Showing papers in "Journal of Aircraft in 1974"
TL;DR: In this paper, a quasi-continuous method is developed for solving thin-wing problems, where the spanwise vortex distribution is assumed to be stepwise-constant, while the chordwise vortex integral is reduced to a finite sum through a modified trapezoidal rule and the theory of Chebyshev polynomials.
Abstract: A quasi-continuous method is developed for solving thin-wing problems. For the purpose of satisfying the wing boundary conditions, the spanwise vortex distribution is assumed to be stepwise-constant, while the chordwise vortex integral is reduced to a finite sum through a modified trapezoidal rule and the theory of Chebyshev polynomials. Wing-edge and Cauchy singularities are acounted for. The total aerodynamic characteristics are obtained by an appropriate quadrature integration. The two-dimensional results for airfoils without flap deflection reproduce the exact solutions in lift and pitching moment coefficients, the leading edge suction, and the pressure difference at a finite number of points. For a flapped airfoil, the present results are more accurate than those given by the vortex-lattice method. The three-dimensional results also show an improvement over the results of the vortex-lattice method. Extension to nonplanar applications is discussed.
173 citations
TL;DR: In this article, a series of extensive studies of the application of leading edge serrations as a device for reducing the vortex noise radiated from stationary and rotating airfoils in low Reynolds number flow are presented and interpreted.
Abstract: This paper presents and interprets a series of extensive studies of the application of leading-edge serrations as a device for reducing the vortex noise radiated from stationary and rotating airfoils in low Reynolds number flow. In these studies, a variety of serrations were attached at selected locations near the leading edge of stationary and rotating airfoils. The noise levels of the airfoils were reduced considerably with the serrations attached. An explanation of the aeroacoustic flow mechanisms involved is given.
131 citations
TL;DR: In this paper, the authors presented a summary of the radial planarity in terms of phase and amplitude of the probe readings from two of the rakes, 180° apart, at the engine IGV plane.
Abstract: Steady-state distortion levels produced during full scale PG operation as evaluated in terms of IDC (circumferential) and IDR (radial) distortion parameters were generally less 0.02 and nominally 0.01 or less as previously realized in the model test. Figure 10 summarizes the circumferential planarity in terms of phase and amplitude. The probe readings from two of the rakes, 180° apart, at the engine IGV plane were averaged to obtain these summary results. Such a comparison indicates excellent circumferential planarity over the complete range of the test data. A similar summary of radial planarity characteristics is presented in Fig. 11. In this case, all the probe readings from 2 selected radial immersions were averaged. Radial immersion B is the tip and D is the hub ring. The hub ring probes had the most deviation from those in any of the other rings, thus the results of Fig. 11 represent the "worst case."
70 citations
TL;DR: In this paper, an analysis is presented which yields an approximate solution for the unsteady aerodynamic response of an infinite swept wing encountering a vertical oblique gust in a compressible stream.
Abstract: An analysis is presented which yields an approximate solution for the unsteady aerodynamic response of an infinite swept wing encountering a vertical oblique gust in a compressible stream. The approximate expressions are of closed form and do not require excessive computer storage or computation time, and further, they are in good agreement with the results of exact theory. This analysis is used to predict the unsteady aerodynamic response of a helicopter rotor blade encountering the trailing vortex from a previous blade. Significant effects of three dimensionality and compressibility are evident in the results obtained.
70 citations
TL;DR: A conceptual wind-tunnel test program has been conducted to verify that blowing a stream of highpressure air over a swept-wing surface in a direction roughly parallel to the leading edge enhances the vortex system as discussed by the authors.
Abstract: A conceptual wind-tunnel-test program has been conducted to verify that blowing a stream of highpressure air over a swept-wing surface in a direction roughly parallel to the leading edge enhances the vortex system. The blowing is shown to intensify the leading-edge vortex and thus delay the deleterious effects of vortex breakdown to higher angle of attack. As a result, the vortex-lift is significantly increased and, as the blowing rate is increased, appears to approach the value predicted by the Polhamus suction-analogy for thin wings.
66 citations
TL;DR: In this article, a method was developed for the calculation of the initial inviscid form of rolled-up wake vortices behind a wing having arbitrary lift distribution, which made use of the Betz assumptions of the conservation of wake voricity and moments of vorticity.
Abstract: A method is developed for the calculation of the initial inviscid form of rolled-up wake vortices behind a wing having arbitrary lift distribution. The method makes use of the Betz assumptions of conservation of wake vorticity and moments of vorticity. It is found that a simple relationship exists between the radial distribution of vorticity in the rolled-up wake and the spanwise lift distribution. Computed tangential velocity profiles for DC-7, DC-9, and C-141 aircraft are shown to compare favorably with profiles measured by the FAA during tower flyby tests of these aircraft in both flapped and unflapped configurations.
62 citations
TL;DR: In this article, a method for predicting deployment and inflation of reefed ribbon parachutes is presented based on integration of axial and radial momentum equations developed in the paper, which is assumed to be describable by drag and radial force coefficients.
Abstract: A method for predicting deployment and inflation of reefed ribbon parachutes is presented. The method is based on integration of axial and radial momentum equations developed in the paper. Axial and radial forces are assumed to be describable by drag and radial force coefficients. Computer solutions of the equations are compared to measured parachute loads and to parachute mouth and maximum diameters from tests of 23and 76-ft-diam conical ribbon parachutes. Comparison of load histories indicates that snatch loads depend to a large extent on deployment bag design and packing influences. Computed loads and parachute size histories for the inflation process compared favorably with flight data. The concept of a radial force coefficient appears to have considerable merit as a means of computing inflation for most types of parachutes.
57 citations
TL;DR: In this paper, a two-dimensionala l laser velocimeter whose focal volume can be rapidly traversed through a flowfield has been used to overcome the problem introduced by excursions of the central vortex filament within a wind-tunnel test section.
Abstract: A two-dimensiona l laser velocimeter whose focal volume can be rapidly traversed through a flowfield has been used to overcome the problem introduced by excursions of the central vortex filament within a wind-tunnel test section. The operation of the instrument is reviewed and data are presented which accurately define the trailing vortex from a square-tipped rectangular wing. Measured axial and tangential velocity distributions are given, both with and without a vortex dissipator panel installed. From the experimental data, circulation and vorticity distributions are obtained and the effect of turbulence injection into the vortex structure is discussed. Nomenclature wing span wing chord lift coefficient radial coordinate from vortex center wind-tunnel mainstream velocity axial velocity component tangential velocity component maximum tangential velocity stream wise ordinate, aft from trailing edge normal ordinate, above upper wing surface from trailing edge normal location of vortex center angle of attack, deg slope of velocity distribution in core region vorticity
50 citations
TL;DR: In this article, an existing vortex-lattice method was modified to include the effects of leading-edge separation, and the modified version was then used to calculate the aerodynamic loads on a highly swept delta wing.
Abstract: Vortex-lattice methods have been used successfully to obtain the aerodynamic coefficients of lifting surfaces without leading-edge separation. It is shown how an existing vortex-lattice method can be modified to include the effects of leading-edge separation. The modified version is then used to calculate the aerodynamic loads on a highly swept delta wing. The results are compared with Peckham's (1958) experimental data.
48 citations
TL;DR: In this paper, a dynamic inflation model for parachutes is presented, which predicts increased dimensionless inflation times and increased dimensioness inflation forces observed at high altitudes at high altitude.
Abstract: This paper describes a dynamic inflation model for parachutes which predicts increased dimensionless inflation times and increased dimensionless inflation forces observed at high altitudes. As altitude is increased, greater relative parachute inertia results in increased inflation times, and greater relative total system inertia results in increased maximum inflation forces. The effect of Mach number on inflation force is also predicted by the inflation model.
45 citations
TL;DR: In this article, it was shown that for a higher order set of equations which has lift and thrust as controls, the necessary condition for optimization including the Generalized Legendre-Clebsch condition are satisfied at the cruise point so that the partial throttle cruise condition is a candidate solution for the minimum fuel-fixed range problem.
Abstract: IN the paper by Schultz and Zagalsky, the solution characteristics for the minimum fuel-fixed range problem are determined for a number of different mathematical models of aircraft dynamics. Different results are obtained for different sets of equations. The energy state equations are shown to not allow a partial throttle, constant velocity, cruise solution; but because the velocity set is not convex, to have a "chattering" solution which provides the best performance but which may not be realizable with plecewise continuous maximum values of the controls. The equation set with throttle and flight path angle as controls was shown to allow a partial throttle cruise solution. This conclusion was shown to be invalid by Speyer by application of the Generalized Legendre-Clebsch condition. The following analysis shows that for a higher order set of equations which has lift and thrust as controls, the necessary condition for optimization including the Generalized Legendre-Clebsch condition are satisfied at the cruise point so that the partial throttle cruise condition is a candidate solution for the minimum fuel-fixed range problem.
TL;DR: In this article, a new, fast, and economical automated procedure for implementing the traditional V-g method of flutter solution is described, which requires as input the generalized aerodynamic forces for a range of reduced frequencies obtained from an aerodynamic program.
Abstract: A new, fast, and economical automated procedure for implementing the traditional V-g method of flutter solution is described. The procedure requires as input the generalized aerodynamic forces for a range of reduced frequencies obtained from an aerodynamic program. These aerodynamic forces are interpolated with respect to reduced frequency using a newly developed, partially tabulated cubic spline that is both fast in execution and economical in storage. The flutter solution is then obtained using an eigenvalue routine that has been developed to take advantage of the parametric nature of the V-g type of solution. Furthermore, the routine takes care of the fundamental and troublesome problem of properly sorting the output eigenvalues. By solving the root-sorting problem, the interpolation for flutter crossings and automatic plotting are accomplished efficiently. The computational techniques used in this new program are described and some sample results are given.
TL;DR: In this paper, a generalized identification method is applied to flight test data analysis based on the maximum likelihood (ML) criterion and includes output error and equation error methods as special cases.
Abstract: Application of a generalized identification method to flight test data analysis. The method is based on the maximum likelihood (ML) criterion and includes output error and equation error methods as special cases. Both the linear and nonlinear models with and without process noise are considered. The flight test data from lateral maneuvers of HL-10 and M2/F3 lifting bodies are processed to determine the lateral stability and control derivatives, instrumentation accuracies, and biases. A comparison is made between the results of the output error method and the ML method for M2/F3 data containing gusts. It is shown that better fits to time histories are obtained by using the ML method. The nonlinear model considered corresponds to the longitudinal equations of the X-22 VTOL aircraft. The data are obtained from a computer simulation and contain both process and measurement noise. The applicability of the ML method to nonlinear models with both process and measurement noise is demonstrated.
TL;DR: In this paper, the performance of an inlet for an integrated scramjet engine concept at Mach 6 was evaluated in terms of integrated performance parameters, and the results of the inlet design and test model were presented.
Abstract: Review of the results of an experimental investigation of the performance of an inlet for an integrated scramjet engine concept at Mach 6. Following a description of the inlet design and test model, the Mach 6 experimental results obtained are presented in terms of integrated performance parameters.
TL;DR: In this paper, the authors considered boundary-layer behavior by applying the Stratford criterion for turbulent separation to numerically computed wall-pressure distributions and determined the minimum-length contraction shapes consistent with fully attached boundary layer flow.
Abstract: Most design methods for aerodynamic contractions are based on an inviscid approach even though the flow quality downstream of the contraction is determined primarily by viscous effects, including possible boundary-layer separation. The present study includes the consideration of boundary-layer behavior by application of the Stratford criterion for turbulent separation to numercially computed wall-pressure distributions. Calculations are presented for a four-parameter family of contractions within which the minimum-length contraction shapes consistent with fully attached boundary-layer flow are determined.
TL;DR: Melzig, H. D., "The Dynamic Pressure Loading on Parachute Canopies," AIAA Aerodynamic Deceleration Systems Conference, AIAa, Houston, Texas, 1966 as mentioned in this paper.
Abstract: Melzig, H. D., "The Dynamic Pressure Loading on Parachute Canopies," AIAA Aerodynamic Deceleration Systems Conference, AIAA, Houston, Texas, 1966. Poppleton, E. D., private communication, Univ. of Sydney, Sydney, Australia. Karmacheti, Principles of Ideal Fluid Aerodynamics, Wiley, New York, 1966. Wood, R. D., "Pressure Distributions on Parachutes During Inflation," Ph.D. thesis, Univ. of Sydney, 1970. Brown, C. E. and Michael, W. H., "On Slender Wings with Leading Edge Separation," TN 3430,1955, NACA. Ibrahim, S. K., "Apparent Added Mass and Moment of Inertia of Cup-Shaped Bodies in Unsteady Incompressible Flow," Ph.D. thesis, University of Minnesota, 1965.
TL;DR: In this article, a collocation technique is used with the nonplanar supersonic kernel function to solve multiple lifting surface problems with interference in steady or oscillatory flow, and the pressure functions used are based on conical flow theory solutions.
Abstract: In the method presented in this paper, a collocation technique is used with the nonplanar supersonic kernel function to solve multiple lifting surface problems with interference in steady or oscillatory flow. The pressure functions used are based on conical flow theory solutions and provide faster solution convergence than is possible with conventional functions. In the application of the nonplanar supersonic kernel function, an improper integral of a 3/2 power singularity along the Mach hyperbola is described and treated. The method is compared with other theories and experiment for two wing-tail configurations in steady and oscillatory flow.
TL;DR: In this paper, a simple computer program has been developed that determines the area development of the equivalent body of revolution required to minimize various sonic boom signature parameters, such as the overpressure signature.
Abstract: Means of reducing or eliminating the sonic boom through aerodynamic design or aircraft operation are discussed. These include designing aircraft to minimize or eliminate certain features of the overpressure signature, operating aircraft at slightly supersonic speeds so that the sonic boom does not reach the ground, and seeking reductions through the high-altitude, high-speed flight conditions of hypersonic transports. A simple computer program has been developed that determines the area development of the equivalent body of revolution required to minimize various sonic boom signature parameters.
TL;DR: Harder et al. as mentioned in this paper proposed a kernel function for nonplanar Oscillating Surfaces in Supersonic Flow to calculate the lift distribution of a single-wing aircraft.
Abstract: 108, March 1972, Air Force Flight Dynamics Lab., Wright-Patterson Air Force Base, Ohio. Harder, R. L. and Rodden, W. P., "Kernel Function for Nonplanar Oscillating Surfaces in Supersonic Flow," Journal of Aircraft, Vol. 8, No. 8, Aug. 1971, pp. 677-679. Cunningham, A. M., Jr., "The Application of General Aerodynamic Lifting Surface Elements to Problems in Unsteady Transonic Flow," CR-112264, Feb. 1973, NASA. Multhopp, H., "Methods for Calculating the Lift Distribution of Wings," (Appendix I, Contributed by W. Mangier), Rept. Aero-2353, January 1950, Royal Aircraft Establishment, Farnborough, Great Britain.
TL;DR: In this paper, the authors reviewed the theory and advantages of the cryogenic tunnel concept and briefly reviewed the characteristics of the Langley 34 cm (13.5 in.) Pilot Cryogenic Transonic Pressure Tunnel and the results of initial tunnel operation.
Abstract: The theory and advantages of the cryogenic tunnel concept are briefly reviewed. The unique ability to vary temperature independently of pressure and Mach number allows, in addition to large reductions in model loads and tunnel power, the independent determination of Reynolds number, Mach number, and aeroelastic effects on the aerodynamic characteristics of the model. Various combinations of Reynolds number and dynamic pressure can be established to accurately represent flight variations of aeroelastic deformation with altitude changes. The consequences of the thermal and caloric imperfections of the test gas under cryogenic conditions have been examined and found to be insignificant for operating pressures up to 5 atm. The characteristics of the Langley 34 cm (13.5 in.) Pilot Cryogenic Transonic Pressure Tunnel are described and the results of initial tunnel operation are presented. Tests of a two-dimensional airfoil at a Mach number of 0.85 show identical pressure distributions for a chord Reynolds number of 8.6 X 106 obtained first at a stagnation pressure of 4.91 atm at a stagnation temperature of +120°F and then at a stagnation pressure of 1.19 atm at a stagnation temperature of -250°F.
TL;DR: In this article, a six-degree-of-freedom nonlinear digital simulation was used to assess the effectiveness of automatic control systems in alleviating vortex wake upsets, and the results of this preliminary study indicate that it is feasible to use an automatic control system to alleviate vortex encounter upsets.
Abstract: The problem of an airplane being upset by encountering the vortex wake of a large transport on takeoff or landing is currently receiving considerable attention. This paper describes the technique and results of a study to assess the effectiveness of automatic control systems in alleviating vortex wake upsets. A six-degree-of-freedom nonlinear digital simulation was used for this purpose. The analysis included establishing the disturbance input due to penetrating a vortex wake from an arbitrary position and angle. Simulations were computed for both a general aviation airplane and a commercial jet transport. Dynamic responses were obtained for the penetrating aircraft with no augmentation and with various command augmentation systems. The results of this preliminary study indicate that it is feasible to use an automatic control system to alleviate vortex encounter upsets.
TL;DR: In this paper, a number of models were tested at subsonic Mach numbers through a range of unseparated, partially separated, and totally separated flow, and static pressure taps were used to determine the surface pressure distribution in the separated region.
Abstract: Separated flow data for axisymmetric afterbodies are presented. A number of models were tested at subsonic Mach numbers through a range of unseparated, partially separated, and totally separated flow. The separation and reattachment locations on the afterbody were determined by oil flow visualization techniques, and static pressure taps were used to determine the surface pressure distribution in the separated region. The data systematically show the effect of shape, Mach number, total pressure and approach boundary-layer thickness upon the onset and extent of flow separation over afterbody models. Existing separation criteria are compared to the data; and an engineering model is proposed for use in predicting the effects of the separated region upon afterbody pressure distribution.
TL;DR: Applications of a multidisciplinary system of computer programs to selected analytical and optimization problems encountered in an aircraft design are described, showing how the depth of the analyses and optimizations permits such detailed output as individual stringer sizes and pressure distributions.
Abstract: Applications of a multidisciplinary system of computer programs to selected analytical and optimization problems encountered in an aircraft design are described The depth of the analyses and optimizations permits such detailed output as individual stringer sizes and pressure distributions Data flow from one program to another is performed in a hands-off manner in the modular and open-ended system Numerical examples show how this automation permits multidisciplinary trade-off studies, typical of a preliminary design process to be based on such a level of detail in each discipline that normally would not be available at this stage of the process
TL;DR: In this article, a simplification in the solution of Prandtl's lifting line equation is presented, which gives an approximate solution along the entire span, so that discontinuities, flap deflections, etc., can be accounted for.
Abstract: The present work offers a simplification in the solution of Prandtl's lifting line equation. The equation for the local circulation is usually solved using a sine series in a collocation method. Using the fact that local circulation and local geometric angle of attack are related by a linear operator, an expression can be obtained for the drag coefficient containing only the square of the unknown constants, which implies that the loadings are orthogonal in Graham's sense (1952). Expressions are derived for the unknown constants, making use of the orthogonality of the loadings. The solution gives an approximate solution along the entire span, so that discontinuities, flap deflections, etc., can be accounted for.
TL;DR: In this article, Love's first-approximation shell theory is used, and a Rayleigh-Ritz solution is obtained for finite-length sandwich cylinders clamped at both edges and loaded in axial compression, pure bending or a combination of these loadings.
Abstract: The buckling analysis reported here is for finite-length sandwich cylinders clamped at both edges and loaded in axial compression, pure bending, or a combination of these loadings. Love's first-approximation shell theory is used, and a Rayleigh-Ritz solution is obtained. For axially compressed sandwich cylinders with clamped boundary conditions, the theory predicts values above published experimental results. For pure bending, predicted results are considerably above the experimental results; however, the experimental data represent a very limited number of specimens. Results for pure bending and combined bending and axial compression are in general agreement with published theoretical results for isotropic thin-walled simply-supported cylinders.
TL;DR: In this paper, the velocity at the outer edge of the bottom plate is estimated by a Bound-Ary layer type approximation with a velocity distribution given by the velocity distribution on the bottom surface.
Abstract: Results Figures 3 and 4 illustrate the plot of the velocity maxi- ma predicted by the above method compared with the measurements of Schauer and Eustis.6 The nondimen- sional static pressure on the bottom wall is also calculated by using Bernoulli's theorem. p + 1/2 pw2 = const on a streamline. The agreement with the experiments of Refs. 3 and 6 is good for the measured static pressure dis- tribution. It is also seen from the above graphs that a nondimensional profile could be obtained for various nozzle distances. The computing time involved in the above calculations were of the order of 5 sees. The bound- ary layer close to the wall could be described by a Bound- ary layer type approximation with a velocity at the outer edge given by the velocity distribution on the bottom plate. Conclusions Considering the normal impinging jet flow as consisting of two different regimes an expeditive method of calcula- tion has been developed. The present method takes into account the viscosity effects producing a rotation in the flow, and also it requires an incomparably shorter com- puter time than needed for the solution of the full Navier- Stokes equation. The accuracy of the predictions by using this method is as good as the predictions assumed by more time consuming methods.
TL;DR: In this article, both the direct and inverse jet flap airfoil potential flow problems are described and compared with the results of previous linear and nonlinear methods as well as with experimental data.
Abstract: Methods for solving both the direct and inverse jet flap airfoil potential flow problems are described. The direct airfoil analysis method is a completely nonlinear iterative method which is applicable to either thick or thin airfoils of arbitrary shape. The very general surface singularity formulation has been extended to include multielement airfoils, ground effects, nonuniform freestreams, inlet flows, jet entrainment effects, etc. Comparisons are given with the results of previous linear and nonlinear methods as well as with experimental data. The inverse (design) method is a more approximate method in which camber and thickness distributions are designed separately. Section shapes are shown for several airfoils designed to have only very small regions of adverse pressure gradient. Nomenclature c = length of airfoil chord cp = coefficient of pressure d = coefficient of lift cu = coefficient of jet momentum h = height of airfoil leading edge above ground plane R = radius of curvature of the jet sheet s = coordinate along the jet sheet t = airfoil thickness V = local flow speed V = average flow speed across a vortex sheet Vj - jet flow speed Vn = component of velocity normal to a surface V = freestream flow speed x = coordinate parallel to the freestream = coordinate perpendicular to the freestream = jet deflection angle at the trailing edge relative to the airfoil chord line 7 = strength of a vortex sheet 0 = local angle of inclination of the jet sheet relative to the freestream 0 = velocity potential