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Showing papers in "Journal of Aircraft in 1986"


Journal ArticleDOI
TL;DR: Aeroelastic tailoring technology is reviewed with reference to the historical background, underlying theory, current trends, and specific applications as mentioned in this paper, and the future of aero-linear tailoring and the development of an automated strength-aero-elastic design tool under the Automated Strength-AeroELastic Design program are examined.
Abstract: Aeroelastic tailoring technology is reviewed with reference to the historical background, the underlying theory, current trends, and specific applications. The specific application discussed include the Transonic Aircraft Technology program, an Advanced Design Composite Aircraft, the Wing/Inlet Advanced Development program, and the forward-swept wing. Finally, the future of aeroelastic tailoring and the development of an aeroelastic tailoring analysis and design tool under the Automated Strength-Aeroelastic Design program are examined.

364 citations


Journal ArticleDOI
TL;DR: The joined wing is a new type of aircraft configuration which employs tandem wings arranged to form diamond shapes in plan view and front view as mentioned in this paper, which provides the following advantages over a comparable wing-plus-tail system; lighter weight and higher stiffness, higher span-efficiency factor, higher trimmed maximum lift coefficient, lower wave drag, plus built-in direct lift and direct sideforce control capability.
Abstract: The joined wing is a new type of aircraft configuration which employs tandem wings arranged to form diamond shapes in plan view and front view. Wind-tunnel tests and finite-element structural analyses have shown that the joined wing provides the following advantages over a comparable wing-plus-tail system; lighter weight and higher stiffness, higher span-efficiency factor, higher trimmed maximum lift coefficient, lower wave drag, plus built-in direct lift and direct sideforce control capability. A summary is given of research performed on the joined wing. Calculated joined wing weights are correlated with geometric parameters to provide simple weight estimation methods. The results of low-speed and transonic wind-tunnel tests are summarized, and guidelines for design of joined-wing aircraft are given. Some example joined-wing designs are presented and related configurations having connected wings are reviewed.

275 citations


Journal ArticleDOI
TL;DR: An approximate analysis of atmospheric effects on wake vortex motion and decay is presented in this article, where the effects of density stratification, turbulence, and Reynolds number are combined in a single model so that the relative importance of different parameters can be determined.
Abstract: An approximate analysis of atmospheric effects on wake vortex motion and decay is presented. The effects of density stratification, turbulence, and Reynolds number are combined in a single model so that the relative importance of different parameters can be determined. Predicted wake motion is shown to be in good agreement with limited data from both ground facility and flight test measurements taken under low turbulence conditions. Wake decay was found to depend strongly on both density stratification and turbulence. For typical levels of turbulence, wake decay was found to result from the 'Crow instability' except under strongly stratified conditions.

269 citations


Journal ArticleDOI
TL;DR: In this paper, a new equilvalent plate analysis formulation is described which is capable of modeling aircraft wing structures with a general planform such as cranked wing boxes, and a direct method to interface this structural analysis procedure with aerodynamic programs for use in aeroelastic calculations is described.
Abstract: A new equilvalent plate analysis formulation is described which is capable of modeling aircraft wing structures with a general planform such as cranked wing boxes. Multiple trapezoidal segments are used to represent such planforms. A Ritz solution technique is used in conjunction with global displacement functions which encompass all the segments. This Ritz solution procedure is implemented efficiently into a computer program so that it can be used by rigorous optimization algorithms for application in early preliminary design. A direct method to interface this structural analysis procedure with aerodynamic programs for use in aeroelastic calculations is described. This equivalent plate analysis procedure is used to calculate the static deflections and stresses and vibration frequencies and modes of an example wing configuration. The numerical results are compared with results from a finite element model of the same configuration to illustrate typical levels of accuracy and computation times resulting from use of this procedure.

129 citations


Journal ArticleDOI
Ilan Kroo1
TL;DR: In this paper, a generalized version of Munk's stagger theorem is used in a rapid, approximate calculation of optimal lift distributions and installed efficiency for single-board and up-outboard designs.
Abstract: An analysis of propeller-win g combinations in inviscid incompressible flow reveals some of the fundamental interactions which affect the performance of an installed propulsion system. A generalized version of Munk's stagger theorem is used in a rapid, approximate calculation of optimal lift distributions and installed efficiency. Results indicate that the distribution of lift over the wing which maximizes overall efficiency differs markedly from elliptic loading. Swirl recovery by the wing leads to increments in net propeller efficiency of 6% in example cases. The maximum installed efficiency is computed for single-rotation (up-inboard and up-outboard designs) and counter-rotating systems. Results suggest that some of the performance advantages attributed to counterrotation may be less dramatic for well-integrated wing-propeller designs than for isolated systems. Nomenclature An = amplitude of nth harmonic of wing lift & = wing aspect ratio b = wingspan c =wing chord CL =wing lift coefficient CT = thrust coefficient, =ir2T/pw2R4 D =drag IU,IW = definite integrals, Eqs. (10) and (11), respectively / = advance ratio, = irU00/uR I = section lift L =lift N = number of blades Obj = objective function Q = propeller torque R = propeller radius r - radial coordinate T = thrust Uw = freestream velocity u = axial perturbation velocity V = induced velocity vt = tangential induced velocity w = induced downwash x = stream wise coordinate y = spanwise coordinate .Vprop = spanwise location of propeller F = circulation 6 =dimensionless spanwise coordinate K =Goldstein's radial velocity correction X =Lagrange multiplier p = density Subscripts int = interference component = wing w = propeller

96 citations


Journal ArticleDOI
TL;DR: In this paper, the effect of ice accretion on airfoil sections is analyzed and experimentally measured using simulation techniques for aerodynamic testing and compared to data with actual ice accretions.
Abstract: Methods of analyzing and experimentally measuring the effect of ice accretion on airfoil sections are presented. Empirical and analytical methods for predicting airfoil performance degradation due to ice are discussed. Ice simulation techniques for aerodynamic testing are presented and compared to data with actual ice accretions. The results show that simulation techniques to imitate the effect of ice on airfoil performance work well in most cases. Comparisons between predicted and measured airfoil performance with ice accretions are presented. For rime ice cases, the predictions compared well with experiments; but for glaze ice, a need for improved methods are seen.

89 citations


Journal ArticleDOI
TL;DR: In this paper, a mathematical model of a downburst for real-time flight simulations of takeoffs and landings in low-altitude severe wind shears due to downbursts or microbursts was developed.
Abstract: A mathematical model of a downburst for real-time flight simulations of takeoffs and landings in low-altitude severe wind shears due to downbursts or microbursts was developed. The downburst mature stage was idealized to resemble the circulatory wind-flow patterns noted in meteorological downburst data from the Joint Airport Weather Studies (JAWS) Project. The idealization produced a three-dimensio nal axisymmetric circulatory flowfield similar to that around a horizontal smoke ring or ring vortex at an appropriate height above the ground. The flowfield around a ring vortex has a stream function expressed in Sir Horace Lamb's classical textbook Hydrodynamics in terms of the complete elliptic integrals combination of [F(k)—E(k)]; which is approximated herein by the expression: [0.788fc2/(0.25 + 0.75Vl - A:2)], in the limited range of 0<£ 2 <1 for the modulus k = (r2 —rl)/(r2-rl}^ where rl and r2 denote the least and greatest distances, respectively, of the point P from the ring vortex. "Digital differentiations" of the downburst stream function at the airplane center of gravity yield both the WZ downdraft and the WR radial wind velocity component. The latter is then resolved into the two horizontal components WX and WY for wind speeds along and across the runway, respectively. Occupying 383 words of memory and having an average real-time execution time of 1.3 ms on a Harris H800 digital computer, the present ring-vortex downburst model provides economical simulation of severe wind-shear flow patterns that closely resemble some of the flow patterns noted in meteorological data from the JAWS Project.

88 citations


Journal ArticleDOI
TL;DR: In this article, a reliable and fast transonic wing flow-field analysis program, TWING, has been coupled with a modified quasi-Newton method, unconstrained optimization algorithm, QNMDIF, to create a new design tool.
Abstract: A computationally efficient and versatile technique for use in the design of advanced transonic wing configurations has been developed. A reliable and fast transonic wing flow-field analysis program, TWING, has been coupled with a modified quasi-Newton method, unconstrained optimization algorithm, QNMDIF, to create a new design tool. Fully three-dimensional wing designs utilizing both specified wing pressure distributions and drag-to-lift ration minimization as design objectives are demonstrated. Because of the high computational efficiency of each of the components of the design code, in particular the vectorization of TWING and the high speed of the Cray X-MP vector computer, the computer time required for a typical wing design is reduced by approximately an order of magnitude over previous methods. In the results presented here, this computed wave drag has been used as the quantity to be optimized (minimized) with great success, yielding wing designs with nearly shock-free (zero wave drag) pressure distributions and very reasonable wing section shapes.

80 citations


Journal ArticleDOI
TL;DR: In this paper, a multibladed helicopter rotor in hovering flight is calculated by solving the threedimensional Euler equations in a rotating coordinate system on body-conforming curvilinear grids around the blades.
Abstract: Aerodynamic loads on a multibladed helicopter rotor in hovering flight are calculated by solving the threedimensional Euler equations in a rotating coordinate system on body-conforming curvilinear grids around the blades Euler equations are recast in the absolute flow variables so that the absolute flow in the far field is uniform but the relative flow is nonuniform Equations are solved for the absolute flow variables employing Jameson's finite-volume explicit Runge-Kutta time-stepping scheme Rotor-wake effects are modeled in the form of a correction applied to the geometric angle of attack along the blades This correction is obtained by computing the local induced downwash with a free-wake analysis program The calculations are performed on a CRAY X/MP-48 for a model helicopter rotor in hover at various collective pitch angles The results compared with experimental data

70 citations



Journal ArticleDOI
TL;DR: In this article, a solution procedure is described for the numerical solution of inviscid rotational flow past fixed and rotor wing configurations, where derivatives along the spanwise direction are lagged by one time step, while all the other terms are treated in a fully implicit manner.
Abstract: A solution procedure is described for the numerical solution of inviscid rotational flow past fixed and rotor wing configurations. This procedure solves the three-dimensional Euler equations in a body-fitted coordinate system and strong conservation form. The derivatives along the spanwise direction are lagged by one time step, while all the other terms are treated in a fully implicit manner. This leads to a semi-implicit scheme that requires two block tridiagonal matrix inversions and one residual evaluation per point at every time step. This procedure also requires the flow variables to be stored at only one time level. A number of fixed wing and rotor wing calculations are presented to demonstrate the efficiency and accuracy of this procedure.


Journal ArticleDOI
TL;DR: In this paper, the problem of safe microburst wind shear encounter during the approach and climb-out flight phases is addressed using flight path optimization, and the best control strategies involved responding to airspeed loss in an unconventional manner by raising the nose to maintain lift.
Abstract: The problem of safe microburst wind shear encounter during the approach and climb-out flight phases is addressed using flight path optimization. The purpose is to investigate the physical limits of safe penetration and to determine control strategies that take full advantage of those limits. Optimal trajectories for both jet transport and general aviation aircraft were computed for encounters with idealized and actual microburst profiles. The results demonstrate that limits on control system design rather than on the aircraft's physical performance may be the deciding factor in an aircraft's capability for safe passage through a wide class of microbursts. The best control strategies involved responding to airspeed loss in an unconventional manner by raising the nose to maintain lift.

Journal ArticleDOI
TL;DR: In this article, an idealized aeroelastic tailoring model is developed to assess the effects of significant changes in directional stiffness orientation upon the flutter and divergence behavior of swept and unswept wings.
Abstract: An idealized aeroelastic tailoring model is developed to assess the effects of significant changes in directional stiffness orientation upon the flutter and divergence behavior of swept and unswept wings. A nondimensional stiffness cross-coupling parameter is used to illustrate the potentially strong influence of stiffness cross-coupling, commonly present in aeroelastically tailored structures, to increase flutter and divergence speeds. Conflicting requirements for flutter and divergence enhancement are indicated. Aeroelastic tailoring for flutter enhancement appears to be less effective when the wing is moderately swept back. However, by combining directional stiffness orientation with inertia balancing, flutter and divergence-free, aft-swept, high-aspect-ratio surfaces are shown to be theoretically possible.

Journal ArticleDOI
TL;DR: In this paper, the support interference associated with rotary rigs used in coning experiments is of a different type, being stationary in nature rather than unsteady, with the coning motion inducing a displacement of the vortex wake similar to that caused by side slip in a static test.
Abstract: Dynamic test results of aircraft models at high angles of attack are analyzed with regard to the support interference effects. Single-degree-of-freedom oscillatory tests in pitch or yaw of aircraft configurations are subject to the same type of support interference through the near-wake recirculatory region as is experienced by slender bodies of revolution. Consequently, the measurements can be corrected for support interference using the same methodology. The support interference associated with rotary rigs used in coning experiments is of a different type, being stationary in nature rather than unsteady, with the coning motion inducing a displacement of the vortex wake similar to that caused by side slip in a static test. Performing static tests at varying incidence and side-slip angles with two alternate supports can provide the information needed to correct coning experiments for support interference.

Journal ArticleDOI
Abstract: The time-dependent flow around ah ogive cylinder undergoing large-amplitude, harmonic pitching motion was investigated using flow visualization techniques. The slender body of revolution was towed in an 18 m water channel at Reynolds numbers up to 1.2xl0. Fluorescent dyes were either introduced uniformly from the body's porous surface or placed as horizontal sheets in the slightly stratified tank prior to a run. The dyes were excited with a sheet of laser light projected in the desired plane to mark the flow in the separation region around the body, the flow in the wake, and the potential flow away from the body. The separation process and the geometry of the leeward vortices were studied under different reduced frequencies and Reynolds numbers. The unsteady separation phenomenon is found to be significantly different from the separation on a body in steady flight. Two distinct separation regions, the forebody and aftbody vortex pairs, evolve in the unsteady case.

Journal ArticleDOI
TL;DR: In this paper, the aerodynamic and aero-elastic response of the B-1 wing is studied using the transonic unsteady code ATRAN3S. The results show that the wing is stable at low-sweep angle for the calculation at the Mach number at which there is a shock wave.
Abstract: The flow over the B-1 wing is studied computationally, including the aeroelastic response of the wing. Computed results are compared with results from wind tunnel and flight tests for both low- and high-sweep cases, at 25.0 and 67.5 deg, respectively, for selected transonic Mach numbers. The aerodynamic and aeroelastic computations are made by using the transonic unsteady code ATRAN3S. Steady aerodynamic computations compare well with wind tunnel results for the 25.0 deg sweep case and also for small angles of attack at 67.5 deg sweep case. The aeroelastic response results show that the wing is stable at the low-sweep angle for the calculation at the Mach number at which there is a shock wave. In the higher-sweep case, for the higher angle of attack at which oscillations were observed in the flight and wind tunnel tests, the calculations do not show any shock waves. Their absence lends support to the hypothesis that the observed oscillations are due to the presence of leading-edge separation vortices and not to shock wave motion, as was previously proposed.

Journal ArticleDOI
TL;DR: Freymuth et al. as mentioned in this paper showed that smoke-generating liquid was released along the leading edge of the wing, marked by the left arrow on top of the first frame.
Abstract: P. Freymuth,* F. Finaish,t and W. BankJ University of Colorado, Boulder, Colorado first frame shown is ^=47/64 s, the time between frame shown is At = 4/64 s for all sequences, and the frame height corresponds to an actual height of h= 16A cm. Smoke-generating liquid was released along the leading edge of the wing, marked by the left arrow on top of the first frame. The trailing edge is covered with smoke and its location is marked by the right arrow at top of the first frame. The tip location can be seen as a horizontal boundary of illumination


Journal ArticleDOI
TL;DR: In this paper, acoustic near field data were collected with model single and twin jet nozzles to determine if they produce higher acoustic loading than do a single nozzle, which was spurred by structural failure of the B-1 exhaust nozzle external flaps and similar damage on the F-15.
Abstract: Acoustic near field data were collected with model single and twin jet nozzles to determine if closely spaced nozzles produce higher acoustic loading than do single nozzles. The tests were spurred by structural failure of the B-1 exhaust nozzle external flaps and similar damage on the F-15. The test was performed using two 5/8 in. ID pipes machined and placed side-by-side to mimic B-1 nozzles. A microphone mounted on the internozzle fairing measured acoustic levels near the nozzle exit plane. The nozzles oscillated significantly more than did a single nozzle over a wide range of nozzle pressure ratios. Acoustic levels in the dual jets exceeded single jet noise by as much as 20 dB, making acoustic resonance a definite candidate for structural damage in the twin jet configuration.

Journal ArticleDOI
TL;DR: Spanwise blowing over the wing and canard of a close-coupled-canard, 60-deg delta fighter-aircraft configuration was investigated experimentally in low-speed flow at angles of attack up to 60 deg and yaw angles of up to 36 deg as mentioned in this paper.
Abstract: Spanwise blowing (SWB) over the wing and canard of a close-coupled-canard, 60-deg delta fighter-aircraft configuration was investigated experimentally in low-speed flow at angles of attack up to 60 deg and yaw angles of up to 36 deg. Significant improvement in lift-curve slope, maximum lift, drag polar, and lateral/directional stability was found, enlarging the usable flight envelope beyond its previous low-speed/maximum-lift limit. It was shown that SWB can achieve the same lift augmentation produced by a canard, without the drag penalty. Contrary to previous experience with 60-deg swept wings, the efficiency of the lift augmentation by SWB was relatively high and was found to increase with increasing jet-momentum coefficient on the close-coupled-canard configuration. Interesting and promising possibilities of obtaining much higher efficiencies with swirling or multiple nonaligned jets were indicated.

Journal ArticleDOI
Ali R. Ahmadi1
TL;DR: In this article, an experimental investigation has been conducted for the blade-vortex interaction (BVI) of a rotor at normal incidence, where the vortex is generally parallel to the rotor axis.
Abstract: An experimental investigation has been conducted for the blade-vortex interaction (BVI) of a rotor at normal incidence, where the vortex is generally parallel to the rotor axis. Tip Mach number, radial BVI station, and free stream velocity were varied during measurements of fluctuating blade pressures, far field sound pressure levels and directivity, incident vortex velocity field, and blade-vortex interaction angles. The experimental setup is representative of the chopping of helicopter main rotor tip vortices by the tail rotor. This interaction is found to generate impulsive noise which radiates primarily ahead of the blade.


Journal ArticleDOI
TL;DR: In this article, the corrected section drag coefficient and effective drag coefficient were calculated for the half-chord chord of the airfoil chord of a single chord in a single-antenna chord.
Abstract: Nomenclature c — airfoil chord cd = corrected section drag coefficient cdg = effective drag coefficient rake = d drag coefficient from wake survey ct = section lift coefficient cm = pitching moment about the half-chord C^ = blowing momentum coefficient, =m-Vj/(qco'S) h = slot height m =jet mass flow A^ = freestream Mach number qx = freestream dynamic pressure s = chordwise slot location S =wing reference area V = velocity a. = angle of attack

Journal ArticleDOI
TL;DR: The objectives of this study were to establish a historical view of the improvement in the state of the art of aircraft-surface dynamic simulation techniques, and to recognize the individuals and organizations that have played a prominent roll in advancing the state- of-the- art.
Abstract: Introduction D the design of an aircraft heavy emphasis is placed on flight-induced motions and loads. However, groundbased operations produce an environment that can generate significant aircraft dynamics uncomfortable to passengers or damaging to the cargo. In addition, high vertical accelerations in the cockpit represent a potential disorientation problem for the pilot, which may cause landing or takeoff accidents. Perhaps more important, the aircraft structure can be subjected to large local deformations leading to stress failure, or the gears could experience loads beyond their design limits. For normal commercial aircraft and airports, ground loads (except for landing impact) should be of secondary concern. But aircrafts such as crop dusters and small private planes, which often operate from unimproved fields, experience a harsh environment during ground operations. Of greatest importance is the growing need of military aircraft to be operational from austere airfields. Most current U.S. Air Force (USAF) aircraft operate on rigid, smooth, paved Main Operating Base (MOB) surfaces. In all of the recent wars that the United States has been involved in, it has enjoyed air superiority, and its airbases were generally well protected and operational under normal procedures. Future conflicts may, however, be fought from MOBs vulnerable to enemy attack, and MOB surface damage or MOB denial is anticipated. Therefore, the USAF is placing greater emphasis on aircraft-surface operations, particularly on bomb damaged repaired (BDR) surfaces, soil, and other emergency surfaces. One effort to meet this challenge, which is presently under way in the United States, is to define the rough surface capabilities of mainline fighters and cargo planes. The USAF is establishing these capabilities through a program called HAVE BOUNCE under which: 1) Simulations are prepared for each aircraft on BDR runways. 2) Aircraft component weaknesses are identified through simulation. 3) Simulations are validated with test data. 4) Operational limitations are developed. In the past 5-10 years a substantial number of computer simulations have been developed to predict aircraft-surface interaction. Many of these programs have been written by USAF personnel or have been contracted to various organizations by the USAF. Others have been developed by aircraft companies to meet their own needs, or by individuals at universities, or in foreign countries (most notably in NATO countries). The objectives of this study were to review the literature concerning aircraft-surface dynamic simulation techniques: 1) to establish a historical view of the improvement in the state of the art, 2) to recognize the individuals and organizations that have played a prominent roll in advancing the state of the art, 3) to develop a knowledge base of physical phenomena that have been simulated, 4) to identify mathematical techniques that have been used, 5) to classify the simulations according to their general purpose, complexity, and accuracy, and 6) to suggest areas in which simulation techniques could be improved, and test could be run to validate the simulations. This report contains a brief summary of the computer programs written to predict the dynamic displacements and forces resulting from nonflight aircraft operations. The capabilities of each program along with their limitations and numerical techniques are cited.

Journal ArticleDOI
TL;DR: In this paper, an analysis of the correlation between the UHF band radar data obtained and TV images of lightning strikes indicates that, with a known aircraft position relative to the radar, the lightning channel motion can be adequately interpreted on the basis of radar echo evolution.
Abstract: The NASA Storm Hazards program was dedicated during the 1984 storm season to a study of lightning strikes on an instrumented F-106B aircraft, during penetrations of thunderstorms at altitudes lower than the 6-8 km center of lightning flash density. These altitudes coincide with the negative charge region of thunderstorms. An analysis of the correlation between the UHF band radar data obtained and TV images of lightning strikes indicates that, with a known aircraft position relative to the radar, the lightning channel motion can be adequately interpreted on the basis of radar echo evolution.

Journal ArticleDOI
TL;DR: In this article, the use of thermal means to control drag under turbulent boundary layer conditions is examined and the authors show that partial wall heating of the forebody can produce almost the same order of total drag reduction as the full body heating case.
Abstract: The use of thermal means to control drag under turbulent boundary layer conditions is examined Numerical calculations are presented for both skin friction and (unseparated) pressure drag for turbulent boundary-layer flows over a fuselage-like body with wall heat transfer In addition, thermal control of separation on a bluff body is investigated It is shown that a total drag reduction of up to 20 percent can be achieved for wall heating with a wall-to-total-freestream temperature ratio of 2 For streamlined slender bodies, partial wall heating of the forebody can produce almost the same order of total drag reduction as the full body heating case For bluff bodies, the separation delay from partial wall cooling of the afterbody is approximately the same as for the fully cooled body

Journal ArticleDOI
TL;DR: In this article, the lagging motion of each helicopter blade is assumed to be of equal amplitude and equally apportioned phase, thus allowing a simplified analytical method to calculate the ground resonance instability of a helicopter model with nonlinear dampers in both the landing gear and blades.
Abstract: The lagging motion of each helicopter blade is assumed to be of equal amplitude and equally apportioned phase, thus allowing a simplified analytical method to be used to calculate the ground resonance instability of a helicopter model with nonlinear dampers in both the landing gear and blades. The geometrical nonlinearities of the blade lag motion and the influence of initial disturbances on ground resonance instability are also discussed. Finally an experiment is carried out using a helicopter scale model. The experimental data agree well with analysis.


Journal ArticleDOI
TL;DR: In this article, a small canard wing was installed in front of a delta wing mounted on a free-to-roll sting balance in a low-speed wind tunnel, and the leading edge vortices originating from the canard enhanced the self-induced roll oscillations at test conditions for which the basic delta wing would otherwise have been stable.
Abstract: A small canard wing was installed in front of a delta wing mounted on a free-to-roll sting balance in a lowspeed wind tunnel. The leading-edge vortices originating from the canard enhanced the self-induced roll oscillations at test conditions for which the basic delta wing would otherwise have been stable. Time-dependent roll angle and normal and side force data recorded during these oscillations are presented along with their phase relations. It was found that, with this combined canard/wing configuration, the range of the angle of attack at which self-induced oscillations occurred was extended up to an angle of attack of about 45 deg.