Showing papers in "Journal of Aircraft in 1980"

Role of Shocks in the "Sub-Transonic" Flutter Phenomenon

Abstract: A semi-quantitative investigation is reported on the influence of partial-chord transonic shocks on flutter of "typical-section" wing models. Unsteady airloads are assumed as the sum of linearized theory and a "shockforce doublet" centered at the measured steady shock location. The shock is shown usually to destabilize singledegree pitching motion; it may affect flexure-torsion flutter either way, often profoundly. Various typicalsection parameters are studied, along with the important phase lag known to be present in the shock oscillation. Energy transfer during flutter is examined. Simplified calculations are presented that are believed relevant to the transonic tests by Farmer & Hanson.

Numerical lifting line theory applied to drooped leading-edge wings below and above stall

Abstract: A numerical iterative solution to the classical Prandtl lifting-line theory, suitably modified for poststall behavior, is used to study the aerodynamic characteristics of straight rectangular finite wings with and without leading-edge droop. This study is prompted by the use of such leading-edge modifications to inhibit stall/spins in light general aviation aircraft. The results indicate that lifting-line solutions at high angle of attack can be obtained that agree with experimental data to within 20%, and much closer for many cases. Therefore, such solutions give reasonable preliminary engineering results for both drooped and undrooped wings in the poststall region. However, as predicted by von Karman, the lifting-line solutions are not unique when sectional negative lift slopes are encountered. In addition, the present numerical results always yield symmetrical lift distributions along the span, in contrast to the asymmetrical solutions observed by Schairer in the late 1930's. Finally, a series of parametric tests at low angle of attack indicate that the effect of drooped leading edges on aircraft cruise performance is minimal.

Divergence of Forward Swept Composite Wings

Abstract: Forward swept wing aircraft may have superior aerodynamic performance for certain missions. Algebraic expressions to predict the static aeroelastic divergence characteristics of forward swept wings constructed of composite materials have been developed using a laminated box beam model to describe the wing structure and aerodynamic strip theory to predict the loads due to wing bending and torsional deformation. The expressions presented show that, because of elastic coupling between wing bending and torsion, wing divergence may be precluded for reasonably large forward sweep angles if the composite structure is properly tailored. The structural parameters that maximize divergence speed are readily identified. Two illustrative examples are presented. ao b

Dynamic Stall at High Frequency and Large Amplitude

Abstract: A previously developed quasisteady analytic method has been shown to give predictions that are in good agreement with experimental stall results as long as the oscillation amplitude and frequency are of moderate magnitudes. In the present paper, this quasisteady method is extended to include the transient effect of the "spilled" leading-edge vortex, thereby providing simple analytic means for prediction of dynamic stall characteristics at high frequency and large amplitudes. The veracity of the method is demonstrated by critical comparisons with the extensive experiments performed by Carr et al. c Ka / m

Rotor blade cooling in high pressure turbines

Abstract: A simplified model of a cooled turbine blade is used to illustrate the important features of cooling. Progress from early convection-cooled blades through to today's blades is traced with reference to the Spey and the RB 211 engines. Consideration of the internal cooling shows that where maximum cooling effectiveness is wanted the Stanton number/friction factor ratio is important. A modified Spalding and Patankar theory predicts the external heat flux in an uncooled static cascade, but empirical factors are needed when this theory is applied to engine film-cooled blades. Operational experience of the Spey HP turbine blade is discussed. A new directionally solidified multipass turbine blade has been designed and developed for the RB 211 offering performance, temperature, and life improvements. The optimization of cooling, aerodynamics, stress, and manufacture requirements plays a crucial part in the design of such a blade. Finally, several new material and process developments have recently appeared which together with cooling advances should insure continued improvements to future cooled blades.

Application of the fast Fourier transform to two-dimensional wind tunnel wall interference

Abstract: Wall interference corrections are evaluated from experimental wind tunnel wall pressure distributions using the Fourier solution for the Dirichlet problem in a rectangle. The series coefficients are computed by the fast Fourier transform, making the method very efficient and suitable as a practical wall correction procedure for two-dimensional wind tunnel data. The method is applicable to arbitrary subcritical wind tunnel walls and the knowledge of their cross-flow properties is not required. A practical example is given for the BGK 1 airfoil, tested at supercritical flow conditions in the 20% perforated wall test section of the NAE high Reynolds number wind tunnel.

Analysis and Design of Strake-Wing Configurations

Abstract: The technology is still evolving for improving the transonic maneuver capability of strake-wing configurations. Much of the work to date has been of an experimental nature; whereas, the theories that are available to handle vortex-flow aerodynamics have mostly treated wings of constant sweep. Hence, two efforts were undertaken: 1) to extend one method—the suction analogy—to more general configurations and evaluate it by using selected critical planforms; and 2) to develop a procedure for strake planform shaping and test the resulting shape in conjunction with a wing-body. The conclusions from this study are that 1) some improvement has been made in estimating high angle-of-attack longitudinal aerodynamics, and 2) the gothic strake designed with the developed procedure does produce a stable vortex system in the presence of a wing body and flat postmaximum lift characteristics.

Vibration Measurements on Planetary Gears of Aircraft Turbine Engines

Abstract: Planetary reduction gear stages are an efficient, compact, and lightweight means of speed reduction and are, therefore, used in many aircraft turbine engines. In the PT6 engines, of which more than 13,000 have been produced, the planetary reduction gearboxes allow a speed reduction from more than 30,000 to about 6000 rpm in turboshaft and 1210-2200 rpm in turboprop applications. These gearboxes have been developed to a high level of reliability. The continuing demand for uprated or new higher powered designs requires a good understanding of the design factors that play a role in the dynamic gear loads and motions. The theoretical dynamic analysis of a planetary gear stage is quite complex due to the multiple, nonlinear gear meshes. Therefore, the development of these gearboxes depends to a high degree upon engine running experience and detailed dynamic measurements. A variety of dynamic measurements taken on PT6 reduction gearboxes over a number of years is reviewed. Peculiar behavior found in these tests is discussed, such as load sharing among planets, responses due to gear errors, and a dynamic instability.

Flutter analysis of a NACA 64A006 airfoil in small disturbance transonic flow

Abstract: Flutter analyses are performed for a NACA 64A006 airfoil pitching and plunging in small-disturbance, unsteady transonic flow. Aerodynamic coefficients are obtained for M^ = 0.7, 0.8, 0.85, 0.8625, and 0.87, and for various values of low reduced frequencies. Two computer codes are used: 1) STRANS2 and UTRANS2 based upon the relaxation method and 2) LTRAN2 based upon the time-integration (indicial) method. Flutter results are presented as plots of flutter speed and corresponding reduced frequency vs one of the four parameters: airfoil/air mass density ratio, position of mass center, position of elastic axis, and freestream Mach number. In each figure, several sets of curves for different values of plunge/pitch frequency ratios are shown. The two sets of results based upon the two separate computer codes are, in general, in good agreement. For a special flutter analysis of a flat plate at M^ =0.7, the present methods agree well with the linear flat plate theory. The effect of each parameter on the trend of each curve of flutter speed is discussed in detail. The examples demonstrate the dip phenomenon of the curves for flutter speed in the transonic regime. Flutter results of transonic codes are compared with those obtained by linear flat plate theory.

Decoupler Pylon: A Simple, Effective Wing/Store Flutter Suppressor

Abstract: As an alternative to alleviating wing/store flutter by conventional passive methods or by more advanced active control methods, a quasi-passive concept, referred to as the decoupler pylon, is investigated which combines desirable features of both methods. Passive soft-spring/damper elements are used to decouple wing modes from store pitch modes, and a low-power control system maintains store alignment under changing mean loads. It is shown by analysis and wind tunnel tests that the decoupler pylon provides substantial increase in flutter speed and makes flutter virtually insensitive to inertia and center-of-gravity location of the store.

Some Features of the Unsteady Pressure Field in Transonic Airfoil Buffeting

Abstract: Comparisons are made between the unsteady transonic flowfields of two airfoils: a Whitcomb supercritical airfoil, and a conventional NACA 0012 section. Wind tunnel experiments on these airfoils included penetration into buffeting as a result of high cl and/or high M^. Fluctuating surface pressure, lift, and shock location were measured on both airfoils. Two-point pressure cross-correlations were used to determine coherence and propagation direction of pressure fluctuation patterns on the upper surface of each airfoil. Between the uppersurface shock and the trailing edge (a region with intense pressure fluctuations), pressure disturbances propagated upstream in attached flow, but traveled downstream when extensive separation existed. In the latter case, convection velocities were found to be frequency dependent. Another cross-correlation, relating surfacepressure fluctuations to unsteady lift, was employed to establish which regions of the pressure fields were of primary importance in producing buffeting forces. While many of the principal features of the pressure/lift cross-correlations were common to both airfoils, some specific differences were found. For example, the supercritical airfoil exhibited less periodicity in its cross-correlation. This result was attributed to the flattopped, aft-cambered shape of the supercritical airfoil section, which reduced the coupling between shock oscillations and lift fluctuations.

Prandtl's biplane theory applied to canard and tandem aircraft

Abstract: Prandtl's biplane theory for elliptic loadings is generalized to apply to nonelliptic spanwise load distributions. The induced drag is calculated by assuming an infinite stagger distance so all of the mutually induced drag acts upon the rear surface which has no effect upon the front surface. Consequently, the mutually induced drag is calculated by integrating the Trefftz-plane downwash of the front surface over the independent load distribution on the rear surface. This procedure is verified by explicit solutions that give the same mutually induced drag irrespective of the fore and aft location of the larger span when carrying either an elliptic or a uniform load distribution. It was found that the mutually induced drag was less when the larger span had a uniform load distribution, but the total induced drag was not decreased because of the additional self-induced drag produced by the change from the ideal elliptic loading to a uniform loading. However, when the larger span carried a uniform loading it allowed the smaller span, when either fore or aft, to support more of the aircraft's weight at the minimum induced drag condition.

Airframe Noise Component Interaction Studies

Abstract: Acoustic wind-tunnel tests were conducted to examine the noise-generating processes of an airframe during approach flight. The airframe model was a two-dimensional wing section, to which high-lift leading and trailing edge devices and landing gear were added. Far-field conventional microphones were utilized to determine component spectrum levels. An acoustic mirror directional microphone was utilized to examine differences in noise source distributions on airframe components extended separately and in combination. Measured spectra are compared with predictions inferred from aircraft flyover data. Aeroacoustic mechanisms for each airframe component are identified. Component interaction effects on total radiated noise generally were small (within about 2 dB). However, some interactions altered local flow velocities and turbulence levels, causing redistribution of local acoustic source strength. Possibilities for noise reduction exist if trailing edge flaps could be modified to decrease their noise radiation caused by incident turbulent flow.

Three-Dimensional Shock Structures for Transonic/Supersonic Compressor Rotors

Abstract: This paper reviews experience at evaluating three-dimensional shock structures for transonic/supersonic compressor rotors, including experimental results obtained by holography, laser velocimetry, and highfrequency pressure transducers. Typical shock wave angles are oblique to the relative flow with angles in the range 60-65 deg range for maximum deflection, rather than the 40-50 deg range predicted by conventional cascade analyses. Results are partially explained by obliquity of the shocks in between-blade-streamsurfaces. Procedures for generating analytical flow patterns consistent with experiment, including supersonic/subsonic transition through oblique shocks, are demonstrated.

Effect of tip vortex structure on helicopter noise due to blade-vortex interaction

Abstract: A potential cause of helicopter impulsive noise, commonly called blade slap, is the unsteady lift fluctuation on a rotor blade due to interaction with the vortex trailed from another blade. The relationship between vortex structure and the intensity of the acoustic signal is investigated. The analysis is based on a theoretical model for blade/vortex interaction. Unsteady lift on the blades due to blade/vortex interaction is calculated using linear unsteady aerodynamic theory, and expressions are derived for the directivity, frequency spectrum, and transient signal of the radiated noise. An inviscid rollup model is used to calculate the velocity profile in the trailing vortex from the spanwise distribution of blade tip loading. A few cases of tip loading are investigated, and numerical results are presented for the unsteady lift and acoustic signal due to blade/vortex interaction. The intensity of the acoustic signal is shown to be quite sensitive to changes in tip vortex structure.

Use of Multiblade Coordinates for Helicopter Flap-Lag Stability with Dynamic Inflow

Abstract: Rotor flap-lag stability in forward flight is studied with and without dynamic inflow feedback via a multiblade coordinate transformation (MCT). The algebra of MCT is found to be so involved that it requires checking the final equations by independent means. Accordingly, an assessment of three derivation methods is given. Numerical results are presented for three- and four-bladed rotors up to an advance ratio of 0.5. While the constant-coefficient approximation under trimmed conditions is satisfactory for low-frequency modes, it is not satisfactory for high-frequency modes or for untrimmed conditions. The advantages of multiblade coordinates are pronounced when the blades are coupled by dynamic inflow.

Finite Element Subvolume Technique for Structural-Borne Interior Noise Prediction

Abstract: Finite element structural and acoustic representations of a vibrating structure and enclosed acoustic volume are used in a study of structural-bo rne interior noise. The direct finite element nodal representations of the equations of motion result in a large system of unsymmetric equations. In this paper, an acoustic subvolume analysis technique is presented which reduces the degrees of freedom of the interior volume to modal form prior to the coupled system dynamic analysis. Analytical predictions are compared to results from an experimental program to verify the analysis procedures. From these comparisons, the acoustic subvolume technique is shown to be a reliable method to reduce the computational requirements for finite element acoustic analysis.

Aeroelastic Tailoring Studies in Fighter Aircraft Design

Abstract: Studies have been conducted on the use of the directional properties of composite material to provide design improvements for fighter aircraft. The TSO (Aeroelastic Tailoring and Structural Optimization) computer program, developed by the Air Force Flight Dynamics Laboratory, was used in these investigations. The configurations evaluated covered a wide spectrum of fighter aircraft aerodynamic surfaces, including the F-15 composite wing, a preliminary design horizontal tail, a prototype aircraft wing, and a future conceptual aircraft wing. Both drag reduction and increased roll effectiveness, with no weight cost, are predicted for the F-15 composite wing. A unique minimum weight design is shown for the preliminary design horizontal tail, in which the anisotropic characteristics of the composite material are used to perform the dual function of strength and flutter balance weight. A more conventional optimum flutter solution, based upon outer panel wing root pitch restraint increases, is shown for the prototype aircraft wing. Finally, significant wing twist, offering potential aerodynamic benefits, can be obtained on the conceptual aircraft wing.

Statistical Estimation of Economic Life for Aircraft Structures

Abstract: An analytical methodology for the statistical prediciton of the economic life of advanced aircraft structures is presented. The approach allows for the determination of the economic life based on either one of the following criteria: 1) a rapid increase of the number of crack damages exceeding the economic repair crack size, and 2) a rapid increase of the maintenance cost, including the costs of inspection and repair. The formulation is general enough for practical applications. While the inspection and repair maintenance procedure has a significant impact on aircraft structural reliability and safety, its effect on the economic life is shown to be limited. Numerical examples are worked out to demonstrate the application of the methodology.

Large Wing-in-Ground Effect Transport Aircraft

Abstract: A promising innovative design concept is wing-in-groun d effect (WIG) where aircraft performance is increased significantly by drag reduction due to ground effect. This paper presents the results of preliminary design studies of large WIG transports utilizing a power-augmented ram system for lift enhancement developed by the David W. Taylor Naval Ship Research and Development Center. These studies include spanloader and fuselage loader designs and cover gross weights up to 1.9 million Ib and payloads up to 661, 500 Ib. Comparison of performance and economics are made among the WIG and several conventional design transport concepts.

Noise transmission into a light aircraft

Abstract: An analytical study on noise transmission into a cabin of a twin engine G/A aircraft is presented. The solution of the governing acoustic-structural equations of motion is developed utilizing modal expansions and a Galerkin type procedure. The exterior noise pressure inputs are taken from available experimental data. A direct comparison between theory and experiments on cabin noise levels is given. Interior noise reduction by stiffening, mass addition, and damping treatments is investigated. It is shown that a combination of added mass and damping could significantly reduce interior noise levels for this aircraft.

Effects of Spanwise Blowing on Two Fighter Airplane Configurations

Abstract: NASA Langley Research Center has recently conducted an investigation to determine the effects of spanwise blowing on two configurations representative of current fighter airplanes. This research examined not only the longitudinal or performance effects but was especially oriented toward determining the lateral-directional effects, particularly in the stall/departure angle-of-attack range. The wind-tunnel tests included measurement of static and forced-oscillation aerodynamic data, visualization of the airflow changes over the wing created by the spanwise blowing, and free-flight model tests. Effects of blowing rate, chordwise location of the blowing ports, asymmetric blowing, and the effects of blowing on the effectiveness of conventional aerodynamic controls were investigated.

Measurements of cabin and ambient ozone on B747 airplanes

Abstract: In response to recent concerns over possibly high ozone levels in the cabins of aircraft flying in the stratosphere, simultaneous measurements of the cabin and ambient ozone levels have been made as part of the NASA Global Atmospheric Sampling Program. Examples of the data taken on commercially operated Boeing 747-100 and 747SP airplanes are given for selected flights, together with summary statistics of over 5600 observations. Cabin ozone levels vary with the ambient level and, for unmodified aircraft, are higher on the 747SP than on the 747-100. Modifications to the ventilation system of the 747SP reduced cabin ozone levels by varying amounts up to a factor of 14.

Theoretical investigation of the aerodynamics of double membrane sailwing airfoil sections

Abstract: The type of sailwing here investigated has a rigid leading-edge spar with a rib at each end, a wire trailing-edge, and membranes which are wrapped around this structure to form upper and lower wing-surfaces. There appears to have been no previous theoretical investigation of a sailwing of this type, although it is expected to have aerodynamic and structural advantages, including controllability by using internal pressure, and to have various applications (e.g. wind-driven generators). Previous theoretical investigations of sailwings have concerned wings with a zero-thickness single membrane. The paper presents a two-dimensional numerical analysis method for obtaining the aerodynamic characteristics of such double-membrane sailwings. Numerical examples are given which show that: i) the pressure distribution on a sailwing with a rounded leading-edge is completely different from that on a single-membrane sailwing. ii) A sailwing with a circular or oval leading-edge has sharp suction peaks in the pressure distribution that may impair performance. iii) A D-spar leading-edge sailwing shows a large improvement in respect of pressure peaks; this is in qualitative agreement with the performance improvement noted in earlier experiments elsewhere with this form of leading edge. See next Abstract.

DYLOFLEX: A Computer Program for Flexible Aircraft Flight Dynamic Loads Analyses with Active Controls

Abstract: This paper describes and illustrates the capabilities of the DYLOFLEX computer program system. DYLOFLEX is an integrated system of computer programs for calculating dynamic loads of flexible airplanes with active control systems. A brief discussion of the engineering formulation of each of the nine DYLOFLEX programs is described. The capabilities of the system are illustrated by the analyses of two example configurations.

Calculation of Transonic Inlet Flowfields Using Generalized Coordinates

Abstract: A method for calculating in viscid supercritical flowfields about axisym metric inlet cowls with centerbodies is presented. A finite-differe nce approximation to the full-potential equation is solved under a general coordinate transformation, using a numerical evaluation of the transformation matrix at each mesh point. For the present problem, a boundary conforming coordinate system was generated by a sequence of conformal and shearing transformations, but this transformation is not essential to the method. Both the quasiconservative and non- conservative forms of Jameson's rotated differencing scheme are used, and the difference equations are solved by relaxation. Numerical results for pressure distributions generally agree well with experiment. INCE the initial success of Murman and ColeJ in applying the type-dependent differencing concept to the transonic small-disturbance equation, this technique has been ex- tensively applied to compute transonic flowfields about airfoils, using both the small-disturbance and full-potential equations. For flowfields around blunt bodies, the full- potential equation is required to properly resolve the solution in the vicinity of the stagnation points. The introduction of the rotated differencing scheme by Jameson 2 enables the use of line relaxation for solving the full-potential equation. There have been many successful applications of the rotated difference scheme to compute inviscid transonic flowfields around blunt-nosed bodies, including nacelle inlets and wing- body configurations. For a flowfield around a nacelle inlet, Arlinger3 applied a conformal mapping technique to trans- form the full-potential equation to a boundary conforming coordinate system. Caughey and Jameson4 solved the same problem using a sequence of simpler transformations, and studied ways to accelerate the iteration scheme used to solve the difference equations. Reyhner5'6 solved the full-potential equation in a Cartesian mesh. This requires an interpolation scheme to accurately treat the surface boundary condition. The mapping techniques generally require transforming the governing equations under the sequence of mappings. For complex geometries, these transformations can become a lengthy and tedious process. The boundary interpolation scheme has the advantage of always differencing the equation in Cartesian coordinates. There are, however, various types of irregular boundary elements, and the effort required to keep track of surface-mesh intersections is a major drawback in the application of this scheme. Caughey and Jameson7 proposed a finite-difference scheme to compute the supercritical flowfield about wing-body configuration using a form of the

Combined pressure and temperature distortion effects on internal flow of a turbofan engine

Abstract: An additional data base for improving and verifying a computer simulation developed by an engine manufacturer was obtained. The multisegment parallel compressor simulation was designed to predict the effects of steady-state circumferential inlet total-pressure and total-temperature distortions on the flows into and through a turbofan compression system. It also predicts the degree of distortion that will result in surge of the compressor. The effect of combined 180 deg square-wave distortion patterns of total pressure and total temperature in various relative positions is reported. The observed effects of the combined distortion on a unitary bypass ratio turbofan engine are presented in terms of total and static pressure profiles and total temperature profiles at stations ahead of the inlet guide vanes as well as through the fan-compressor system. These observed profiles are compared with those predicted by the complex multisegment model. The effects of relative position of the two components comprising the combined distortion on the degree resulting in surge are discussed. Certain relative positions required less combined distortion than either a temperature or pressure distortion by itself.

Studies of Leading-Edge Thrust Phenomena

Abstract: A study of practical limitations on achievement of theoretical leading-edge thrust has been made and an empirical method for estimation of attainable thrust has been developed. The method is based on a theoretical analysis of a set of two-dimensional airfoils to define thrust dependence on airfoil geometric characteristics and arbitrarily defined limiting pressures, an examination of two-dimensional airfoil experimental data to provide an estimate of limiting pressure dependence on local Mach number and Reynolds number, and employment of simple sweep theory to adapt the method to three-dimensional wings. Because the method takes into account the spanwise variation of airfoil section characteristics, an opportunity is afforded for design by iteration to maximize the attainable thrust and the attendant performance benefits. The applicability of the method was demonstrated by comparisons of theoretical and experimental aerodynamic characteristics for a series of wing-body configurations. Generally, good predictions of the attainable thrust and its influence on lift and drag characteristics were obtained over a range of Mach numbers from 0.24 to 2.0.