Showing papers in "Journal of Aircraft in 1973"

Vortex Noise of Isolated Airfoils

Abstract: Hohenemser, K. H. and Yin, Shen-Kuang, "Some Applications of the Method of Multi-blade Coordinates," Journals of the American Helicopter Society, Vol. 17, No. 3, July 1972, pp. 3-12. Coleman, R. P., "Theory of Self-Excited Mechanical Oscillations of Hinged Rotor Blades," NACA Advanced Restricted Rept. 3G29, Republished as Rept. 1351,1943, NACA. Miller, R. H., "Helicopter Control and Stability in Hovering Flight," Journal of the Aeronautical Sciences, Vol. 15, No. 8, Aug. 1948, pp. 453-472. Azuma, A., "Dynamic Analysis of the Rigid Rotor System," Journal of Aircraft, Vol. 4, No. 3, May-June 1967, pp. 203-209. Curtiss, H. C., Jr. and Shupe, N. K., "A Stability and Control Theory for Hingeless Rotors," American Helicopter 27th Annual National V/STOL Forum, Washington, D. C., May 1971. Ward, J. F., "Exploratory Flight Investigation and Analysis of Structural Loads Encountered By A Helicopter Hingeless Rotor System," D-3676, Nov. 1966, NASA. Harris, F. D., "Articulated Rotor Blade Flapping Motion at Low Advance Ration," Journal of the American Helicopter Society, Vol. 17, No. 1, Jan. 1972, pp. 41-48. Ormiston, R. A. and Peters, D. A., "Hingeless Helicopter Rotor Response with Nonuniform Inflow and Elastic Blade Bending," Journal of Aircraft, Vol. 9, No. 10, Oct. 1972, pp. 730736.

Identification and Optimization of Aircraft Dynamics

Abstract: A technique is described for the design of an adaptive controller for multivariable systems and is based on recently developed methods for identification and optimization. An application of the method to a helicopter system with time-varying parameters is considered in detail. The response of the adaptive system is compared with the corresponding response of a system with a fixed controller and a system using optimal control. The comparison reveals the almost optimal character of the adaptive system. Nomenclature A = n X n, system matrix B = n X m, input matrix C = n X n, model matrix F = m X n, feedback matrix G — n X n, model matrix (estimate of A) H = n X m, model matrix (estimate of B) K = n x n, symmetric Riccati matrix P = n X n, symmetric positive definite matrix used in the model Q, S — n X n, symmetric positive semidefinite matrices of the performance index Qi = n X n, symmetric positive definite Lyapunov matrix R = m X m, symmetric positive definite matrix of the

Gas Path Analysis Applied to Turbine Engine Condition Monitoring

Abstract: An approach is presented to turbine engine gas path analysis and monitoring, which permits the isolation of single or simultaneous multiple engine faults, with a quantitative assessment of their relative' severity. The software approach is described, showing features of its mathematical development and thermodynamic justification. Measureable engine parameters are treated as dependent variables, changes in which are mathematically interrelated to changes in component performance brought about by physical engine faults. Typical results are presented from real programs, wherein engine data were analyzed to provide meaningful and verified diagnoses of single and multiple engine faults.

A modified wall wake velocity profile for turbulent compressible boundary layers.

Abstract: The law of the wake and the law of the wall in incompressible turbulent boundary layers formulated by Coles (1956) and their use by Mathews et al. (1970) in the development of a wall-wake representation of the velocity profile in a form applicable for isoenergetic compressible boundary layers are extended to a modified wall-wake velocity profile for turbulent compressible boundary layers. The modified wall-wake profile is shown to provide good representations of experimental velocity distributions.

Rapid Scanning, Three-Dimensional Hot-Wire AnemometerSurveys of Wing-Tip Vortices

Abstract: The use of CADD has effected many changes in traditional work methods of the various interfacing disciplines at MCAIR. We have streamlined our tasks by utilizing the cost/time saving techniques presently afforded by this system, and plan to continue additional development efforts which will further improve it. As we discover more and more practical applications for CADD, new or modified software routines will be formulated to make the system as responsive and reliable as possible. Since the potential for substantial dollar savings exists, particularly in the area of manufacturing, we are consciously attempting to align our techniques to make engineering data as usable as possible by manufacturing personnel in its existing form, thus eliminating as many intermediate manual processes as possible. Our final objectives, however, cannot now be defined. CADD's utility will increase as we continue to conceive new ideas for its application. Its ultimate benefits can only be determined by the limits of our creative imagination.

A Class of Airfoils Designed for High Lift in Incompressible Flow

Abstract: The problem studied is that of designing a single element airfoil which provides the maximum possible lift in an unseparated incompressible flow. First, an airfoil velocity distribution is defined and optimized using boundary-layer theory and the calculus of variations. The resulting velocity distribution is then used as an input for an inverse airfoil design program which provides the corresponding airfoil shape. Since there is no guarantee that an arbitrarily defined velocity distribution will yield a physically possible airfoil shape, some parametric adjustments in the optimized distributions are required in order to obtain realistic and practical airfoil geometries. Wind-tunnel tests of two different airfoils (one assuming a laminar rooftop and the other a turbulent rooftop) have been conducted and in both cases the results met the theoretically predicted performance; for example, the laminar section exhibited a low drag range of CD — 0.0085 from CL - 0.8 to CL = 2.2.

Observations of Atmospheric Effects on Vortex Wake Behavior

Abstract: Smoke-marked trailing vortices were generated by a light aircraft under a hierarchy of measured atmospheric stability and turbulence levels and their motion and decay was recorded photographically. Decay from both sinuous vortex interaction and core bursting type instabilities occurred, with bursting being the dominant mode. Turbulence had a strong effect on wake life, with time-to-breakup for both modes varying as e 173, where e is the turbulent dissipation rate. Observed lifetimes ranged from 6 sec in light-to-moderate turbulence to more than SO sec in calm, stable air. One exceptionally long-lived solitary vortex was observed for more than 3 min. Atmospheric stratification had a weak influence on wake life and its effect on wake descent could not be determined, since descent was often stopped by a rolling of the plane of the vortices. The observed data correlates well with a new theory for time-to-breakup.

On the inviscid rolled-up structure of lift-generated vortices

Abstract: A simple form is presented of the relationships for the inviscid, fully developed structure of lift-generated vortices behind aircraft wings. The method is then extended to arbitrary span-load distributions by inferring guidelines for the selection of rollup centers for the vortex sheet, along with rules for calculating the fully developed structure of the resulting multiple vortices. These techniques yield realistic estimates of the rolled-up structure of vortices produced by a wider variety of span-load distributions than possible with the original form of the theory.

Inclusion of Rotor Dynamics in Controller Design for Helicopters

Abstract: State-feedback-controllers and state-estimators (filters) are designed for the roll-pitch-horizontal motions of a helicopter near hover, using a new quadratic synthesis technique. One model (tenth order) uses a dynamic model of the rotor, whereas the other model (sixth order) assumes the rotor can be tilted instantaneously. It is shown that, for tight control, neglecting the rotor dynamics in designing the autopilot can produce unstable closed-loop response on the model that includes rotor dynamics. Two filters are designed to use only fuselage sensors and two are designed to use both fuselage and rotor sensors. It is shown that rotor states can be estimated with sufficient accuracy using only fuselage sensors so that it does not seem worthwhile to use rotor sensors. The mean square response of the vehicle to a gusty, random wind, using several different filter/state-feedback compensators, is shown to be satisfactory.

Flight Test Studies of the Formation and Dissipation of Trailing Vortices

Abstract: A flight test program has been conducted to measure experimentally parameters which describe the characteristics of vortex behavior and their instability. Three basic atmospheric flight conditions were investigated: steady level flight in calm air, steady level flight in light gusting air and light winds, and unsteady flight produced by control surface oscillations in calm air. A DeHavilland Beaver DHC-2 airplane and a Beechcraft T-34B airplane were used in the investigation. Smoke grenades were located near the wing tips of each airplane such that the vortices could be seeded with smoke and thus made visible. This made it possible to photograph the vortices and make measurements from the photographs obtained. The experimental results show wavelengths of vortex instabilities, dissipation time of trailing vortices, effects of atmospheric current or gust, and the effect of control surface oscillation. These results provide additional experimental verification of the existence of vortex wake instability predicted by theory and show that small oscillations in pitch at a critical frequency accelerate the dissipation of high-intensity vortices.

Compact Ejector Thrust Augmentation

Abstract: Ejectors offer interesting means for resolving problems arising from the additional power requirements of V/STOL aircraft. They are capable not only of turning and augmenting the cruise engine's thrust vector, but their efflux can also serve to control circulation lift. The feasibility of ejector propulsion-lift concepts requires the simultaneous attainment of two conflicting objectives: high performance and compactness. Performance is degraded by losses occurring in the inlet, the primary nozzle, and the duct-diffuser of the ejector. Analytic results identified practical loss trade-offs that led to the design of the ejector's components. Static experiments with independently varied duct and diffuser lengths showed, surprisingly, that skewed flows can be diffused effectively. Augmentation ratios in excess of 1.85 were measured with a duct-diffuser 28 in. long. Increasing the length to 50 in. caused augmentation ratios to exceed 2.0.

Design and Development of the United Aircraft Research Laboratories Acoustic Research Tunnel

Abstract: This new wind tunnel, designed specifically for aerodynamic noise research, has a variable area, open jet test section enclosed in an anechoic chamber. Speeds up to a Mach number of 0.65 are possible with a test section area of 5 ft2. A separate, low noise level, high pressure air source is provided for jet noise studies. Design and construction considerations for the major tunnel sections (i.e., inlet, anechoic chamber, diffuser, muffler, driver) are described. Aerodynamic and acoustic calibration of the tunnel is discussed in relation to design criteria. Problems applicable to other acoustic tunnels such as background noise, edge tone suppression, deflected jet noise and distortion of noise directivity patterns by shear layer refraction are discussed. Initial testing demonstrated that a sufficiently low background noise level has been obtained to permit measurement of the noise radiated by the relatively weak noise source of an isolated airfoil placed in a uniform, low turbulence stream.

Bleed System Design Technology for Supersonic Inlets

Abstract: A boundary-layer bleed system design procedure for supersonic inlets, with emphasis on the selection of bleed hole geometry, is described. Available experimental bleed hole performance data, coupled with bleed drag calculations, show that holes with shallow inclination are superior to holes normal to the surface in terms of over-all inlet performance. Recent test results from large-scale inlet models indicate that bleed hole size, bleed hole length, and boundary-layer velocity profile upstream of the bleed region are important parameters in the design of an effective and efficient bleed system. Nomenclature A = area A/A* = sonic area ratio Ca - drag coefficient D = bleed hole diameter HI - boundary-layer shape factor L/D = bleed hole length to diameter ratio m = mass flow M = Mach number P = pressure Q = sonic mass flow coefficient; ratio of actual mass flow to theoretical maximum mass flow at local total pressure and temperature X/D = spacing between rows of bleed holes Y/D - bleed hole spacing a = bleed hole inclination to inlet surface /8 = exit nozzle ramp angle 6* = boundary-layer displacement thickness rj = bleed efficiency factor Subscripts bl = bleed epl = bleed exit plenum / = cowl lip loc = local condition pi — bleed plenum se = sonic exit to — freestream stagnation condition 00 = freestream

On the Fuel Optimality of Cruise

Abstract: longitudinal stability in Fig. 2 will not change; however, lateral stability will be greatly affected. Figures 4-6 show the effect that the reaction wheel has on lateral stability. From Fig. 4 it is seen that as the wheel size increases, the cable length required for stability decreases. Similar results are obtained in Figs. 5 and 6 when the wheel speed or weight is increased. These plots show that the cable length can be decreased to almost any length such that lateral stability is no longer the controlling factor, but longitudinal stability is. While not completely satisfactory, the addition of the reaction wheel is seen to have a tremondous effect on lat-

Dynamic Behavior of Solid Particles Suspended by Polluted Flow in a Turbine Stage

Abstract: The equations that govern the three dimensional motion of solid particles suspended by a compressible gas flow through a rotating cascade of a turbomachine are formulated. These equations are solved for the case of flow through a turbine stage. The solution takes into account the loss in particle momentum due to their collision with the turbine blades or casing. The dynamic characteristics of the solid particles; namely, their absolute trajectories, paths relative to the turbine rotor, velocity distributions, and the combined stage velocity diagrams, are calculated. The effects of changing the particles mean diameter, material density, and initial particle and gas velocities at the stator inlet on the dynamic characteristics of the solid particles are investigated. The results obtained from this study indicate the locations on the turbine blades subjected to severe erosion damage. Nomenclature a = particle absolute acceleration B = frame fixed in blades /# = angle between particle relative velocity and tangent to surface C = absolute velocity C" = absolute velocity component in the x,0 plane Cpg , = specific heat of gas at constant pressure

Fluctuating Lift and Moment Coefficients for Cascaded Airfoils in a Nonuniform Compressible Flow

Abstract: The effects of compressibility on both the fluctuating lift and the fluctuating moment coefficients for cascaded airfoils due to an upstream nonuniformity are determined by obtaining a solution to the time-dependent, compressible, two-dimensional partial differential equation which describes the perturbation velocity potential. This is accomplished through an application of Fourier-trans form theory, with the resulting integral solution equation evaluated numerically by a matrix-inversion technique. The results presented show the variation in both the fluctuating lift and the fluctuating moment coefficients over the mean cascade inlet Mach number range of 0.0 (incompressible) to 0.9 with the cascade solidity, cascade stagger angle, interblade phase angle and reduced frequency as parameters.

Decoupling of a Class of Nonlinear Systems and Its Application to an Aircraft Control Problem

Abstract: The necessary and sufficient condition for decoupling a nonlinear system with state feedback is obtained. It is shown that when this condition is satisfied there exists a control law which makes each output variable of the dynamical system independently controllable with a separate input. The theory is then applied to an aircraft control problem where the implication of the theoretical results is discussed. The objective of the aircraft control problem is to decouple the vertical and the horizontal path angles of the flight trajectory relative to earth-fixed axes. The aircraft equations are simplified by postulating a rudder control law which maintains zero sideslip velocity in flight. The control laws for the elevator and aileron which decouple the simplified aircraft model are obtained. The control system is then evaluated in a simulation study to show that it indeed decouples the flight path angles.

Remarks on vortex-lattice methods.

Abstract: Results are presented of some numerical experiments on simple planar configurations. The experiments serve to establish more precisely some ground rules for optimum lattice arrangements. In particular, the location of both the horseshoe vortex elements and the control points at which the surface boundary conditions are to be satisfied is uniquely determined. Questions of lattice arrangement are discussed together with numerical results and problems of control point location.

Exact Method of Designing Airfoils With Given Velocity Distribution in Incompressible Flow

Abstract: The inverse problem of airfoil theory, i.e., from a given surface velocity distribution determine the airfoil shape, is solved by conformal mapping procedures. The method is based upon prior work by Arlinger, which in turn is an extension of Lighthill's basic development. It involves the use of least squares and Lagrangian multipliers to modify the prescribed velocity distribution along a portion of the lower surface of the airfoil, thus ensuring that the modifications required for profile closure are minimized. The method developed should be of particular importance for calculating the shapes of new types of airfoils with high design lift coefficients, i.e., under conditions when conventional linearized theory breaks down. The method is exact in the sense of potential flow theory. Sample calculations are presented for a prescribed velocity distribution having an upper-surface constant-velocity region, followed by a Stratford-type zero-skin-friction portion, designed for Reynolds number = 3.10 and turbulent flow on both upper and lower surfaces.

Aerodynamic Boundary Layer Effects on Flutter and Damping of Plates

Abstract: References Stable, C. V., "Transonic Effects on T-tail Flutter," RM-24, 1959, The Martin Company, Baltimore, Md. Stark, V. J. E., "Aerodynamic Forces on a Combination of a Wing and a Fin Oscillating in Subsonic Flow," SAAB TN54, Feb. 1964, Likoping, Sweden. Laschka, B. and Schmid, H., "Unsteady Aerodynamic Forces on Coplanar Lifting Surfaces in Subsonic Flow (Wing-Horizontal Tail Interference)," AGARD Structures and Materials Panel Meeting, Sept. 25-27,1967, Ottawa, Canada. Topp, L. J., Rowe, W. S., and Shattuck, A. W., "Aeroelastic Considerations in the Design of Variable Sweep Airplanes," 1C AS Paper 66-12, Sept. 1966, London, England. Sensberg, O. and Laschka, B., "Flutter Induced by Aerodynamic Interference between Wing and Tail," Journal of Aircraft, Vol. 7, No. 4, July-Aug. 1970, p. 319-324. Mykytow, W. J., Noll, T. E., Huttsell, L. J., and Shirk, M. H., "Investigations Concerning the Coupling Wing-Fuselage-Tail Flutter Phenomenon," Journal of Aircraft, Vol. 9, No. 1, Jan. 1972, pp. 48-54. Albano, E. and Rodden, W. P., "A Doublet Lattice Method for Calculating Lift Distributions on Oscillating Surfaces in Subsonic Flow," AIAA Journal, Vol. 7, No. 2, Feb. 1969, pp. 279-285. Albano, E., Perkinson, F., and Rodden, W. P., "Subsonic Lifting Surface Theory Aerodynamics and Flutter Analysis of Interfering Wing/Horizontal Tail Configurations," AFFDL-TR-70-59, Sept. 1970, Air Force Flight Dynamics Lab., Wright-Patterson Air Force Base, Ohio. Triplett, W. E., Burkhart, T. H., and Birchfield, E. B., "A Comparison of Methods for the Analysis of Wing-Tail Interaction Flutter," AIAA Paper 70-80, 1970, New York. Goetz, R. C., "Lifting and Control Surface Flutter," TMX52876, Vol. Ill, July 1970, NASA. ^Rodden, W. P., Giesing, J. P., and Kalman, T. P., "New Developments and Applications of the Subsonic Doublet-Lattice Method for Nonplanar Configurations," AGARD Symposium, Tonsberg, Oslofjorden, Norway, Nov. 1970. Chipman, R. R., "Interference Flutter on Space Shuttle Configurations," Paper presented at Aerospace Flutter and Dynamics Council Meeting, May 1972, New York. Bisplinghoff, R. L. and Ashley, H., Principles of Aeroelasticity, J. Wiley and Sons, New York, 1962.

Stall-Flutter Analysis

Abstract: An analysis of unsteady airfoil stall and stall flutter is presented that is based on a series of approximations. Unsteady aerodynamic characteristics are related theoretically to static aerodynamic characteristics. Preliminary results show good agreement with experimental dynamic stall data. The analysis is applied to determine the boundaries for stall flutter, particularly for the straight wing of one candidate space-shuttle configuration. As formulated, the analysis should provide a conservative estimate - i.e., the predicted stall flutter region is slightly larger than the expected one, as demonstrated by comparison with experiments.

Dynamics of Slung Bodies Using a Single-Point Suspension System

Abstract: "Watteirdorf, F., "Study of the Effect of Curvature on Fully Developed Turbulent Flow," Proceedings of The Royal Society, Vol. A148,1935, pp. 568-598. Spangenberg, W. G., Rowland, W. R., and Mease, N. E., "Measurements in a Turbulent Boundary Layer Maintained in a Nearly Separated Condition," Fluid Mechanics of Internal Flow, Elsevier, New York, 1967, pp. 110-151. Clauser, F. H., "Turbulent Boundary Layers in Adverse Pressure Gradients," Journal of Aeronautical Sciences, Vol. 21, Feb. 1954, pp.91-108. Ludwieg, H. and Tillmann, W., "Investigations of the Wall Shearing Stress in Turbulent Boundary Layers," TM 1285, 1950, NACA. Townsend, A. E., "The Behaviour of a Turbulent Boundary Layer near Separation," Journal of Fluid Mechanics, Vol. 12, April 1962, pp. 536-554. Thompson, B. G. J., "A New Two-Parameter Family of Mean Velocity Profiles for Incompressible Turbulent Boundary Layers on Smooth Walls," R&M 3463, 1967, Aeronautical Research Council, London, England.

Development of pilot-in-the-loop analysis.

Abstract: PILOTED aircraft, to be effectively used, have always required a satisfactory match of the aircraft characteristics (including vehicle dynamics, control manipulator, stability augmentors, displays, etc.) with the controller properties of the human pilot. An agreeable marriage is not intrinsically achieved in the design process, so the provision of proper aircraft flying qualities has often posed serious problems which the designer must solve. Until fairly recently these solutions relied very heavily on intuitive cutand-try procedures. Over the years this approach fostered many of the adventures and uncertainties of flight testing. The desire to handle aircraft stability and control problems in a more analytical fashion was recognized long ago. For example, before World War II Koppen stated: "Since the controlled motion of an airplane is a combination of airplane and pilot characteristics, it is necessary to know something about both airplane and pilot characteristics before a satisfactory job of airplane design can be done." But the central difficulty in accomplishing a pilot/vehicle analysis was recognized earlier still. For example, W. Crowley and Sylvia Skan remarked in a 1930 Aeronautical Research Committee report: "A mathematical investigation of the controlled motion is rendered almost impossible on account of the adaptability of the pilot. Thus if it is found that the pilot operates the controls of a certain machine according to certain laws, and so obtains the best performance, it cannot be assumed that the same pilot would apply the same laws to another machine. He would subconsciously, if not intentionally, change his methods to suite the new conditions, and the various laws possible to a pilot are too numerous for a general analysis." Actually, matters are even worse than Crowley and Skan recognized; for while much of the pilot's dynamic behavior is governed by the aircraft dynamics, many additional factors also affect his properties.

Static Aeroelasticity and the Flying Wing

Abstract: Smith, M. J. T. and House, M. E., "Internally Generated Noise From Gas Turbine Engines, Measurement and Prediction," Transactions ofASME: Journal of Engineering for Power, 1967. Smith, E. B. et al., "Study and Tests to Reduce Compressor Sounds of Jet Aircraft," TR DS-68-7, contract FA65WA-1236 to FAA, 1968, General Electric Co., Cincinnati, Ohio. Schaut, L. A., "Results of an Experimental Investigation of Total Pressure Performance and Noise Reduction of an Airfoil Grid Inlet," document D6-23276, 1969, Boeing Co., Seattle, Wash.

Dynamics of Slung Bodies Utilizing a Rotating Wheel for Stability

Abstract: filler, R. H., "Rotor Blade Harmonic Airloadings," AIAA Journal, Vol. 2, No. 7, July 1964, p. 1260. Curtiss, H. C. Jr., "The Use of Complex Coordinates in the study of Rotor Dynamics," Journal of Aircraft, Vol. 10, No. 5, May 1973, pp. 285-295. Ormiston, R; A. and Peters, D. A., "Hingeless Rotor Response with Non-Uniform Inflow and Elastic Blade Bending," Journal of Aircraft, Vol. 9, No. 10, Oct. 1972, pp. 730-736. Kerr, A. W., Potthast, A. J., and Anderson, W. D., "An Interdisciplinary Approach to Integrated Rotor/Body Mathematical Modeling," Mideast Region Symposium, American Helicopter Society, 1972. Hohenemser, K. H. and Crews, S. T., "Model Tests on Unsteady Rotor Wake Effects," Journal of Aircraft, Vol. 10, No. 1, Jan. 1973, pp. 58-60. Carpenter, P. J. and Fridovich, B., "Effect of a Rapid Blade Pitch Increase on the Thrust and Induced Velocity Response of a Full Scale Helicopter Rotor," TN 3044, Nov. 1953, NACA. Hohenemser, K. H. and Yin, S. K., "Some Applications of the Method of Miiltiblade Coordinates," Journal American Helicopter Society, Vol. 17, No. 4, July 1972, pp. 3-12. Dynamics of Slung Bodies Utilizing a Rotating Wheel for Stability

Measurements of the Boundary-Layer Growth in Annular Diffusers

Abstract: proper consideration must however by given to the added weight, reliability and cost factors. The data just presented covers Engine "A" with a relatively low tip speed fan. The next major step in the NASA/GE experimental Quiet Engine Program is to evaluate an engine incorporating Fan "C" which is representative of a high speed fan. Such an engine offers the advantage of having shorter length and fewer components, both of which permit important improvements in engine weight and engine cost. It is expected that the acoustic investigation of the Engine "C" technology demonstrator Will be pursued into the Spring of 1973. At that time comparable data will have been obtained on both "A" and "C" technology demonstrators and tradeoff studies will be performed to evaluate their relative merits acoustically and in terms of direct operating costs on a number of airplane applications. References

Design for Air Combat

Abstract: aircraft loss rates during a military conflict. This fleet de- terioration has been calculated by application of suitable Lancaster equations2 and a computerized war game. Points 2 and 3 are, in this paper, completely covered only for the battlefield air superiority role, in particular the combat air patrol mission. Section 2 of this paper de- scribes some weight-cost performance relationships de- rived from the systematic variation of important design parameters, such as wing loading, wing shape and thrust to weight ratio. Computations were based on state of the art rubberized engines. Design parameter values range from 0.5 to 1.7 thrust to weight ratio, 40-90 lb/ft2 wing loading and 1.2-2.4 max Mach number. A large group of consistent parametric fighter designs has been evaluated in a sequence of detection, intercept and close-in combat. Some of the relevant system param- eter combinations and their impact on these three phases of air-to-air combat are discussed in Sec. 3. Fighter designs with parameter combinations which achieve maximum individual combat results at a giveri level of system cost are discussed in Sec. 4. The effective- ness of these fighter designs has been analyzed in a fleet environment under constant budget cost. It is shown that an optimum combination of thrust/weight and wing loading can be found for a defined scenario. The concluding sec- tion tries to answer some of the questions which are vital in the design for air combat: 1) Can missile maneuver- ability compensate for lack of air vehicle maneuverabili- ty? 2) Are requirements for offensive and defensive com-

Preliminary Criteria for Predicting Departure Characteristics/ Spin Susceptibility of Fighter-Type Aircraft

Abstract: The results of a study conducted to determine criteria for predicting loss-of-control characteristics at near-stall angles of attack based on lateral-directional static stability parameters are presented. The Cn0,dynamic and aileron-alone divergence parameters are calculated for variations in lateral-directional static stability and departure characteristics, as predicted by these parameters, are then correlated with high angle-of-attack stability and control characteristics determined from model and full-scale flight tests. Results of this limited correlation indicate that the Cnp,dynamic and aileron-alone divergence parameters are promising as a preliminary criteria for predicting departure characteristics at near-stall angles of attack as well as spin susceptibility.

An unsteady wake model for a hingeless rotor

Abstract: A simple nonsteady wake model derived from the unsteady moment of the momentum equation for zero advance ratio is correlated with cyclic pitch frequency response tests conducted with a small hingeless rotor model. Two and three or more bladed rotor analyses are presented.