred operational concepts, the design characteristics, and the economic viability. Three central issues governing economic viability are as follows: (1) operator economics given the market, (2) the required transportation facilities, and (3) the external economic effects of a set of regional STOL transportation systems.

VIEW OF 77710A REACTOR WING, LOOKING NORTHEAST,SHOWING LOADING DOOR TO ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

VIEW OF 777-10A REACTOR WING, LOOKING NORTHEAST,SHOWING LOADING DOOR TO THE PROCESS DEVELOPMENT PILE ROOM. BUILDING 305-A IN BACKGROUND ON LEFT - Physics Assembly Laboratory, Area A/M, Savannah River Site, Aiken, Aiken County, SC

Development of an aerodyanmic theory capable of predicting surface loads on slender wings with vortex flow

NASA Technical Reports Server (NTRS)

Gloss, B. B.; Johnson, F. T.

1976-01-01

The Boeing Commercial Airplane Company developed an inviscid three-dimensional lifting surface method that shows promise in being able to accurately predict loads, subsonic and supersonic, on wings with leading-edge separation and reattachment.

A Comparison of the Experimental and Theoretical Loading over Triangular Wings in Sideslip at Supersonic Speeds

NASA Technical Reports Server (NTRS)

Boyd, John W

1951-01-01

The results of an experimental investigation of the load distribution over two triangular wings in sideslip at Mach numbers from 1.20 to 1.79 are presented and compared with theory. The two wings tested have identical plan form, 45 degrees sweepback of the leading edge, and an aspect ratio of 4.0. One model was composed of round-nose airfoil sections and the other of sharp-nose, biconvex sections. For both wings the maximum thickness of streamwise sections was 6 percent and was located at the 30-percent chord.

An analysis of life expectancy of airplane wings in normal cruising flight

NASA Technical Reports Server (NTRS)

Putnam, Abbott A

1945-01-01

In order to provide a basis for judging the relative importance of wing failure by fatigue and by single intense gusts, an analysis of wing life for normal cruising flight was made based on data on the frequency of atmospheric gusts. The independent variables considered in the analysis included stress-concentration factor, stress-load relation, wing loading, design and cruising speeds, design gust velocity, and airplane size. Several methods for estimating fatigue life from gust frequencies are discussed. The procedure selected for the analysis is believed to be simple and reasonably accurate, though slightly conservative.

Cabin-fuselage-wing structural design concept with engine installation

NASA Technical Reports Server (NTRS)

Ariotti, Scott; Garner, M.; Cepeda, A.; Vieira, J.; Bolton, D.

1993-01-01

The purpose of this project is to provide a fuselage structural assembly and wing structural design that will be able to withstand the given operational parameters and loads provided by Federal Aviation Regulation Part 23 (FAR 23) and the Statement of Work (SOW). The goal is to provide a durable lightweight structure that will transfer the applied loads through the most efficient load path. Areas of producibility and maintainability of the structure will also be addressed. All of the structural members will also meet or exceed the desired loading criteria, along with providing adequate stiffness, reliability, and fatigue life as stated in the SOW. Considerations need to be made for control system routing and cabin heating/ventilation. The goal of the wing structure and carry through structure is also to provide a simple, lightweight structure that will transfer the aerodynamic forces produced by the wing, tailboom, and landing gear. These forces will be channeled through various internal structures sized for the pre-determined loading criteria. Other considerations were to include space for flaps, ailerons, fuel tanks, and electrical and control system routing. The difficulties encountered in the fuselage design include expanding the fuselage cabin to accept a third occupant in a staggered configuration and providing ample volume for their safety. By adding a third person the CG of aircraft will move forward so the engine needs to be moved aft to compensate for the difference in the moment. This required the provisions of a ring frame structure for the new position of the engine mount. The difficulties encountered in the wing structural design include resizing the wing for the increased capacity and weight, and compensating for a large torsion produced by the tail boom by placing a great number of stiffeners inside the boom, which will result in the relocation of the fuel tank. Finally, an adequate carry through structure for the wing and fuselage interface will be designed to effectively transmit loads through the fuselage.

Free-Spinning Wind-Tunnel Tests of a Low-Wing Monoplane with Systematic Changes in Wings and Tails V : Effect of Airplane Relative Density

NASA Technical Reports Server (NTRS)

Seidman, Oscar; Neihouse, A I

1940-01-01

The reported tests are a continuation of an NACA investigation being made in the free-spinning wind tunnel to determine the effects of independent variations in load distribution, wing and tail arrangement, and control disposition on the spin characteristics of airplanes. The standard series of tests was repeated to determine the effect of airplane relative density. Tests were made at values of the relative-density parameter of 6.8, 8.4 (basic), and 12.0; and the results were analyzed. The tested variations in the relative-density parameter may be considered either as variations in the wing loading of an airplane spun at a given altitude, with the radii of gyration kept constant, or as a variation of the altitude at which the spin takes place for a given airplane. The lower values of the relative-density parameter correspond to the lower wing loadings or to the lower altitudes of the spin.

Drones for aerodynamic and structural testing /DAST/ - A status report

NASA Technical Reports Server (NTRS)

Murrow, H. N.; Eckstrom, C. V.

1978-01-01

A program for providing research data on aerodynamic loads and active control systems on wings with supercritical airfoils in the transonic speed range is described. Analytical development, wind tunnel tests, and flight tests are included. A Firebee II target drone vehicle has been modified for use as a flight test facility. The program currently includes flight experiments on two aeroelastic research wings. The primary purpose of the first flight experiment is to demonstrate an active control system for flutter suppression on a transport-type wing. Design and fabrication of the wing are complete and after installing research instrumentation and the flutter suppression system, flight testing is expected to begin in early 1979. The experiment on the second research wing - a fuel-conservative transport type - is to demonstrate multiple active control systems including flutter suppression, maneuver load alleviation, gust load alleviation, and reduce static stability. Of special importance for this second experiment is the development and validation of integrated design methods which include the benefits of active controls in the structural design.

Modeling, Control, and Estimation of Flexible, Aerodynamic Structures

NASA Astrophysics Data System (ADS)

Ray, Cody W.

Engineers have long been inspired by nature’s flyers. Such animals navigate complex environments gracefully and efficiently by using a variety of evolutionary adaptations for high-performance flight. Biologists have discovered a variety of sensory adaptations that provide flow state feedback and allow flying animals to feel their way through flight. A specialized skeletal wing structure and plethora of robust, adaptable sensory systems together allow nature’s flyers to adapt to myriad flight conditions and regimes. In this work, motivated by biology and the successes of bio-inspired, engineered aerial vehicles, linear quadratic control of a flexible, morphing wing design is investigated, helping to pave the way for truly autonomous, mission-adaptive craft. The proposed control algorithm is demonstrated to morph a wing into desired positions. Furthermore, motivated specifically by the sensory adaptations organisms possess, this work transitions to an investigation of aircraft wing load identification using structural response as measured by distributed sensors. A novel, recursive estimation algorithm is utilized to recursively solve the inverse problem of load identification, providing both wing structural and aerodynamic states for use in a feedback control, mission-adaptive framework. The recursive load identification algorithm is demonstrated to provide accurate load estimate in both simulation and experiment.

Structural modeling and optimization of a joined-wing configuration of a High-Altitude Long-Endurance (HALE) aircraft

NASA Astrophysics Data System (ADS)

Kaloyanova, Valentina B.

Recent research trends have indicated an interest in High-Altitude, Long-Endurance (HALE) aircraft as a low-cost alternative to certain space missions, such as telecommunication relay, environmental sensing and military reconnaissance. HALE missions require a light vehicle flying at low speed in the stratosphere at altitudes of 60,000-80,000 ft, with a continuous loiter time of up to several days. To provide high lift and low drag at these high altitudes, where the air density is low, the wing area should be increased, i.e., high-aspect-ratio wings are necessary. Due to its large span and lightweight, the wing structure is very flexible. To reduce the structural deformation, and increase the total lift in a long-spanned wing, a sensorcraft model with a joined-wing configuration, proposed by AFRL, is employed. The joined-wing encompasses a forward wing, which is swept back with a positive dihedral angle, and connected with an aft wing, which is swept forward. The joined-wing design combines structural strength, high aerodynamic performance and efficiency. As a first step to study the joined-wing structural behavior an 1-D approximation model is developed. The 1-D approximation is a simple structural model created using ANSYS BEAM4 elements to present a possible approach for the aerodynamics-structure coupling. The pressure loads from the aerodynamic analysis are integrated numerically to obtain the resultant aerodynamic forces and moments (spanwise lift and pitching moment distributions, acting at the aerodynamic center). These are applied on the 1-D structural model. A linear static analysis is performed under this equivalent load, and the deformed shape of the 1-D model is used to obtain the deformed shape of the actual 3-D joined wing, i.e. deformed aerodynamic surface grid. To date in the existing studies, only simplified structural models have been examined. In the present work, in addition to the simple 1-D beam model, a semi-monocoque structural model is developed. All stringers, skin panels, ribs and spars are represented by appropriate elements in a finite-element model. Also, the model accounts for the fuel weight and sensorcraft antennae housed within the wings. Linear and nonlinear static analyses under the aerodynamic load are performed. The stress distribution in the wing as well as deformation is explored. Starting with a structural model with uniform mass distribution, a design optimization is performed to achieve a fully stressed design. As the joined-wing structure is prone to buckling, after the design optimization is complete linear and nonlinear bucking analyses are performed to study the global joined-wing structural instability, the load magnitude at which it is expected to occur, and the buckling mode. The buckled shape of the aft wing (which is subjected to compression) is found to resemble that of a fixed-pinned column. The linear buckling analysis overestimates the buckling load. However, even the nonlinear buckling analysis results in a load factor higher than 3, i.e. the wing structure is buckling safe under its current loading conditions. As the region of the joint has a very complicated geometry that has adverse effects in the flow and stress behavior an independent, more finely meshed model (submodel) of the joint region is generated and analyzed. A detailed discussion of the stress distribution obtained in the joint region via the submodeling technique is presented in this study as well. It is found out that compared to its structural response, the joint adverse effects are much more pronounced in its aerodynamic response, so it is suggested for future studies the geometry of the joint to be optimized based on its aerodynamic performance. As this design and analysis study is aimed towards developing a realistic structural representation of the innovative joined-wing configuration, in addition to the "global", or upper-level optimization, a local level design optimization is performed as well. At the lower (local) level detailed models of wing structural panels are used to compute more complex failure modes and to design the details that are not included in the upper (global) level model. Proper coordination between local skin-stringer panel models and the global joined-wing model prevents inconsistency between the upper- (global) and lower- (local) level design models. (Abstract shortened by UMI.)

The Calculated Effect of Various Hydrodynamic and Aerodynamic Factors on the Take-Off of a Large Flying Boat

NASA Technical Reports Server (NTRS)

Olson, R.E.; Allison, J.M.

1939-01-01

Present designs for large flying boats are characterized by high wing loading, high aspect ratio, and low parasite drag. The high wing loading results in the universal use of flaps for reducing the takeoff and landing speeds. These factors have an effect on takeoff performance and influence to a certain extent the design of the hull. An investigation was made of the influence of various factors and design parameters on the takeoff performance of a hypothetical large flying boat by means of takeoff calculations. The parameters varied in the calculations were size of hull (load coefficient), wing setting, trim, deflection of flap, wing loading, aspect ratio, and parasite drag. The takeoff times and distances were calculated to the stalling speeds and the performance above these speeds was studied separately to determine piloting technique for optimum takeoff. The advantage of quick deflection of the flap at high water speeds is shown.

Novel deployable morphing wing based on SMP composite

NASA Astrophysics Data System (ADS)

Yu, Kai; Sun, Shouhua; Liu, Liwu; Zhang, Zhen; Liu, Yanju; Leng, Jinsong

2009-07-01

In this paper, a novel kind of deployable morphing wing base on shape memory polymer (SMP) composite is designed and tested. While the deployment of the morphing wing still relies on the mechanisms to ensure the recovery force and the stability performance, the deploying process tends to be more steady and accurate by the application of SMP composite, which overcomes the inherent drawbacks of the traditional one, such as harmful impact to the flight balance, less accuracy during the deployment and complex mechanical masses. On the other hand, SMP composite is also designed as the wing's filler. During its shape recovery process, SMP composite stuffed in the wing helps to form an aerofoil for the wing and withstand the aerodynamic loads, leading to the compressed aerofoil recovering its original shape. To demonstrate the feasibility and the controllability of the designed deployable morphing wing, primary tests are also conducted, including the deploying speed of the morphing wing and SMP filler as the main testing aspects. Finally, Wing's deformation under the air loads is also analyzed by using the finite element method to validate the flight stability.

Dynamics of ultralight aircraft: Motion in vertical gusts

NASA Technical Reports Server (NTRS)

Jones, R. T.

1977-01-01

Gust load calculations are extended to the range of conditions encountered by ultralight aircraft such as hang gliders. Having wing loadings of the order of 5 kg/sq m, these gliders acquire a substantial fraction of the motion of a gust within a distance of 1 or 2 m. Comparative loads and displacements for a small powered airplane having a wing loading of 50 kg sq m and for a commercial jet with 500 kg sq m are shown.

Economic and environmental aspects of STOL transportation

NASA Technical Reports Server (NTRS)

Stewart, E. C.

1972-01-01

A system study to analyze the question of the impact of advanced STOL aircraft in meeting the needs of short-haul air transportation systems is discussed. The study is concerned with the following aspects: (1) service to the passenger, (2) economic viability, and (3) economic criteria to include community noise, ground and air decongestion, and air pollution. The STOL aircraft parameters are defined. Preliminary conclusions concerning the feasibility of short-haul air transportation are presented.

Effect of Winglets on a First-Generation Jet Transport Wing. 2: Pressure and Spanwise Load Distributions for a Semispan Model at High Subsonic Speeds. [in the Langley 8 ft transonic tunnel

NASA Technical Reports Server (NTRS)

Montoya, L. C.; Flechner, S. G.; Jacobs, P. F.

1977-01-01

Pressure and spanwise load distributions on a first-generation jet transport semispan model at high subsonic speeds are presented for the basic wing and for configurations with an upper winglet only, upper and lower winglets, and a simple wing-tip extension. Selected data are discussed to show the general trends and effects of the various configurations.

Some considerations in the design of transport aircraft /The W. Rupert Turnbull Lecture for 1975/

NASA Technical Reports Server (NTRS)

Jones, R. T.

1975-01-01

The slow landing speeds (30 mph, 65 mph) and light wing loading required for safety in the early days of aviation are shown to be irrelevant to safe landings of propeller-driven aircraft, while increases in wing loading and landing speed have been accompanied by improved safety records. This is attributed to length of runway and time available for approach maneuvers, plus immunity to wind gusts and turbulence conferred by higher wing loadings. Aerodynamical and mechanical aspects of safe landing are discussed, with no mention of instruments. Fuel savings achievable through high aspect ratio, variable sweep angle, and supercritical airfoils are also considered.

Nano-mechanical properties and structural of a 3D-printed biodegradable biomimetic micro air vehicle wing

NASA Astrophysics Data System (ADS)

Salami, E.; Montazer, E.; Ward, T. A.; Ganesan, P. B.

2017-06-01

The biomimetic micro air vehicles (BMAV) are unmanned, micro-scaled aircraft that are bio-inspired from flying organisms to achieve the lift and thrust by flapping their wings. The main objectives of this study are to design a BMAV wing (inspired from the dragonfly) and analyse its nano-mechanical properties. In order to gain insights into the flight mechanics of dragonfly, reverse engineering methods were used to establish three-dimensional geometrical models of the dragonfly wings, so we can make a comparative analysis. Then mechanical test of the real dragonfly wings was performed to provide experimental parameter values for mechanical models in terms of nano-hardness and elastic modulus. The mechanical properties of wings were measured by nanoindentre. Finally, a simplified model was designed and the dragonfly-like wing frame structure was bio-mimicked and fabricated using a 3D printer. Then mechanical test of the BMAV wings was performed to analyse and compare the wings under a variety of simplified load regimes that are concentrated force, uniform line-load and a torque. This work opened up the possibility towards developing an engineering basis for the biomimetic design of BMAV wings.

Food load manipulation ability shapes flight morphology in females of central-place foraging Hymenoptera

PubMed Central

2013-01-01

Background Ecological constraints related to foraging are expected to affect the evolution of morphological traits relevant to food capture, manipulation and transport. Females of central-place foraging Hymenoptera vary in their food load manipulation ability. Bees and social wasps modulate the amount of food taken per foraging trip (in terms of e.g. number of pollen grains or parts of prey), while solitary wasps carry exclusively entire prey items. We hypothesized that the foraging constraints acting on females of the latter species, imposed by the upper limit to the load size they are able to transport in flight, should promote the evolution of a greater load-lifting capacity and manoeuvrability, specifically in terms of greater flight muscle to body mass ratio and lower wing loading. Results Our comparative study of 28 species confirms that, accounting for shared ancestry, female flight muscle ratio was significantly higher and wing loading lower in species taking entire prey compared to those that are able to modulate load size. Body mass had no effect on flight muscle ratio, though it strongly and negatively co-varied with wing loading. Across species, flight muscle ratio and wing loading were negatively correlated, suggesting coevolution of these traits. Conclusions Natural selection has led to the coevolution of resource load manipulation ability and morphological traits affecting flying ability with additional loads in females of central-place foraging Hymenoptera. Release from load-carrying constraints related to foraging, which took place with the evolution of food load manipulation ability, has selected against the maintenance of a powerful flight apparatus. This could be the case since investment in flight muscles may have to be traded against other life-history traits, such as reproductive investment. PMID:23805850

A predictive pilot model for STOL aircraft landing

NASA Technical Reports Server (NTRS)

Kleinman, D. L.; Killingsworth, W. R.

1974-01-01

An optimal control approach has been used to model pilot performance during STOL flare and landing. The model is used to predict pilot landing performance for three STOL configurations, each having a different level of automatic control augmentation. Model predictions are compared with flight simulator data. It is concluded that the model can be effective design tool for studying analytically the effects of display modifications, different stability augmentation systems, and proposed changes in the landing area geometry.

Global and Local Stress Analyses of McDonnell Douglas Stitched/RFI Composite Wing Stub Box

NASA Technical Reports Server (NTRS)

Wang, John T.

1996-01-01

This report contains results of structural analyses performed in support of the NASA structural testing of an all-composite stitched/RFI (resin film infusion) wing stub box. McDonnell Douglas Aerospace Company designed and fabricated the wing stub box. The analyses used a global/local approach. The global model contains the entire test article. It includes the all-composite stub box, a metallic load-transition box and a metallic wing-tip extension box. The two metallic boxes are connected to the inboard and outboard ends of the composite wing stub box, respectively. The load-transition box was attached to a steel and concrete vertical reaction structure and a load was applied at the tip of the extension box to bend the wing stub box upward. The local model contains an upper cover region surrounding three stringer runouts. In that region, a large nonlinear deformation was identified by the global analyses. A more detailed mesh was used for the local model to obtain more accurate analysis results near stringer runouts. Numerous analysis results such as deformed shapes, displacements at selected locations, and strains at critical locations are included in this report.

The power induced effects module: A FORTRAN code which estimates lift increments due to power induced effects for V/STOL flight

NASA Technical Reports Server (NTRS)

Sandlin, Doral R.; Howard, Kipp E.

1991-01-01

A user friendly FORTRAN code that can be used for preliminary design of V/STOL aircraft is described. The program estimates lift increments, due to power induced effects, encountered by aircraft in V/STOL flight. These lift increments are calculated using empirical relations developed from wind tunnel tests and are due to suckdown, fountain, ground vortex, jet wake, and the reaction control system. The code can be used as a preliminary design tool along with NASA Ames' Aircraft Synthesis design code or as a stand-alone program for V/STOL aircraft designers. The Power Induced Effects (PIE) module was validated using experimental data and data computed from lift increment routines. Results are presented for many flat plate models along with the McDonnell Aircraft Company's MFVT (mixed flow vectored thrust) V/STOL preliminary design and a 15 percent scale model of the YAV-8B Harrier V/STOL aircraft. Trends and magnitudes of lift increments versus aircraft height above the ground were predicted well by the PIE module. The code also provided good predictions of the magnitudes of lift increments versus aircraft forward velocity. More experimental results are needed to determine how well the code predicts lift increments as they vary with jet deflection angle and angle of attack. The FORTRAN code is provided in the appendix.

Multi-Objective Flight Control for Drag Minimization and Load Alleviation of High-Aspect Ratio Flexible Wing Aircraft

NASA Technical Reports Server (NTRS)

Nguyen, Nhan; Ting, Eric; Chaparro, Daniel; Drew, Michael; Swei, Sean

2017-01-01

As aircraft wings become much more flexible due to the use of light-weight composites material, adverse aerodynamics at off-design performance can result from changes in wing shapes due to aeroelastic deflections. Increased drag, hence increased fuel burn, is a potential consequence. Without means for aeroelastic compensation, the benefit of weight reduction from the use of light-weight material could be offset by less optimal aerodynamic performance at off-design flight conditions. Performance Adaptive Aeroelastic Wing (PAAW) technology can potentially address these technical challenges for future flexible wing transports. PAAW technology leverages multi-disciplinary solutions to maximize the aerodynamic performance payoff of future adaptive wing design, while addressing simultaneously operational constraints that can prevent the optimal aerodynamic performance from being realized. These operational constraints include reduced flutter margins, increased airframe responses to gust and maneuver loads, pilot handling qualities, and ride qualities. All of these constraints while seeking the optimal aerodynamic performance present themselves as a multi-objective flight control problem. The paper presents a multi-objective flight control approach based on a drag-cognizant optimal control method. A concept of virtual control, which was previously introduced, is implemented to address the pair-wise flap motion constraints imposed by the elastomer material. This method is shown to be able to satisfy the constraints. Real-time drag minimization control is considered to be an important consideration for PAAW technology. Drag minimization control has many technical challenges such as sensing and control. An initial outline of a real-time drag minimization control has already been developed and will be further investigated in the future. A simulation study of a multi-objective flight control for a flight path angle command with aeroelastic mode suppression and drag minimization demonstrates the effectiveness of the proposed solution. In-flight structural loads are also an important consideration. As wing flexibility increases, maneuver load and gust load responses can be significant and therefore can pose safety and flight control concerns. In this paper, we will extend the multi-objective flight control framework to include load alleviation control. The study will focus initially on maneuver load minimization control, and then subsequently will address gust load alleviation control in future work.

The Effects of Warhead-Induced Damage on the Aeroelastic Characteristics of Lifting Surfaces. Volume I. Aeroelastic Effects.

DTIC Science & Technology

1980-07-01

Arnold. Some further insight into the problem is obtained here, however, when it is demonstrated that highly optimized structural designs may...aircraft of this type are normally designed to withstand loads up to 1.5 times the maximum limit load (load factor 8.0 to 8.67), the structure should...on the wing, for example, give rise to concentrated drag and chordwise loadings as does the recoil from firing wing mounted gun systems . The drag on

Charts for Determining Preliminary Values of Span-load, Shear, Bending-moment, and Accumulated-torque Distributions of Swept Wings of Various Taper Ratios

NASA Technical Reports Server (NTRS)

Wollner, Bertram C

1948-01-01

Contains charts for use in determining preliminary values of the spanwise-load, shear, bending-moment, and accumulated-torque distributions of swept wings. The charts are based on strip theory and include four aerodynamic-load distributions, two section-moment distributions, and two inertia-load distributions. The taper ratios considered cover the range from 1.0 to 0 and the results are applicable to any angle of sweep.

Buckling behavior of Rene 41 tubular panels for a hypersonic aircraft wing

NASA Technical Reports Server (NTRS)

Ko, W. L.; Fields, R. A.; Shideler, J. L.

1986-01-01

The buckling characteristics of Rene 41 tubular panels for a hypersonic aircraft wing were investigated. The panels were repeatedly tested for buckling characteristics using a hypersonic wing test structure and a universal tension/compression testing machine. The nondestructive buckling tests were carried out under different combined load conditions and in different temperature environments. The force/stiffness technique was used to determine the buckling loads of the panels. In spite of some data scattering resulting from large extrapolations of the data-fitting curve (because of the termination of applied loads at relatively low percentages of the buckling loads), the overall test data correlate fairly well with theoretically predicted buckling interaction curves. Also, the structural efficiency of the tubular panels was found to be slightly higher than that of beaded panels.

Buckling behavior of Rene 41 tubular panels for a hypersonic aircraft wing

NASA Technical Reports Server (NTRS)

Ko, W. L.; Shideler, J. L.; Fields, R. A.

1986-01-01

The buckling characteristics of Rene 41 tubular panels for a hypersonic aircraft wing were investigated. The panels were repeatedly tested for buckling characteristics using a hypersonic wing test structure and a universal tension/compression testing machine. The nondestructive buckling tests were carried out under different combined load conditions and in different temperature environments. The force/stiffness technique was used to determine the buckling loads of the panel. In spite of some data scattering, resulting from large extrapolations of the data fitting curve (because of the termination of applied loads at relatively low percentages of the buckling loads), the overall test data correlate fairly well with theoretically predicted buckling interaction curves. Also, the structural efficiency of the tubular panels was found to be slightly higher than that of beaded panels.

A Flight Examination of Operating Problems of V/STOL Aircraft in STOL-Type Landing and Approach

NASA Technical Reports Server (NTRS)

Innis, Robert C.; Quigley, Hervey C.

1961-01-01

A flight investigation has been conducted using a large twin-engine cargo aircraft to isolate the problems associated with operating propeller-driven aircraft in the STOL speed range where appreciable engine power is used to augment aerodynamic lift. The problems considered would also be representative of those of a large overloaded VTOL aircraft operating in an STOL manner with comparable thrust-to-weight ratios. The study showed that operation at low approach speeds was compromised by the necessity of maintaining high thrust to generate high lift and yet achieving the low lift-drag ratios needed for steep descents. The useable range of airspeed and flight path angle was limited by the pilot's demand for a positive climb margin at the approach speed, a suitable stall margin, and a control and/or performance margin for one engine inoperative. The optimum approach angle over an obstacle was found to be a compromise between obtaining the shortest air distance and the lowest touchdown velocity. In order to realize the greatest low-speed potential from STOL designs, the stability and control characteristics must be satisfactory.

Aerostructural optimization of a morphing wing for airborne wind energy applications

NASA Astrophysics Data System (ADS)

Fasel, U.; Keidel, D.; Molinari, G.; Ermanni, P.

2017-09-01

Airborne wind energy (AWE) vehicles maximize energy production by constantly operating at extreme wing loading, permitted by high flight speeds. Additionally, the wide range of wind speeds and the presence of flow inhomogeneities and gusts create a complex and demanding flight environment for AWE systems. Adaptation to different flow conditions is normally achieved by conventional wing control surfaces and, in case of ground generator-based systems, by varying the reel-out speed. These control degrees of freedom enable to remain within the operational envelope, but cause significant penalties in terms of energy output. A significantly greater adaptability is offered by shape-morphing wings, which have the potential to achieve optimal performance at different flight conditions by tailoring their airfoil shape and lift distribution at different levels along the wingspan. Hence, the application of compliant structures for AWE wings is very promising. Furthermore, active gust load alleviation can be achieved through morphing, which leads to a lower weight and an expanded flight envelope, thus increasing the power production of the AWE system. This work presents a procedure to concurrently optimize the aerodynamic shape, compliant structure, and composite layup of a morphing wing for AWE applications. The morphing concept is based on distributed compliance ribs, actuated by electromechanical linear actuators, guiding the deformation of the flexible—yet load-carrying—composite skin. The goal of the aerostructural optimization is formulated as a high-level requirement, namely to maximize the average annual power production per wing area of an AWE system by tailoring the shape of the wing, and to extend the flight envelope of the wing by actively alleviating gust loads. The results of the concurrent multidisciplinary optimization show a 50.7% increase of extracted power with respect to a sequentially optimized design, highlighting the benefits of morphing and the potential of the proposed approach.

Design & fabrication of two seated aircraft with an advanced rotating leading edge wing

NASA Astrophysics Data System (ADS)

Al Ahmari, Saeed Abdullah Saeed

The title of this thesis is "Design & Fabrication of two Seated Aircraft with an Advanced Rotating Leading Edge Wing", this gives almost a good description of the work has been done. In this research, the moving surface boundary-layer control (MSBC) concept was investigated and implemented. An experimental model was constructed and tested in wind tunnel to determine the aerodynamic characteristics using the leading edge moving surface of modified semi-symmetric airfoil NACA1214. The moving surface is provided by a high speed rotating cylinder, which replaces the leading edge of the airfoil. The angle of attack, the cylinder surfaces velocity ratio Uc/U, and the flap deflection angle effects on the lift and drag coefficients and the stall angle of attack were investigated. This new technology was applied to a 2-seat light-sport aircraft that is designed and built in the Aerospace Engineering Department at KFUPM. The project team is led by the aerospace department chairman Dr. Ahmed Z. AL-Garni and Dr. Wael G. Abdelrahman and includes graduate and under graduate student. The wing was modified to include a rotating cylinder along the leading edge of the flap portion. This produced very promising results such as the increase of the maximum lift coefficient at Uc/U=3 by 82% when flaps up and 111% when flaps down at 40° and stall was delayed by 8degrees in both cases. The laboratory results also showed that the effective range of the leading-edge rotating cylinder is at low angles of attack which reduce the need for higher angles of attack for STOL aircraft.

Pressure Distribution Over a Thick, Tapered and Twisted Monoplane Wing Model-NACA 81-J

NASA Technical Reports Server (NTRS)

Wenzinger, Carl J

1932-01-01

This reports presents the results of pressure distribution tests on a thick, tapered and twisted monoplane wing model. The investigation was conducted for the purpose of obtaining data on the aerodynamic characteristics of the new wing and to provide additional information suitable for use in the design of tapered cantilever wings. The tests included angles of attack up to 90 degrees. The span loading over the wing was approximately of elliptical shape, which gave rise to relatively small bending moments about the root. The angle of zero lift for all sections along the span varied only within plus or minus 0.4 degree of the angle of zero lift for the whole wing, resulting in small leading edge loads for the high-speed condition of flight. The results also add to the available information for the study of large angles of attack.

A Method of Determining Aerodynamic-Influence Coefficients from Wind-Tunnel Data for Wings at Supersonic Speeds

NASA Technical Reports Server (NTRS)

Gainer, Patrick A.

1961-01-01

A method is described for determining aerodynamic-influence coefficients from wind-tunnel data for calculating the steady-state load distribution on a wing with arbitrary angle-of-attack distribution at supersonic speeds. The method combines linearized theory with empirical adjustments in order to give accurate results over a wide range of angles of attack. The experimented data required are pressure distributions measured on a flat wing of the desired planform at the desired Mach number and over the desired range of angles of attack. The method has been tested by applying it to wind-tunnel data measured at Mach numbers of 1.61 and 2.01 on wings of the same planform but of different surface shapes. Influence coefficients adjusted to fit the flat wing gave good predictions of the spanwise and chord-wise distributions of loadings measured on twisted and cambered wings.

Residual strength and crack propagation tests on C-130 airplane center wings with service-imposed fatigue damage

NASA Technical Reports Server (NTRS)

Snider, H. L.; Reeder, F. L.; Dirkin, W. J.

1972-01-01

Fourteen C-130 airplane center wings, each containing service-imposed fatigue damage resulting from 4000 to 13,000 accumulated flight hours, were tested to determine their fatigue crack propagation and static residual strength characteristics. Eight wings were subjected to a two-step constant amplitude fatigue test prior to static testing. Cracks up to 30 inches long were generated in these tests. Residual static strengths of these wings ranged from 56 to 87 percent of limit load. The remaining six wings containing cracks up to 4 inches long were statically tested as received from field service. Residual static strengths of these wings ranged from 98 to 117 percent of limit load. Damage-tolerant structural design features such as fastener holes, stringers, doublers around door cutouts, and spanwise panel splices proved to be effective in retarding crack propagation.

Effect of winglets on a first-generation jet transport wing. 3: Pressure and spanwise load distributions for a semispan model at Mach 0.30. [in the Langley 8 ft transonic tunnel

NASA Technical Reports Server (NTRS)

Montoya, L. C.; Jacobs, P. F.; Flechner, S. G.

1977-01-01

Pressure and spanwise load distributions on a first-generation jet transport semispan model at a Mach number of 0.30 are given for the basic wing and for configurations with an upper winglet only, upper and lower winglets, and a simple wing-tip extension. To simulate second-segment-climb lift conditions, leading- and/or trailing-edge flaps were added to some configurations.

Adaptive wing static aeroelastic roll control

NASA Astrophysics Data System (ADS)

Ehlers, Steven M.; Weisshaar, Terrence A.

1993-09-01

Control of the static aeroelastic characteristics of a swept uniform wing in roll using an adaptive structure is examined. The wing structure is modeled as a uniform beam with bending and torsional deformation freedom. Aerodynamic loads are obtained from strip theory. The structure model includes coefficients representing torsional and bending actuation provided by embedded piezoelectric material layers. The wing is made adaptive by requiring the electric field applied to the piezoelectric material layers to be proportional to the wing root loads. The proportionality factor, or feedback gain, is used to control static aeroelastic rolling properties. Example wing configurations are used to illustrate the capabilities of the adaptive structure. The results show that rolling power, damping-in-roll and aileron effectiveness can be controlled by adjusting the feedback gain. And that dynamic pressure affects the gain required. Gain scheduling can be used to set and maintain rolling properties over a range of dynamic pressures. An adaptive wing provides a method for active aeroelastic tailoring of structural response to meet changing structural performance requirements during a roll maneuver.

Sonic environment of aircraft structure immersed in a supersonic jet flow stream

NASA Technical Reports Server (NTRS)

Guinn, W. A.; Balena, F. J.; Soovere, J.

1976-01-01

Test methods for determining the sonic environment of aircraft structure that is immersed in the flow stream of a high velocity jet or that is subjected to the noise field surrounding the jet, were investigated. Sonic environment test data measured on a SCAT 15-F model in the flow field of Mach 1.5 and 2.5 jets were processed. Narrow band, lateral cross correlation and noise contour plots are presented. Data acquisition and reduction methods are depicted. A computer program for scaling the model data is given that accounts for model size, jet velocity, transducer size, and jet density. Comparisons of scaled model data and full size aircraft data are made for the L-1011, S-3A, and a V/STOL lower surface blowing concept. Sonic environment predictions are made for an engine-over-the-wing SST configuration.

Forward velocity effects on fan noise and the suppression characteristics of advanced inlets as measured in the NASA Ames 40 by 80 foot wind tunnel: Acoustic data report

NASA Technical Reports Server (NTRS)

Moore, M. T.

1981-01-01

Forward velocity effects on the forward radiated fan noise and on the suppression characteristics of three advanced inlets relative to a baseline cylindrical inlet were measured in a wind tunnel. A modified JT15D turbofan engine in a quiet nacelle was the source of fan noise; the advanced inlets were a CTOL hybrid inlet, an STOL hybrid inlet, and a treated deflector inlet. Also measured were the static to flight effects on the baseline inlet noise and the effects on the fan noise of canting the baseline inlet 4 deg downward to simulate typical wing mounted turbofan engines. The 1/3 octave band noise data from these tests are given along with selected plots of 1/3 octave band spectra and directivity and full scale PNL directivities. The test facilities and data reduction techniques used are also described.

The oblique impingement of an axisymmetric jet. [flow characteristics of jet flow over flat plates

NASA Technical Reports Server (NTRS)

Foss, J. F.; Kleis, S. J.

1976-01-01

The mechanics of the oblique impingement of an axisymmetric jet on a plane surface are examined in detail. The stagnation point is discussed. A schematic drawing of the problem and coordinate system used to describe the flow field are given. The kinematic features of the flow above the plate are examined in the context of the conservation of mass, the vorticity of the jet, and the vorticity introduced by the jetplate interaction. The dynamic features of the flow are examined in terms of the surface pressure distribution and the cause-effect relationships which exist between the pressure and velocity/vorticity distributions. Flow calculations performed are given. The investigation is relevant to the flow resulting from the interaction of the propulsion jet with the main airfoil (STOL aircraft), and is appropriate to an over- or under- wing configuration.

Fiber Optic System Test Results In A Tactical Military Aircraft

NASA Astrophysics Data System (ADS)

Uhlhorn, Roger W.; Greenwell, Roger A.

1980-09-01

The YAV-8B Electromagnetic Immunity and Flight-Test Program was established to evaluate the susceptibility of wire and optical fiber signal transmission lines to electromagnetic interference when these lines are installed in a graphite/epoxy composite wing and to demonstrate the flightworthiness of fiber optics interconnects in the vertical/ short takeoff and landing aircraft environment. In response, two fiber optic systems were designed, fabricated, and flight tested by McDonnell Aircraft Co. (MCAIR), a division of the McDonnell Douglas Corporation, on the two YAV-8B V/STOL flight test aircraft. The program successfully demonstrated that fiber optics are compatible with the attack aircraft environment. As a result, the full scale development AV-8B will incorporate fiber optics in a point-to-point data link. We describe here the fiber optic systems designs, test equipment development, cabling and connection requirements, fabrication and installation experience, and flight test program results.

Interface concerns of ejector integration in V/STOL aircraft

NASA Technical Reports Server (NTRS)

Lowry, R. B.

1979-01-01

A number of areas which have in the past contributed to weight, complexity, and thrust losses in the ejector-powered V/STOL vehicle were identified. Most of these interfaces taken singly do not represent a severe compromise to the vehicle; however, the bottom line is that the sum of compromises and the subsequent effects on performance, flight operations and maintenance have rendered the ejector V/STOL aircraft unattractive. In addition to some of the unique ejector/aircraft integration problems, the vehicle by virtue of having a V/STOL capability, is compromised in other areas. To be successful and acceptable, the advantages must outweight the disadvantages and simplicity with minimum penalties must be the rule. It is concluded that more emphasis must be placed on the ejector/aircraft interface for the concept to be successful.

Historical overview of V/STOL aircraft technology

NASA Technical Reports Server (NTRS)

Anderson, S. B.

1981-01-01

The requirements for satisfactory characteristics in several key technology areas are discussed and a review is made of various V/STOL aircraft for the purpose of assessing the success or failure of each design in meeting design requirements. Special operating techniques were developed to help circumvent deficiencies. For the most part performance and handling qualities limitations restricted operational evaluations. Flight operations emphasized the need for good STOL performance, good handling qualities, and stability and control augmentation. The majority of aircraft suffered adverse ground effects.

Experimental and Theoretical Study of a Rectangular Wing in a Vortical Wake at Low Speed

NASA Technical Reports Server (NTRS)

Smith, Willard G.; Lazzeroni, Frank A.

1960-01-01

A systematic study has been made, experimentally and theoretically, of the effects of a vortical wake on the aerodynamic characteristics of a rectangular wing at subsonic speed. The vortex generator and wing were mounted on a reflection plane to avoid body-wing interference. Vortex position, relative to the wing, was varied both in the spanwise direction and normal to the wing. Angle of attack of the wing was varied from -40 to +60. Both chordwise and spanwise pressure distributions were obtained with the wing in uniform and vortical flow fields. Stream surveys were made to determine the flow characteristics in the vortical wake. The vortex-induced lift was calculated by several theoretical methods including strip theory, reverse-flow theory, and reverse-flow theory including a finite vortex core. In addition, the Prandtl lifting-line theory and the Weissinger theory were used to calculate the spanwise distribution of vortex-induced loads. With reverse-flow theory, predictions of the interference lift were generally good, and with Weissinger's theory the agreement between the theoretical spanwise variation of induced load and the experimental variation was good. Results of the stream survey show that the vortex generated by a lifting surface of rectangular plan form tends to trail back streamwise from the tip and does not approach the theoretical location, or centroid of circulation, given by theory. This discrepancy introduced errors in the prediction of vortex interference, especially when the vortex core passed immediately outboard of the wing tip. The wake produced by the vortex generator in these tests was not fully rolled up into a circular vortex, and so lacked symmetry in the vertical direction of the transverse plane. It was found that the direction of circulation affected the induced loads on the wing either when the wing was at angle of attack or when the vortex was some distance away from the plane of the wing.

Asymmetry costs: effects of wing damage on hovering flight performance in the hawkmoth Manduca sexta.

PubMed

Fernández, María José; Driver, Marion E; Hedrick, Tyson L

2017-10-15

Flight performance is fundamental to the fitness of flying organisms. Whilst airborne, flying organisms face unavoidable wing wear and wing area loss. Many studies have tried to quantify the consequences of wing area loss to flight performance with varied results, suggesting that not all types of damage are equal and different species may have different means to compensate for some forms of wing damage with little to no cost. Here, we investigated the cost of control during hovering flight with damaged wings, specifically wings with asymmetric and symmetric reductions in area, by measuring maximum load lifting capacity and the metabolic power of hovering flight in hawkmoths ( Manduca sexta ). We found that while asymmetric and symmetric reductions are both costly in terms of maximum load lifting and hovering efficiency, asymmetric reductions are approximately twice as costly in terms of wing area lost. The moths also did not modulate flapping frequency and amplitude as predicted by a hovering flight model, suggesting that the ability to do so, possibly tied to asynchronous versus synchronous flight muscles, underlies the varied responses found in different wing clipping experiments. © 2017. Published by The Company of Biologists Ltd.

Analytical model of the structureborne interior noise induced by a propeller wake

NASA Technical Reports Server (NTRS)

Junger, M. C.; Garrelick, J. M.; Martinez, R.; Cole, J. E., III

1984-01-01

The structure-borne contribution to the interior noise that is induced by the propeller wake acting on the wing was studied. Analytical models were developed to describe each aspect of this path including the excitation loads, the wing and fuselage structures, and the interior acoustic space. The emphasis is on examining a variety of parameters, and as a result different models were developed to examine specific parameters. The excitation loading on the wing by the propeller wake is modeled by a distribution of rotating potential vortices whose strength is related to the thrust per blade. The response of the wing to this loading is examined using beam models. A model of a beam structurally connected to a cylindrical shell with an internal acoustic fluid was developed to examine the coupling of energy from the wing to the interior space. The model of the acoustic space allows for arbitrary end conditions (e.g., rigid or vibrating end caps). Calculations are presented using these models to compare with a laboratory test configuration as well as for parameters of a prop-fan aircraft.

Numerical Characterization of a Composite Bonded Wing-Box

NASA Technical Reports Server (NTRS)

Smeltzer, Stanley S., III; Lovejoy, Andrew E.; Satyanarayana, Arunkumar

2008-01-01

The development of composite wing structures has focused on the use of mechanical fasteners to join heavily-loaded areas, while bonded joints have been used only for select locations. The focus of this paper is the examination of the adhesive layer in a generic bonded wing box that represents a "fastenerless" or unitized structure in order to characterize the general behavior and failure mechanisms. A global/local approach was applied to study the response of the adhesive layer using a global shell model and a local shell/solid model. The wing box was analyzed under load to represent a high-g up-bending condition such that the strains in the composite sandwich face sheets are comparable to an expected design allowable. The global/local analysis indicates that at these wing load levels the strains in the adhesive layer are well within the adhesive's elastic region, such that yielding would not be expected in the adhesive layer. The global/local methodology appears to be a promising approach to evaluate the structural integrity of the adhesively bonded structures.

Aeroacoustic theory for noncompact wing-gust interaction

NASA Technical Reports Server (NTRS)

Martinez, R.; Widnall, S. E.

1981-01-01

Three aeroacoustic models for noncompact wing-gust interaction were developed for subsonic flow. The first is that for a two dimensional (infinite span) wing passing through an oblique gust. The unsteady pressure field was obtained by the Wiener-Hopf technique; the airfoil loading and the associated acoustic field were calculated, respectively, by allowing the field point down on the airfoil surface, or by letting it go to infinity. The second model is a simple spanwise superposition of two dimensional solutions to account for three dimensional acoustic effects of wing rotation (for a helicopter blade, or some other rotating planform) and of finiteness of wing span. A three dimensional theory for a single gust was applied to calculate the acoustic signature in closed form due to blade vortex interaction in helicopters. The third model is that of a quarter infinite plate with side edge through a gust at high subsonic speed. An approximate solution for the three dimensional loading and the associated three dimensional acoustic field in closed form was obtained. The results reflected the acoustic effect of satisfying the correct loading condition at the side edge.

14 CFR 23.573 - Damage tolerance and fatigue evaluation of structure.

Code of Federal Regulations, 2012 CFR

2012-01-01

... result in catastrophic loss of the airplane, the limit load capacity must be substantiated by one of the... equal to 1.1 times the normal operating differential pressure without any other load. [Doc. No. 26269... would result in catastrophic loss of the airplane, in each wing (including canards, tandem wings, and...

14 CFR 23.573 - Damage tolerance and fatigue evaluation of structure.

Code of Federal Regulations, 2010 CFR

2010-01-01

... result in catastrophic loss of the airplane, the limit load capacity must be substantiated by one of the... equal to 1.1 times the normal operating differential pressure without any other load. [Doc. No. 26269... would result in catastrophic loss of the airplane, in each wing (including canards, tandem wings, and...

14 CFR 23.573 - Damage tolerance and fatigue evaluation of structure.

Code of Federal Regulations, 2014 CFR

2014-01-01

... result in catastrophic loss of the airplane, the limit load capacity must be substantiated by one of the... equal to 1.1 times the normal operating differential pressure without any other load. [Doc. No. 26269... would result in catastrophic loss of the airplane, in each wing (including canards, tandem wings, and...

14 CFR 23.573 - Damage tolerance and fatigue evaluation of structure.

Code of Federal Regulations, 2013 CFR

2013-01-01

... result in catastrophic loss of the airplane, the limit load capacity must be substantiated by one of the... equal to 1.1 times the normal operating differential pressure without any other load. [Doc. No. 26269... would result in catastrophic loss of the airplane, in each wing (including canards, tandem wings, and...

14 CFR 23.573 - Damage tolerance and fatigue evaluation of structure.

Code of Federal Regulations, 2011 CFR

2011-01-01

... result in catastrophic loss of the airplane, the limit load capacity must be substantiated by one of the... equal to 1.1 times the normal operating differential pressure without any other load. [Doc. No. 26269... would result in catastrophic loss of the airplane, in each wing (including canards, tandem wings, and...

Range Performance of Bombers Powered by Turbine-Propeller Power Plants

NASA Technical Reports Server (NTRS)

Cline, Charles W.

1950-01-01

Calculations have been made to find range? attainable by bombers of gross weights from l40,000 to 300,000 pounds powered by turbine-propeller power plants. Only conventional configurations were considered and emphasis was placed upon using data for structural and aerodynamic characteristics which are typical of modern military airplanes. An effort was made to limit the various parameters invoked in the airplane configuration to practical values. Therefore, extremely high wing loadings, large amounts of sweepback, and very high aspect ratios have not been considered. Power-plant performance was based upon the performance of a typical turbine-propeller engine equipped with propellers designed to maintain high efficiencies at high-subsonic speeds. Results indicated, in general, that the greatest range, for a given gross weight, is obtained by airplanes of high wing loading, unless the higher cruising speeds associated with the high-wing-loading airplanes require-the use of thinner wing sections. Further results showed the effect of cruising at-high speeds, of operation at very high altitudes, and of carrying large bomb loads.

A Comparison of Three Theoretical Methods of Calculating Span Load Distribution on Swept Wings

NASA Technical Reports Server (NTRS)

VanDorn, Nicholas H.; DeYoung, John

1947-01-01

Three methods for calculating span load distribution, those developed by V.M Falkner, Wm. Mutterperl, and J. Weissinger, have been applied to five swept wings. The angles of sweep ranged from -45 degrees to +45 degrees. These methods were examined to establish their relative accuracy and case of application. Experimentally determined loadings were used as a basis for judging accuracy. For the convenience of the readers the computing forms and all information requisite to their application are included in appendixes. From the analysis it was found that the Weissinger method would be best suited to an over-all study of the effects of plan form on the span loading and associated characteristics of wings. The method gave good, but not best, accuracy and involved by far the least computing effort. The Falkner method gave the best accuracy but at a considerable expanse in computing effort and hence appeared to be most useful for a detailed study of a specific wing. The Mutterperl method offered no advantages in accuracy of facility over either of the other methods and hence is not recommended for use.

Test results from large wing and fuselage panels

NASA Technical Reports Server (NTRS)

Madan, Ram C.; Voldman, Mike

1993-01-01

This paper presents the first results in an assessment of the strength, stiffness, and damage tolerance of stiffened wing and fuselage subcomponents. Under this NASA funded program, 10 large wing and fuselage panels, variously fabricated by automated tow placement and dry-stitched preform/resin transfer molding, are to be tested. The first test of an automated tow placement six-longeron fuselage panel under shear load was completed successfully. Using NASTRAN finite-element analysis the stiffness of the panel in the linear range prior to buckling was predicted within 3.5 percent. A nonlinear analysis predicted the buckling load within 10 percent and final failure load within 6 percent. The first test of a resin transfer molding six-stringer wing panel under compression was also completed. The panel failed unexpectedly in buckling because of inadequate supporting structure. The average strain was 0.43 percent with a line load of 20.3 kips per inch of width. This strain still exceeds the design allowable strains. Also, the stringers did not debond before failure, which is in contrast to the general behavior of unstitched panels.

Computation of spanwise distribution of circulation and lift coefficient for flapped wings of arbitrary planform

NASA Technical Reports Server (NTRS)

Razak, K.

1980-01-01

The question of the effect of distribution and magnitude of spanwise circulation and shed vorticity from an airplane wing on the distribution pattern of agricultural products distributed from an airplane was studied. The first step in an analysis of this question is the determination of the actual distribution of lift along an airplane wing, from which the pattern of shed vorticity can be determined. A procedure is developed to calculate the span loading for flapped and unflapped wings of arbitrary aspect ratio and taper ratio. The procedure was programmed on a small programmable calculator, the Hewlett Packard HP-97, and also was programmed in BASIC language. They could be used to explore the variations in span loading that can be secured by variable flap deflections or the effect of flying at varying air speeds at different airplane gross weights. Either an absolute evaluation of span loading can be secured or comparative span loading can be evaluated to determine their effect on swath width and swath distribution pattern. The programs are intended to assist the user in evaluating the effect of a given spanload distribution.

Aerodynamic Loads on an External Store Adjacent to a 45 Degree Sweptback Wing at Mach Numbers from 0.70 to 1.96, Including an Evaluation of Techniques Used

NASA Technical Reports Server (NTRS)

Guy, Lawrence D; Hadaway, William M

1955-01-01

Aerodynamic forces and moments have been obtained in the Langley 9- by 12-inch blowdown tunnel on an external store and on a 45 degree swept-back wing-body combination measured separately at Mach numbers from 0.70 to 1.96. The wing was cantilevered and had an aspect ratio of 4.0; the store was independently sting-mounted and had a Douglas Aircraft Co. (DAC) store shape. The angle of attack range was from -3 degrees to 12 degrees and the Reynolds number (based on wing mean aerodynamic chord) varied from 1.2 x10(6) to 1.7 x 10(6). Wing-body transonic forces and moments have been compared with data of a geometrically similar full-scale model tested in the Langley 16-foot and 8-foot transonic tunnels in order to aid in the evaluation of transonic-tunnel interference. The principal effect of the store, for the position tested, was that of delaying the wing-fuselage pitch-up tendency to higher angles of attack at Mach numbers from 0.70 to 0.90 in a manner similar to that of a wing chord extension. The most critical loading condition on the store was that due to side force, not only because the loads were of large magnitude but also because they were in the direction of least structural strength of the supporting pylon. These side loads were greatest at high angles of attack in the supersonic speed range. Removal of the supporting pylon (or increasing the gap between the store and wing) reduced the values of the variation of side-force coefficientwith angle of attack by about 50 percent at all test Mach numbers, indicating that important reductions in store side force may be realized by proper design or location of the necessary supporting pylon. A change of the store skew angle (nose inboard) was found to relieve the excessive store side loads throughout the Mach number range. It was also determined that the relative position of the fuselage nose to the store can appreciably affect the store side forces at supersonic speeds.

Technical Data Requirements for Shipboard and Shorebased Vertical/Short Takeoff and Landing (V/STOL) Aircraft. Revision A

DTIC Science & Technology

1978-04-26

Geometry 11-13 13-12 Shipboard Heavw Weather Tiedown 11-14 11-13 Nose & ’Main Gear Load Deflection Curves 11-15 11-14 Main Wheel Tire Span vs Aircraft...sustained taxi roll under conditions of 40-knot headwind and for wheel roll over 1-1/2 inch cable immediately after initial forward motion? 9. Planform...rolling/roll-oG vertical takeoff versus VTO. Discuss various methods of approach (e. g., stern, offset, cross axial). A Define minimum wheel -to-deck

Selected advanced aerodynamics and active controls technology concepts development on a derivative B-747

NASA Technical Reports Server (NTRS)

1980-01-01

The feasibility of applying wing tip extensions, winglets, and active control wing load alleviation to the Boeing 747 is investigated. Winglet aerodynamic design methods and high speed wind tunnel test results of winglets and of symmetrically deflected ailerons are presented. Structural resizing analyses to determine weight and aeroelastic twist increments for all the concepts and flutter model test results for the wing with winglets are included. Control law development, system mechanization/reliability studies, and aileron balance tab trade studies for active wing load alleviation systems are discussed. Results are presented in the form of incremental effects on L/D, structural weight, block fuel savings, stability and control, airplane price, and airline operating economics.

Qualitative comparison of calculated turbulence responses with wind-tunnel measurements for a DC-10 derivative wing with an active control system

NASA Technical Reports Server (NTRS)

Perry, B., III

1981-01-01

Comparisons are presented analytically predicted and experimental turbulence responses of a wind tunnel model of a DC-10 derivative wing equipped with an active control system. The active control system was designed for the purpose of flutter suppression, but it had additional benefit of alleviating gust loads (wing bending moment) by about 25%. Comparisions of various wing responses are presented for variations in active control system parameters and tunnel speed. The analytical turbulence responses were obtained using DYLOFLEX, a computer program for dynamic loads analyses of flexible airplanes with active controls. In general, the analytical predictions agreed reasonably well with the experimental data.

Hovering hummingbird wing aerodynamics during the annual cycle. I. Complete wing.

PubMed

Achache, Yonathan; Sapir, Nir; Elimelech, Yossef

2017-08-01

The diverse hummingbird family (Trochilidae) has unique adaptations for nectarivory, among which is the ability to sustain hover-feeding. As hummingbirds mainly feed while hovering, it is crucial to maintain this ability throughout the annual cycle-especially during flight-feather moult, in which wing area is reduced. To quantify the aerodynamic characteristics and flow mechanisms of a hummingbird wing throughout the annual cycle, time-accurate aerodynamic loads and flow field measurements were correlated over a dynamically scaled wing model of Anna's hummingbird ( Calypte anna ). We present measurements recorded over a model of a complete wing to evaluate the baseline aerodynamic characteristics and flow mechanisms. We found that the vorticity concentration that had developed from the wing's leading-edge differs from the attached vorticity structure that was typically found over insects' wings; firstly, it is more elongated along the wing chord, and secondly, it encounters high levels of fluctuations rather than a steady vortex. Lift characteristics resemble those of insects; however, a 20% increase in the lift-to-torque ratio was obtained for the hummingbird wing model. Time-accurate aerodynamic loads were also used to evaluate the time-evolution of the specific power required from the flight muscles, and the overall wingbeat power requirements nicely matched previous studies.

ARC-1969-A-31322

NASA Image and Video Library

1963-03-10

3/4 REAR VIEW OF Breguet 941 AIRPLANE; FLIGHT EVALUATION, MAY 1963. Boundary Layer Control, STOL, and V/STOL Research. Fig. 105 NASA SP Flight Research at Ames: 57 Years of Development and Validation of Aeronautical Technology

V/STOL Dynamics, Control, and Flying Qualities

NASA Technical Reports Server (NTRS)

Franklin, James A.

2000-01-01

This publication presents material that constituted the lectures presented by the author as part of Course AA 234, Dynamics, Control, and Flying Qualities of Vertical/Short Takeoff and Landing (V/STOL) Aircraft that was taught in the Department of Aeronautics and Astronautics at Stanford University. It covers representative operations of vertical and short takeoff and landing (V/STOL) aircraft, a discussion of the pilot's strategy in controlling these aircraft, the equations of motion pertinent to V/STOL tasks, and their application in the analysis of longitudinal and lateral-directional control in hover and forward flight. Following that development, which applies to the characteristics of the basic airframe and propulsion system, the text concludes with a discussion of the contributions of control augmentation in specific flight tasks and of the integration of modern electronic displays with these controls.

Computed lateral rate and acceleration power spectral response of conventional and STOL airplanes to atmospheric turbulence

NASA Technical Reports Server (NTRS)

Lichtenstein, J. H.

1975-01-01

Power-spectral-density calculations were made of the lateral responses to atmospheric turbulence for several conventional and short take-off and landing (STOL) airplanes. The turbulence was modeled as three orthogonal velocity components, which were uncorrelated, and each was represented with a one-dimensional power spectrum. Power spectral densities were computed for displacements, rates, and accelerations in roll, yaw, and sideslip. In addition, the power spectral density of the transverse acceleration was computed. Evaluation of ride quality based on a specific ride quality criterion was also made. The results show that the STOL airplanes generally had larger values for the rate and acceleration power spectra (and, consequently, larger corresponding root-mean-square values) than the conventional airplanes. The ride quality criterion gave poorer ratings to the STOL airplanes than to the conventional airplanes.

Design and fabrication of composite wing panels containing a production splice

NASA Technical Reports Server (NTRS)

Reed, D. L.

1975-01-01

Bolted specimens representative of both upper and lower wing surface splices of a transport aircraft were designed and manufactured for static and random load tension and compression fatigue testing including ground-air-ground load reversals. The specimens were fabricated with graphite-epoxy composite material. Multiple tests were conducted at various load levels and the results were used as input to a statistical wearout model. The statically designed specimens performed very well under highly magnified fatigue loadings. Two large panels, one tension and compression, were fabricated for testing by NASA-LRC.

Aerodynamics on a transport aircraft type wing-body model

NASA Technical Reports Server (NTRS)

Schmitt, V.

1982-01-01

The DFLR-F4 wing-body combination is studied. The 1/38 model is formed by a 9.5 aspect ratio transonic wing and an Airbus A 310 fuselage. The F4 wing geometrical characteristics are described and the main experimental results obtained in the S2MA wind tunnel are discussed. Both wing-fuselage interferences and viscous effects, which are important on the wing due to a high rear loading, are investigated by performing 3D calculations. An attempt is made to find their limitations.

PRSEUS Acoustic Panel Fabrication

NASA Technical Reports Server (NTRS)

Nicolette, Velicki; Yovanof, Nicolette P.; Baraja, Jaime; Mathur, Gopal; Thrash, Patrick; Pickell, Robert

2011-01-01

This report describes the development of a novel structural concept, Pultruded Rod Stitched Efficient Unitized Structure (PRSEUS), that addresses the demanding fuselage loading requirements for the Hybrid Wing or Blended Wing Body (BWB) airplane configuration with regards to acoustic response. A PRSEUS panel was designed and fabricated and provided to NASA-LaRC for acoustic response testing in the Structural Acoustics Loads and Transmission (SALT) facility). Preliminary assessments of the sound transmission characteristics of a PRSEUS panel subjected to a representative Hybrid Wing Body (HWB) operating environment were completed for the NASA Environmentally Responsible Aviation (ERA) Program.

Testing and Analysis of a Composite Non-Cylindrical Aircraft Fuselage Structure . Part II; Severe Damage

NASA Technical Reports Server (NTRS)

Przekop, Adam; Jegley, Dawn C.; Lovejoy, Andrew E.; Rouse, Marshall; Wu, Hsi-Yung T.

2016-01-01

The Environmentally Responsible Aviation Project aimed to develop aircraft technologies enabling significant fuel burn and community noise reductions. Small incremental changes to the conventional metallic alloy-based 'tube and wing' configuration were not sufficient to achieve the desired metrics. One airframe concept identified by the project as having the potential to dramatically improve aircraft performance was a composite-based hybrid wing body configuration. Such a concept, however, presented inherent challenges stemming from, among other factors, the necessity to transfer wing loads through the entire center fuselage section which accommodates a pressurized cabin confined by flat or nearly flat panels. This paper discusses a finite element analysis and the testing of a large-scale hybrid wing body center section structure developed and constructed to demonstrate that the Pultruded Rod Stitched Efficient Unitized Structure concept can meet these challenging demands of the next generation airframes. Part II of the paper considers the final test to failure of the test article in the presence of an intentionally inflicted severe discrete source damage under the wing up-bending loading condition. Finite element analysis results are compared with measurements acquired during the test and demonstrate that the hybrid wing body test article was able to redistribute and support the required design loads in a severely damaged condition.

An experimental study of pilots' control characteristics for flight of an STOL aircraft in backside of drag curve at approach and landing.

PubMed

Ema, T

1992-01-01

In general, most vehicles can be modelled by a multi-variable system which has interactive variables. It can be clearly shown that there is an interactive response in an aircraft's velocity and altitude obtained by stick control and/or throttle control. In particular, if the flight conditions fall to backside of drag curve in the flight of an STOL aircraft at approach and landing then the ratio of drag variation to velocity change has a negative value (delta D/delta u less than 0) and the system of motion presents a non-minimum phase. Therefore, the interaction between velocity and altitude response becomes so complicated that it affects to pilot's control actions and it may be difficult to control the STOL aircraft at approach and landing. In this paper, experimental results of a pilot's ability to control the STOL aircraft are presented for a multi-variable manual control system using a fixed ground base simulator and the pilot's control ability is discussed for the flight of an STOL aircraft at backside of drag curve at approach and landing.

Proceedings of the 1985 NASA Ames Research Center's Ground-Effects Workshop

NASA Technical Reports Server (NTRS)

Mitchell, Kerry (Editor)

1987-01-01

The purpose of the workshop was to discuss the current technology base for aerodynamic ground effects and to establish directions for further research of advanced, high performance aircraft designs, particularly those concepts utilizing powered lift systems; e.g., V/STOL, ASTOVL, and STOL aircraft. Fourteen papers were presented in the following areas: suckdown and fountain effects in hover; STOL ground vortex and hot gas ingestion; and vortex lift and jet flaps in ground effect. These subject areas were chosen with regard to current activities in the field of aircraft ground effects research.

Design of a V/STOL propulsion system for a large-scale fighter model

NASA Technical Reports Server (NTRS)

Willis, W. S.

1981-01-01

Modifications were made to the existing Large-Scale STOL fighter model to simulate a V/STOL configuration. Modifications include the substitutions of two dimensional lift/cruise exhaust nozzles in the nacelles, and the addition of a third J97 engine in the fuselage to suppy a remote exhaust nozzle simulating a Remote Augmented Lift System. A preliminary design of the inlet and exhaust ducting for the third engine was developed and a detailed design was completed of the hot exhaust ducting and remote nozzle.

Conceptual design studies of candidate V/STOL lift fan commercial short haul transport for 1980 - 1985 V/STOL lift fan study

NASA Technical Reports Server (NTRS)

Eldridge, W. M.; Ferrell, J. A.; Mckee, J. W.; Wayne, J. E., Jr.; Zabinsky, J. M.

1973-01-01

Conceptual designs of V/STOL lift fan commercial short haul transport aircraft for the 1980-85 time period were studied to determine their technical and economic feasibility. The engine concepts included both integral and remote fans. The scope of the study included definition of the hover control concept for each propulsion system, aircraft design, aircraft mass properties, cruise performance, noise and ride qualities evaluation. Economic evaluating was also studied on a basis of direct operating costs and route structure.

Development of the Main Wing Structure of a High Altitude Long Endurance UAV

NASA Astrophysics Data System (ADS)

Park, Sang Wook; Shin, Jeong Woo; Kim, Tae-Uk

2018-04-01

To enhance the flight endurance of a HALE UAV, the main wing of the UAV should have a high aspect ratio and low structural weight. Since a main wing constructed with the thin walled and slender components needed for low structural weight can suffer catastrophic failure during flight, it is important to develop a light-weight airframe without sacrificing structural integrity. In this paper, the design of the main wing of the HALE UAV was conducted using spars which were composed of a carbon-epoxy cylindrical tube and bulkheads to achieve both the weight reduction and structural integrity. The spars were sized using numerical analysis considering non-linear deformation under bending moment. Static strength testing of the wing was conducted under the most critical load condition. Then, the experimental results obtained for the wing were compared to the analytical result from the non-linear finite-element analysis. It was found that the developed main wing reduced its structural weight without any failure under the ultimate load condition of the static strength testing.

In-flight investigation of shuttle tile pressure orifice installations

NASA Technical Reports Server (NTRS)

Moes, Timothy R.; Meyer, Robert R., Jr.

1990-01-01

To determine shuttle orbiter wing loads during ascent, wing load instrumentation was added to Columbia (OV-102). This instrumentation included strain gages and pressure orifices on the wing. The loads derived from wing pressure measurements taken during STS 61-C did not agree with those derived from strain gage measurements or with the loads predicted from the aerodynamic database. Anomalies in the surface immediately surrounding the pressure orifices in the thermal protection system (TPS) tiles were one possible cause of errors in the loads derived from wing pressure measurements. These surface anomalies were caused by a ceramic filler material which was installed around the pressure tubing. The filler material allowed slight movement of the TPS tile and pressure tube as the airframe flexed and bent under aerodynamic loads during ascent and descent. Postflight inspection revealed that this filler material had protruded from or receeded beneath the surface, causing the orifice to lose its flushness. Flight tests were conducted at NASA Ames Research Center Dryden Flight Research Facility to determine the effects of any anomaly in surface flushness of the orifice installation on the measured pressures at Mach numbers between 0.6 and 1.4. An F-104 aircraft with a flight test fixture mounted beneath the fuselage was used for these flights. Surface flushness anomalies typical of those on the orbiter after flight (STA 61-C) were tested. Also, cases with excessive protrusion and recession of the filler material were tested. This report shows that the anomalies in STS 61-C orifice installations adversely affected the pressure measurements. But the magnitude of the affect was not great enough to account for the discrepancies with the strain gage measurements and the aerodynamic predictions.

Analysis of a Hybrid Wing Body Center Section Test Article

NASA Technical Reports Server (NTRS)

Wu, Hsi-Yung T.; Shaw, Peter; Przekop, Adam

2013-01-01

The hybrid wing body center section test article is an all-composite structure made of crown, floor, keel, bulkhead, and rib panels utilizing the Pultruded Rod Stitched Efficient Unitized Structure (PRSEUS) design concept. The primary goal of this test article is to prove that PRSEUS components are capable of carrying combined loads that are representative of a hybrid wing body pressure cabin design regime. This paper summarizes the analytical approach, analysis results, and failure predictions of the test article. A global finite element model of composite panels, metallic fittings, mechanical fasteners, and the Combined Loads Test System (COLTS) test fixture was used to conduct linear structural strength and stability analyses to validate the specimen under the most critical combination of bending and pressure loading conditions found in the hybrid wing body pressure cabin. Local detail analyses were also performed at locations with high stress concentrations, at Tee-cap noodle interfaces with surrounding laminates, and at fastener locations with high bearing/bypass loads. Failure predictions for different composite and metallic failure modes were made, and nonlinear analyses were also performed to study the structural response of the test article under combined bending and pressure loading. This large-scale specimen test will be conducted at the COLTS facility at the NASA Langley Research Center.

Ground noise measurements during landing, take-off, and flyby operations of a four-engine turbopropeller STOL airplane

NASA Technical Reports Server (NTRS)

Hilton, D. A.; Henderson, H. R.; Maglieri, D. J.

1971-01-01

Noise measurements were obtained for a four-engine turbopropeller STOL airplane during a Federal Aviation Administration flight evaluation program at the National Aviation Facilities Experimental Center. These noise measurements involved landing-approach, takeoff-climbout, and flyby operations of the airplane. A total of 13 measuring positions were used to define the noise characteristics around a simulated STOL port. The results are presented in the form of both physical and subjective measurements. An appendix is included to present tabulated values of various subjective reaction units which may be significant for the planning and operation of STOL ports. The main source of noise produced by this vehicle was found to be the propeller, and noise levels decrease generally in accordance with the inverse-distance law for distances up to about 457 meters. For similar slant ranges, somewhat lower noise levels were experienced during flyby than during takeoff or landing.

Unsteady aerodynamic flow field analysis of the space shuttle configuration. Part 1: Orbiter aerodynamics

NASA Technical Reports Server (NTRS)

Ericsson, L. E.; Reding, J. P.

1976-01-01

An analysis of the steady and unsteady aerodynamics of the space shuttle orbiter has been performed. It is shown that slender wing theory can be modified to account for the effect of Mach number and leading edge roundness on both attached and separated flow loads. The orbiter unsteady aerodynamics can be computed by defining two equivalent slender wings, one for attached flow loads and another for the vortex-induced loads. It is found that the orbiter is in the transonic speed region subject to vortex-shock-boundary layer interactions that cause highly nonlinear or discontinuous load changes which can endanger the structural integrity of the orbiter wing and possibly cause snap roll problems. It is presently impossible to simulate these interactions in a wind tunnel test even in the static case. Thus, a well planned combined analytic and experimental approach is needed to solve the problem.

The Pressure Distribution over the Wings and Tail Surfaces of a PW-9 Pursuit Airplane in Flight

NASA Technical Reports Server (NTRS)

Rhode, Richard

1931-01-01

This report presents the results of an investigation to determine (1) the magnitude and distribution of aerodynamic loads over the wings and tail surfaces of a pursuit-type airplane in the maneuvers likely to impose critical loads on the various subassemblies of the airplane structure. (2) To study the phenomenon of center of pressure movement and normal force coefficient variation in accelerated flight, and (3) to measure the normal accelerations at the center of gravity, wing-tip, and tail, in order to determine the nature of the inertia forces acting simultaneously with the critical aerodynamic loads. The results obtained throw light on a number of important questions involving structural design. Some of the more interesting results are discussed in some detail, but in general the report is for the purpose of making this collection of airplane-load data obtained in flight available to those interested in airplane structures.

Hovering hummingbird wing aerodynamics during the annual cycle. I. Complete wing

PubMed Central

Sapir, Nir; Elimelech, Yossef

2017-01-01

The diverse hummingbird family (Trochilidae) has unique adaptations for nectarivory, among which is the ability to sustain hover-feeding. As hummingbirds mainly feed while hovering, it is crucial to maintain this ability throughout the annual cycle—especially during flight-feather moult, in which wing area is reduced. To quantify the aerodynamic characteristics and flow mechanisms of a hummingbird wing throughout the annual cycle, time-accurate aerodynamic loads and flow field measurements were correlated over a dynamically scaled wing model of Anna’s hummingbird (Calypte anna). We present measurements recorded over a model of a complete wing to evaluate the baseline aerodynamic characteristics and flow mechanisms. We found that the vorticity concentration that had developed from the wing’s leading-edge differs from the attached vorticity structure that was typically found over insects’ wings; firstly, it is more elongated along the wing chord, and secondly, it encounters high levels of fluctuations rather than a steady vortex. Lift characteristics resemble those of insects; however, a 20% increase in the lift-to-torque ratio was obtained for the hummingbird wing model. Time-accurate aerodynamic loads were also used to evaluate the time-evolution of the specific power required from the flight muscles, and the overall wingbeat power requirements nicely matched previous studies. PMID:28878971

Piezoelectric energy harvesting from multifunctional wing spars for UAVs: Part 1. Coupled modeling and preliminary analysis

NASA Astrophysics Data System (ADS)

Erturk, A.; Anton, S. R.; Inman, D. J.

2009-03-01

This paper discusses the basic design factors for modifying an original wing spar to a multifunctional load-bearing - energy harvester wing spar. A distributed-parameter electromechanical formulation is given for modeling of a multilayer piezoelectric power generator beam for different combinations of the electrical outputs of piezoceramic layers. In addition to the coupled vibration response and voltage response expressions for a multimorph, strength formulations are given in order to estimate the maximum load input that can be sustained by the cantilevered structure without failure for a given safety factor. Embedding piezoceramics into an original wing spar for power generation tends to reduce the maximum load that can be sustained without failure and increase the total mass due to the brittle nature and large mass densities of typical piezoelectric ceramics. Two case studies are presented for demonstration. The theoretical case study discusses modification of a rectangular wing spar to a 3-layer generator wing spar with a certain restriction on mass addition for fixed dimensions. Power generation and strength analyses are provided using the electromechanical model. The experimental case study considers a 9-layer generator beam with aluminum, piezoceramic, Kapton and epoxy layers and investigates its power generation and load-bearing performances experimentally and analytically. This structure constitutes the main body of the multifunctional self-charging structure concept proposed by the authors. The second part of this work (experiments and storage applications) employs this multi-layer generator along with the thin-film battery layers in order to charge the battery layers using the electrical outputs of the piezoceramic layers.

Experimental study on thrust and power of flapping-wing system based on rack-pinion mechanism.

PubMed

Nguyen, Tuan Anh; Vu Phan, Hoang; Au, Thi Kim Loan; Park, Hoon Cheol

2016-06-20

This experimental study investigates the effect of three parameters: wing aspect ratio (AR), wing offset, and flapping frequency, on thrust generation and power consumption of a flapping-wing system based on a rack-pinion mechanism. The new flapping-wing system is simple but robust, and is able to create a large flapping amplitude. The thrust measured by a load cell reveals that for a given power, the flapping-wing system using a higher wing AR produces larger thrust and higher flapping frequency at the wing offset of 0.15[Formula: see text] or 0.20[Formula: see text] ([Formula: see text] is the mean chord) than other wing offsets. Of the three parameters, the flapping frequency plays a more significant role on thrust generation than either the wing AR or the wing offset. Based on the measured thrusts, an empirical equation for thrust prediction is suggested, as a function of wing area, flapping frequency, flapping angle, and wing AR. The difference between the predicted and measured thrusts was less than 7%, which proved that the empirical equation for thrust prediction is reasonable. On average, the measured power consumption to flap the wings shows that 46.5% of the input power is spent to produce aerodynamic forces, 14.0% to overcome inertia force, 9.5% to drive the rack-pinion-based flapping mechanism, and 30.0% is wasted as the power loss of the installed motor. From the power analysis, it is found that the wing with an AR of 2.25 using a wing offset of 0.20[Formula: see text] showed the optimal power loading in the flapping-wing system. In addition, the flapping frequency of 25 Hz is recommended as the optimal frequency of the current flapping-wing system for high efficiency, which was 48.3%, using a wing with an AR of 2.25 and a wing offset of 0.20[Formula: see text] in the proposed design.

Structural Testing of a Stitched/Resin Film Infused Graphite-Epoxy Wing Box

NASA Technical Reports Server (NTRS)

Jegley, Dawn C.; Bush, Harold G.

2001-01-01

The results of a series of tests conducted at the NASA Langley Research Center to evaluate the behavior of an all-composite full-scale wing box are presented. The wing box is representative of a section of a 220-passenger commercial transport aircraft wing box and was designed and constructed by The Boeing Company as part of the NASA Advanced Subsonics Technology (AST) program. The semi-span wing was fabricated from a graphite-epoxy material system with cover panels and spars held together using Kevlar stitches through the thickness. No mechanical fasteners were used to hold the stiffeners to the skin of the cover panels. Tests were conducted with and without low-speed impact damage, discrete source damage and repairs. Up-bending, down-bending and brake roll loading conditions were applied. The structure with non-visible impact damage carried 97% of Design Ultimate Load prior to failure through a lower cover panel access hole.

Structural Response and Failure of a Full-Scale Stitched Graphite-Epoxy Wing

NASA Technical Reports Server (NTRS)

Jegley, Dawn C.; Lovejoy, Andrew E.; Bush, Harold G.

2001-01-01

Analytical and experimental results of the test for an all-composite full-scale wing box are presented. The wing box is representative of a section of a 220-passenger commercial transport aircraft wing box and was designed and constructed by The Boeing Company as part of the NASA Advanced Subsonics Technology (AST) program. The semi-span wing was fabricated from a graphite-epoxy material system with cover panels and spars held together using Kevlar stitches through the thickness. No mechanical fasteners were used to hold the stiffeners to the skin of the cover panels. Tests were conducted with and without low-speed impact damage, discrete source damage and repairs. Up-bending down-bending and brake roll loading conditions were applied. The structure with nonvisible impact damage carried 97% of Design Ultimate Load prior to failure through a lower cover panel access hole. Finite element and experimental results agree for the global response of the structure.

The spanwise distribution of lift for minimum induced drag of wings having a given lift and a given bending moment

NASA Technical Reports Server (NTRS)

Jones, R. T.

1950-01-01

The problem of the minimum induced drag of wings having a given lift and a given span is extended to include cases in which the bending moment to be supported by the wing is also given. The theory is limited to lifting surfaces traveling at subsonic speeds. It is found that the required shape of the downwash distribution can be obtained in an elementary way which is applicable to a variety of such problems. Expressions for the minimum drag and the corresponding spanwise load distributions are also given for the case in which the lift and the bending moment about the wing root are fixed while the span is allowed to vary. The results show a 15-percent reduction of the induced drag with a 15-percent increase in span as compared with results for an elliptically loaded wing having the same total lift and bending moment.

Probabilistic Analysis and Design of a Raked Wing Tip for a Commercial Transport

NASA Technical Reports Server (NTRS)

Mason Brian H.; Chen, Tzi-Kang; Padula, Sharon L.; Ransom, Jonathan B.; Stroud, W. Jefferson

2008-01-01

An approach for conducting reliability-based design and optimization (RBDO) of a Boeing 767 raked wing tip (RWT) is presented. The goal is to evaluate the benefits of RBDO for design of an aircraft substructure. A finite-element (FE) model that includes eight critical static load cases is used to evaluate the response of the wing tip. Thirteen design variables that describe the thickness of the composite skins and stiffeners are selected to minimize the weight of the wing tip. A strain-based margin of safety is used to evaluate the performance of the structure. The randomness in the load scale factor and in the strain limits is considered. Of the 13 variables, the wing-tip design was controlled primarily by the thickness of the thickest plies in the upper skins. The report includes an analysis of the optimization results and recommendations for future reliability-based studies.

Evaluation of the Structural Response and Failure of a Full-Scale Stitched Graphite-Epoxy Wing

NASA Astrophysics Data System (ADS)

Jegley, Dawn C.; Bush, Harold G.; Lovejoy, Andrew E.

2001-01-01

Analytical and experimental results for an all-composite full-scale wing box are presented. The wing box is representative of a section of a 220-passenger commercial transport aircraft wing box and was designed and constructed by The Boeing Company as part of the NASA Advanced Subsonics Technology (AST) program. The semi-span wing was fabricated from a graphite-epoxy material system with cover panels and spars held together using Kevlar stitches through the thickness. No mechanical fasteners were used to hold the stiffeners to the skin of the cover panels. Tests were conducted with and without low-speed impact damage, discrete source damage and repairs. Upbending, down-bending and brake roll loading conditions were applied. The structure with nonvisible impact damage carried 97% of Design Ultimate Load prior to failure through a lower cover panel access hole. Finite element and experimental results agree for the global response of the structure.

Theoretical parametric study of the relative advantages of winglets and wing-tip extensions

NASA Technical Reports Server (NTRS)

Heyson, H. H.; Riebe, G. D.; Fulton, C. L.

1977-01-01

It was found that for identical increases in bending moment, a winglet provides a greater gain in induced efficiency than a tip extension. Winglet toe-in angle allows design trades between efficiency and root moment. A winglet showed the greatest benefit when the wing loads were heavy near the tip. Washout diminished the benefit of either tip modification, and the gain in induced efficiency became a function of lift coefficient; heavy wing loadings obtained the greatest benefit from a winglet, and low speed performance was enhanced even more than cruise performance. Both induced efficiency and bending moment increased with winglet length and outward cant. The benefit of a winglet relative to a tip extension was greatest for a nearly vertical winglet. Root bending moment was proportional to the minimum weight of bending material required in the wing; it is a valid index of the impact of tip modifications on a new wing design.

Theoretical Parametric Study of the Relative Advantages of Winglets and Wing-Tip Extensions

NASA Technical Reports Server (NTRS)

Heyson, H. H.; Riebe, G. D.; Fulton, C. L.

1977-01-01

For identical increases in bending moment, a winglet provides a greater gain in induced efficiency than tip extension. Winglet toe angle allows design trades between efficiency and root moment. A winglet shows the greatest benefit when the wing loads are heavy near the tip. Washout diminishes the benefit of either tip modification, and the gain in induced efficiency becomes a function of lift coefficient; thus, heavy wing loadings obtain the greatest benefit from a winglet, and low-speed performance is enhanced even more than cruise performance. Both induced efficiency and bending moment increase with winglet length and outward cant. The benefit of a winglet relative to a tip extension is greatest for a nearly vertical winglet. Root bending moment is proportional to the minimum weight of bending material required in the wing; thus, it is a valid index of the impact of tip modifications on a new wing design.

Aerostructural Level Set Topology Optimization for a Common Research Model Wing

NASA Technical Reports Server (NTRS)

Dunning, Peter D.; Stanford, Bret K.; Kim, H. Alicia

2014-01-01

The purpose of this work is to use level set topology optimization to improve the design of a representative wing box structure for the NASA common research model. The objective is to minimize the total compliance of the structure under aerodynamic and body force loading, where the aerodynamic loading is coupled to the structural deformation. A taxi bump case was also considered, where only body force loads were applied. The trim condition that aerodynamic lift must balance the total weight of the aircraft is enforced by allowing the root angle of attack to change. The level set optimization method is implemented on an unstructured three-dimensional grid, so that the method can optimize a wing box with arbitrary geometry. Fast matching and upwind schemes are developed for an unstructured grid, which make the level set method robust and efficient. The adjoint method is used to obtain the coupled shape sensitivities required to perform aerostructural optimization of the wing box structure.

Speed limits of aircraft

NASA Technical Reports Server (NTRS)

Everling, E

1923-01-01

This paper is restricted to the question of attainable speed limits and attacks the problem from different angles. Theoretical limits due to air resistance are presented along with design factors which may affect speed such as wing loads, wing areas, wing section shifting, landing speeds, drag-lift ratios, and power coefficients.

American Airlines Propeller STOL Transport Economic Risk Analysis

NASA Technical Reports Server (NTRS)

Ransone, B.

1972-01-01

A Monte Carlo risk analysis on the economics of STOL transports in air passenger traffic established the probability of making the expected internal rate of financial return, or better, in a hypothetical regular Washington/New York intercity operation.

ACEE composite structures technology

NASA Technical Reports Server (NTRS)

Quinlivan, John T.; Wilson, Robert D.; Smith, Peter J.; Johnson, Ronald W.

1984-01-01

Toppics addressed include: advanced composites on Boeing commercial aircraft; composite wing durability; damage tolerance technology development; heavily loaded wing panel design; and pressure containment and damage tolerance in fuselages.

Development of a nonlinear vortex method. [steady and unsteady aerodynamic loads of highly sweptback wings

NASA Technical Reports Server (NTRS)

Kandil, O. A.

1981-01-01

Progress is reported in the development of reliable nonlinear vortex methods for predicting the steady and unsteady aerodynamic loads of highly sweptback wings at large angles of attack. Abstracts of the papers, talks, and theses produced through this research are included. The modified nonlinear discrete vortex method and the nonlinear hybrid vortex method are highlighted.

Relative loading on biplane wings

NASA Technical Reports Server (NTRS)

Diehl, Walter S

1934-01-01

Recent improvements in stress analysis methods have made it necessary to revise and to extend the loading curves to cover all conditions of flight. This report is concerned with a study of existing biplane data by combining the experimental and theoretical data to derive a series of curves from which the lift curves of the individual wings of a biplane may be obtained.

Biologically Inspired, Anisoptropic Flexible Wing for Optimal Flapping Flight

DTIC Science & Technology

2013-01-31

Anisotropic Flexible Wing for Optimal Flapping Flight FA9550-07-1-0547 Sb. GRANT NUMBER Sc. PROGRAM ELEMENT NUMBER 6. AUTHOR(S) Sd. PROJECT NUMBER...anisotropic structural flexibility ; c) Conducted coordinated experimental and computational modeling to determine the roles of aerodynamic loading, wing inertia...and structural flexibility and elasticity; and d) Developed surrogate tools for flapping wing MA V design and optimization. Detailed research

Computing Trimmed, Mean-Camber Surfaces At Minimum Drag

NASA Technical Reports Server (NTRS)

Lamar, John E.; Hodges, William T.

1995-01-01

VLMD computer program determines subsonic mean-camber surfaces of trimmed noncoplanar planforms with minimum vortex drag at specified lift coefficient. Up to two planforms designed together. Method used that of subsonic vortex lattice method of chord loading specification, ranging from rectangular to triangular, left specified by user. Program versatile and applied to isolated wings, wing/canard configurations, tandem wing, and wing/-winglet configuration. Written in FORTRAN.

Validation of Vortex-Lattice Method for Loads on Wings in Lift-Generated Wakes

NASA Technical Reports Server (NTRS)

Rossow, Vernon J.

1995-01-01

A study is described that evaluates the accuracy of vortex-lattice methods when they are used to compute the loads induced on aircraft as they encounter lift-generated wakes. The evaluation is accomplished by the use of measurements made in the 80 by 120 ft Wind Tunnel of the lift, rolling moment, and downwash in the wake of three configurations of a model of a subsonic transport aircraft. The downwash measurements are used as input for a vortex-lattice code in order to compute the lift and rolling moment induced on wings that have a span of 0.186, 0.510, or 1.022 times the span of the wake-generating model. Comparison of the computed results with the measured lift and rolling-moment distributions the vortex-lattice method is very reliable as long as the span of the encountering or following wing is less than about 0.2 of the generator span. As the span of the following wing increases above 0.2, the vortex-lattice method continues to correctly predict the trends and nature of the induced loads, but it overpredicts the magnitude of the loads by increasing amounts.

Aeroelastic loads prediction for an arrow wing. Task 1: Evaluation of R. P. White's method

NASA Technical Reports Server (NTRS)

Borland, C. J.; Manro, M. E.

1983-01-01

The separated flow method is evaluated. This method was developed for moderately swept wings with multiple, constant strength vortex systems. The flow on the highly swept wing used in this evaluation is characterized by a single vortex system of continuously varying strength.

Path changing methods applied to the 4-D guidance of STOL aircraft.

DOT National Transportation Integrated Search

1971-11-01

Prior to the advent of large-scale commercial STOL service, some challenging navigation and guidance problems must be solved. Proposed terminal area operations may require that these aircraft be capable of accurately flying complex flight paths, and ...

Flight Path Synthesis and HUD Scaling for V/STOL Terminal Area Operations

DOT National Transportation Integrated Search

1995-04-01

A two circle horizontal flightpath synthesis algorithm for Vertical/Short : Takeoff and Landing (V/STOL) terminal area operations is presented. This : algorithm provides a flight-path that is tangential to the aircraft's velocity : vector at the inst...

Preliminary survey of potential STOL terminal area operational requirements

DOT National Transportation Integrated Search

1971-06-01

A preliminary survey of potential operational requirements for STOL in the terminal area was made. The presentation of the survey is in three sections. The first section presents the motivation for the survey. The second section discusses the markets...

Propulsion system mathematical model for a lift/cruise fan V/STOL aircraft

NASA Technical Reports Server (NTRS)

Cole, G. L.; Sellers, J. F.; Tinling, B. E.

1980-01-01

A propulsion system mathematical model is documented that allows calculation of internal engine parameters during transient operation. A non-realtime digital computer simulation of the model is presented. It is used to investigate thrust response and modulation requirements as well as the impact of duty cycle on engine life and design criteria. Comparison of simulation results with steady-state cycle deck calculations showed good agreement. The model was developed for a specific 3-fan subsonic V/STOL aircraft application, but it can be adapted for use with any similar lift/cruise V/STOL configuration.

Study of quiet turbofan STOL aircraft for short haul transportation

NASA Technical Reports Server (NTRS)

Higgins, T. P.; Stout, E. G.; Sweet, H. S.

1973-01-01

Conceptual designs of Quiet Turbofan STOL Short-Haul Transport Aircraft for the mid-1980 time period are developed and analyzed to determine their technical, operational, and economic feasibility. A matrix of aircraft using various high-lift systems and design parameters are considered. Variations in aircraft characteristics, airport geometry and location, and operational techniques are analyzed systematically to determine their effects on the market, operating economics, and community acceptance. In these studies, the total systems approach is considered to be critically important in analyzing the potential of STOL aircraft to reduce noise pollution and alleviate the increasing air corridor and airport congestion.

Study of quiet turbofan STOL aircraft for short-haul transportation. Volume 1: Summary

NASA Technical Reports Server (NTRS)

1974-01-01

Conceptual designs of Quiet Turbofan STOL Short-Haul Transport Aircraft for the mid-1980 time period are developed and analyzed to determine their technical, operational, and economic feasibility. A matrix of aircraft using various high-lift systems and design parameters are considered. Variations in aircraft characteristics, airport geometry and location, and operational techniques are analyzed systematically to determine their effects on the market, operating economics, and community acceptance. The total systems approach is considered to be critically important in analyzing the potential of STOL aircraft to reduce noise pollution and alleviate the increasing air corridor and airport congestion.

An experimental study of several wind tunnel wall configurations using two V/STOL model configurations. [low speed wind tunnels

NASA Technical Reports Server (NTRS)

Binion, T. W., Jr.

1975-01-01

Experiments were conducted in the low speed wind tunnel using two V/STOL models, a jet-flap and a jet-in-fuselage configuration, to search for a wind tunnel wall configuration to minimize wall interference on V/STOL models. Data were also obtained on the jet-flap model with a uniform slotted wall configuration to provide comparisons between theoretical and experimental wall interference. A test section configuration was found which provided some data in reasonable agreement with interference-free results over a wide range of momentum coefficients.

Aircraft Configuration Study for Experimental 2-Place Aircraft and RPVs

DTIC Science & Technology

1990-03-01

area (sq. ft.) 84.24 82.86 Wing airfoil section Eppler Wing aspect ratio 8.09 Wing loading (GW) (lb./sq. ft.: 7.30 7.24 Canard span (ft.) 11.70 11.60...ESTIMATION FOR THE CANARD DRAG POLAR BUILDUP Aircraft Canard FG Input italicized data Wing for Eppler airfoil Cdmin = .0080 S = 82.9 Canard from Eppler ...DRAG POLAR BUILDUP Aircraft Canard FG Input italicized data Wing for Eppler airfoil Cdmin = .0080 S = 82.9 Canard from Eppler for GA(A)-1 airfoil Cdmin

Study of V/STOL aircraft implementation. Volume 2: Appendices

NASA Technical Reports Server (NTRS)

Androsky, A.; Miller, S. C.; Neiss, J. A.; Portenier, W. J.; Webb, H. M.

1972-01-01

An analysis of V/STOL aircraft implementation and utilization is presented. The subjects discussed are: (1) short haul air transportation requirements, (2) available aircraft technology, (3) aircraft production requirements, (4) airport requirements, (5) roles and responsibilities, and (6) cost and funding.

Real-time simulation program for De Havilland (Canada) "Buffalo" and "Twin Otter" STOL transports

DOT National Transportation Integrated Search

1971-06-25

Simulation models of two representative STOL aircraft - the DeHavilland (Canada) "Buffalo" and "Twin Otter" transports - have been generated. The aircraft are described by means of nonlinear equations that will accommodate gross changes in angle of a...

ARC-1969-A-20485

NASA Image and Video Library

1955-06-22

Grumman F9F-6 (Bu. No. 128138) Cougar airplane. EVALUATION OF CARRIER APPROACH TECHNIQUES Boundary Layer Control, STOL, and V/STOL Note: Used in publication in Flight Research at Ames; 57 Years of Development and Validation of Aeronautical Technology NASA SP-1998-3300 fig. 101

Linearized mathematical models for De Havilland Canada "Buffalo & Twin Otter" STOL transports.

DOT National Transportation Integrated Search

1971-06-01

Linearized six degree of freedom rigid body aircraft equations of motion are presented in a stability axes system. Values of stability derivatives are estimated for two representative STOL aircraft - the DeHavilland of Canada 'Buffalo' and 'Twin Otte...

Understanding the unsteady aerodynamics of a revolving wing with pitching-flapping perturbations

NASA Astrophysics Data System (ADS)

Chen, Long; Wu, Jianghao; Zhou, Chao; Hsu, Shih-Jung; Eslam Panah, Azar; Cheng, Bo

2017-11-01

Revolving wings become less efficient for lift generation at low Reynolds numbers. Unlike flying insects using reciprocating revolving wings to exploit unsteady mechanisms for lift enhancement, an alternative that introduces unsteadiness through vertical flapping perturbation, is studied via experiments and simulations. Substantial drag reduction, linearly dependent on Strouhal number, is observed for a flapping-perturbed revolving wing at zero angle of attack (AoA), which can be explained by changes in the effective angle of attack and formation of reverse Karman vortex streets. When the AoA increases, flapping perturbations improve the maximum lift coefficient attainable by the revolving wing, with minor increases of drag or even minor drag reductions depending on Strouhal number and normalized flapping amplitude. When the pitching perturbations are further introduced, more substantial drag reduction and lift enhancement can be achieved in zero and positive AoAs, respectively. As the flapping-perturbed wings are less efficient compared with revolving wings in terms of power loading, the pitching-flapping perturbations can achieve a higher power loading at 20°AoA and thus have potential applications in micro air vehicle designs. This research was supported by NSF, DURIP, NSFC and Penn State Multi-Campus SEED Grant.

Projection Moire Interferometry Measurements of Micro Air Vehicle Wings

NASA Technical Reports Server (NTRS)

Fleming, Gary A.; Bartram, Scott M.; Waszak, Martin R.; Jenkins, Luther N.

2001-01-01

Projection Moire Interferometry (PMI) has been used to measure the structural deformation of micro air vehicle (MAV) wings during a series of wind tunnel tests. The MAV wings had a highly flexible wing structure, generically reminiscent of a bat s wing, which resulted in significant changes in wing shape as a function of MAV angle-of-attack and simulated flight speed. This flow-adaptable wing deformation is thought to provide enhanced vehicle stability and wind gust alleviation compared to rigid wing designs. Investigation of the potential aerodynamic benefits of a flexible MAV wing required measurement of the wing shape under aerodynamic loads. PMI was used to quantify the aerodynamically induced changes in wing shape for three MAV wings having different structural designs and stiffness characteristics. This paper describes the PMI technique, its application to MAV testing, and presents a portion of the PMI data acquired for the three different MAV wings tested.

Aerodynamic characteristics of a wing with Fowler flaps including flap loads, downwash, and calculated effect on take-off

NASA Technical Reports Server (NTRS)

Platt, Robert C

1936-01-01

This report presents the results of wind tunnel tests of a wing in combination with each of three sizes of Fowler flap. The purpose of the investigation was to determine the aerodynamic characteristics as affected by flap chord and position, the air loads on the flaps, and the effect of flaps on the downwash.

14 CFR Appendix A to Part 23 - Simplified Design Load Criteria

Code of Federal Regulations, 2011 CFR

2011-01-01

... quarter-chord), delta planforms, or slatted lifting surfaces; or (5) Winglets or other wing tip devices... single engine excluding turbine powerplants; (2) A main wing located closer to the airplane's center of gravity than to the aft, fuselage-mounted, empennage; (3) A main wing that contains a quarter-chord sweep...

14 CFR Appendix A to Part 23 - Simplified Design Load Criteria

Code of Federal Regulations, 2010 CFR

2010-01-01

... quarter-chord), delta planforms, or slatted lifting surfaces; or (5) Winglets or other wing tip devices... single engine excluding turbine powerplants; (2) A main wing located closer to the airplane's center of gravity than to the aft, fuselage-mounted, empennage; (3) A main wing that contains a quarter-chord sweep...

Overview: Performance Adaptive Aeroelastic Wing

NASA Technical Reports Server (NTRS)

Hashemi, Kelley

2017-01-01

An overview of recent aeroelasitc wing-shaping work at the NASA Ames Research Center is presented. The highlight focuses on activity related to the Performance Adaptive Aeroelastic Wing concept and related Variable Camber Continuous Trailing Edge Flap actuation system. Topics covered include drag-reducing configurations and online algorithms, gust and maneuver load techniques, and wind tunnel demonstrations.

Chordwise load distribution of a simple rectangular wing

NASA Technical Reports Server (NTRS)

Wieghardt, Karl

1940-01-01

The chordwise distribution theory was taken over from the theory of the infinite wing. Since in this work a series expansion in b/t was used, the computation converges only for large aspect ratios. In this paper a useful approximate solution will be found also for wings with large chord - i.e., small aspect ratio.

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility (OPF), a United Space Alliance technician examines the attachment points for the spars on the exterior of a wing of Space Shuttle Atlantis. Reinforced Carbon Carbon (RCC) panels are mechanically attached to the wing with a series of floating joints - spars - to reduce loading on the panels caused by wing deflections. The aluminum and the metallic attachments are protected from exceeding temperature limits by internal insulation. The next launch of Atlantis will be on mission STS-114, a utilization and logistics flight to the International Space Station.

NASA Image and Video Library

2003-09-03

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility (OPF), a United Space Alliance technician examines the attachment points for the spars on the exterior of a wing of Space Shuttle Atlantis. Reinforced Carbon Carbon (RCC) panels are mechanically attached to the wing with a series of floating joints - spars - to reduce loading on the panels caused by wing deflections. The aluminum and the metallic attachments are protected from exceeding temperature limits by internal insulation. The next launch of Atlantis will be on mission STS-114, a utilization and logistics flight to the International Space Station.

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility (OPF), United Space Alliance technicians replace the attachment points for the spars on the interior of a wing of Space Shuttle Atlantis. Reinforced Carbon Carbon (RCC) panels are mechanically attached to the wing with a series of floating joints - spars - to reduce loading on the panels caused by wing deflections. The aluminum and the metallic attachments are protected from exceeding temperature limits by internal insulation. The next launch of Atlantis will be on mission STS-114, a utilization and logistics flight to the International Space Station.

NASA Image and Video Library

2003-09-03

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility (OPF), United Space Alliance technicians replace the attachment points for the spars on the interior of a wing of Space Shuttle Atlantis. Reinforced Carbon Carbon (RCC) panels are mechanically attached to the wing with a series of floating joints - spars - to reduce loading on the panels caused by wing deflections. The aluminum and the metallic attachments are protected from exceeding temperature limits by internal insulation. The next launch of Atlantis will be on mission STS-114, a utilization and logistics flight to the International Space Station.

KENNEDY SPACE CENTER, FLA. -In the Orbiter Processing Facility (OPF), a United Space Alliance technician examines the attachment points for the spars on the exterior of a wing of Space Shuttle Atlantis. Reinforced Carbon Carbon (RCC) panels are mechanically attached to the wing with a series of floating joints - spars - to reduce loading on the panels caused by wing deflections. The aluminum and the metallic attachments are protected from exceeding temperature limits by internal insulation. The next launch of Atlantis will be on mission STS-114, a utilization and logistics flight to the International Space Station.

NASA Image and Video Library

2003-09-03

KENNEDY SPACE CENTER, FLA. -In the Orbiter Processing Facility (OPF), a United Space Alliance technician examines the attachment points for the spars on the exterior of a wing of Space Shuttle Atlantis. Reinforced Carbon Carbon (RCC) panels are mechanically attached to the wing with a series of floating joints - spars - to reduce loading on the panels caused by wing deflections. The aluminum and the metallic attachments are protected from exceeding temperature limits by internal insulation. The next launch of Atlantis will be on mission STS-114, a utilization and logistics flight to the International Space Station.

Challenges, Ideas, and Innovations of Joined-Wing Configurations: A Concept from the Past, an Opportunity for the Future

NASA Astrophysics Data System (ADS)

Cavallaro, Rauno; Demasi, Luciano

2016-11-01

Diamond Wings, Strut- and Truss-Braced Wings, Box Wings, and PrandtlPlane, the so-called "JoinedWings", represent a dramatic departure from traditional configurations. Joined Wings are characterized by a structurally overconstrained layout which significantly increases the design space with multiple load paths and numerous solutions not available in classical wing systems. A tight link between the different disciplines (aerodynamics, flight mechanics, aeroelasticity, etc.) makes a Multidisciplinary Design and Optimization approach a necessity from the early design stages. Researchers showed potential in terms of aerodynamic efficiency, reduction of emissions and superior performances, strongly supporting the technical advantages of Joined Wings. This review will present these studies, with particular focus on the United States joined-wing SensorCraft, Strut- and Truss- Braced Wings, Box Wings and PrandtlPlane.

Steady-State Solution of a Flexible Wing

NASA Technical Reports Server (NTRS)

Karkehabadi, Reza; Chandra, Suresh; Krishnamurthy, Ramesh

1997-01-01

A fluid-structure interaction code, ENSAERO, has been used to compute the aerodynamic loads on a swept-tapered wing. The code has the capability of using Euler or Navier-Stokes equations. Both options have been used and compared in the present paper. In the calculation of the steady-state solution, we are interested in knowing how the flexibility of the wing influences the lift coefficients. If the results of a flexible wing are not affected by the flexibility of the wing significantly, one could consider the wing to be rigid and reduce the problem from fluid-structure interaction to a fluid problem.

Lessons learned in the development of the STOL intelligent tutoring system

NASA Technical Reports Server (NTRS)

Seamster, Thomas; Baker, Clifford; Ames, Troy

1991-01-01

Lessons learned during the development of the NASA Systems Test and Operations Language (STOL) Intelligent Tutoring System (ITS), being developed at NASA Goddard Space Flight Center are presented. The purpose of the intelligent tutor is to train STOL users by adapting tutoring based on inferred student strengths and weaknesses. This system has been under development for over one year and numerous lessons learned have emerged. These observations are presented in three sections, as follows. The first section addresses the methodology employed in the development of the STOL ITS and briefly presents the ITS architecture. The second presents lessons learned, in the areas of: intelligent tutor development; documentation and reporting; cost and schedule control; and tools and shells effectiveness. The third section presents recommendations which may be considered by other ITS developers, addressing: access, use and selection of subject matter experts; steps involved in ITS development; use of ITS interface design prototypes as part of knowledge engineering; and tools and shells effectiveness.

Full-Field Reconstruction of Structural Deformations and Loads from Measured Strain Data on a Wing Using the Inverse Finite Element Method

NASA Technical Reports Server (NTRS)

Miller, Eric J.; Manalo, Russel; Tessler, Alexander

2016-01-01

A study was undertaken to investigate the measurement of wing deformation and internal loads using measured strain data. Future aerospace vehicle research depends on the ability to accurately measure the deformation and internal loads during ground testing and in flight. The approach uses the inverse Finite Element Method (iFEM). The iFEM is a robust, computationally efficient method that is well suited for real-time measurement of real-time structural deformation and loads. The method has been validated in previous work, but has yet to be applied to a large-scale test article. This work is in preparation for an upcoming loads test of a half-span test wing in the Flight Loads Laboratory at the National Aeronautics and Space Administration Armstrong Flight Research Center (Edwards, California). The method has been implemented into an efficient MATLAB® (The MathWorks, Inc., Natick, Massachusetts) code for testing different sensor configurations. This report discusses formulation and implementation along with the preliminary results from a representative aerospace structure. The end goal is to investigate the modeling and sensor placement approach so that the best practices can be applied to future aerospace projects.

An in-flight simulator investigation of roll and yaw control power requirements for STOL approach and landing: Development of capability and preliminary results

NASA Technical Reports Server (NTRS)

Ellis, D. R.; Raisinghani, S. C.

1979-01-01

A six-degree-of-freedom variable-response research aircraft was used to determine the minimum lateral-directional control power required for desirable and acceptable levels of handling qualities for the STOL landing approach task in a variety of simulated atmospheric disturbance conditions for a range of lateral-directional response characteristics. Topics covered include the in-flight simulator, crosswind simulation, turbulence simulation, test configurations, and evaluation procedures. Conclusions based on a limited sampling of simulated STOL transport configurations flown to touchdown out of 6 deg, 75 kt MLS approaches, usually with a sidestep maneuver are discussed.

GSFC Systems Test and Operation Language (STOL) functional requirements and language description

NASA Technical Reports Server (NTRS)

Desjardins, R.; Hall, G.; Mcguire, J.; Merwarth, P.; Mocarsky, W.; Truszkowski, W.; Villasenor, A.; Brosi, F.; Burch, P.; Carey, D.

1978-01-01

The Systems Tests and Operation Language (STOL) provides the means for user communication with payloads, applications programs, and other ground system elements. It is a systems operation language that enables an operator or user to communicate a command to a computer system. The system interprets each high level language directive from the user and performs the indicated action, such as executing a program, printing out a snapshot, or sending a payload command. This document presents the following: (1) required language features and implementation considerations; (2) basic capabilities; (3) telemetry, command, and input/output directives; (4) procedure definition and control; (5) listing, extension, and STOL nucleus capabilities.

Conceptual design study of improved 1985 remote lift-fan V/STOL commercial transports

NASA Technical Reports Server (NTRS)

Cavage, R. L.

1975-01-01

A design study was conducted for a remote lift-fan commercial V/STOL transport for the 1985 time period. The investigation centered on the commercial short haul transportation application to carry 100 passengers over trip distances of 400 nautical miles from a vertical takeoff and landing, and 800 nautical miles after a 1600 foot STOL takeoff. The study included investigation of alternate numbers and arrangements of lift fans and gas generators, fan control margins, and structural concepts. The sensitivity of direct operating costs to major airframe parameters, airframe costs, propulsion costs, yearly aircraft utilization rate, and trip distances are evaluated.

The Influence of Tip Shape on the Wing Load Distribution as Determined by Flight Tests

NASA Technical Reports Server (NTRS)

Rhode, Richard V

1935-01-01

Pressure measurements were made in flight on the right upper wing of an M-3 airplane. The effects of tip plan form, washout, and transverse camber were investigated with eight tip forms in unyawed conditions through the range of positive lift coefficients from zero lift to the stall. The conclusion is that the tip plan form does not influence the span distribution of the coefficients of normal force and moment. It is shown inferentially that temperature, humidity, and the aging of the wood and fabric wing structure used on the M-3 airplane have an appreciable influence on the load distribution.

Continued development and correlation of analytically based weight estimation codes for wings and fuselages

NASA Technical Reports Server (NTRS)

Mullen, J., Jr.

1978-01-01

The implementation of the changes to the program for Wing Aeroelastic Design and the development of a program to estimate aircraft fuselage weights are described. The equations to implement the modified planform description, the stiffened panel skin representation, the trim loads calculation, and the flutter constraint approximation are presented. A comparison of the wing model with the actual F-5A weight material distributions and loads is given. The equations and program techniques used for the estimation of aircraft fuselage weights are described. These equations were incorporated as a computer code. The weight predictions of this program are compared with data from the C-141.

A computer program to calculate the longitudinal aerodynamic characteristics of wing-flap configurations with externally blown flaps

NASA Technical Reports Server (NTRS)

Mendenhall, M. R.; Goodwin, F. K.; Spangler, S. B.

1976-01-01

A vortex lattice lifting-surface method is used to model the wing and multiple flaps. Each lifting surface may be of arbitrary planform having camber and twist, and the multiple-slotted trailing-edge flap system may consist of up to ten flaps with different spans and deflection angles. The engine wakes model consists of a series of closely spaced vortex rings with circular or elliptic cross sections. The rings are normal to a wake centerline which is free to move vertically and laterally to accommodate the local flow field beneath the wing and flaps. The two potential flow models are used in an iterative fashion to calculate the wing-flap loading distribution including the influence of the waves from up to two turbofan engines on the semispan. The method is limited to the condition where the flow and geometry of the configurations are symmetric about the vertical plane containing the wing root chord. The calculation procedure starts with arbitrarily positioned wake centerlines and the iterative calculation continues until the total configuration loading converges within a prescribed tolerance. Program results include total configuration forces and moments, individual lifting-surface load distributions, including pressure distributions, individual flap hinge moments, and flow field calculation at arbitrary field points.

Status and future plans of the Drones for Aerodynamic and Structural Testing (DAST) program. [Aeroelastic Research Wing (ARW)

NASA Technical Reports Server (NTRS)

Murrow, H. N.

1981-01-01

Results from flight tests of the ARW-1 research wing are presented. Preliminary loads data and experiences with the active control system for flutter suppression are included along with comparative results of test and prediction for the flutter boundary of the supercritical research wing and on performance of the flutter suppression system. The status of the ARW-2 research wing is given.

Prediction of V/STOL Noise for Application to Community Noise Exposure

DOT National Transportation Integrated Search

1973-05-01

A computer program to predict the Effective Perceived Noise Level (EPNL), the tone corrected Perceived Noise Level (PNLT) and the A-Weighted Sound Level (dBA) radiated by a V/STOL vehicle as it flies along a prescribed takeoff, landing, or cruise fli...

The 1987 Ground Vortex Workshop

NASA Technical Reports Server (NTRS)

Margason, Richard J. (Editor)

1988-01-01

The purpose of this workshop was to discuss the current understanding of the ground vortex phenomena and their effects on aircraft, and to establish directions for further research on advanced, high-performance aircraft designs, particularly those concepts utilizing powered-lift systems; e.g., V/STOL. ASTOVL, and STOL aircraft.

STOL terminal area operating systems (aircraft and onboard avionics, ATC, navigation aids)

NASA Technical Reports Server (NTRS)

Burrous, C.; Erzberger, H.; Johnson, N.; Neuman, F.

1974-01-01

Operational procedures and systems onboard the STOL aircraft which are required to enable the aircraft to perform acceptably in restricted airspace in all types of atmospheric conditions and weather are discussed. Results of simulation and flight investigations to establish operational criteria are presented.

A computer program to calculate the longitudinal aerodynamic characteristics of upper-surface-blown wing-flap configurations

NASA Technical Reports Server (NTRS)

Mendenhall, M. R.

1978-01-01

A user's manual is presented for a computer program in which a vortex-lattice lifting-surface method is used to model the wing and multiple flaps. The engine wake model consists of a series of closely spaced vortex rings with rectangular cross sections. The jet wake is positioned such that the lower boundary of the jet is tangent to the wing and flap upper surfaces. The two potential flow models are used to calculate the wing-flap loading distribution including the influence of the wakes from up to two engines on the semispan. The method is limited to the condition where the flow and geometry of the configurations are symmetric about the vertical plane containing the wing root chord. The results include total configuration forces and moments, individual lifting-surface load distributions, pressure distributions, flap hinge moments, and flow field calculation at arbitrary field points. The use of the program, preparation of input, the output, program listing, and sample cases are described.

A Preliminary Study of V/STOL Transport Aircraft and Bibliography of NASA Research in the VTOL-STOL Field

NASA Technical Reports Server (NTRS)

1961-01-01

This group of papers was prepared by the staff of the Langley Research Center to assist in planning for future commercial air-transport facilities in the New York metropolitan area. Areas of particular interest were predictions regarding the types of V/STOL aircraft that are likely to be developed for various commercial transport applications, estimates of the performance and probable operating procedures for such aircraft, and the approximate dates these aircraft could be available for use. Although the NASA has made no comprehensive studies of this type, the extensive research program in the VTOL-STOL field during the last 10 years appeared to provide a source for some of the desired information . The five papers included herein were therefore prepared to summarize pertinent available material in a form suitable for the intended use. In several instances, new studies and analysis were required to provide the necessary information, but because of a time deadline, many of the significant points received only a cursory examination. For example, much of the quantitative data used in the papers for making generalized comparisons was obtained by approximate methods and is not considered appropriate for use in applications where precise estimates are required. It should be recognized, then, that the treatment of the V/STOL transport provided by this group of papers is necessarily of a preliminary nature.

Aerodynamic Characteristics of a 45 Degree Swept-wing Fighter-Airplane Model and Aerodynamic Loads on Adjacent Stores and Missiles at Mach Numbers of 1.57, 1.87, 2.16, and 2.53

NASA Technical Reports Server (NTRS)

Oehman, Waldo I; Turner, Kenneth L

1958-01-01

An investigation was performed in the Langley Unitary Plan wind tunnel to determine the aerodynamic characteristics of a model of a 450 swept-wing fighter airplane, and to determine the loads on attached stores and detached missiles in the presence of the model. Also included was a determination of aileron-spoiler effectiveness, aileron hinge moments, and the effects of wing modifications on model aerodynamic characteristics. Tests were performed at Mach numbers of 1.57, 1.87, 2.16, and 2.53. The Reynolds numbers for the tests, based on the mean aerodynamic chord of the wing, varied from about 0.9 x 10(exp 6) to 5 x 10(exp 6). The results are presented with minimum analysis.

Limits to Open Class Performance?

NASA Technical Reports Server (NTRS)

Bowers, Albion H.

2008-01-01

This presentation discusses open or unlimited class aircraft performance limitations and design solutions. Limitations in this class of aircraft include slow climbing flight which requires low wing loading, high cruise speed which requires high wing loading, gains in induced or viscous drag alone which result in only half the gain overall and other structural problems (yaw inertia and spins, flutter and static loads integrity). Design solutions include introducing minimum induced drag for a given span (elliptical span load or winglets) and introducing minimum induced drag for a bell shaped span load. It is concluded that open class performance limits (under current rules and technologies) is very close to absolute limits, though some gains remain to be made from unexplored areas and new technologies.

77 FR 24835 - Airworthiness Directives; The Boeing Company Airplanes

Federal Register 2010, 2011, 2012, 2013, 2014

2012-04-26

... inspections must be used for that wing for that visit and for all subsequent repeat inspections. Delta stated... Company Model 767 airplanes. This AD was prompted by reports of cracking in the upper wing skin at the... in the loss of the strut-to-wing upper link load path and possible separation of a strut and engine...

Biologically-Inspired Anisotropic Flexible Wing for Optimal Flapping Flight

DTIC Science & Technology

2013-07-01

AFRL-OSR-VA-TR-2013-0400 Biologically-Inspired, Anisotropic Flexible Wing for Optimal Flapping Flight Luis Bernal, Wei Shyy...Final Report Contract Number: FA9550-07-1-0547 Biologically-Inspired, Anisotropic Flexible Wing for Optimal Flapping Flight University of...minimize power consumption; 2. The interactions of unsteady aerodynamic loading with flexible structures; 3. Flexible , light-weight, multifunctional

Wing Download Results from a Test of a 0.658-Scale V-22 Rotor and Wing

NASA Technical Reports Server (NTRS)

Felker, Fort F.

1992-01-01

A test of a 0.658-scale V-22 rotor and wing was conducted in the 40 x 80 Foot Wind Tunnel at Ames Research Center. One of the principal objectives of the test was to measure the wing download in hover for a variety of test configurations. The wing download and surface pressures were measured for a wide range of thrust coefficients, with five different flap angles, two nacelle angles, and both directions or rotor rotation. This paper presents these results, and describes a new method for interpreting wing surface pressure data in hover. This method shows that the wing flap can produce substantial lift loads in hover.

Proceedings of the 2004 NASA/ONR Circulation Control Workshop, Part 1

NASA Technical Reports Server (NTRS)

Jones, Gregory S. (Editor); Joslin, Ronald D. (Editor)

2005-01-01

As technological advances influence the efficiency and effectiveness of aerodynamic and hydrodynamic applications, designs and operations, this workshop was intended to address the technologies, systems, challenges and successes specific to Coanda driven circulation control in aerodynamics and hydrodynamics. A major goal of this workshop was to determine the 2004 state-of-the-art in circulation control and understand the roadblocks to its application. The workshop addressed applications, CFD, and experiments related to circulation control, emphasizing fundamental physics, systems analysis, and applied research. The workshop consisted of 34 single session oral presentations and written papers that focused on Naval hydrodynamic vehicles (e.g. submarines), Fixed Wing Aviation, V/STOL platforms, propulsion systems (including wind turbine systems), ground vehicles (automotive and trucks) and miscellaneous applications (e.g., poultry exhaust systems and vacuum systems). Several advanced CFD codes were benchmarked using a two-dimensional NCCR circulation control airfoil. The CFD efforts highlighted inconsistencies in turbulence modeling, separation and performance predictions.

A numerical study of the controlled flow tunnel for a high lift model

NASA Technical Reports Server (NTRS)

Parikh, P. C.

1984-01-01

A controlled flow tunnel employs active control of flow through the walls of the wind tunnel so that the model is in approximately free air conditions during the test. This improves the wind tunnel test environment, enhancing the validity of the experimentally obtained test data. This concept is applied to a three dimensional jet flapped wing with full span jet flap. It is shown that a special treatment is required for the high energy wake associated with this and other V/STOL models. An iterative numerical scheme is developed to describe the working of an actual controlled flow tunnel and comparisons are shown with other available results. It is shown that control need be exerted over only part of the tunnel walls to closely approximate free air flow conditions. It is concluded that such a tunnel is able to produce a nearly interference free test environment even with a high lift model in the tunnel.

Pitch, roll, and yaw moment generator for insect-like tailless flapping-wing MAV

NASA Astrophysics Data System (ADS)

Phan, Hoang Vu; Park, Hoon Cheol

2016-04-01

In this work, we proposed a control moment generator, which is called Trailing Edge Change (TEC) mechanism, for attitudes change in hovering insect-like tailless flapping-wing MAV. The control moment generator was installed to the flapping-wing mechanism to manipulate the wing kinematics by adjusting the wing roots location symmetrically or asymmetrically. As a result, the mean aerodynamic force center of each wing is relocated and control moments are generated. The three-dimensional wing kinematics captured by three synchronized high-speed cameras showed that the flapping-wing MAV can properly modify the wing kinematics. In addition, a series of experiments were performed using a multi-axis load cell to evaluate the forces and moments generation. The measurement demonstrated that the TEC mechanism produced reasonable amounts of pitch, roll and yaw moments by shifting position of the trailing edges at the wing roots of the flapping-wing MAV.

Study of short haul high-density V/STOL transportation systems. Volume 2: Appendices

NASA Technical Reports Server (NTRS)

Solomon, H. L.

1972-01-01

Essential supporting data to the short haul transportation study are presented. The specific appendices are arena characteristics, aerospace transportation analysis computer program, economics, model calibration, STOLport siting and services path selection, STOL schedule definition, tabulated California corridor results, and tabulated Midwest arena results.

STOL ride control feasibility study

NASA Technical Reports Server (NTRS)

Gordon, C. K.; Dodson, R. O.

1973-01-01

The feasibility of developing a ride-smoothing control system for a 20-passenger turboprop STOL transport was assessed. Five different ride-control system configurations with varying degrees of complexity, performance, and cost were investigated. Results indicate that a satisfactory ride-control system can be practically implemented on the aircraft with minimum flight performance degradation.

Study of industry information requirements for flight control and navigation systems of STOL aircraft

NASA Technical Reports Server (NTRS)

Gorham, J. A.

1976-01-01

Answers to specific study questions are used to ascertain the data requirements associated with a guidance, navigation and control system for a future civil STOL airplane. Results of the study were used to recommend changes for improving the outputs of the STOLAND flight experiments program.

Computation of wind tunnel model deflections. [for transport type solid wing

NASA Technical Reports Server (NTRS)

Mehrotra, S. C.; Gloss, B. B.

1981-01-01

The experimental deflections for a transport type solid wing model were measured for several single point load conditions. These deflections were compared with those obtained by structural modeling of the wing by using plate and solid elements of Structural Performance Analysis and Redesign (SPAR) program. The solid element representation of the wing showed better agreement with the experimental deflections than the plate representation. The difference between the measured and calculated deflections is about 5 percent.

Aircraft

DOEpatents

Hibbs, B.D.; Lissaman, P.B.S.; Morgan, W.R.; Radkey, R.L.

1998-09-22

This disclosure provides a solar rechargeable aircraft that is inexpensive to produce, is steerable, and can remain airborne almost indefinitely. The preferred aircraft is a span-loaded flying wing, having no fuselage or rudder. Travelling at relatively slow speeds, and having a two-hundred foot wingspan that mounts photovoltaic cells on most all of the wing`s top surface, the aircraft uses only differential thrust of its eight propellers to turn. Each of five sections of the wing has one or more engines and photovoltaic arrays, and produces its own lift independent of the other sections, to avoid loading them. Five two-sided photovoltaic arrays, in all, are mounted on the wing, and receive photovoltaic energy both incident on top of the wing, and which is incident also from below, through a bottom, transparent surface. The aircraft is capable of a top speed of about ninety miles per hour, which enables the aircraft to attain and can continuously maintain altitudes of up to sixty-five thousand feet. Regenerative fuel cells in the wing store excess electricity for use at night, such that the aircraft can sustain its elevation indefinitely. A main spar of the wing doubles as a pressure vessel that houses hydrogen and oxygen gases for use in the regenerative fuel cell. The aircraft has a wide variety of applications, which include weather monitoring and atmospheric testing, communications, surveillance, and other applications as well. 31 figs.

A wrinkle in flight: the role of elastin fibres in the mechanical behaviour of bat wing membranes

PubMed Central

Cheney, Jorn A.; Konow, Nicolai; Bearnot, Andrew; Swartz, Sharon M.

2015-01-01

Bats fly using a thin wing membrane composed of compliant, anisotropic skin. Wing membrane skin deforms dramatically as bats fly, and its three-dimensional configurations depend, in large part, on the mechanical behaviour of the tissue. Large, macroscopic elastin fibres are an unusual mechanical element found in the skin of bat wings. We characterize the fibre orientation and demonstrate that elastin fibres are responsible for the distinctive wrinkles in the surrounding membrane matrix. Uniaxial mechanical testing of the wing membrane, both parallel and perpendicular to elastin fibres, is used to distinguish the contribution of elastin and the surrounding matrix to the overall membrane mechanical behaviour. We find that the matrix is isotropic within the plane of the membrane and responsible for bearing load at high stress; elastin fibres are responsible for membrane anisotropy and only contribute substantially to load bearing at very low stress. The architecture of elastin fibres provides the extreme extensibility and self-folding/self-packing of the wing membrane skin. We relate these findings to flight with membrane wings and discuss the aeromechanical significance of elastin fibre pre-stress, membrane excess length, and how these parameters may aid bats in resisting gusts and preventing membrane flutter. PMID:25833238

Inspiration for wing design: how forelimb specialization enables active flight in modern vertebrates

PubMed Central

2017-01-01

Harnessing flight strategies refined by millions of years of evolution can help expedite the design of more efficient, manoeuvrable and robust flying robots. This review synthesizes recent advances and highlights remaining gaps in our understanding of how bird and bat wing adaptations enable effective flight. Included in this discussion is an evaluation of how current robotic analogues measure up to their biological sources of inspiration. Studies of vertebrate wings have revealed skeletal systems well suited for enduring the loads required during flight, but the mechanisms that drive coordinated motions between bones and connected integuments remain ill-described. Similarly, vertebrate flight muscles have adapted to sustain increased wing loading, but a lack of in vivo studies limits our understanding of specific muscular functions. Forelimb adaptations diverge at the integument level, but both bird feathers and bat membranes yield aerodynamic surfaces with a level of robustness unparalleled by engineered wings. These morphological adaptations enable a diverse range of kinematics tuned for different flight speeds and manoeuvres. By integrating vertebrate flight specializations—particularly those that enable greater robustness and adaptability—into the design and control of robotic wings, engineers can begin narrowing the wide margin that currently exists between flying robots and vertebrates. In turn, these robotic wings can help biologists create experiments that would be impossible in vivo. PMID:28592663

Wind-tunnel investigation of aerodynamic loading on a 0.237-scale model of a remotely piloted research vehicle with a thick, high-aspect-ratio supercritical wing

NASA Technical Reports Server (NTRS)

Byrdsong, T. A.; Brooks, C. W., Jr.

1983-01-01

Wind-tunnel measurements were made of the wing-surface static-pressure distributions on a 0.237 scale model of a remotely piloted research vehicle equipped with a thick, high-aspect-ratio supercritical wing. Data are presented for two model configurations (with and without a ventral pod) at Mach numbers from 0.70 to 0.92 at angles of attack from -4 deg to 8 deg. Large variations of wing-surface local pressure distributions were developed; however, the characteristic supercritical-wing pressure distribution occurred near the design condition of 0.80 Mach number and 2 deg angle of attack. The significant variations of the local pressure distributions indicated pronounced shock-wave movements that were highly sensitive to angle of attack and Mach number. The effect of the vertical pod varied with test conditions; however at the higher Mach numbers, the effects on wing flow characteristics were significant at semispan stations as far outboard as 0.815. There were large variations of the wing loading in the range of test conditions, both model configurations exhibited a well-defined peak value of normal-force coefficient at the cruise angle of attack (2 deg) and Mach number (0.80).

A wrinkle in flight: the role of elastin fibres in the mechanical behaviour of bat wing membranes.

PubMed

Cheney, Jorn A; Konow, Nicolai; Bearnot, Andrew; Swartz, Sharon M

2015-05-06

Bats fly using a thin wing membrane composed of compliant, anisotropic skin. Wing membrane skin deforms dramatically as bats fly, and its three-dimensional configurations depend, in large part, on the mechanical behaviour of the tissue. Large, macroscopic elastin fibres are an unusual mechanical element found in the skin of bat wings. We characterize the fibre orientation and demonstrate that elastin fibres are responsible for the distinctive wrinkles in the surrounding membrane matrix. Uniaxial mechanical testing of the wing membrane, both parallel and perpendicular to elastin fibres, is used to distinguish the contribution of elastin and the surrounding matrix to the overall membrane mechanical behaviour. We find that the matrix is isotropic within the plane of the membrane and responsible for bearing load at high stress; elastin fibres are responsible for membrane anisotropy and only contribute substantially to load bearing at very low stress. The architecture of elastin fibres provides the extreme extensibility and self-folding/self-packing of the wing membrane skin. We relate these findings to flight with membrane wings and discuss the aeromechanical significance of elastin fibre pre-stress, membrane excess length, and how these parameters may aid bats in resisting gusts and preventing membrane flutter. © 2015 The Author(s) Published by the Royal Society. All rights reserved.

Investigation of a robust tendon-sheath mechanism for flexible membrane wing application in mini-UAV

NASA Astrophysics Data System (ADS)

Lee, Shian; Tjahjowidodo, Tegoeh; Lee, Hsuchew; Lai, Benedict

2017-02-01

Two inherent issues manifest themselves in flying mini-unmanned aerial vehicles (mini-UAV) in the dense area at tropical climate regions, namely disturbances from gusty winds and limited space for deployment tasks. Flexible membrane wing (FMW) UAVs are seen to be potentials to mitigate these problems. FMWs are adaptable to gusty airflow as the wings are able to flex according to the gust load to reduce the effective angle-of-attack, thus, reducing the aerodynamic loads on the wing. On the other hand, the flexible structure is allowing the UAV to fold in a compact package, and later on, the mini-UAV can be deployed instantly from the storage tube, e.g. through a catapult mechanism. This paper discusses the development of an FMW UAV actuated by a tendon-sheath mechanism (TSM). This approach allows the wing to morph to generate a rolling moment, while still allowing the wing to fold. Dynamic characteristics of the mechanism that exhibits the strong nonlinear phenomenon of friction on TSM are modeled and compensated for. A feed-forward controller was implemented based on the identified nonlinear behavior to control the warping position of the wing. The proposed strategy is validated experimentally in a wind tunnel facility by creating a gusty environment that is imitating a realistic gusty condition based upon the results of computational fluid dynamics (CFD) simulation. The results demonstrate a stable and robust wing-warping actuation, even in gusty conditions. Accurate wing-warping can be achieved via the TSM, while also allowing the wings to fold.

Intra-specific variation in wing morphology and its impact on take-off performance in blue tits (Cyanistes caeruleus) during escape flights

PubMed Central

McFarlane, Laura; Altringham, John D.; Askew, Graham N.

2016-01-01

ABSTRACT Diurnal and seasonal increases in body mass and seasonal reductions in wing area may compromise a bird's ability to escape, as less of the power available from the flight muscles can be used to accelerate and elevate the animal's centre of mass. Here, we investigated the effects of intra-specific variation in wing morphology on escape take-off performance in blue tits (Cyanistes caeruleus). Flights were recorded using synchronised high-speed video cameras and take-off performance was quantified as the sum of the rates of change of the kinetic and potential energies of the centre of mass. Individuals with a lower wing loading, WL (WL=body weight/wing area) had higher escape take-off performance, consistent with the increase in lift production expected from relatively larger wings. Unexpectedly, it was found that the total power available from the flight muscles (estimated using an aerodynamic analysis) was inversely related to WL. This could simply be because birds with a higher WL have relatively smaller flight muscles. Alternatively or additionally, variation in the aerodynamic load on the wing resulting from differences in wing morphology will affect the mechanical performance of the flight muscles via effects on the muscle's length trajectory. Consistent with this hypothesis is the observation that wing beat frequency and relative downstroke duration increase with decreasing WL; both are factors that are expected to increase muscle power output. Understanding how wing morphology influences take-off performance gives insight into the potential risks associated with feather loss and seasonal and diurnal fluctuations in body mass. PMID:26994175

Optimal pitching axis location of flapping wings for efficient hovering flight.

PubMed

Wang, Q; Goosen, J F L; van Keulen, F

2017-09-01

Flapping wings can pitch passively about their pitching axes due to their flexibility, inertia, and aerodynamic loads. A shift in the pitching axis location can dynamically alter the aerodynamic loads, which in turn changes the passive pitching motion and the flight efficiency. Therefore, it is of great interest to investigate the optimal pitching axis for flapping wings to maximize the power efficiency during hovering flight. In this study, flapping wings are modeled as rigid plates with non-uniform mass distribution. The wing flexibility is represented by a linearly torsional spring at the wing root. A predictive quasi-steady aerodynamic model is used to evaluate the lift generated by such wings. Two extreme power consumption scenarios are modeled for hovering flight, i.e. the power consumed by a drive system with and without the capacity of kinetic energy recovery. For wings with different shapes, the optimal pitching axis location is found such that the cycle-averaged power consumption during hovering flight is minimized. Optimization results show that the optimal pitching axis is located between the leading edge and the mid-chord line, which shows close resemblance to insect wings. An optimal pitching axis can save up to 33% of power during hovering flight when compared to traditional wings used by most of flapping wing micro air vehicles (FWMAVs). Traditional wings typically use the straight leading edge as the pitching axis. With the optimized pitching axis, flapping wings show higher pitching amplitudes and start the pitching reversals in advance of the sweeping reversals. These phenomena lead to higher lift-to-drag ratios and, thus, explain the lower power consumption. In addition, the optimized pitching axis provides the drive system higher potential to recycle energy during the deceleration phases as compared to their counterparts. This observation underlines the particular importance of the wing pitching axis location for energy-efficient FWMAVs when using kinetic energy recovery drive systems.

Estimation of wing nonlinear aerodynamic characteristics at supersonic speeds

NASA Technical Reports Server (NTRS)

Carlson, H. W.; Mack, R. J.

1980-01-01

A computational system for estimation of nonlinear aerodynamic characteristics of wings at supersonic speeds was developed and was incorporated in a computer program. This corrected linearized theory method accounts for nonlinearities in the variation of basic pressure loadings with local surface slopes, predicts the degree of attainment of theoretical leading edge thrust, and provides an estimate of detached leading edge vortex loadings that result when the theoretical thrust forces are not fully realized.

An in-flight simulation of approach and landing of a STOL transport with adverse ground effect

NASA Technical Reports Server (NTRS)

Ellis, D. R.

1976-01-01

The results of an in-flight simulation program undertaken to study the problems of landing a representative STOL transport in the presence of adverse ground effects are presented. Landings were performed with variations in ground effect magnitude, ground effect lag, and thrust response. Other variations covered the effects of augmented lift response, SAS-failures, turbulence, segmented approach, and flare warning. The basic STOL airplane required coordinated use of both stick and throttle for consistently acceptable landings, and the presence of adverse ground effects made the task significantly more difficult. Ground effect lag and good engine response gave noticeable improvement, as did augmented lift response.

Hypermedia and intelligent tutoring applications in a mission operations environment

NASA Technical Reports Server (NTRS)

Ames, Troy; Baker, Clifford

1990-01-01

Hypermedia, hypertext and Intelligent Tutoring System (ITS) applications to support all phases of mission operations are investigated. The application of hypermedia and ITS technology to improve system performance and safety in supervisory control is described - with an emphasis on modeling operator's intentions in the form of goals, plans, tasks, and actions. Review of hypermedia and ITS technology is presented as may be applied to the tutoring of command and control languages. Hypertext based ITS is developed to train flight operation teams and System Test and Operation Language (STOL). Specific hypermedia and ITS application areas are highlighted, including: computer aided instruction of flight operation teams (STOL ITS) and control center software development tools (CHIMES and STOL Certification Tool).

Maintenance problems associated with the operation of the F402 /Pegasus/ engine in the AV-8A /Harrier/ aircraft

NASA Technical Reports Server (NTRS)

Stanley, C. W.; Hood, W. E.

1981-01-01

The U.S. Marine Corp (USMC) has been operating the only V/STOL attack aircraft in the western world since 1971. Some of the maintenance problems experienced are related to the unique V/STOL design criteria of the Pegasus engine. However, the major part of the required maintenance effort is found to involve the more conventional engine problems. A description of the aircraft engine is provided and the problems resulting from V/STOL design demands are examined. Attention is given to the fuel system control, the engine air bleed, foreign object damage to the hp compressor, and the engine exhaust system.

Measurement and Analysis of Wing and Tail Buffeting Loads on a Fighter Airplane

NASA Technical Reports Server (NTRS)

Huston, Wilber B; Skopinski, T H

1955-01-01

The buffeting loads measured on the wing and tail of a fighter airplane during 194 maneuvers are given in tabular form, along with the associated flight conditions. Measurements were made at altitudes of 30,000 to 10,000 feet and at speeds up to a Mach number of 0.8. Least-squares methods have been used for a preliminary analysis of the data. The agreement between the results of this analysis and the loads measured in stalls is sufficiently good to suggest the examination of the buffeting of other airplanes on the same basis.

Buckling characteristics of hypersonic aircraft wing tubular panels

NASA Technical Reports Server (NTRS)

Ko, William L.; Shideler, John L.; Fields, Roger A.

1986-01-01

The buckling characteristics of Rene 41 tubular panels installed as wing panels on a hypersonic wing test structure (HWTS) were determined nondestructively through use of a force/stiffness technique. The nondestructive buckling tests were carried out under different combined load conditions and different temperature environments. Two panels were subsequently tested to buckling failure in a universal tension compression testing machine. In spite of some data scattering because of large extrapolations of data points resulting from termination of the test at a somewhat low applied load, the overall test data correlated fairly well with theoretically predicted buckling interaction curves. The structural efficiency of the tubular panels was slightly higher than that of the beaded panels which they replaced.

Player Load in Elite Netball: Match, Training, and Positional Comparisons.

PubMed

Young, Christopher M; Gastin, Paul B; Sanders, Nick; Mackey, Luke; Dwyer, Dan B

2016-11-01

The activity profile of competition and training in elite netball has not been comprehensively reported in the literature. To measure and analyze player load in elite netballers during matches and training sessions. The primary research question was, How does player load vary between playing positions in a match and between matches and training sessions? Various measures of player load were recorded in 12 elite professional netballers with a mean ± SD age of 26 ± 4.9 y and height of 183.2 ± 8.7 cm. Player load was assessed using a published method that uses accelerometry. Load was represented as total load in arbitrary units (au), playing intensity (au/min), and relative time spent in each of 4 playing intensity zones (low, low to moderate, moderate, and high). Data from 15 games and up to 17 training sessions were analyzed for each player. Player load in matches for the goal-based positions (goal shooter, goal keeper, and goal defense) tended to be lower than the attacking and wing-based positions (goal attack, wing attack, wing defense, and center). The difference was largely due to the amount of time spent in low-intensity activity. Playing intensity of matches was greater than in training sessions; however, the total time spent in moderate- to high-intensity activities was not practically different. Accelerometry is a valuable method of measuring player load in netball, and the present results provide new information about the activity profile of different playing positions.

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility, Jerry Belt, with United Space Alliance, checks a spar attachment on the wing of the orbiter Atlantis before installing Reinforced Carbon Carbon (RCC) panels on the wing. The spars - floating joints - reduce loading on the panels caused by wing deflections. The gray carbon composite RCC panels have sufficient strength to withstand the aerodynamic forces experienced during launch and reentry, which can reach as high as 800 pounds per square foot. The operating range of RCC is from minus 250º F to about 3,000º F, the temperature produced by friction with the atmosphere during reentry.

NASA Image and Video Library

2003-09-05

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility, Jerry Belt, with United Space Alliance, checks a spar attachment on the wing of the orbiter Atlantis before installing Reinforced Carbon Carbon (RCC) panels on the wing. The spars - floating joints - reduce loading on the panels caused by wing deflections. The gray carbon composite RCC panels have sufficient strength to withstand the aerodynamic forces experienced during launch and reentry, which can reach as high as 800 pounds per square foot. The operating range of RCC is from minus 250º F to about 3,000º F, the temperature produced by friction with the atmosphere during reentry.

DC-9 V/STOL Transport Model in the 40x80 Foot Wind Tunnel.

NASA Image and Video Library

1971-04-28

3/4 front view of McDonnell-Douglas Large-Scale lift fan, vertical and/or short take-off and landing (V/STOL), transport model. Francis Malerick in photograph. The McDonnell Douglas DC-9 (initially known as the Douglas DC-9) is a twin-engine, single-aisle jet airliner.

A flight-test methodology for identification of an aerodynamic model for a V/STOL aircraft

NASA Technical Reports Server (NTRS)

Bach, Ralph E., Jr.; Mcnally, B. David

1988-01-01

Described is a flight test methodology for developing a data base to be used to identify an aerodynamic model of a vertical and short takeoff and landing (V/STOL) fighter aircraft. The aircraft serves as a test bed at Ames for ongoing research in advanced V/STOL control and display concepts. The flight envelope to be modeled includes hover, transition to conventional flight, and back to hover, STOL operation, and normaL cruise. Although the aerodynamic model is highly nonlinear, it has been formulated to be linear in the parameters to be identified. Motivation for the flight test methodology advocated in this paper is based on the choice of a linear least-squares method for model identification. The paper covers elements of the methodology from maneuver design to the completed data base. Major emphasis is placed on the use of state estimation with tracking data to ensure consistency among maneuver variables prior to their entry into the data base. The design and processing of a typical maneuver is illustrated.

A subjective evaluation of synthesized STOL airplane noises

NASA Technical Reports Server (NTRS)

Powell, C. A., Jr.

1973-01-01

A magnitude-estimation experiment was conducted to evaluate the subjective annoyance of the noise generated by possible future turbofan STOL aircraft as compared to that of several current CTOL aircraft. In addition, some of the units used to scale the magnitude of aircraft noise were evaluated with respect to their applicability to STOL noise. Twenty test subjects rated their annoyance to a total of 119 noises over a range of 75 PNdb to 105 PNdb. Their subjective ratings were compared with acoustical analysis of the noises in terms of 28 rating scale units. The synthesized STOL noises of this experiment were found to be slightly more annoying than the conventional CTOL noises at equal levels of PNL and EPNL. Over the range of levels investigated the scaling units, with a few exceptions, were capable of predicting the points of equal annoyance for all of the noises with plus or minus 3 dB. The inclusion of duration corrections, in general, improved the predictive capabilities of the various scaling units; however, tone corrections reduced their predictive capabilities.

Development of the Rules Governing the Strength of Airplanes. Part I : German Loading Conditions up to 1926

NASA Technical Reports Server (NTRS)

Kussner, H G; Thalau, Karl

1933-01-01

Load factors and loading conditions are presented for German aircraft. Loading conditions under various stress factors are presented along with a breakdown of individual aircraft components such as landing gear, wings, etc.

Aircraft Design

NASA Technical Reports Server (NTRS)

Uden, Edward (Inventor); Bowers, Albion H. (Inventor)

2016-01-01

The present invention is an aircraft wing design that creates a bell shaped span load, which results in a negative induced drag (induced thrust) on the outer portion of the wing; such a design obviates the need for rudder control of an aircraft.

Determination of stores pointing error due to wing flexibility under flight load

NASA Technical Reports Server (NTRS)

Lokos, William A.; Bahm, Catherine M.; Heinle, Robert A.

1995-01-01

The in-flight elastic wing twist of a fighter-type aircraft was studied to provide for an improved on-board real-time computed prediction of pointing variations of three wing store stations. This is an important capability to correct sensor pod alignment variation or to establish initial conditions of iron bombs or smart weapons prior to release. The original algorithm was based upon coarse measurements. The electro-optical Flight Deflection Measurement System measured the deformed wing shape in flight under maneuver loads to provide a higher resolution database from which an improved twist prediction algorithm could be developed. The FDMS produced excellent repeatable data. In addition, a NASTRAN finite-element analysis was performed to provide additional elastic deformation data. The FDMS data combined with the NASTRAN analysis indicated that an improved prediction algorithm could be derived by using a different set of aircraft parameters, namely normal acceleration, stores configuration, Mach number, and gross weight.

Results of a Cyclic Load Test of an RB-47E Airplane

NASA Technical Reports Server (NTRS)

Huston, Wilber B.

1959-01-01

Results of a cyclic load test made by NASA on an EB-47E airplane are given. The test reported on is for one of three B-47 airplanes in a test program set up by the U. S. Air Force to evaluate the effect of wing structural reinforcements on fatigue life. As a result of crack development in the upper fuselage longerons of the other two airplanes in the program, a longeron and fuselage skin modification was incorporated early in the test. Fuselage strain-gage measurements made before and after the longeron modification and wing strain-gage measurements made only after wing reinforcement are summarized. The history of crack development and repair is given in detail. Testing was terminated one sequence short of the planned end of the program with the occurrence of a major crack in the lower right wing skin.

Chordwise and compressibility corrections to slender-wing theory

NASA Technical Reports Server (NTRS)

Lomax, Harvard; Sluder, Loma

1952-01-01

Corrections to slender-wing theory are obtained by assuming a spanwise distribution of loading and determining the chordwise variation which satisfies the appropriate integral equation. Such integral equations are set up in terms of the given vertical induced velocity on the center line or, depending on the type of wing plan form, its average value across the span at a given chord station. The chordwise distribution is then obtained by solving these integral equations. Results are shown for flat-plate rectangular, and triangular wings.

Fatigue Testing of Wing Beam by the Resonance Method

NASA Technical Reports Server (NTRS)

Bleakney, William M

1938-01-01

Preliminary fatigue tests on two aluminum-alloy wing-beam specimens subjected to reversed axial loading are described. The motion used consists in incorporating one or two reciprocating motors in a resonance system of which the specimen is the spring element. A description is given of the reciprocating motors, and of the method of assembling and adjusting the vibrating system. The results indicate that the method is well adapted to fatigue tests of not only uniform wing beams but also wing beams with asymmetrical local reinforcements.

Active In-Flight Load Redistribution Utilizing Fiber-Optic Shape Sensing and Multiple Control Surfaces

NASA Technical Reports Server (NTRS)

Pena, Francisco; Martins, Benjamin L.; Richards, W. Lance

2018-01-01

Morphing wing technologies have gained research interest in recent years as technological advancements pave the way for such innovations. A key benefit of such a morphing wing concept is the ability of the wing to transition into an optimal configuration at multiple flight conditions. Such a morphing wing will have applications not only in drag reduction but also in flutter suppression and gust alleviation. By manipulating the wing geometry to match a given flight profile it is likely that the wing will yield increases in not just aerodynamic efficiency but also structural efficiency. These structurally efficient designs will likely rely on some type of structural sensing system which will ensure the wing maintains positive margins throughout its flight profile.

Detailed pressure distribution measurements obtained on several configurations of an aspect-ratio-7 variable twist wing

NASA Technical Reports Server (NTRS)

Holbrook, G. T.; Dunham, D. M.

1985-01-01

Detailed pressure distribution measurements were made for 11 twist configurations of a unique, multisegmented wing model having an aspect ratio of 7 and a taper ratio of 1. These configurations encompassed span loads ranging from that of an untwisted wing to simple flapped wings both with and without upper-surface spoilers attached. For each of the wing twist configurations, electronic scanning pressure transducers were used to obtain 580 surface pressure measurements over the wing in about 0.1 sec. Integrated pressure distribution measurements compared favorably with force-balance measurements of lift on the model when the model centerbody lift was included. Complete plots and tabulations of the pressure distribution data for each wing twist configuration are provided.

Optimum Design of a Compound Helicopter

NASA Technical Reports Server (NTRS)

Yeo, Hyeonsoo; Johnson, Wayne

2006-01-01

A design and aeromechanics investigation was conducted for a 100,000-lb compound helicopter with a single main rotor, which is to cruise at 250 knots at 4000 ft/95 deg F condition. Performance, stability, and control analyses were conducted with the comprehensive rotorcraft analysis CAMRAD II. Wind tunnel test measurements of the performance of the H-34 and UH-1D rotors at high advance ratio were compared with calculations to assess the accuracy of the analysis for the design of a high speed helicopter. In general, good correlation was obtained when an increase of drag coefficients in the reverse flow region was implemented. An assessment of various design parameters (disk loading, blade loading, wing loading) on the performance of the compound helicopter was conducted. Lower wing loading (larger wing area) and higher blade loading (smaller blade chord) increased aircraft lift-to-drag ratio. However, disk loading has a small influence on aircraft lift-to-drag ratio. A rotor parametric study showed that most of the benefit of slowing the rotor occurred at the initial 20 to 30% reduction of the advancing blade tip Mach number. No stability issues were observed with the current design. Control derivatives did not change significantly with speed, but the did exhibit significant coupling.

Structural Concepts Study of Non-circular Fuselage Configurations

NASA Technical Reports Server (NTRS)

Mukhopadhyay, Vivel

1996-01-01

A preliminary study of structural concepts for noncircular fuselage configurations is presented. For an unconventional flying-wing type aircraft, in which the fuselage is inside the wing, multiple fuselage bays with non-circular sections need to be considered. In a conventional circular fuselage section, internal pressure is carried efficiently by a thin skin via hoop tension. If the section is non-circular, internal pressure loads also induce large bending stresses. The structure must also withstand additional bending and compression loads from aerodynamic and gravitational forces. Flat and vaulted shell structural configurations for such an unconventional, non-circular pressurized fuselage of a large flying-wing were studied. A deep honeycomb sandwich-shell and a ribbed double-wall shell construction were considered. Combinations of these structural concepts were analyzed using both analytical and simple finite element models of isolated sections for a comparative conceptual study. Weight, stress, and deflection results were compared to identify a suitable configuration for detailed analyses. The flat sandwich-shell concept was found preferable to the vaulted shell concept due to its superior buckling stiffness. Vaulted double-skin ribbed shell configurations were found to be superior due to their weight savings, load diffusion, and fail-safe features. The vaulted double-skin ribbed shell structure concept was also analyzed for an integrated wing-fuselage finite element model. Additional problem areas such as wing-fuselage junction and pressure-bearing spar were identified.

Calculation of the aerodynamic loading of swept and unswept flexible wings of arbitrary stiffness

NASA Technical Reports Server (NTRS)

Diederich, Franklin W

1950-01-01

A method is presented for calculating the aerodynamic loading, the divergence speed, and certain stability derivatives of swept and unswept wings and tail surfaces of arbitrary stiffness. Provision is made for using either stiffness curves and root rotation constants or structural influence coefficients in the analysis. Computing forms, tables of numerical constants required in the analysis, and an illustrative example are included to facilitate calculations by means of the method.

DOE Office of Scientific and Technical Information (OSTI.GOV)

Huang, J. Y.; E, J. C.; Huang, J. W.

Impact fracture of single-crystal Si is critical to long-term reliability of electronic devices and solar cells for its wide use as components or substrates in semiconductor industry. Single-crystal Si is loaded along two different crystallographic directions with a split Hopkinson pressure bar integrated with an in situ x-ray imaging and diffraction system. Bulk stress histories are measured, simultaneously with x-ray phase contrast imaging (XPCI) and Laue diffraction. Damage evolution is quantified with grayscale maps from XPCI. Single-crystal Si exhibits pronounced anisotropy in fracture modes, and thus fracture strengths and damage evolution. For loading along [11¯ 0] and viewing along [001],more » (1¯1¯0)[11¯ 0] cleavage is activated and induces horizontal primary cracks followed by perpendicular wing cracks. However, for loading along [011¯] and viewing along [111], random nucleation and growth of shear and tensile-splitting crack networks lead to catastrophic failure of materials with no cleavage. The primary-wing crack mode leads to a lower characteristic fracture strength due to predamage, but a more concentrated strength distribution, i.e., a higher Weibull modulus, compared to the second loading case. Furthermore, the sequential primary cracking, wing cracking and wing-crack coalescence processes result in a gradual increase of damage with time, deviating from theoretical predictions. Particle size and aspect ratios of fragments are discussed with postmortem fragment analysis, which verifies fracture modes observed in XPCI.« less

Analytical Fuselage and Wing Weight Estimation of Transport Aircraft

DOT National Transportation Integrated Search

1996-05-01

A method of estimating the load-bearing fuselage weight and wing weight of transport aircraft based on fundamental structural principles has been developed. This method of weight estimation represents a compromise between the rapid assessment of comp...

A design approach and selected wind tunnel results at high subsonic speeds for wing-tip mounted winglets

NASA Technical Reports Server (NTRS)

Whitcomb, R. T.

1976-01-01

Winglets, which are small, nearly vertical, winglike surfaces, substantially reduce drag coefficients at lifting conditions. The primary winglet surfaces are rearward above the wing tips; secondary surfaces are forward below the wing tips. This report presents a discussion of the considerations involved in the design of the winglets; measured effects of these surfaces on the aerodynamic forces, moments, and loads for a representative first generation, narrow body jet transport wing; and a comparison of these effects with those for a wing tip extension which results in approximately the same increase in bending moment at the wing-fuselage juncture as did the addition of the winglets.

Conceptual design of high speed supersonic aircraft: A brief review on SR-71 (Blackbird) aircraft

NASA Astrophysics Data System (ADS)

Xue, Hui; Khawaja, H.; Moatamedi, M.

2014-12-01

The paper presents the conceptual design of high-speed supersonic aircraft. The study focuses on SR-71 (Blackbird) aircraft. The input to the conceptual design is a mission profile. Mission profile is a flight profile of the aircraft defined by the customer. This paper gives the SR-71 aircraft mission profile specified by US air force. Mission profile helps in defining the attributes the aircraft such as wing profile, vertical tail configuration, propulsion system, etc. Wing profile and vertical tail configurations have direct impact on lift, drag, stability, performance and maneuverability of the aircraft. A propulsion system directly influences the performance of the aircraft. By combining the wing profile and the propulsion system, two important parameters, known as wing loading and thrust to weight ratio can be calculated. In this work, conceptual design procedure given by D. P. Raymer (AIAA Educational Series) is applied to calculate wing loading and thrust to weight ratio. The calculated values are compared against the actual values of the SR-71 aircraft. Results indicates that the values are in agreement with the trend of developments in aviation.

Fiber-optically sensorized composite wing

NASA Astrophysics Data System (ADS)

Costa, Joannes M.; Black, Richard J.; Moslehi, Behzad; Oblea, Levy; Patel, Rona; Sotoudeh, Vahid; Abouzeida, Essam; Quinones, Vladimir; Gowayed, Yasser; Soobramaney, Paul; Flowers, George

2014-04-01

Electromagnetic interference (EMI) immune and light-weight, fiber-optic sensor based Structural Health Monitoring (SHM) will find increasing application in aerospace structures ranging from aircraft wings to jet engine vanes. Intelligent Fiber Optic Systems Corporation (IFOS) has been developing multi-functional fiber Bragg grating (FBG) sensor systems including parallel processing FBG interrogators combined with advanced signal processing for SHM, structural state sensing and load monitoring applications. This paper reports work with Auburn University on embedding and testing FBG sensor arrays in a quarter scale model of a T38 composite wing. The wing was designed and manufactured using fabric reinforced polymer matrix composites. FBG sensors were embedded under the top layer of the composite. Their positions were chosen based on strain maps determined by finite element analysis. Static and dynamic testing confirmed expected response from the FBGs. The demonstrated technology has the potential to be further developed into an autonomous onboard system to perform load monitoring, SHM and Non-Destructive Evaluation (NDE) of composite aerospace structures (wings and rotorcraft blades). This platform technology could also be applied to flight testing of morphing and aero-elastic control surfaces.

Wind tunnel tests of a free-wing/free-trimmer model

NASA Technical Reports Server (NTRS)

Sandlin, D. R.

1982-01-01

The riding qualities of an aircraft with low wing loading can be improved by freeing the wing to rotate about its spanwise axis. A trimming surface also free to rotate about its spanwise axis can be added at the wing tips to permit the use of high lift devices. Wind tunnel tests of the free wing/free trimmer model with the trimmer attached to the wing tips aft of the wing chord were conducted to validate a mathematical model developed to predict the dynamic characteristics of a free wing/free trimmer aircraft. A model consisting of a semispan wing with the trimmer mounted on with the wing on an air bearing and the trimmer on a ball bearing was displaced to various angles of attack and released. The damped oscillations of the wing and trimmer were recorded. Real and imaginary parts of the characteristic equations of motion were determined and compared to values predicted using the mathematical model.

Vortex maneuver lift for super-cruise configurations

NASA Technical Reports Server (NTRS)

Campbell, J. F.; Gloss, B. B.; Lamar, J. E.

1976-01-01

Some of the theoretical and experimental research conducted at the NASA Langley Research Center is presented to investigate the subsonic vortex-lift producing capabilities for two classes of Super-Cruise designs: a close-coupled wing-canard arrangement and a slender wing configuration. In addition, several analytical methods are discussed for estimating critical structural design loads for thin, highly swept wings having separated leading-edge vortex flows.

High-Lift Systems on Commercial Subsonic Airliners

NASA Technical Reports Server (NTRS)

Rudolph, Peter K. C.

1996-01-01

The early breed of slow commercial airliners did not require high-lift systems because their wing loadings were low and their speed ratios between cruise and low speed (takeoff and landing) were about 2:1. However, even in those days the benefit of high-lift devices was recognized. Simple trailing-edge flaps were in use, not so much to reduce landing speeds, but to provide better glide-slope control without sideslipping the airplane and to improve pilot vision over the nose by reducing attitude during low-speed flight. As commercial-airplane cruise speeds increased with the development of more powerful engines, wing loadings increased and a real need for high-lift devices emerged to keep takeoff and landing speeds within reasonable limits. The high-lift devices of that era were generally trailing-edge flaps. When jet engines matured sufficiently in military service and were introduced commercially, airplane speed capability had to be increased to best take advantage of jet engine characteristics. This speed increase was accomplished by introducing the wing sweep and by further increasing wing loading. Whereas increased wing loading called for higher lift coefficients at low speeds, wing sweep actually decreased wing lift at low speeds. Takeoff and landing speeds increased on early jet airplanes, and, as a consequence, runways worldwide had to be lengthened. There are economical limits to the length of runways; there are safety limits to takeoff and landing speeds; and there are speed limits for tires. So, in order to hold takeoff and landing speeds within reasonable limits, more powerful high-lift devices were required. Wing trailing-edge devices evolved from plain flaps to Fowler flaps with single, double, and even triple slots. Wing leading edges evolved from fixed leading edges to a simple Krueger flap, and from fixed, slotted leading edges to two- and three-position slats and variable-camber (VC) Krueger flaps. The complexity of high-lift systems probably peaked on the Boeing 747, which has a VC Krueger flap and triple-slotted, inboard and outboard trailing-edge flaps. Since then, the tendency in high-lift system development has been to achieve high levels of lift with simpler devices in order to reduce fleet acquisition and maintenance costs. The intent of this paper is to: (1) review available high-lift devices, their functions, and design criteria; (2) appraise high-lift systems presently in service on commercial air liners; (3) present personal study results on high-lift systems; (4) develop a weight and cost model for high-lift systems; and (5) discuss the development tendencies of future high-lift systems.

Ground effects and control effectiveness tests of a .095 scale powered model of a modified T-39 lift/cruise fan V/STOL research airplane

NASA Technical Reports Server (NTRS)

Dawson, C. R.; Omar, E.

1977-01-01

Wind tunnel test data are analysed to determine ground effects and the effectiveness of the aerodynamic control surfaces to provide a technology base for a Navy type A V/STOL airplane. Three 14CM (5.5 inch) turbopowered simulators were used to power the model which was tested primarily in the following configurations: (1) VTOL with flaps deployed, gear down, and engines tilted to 80 deg, 90 deg, and 95 deg, (2) STOL with flap and gear down and engines tilted to 50 deg; and (3) Loiter with flaps and gear up and L/C nacelles off. Data acquired during the tests are included as an appendix.

V/STOL wind-tunnel testing

NASA Technical Reports Server (NTRS)

Koenig, D. G.

1984-01-01

Factors influencing effective program planning for V/STOL wind-tunnel testing are discussed. The planning sequence itself, which includes a short checklist of considerations that could enhance the value of the tests, is also described. Each of the considerations, choice of wind tunnel, type of model installation, model development and test operations, is discussed, and examples of appropriate past and current V/STOL test programs are provided. A short survey of the moderate to large subsonic wind tunnels is followed by a review of several model installations, from two-dimensional to large-scale models of complete aircraft configurations. Model sizing, power simulation, and planning are treated, including three areas is test operations: data-acquisition systems, acoustic measurements in wind tunnels, and flow surveying.

Investigation of a laser Doppler velocimeter system to measure the flow field around a large scale V/STOL aircraft in ground effect

NASA Technical Reports Server (NTRS)

Zalay, A. D.; Brashears, M. R.; Jordan, A. J.; Shrider, K. R.; Vought, C. D.

1979-01-01

The flow field measured around a hovering 70 percent scale vertical takeoff and landing (V/STOL) aircraft model is described. The velocity measurements were conducted with a ground based laser Doppler velocimeter. The remote sensing instrumentation and experimental tests of the velocity surveys are discussed. The distribution of vertical velocity in the fan jet and fountain; the radial velocity in the wall jet and the horizontal velocity along the aircraft underside are presented for different engine rpms and aircraft height above ground. Results show that it is feasible to use a mobile laser Doppler velocimeter to measure the flow field generated by a large scale V/STOL aircraft operating in ground effect.

Noise Considerations for V/STOL Transports

NASA Technical Reports Server (NTRS)

Kenyon, George C.

1968-01-01

Noise consideration may well be as important a factor in future aircraft concept selection as such economic factors as operating cost and profitability. The impact of noise on some of the design and operational aspects of future V/STOL transports is examined in detail, including consideration of configuration, attitude-control system, lift system, and terminal flight pattern. Extended vertical rise of VTOL aircraft as a method of limiting the intense noise exposure to the terminal area is shown to be only partially effective as well as costly. Comparisons are made of noise contours for conceptual V/STOL transports for several PNdB criteria. The variation in extent of affected area with configuration and criterion emphasizes the importance of establishing an "acceptable" noise level for "city-center" operation.

Reliability aspects of a composite bolted scarf joint. [in wing skin splice

NASA Technical Reports Server (NTRS)

Reed, D. L.; Eisenmann, J. R.

1975-01-01

The design, fabrication, static test, and fatigue test of both tension and compression graphite-epoxy candidates for a wing splice representative of a next-generation transport aircraft was the objective of the reported research program. A single-scarf bolted joint was selected as the design concept. Test specimens were designed and fabricated to represent an upper-surface and a lower-surface panel containing the splice. The load spectrum was a flight-by-flight random-load history including ground-air-ground loads. The results of the fatigue testing indicate that, for this type of joint, the inherent fatigue resistance of the laminate is reflected in the joint behavior and, consequently, the rate of damage accumulation is very slow under realistic fatigue loadings.

Active load control during rolling maneuvers. [performed in the Langley Transonic Dynamics Tunnel

NASA Technical Reports Server (NTRS)

Woods-Vedeler, Jessica A.; Pototzky, Anthony S.; Hoadley, Sherwood T.

1994-01-01

A rolling maneuver load alleviation (RMLA) system has been demonstrated on the active flexible wing (AFW) wind tunnel model in the Langley Transonic Dynamics Tunnel (TDT). The objective was to develop a systematic approach for designing active control laws to alleviate wing loads during rolling maneuvers. Two RMLA control laws were developed that utilized outboard control-surface pairs (leading and trailing edge) to counteract the loads and that used inboard trailing-edge control-surface pairs to maintain roll performance. Rolling maneuver load tests were performed in the TDT at several dynamic pressures that included two below and one 11 percent above open-loop flutter dynamic pressure. The RMLA system was operated simultaneously with an active flutter suppression system above open-loop flutter dynamic pressure. At all dynamic pressures for which baseline results were obtained, torsion-moment loads were reduced for both RMLA control laws. Results for bending-moment load reductions were mixed; however, design equations developed in this study provided conservative estimates of load reduction in all cases.

The Flying Diamond: A joined aircraft configuration design project, volume 1

NASA Technical Reports Server (NTRS)

Ball, Chris; Czech, Joe; Lentz, Bryan; Kobashigawa, Daryl; Oishi, Curtis; Poladian, David

1988-01-01

The results of the analysis conducted on the Joined Wing Configuration study are presented. The joined wing configuration employs a conventional fuselage and incorporates two wings joined together near their tips to form a diamond shape in both plan view and front view. The arrangement of the lifting surfaces uses the rear wing as a horizontal tail and as a forward wing strut. The rear wing has its root at the tip of the vertical stabilizer and is structurally attached to the trailing edge of the forward wing. This arrangement of the two wings forms a truss structure which is inherently resistant to the aerodynamic bending loads generated during flight. This allows for a considerable reduction in the weight of the lifting surfaces. With smaller internal wing structures needed, the Joined Wing may employ thinner wings which are more suitable for supersonic and hypersonic flight, having less induced drag than conventional cantilever winged aircraft. Inherent in the Joined Wing is the capability of the generation of direct lift and side force which enhance the performance parameters.

Outperforming hummingbirds' load-lifting capability with a lightweight hummingbird-like flapping-wing mechanism.

PubMed

Leys, Frederik; Reynaerts, Dominiek; Vandepitte, Dirk

2016-08-15

The stroke-cam flapping mechanism presented in this paper closely mimics the wing motion of a hovering Rufous hummingbird. It is the only lightweight hummingbird-sized flapping mechanism which generates a harmonic wing stroke with both a high flapping frequency and a large stroke amplitude. Experiments on a lightweight prototype of this stroke-cam mechanism on a 50 mm-long wing demonstrate that a harmonic stroke motion is generated with a peak-to-peak stroke amplitude of 175° at a flapping frequency of 40 Hz. It generated a mass lifting capability of 5.1 g, which is largely sufficient to lift the prototype's mass of 3.39 g and larger than the mass-lifting capability of a Rufous hummingbird. The motor mass of a hummingbird-like robot which drives the stroke-cam mechanism is considerably larger (about five times) than the muscle mass of a hummingbird with comparable load-lifting capability. This paper presents a flapping wing nano aerial vehicle which is designed to possess the same lift- and thrust-generating principles of the Rufous hummingbird. The application is indoor flight. We give an overview of the wing kinematics and some specifications which should be met to develop an artificial wing, and also describe the applications of these in the mechanism which has been developed in this work. © 2016. Published by The Company of Biologists Ltd.

Inspiration for wing design: how forelimb specialization enables active flight in modern vertebrates.

PubMed

Chin, Diana D; Matloff, Laura Y; Stowers, Amanda Kay; Tucci, Emily R; Lentink, David

2017-06-01

Harnessing flight strategies refined by millions of years of evolution can help expedite the design of more efficient, manoeuvrable and robust flying robots. This review synthesizes recent advances and highlights remaining gaps in our understanding of how bird and bat wing adaptations enable effective flight. Included in this discussion is an evaluation of how current robotic analogues measure up to their biological sources of inspiration. Studies of vertebrate wings have revealed skeletal systems well suited for enduring the loads required during flight, but the mechanisms that drive coordinated motions between bones and connected integuments remain ill-described. Similarly, vertebrate flight muscles have adapted to sustain increased wing loading, but a lack of in vivo studies limits our understanding of specific muscular functions. Forelimb adaptations diverge at the integument level, but both bird feathers and bat membranes yield aerodynamic surfaces with a level of robustness unparalleled by engineered wings. These morphological adaptations enable a diverse range of kinematics tuned for different flight speeds and manoeuvres. By integrating vertebrate flight specializations-particularly those that enable greater robustness and adaptability-into the design and control of robotic wings, engineers can begin narrowing the wide margin that currently exists between flying robots and vertebrates. In turn, these robotic wings can help biologists create experiments that would be impossible in vivo . © 2017 The Author(s).

Outperforming hummingbirds’ load-lifting capability with a lightweight hummingbird-like flapping-wing mechanism

PubMed Central

Reynaerts, Dominiek; Vandepitte, Dirk

2016-01-01

ABSTRACT The stroke-cam flapping mechanism presented in this paper closely mimics the wing motion of a hovering Rufous hummingbird. It is the only lightweight hummingbird-sized flapping mechanism which generates a harmonic wing stroke with both a high flapping frequency and a large stroke amplitude. Experiments on a lightweight prototype of this stroke-cam mechanism on a 50 mm-long wing demonstrate that a harmonic stroke motion is generated with a peak-to-peak stroke amplitude of 175° at a flapping frequency of 40 Hz. It generated a mass lifting capability of 5.1 g, which is largely sufficient to lift the prototype's mass of 3.39 g and larger than the mass-lifting capability of a Rufous hummingbird. The motor mass of a hummingbird-like robot which drives the stroke-cam mechanism is considerably larger (about five times) than the muscle mass of a hummingbird with comparable load-lifting capability. This paper presents a flapping wing nano aerial vehicle which is designed to possess the same lift- and thrust-generating principles of the Rufous hummingbird. The application is indoor flight. We give an overview of the wing kinematics and some specifications which should be met to develop an artificial wing, and also describe the applications of these in the mechanism which has been developed in this work. PMID:27444790

Aircraft

DOEpatents

Hibbs, Bart D.; Lissaman, Peter B. S.; Morgan, Walter R.; Radkey, Robert L.

1998-01-01

This disclosure provides a solar rechargeable aircraft that is inexpensive to produce, is steerable, and can remain airborne almost indefinitely. The preferred aircraft is a span-loaded flying wing, having no fuselage or rudder. Travelling at relatively slow speeds, and having a two-hundred foot wingspan that mounts photovoltaic cells on most all of the wing's top surface, the aircraft uses only differential thrust of its eight propellers to turn. Each of five sections of the wing has one or more engines and photovoltaic arrays, and produces its own lift independent of the other sections, to avoid loading them. Five two-sided photovoltaic arrays, in all, are mounted on the wing, and receive photovoltaic energy both incident on top of the wing, and which is incident also from below, through a bottom, transparent surface. The aircraft is capable of a top speed of about ninety miles per hour, which enables the aircraft to attain and can continuously maintain altitudes of up to sixty-five thousand feet. Regenerative fuel cells in the wing store excess electricity for use at night, such that the aircraft can sustain its elevation indefinitely. A main spar of the wing doubles as a pressure vessel that houses hydrogen and oxygen gasses for use in the regenerative fuel cell. The aircraft has a wide variety of applications, which include weather monitoring and atmospheric testing, communications, surveillance, and other applications as well.

Automated Wing Twist And Bending Measurements Under Aerodynamic Load

NASA Technical Reports Server (NTRS)

Burner, A. W.; Martinson, S. D.

1996-01-01

An automated system to measure the change in wing twist and bending under aerodynamic load in a wind tunnel is described. The basic instrumentation consists of a single CCD video camera and a frame grabber interfaced to a computer. The technique is based upon a single view photogrammetric determination of two dimensional coordinates of wing targets with a fixed (and known) third dimensional coordinate, namely the spanwise location. The measurement technique has been used successfully at the National Transonic Facility, the Transonic Dynamics Tunnel, and the Unitary Plan Wind Tunnel at NASA Langley Research Center. The advantages and limitations (including targeting) of the technique are discussed. A major consideration in the development was that use of the technique must not appreciably reduce wind tunnel productivity.

Finite Element Based HWB Centerbody Structural Optimization and Weight Prediction

NASA Technical Reports Server (NTRS)

Gern, Frank H.

2012-01-01

This paper describes a scalable structural model suitable for Hybrid Wing Body (HWB) centerbody analysis and optimization. The geometry of the centerbody and primary wing structure is based on a Vehicle Sketch Pad (VSP) surface model of the aircraft and a FLOPS compatible parameterization of the centerbody. Structural analysis, optimization, and weight calculation are based on a Nastran finite element model of the primary HWB structural components, featuring centerbody, mid section, and outboard wing. Different centerbody designs like single bay or multi-bay options are analyzed and weight calculations are compared to current FLOPS results. For proper structural sizing and weight estimation, internal pressure and maneuver flight loads are applied. Results are presented for aerodynamic loads, deformations, and centerbody weight.

Damage Arresting Composites for Shaped Vehicles

NASA Technical Reports Server (NTRS)

Velicki, Alex

2009-01-01

This report describes the development of a novel structural solution that addresses the demanding fuselage loading requirements for the Hybrid Wing or Blended Wing Body configurations that are described in NASA NRA subtopic A2A.3, "Materials and Structures for Wing Components and Non-Circular Fuselage." The phase I portion of this task includes a comprehensive finite element model-based structural sizing exercise performed using the BWB airplane configuration to generate internal loads and fuselage panel weights for an advanced Pultruded Rod Stitched Efficient Unitized Structure (PRSEUS) structural concept. An accompanying element-level test program is also described which substantiates the analytical results and calculation methods used in the trade study. The phase II plan for the continuation of this research is also included herein.

Effect of outboard vertical-fin position and orientation on the low-speed aerodynamic performance of highly swept wings. [supersonic cruise aircraft research

NASA Technical Reports Server (NTRS)

Johnson, V. S.; Coe, P. L., Jr.

1979-01-01

A theoretical study was conducted to determine the potential low-speed performance improvements which can be achieved by altering the position and orientation of the outboard vertical fins of low-aspect-ratio highly swept wings. Results show that the magnitude of the performance improvements is solely a function of the span-load distribution. Both the vertical-fin-chordwise position and toe angle provided effective means for adjusting the overall span-load distribution.

Heat transfer distributions induced by elevon deflections on swept wings and adjacent surfaces at Mach 6

NASA Technical Reports Server (NTRS)

Johnson, C. B.; Kaufman, L. G., II

1978-01-01

Surface heat transfer distributions are presented for swept wing semispan models having trailing edge elevon ramp angles of 0, 10, 20, and 30 degrees. The wing sweepback angles are 0, 50, and 70 degrees. The models have attachable cylindrical and flat plate center bodies and various attachable wing-tip fins. The data, obtained for a 0 degree angle of attack, a free stream Mach number of 6, and a wing root chord Reynolds number of about 17,000,000, reveal considerably larger regions of elevon induced thermal loads on adjacent surfaces than would be suggested by fully attached flow analyses.

Effect of wing flexibility on the experimental aerodynamic characteristics of an oblique wing

NASA Technical Reports Server (NTRS)

Hopkins, E. J.; Yee, S. C.

1977-01-01

A solid-aluminum oblique wing was designed to deflect considerably under load so as to relieve the asymmetric spanwise stalling that is characteristic of this type of wing by creating washout on the trailing wing panel and washin on the leading wing panel. Experimental forces, and pitching, rolling and yawing moments were measured with the wing mounted on a body of revolution. In order to vary the dynamic pressure, measurements were made at several unit Reynolds numbers, and at Mach numbers. The wing was investigated when unswept (at subsonic Mach numbers only) and when swept 45 deg, 50 deg, and 60 deg. The wing was straight tapered in planform, had an aspect ratio of 7.9 (based on the unswept span), and a profile with a maximum thickness of 4 percent chord. The results substantiate the concept that an oblique wing designed with the proper amount of flexibility self relieves itself of asymmetric spanwise stalling and the associated nonlinear moment curves.

Development of a Non-autorotative Airplane Capable of Steep Landing

NASA Technical Reports Server (NTRS)

Schmidt, Wilhelm

1931-01-01

In the following we develop a non-autorotating monoplane wing. The conditions imposed on such a wing, aside from its freedom from autorotation,with respect to its polars and its construction, are taken into account as far as possible. It is indicated that the autorotation characteristics of a wing are dependent upon the speed of air flow as well as on the angle of yaw. This report postulates the knowledge of the behavior of certain conventional wings of different chords and cambers with respect to their air loads at large angles of attack.

Selected advanced aerodynamics and active controls technology concepts development on a derivative B-747 aircraft

NASA Technical Reports Server (NTRS)

1983-01-01

Analytical design and wind tunnel test evaluations covering the feasibility of applying wing tip extensions, winglets, and active control wing had alleviation to the model B747 are described. Aerodynamic improvement offered by wing tip extension and winglet individually, and the combined aerodynamic and weight improvements when wing load alleviation is combined with the tip extension or the winglet are evaluated. Results are presented in the form of incremental effects on weight mission range, fuel usage, cost, and airline operating economics.

Axial propulsion with flapping and rotating wings, a comparison of potential efficiency.

PubMed

Kroninger, Christopher M

2018-04-18

Interest in biological locomotion and what advantages the principles governing it might offer in the design of manmade vehicles prompts one to consider the power requirements of flapping relative to rotary propulsion. The amount of work performed on the fluid surrounding a thrusting surface (wing or blade) is reflected in the kinetic energy of the wake. Consideration of the energy in the wake is sufficient to define absolute minimum limitations on the power requirement to generate a particular thrust. This work applies wake solutions to compare the minimum inviscid propulsive power requirement of wings flapping and in rotation at wing loading conditions reflective of hover through a state of lightly-loaded cruise. It is demonstrated that hovering flapping flight is less efficient than rotary wing propulsion except for the most extreme flap amplitude strokes ([Formula: see text]   >  160°) if operating at large wake wavelength. In cruise, a larger range of flap amplitude kinematics ([Formula: see text]  >  140°) can be aerodynamically more energy efficient for wake wavelengths reflective of biological propulsion. These results imply, based on the observed wing kinematics of continuous steady flight, that flapping propulsion in animals is unlikely to be more efficient than rotary propulsion.

Rolling Moments Due to Rolling and Yaw for Four Wing Models in Rotation

NASA Technical Reports Server (NTRS)

Knight, Montgomery; Wenzinger, Carl J

1932-01-01

This report presents the results of a series of autorotation and torque tests on four different rotating wing systems at various rates of roll and at several angles of yaw. The investigation covered an angle of attack range up to 90 degrees and angles of yaw of 0 degree, 5 degrees, 10 degrees, and 20 degrees. The tests were made in a 5-foot, closed-throat atmospheric wind tunnel. The object of the tests was primarily to determine the effects of various angles of yaw on the rolling moments of the rotating wings up to large angles of attack. It was found that at angles of attack above that of maximum lift the rolling moments on the wings due to yaw (or side slip) from 5 degrees to 20 degrees were roughly of the same magnitude as those due to rolling. There was a wide variation in magnitude of the rolling moment due to yaw angle. The rates and ranges of stable autorotation for the monoplane models were considerably increased by yaw, whereas for an unstaggered biplane they were little affected. The immediate cause of the rolling moment due to yaw is apparently the building up of large loads on the forward wing tip and the reduction of loads on the rearward wing tip.