Showing papers in "Journal of Aircraft in 1968"

Analysis and design of wing-body combinations at subsonic and supersonic speeds.

Abstract: The method of aerodynamic influence coefficients has proved to be an effective tool for the analysis and design of wings, bodies, and wing-body combinations at supersonic speeds. This paper describes the extension of this method into the subsonic flow regime, and correlates the theory with experiment over a wide speed range. The method may be applied to the calculation of the pressures and forces acting on arbitrary wing-body combinations in steady flight, including aeroelastic effects, and to the design of wing camber surfaces in the presence of a body. Nomenclature A = aspect ratio b = span c =. chord C = aerodynamic coefficient d = distance / = singularity distribution function F = distribution function K = constant I = body length L — panel sweep M = Mach number r = radius Re = real part u, v, w = perturbation velocities U = freestream velocity x, y, z = Cartesian coordinates a. = angle of attack 0 = angular coordinate, panel inclination X = taper ratio A = leading edge sweep £, rj = integration variables

Subsonic diffusers designed integrally with vortex generators.

Abstract: A short subsonic diffuser has been developed incorporating vortex generators as an integral design feature. The principle of operation is that line vortices, when suitably arranged, mutually interact to drive each other towards an adjacent plane wall. The wall may then be pulled away from the vortices at such a rate that the vortices remain a constant distance from it. The ideal arrangement, having the vortex lines running along the edge of the boundary layer, can thus be obtained by design. A simple, two-dimensional diffuser was developed at Lockheed using these principles. It was compared experimentally with a conventional high-performance trumpet-shaped diffuser. Two significant results were observed in this test series. One was that both pressure recovery losses and distortion were reduced by about 40% by the new design. The other was that vortex generator design mismatches carry larger penalties with the integrally designed diffuser than with the conventional type. The subsonic diffuser of an inlet for a Mach 2.7 supersonic transport airplane was shortened and redesigned in two stages, using the integrated vortex generator approach. First, vortex generators were tailored to a short subsonic diffuser of conventional design, and then a new inlet was tested which included a subsonic diffuser designed on the integral basis. The successive changes improved both pressure recovery and flow uniformity. Thus, the basic concept was validated in a practical application.

A theory of three-dimensional parachute dynamic stability

Abstract: The three-dimensional motion of a freely descending parachute is studied with a five-degreeoffreedom analysis (the roll motion is neglected). The equations of motion are nondimensionalized and the resulting parameters discussed. Exact expressions are given for the longitudinal and lateral small-disturbance stability of the familiar gliding motion of parachutes. The breakdown of these small-disturbance expressions is illustrated by exact large-disturbance studies. A large longitudinal disturbance of most parachutes will result in a large pitching motion, whereas a large lateral (out of the glide plane) disturbance will usually cause a large angle vertical coning motion. Exact algebraic expressions are given for the coning mode, which is a stable rotation, and a small amount of available coning data is included for comparison. Some parametric computer studies of the various motions are also shown.

Turbulent Boundary-Layer Characteristics of Compliant Surfaces

Abstract: A hot-wire anemometer study was made of the turbulent boundary layer on a compliant coating. A compliant coated flat plate was made by covering a y^-in.-deep reservoir of fluid with a thin sheet of poly vinyl chloride. This compliant test plate which measured 26^ in. X 8^ in. was inserted flush in the floor of a low-speed wind tunnel. The hot-wire anemometer was used to measure velocity profiles, Reynolds stresses and turbulence intensity in the boundary layer. All tests were run at 38 fps and the skin tension, skin thickness, and reservoir fluid properties were varied during the tests. The universal velocity profile of the compliant coating indicated no change in mixing length from that of a hard plate but seemed to indicate a thicker laminar sublayer. The Reynolds stress and turbulence intensity were smaller for the compliant coating than for the hard plate and they seemed to correlate with previously measured skin-friction reductions.

A numerical solution for the minimum induced drag of nonplanar wings.

Abstract: A procedure has been developed for the accurate computation of the minimum induced drag of nonplanar wings with pylonlike panels, provided the wing front view consists of straight line segments. As is well known, the induced drag may be expressed as an integral in an auxiliary mapping plane. Previously, the main computational difficulty had been the determination of the Schwarz-Christoffel mapping between the real and the auxiliary planes. By means of the electrostatic analogy to potential flow, the constants of the mapping are determined with a small experimental error by using an analog field plotter. The mapping is then integrated by numerical techniques, and the constants are adjusted until the desired geometry is achieved to any order of accuracy. The induced drag is determined by quadrature and is shown by comparison with known test cases to be accurate to 10~ 4. Comparison of results with earlier approximate solutions (Mangier, Cone) shows that some of the earlier approximate solutions give more favorable predictions (less drag) than the solution derived here. The discrepancies in the earlier work are shown to be due to improper boundary conditions, and some suggestions are made to minimize these effects. The results show a potential reduction of minimum induced drag of less than 1% for a current subsonic jet transport when the pylons are properly loaded.

Controlled Recovery of Payloads at Large Glide Distances, Using the Para-Foil

Abstract: Recent progress of Sandia Laboratory in the development of a small guidance and control system and by the University of Notre Dame on the Para-Foil has indicated that the design of a system capable of recovering a 150-Ib payload from altitudes greater than 300,000 ft and ranges of from three to five times the deployment height (60,000 ft) is quite feasible. The Para-Foil is a completely nonrigid, self-inflating flying wing, capable of being packed and deployed like a conventional parachute and able to glide large distances. The guidance package is an electromechanical control system employing a direction-finding antenna to control the direction of the Para-Foil's glide path with respect to a ground or shipboard transmitter. The recovery system design program includes such parameters as size and weight of the recovery unit, size of the payload (W/S wing loading), glide ratio and wind structure, and flare-out capabilities in the recovery area. The full-scale recovery program includes numerous packing, deployment, and glide tests. These tests demonstrated successful deployment, excellent gliding performance (JL/D = 3.88) and dynamic flight stability, and very low impact velocities during the final recovery phases. Recent guidance and control tests also demonstrated excellent response in maneuverability of the Para-Foil. This paper summarizes the design program and the results of full-scale deployment and guided recovery flight tests and demonstrates the potential of the Para-Foil for use in many space-age recovery programs.

Attenuation of circumferential inlet distortion in multistage axial compressors.

Abstract: Circumferential inlet distortion in a high hub-tip ratio multistage machine is treated by analyzing the compressor as a region in which a large number of small stages produce a pressure rise that is a function of the local mass flow rate. The resistance to circumferential flow due to the blading is included through an empirical factor. It is found that the over-all attenuation of both total pressure distortion and axial velocity distortion is mainly dependent on the slope of the compressor pressure rise vs flow rate characteristic. The attenuation increases when the slope of the characteristic is made more negative. In addition, considerable flow redistribution is found to occur upstream of the compressor. The theory has been compared with interstage data obtained on a three-stage, low-speed compressor with axial clearances that are 26% of the total length and a hub-tip ratio of 0.675. It is found that the approximation of zero axial clearance (infinite resistance to circumferential flow) gives excellent results. In consequence, it appears that for the normal range of axial clearances, the circumferential flow within the compressor can be neglected in a first-order analysis of the effects of inlet distortion.

Atmospheric turbulence considerations for future aircraft designed to operate at low altitudes.

Abstract: A research program (LO-LOCAT) is presently in progress to investigate the characteristics of atmospheric turbulence at low altitudes. A significant quantity of gust-velocity and meteorological data are being measured by instrumented aircraft. Resultant information will be incorporated in design criteria for those aircraft slated for operation in the next two decades. One of the items required for these criteria, and one objective of the program, is an analysis of atmospheric turbulence spectra. Results obtained thus far, regarding this aspect of the research, are discussed. Good correlation has been found between experimentally determined spectra and mathematical expressions as advanced by von Karman, and Lumley and Panofsky. Conversely, the experimental data have shown poor agreement with expressions developed by Dryden and by Lappe. This finding is pertinent, since the Dryden expressions have been used extensively in the past and are, in fact, recommended by some agencies for use in present aircraft design. An analysis of turbulence scale lengths associated with 1200 of the experimentally measured spectra is included. Trends in scale length variation with associated geophysical characteristics and gust-velocity statistics are shown. These data give an indication of the low-altitude atmospheric turbulence model which is evolving from the LO-LOCAT program.

Calculation of Profile Drag of Airfoils at Low Mach Numbers

Abstract: This paper investigates the accuracy of a particular method for calculating the profile drag of airfoils at low Mach numbers. The method consists of 1) calculation of the pressure distribution by any suitable method, 2) calculation of laminar flow near the nose by Thwaites' method, 3) calculation of transition by Michel's method (if it is not known a priori), 4) calculation of turbulent boundary-layer flow by Head's method, and, finally, 5) calculation of momentum deficiency in the far wake by means of the Squire-Young relation. Profile drag has been calculated by this method for several airfoils at various angles of attack and Reynolds numbers. The effects of transition and airfoil thickness on the profile drag are studied. Comparison of calculated and experimental values show generally good agreement. The rms error based on 88 calculated drag values is 2.7%.

Design criterion of panel structure excited by turbulent boundary layer.

Abstract: Using a relatively simple functional representation of the space-time correlation of the wallpressure fluctuation, the motion of a simply supported panel and the resulting acoustic radiation can be predicted within 1 order of magnitude from the experimental results. Criterion for designing a panel for given flow conditions by this method is considered. The governing parameter is the so-called coherence distance, the distance over which a given turbulent pattern remains distinguishable. Calculations indicate that, when coherence distance is much smaller than the panel length, the response is mostly due to coincidence. From knowledge of the panel motion, the radiated sound intensity is obtained. For a panel much longer than the coherence distance, the acoustic power radiated is considerably reduced. Significant results were obtained from suggested practical methods of lessening the panel response and vibration noise level. Structural excitation by separated flow is localized on an airplane, but since it is severe it is given careful consideration.

Induced Flow Variation of the Helicopter Rotor Operating in the Vortex Ring State

Abstract: By using small, quickly responsive, and very sensitive windmills we can measure the inflow variation of the model-helicopter rotor operating in the vortex ring state. The wind-tunnel tests show that 1) as the rotor starts to descend vertically, the periodic-induced-flow variation is observed at the rotor tip without any notable thrust change; 2) when the rate of descent approaches the induced velocity generated at hovering state, a thrust reduction appears, this being closely correlated with the increment of the downwash component near the rotor tip induced by a strong vortex ring; 3) as the rate of descent'increases beyond the aforementioned induced velocity, the downwash variation becomes predominant near the center of the rotor and the thrust variation tends to decrease; 4) torque variation is not observable for low collective pitch operation, but for high pitch the torque fluctuates with thrust variation. The preceding final result is explained by the blade element theory with measured inflow.

Fabrication and full-scale structural evaluation of glass-fabric reinforced plastic shells.

Abstract: This paper reports the initial investigations on full-size sandwich shell structures consisting of cylindrically curved panels and full cylinders, although some previous work in fabrication and basic buckling tests on small flat panels is presented as background information. The large thin-walled sandwich shells were fabricated by bonding 2-ply epoxy-fiber glass laminate facings to aluminum honeycomb cores. The curved panels were tested in axial compression and in torsional shear. The complete cylinders were loaded in a special torsionbending testing frame capable of applying bending moments and torques of 3,000,000 in.Ib. each. In general, the buckling test results compared as well with calculated values as do those for buckling of homogeneous isotropic shells.

Shock-wave reflection from a turbulent boundary layer with mass bleed.

Abstract: Methods for calculating the changes in turbulent boundary-layer characteristics across an oblique shock reflection based on integral flow models are developed for two-dimensional and axially symmetric flows with solid boundaries and mass bleed through porous walls, slots, or scoops. For each model, a control surface is assumed about the region of interaction; the velocity profiles upstream and downstream of the interaction are assumed to be power laws, and the integral continuity and momentum equations are written for the control volume. For a given mass-bleed configuration, specification of the upstream conditions and the bleed rate permits solution for the downstream boundary-layer thickness and velocity-profile exponent. Numerical results are presented for a range of upstream Mach numbers, mass-bleed rates, and incident-shock strengths. Comparison with data for two-dimensional shock reflections with zero bleed indicates that the analysis yields the correct trends.

A frequency domain approach to linear optimal control.

Abstract: To find the control that minimizes a quadratic performance index requires, in the frequency domain, the solution of a matrix Wiener-Hopf equation A simple algebraic method is given for solving this equation in the linear mult icon troller case The solution yields not only the optimal control but the feedback gains as well This obviates the need to solve a steadystate Riccati equation The ideas involved are illustrated with a simplified set of equations of motion representing a small jet aircraft in a power approach Thrust and elevator are both used as active controllers A method that specifies a set of desirable closed-loop poles and yet retains some freedom for positioning the zeros of the system is discussed

Aircraft design augmented by a man-computer graphic system.

Abstract: A status report is presented on one facet of a continuing research and development program at the Lockheed Georgia Research Laboratory in the application of man-computer graphic systems to the design and manufacture of aircraft. The program described is one of the initial steps contributing to the eventual evolution of a computercentric engineering/manufacturing/management system. It is being developed as a prototype covering the aircraft preliminary design and performance estimation process. This prototype will enable a creative design engineer to evolve and refine an aircraft configuration rapidly, by graphically and numerically describing aircraft components to a computer, using interactive, on-line, graphic display devices and unique, generalized graphic programs, by directing the computer through analyses of the configuration or components, and by reviewing directly on the computerdriven display the results of the analysis for approval or modification. A time compression of the preliminary design process of perhaps 10-to-l is one of the goals of this system. The system is discussed in general, and the graphic and analytical routines are described in some detail. Photographs show the graphic terminals being used in the execution of key steps in the preliminary design process.

Surfaces for computer-aided aircraft design.

Abstract: A simple but general way is described to define freeform surfaces such as airplane fuselages, wings, fillets, ducts, and other shapes by means of man-machine graphical interaction with a computer. In the past, much attention has been directed toward fitting mathematical functions to surfaces already defined by a mesh of points. The present discussion centers around the philosophy that in the preliminary phase of shape description the computer's aid should be enlisted at the very beginning and that in this way the results of preliminary surface design become the first "master dimensions" of the airplane directly, without the necessity of refairing or other subsequent treatment. These surfaces and curve formulations can be incorporated into computer data structures, where they will serve as the skeleton upon which other associated data can be hung, such as velocity fields, pressures, temperatures, forces, and other physical quantities that arise in connection with analytical and design procedures.

Some contributions on a control-surface buzz at high subsonic speeds.

Abstract: of Minnesota; also reprint, Univ. of Notre Dame, Notre Dame, Ind. 2 Nicolaides, J. D., "On the Discovery and Research of the Para-Foil," International Congress on Air Technology, Nov. 1965, Little Rock, Ark.; also reprint, Univ. of Notre I3ame, Notre Dame, Ind. Nathe, G. A., Knapp, C. F., and Nicolaides, J. D., "The Para-Foil and Targetry Applications," Feb. 8, 1966, Target Systems Planning Group, Hill Air Force Base, Ogden, Utah. 4 Nathe, G. A., "An Analysis of the Para-Foil," AIAA Student Journal Vol. 5, No. 1, Feb. 1967, pp. 4-9. 5 Knapp, C. F., Nathe, G. A., and Nicolaides, J. D., "Launching Technique for Tethered Para-Foils," Manual, May 9, 1966, NASA, Cape Kennedy, Fla. 6 Nathe, G. A., Knapp, C. F., and Hall, C. R., "Wind Tunnel and Free Flight Testing of Para-Foil Number 125, "Departmental Paper, June 1, 1966, Aero-Space Engineering Dept., Univ. of Notre Dame, Notre Dame, Ind. 7 Nathe, G. A. and Knapp, C. F., "A Qualitative Discussion of the Lateral-Directional Stability of Tethered and Towed Para-Foils," Departmental Paper, June 6, 1966, Aero-Space Engineering Dept.. Univ. of Notre Dame, Notre Dame, Ind. 8 Gorin, B. F., "An Approximation of Spanwise Para-Foil Collapse," Departmental Paper, June 1966, Aero-Space Engineering Dept., Univ. of Notre Dame, Notre Dame, Ind. 9 Nicolaides, J. D., Nathe, G. A., and Knapp, C. F., "A Summary of the Tests Conducted on the Para-Foil," June 8, 1966, Aero-Space Engineering Dept., Univ. of Notre Dame, Notre Dame, Ind. 10 Nicolaides, J. D. and Knapp, C. F., "Para-Foil Design," UNDAS-866 JDN Kept., Contract AF 33(615)-5004, Aug. 1967, U.S. Air Force Flight Dynamics Lab., Wright Patterson Air Force Base, Ohio. 11 Terman, F. E., Radio Engineers Handbook, McGraw-Hill, New York, 1943. 12 "Ground Base VHF Automatic Direction Finder Equipment," Final Engineering Rept., Contract AF-08(606)-364, Aug. 1954, College of Engineering, Univ. of Florida, Gainesville, Fla. 13 Kane, M. T., Dicken, H. D., and Buehler, R. C., "A Homing Parachute System," SC-4537(RR), Jan. 1961, Sandia Corp., Albuquerque, N.Mex. Coonce, C. A., "Para-Foil Free-Flight Test Data, SCTM 66-2616, Dec. 1966, Sandia Lab., Albuquerque, N.Mex.

In-flight measurement of human response characteristics.

Abstract: : This paper deals with flight programs in which data have been obtained and analyzed to define the transfer characteristics of the human pilot. The first part of the paper is a general review of previous programs and data that are documented in the literature. Some of the techniques of these past programs are described. However, more in-flight data are needed, than has been obtained in the past, to more fully define the pilot in the actual flight situation. To initiate obtaining more flight data a program has recently been done in which the Air Force variable stability T-33 was used as both a ground-based simulator and an in-flight simulator. The second part of the paper discusses this current program in which pilot transfer characteristics have been obtained for the compensatory roll tracking task. Data are presented which describe the pilot in both the ground simulator and actual flight situations.

Elastic wind-tunnel models for predicting longitudinal stability derivatives of elastic airplanes.

Abstract: The objective of this investigation was the development of elastic wind-tunnel models to be used in predicting stability derivatives for elastic airplanes, in particular the SST. A novel feature of these models is that they were tested at full-scale dynamic pressures. Scaling laws, manufacturing problems, and testing problems encountered during this development are brought out. The differences between jig shape and design shape of an elastic airplane are illustrated, using the Boeing SST as an example. Application and use of elastic model data to the prediction of full-scale airplane lift and pitching-moment characteristics are presented. Correlation data between theoretical predictions for longitudinal stability and control derivatives and experimental measurements on elastic models are given. These results clearly demonstrate the practicality of elastic models for tests at full-scale dynamic pressures.

Takeoff performance of jet-propelled conventional and vectored-thrust STOL aircraft.

Abstract: An analytical method for the determination of the air distance of jet-propelled conventional (CTOL) and vectored-thrust short takeoff and landing (STOL) aircraft has been developed. The method assumes constant lift and drag coefficients during the climb and a constant value of the horizontal acceleration based on the aircraft's average velocity from touchoff to the 35- or 50-ft obstacle. It is indicated by this method that some "classical" or specification methods for computing CTOL air distances are not generally applicable. Design charts are developed for determining the ground, air, and total takeoff distance in terms of: an STOL thrust-to-weight ratio, (Tx/W)/(\ - TV/W)\ an effective wing loading (W/S)(l/aCLo)', and the L/D ratio. A parametric study of approximately 100 hypothetical CTOL and STOL aircraft was made, and graphs of the takeoff performance presented. It is shown that thrustto-weight ratios greater than approximately 0.6 are required in order to show performance gains by thrust vectoring, that the major improvement in takeoff distance derives from the reduction in ground roll while air distance is relatively unaffected, and that reduced wing loadings or improved high-lift capabilities are equally as beneficial.

Design and development of an air intake for a supersonic transport aircraft.

Abstract: The reasons for choice and the characteristics of an external-compression intake geometry for operation at Mach numbers up to and beyond 2.0 are described, and the problems of application to SST aircraft are discussed. Mounted under the wing of the Concorde, this inlet is divided to provide independent supplies of air to a pair of engines, an arrangement that introduces particular problems in allowing for the wing flowfield and avoiding interaction between the twin inlets. The precise definition of an intake geometry for a supersonic transport should have regard for the over-all performance of the propulsion system. The choice of controlling parameters and the design of the control system must give good performance and engine handling in a wide range of off-design conditions without demanding excessive complexity. The aerodynamic and other development tests required to make the appropriate decisions are described in detail. The results underline the suitability of this basic geometry in association with the other components of the propulsion system for SST operation.

Linear Optimal Theory Applied to Active Structural Bending Control

Abstract: To improve ride qualities and structural fatigue life of high-performance aircraft, control of structural flexing is desirable. Linear optimal control via the root-square locus was employed to design a simple, effective bending-control system for four XB-70 coupled longitudinal bending modes. The weighted sum of the mean-square differential angular acceleration and the control surface position was minimized for random wind-gust inputs. Approximation of frequency response characteristics of high-order optimal compensations led to a suboptimal mechanization (fixed-parameter, third-order transfer function with a programmed gain) which produced nearly optimal performance for three diverse low-altitude flight cases. The rms differential angular acceleration was reduced by a factor of 5 for a control expenditure of approximately 13°/sec rms elevon rate per fps of rms wind-gust velocity. Changed vehicle parameters for a fourth flight case require different sensor locations for effective control. Extensive design data show correlation of performance with plant gain and performance-index weighting factors, indicating a potential for predesign prediction of results. Comparable performance is provided by simple, uncompensated feedbacks for some cases, but stability margins are lower than for the near-optimal systems.

Effects of manual altitude control and other factors on short-periodhandling quality requirements.

Abstract: Several factors that appear to affect short-period handling quality requirements are reviewed with particular attention paid to manual control of pitch attitude and altitude. The effects of the various short-period parameters on the pilot's closures of these two loops are examined. Other factors that are considered include attitude overshoots, flight-path and attitude consonance, gust responses, flight condition, and vehicle type. It is found that in several cases the factors produce conflicting requirements. For example, attitude and altitude-control requirements can be conflicting. A very encouraging correlation between the expectations based on analysis and existing experimental data is found. From this correlation, several handling quality requirements for landing and cruise conditions are formulated. It is also shown that short-period requirements cannot, in general, be reduced to two or three simple parameters.