Showing papers on "Pitching moment published in 1987"

Enhancements to Program LAURA for computation of three-dimensional hypersonic flow

Abstract: Changes to Program Laura (Langley Aerothermodynamic Upwind Relaxation Algorithm) are presented which enhance both stability and accuracy of the algorithm. A discussion of iteration/sweeping strategies and their relation to computer architectures is included to best exploit the capabilities of serial, vector, and parallel processor machines. Test cases for Mach 10 perfect gas flow and Mach 32 real gas flow in chemical nonequilibrium over a blunt, raked elliptic cone using the thin-layer Navier-Stokes equations are presented in order to demonstrate the current improved capabilities. Algorithm changes include the use of volume averaging, application of a symmetric total variation diminishing (TVD) scheme, and stronger interaction between the grid/shock alignment routine and the relaxation algorithm. Good comparisons with heat transfer and pitching moment data at three different angles of attack for the Mach 10 tests serve to further validate the present algorithm. Parameters are defined which control the coupling of the specie continuity equations with the solution of the mixture conservation equations. A discussion of the consequences involved in the choice of strong versus weak coupling is presented, and a sample nonequilibrium calculation on a fine grid over a full scale model of the Aeroassist Flight Experiment (AFE) demonstrates current capabilities.

Unsteady Stall Penetration Experiments at High Reynolds Number

Abstract: : An experiment was performed to examine the unsteady aerodynamics of stall penetration at constant pitch rate and high Reynolds number, in an attempt to more accurately model conditions during aircraft post-stall maneuvers and during helicopter high speed forward flight The model spanned the 8 ft wind tunnel and consisted of a 173 in chord wing with a Sikorsky SSC-AOQ airfoil section Two forms of pitching motion were used: constant pitch rate ramps and sinusoidal oscillations Ramp data were obtained for 36 test points at pitch rates between 0001 and 0020, Mach numbers between 02 and 04, and Reynolds numbers between 2 and 4 million Sinusoidal data were obtained for an additional 9 conditions The results demonstrate the influence of the leading edge stall vortex on the unsteady aerodynamic response during and after stall The vortex- related unsteady increments to the lift, drag, and pitching moment increase with pitch rate; the maximum delta C sub L is 12 at A =002 Angular delays in stall events also increase with pitch rate Vortex strength and propagation velocity were determined from pressures induced on the airfoil surface The vortex is strengthened by increasing the pitch rate, and is weakened both by increasing the Mach number and by starting the motion close to the steady-state stall angle Propagation velocity increases linearly with pitch rate

Forebody vortex management for yaw control at high angles of attack

Abstract: The yaw-control potential of deployable forebody strakes at angles of attack above the range of conventional rudder effectiveness has been investigated. The conformally-stored strakes when deployed force asymmetrical vortex shedding from the forebody, thereby generating a controlled yawing moment. The concept was explored through low-speed wind tunnel tests on a conical forebody in isolation and in a generic fighter configuration. Force and moment measurements on the complete model were supplemented with circumferential pressure and flow-visualization surveys on an isolated forebody, in order to gain insight into the vortex flow mechanisms resulting from forced asymmetrical separations and to quantify the obtainable yaw power at angles of attack to 80 deg. This preliminary, low-Reynolds-number study showed asymmetrically-deployed forebody strakes to have considerable yaw control potential, whose sensitivity to scale effects needs further investigation.

Optimum Release Conditions for the New Rules Javelin

Abstract: Changes in the rules for construction of the men's javelin have dramatically altered the pitching moment profile as a function of angle of attack. Thus the optimal release conditions are different for the new javelin. Optimal release conditions are presented for nominal release velocities in the range 20 < vn < 35 m/s. Although the optimal release angle remains roughly constant near 30° over this speed range, the optimal angle of attack and pitching angular velocity change substantially with speed. The main effects of the rule change have been (a) to decrease the achievable range at a nominal velocity vn = 30 m/s by about 10% by making it impossible to take advantage of the javelin's potentially large aerodynamic lift forces, and (b) to make the flight much less sensitive to initial conditions.

A study of high-lift airfoils at high Reynolds numbers in the Langley low-turbulence pressure tunnel

Abstract: An experimental study was conducted in the Langley Low Turbulence Pressure Tunnel to determine the effects of Reynolds number and Mach number on the two-dimensional aerodynamic performance of two supercritical type airfoils, one equipped with a conventional flap system and the other with an advanced high lift flap system. The conventional flap system consisted of a leading edge slat and a double slotted, trailing edge flap with a small chord vane and a large chord aft flap. The advanced flap system consisted of a leading edge slat and a double slotted, trailing edge flap with a large chord vane and a small chord aft flap. Both models were tested with all elements nested to form the cruise airfoil and with the leading edge slat and with a single or double slotted, trailing edge flap deflected to form the high lift airfoils. The experimental tests were conducted through a Reynolds number range from 2.8 to 20.9 x 1,000,000 and a Mach number range from 0.10 to 0.35. Lift and pitching moment data were obtained. Summaries of the test results obtained are presented and comparisons are made between the observed aerodynamic performance trends for both models. The results showing the effect of leading edge frost and glaze ice formation is given.

Multiaxis control power from thrust vectoring for a supersonic fighter aircraft model at Mach 0.20 to 2.47

Abstract: The aeropropulsive characteristics of an advanced twin-engine fighter aircraft designed for supersonic cruise have been studied in the Langley 16-Foot Tansonic Tunnel and the Lewis 10- by 10-Foot Supersonic Tunnel The objective was to determine multiaxis control-power characteristics from thrust vectoring A two-dimensional convergent-divergent nozzle was designed to provide yaw vector angles of 0, -10, and -20 deg combined with geometric pitch vector angles of 0 and 15 deg Yaw thrust vectoring was provided by yaw flaps located in the nozzle sidewalls Roll control was obtained from differential pitch vectoring This investigation was conducted at Mach numbers from 020 to 247 Angle of attack was varied from 0 to about 19 deg, and nozzle pressure ratio was varied from about 1 (jet off) to 28, depending on Mach number Increments in force or moment coefficient that result from pitch or yaw thrust vectoring remain essentially constant over the entire angle-of-attack range of all Mach numbers tested There was no effect of pitch vectoring on the lateral aerodynamic forces and moments and only very small effects of yaw vectoring on the longitudinal aerodynamic forces and moments This result indicates little cross-coupling of control forces and moments for combined pitch-yaw vectoring

Transition effects on airfoil dynamics and the implications for subscale tests

Abstract: p Re 5 t At U x z a a. ot0 A e 6 A v £ p \l/ co, co Nomenclature disturbance amplitude disturbance amplitude in freestream initial disturbance amplitude in boundary layer reference initial disturbance amplitude in boundary layer at Mr(= 1.3) airfoil chord mean aerodynamic chord drag, coefficient CD = D/(p00Ux>/2)S frequency proportionality constant, Eq. (1) lift, coefficient CL = L /(p^ Ul/2)S sectional lift, coefficient cl = l/(p00(/i/2)c Mach number mass flow injected calculated mass flow in boundary layer just upstream of air holes sectional pitching moment, coefficient cm-mp/

High velocity aerodynamic body having telescopic pivotal tip

Abstract: A high velocity aerodynamic body, particularly a shell flying at supersonic velocity, having a device for stabilizing the aerodynamic body and for reducing its oscillation. The aerodynamic body has in the vicinity of its tip, a rotation symmetrical tip enclosure which is supported, with balanced mass, about a support point located on the longitudinal axis of the aerodynamic body, freely tiltably on all sides.

Wind tunnel dynamic flying study of the pitching moment derivatives of the Standard Dynamics Model in active control

Abstract: This paper presents the results of a dynamic wind tunnel flying study of the pitching moment coefficients of the Standard Dynamics Model as a function of angle-of-attack, ranging from about -10 deg to +30 deg. The fully instrumented model, with its servoed elevator, has been flown in a single-pitch degree-of-freedom at low speeds under moment trim conditions covering both the stable CG locations of 19.2 percent, 25 percent, and the unstable CG location of 35 percent of mean aerodynamic chord. In the unstable case an active control law has been invoked to fly the model in an angle of attack control mode. The freely flying model has been disturbed with flight testlike inputs, and the responses have been analyzed to generate the aerodynamic pitching moment coefficients using a maximum likelihood estimation procedure. The analysis includes the closed loop active controlled system parameter estimation and the results are compared with the available SDM pitching moment data.

Euler analysis of the AFWAL 65-deg delta wing

Abstract: The "International Vortex Flow Experiment on Euler Code Validation" program experimental data has been compared to Euler code predictions for a range of Mach numbers and angles of attack. The geometry used in this study is a generic 65 degree swept leading edge delta wing with both round and sharp l e a d i n g e d g e s . C o m p a r i s o n s o f experimental and predicted coefficients of lift, drag, pitching moment and pressure are presented at Mach numbers of 0.4, 0.85 and 1 . 2 for angles of attack from 0 to 25 degrees. The comparison of force and moment data is in good agreement for all cases, except for subsonic pitching moment coefficient. Pressure coefficients are in excellent agreement in attached flow regions and in poor agreement in separated flow regions. The effect of vortex breakdown o n lift and drag has been accurately predicted using the Euler equations at a free stream Mach number of 0.85. An examination of a predicted complex shock structure interacting with the vortex breakdown s t r u c t u r e i s presented.

Steady and unsteady aerodynamic forces from the SOUSSA surface-panel method for a fighter wing with tip missile and comparison with experiment and PANAIR

Abstract: The body surface-panel method SOUSSA is applied to calculate steady and unsteady lift and pitching moment coefficients on a thin fighter-type wing model with and without a tip-mounted missile. Comparisons are presented with experimental results and with PANAIR and PANAIR-related calculations for Mach numbers from 0.6 to 0.9. In general the SOUSSA program, the experiments, and the PANAIR (and related) programs give lift and pitching-moment results which agree at least fairly well, except for the unsteady clean-wing experimental moment and the unsteady moment on the wing tip body calculated by a PANAIR-predecessor program at a Mach number of 0.8.

Transonic wing design procedure

Abstract: The conventional steps of defining performance requirements of an aircraft wing and conducting general sizing of the wing are followed by the new steps of determining the aerodynamic sweep angle of the wing with respect to an actual location of a shock wave on the wing. A two dimensional Mach number for the wing airfoil is calculated by multiplying the three dimensional Mach number by the cosine of the aerodynamic sweep angle. A two dimensional lift coefficient for the wing airfoil is calculated by dividing the three dimensional lift coefficient by the square of the cosine of the aerodynamic sweep angle. Airfoil shape in two dimensions is determined on the basis of the two dimensions is determined on the basis of the two dimensional Mach number and the two dimensional lift coefficient. The shape of the wing in three dimensions is then defined by placing the airfoil in the wing along an arc constructed by skewed chord lines perpendicular to local sweep lines of the wing at a series of locations along a chord of the wing.

Jet spoiler as a yaw control device

Abstract: = span of the semispan model = wing chord = jet slot length = rolling moment coefficient = lift coefficient = yawing moment coefficient = pressure coefficient = jet momentum coefficient = vertical projection of equivalent solid spoiler, percent chord = proportionality constant = freestream velocity = jet exit velocity averaged over slot area = wind-off jet velocity along slot center line = chordwise coordinate from leading edge = spanwise coordinate from wing center = coordinate normal to wing upper surface = change of local drag coefficient = change of local lift coefficient = angle of attack = jet slot width = nondimensional spanwise coordinate (y/b) from wing root = nondimensional chordwise coordinate (x/c) from wing leading edge

Two-dimensional NACA 66(MOD) hydrofoil High Speed Water Tunnel tests

Abstract: Two-dimensional tests were conducted on a NACA 66(MOD) hydrofoil in the the GALCIT Hydrolab High Speed Water Tunnel (HSWT) . These tests were conducted using the hydrofoil with a. a rough leading edge, and b. a smooth leading edge, covering the following range of conditions: 1. Speed range of 30 ft/s to 60 ft/s 2. Angles of attack of 0° to 6° and 3. static pressures of 3. 03 psiA to 33. 54 psiA, corresponding to cavitating, incipient cavitation thru fully wetted flow conditions. These tests were performed in the two-dimensional test section of the HSWT and included measurements of: --Tunnel velocity. --Tunnel static pressure. --Lift, Drag and Pitching Moment forces (with tare forces removed). --Pressure coefficients on 13 taps, 12 at selected locations on the lifting surface, plus 1 location on the bottom surface. --High speed (strobe) flow visualization photography under flow cavitation conditions. --Airfoil gap dependence on static pressure.

Subsonic maneuver capability of a supersonic cruise fighter wing concept

Abstract: A theoretical and experimental investigation was conducted of the subsonic maneuver capability of a fighter wing concept designed for supersonic cruise. To improve the subsonic maneuver capability, the wing utilized full-span leading- and trailing-edge flaps that were designed with the aid of a subsonic-analysis computer program. Wind-tunnel tests were made at Mach numbers of 0.3, 0.5, and 0.7. Force and moment data obtained were compared with theoretical predictions of Mach 0.5 from two subsonic-analysis computer programs. The two theoretical programs gave a good prediction of the lift and drag characteristics but only a fair prediction of the pitching moment. The experimental results of this study show that with the proper combination of leading- and trailing-edge flap deflections, a suction parameter of nearly 90 percent can be attained at a Mach number of 0.5 and a lift coefficient of 0.73; this is a three-fold improvement from 30 percent for the basic wing.

VSAERO analysis of structurally decoupled tip planforms for a semispan wing

Abstract: A numerical analysis has been conducted with the three-dimensional panel code VSAERO for two interacting lifting surfaces that are separated in the spanwise direction by a narrow gap, with the angle of attack of the outboard section being set independently of the inboard section, as in the 'free tip' rotor blade system proposed for helicopters. Computed values of tip surface lift and pitching moment coefficients are correlated with experimental data to determine the most suitable method for modeling the gap region between the surfaces. It is shown that the induced drag of the tip surface is reduced for negative incidence angles relative to the inboard section.

An Assessment of the Use of Low-Order Panel Methods for the Calculation of Supersonic Flows,

Abstract: : Three low order panel methods developed for the analysis of supersonic flows have been used to make calculations for a number of test configurations, with the aim of establishing the range of problems for which this class of computational method yields adequate solutions. The programs investigated were the Woodward USSAERO/C program, the related NLRAERO program, and the British Aerospace Warton supersonic panel program. Results obtained using these programs have been evaluated against theoretical and experimental data for a number of test cases covering wing alone, body alone and wing body geometries. It is concluded that low-order panel methods can provide adequate solutions for supersonic flows about wings, bodies and wing body combinations, provided the assumptions implicit in the linearised potential flow model are not violated. Examples show that the prediction of lift and pitching moment curve slopes for quite complex configurations may be acceptable, but that the detailed pressure distributions are not always predicted satisfactorily. In particular, serious problems are encountered in calculating the flow about wings with rounded supersonic leading edges due to the linearised flow model which is used.

Investigation of Periodic Pitching through the Static Stall Angle of Attack.

Abstract: : The flow over an NACA 0015 airfoil undergoing periodic pitching motions using constant pitch rate ramps was experimentally stuided over a range of pitch rates and angles of attack. Surface pressure transducers coupled with a microcomputer-based data acquistion system were used to collect surface pressure data a rate of 4000 samples per second. The data was reduced through numerical integration of the pressure data to provide graphs of the coefficients of lift, pressure drag and pitching moment versus time. Each point on the graphs represents an average of five runs. The results were compared according to their nondimensional pitch rates (defined as the product of one-half the chord length and the pitch rate divided by the freestream velocity). Data was collected in the range of nondimensional pitch up rates between 0.104 and 0.384 and the range of angles of attack between 0 and 30 degrees. Keywords: Dynamic stall. (Theses).

Contributions to the prediction of low Reynolds number aerofoil performance

Abstract: The dominant aspects of low Reynolds number flows are identified and their relevance to aerofoil performance discussed A method for assessing two-dimensional aerofoil performance characteristics, including trailing edge and gross laminar separation, is developed along with a subsidiary direct boundary layer calculation scheme capable of accounting for short laminar separation bubbles The constituent parts of the performance prediction scheme, which consists a vortex panel method with boundary layer corrections and an inviscidly modelled wake, are described in some detail Predictions obtained for both laminar and turbulent separation are also presented For laminar separation, an inviscid Wake Factor Increment correlation is developed to account for the effects of the free laminar shear layers Generally, the predictions of lift and pitching moment may be considered to be within the experimental error, but where this is not the case the applicability of the modelling technique is discussed The developed direct boundary layer calculation technique is demonstrated to provide an indication of the boundary layer growth through a separation bubble for a prescribed pressure distribution whilst encountering no difficulty at separation Comparisons with inverse calculations are made and exhibit good general agreement Finally, the general applicability of the predictive scheme is discussed along with possible future enhancements.

Low Reynolds Number Performance of a NACA 0015 and a GA(W)-1 aerofoil

Abstract: Due to the recent interest in a wide variety of low Reynolds number applications such as mini RPV's, wind tubrine fans, sailplanes etc., attention has been focused on the evaluation of efficient airfoil sections at chord Reynolds numbers from about 50,000 to about 1,000,00. In this experimental study, the low Reynolds number aerodynamic characteristics of two airfoil sections, namely the NASA GA(W)-1 and NACA-0015 were examined. These were compared to the ones obtained from previous investigations for the GU25-5(11)8 airfoil section. The airfoils were tested in the Reynolds number range from about 50,000 to about 500,000 and for incidences of 0 to 22. An automated pressure measuring system was developed to improve the speed and facilitate the measurements of pressures around the airfoil sections. The pressure measurements were converted to pressure coefficients and these were in turn integrated to provide normal force and pitching moment coefficients. Oil flow visualisation was used to obtain a better picture of the different flow phenomena around the airfoil sections. It proved to be an essential tool for obtaining information about the different flow fields which occur around the airfoil models when these were not apparent by pressure distributions. Many of the significant aerodynamic problems which occur in this low Reynolds number regime such as the creation and behaviour of laminar separation bubbles and the extreme sensitivity of the boundary layer to the test environment (i.e. free stream turbulence level, mechanical vibrations and noise levels) were highlighted. Significant differences were found in the behaviour of the boundary layer and subsequently in the aerodynamic characteristics of the three airfoil sections. This resulted in marginal differences as far as the Reynolds number operational range is concerned.

Lift and pitching moment coefficient changes from low-level freestream turbulence

Abstract: Etude experimentale des variations du moment de tangage et de la portance en fonction de la turbulence. Modele de variation du coefficient de portance

A Numerical Study of Flare Stabilized Projectiles at Mach 7.0.

Abstract: The effect of nose shape on two flare stabilized projectiles was studied using a Parabolized Navier-Stokes code. Pressure coefficients, forces and moments, skin friction coefficients, and Stanton number calculations are presented for the hemisphere-cylinder-flare and the cone-cylinder-flare configurations. Pitching moment and static margin plots versus flare angles and flare lengths are presented in a parametric study to show aerodynamic stability effects. Distinctly different flow field values of pressure, local Mach number, and dynamic pressure were generated by the two different nose configurations. These flow field values just upstream of the flare will be examined. The code demonstrated its value as a design tool by making a clear distinction between aerodynamically stable characteristics for this variety of nose shapes, flare angles, and flare lengths.

Aerodynamic analysis of a complex configuration using a full potential code

Abstract: A detailed comparison between supersonic full potential calculations and experimental surf ace pressures and force and moment coefficients for the F111/TACT configuration at Mach numbers of 1.6 and 2.1 has been performed. The agreement between the predicted and measured surface pressure distributions, the lift curve slope and the lift-todrag ratio was good to excellent. The zero-lift pitching moment for both Mach numbers examined was underpredicted. The drag polar for Mach 1.6 was in good agreement with experiment. For Mach 2.1 the drag was underpredicted by about forty counts but the shape of the polar was in good agreement with experiment. The disagreement between the predicted and measured zero-lift pitching moments and drag levels was attributed to the use of actual geometry rather than the distorted configuration required for incorporation of the wind tunnel sting.

Summary of studies to reduce wing-mounted propfan installation drag on an M = 0.8 transport

Abstract: Powerplant installation losses for an advanced, high-speed, turboprop transport have been investigated in the Ames Research Center Transonic Wind Tunnels as a part of the NASA Advanced Turboprop Program (ATP) Force and pressure tests have been completed at Mach numbers from 06 to 082 on baseline and modified powered-model configurations to determine the magnitude of the losses and to what extent current design tools could be used to optimize the installed performance of turboprop propulsion systems designed to cruise at M = 08 Results of the tests indicate a large reduction in installed drag for the modified configuration The wing-mounted power plant caused destabilizing pitching moments and a negative shift in the zero-lift pitching moment

The effect of periodic ballonet jet exhaust on the stability of a tethered aerostat AIAA-87-2504

Abstract: Tethered aerostats operating under wide13; range of altitudes and atmospheric pressure conditions13; employ ballonets within the hull filled with13; air. These ballonets exhaust the air into the13; atmosphere whenever necessary to maintain the13; difference between the atmospheric pressure and13; the internal pressure within a desired range.13; Assuming that the air exhaust is made periodic,13; studies are made on the effect of the periodic13; jet force and the resulting pitching moment on13; the responce of the aerostat system. In the first13; phase, the stability of the system is studied by13; analyzing the eigenvalues and eigenvectors. The13; result shows that the aerostat system is unstable13; in the wind speed range of 0-15 fps and stable13; in the windepesd range of 15-100 fpamp; When13; the system is subbcted to a suitable periodic13; forcing from the exhaust jet, it is observed that13; It is not only possible to extend the stability13; boundary of 15 fps down to typically 5 fps, but13; much more interestingly, there exist forcing periods13; of the extraust jet which render the system completely13; stationary.13;

Predicted Pitching Moment Characteristics of X=29A Aircraft

Abstract: The predicted pitching moment characteristics of the X-29A aircraft are presented for angles of attack from 0 to 20 deg. and Mach numbers of 0.2, 0.6, 0.9, 1.2, and 1.5 for altitudes of sea level, 4572 m (15,000 ft), 9144 m (30,000 ft), and 12,192 m (40,000 ft). These data are for both rigid and flexible aircraft and for the full range of control-surface positions. The characteristics were extracted from a nonlinear, symmetric, flexibilized wind tunnel data base.