ALBERT

Description: The flows around highly sweptback wings and bodies of revolution at high angle of attack are described, and inviscid model approximations and mathematical formulation of the problem are given to steady and unsteady incompressible flows. A general presentation of the methods of solution is given, with emphasis on current computational techniques. Detailed descriptions of the nonlinear vortex-lattice and vortex-panel techniques are presented to show how the boundary conditions are enforced using iteration. Typical numerical results are compared with the available experimental data.

Description: A brief review is presented of various problems which are confronted in the development of an unsteady finite difference potential code. This review is conducted mainly in the context of what is done for a typical small disturbance and full potential methods. The issues discussed include choice of equation, linearization and conservation, differencing schemes, and algorithm development. A number of applications including unsteady three-dimensional rotor calculation, are demonstrated.

Description: The computational treatment of unsteady transonic flows is discussed, reviewing the historical development and current techniques. The fundamental physical principles are outlined; the governing equations are introduced; three-dimensional linearized and two-dimensional linear-perturbation theories in frequency domain are described in detail; and consideration is given to frequency-domain FEMs and time-domain finite-difference and integral-equation methods. Extensive graphs and diagrams are included.

Description: This lecture is introductory to the subject of unsteady subsonic and supersonic flows. The primary objective is to present fundamental concepts in order to promote an understanding of the relations between the basic physical problems and their mathematical formulation as well as to establish a common foundation for the more detailed presentations of subsequent lectures in this session. Linearized (small-perturbation) potential flow is emphasized, although needs beyond that limit are indicated. The basic equations, concepts, and procedures common to all the methods are reviewed first, followed by the development, discussion, and status of methods for creating two-dimensional incompressible flow, strip theory, subsonic lifting-surface theory, subsonic/supersonic surface-panel methods, and supersonic lifting-surface theory.

Description: An overview of helicopter aerodynamics technology is presented with emphasis on rotor wake and airloads methodology developed at the United Technologies Research Center (UTRC). The evolution over the past twenty years of various levels of computerized wake geometry models at UTRC, such as undistorted wake, prescribed empirical wake, predicted distorted wake, and generalized wake models for the hover and forward flight regimes, is reviewed. The requirement for accurate wake modeling for flow field and airload prediction is demonstrated by comparisons of theoretical and experimental results. These results include blade pressure distributions predicted from a recently developed procedure for including the rotor wake influence in a full potential flow analysis. Predictions of the interactional aerodynamics of various helicopter components (rotor, fuselage, and tail) are also presented. It is concluded that, with advanced computers and the rapidly progressing computational aerodynamics technology, significant progress toward reliable prediction of helicopter airloads is forseeable in the near future.

Description: Interferometry methods were applied to the investigation of steady and unsteady flows in large scale transonic wind tunnels. Holographic interferometry was demonstrated to provide reliable flow visualization and quantitative results for a number of two-dimensional flows. These conclusions were based on extensive comparisons with results obtained by other means. Data obtained on a NACA 64A010 airfoil with an oscillating flap installed in the Ames 11-foot transonic tunnel are presented. Interferograms were recorded at a free stream Mach number of 0.8, flap frequency of 30 Hertz and chord Reynolds numbers of 6.6 x 10 to the 6th and 12.3 x 10 to the 6th. The interferometric results were reduced to dynamic surface pressures, Mach contours and wake flow profiles. A new interferometry method that is capable of providing real-time interferometry data is also discussed.

Description: Through a series of flights in artificial clouds, ice accretions on the main rotor of a UH-1H helicopter were documented in detail upon landing by silicone-rubber molds for both hover and level flights. Full scale reproductions of typical accretions in hover were fabricated by means of epoxy castings and used for a wind-tunnel test program. Surface static pressure distributions were recorded and used to evaluate lift and pitching moment increments while drag was determined by wake surveys. For comparison, accreted ice shapes are presented for two level flight cases as well as preliminary analytical predictions.

Description: The three-dimensional inviscid DENTON code is used to analyze flow through a radial-inflow turbine rotor. Experimental data from the rotor are compared with analytical results obtained by using the code. The experimental data available for comparison are the radial distributions of circumferentially averaged values of absolute flow angle and total pressure downstream of the rotor exit. The computed rotor-exit flow angles are generally underturned relative to the experimental values, which reflect the boundary-layer separation at the trailing edge and the development of wakes downstream of the rotor. The experimental rotor is designed for a higher-than-optimum work factor of 1.126 resulting in a nonoptimum positive incidence and causing a region of rapid flow adjustment and large velocity gradients. For this experimental rotor, the computed radial distribution of rotor-exit to turbine-inlet total pressure ratios are underpredicted due to the errors in the finite-difference approximations in the regions of rapid flow adjustment, and due to using the relatively coarser grids in the middle of the blade region where the flow passage is highly three-dimensional. Additional results obtained from the three-dimensional inviscid computation are also presented, but without comparison due to the lack of experimental data. These include quasi-secondary velocity vectors on cross-channel surfaces, velocity components on the meridional and blade-to-blade surfaces, and blade surface loading diagrams. Computed results show the evolution of a passage vortex and large streamline deviations from the computational streamwise grid lines. Experience gained from applying the code to a radial turbine geometry is also discussed.

Description: This paper demonstrates the current and future potential of finite-difference methods for solving real rotor problems which now rely largely on empiricism. The demonstration consists of a simple means of combining existing finite-difference, integral, and comprehensive loads codes to predict real transonic rotor flows. These computations are performed for hover and high-advance-ratio flight. Comparisons are made with experimental pressure data.

Description: A computational procedure is presented to simulate transonic unsteady flows and corresponding aeroelasticity of wings at low-supersonic freestreams. The flow is modeled by using the transonic small-perturbation theory. The structural equations of motions are modeled using modal equations of motion directly coupled with aerodynamics. Supersonic freestreams are simulated by properly accounting for the boundary conditions based on pressure waves along the flow characteristics in streamwise planes. The flow equations are solved using the time-accurate, alternating-direction implicit finite-difference scheme. The coupled aeroelastic equations of motion are solved by an integration procedure based on the time-accurate, linear-acceleration method. The flow modeling is verified by comparing calculations with experiments for both steady and unsteady flows at supersonic freestreams. The unsteady computations are made for oscillating wings. Comparisons of computed results with experiments show good agreement. Aeroelastic responses are computed for a rectangular wing at Mach numbers ranging from subtransonic to upper-transonic (supersonic) freestreams. The extension of the transonic dip into the upper transonic regime is illustrated.

Description: The effect of a simulated glaze-ice accretion on the aerodynamic performance of a NACA 0012 airfoil was studied experimentally. Two ice shapes were tested: one from an experimentally measured accretion, and one from an accretion predicted using a computer model given the same icing conditions. Lift, drag, and pitching moment were measured for the airfoil with both smooth and rough ice shapes. The ice shapes caused large lift and drag penalties, primarily due to large separation bubbles. Surface pressure distributions clearly showed the regions of separated flow. The aerodynamic performance of the two shapes compared well at positive, but not negative, angles of attack.

Description: The drag of airfoils in transonic flow can be reduced through the use of a passive venting system that employs a porous plate for part of the airfoil upper surface with a vent chamber underneath the porous plate Attention is given to the results obtained with a wind tunnel model employing such a porous floor system. This passive venting system has been used to extend the length/height value before the onset of high drag-producing closed cavity flow at supersonic speeds.

Description: The wakes of highly loaded compressor blades are generally considered to be turbulent flows. Recent work has suggested that the blade wakes are dominated by a vortex streetlike structure. The experimental evidence supporting the wake vortex structure is reviewed. This structure is shown to redistribute thermal energy within the flowfield. The effect of the wake structure on conventional aerodynamic measurements of compressor performance is noted. A two-dimensional, time-accurate, viscous numerical simulation of the flow exhibits both vortex shedding in the wake and a lower-frequency flow instability that modulates the shedding. The numerical results are shown to agree quite well with the measurement from transonic compressor rotors.

Description: The effects of mean-flow incidence, airfoil camber, and airfoil thickness on the incompressible aerodynamics of an oscillating airfoil are investigated theoretically, developing and applying a first-order FEM based on locally analytical solutions (LASs). Laplace equations are used to describe the steady and unsteady harmonic velocity potentials; a body-fitted computational grid is employed; grid-element solutions for both potentials are determined using a numerical LAS method; and the LASs are then assembled to obtain a complete solution. Results for a series of flat-plate and Joukowski airfoils are presented in extensive graphs and discussed in detail.

Description: A complete mathematical model is formulated to analyze the effects of mean-flow incidence angle on the unsteady aerodynamics of an oscillating airfoil in an incompressible flow field. A velocity potential formulation is utilized. The steady flow is independent of the unsteady flow field but coupled to it through the boundary conditions on the oscillating airfoil. The numerical solution technique for both the steady and unsteady flow fields is based on a locally analytical method. The flow model and solution method are then verified through the excellent correlation obtained with the Theodorsen oscillating-flat-plate and Sears transverse-gust classical solutions. The effects of mean flow incidence on the steady and oscillating airfoil aerodynamics are then investigated.

Description: Computational results are presented for the transitional or turbulent flow about a prolate spheroid, at alpha = 10 deg or 30 deg, correspondingly, using an implicit, approximately factored, partially flux-split algorithm, based on the thin-layer equations. The computed flow field is in good agreement with available experimental data.

Description: Thin-element riblets for aircraft aerodynamic surface turbulent viscous drag reduction are presently found to be as effective as symmetric V-grooves in this role, while possessing a greater range of admissible spacings. The thin-element geometry shows the qualitatively predictable influence of independent riblet height and spacing variations. The evidence for more than one drag-reduction mechanism in thin-element riblets is found to be inconclusive.

Description: A theoretical investigation of the aerodynamics of sharp leading-edge delta wings at supersonic speeds has been conducted. The primary objective of this was to determine the applicability of existing theoretical methods to predict wing leading-edge separated-flow characteristics at conditions conductive to high-lift supersonic flight. Predicted results from two modified linear-theory methods (LTSTAR and VORCAM) are compared with experimental data. Comparison of the two methods for uncambered wings revealed that the LTSTAR code is in much better agreement with experimentally measured vortex strength, vortex position, and total lifting characteristics than the VORCAM code. Selected analysis was also performed with an Euler code, SWINT. The results of this study indicated that the SWINT code was not well suited to the analysis of wings with separated flow at high lift and low supersonic speeds.

Description: In the present determination of the free molecule flow drag coefficient for a cylindrical spacecraft flying parallel to its principal axis, the lateral surface effects of thermal motion are explicitly included in terms of the average impact angle of the incident gas momentum vector. Kinetic theory is used to characterize self-shadowing, as well as to obtain an expression for the lateral surface coefficient in terms of the average impact angle of the incident momentum vector and the fractional momentum transfer along the line of impact. It is found that, for a length/diameter ratio of about 5, the lateral surface contribution to the drag coefficient is comparable to that of the front face.

Description: The effect of shoulder radiusing and grooving (longitudinally and circumferentially) the afterbodies of bluff bodies to reduce the base drag at low speeds is investigated experimentally. Shoulder radii as large as 2.75 body diameters are examined. Reynolds number (ReD) based on body diameter varied from 20,000 to 200,000. Results indicate that increasing the shoulder radius to 2.00 body diameters can reduce the drag levels to those of a streamline body having 67 percent greater fineness ratio. For the relatively sharp shoulder case, body drag reductions as large as 50 and 33 percent are obtained using circumferential or longitudinal grooves, respectively.

Description: The novel implicit and unconditionally stable, high resolution Total Variation Diminishing (TVD) scheme whose application to steady state calculations is presently examined is a member of a one-parameter family of implicit, second-order accurate systems developed by Harten (1983) for the computation of weak solutions for one-dimensional hyperbolic conservation laws. The scheme will not generate spurious oscillations for a nonlinear scalar equation and a constant coefficient system. Numerical experiments for a quasi-one-dimensional nozzle problem show that the experimentally determined stability limit correlates exactly with the theoretical stability limit for the nonlinear scalar hyberbolic conservation laws.

Description: The Stalled Airfoil Analysis Program (SAAP) is a computer code for predicting the aerodynamic characteristics of an airfoil up to, and beyond, stall. SAAP is presently evaluated through comparisons with experiments and with two other theoretical methods over an extensive range of airfoils and Reynolds number conditions. SAAP modeled drag more accurately than either of the other methods, and at angles of attack below stall yielded a smoother lift variation with angle of attack.

Description: This paper briefly reviews some national studies and new programs concerning hypersonic flight. The flight environment that will be encountered by this new class of hypersonic vehicles is described, and the fluid-dynamic and chemical phenomena that occur in hypersonic flight are examined. Ground-based facilities are briefly described, and their use in helping to validate the codes is examined.

Description: Different models for inviscid transonic flows are examined. The common assumptions that the flow is isentropic and irrotational are critically evaluated. Entropy and vorticity correction procedures for potential and stream function formulations are presented, together with the details of the treatment of shocks and wakes, and drag and lift calculations. The non-uniqueness problem of the potential formulation is studied using different artificial viscosity forms. Numerical results are compared with Euler solutions.

Description: The dynamics of unsteady transonic small disturbance flows about two-dimensional airfoils is examined, with emphasis on the behavior in the region where the steady state flow is nonunique. It is shown that nonuniqueness results from an extremely long time scale instability which occurs in a finite Mach number and angle of attack range. The similarity scaling rules for the instability are presented and the possibility of similar behvior in the Euler equations is discussed.

Description: An investigation of the aerodynamics of sharp leading-edge delta wings at supersonic speeds has been conducted. The supporting experimental data for this investigation were taken from published force, pressure, and flow-visualization data in which the Mach number normal to the wing leading edge is always less than 1.0. The individual upper- and lower-surface nonlinear characteristics for uncambered delta wings are determined and presented in three charts. The upper-surface data show that both the normal-force coefficient and minimum pressure coefficient increase nonlinearly with a decreasing slope with increasing angle of attack. The lower-surface normal-force coefficient was shown to be independent of Mach number and to increase nonlinearly, with an increasing slope, with increasing angle of attack. These charts are then used to define a wing-design space for sharp leading-edge delta wings.

Description: A wind tunnel investigation was conducted in which independent, steady state aerodynamic forces and moments were measured on a 2.24 m diam. two bladed helicopter rotor and on several different bodies. The mutual interaction effects for variations in velocity, thrust, tip-path-plane angle of attack, body angle of attack, rotor/body position, and body geometry were determined. The results show that the body longitudinal aerodynamic characteristics are significantly affected by the presence of a rotor and hub, and that the hub interference may be a major part of such interaction. The effects of the body on the rotor performance are presented.

Description: The aerodynamic characteristics of airfoils with several flap configurations were studied theoretically and experimentally in environments that simulate a wing immersed in the downwash of a hovering rotor. Special techniques were developed for correcting and validating the wind tunnel data for large blockage effects, and the test results were used to evaluate two modern blockage effects, and the test results were used to evaluate two modern computational aerodynamics codes. The combined computed and measured results show that improved flap and leading-edge configurations can be designed which will achieve large reductions in the downloads of tilt-rotor aircraft, and thereby improve their hover efficiency.

Description: A computational method for designing shock-free, quasi-three-dimensional, transonic, turbomachinery blades is described. Shock-free designs are found by implementing Sobieczky's fictitious gas principle in the analysis of a baseline shape, resulting in an elliptic solution that is incorrect in the supersonic domain. Shock-free designs are obtained by combining the subsonic portion of this solution with a characteristic calculation of the correct supersonic flow using the sonic line data from the fictitious elliptic solution. This provides a new, shock-free blade design. Examples presented include the removal of shocks from two blades in quasi-three-dimensional flow and the development of a series of shock-free two-dimensional stators. The new designs all include modifications to the upper surface of an experimental stator blade developed at NASA Lewis Research Center. While the designs presented here are for inviscid flow, the same concepts have been successfully applied to the shock-free design of airfoils and three-dimensional wings with viscous effects. The extension of the present method to viscous flows is straightforward given a suitable analysis algorithm for the flow.

Description: Low-speed wind tunnel drag force measurements were taken on a laminar flow body of revolution free of support interference. This body was tested at zero incidence in the NASA Langley 13 inch Magnetic Suspension and Balance System (MSBS). The primary objective of these tests was to substantiate the drag force measuring capabilities of the 13 inch MSBS. A secondary objective was to obtain support interference free drag measurements on an axisymmetric body of interest. Both objectives were met. The drag force calibrations and wind-on repeatability data provide a means of assessing the drag force measuring capabilities of the 13 inch MSBS. The measured drag coefficients for this body are of interest to researchers actively involved in designing minimum drag fuselage shapes. Additional investigations included: the effects of fixing transition; the effects of fins installed in the tail; surface flow visualizations using both liquid crystals and oil flow; and base pressure measurements using a one-channel telemetry system. Two drag prediction codes were used to assess their usefulness in estimating overall body drag. These theoretical results did not compare well with the measured values because of the following: incorrect or non-existent modeling of a laminar separation bubble on the body and incorrect of non-existent estimates of base pressure drag.

Description: It is observed that the center of pressure on a wing shifts as the Mach number is changed. Such shifts are in general undesirable and are sometimes compensated for by actively shifting the center of gravity of the aircraft or by using active stability controls. To avoid this complication, it is desirable to design the wings of a high speed aircraft so as to minimize the extent of the center-of-pressure shifts. This, together with a desire to minimize the center-of-pressure shifts in missile control surfaces, provides the motivation for this project. There are many design parameters which affect center-of-pressure shifts, but it is expected that the largest effects are due to the wing planform. Thus, for the sake of simplicity, this study is confined to an investigation of thin, flat, (i.e., no camber or twist), relatively slender, pointed wings flying at a small angle of attack. Once the dependence of the center of pressure on planform and Mach number is understood, we can expect to investigate the sensitivity of the center-of-pressure shifts to various other parameters.

Description: Within the last two decades, there has been increasing emphasis on developing more sophisticated computational fluid dynamics (CFD) methods to handle a wide range of problems of interest to the aerospace community. The comprehensive picture of the status of CFD development and capability as well as an assessment of requirements and future directions are given. An independent review and assessment was also carried out by the author as part of the current assignment and the results are outlined herein.

Description: Two dimensional problems are solved using numerical techniques. Navier-Stokes equations are studied both in the vorticity-stream function formulation which appears to be the optimal choice for two dimensional problems, using a storage approach, and in the velocity pressure formulation which minimizes the number of unknowns in three dimensional problems. Analysis shows that compact centered conservative second order schemes for the vorticity equation are the most robust for high Reynolds number flows. Serious difficulties remain in the choice of turbulent models, to keep reasonable CPU efficiency.

Description: Supersonic external compression inlets are introduced, and the computational fluid dynamics (CFD) codes and tests needed to study flow associated with these inlets are outlined. Normal shock wave turbulent boundary layer interaction is discussed. Boundary layer control is considered. Glancing sidewall shock interaction is treated. The CFD validation of hypersonic inlet configurations is explained. Scramjet inlet modules are shown.

Description: The fluid dynamics of curved diffuser duct flows of military aircraft is discussed. Three-dimensional parabolized Navier-Stokes analysis, and experiment techniques are reviewed. Flow measurements and pressure distributions are shown. Velocity vectors, and the effects of vortex generators are considered.

Description: Fundamental equations encountered in computational fluid dynamics (CFD), and analyses used for internal flow are introduced. Irrotational flow; Euler equations; boundary layers; parabolized Navier-Stokes equations; and time averaged Navier-Stokes equations are treated. Assumptions made and solution methods are outlined, with examples. The overall status of CFD in propulsion is indicated.

Description: One of the flows inherent in VSTOL operations, the jet in ground effect with a crossflow, is studied using the Fortified Navier-Stokes (FNS) scheme. Through comparison of the simulation results and the experimental data, and through the variation of the flow parameters (in the simulation) a number of interesting characteristics of the flow have been observed. For example, it appears that the forward penetration of the ground vortex is a strong inverse function of the level of mixing in the ground vortex. Also, an effort has been made to isolate issues which require additional work in order to improve the numerical simulation of the jet in ground effect flow. The FNS approach simplifies the simulation of a single jet in ground effect, but it will be even more effective in applications to more complex topologies.

Description: Key results from low speed wind tunnel testing of the F-15 STOL and Maneuver Technology Demonstrator (SMDT) with thrust reversers are presented. Longitudinally, the largest induced increments in the stability and control occur at landing gear height. These generally reflect an induced lift loss and a nose-up pitching moment, and vary with sideslip. Directional stability is reduced at landing gear height with full reverse thrust. Nonlinearities in the horizontal tail effectiveness are found in free air and at landing gear height.

Description: A new testing technique was developed wherein the rate of descent can be included as a parameter in ground effects investigations. This technique simulates the rate of descent by horizontal motion of a model over an inclined ground board in the Langley Vortex Research Facility (VRF) During initial evaluations of the technique, dynamic ground effects data were obtained over the inclined ground board, steady state ground effects data were obtained over a flat portion of the ground board, and the results were compared to conventional static wind tunnel ground effect data both with and without a moving belt ground plane simulation. Initial testing and analysis led to the following conclusions: the moving belt ground plane had little effect on static ground effects for the configurations tested unless thrust reversers were employed; in general, rate-of-descent reduced ground effects to the point that for reversed thrust cases an expected loss of lift due to ground effects was eliminated at approach conditions; and, in general, the steady state results from the VRF matched static results obtained from the wind tunnel once the flow field stabilized over the flat portion of the ground board.

Description: The results of an experimental investigation into the position and characteristics of the ground vortex are summarized. A 48-inch wind tunnel was modified to create a testing environment suitable for the ground vortex study. Flow visualization was used to document the jet-crossflow interaction and a two-component Laser Doppler Velocimeter (LDV) was used to survey the flowfield in detail. Measurements of the ground vortex characteristics and location as a function of freestream-to-jet velocity ratio, jet height, pressure gradient and upstream boundary layer thickness were obtained.

Description: Flow field investigations were conducted at the NASA Ames-Dryden Flow Visualization Facility (water tunnel) to investigate the ground effect produced by the impingement of jets from aircraft nozzles on a ground board in a STOL operation. Effects on the overall flow field with both a stationary and a moving ground board were photographed and compared with similar data found in other references. Nozzle jet impingement angles, nozzle and inlet interaction, side-by-side nozzles, nozzles in tandem, and nozzles and inlets mounted on a flat plate model were investigated. Results show that the wall jet that generates the ground effect is unsteady and the boundary between the ground vortex flow field and the free-stream flow is unsteady. Additionally, the forward projection of the ground vortex flow field with a moving ground board is one-third less than that measured over a fixed ground board. Results also showed that inlets did not alter the ground vortex flow field.

Description: The interaction of the free stream velocity on the wall jet formed by the impingement of deflected engine thrust results in a rolled up vortex which exerts sizable forces on a short takeoff (STOL) airplane configuration. Some data suggest that the boundary layer under the free stream ahead of the configuration may be important in determining the extent of the travel of the wall jet into the oncoming stream. Here, early studies of the ground vortex are examined, and those results are compared to some later data obtained with moving a model over a fixed ground board. The effect of the ground vortex on the aerodynamic characteristics are discussed.

Description: Numerical and experimental simulation of unsteady airflow through the control valve and slotted air duct of a circulation control rotor is described. The numerical analysis involves the solution of the quasi-one-dimensional compressible fluid-dynamic equations in the blade air duct together with the coupled isentropic flow equations for flow into the blade through the valve and out of the blade through the Coanda slot. Numerical solutions are compared with basic experimental results obtained for a mockup of a circulation control rotor and its pneumatic valving system. The pneumodynamic phenomena that were observed are discussed with particular emphasis on the characteristic system time lags associated with the response of the flow variables to transient and periodic control valve inputs.

Description: A review is given of the relationship between experimental data and the development of modern dynamic-inflow theory. Some of the most interesting data, first presented 10 years ago at the Dynamic Specialist's Meeting, is now reviewed in light of the newer theories. These pure blade-flapping data correlate very well with analyses that include the new dynamic inflow theory, thus verifying the theory. Experimental data are also presented for damping with coupled inplane and body motions. Although inclusion of dynamic inflow is often required to correlate this coupled data, the data cannot be used to verify any particular dynamic inflow theory due to the uncertainties in modeling the inplane degree of freedom. For verification, pure flapping is required. However, the coupled data do show that inflow is often important in such computations.

Description: A research study was initiated to systematically determine the impact of selected blade tip geometric parameters on conformable rotor performance and loads characteristics. The model articulated rotors included baseline and torsionally soft blades with interchangeable tips. Seven blade tip designs were evaluated on the baseline rotor and six tip designs were tested on the torsionally soft blades. The designs incorporated a systemmatic variation in geometric parameters including sweep, taper, and anhedral. The rotors were evaluated in the NASA Langley Transonic Dynamics Tunnel at several advance ratios, lift and propulsive force values, and tip Mach numbers. A track sensitivity study was also conducted at several advance ratios for both rotors. Based on the test results, tip parameter variations generated significant rotor performance and loads differences for both baseline and torsionally soft blades.

Description: A reliable rotor aeroelastic analysis operational that correctly predicts the vibration levels for a helicopter is utilized to test various unsteady aerodynamics models with the objective of improving the correlation between test and theory. This analysis called Rotor Aeroelastic Vibration (RAVIB) computer program is based on a frequency domain forced response analysis which utilizes the transfer matrix techniques to model helicopter/rotor dynamic systems of varying degrees of complexity. The results for the AH-1G helicopter rotor were compared with the flight test data during high speed operation and they indicated a reasonably good correlation for the beamwise and chordwise blade bending moments, but for torsional moments the correlation was poor. As a result, a new aerodynamics model based on unstalled synthesized data derived from the large amplitude oscillating airfoil experiments was developed and tested.

Description: A lifting surface theory was developed for a helicopter rotor in forward flight for compressible and incompressible flow. The method utilizes the concept of the linearized acceleration potential and makes use of the vortex lattice procedure. Calculations demonstrating the application of the method are given in terms of the lift distribution on a single rotor, a two-bladed rotor, and a rotor with swept-forward and swept-back tips. In addition, the lift on a rotor which is vibrating in a pitching mode at 4/rev is given. Compressibility effects and interference effects for a two-bladed rotor are discussed.

Description: The challenge in the definition of the entry aerothermodynamic environment arising from the challenge of a reliable and reusable Orbiter is reviewed in light of the existing technology. Select problems pertinent to the orbiter development are discussed with reference to comprehensive treatments. These problems include boundary layer transition, leeward-side heating, shock/shock interaction scaling, tile gap heating, and nonequilibrium effects such as surface catalysis. Sample measurements obtained from test flights of the Orbiter are presented with comparison to preflight expectations. Numerical and wind tunnel simulations gave efficient information for defining the entry environment and an adequate level of preflight confidence. The high quality flight data provide an opportunity to refine the operational capability of the orbiter and serve as a benchmark both for the development of aerothermodynamic technology and for use in meeting future entry heating challenges.

Description: Some of the single rotation propfan model wind tunnel aeroelastic findings from the experimental part of this research program are described. These findings include results for unstalled or classical flutter, blade response from separated flow excitations, and blade response from aerodynamic excitations at angled inflow conditions.

Description: The ability of existing data reduction techniques to determine frequency and damping from transient time-history records was evaluated. Analog data records representative of small-scale helicopter aeroelastic stability tests were analyzed. The data records were selected to provide information on the accuracy of reduced frequency and decay coefficients as a function of modal damping level, modal frequency, number of modes present in the time history record, proximity to other modes with different frequencies, steady offset in time history, and signal-to-noise ratio. The study utilized the results from each of the major U.S. helicopter manufacturers, the U.S. Army Aeroflightdynamics Directorate, and NASA Ames Research Center using their inhouse data reduction and analysis techniques. Consequently, the accuracy of different data analysis techniques and the manner in which they were implemented were also evaluated. It was found that modal frequencies can be accurately determined even in the presence of significant random and periodic noise. Identified decay coefficients do, however, show considerable variation, particularly for highly damped modes. The manner in which the data are reduced and the role of the data analyst was shown to be important. Although several different damping determination methods were used, no clear trends were evident for the observed differences between the individual analysis techniques. It is concluded that the data reduction of modal-damping characteristics from transient time histories results in a range of damping values.

Description: In response to a systematic methodology assessment program directed to the aeroelastic stability of hingeless helicopter rotor blades, improved basic aeroelastic reformulations and new formulations relating to structural sweep were achieved. Correlational results are presented showing the substantially improved performance of the G400 aeroelastic analysis incorporating these new formulations. The formulations pertain partly to sundry solutions to classic problem areas, relating to dynamic inflow with vortex-ring state operation and basic blade kinematics, but mostly to improved physical modeling of elastic axis offset (structural sweep) in the presence of nonlinear structural twist. Specific issues addressed are an alternate modeling of the delta EI torsional excitation due to compound bending using a force integration approach, and the detailed kinematic representation of an elastically deflected point mass of a beam with both structural sweep and nonlinear twist.

Description: The development and application of comprehensive rotorcraft analysis methods in the field of rotorcraft technology are described. These large scale analyses and the resulting computer programs are intended to treat the complex aeromechanical phenomena that describe the behavior of rotorcraft. They may be used to predict rotor aerodynamics, acoustic, performance, stability and control, handling qualities, loads and vibrations, structures, dynamics, and aeroelastic stability characteristics for a variety of applications including research, preliminary and detail design, and evaluation and treatment of field problems. The principal comprehensive methods developed or under development in recent years and generally available to the rotorcraft community because of US Army Aviation Research and Technology Activity (ARTA) sponsorship of all or part of the software systems are the Rotorcraft Flight Simulation (C81), Dynamic System Coupler (DYSCO), Coupled Rotor/Airframe Vibration Analysis Program (SIMVIB), Comprehensive Analytical Model of Rotorcraft Aerodynamics and Dynamics (CAMRAD), General Rotorcraft Aeromechanical Stability Program (GRASP), and Second Generation Comprehensive Helicopter Analysis System (2GCHAS).

Description: The present paper is concerned with two methods for the accelerated solution of the steady Euler equations. One method makes use of a second-order embedding to facilitate the derivation of the relaxation solution of the steady equations of motion, while the other method employs a multile-gridding concept to accelerate the convergence of a simple, explicit, time-marching scheme applied to the unsteady equations. It is pointed out that the surrogate equation technique provides a means for formulating problems involving the full steady Euler equations in such a way as to allow the use of relaxation solution procedures. It is, therefore, possible to solve either irrotational or rotational flow problems spanning the entire spectrum of subsonic, transonic, and supersonic conditions. The solutions can be obtained without an employement of either derived dependent variables, semidirect methods, or an unsteady formulation.

Description: It is pointed out that the supercritical wing is one of the most important features of modern transonic aerodynamics. The design of its shock free airfoil section depends on potential flow calculations. The present paper is concerned with the development of inviscid flow simulation methods based on potential formulations, taking into account also the problem of nonuniqueness of the potential solution. Nonisentropic and nonisoenergetic models are considered, and an alternative approach using the stream function is discussed. Attention is given to transonic small disturbance calculations, calculations based on the full potential equation, iterative methods, wave drag calculations, and an alternative form of Euler equations.

Description: The surface integral terms in Green's third identity are often used to solve the Prandtl-Glauert (linear potential-flow) equation with panel methods. This can be done, as in the PAN AIR code, for either subsonic or supersonic flow about complete aircraft. The extension to transonic flow is suggested by the volume integral terms of Green's third identity. The mathematical basis for this extension, without the use of body-fitted grids, is presented. Supercritical transonic results computed from a two-dimensional transonic PAN AIR research code demonstrate the method.

Description: Attempts have been made to explain why finite difference solutions of the Euler equations can describe flows with large vortical structures around sharp-edged bodies. The present paper is concerned with the influence of a singular sharp edge on the truncation error for a set of discretized Euler equations. An analysis is conducted of the distribution of the truncation error of one finite difference approximation of the Euler equations near a sharp edge of a thin plate. The analysis leads to a determination of the size of the region of the neighborhood of such a singularity. Attention is given to the consistency of a discretization of the Euler equations, and numerical experiments.