Showing papers in "Journal of Aircraft in 1972"

Interpolation using surface splines.

Abstract: A surface spline is a mathematical tool for interpolating a function of two variables. It is based upon the small deflection equation of an infinite plate. The surface spline depends upon the solution of a system of linear equations, and thus, will ordinarily require the use of a digital computer. The closed form solution involves no functions more complicated than logarithms, and is easily coded. Several modifications which can be incorporated are discussed.

Analysis and Correlation of Data on Pressure Fluctuations in Separated Flow

Abstract: The surface pressure fluctuations caused by bubbles at subsonic speeds are described and related with the mean pressure within the bubble and the development of the mixing layer. The pressure fluctuations caused by bubbles increase gradually from the separation line, reach a maximum near the reattachment line and then decrease gradually downstream of the reattachment line. Spectra of the pressure fluctuations near the reattachment line are similar for bubbles caused by leading-edge separation on wings, by forward facing steps, by rearward facing steps, by sudden enlargements in pipes and by cavities, if all the spectra are expressed in terms of a frequency parameter based on the bubble length. These observations should give a fairly good preliminary design method for evaluating fluctuating pressures.

Aircraft performance optimization.

Abstract: Using the calculus of variations, the solutions to various fixed end point flight-path optimization problems are developed These include the minimum fuel-fixed range problem, the minimum time-fixed range problem, and the minimum fuel-fixed range-fixed time problem Altitude profiles and throttle control laws are presented A variety of aircraft mathematical models is initially examined, and the existence of a classically optimal controller is verified for a simple model For this model, the first integral condition is used to eliminate the requirement of integrating the Euler Lagrange adjoint differential equations The resulting computational algorithms are attractive for both laboratory investigations and airborne implementations

Calculation of Separation Points in Incompressible Turbulent Flows

Abstract: The purpose ohis paper ivaluate tccuracy with which tocation ourbulent separation c be predicted owo-dimensional ac bodies. Tvaluation wade btudying aonsiderable number of flows that had separation. Calculate d separation points were compared with the experimentally measured location. Fou r methods of predicting separation in turbulent flow were evaluated. The y were Goldschmied' s method, Stratford's method, Head's method, and the Cebeci-Smit h method. I t wa s concluded from the study that the last three listed methods predict separation points with the reliability and accuracy needed for aerodynamic design purposes.

Refinement of the nonplanar aspects of the subsonic doublet-lattice lifting surface method.

Abstract: The initial formulation of the Doublet-Lattice Method has proven theoretically accurate for calculating the interference effects on arbitrary classes of oscillating nonplanar configurations with one known exception: nearly coplanar wing/horizontal-tail combinations. For this class of problems the integration of the kernel across the element loses accuracy. The reason for this loss of accuracy is explained, and a refined method for performing the required integration is presented. Numerical studies using wing-tail configurations are presented to illustrate how well the refined method works. Also a T-tail calculation is repeated to show that calculations using other configurations are unaffected by the refinement.

Wind-Tunnel Studies of Wing Wake Turbulence

Abstract: Description of velocity measurements made in the wake of wings in the Ames 7 x 10 ft wind tunnel. Distributions of velocity components were measured with a three-wire anemometer up to 12 chord lengths downstream of a CV-990 aircraft model and a rectangular wing. Results show that increasing the drag increases the vortex core radius, reduces the maximum tangential velocities, and increases the magnitude of axial velocity defects. For the rectangular wing, axial velocity changes from a defect (wake flow) for angles of attack less than 9 deg to an excess (jet flow) for angles of attack greater than 9 deg. Wind-tunnel measurements of the near flowfield are compared with flight measurements of the far flowfield.

Subsonic Steady and Oscillatory Aerodynamics for Multiple Interfering Wings and Bodies

Abstract: A technique for predicting steady and oscillatory aerodynamic loadings on general configurations is presented for use in flutter, gust, and static aeroelastic analyses and estimation of static and dynamic stability derivatives. The procedures are based on the Doublet-Lattice Method and the method of images. Chordwise and spanwise load distributions on lifting surfaces and longitudinal load distributions on bodies are determined for configurations that consist of an assemblage of bodies (with variable circular or elliptic cross sections) and lifting surfaces (with arbitrary planform and dihedral, with or without control surfaces). Extensive comparisons are made with steady and oscillatory experimental lifting pressure data.

A Mixed Optimization Method for Automated Design of Fuselage Structures

Abstract: A procedure for automating the design of transport aircraft fuselage structures has been developed and implemented in the form of an operational program. The structure is designed in two stages. First, an over-all distribution of structural material is obtained by means of optimality criteria to meet strength and displacement constraints. Subsequently, the detailed design of selected rings and panels consisting of skin and stringers is performed by mathematical optimization accounting for a set of realistic design constraints. As a result one obtains a procedure whose practicality and computer efficiency is demonstrated on cylindrical and area-ruled large transport fuselage.

A Stability Analysis for Tethered Aerodynamically Shaped Balloons

Abstract: This work investigates the dynamic stability of tethered, aerodynamically shaped balloons by considering the system to pose essentially a cable problem, with the balloon's dynamics giving end and auxiliary conditions. This physical model gives a first-order problem in a sequence of partial differential wave equations with nonhomogeneous boundary conditions. Further, these equations uncouple to give a "lateral" problem and a "longitudinal" problem—as in first-order airplane dynamics. The solution of either problem takes the form of a transcendental characteristic equation for the stability roots, from which these roots are extracted by using an electronic computer and a roots locus plot. Further, this theory was applied toward the development of a highperformance tethered balloon design, and the results showed that good stability was attainable by the use of large and aerodynamically efficient fins.

Effect of Streamwise Vortices on Wake Properties Associated with Sound Generation

Abstract: T important aspects of the viscous wake leading to the generation of sound sources are: 1) unsteady twodimensional disturbances such as the Karman vortex street, and 2) unsteady forces generated by the passage of a rotor blade through the two-dimensional wake of a stator, or vice versa. The unsteady wake of an aircraft makes a significant contribution to the total noise during the landing approach, whereas the unsteady forces due to wake impingement on stator and rotor blades contribute significantly during takeoff and cruise. The latter effect, analyzed by Kemp and Sears, is responsible for about half of the pure-tone noise generated by turbomachines; the other half originates with the mutual interference between the rotor and stator flowfields. Wind-tunnel tests reported here show that the streamwise vortices shed from vortex generators near the trailing edges of a flat plate and an airfoil have a strong suppression effect on the formation of the Karman vortex street, and reduce the "effective" area of the velocity deficit in the wake, as seen by the following blade in a turbomachine. Both effects tend to suppress significantly the formation of sound sources.

Three-dimensional, minimum-time turns for a supersonic aircraft

Abstract: Altitude, bank angle and thrust program for minimizing time required by supersonic aircraft to turn through specified heading angle and reach required energy

Investigation of Nonlinear Motion Experienced on a Slender-Wing Research Aircraft

Abstract: A combined experimental and theoretical study is described of the nature of the nonlinear lateral oscillation experienced by a slender-wing aircraft at high angles of attack. Flight tests of the Handley Page 115 research aircraft confirmed the expectation that the Dutch roll oscillation is undamped at high angles of attack, and also showed that a limit cycle develops, with steady amplitude in bank angle of about 30°. The measured stability derivatives are given, together with the responses obtained from the RAE flight dynamics simulator, in which the digital computation of the equations of motion uses the static aerodynamic data directly from a controlled model in a wind tunnel. The motion is also analysed theoretically using a new approximate method for obtaining solutions for nonlinear differential equations. The analysis gives the conditions sufficient for the existence of a sustained oscillation, and its amplitude and frequency in terms of the aerodynamic and inertia properties of the aircraft.

Structural Considerations of the Human Vertebral Column Under +Gz Impact Acceleration

Abstract: : Among the major limitations on manned aerospace vehicular-and escape system designs are the structural limits of the human body. One of the lower limits is the strength of the vertebral body under +G (eyeballs down) impact acceleration. When the vertebral column is considered as a structural member, a finite limit on nonfracturing accelerations can be specified, as has previously been the case. An hypothesis as to the mechanism of fracture, which suggests an approach capable of raising the limit, and experimental evidence in support of the hypothesis are presented. A crude device based upon the approach was designed and tested experimentally on cadaveric exposures to +G acceleration. A statistically significant increase in the level of acceleration required to cause fracture was measured.

Hingeless Helicopter Rotor Response with Nonuniform Inflow and Elastic Blade Bending

Abstract: A generalized harmonic balance theory for the steady-state, linear, response characteristics of a hingeless rotor in forward flight is presented and evaluated with the aid of recent experimental data. Comparisons of several approximate representations for the rotor blade revealed that the simple rigid hinged blade is inadequate except at very low advance ratios, or high flap frequencies. For typical values of flap frequency the first two elastic flap bending modes are required for accurate response predictions. Simplified models of the nonuniform induced inflow were derived, using momentum arid vbrtex theory, and found to be the most important factor in improving correlation with the data. An empirical inflow model was developed from the experimental data by obtaining an inverse solution of the present theory. This inflow model was found to be in reasonable agreement with the simple inflow theories for low advance ratios but revealed large unexpected variations in the induced inflow of the rotor for advance ratios near 0.8. a B b c CT d Cm e epc El J [L] m [M] [N] n p r, R U v'i v0 w

Adaptive Model Following Systems for Flight Control and Simulation

Abstract: A method for the design of adaptive model following control systems has been developed using a hyperstability approach. The derivation of the adaptation algorithms is presented. The implementation of the algorithm is realized by a combination of linear filters with some positivity properties and of the multipliers which processes the model-plant error. The application of this method to the design of flight control and simulation systems is discussed. The feasibility and advantages of the procedure are illustrated by applying it to an aircraft longitudinal control problem.

Prediction of the Spectral Content of Combination Tone Noise

Abstract: Combination tone noise is generated by a pattern of rotating shock waves in the inlet of turbofan engines when the relative tip speed of the fan blades is supersonic. A method is presented for determining the expected distribution of sound power among the harmonics of engine rotation frequency, based on the spectral analysis of an almost periodic succession of pulses. The spectral distribution of combination tone noise is found to depend on two statistical — crfl, the standard deviations of the sequence of shock wave amplitudes; and cr£, the standard deviation of the sequence of time intervals between successive shock waves. The spectral distribution of sound power is found to depend more critically on

Theory and Flight Verification of the TIFS Model-Following System

Abstract: The content of this paper describes the theoretical development and flight-test results of the model-following control system of the Air Force Total In-Flight Simulator (TIFS). A discussion of the conceptual design and detailed development of the system configuration is given. The manner in which the feedforward, gust compensator and lateral-directional feedback gains are obtained is developed. The feedforward and gust compensator gains are obtained by simple matrix algebra calculations. A sensitivity minimization approach using modern control theory is used to obtain the lateral-directional feedback gains. Digital simulation results are included to show the improvement in model-followi ng achieved with the feedback gains determined by this approach. Time histories of the model and TIFS responses from flight test are also included to show the quality of modelfollowing obtained with the system for both the lateral-directional and longitudinal modes of operation. These results verify the theory and design procedure used to obtain the TIFS model-following control system.

Low-Area Ratio, Thrust-Augmenting Ejectors

Abstract: The thrust augmentation, lift augmentation, and noise reduction characteristics of compact ejectors make them potentially attractive for propulsion lift systems; however, in the past, poor thrust augmentation results have negated the other benefits. A synthesis of an ejector's internal flow phenomena developed in this paper indicates that improved mixing and diffusion can significantly increase thrust augmentation. Experiments with an ejector employing a "hypermixing" primary nozzle confirm the analytical model's augmentation predictions and show reasonable agreement with other predicted flow characteristics. The two-dimensional primary nozzle has its slotshaped exit throat divided into alternating segments that impart left and right deflections to the primary flow to accelerate mixing. The results indicate that a properly designed ejector should give a thrust augmentation ratio near 1.075 + O.Q25(A3/AQ\ for 5 < A3/A0 < 14, where A3/A0 is the ratio of the ejector's exit area to the primary nozzle area.

Automated procedures for sizing aerospace vehicle structures /SAVES/

Abstract: Results from a continuing effort to develop automated methods for structural design are described. A system of computer programs presently under development called SAVES is intended to automate the preliminary structural design of a complete aerospace vehicle. Each step in the automated design process of the SAVES system of programs is discussed, with emphasis placed on use of automated routines for generation of finite-element models. The versatility of these routines is demonstrated by structural models generated for a space shuttle orbiter, an advanced technology transport,n hydrogen fueled Mach 3 transport. Illustrative numerical results are presented for the Mach 3 transport wing.

Aeroelastic Optimization of a Panel in High Mach Number Supersonic Flow

Abstract: Solution for a least-weight skin thickness distribution for a panel with a flutter parameter constraint. This panel weighs less than any similar constant thickness panel, but has the same critical supersonic panel flutter parameter. The panel rests on simple supports and is of sandwich construction. The span to chord ratio is large enough that the inertial, elastic, and aerodynamic behavior is one-dimensional. The Mach number is great enough that the aerodynamic forces acting on the upper panel surface may be accurately described by quasi-steady, linearized, supersonic aerodynamic theory. The final optimum design is obtained from theoretical and numerical methods adapted from optimal control theory. The results of this investigation show that the optimal panel thickness distribution is symmetric about the panel chord midpoint. Compared to a reference panel with constant thickness, optimum panels are found to be nearly 12% lighter.

A Feasibility Study of Active Wing/Store Flutter Control

Abstract: Analytical investigations of active feedback flutter control for fighter type aircraft, specifically with respect to wing/store flutter control, show promise of significant benefits for both contemporary and future aircraft. The F-4 Phantom aircraft with an external store is idealized for a flutter critical configuration. Computer programs, based on both frequency and time domains, are used with conventional control system design techniques to generate feedback compensation for active control of flutter for this configuration. Results of linear analyses indicate the possibility of expanding the permissible flight envelope by 150 knots using the existing aileron control surfaces and establish preliminary requirements for control system hardware.

Alleviation of the Sonic Boom by Thermal Means

Abstract: Previous studies suggest that reduction of SST sonic boom noise to levels acceptable to the public requires elimination of shocks and creation of pressure rise times of 10 msec or more at the ground. Because it appears impractical to lengthen the aircraft sufficiently to achieve this objective, a feasibility investigation of the use of heat to simulate a long body was undertaken. The large power requirement dictates the use of direct combustion of cheap fuel for this purpose, and practical considerations motivate off-axis heating. This can be accomplished in principle by the use of either a thermal spike with its axis parallel to the airstream or a thermal keel that extends downward from the aircraft.

Analysis of Supersonic-Hypersonic Flutter of Lifting Surfaces at Angle of Attack

Abstract: Several important features of supersonic-hypersonic flutter of lifting surfaces at angle of attack are highlighted in an exploratory study. Three simple analytical methods—a modified strip analysis incorporating shock expansion theory, modified Newtonian flow theory, and local flow piston theory—are employed in order to avoid the usual limitations of piston theory and Newtonian flow theory with regard to Mach number range, angle-of-attack range, and airfoil section. Illustrative flutter calculations have been made for two rectangular wings with diamond airfoils and with pitch and translational degrees of freedom. Results from the modified strip analysis and from the modified Newtonian flow theory are in reasonable agreement with limited available experimental data for Mach number 10, but results from the local flow piston theory are unconservative. The calculations show a typical degradation of flutter-speed index with increasing angle of attack, which at low angles of attack is most pronounced for thin sections. For some airfoil shapes, however, a forward location of the center of gravity may mitigate the degradation at low to moderate angles of attack and essentially postpone it until shock detachment conditions are approached. As angle of attack is increased, or Mach number is decreased toward the shock detachment condition, a sharp drop in flutter speed index is predicted by the methods that account for detachment. Thus, the vicinity of shock detachment is indicated to be a critical region for supersonic-hypersonic flutter.

Negative Magnus forces in the critical Reynolds number regime.

Abstract: A spinning two-dimensional cylinder experiences a negative Magnus lift force at small spin ratios (peripheral speed/freestream speed) in the critical Reynolds number regime in steady incompressible flow. A comprehensive explanation of this phenomenon is offered. A negative Magnus side force at a small spin ratio has been measured on an inclined, spinning ogive-cylinder at a certain subcritical cross-flow Reynolds number. Through use of the impulsive flow analogy, the two-dimensional explanation can be applied to the negative side force experienced by the inclined spinning ogive-cylinder if the effect of the body vortices is allowed for. At both critical and subcritical cross-flow Reynolds numbers, the body vortices produce a side force which is opposed to, and larger than, the side force associated with the cross-flow separation positions. The cross-flow separation positions are similar to those on a spinning two-dimensional cylinder. Starting with the measured separation and body vortex positions, a simple potential flow model is developed, which predicts total side forces in qualitative agreement with those measured.

Improved Navigation by Combining VOR/DME Information and Air Data

Abstract: Navigational accuracy improvement by combining VOR/DME information with airspeed and heading data via maximum likelihood filter, using small airborne computer