ENGINEERING RESEARCH INSTITUTE THE UNIVERSITY OF MICHIGAN ANN ARBOR Final Report THE GENERATION OF GUSTS IN A WIND TUNNEL AND MEASUREMENT OF UNSTEADY LIFT ON AN AIRFOIL WADC Technical Report No. 57-401o. L. C.Garby A..M. Kuethe J..D. Shetzer Project 2099 DEPARTMENT OF THE AIR FORCE WRIGHT AIR DEVELOPMENT CENTER WRIGHT-PATTERSON AIR FORCE BASE, OHIO CONTRACT NO. AF 33(616)316 PROJECT NO. 53-670A-686 AND 52-670A-86 June 1957

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TABLE OF CONTENTB Page LIST OF ILLUSTRATIONS iv ABSTRACT vi OBJECTIVE vi INTRODUCTION 1 UNSTEADY AERODYNAMIC EFFECTS 1 PREVIOUS WORK ON GUST SIMULATION AND RELATED PROBLEMS 2 EXPERIMENTAL TECHNIQUES 3 Repeatability 3 Gust Testing 3 Gust Structure 3 Dynamic Velocity 3 Quasi-Steady Effect 3 LIFT-LAG THEORY 5 QUASI-STEADY AND APPARENT MASS LIFTS 6 WAKE LIFT 7 TESTS RESULTS 7 Moving Bump Gust Generator 7 Vortex Gust Generator 9 CONCLUSIONS 9 REFERENCES 11 APPENDIX I. DESCRIPTION OF TEST TUNNELS 14 A. Open-Return Tunnel 14 B. Closed Return: 5-Foot x 7-Foot Tunnel 14 APPENDIX II. INSTRUMENTATION DETAILS 16 A. Bump-Position Indicator 16 Bo Angle-of-Attack Measurements 16 C. Speed Measurements 16 D, Balance System 16 WADC TR 57-401 iii

LIST OF ILLUSTRATIONS Figure Page la Schematic diagram of Vortex Gust Generator. 19 lb Sketch showing Vortex pattern formed by action of the slats. 20 2 Normal velocity distribution for three positions of the bump. 21 3 Open-return tunneil. 22 4a Plan view of 5-foot x 7-foot gust-generator model. 23 4b Perspective drawing of 5-foot x 7yfoot gust-generator model. 24 5a Schematic drawing of Vortex Gust Generator mechanism. 25 5b Photograph showing bench calibration of Vortex Gust Generator and the associated electrical equipment 26 5c View of linkage mechanism of Vortex Gust Generator. 27 6a Schematic diagram of Moving Bump Gust Generator. 28 6b Photographs of moving bump and its dimensions as used in 21 —inch x 29-inch open-return tunnel.o 29 6c Photograph of moving bump in 5-foot x 7-foot tunnel. 30 6d Sketch of moving bump in 5-foot x 7-foot tunnel. 30 7 Bump position in chord lengths versus timeo 31 8 Flow inclination in degrees versus bump position in wing chords. 32 9 Flow deflection for Vortex Gust Generator versus time for various stations along tunnel axiS, 33 10 Sketch showing mathematical model of lift-lag system, 34 11..antitative effect of bump motion upon flow-streamline distribution. 35 12 Measured dynamic lift, quasi-steady lift, and bump position versus time for a test-section velocity of 44.4 feet per second -36 t3. Coparison of eperimental ard theoretial lift lag versus time for test-s:ection velocity of 4404 feet per second ~ 37 14 Measured dyna;mi lift, quasi-steady lift, and bup position versus time for test-section velocity of 72.5 feet per second. 38 WADC TR 57-401 ivr

LIST OF ILLUSTRATIONS (Concluded) Page 15 Comparison of theoretical and experimental lift lag for Moving Bump Gust Generator. 39 16 Sketch showing flow patterns at three bump positions for liftovershoot exeriment. 40 7 Lfift response at two bump speeds in region of wing stall. 41 1 Effeet of rate of chane of angle of attack upon. maximum normal force. oefficient. 41 19 Oscillograms. of the lifting-surface and monitor-probe responses to the Vortex Gust Geerator. 42 20.esponses of lifting-surface and hot-wire monitor probe to.Vortex Gust Generator.for three wind speeds. 43 21:Sketch.of bunp-position mechanism. 44 2:2 Electrical circuit used to measure flow angles. 4' 23a Exploded view of balance as used in Moving Bump Gust Generator 46 23b View showing flexure-beam block and Schaevitz Transformer mounting. 47 WADC DTR 57-401 v

ABSTRACT This report covers the work done under Contract No. AF 33(616)316 between The University of Michigan and Wright Air Development Center of the United States Air Force. The work includes (1) the development of the-equipment first on a small scale and measurements in a small wind tunnel; (2) the design and fabrication of a low-turbulence wind tunnel with a 5-foot x 7-foot working section [according to specifications developed in the low-speed work under (1)]; (3) construction of equipment for tests in the large tunnel; (4) tests of an airfoil at low and high angles of attack during passage through a gust; and (5) analysis of the results and comparison with theory where possible OBJECTIVE The object of the work was to develop facilities for the measurement of gust loads and to demonstrate their feasibility at Reynolds numbers high enough to permit application of the results to full-scale flight. WADC TR 57-401 vi

INTRODUCTION With inreasing.size range, and complexity of aircraft, it is becoming important to determine more closely the effects of atmospheric gusts upon aircraft. Exterimental information has been obtained largely in two ways: by full-scale flight tests through turbulent airy and from experimental model techniques in which the tt i s.imulated. Atmospheric turbulence is made up of a large number of gusts statibstically distributed in intensity and scale, A particular cross; sction of the atmosphere through which an aircraft flies is repeatable, therefore, only in its statistical aspects. Flight-test-work can, in general, only give a gross response, and will not easily yield details as to the exact manner in which the gust acts upon the aircraft. To provide the designer with more information, and to point e way for theoretical work, an experimental approach is needed. UNSTEADY AERODYNAMIC EFFECTS If the flow field about an aircraft is changing rapidly, certain nonstationary effects come into play that make the lift and moment on the aircraft depart from th ir "quasi-steady values. The term "quasi-steady" is used to denote the timevarying force that, at any instant, is equal to the steady-state value correspondI:g to the flow configurti ation at that instant. The difference between the true dyiaic' lift and the quasi-steady value is called the lift lag. The details of the various contributions to the lift lag are not understood for wing-body-tail combinations in which interference effects play a prominent role. For wings alone at subsonic speed, the lift lag is produced by the wake that accompanies the time-varying circulation and by the inertial reaction of the air to local acceleration. For simple plan forms, the so-called "wake lift" and "apparent-mass" lift can be predicted from the theory of thin airfoils. For a wing moving through al severe gust, the lift lag is known to be important. For a complete airplane with all interference effects considered, it is expected -that the lift lag will also be a prominent factor. The lift lag is an essential feature of the gust problem and, therefore, gust-simulation equipnlent must include instrumentation that is adequate for measurement of the lift lag. Tests on the pilot model of gust-simulation equipment which is described in the following sections have been run on a.constrained two-dimensional wing. The theory for this ease is well understood and has served as a guide for developing the instrumenation necessary to measure the lift. Furthertests in a 5-foot x 7-oot wind tunnel have verified the theory for two0odimensional flow and also indicate unsteady effects near maximum lifta WADC TRE 57401 1

PREVIOUS WORK ON GUST SIMULATION AND RELATED PROBLEMS Typical references to literature on problems related to the dynamics of air-:craft flying,through gusts are cited in thiS section. No attempt is mad.e to give a -ofmplete bibliography on the sub Secto..: V -:It is believed that the only facilities presently in use for simulating the effet -of atmospheric gusts are of the NACA Langley Field type (References 1 and 2), -.- which. amodel is flown through an updraft of controlled profile. and the resulting:. motion is observed photographically Experiments in unsteady aerodynamics have been confined very largely to the measurement of forces and moments on models that are oscillated in a wind tunnel.:Th. frequency information so obtained is used for arbitrary periodic motions by applying the superposition principle. Experiments of this type have been reported in the literature as early as 20 years ago, and they are currently being performed in ever-increasing numbers. Representative experiments on wings, bodies, and wingbody combinations are described in References 3 through 6. Unsteady forces and moments have been calculated from the theory of thin air foils for two- and three-dimensional wings moving through an incompressible fluid. and.also for wings moving.subsonically and supersonically through a.compressible fluid-. Numerical data appear in the literature in the form of the indicial admittance (response to a unit step input) and the aerodynamic transfer function (repone to a sinusoidal input). The response to an arbitrary input is found by Iuhamel integration, using the indicial admittance, or Fourier synthesis, using the. transfer function. Representative theory and calculations on nonstationary aerodynamids appear in References 7 through 16.:,:,;:Among the earlier experiments on unsteady flow are those described in Refer-:nce 17, in which an airfoil was set in motion in a water tank and the flow pattern made visible by..hi ng light on suspended oil globules. The circulation build-up duing the first few chords of airfoil travel was calculated from the motion of the globules. Direct force measurements were obtained. on a wing near the stall and rotated at a rate of one degree in 2.5 chords of wing travel (Reference 18). The circulation build-up on a.wing was determined by taking hot-wire measurements of the flow pattern around the tip-trailing vortices as the wing passed through an updraft of known profile (Reference 19).;: Calculations of the response of aircraft to gusts of given profile have appeared abundantly in the literature. A complete investigation requires that the rigid and elastic degrees of freedom of the aircraft be considered in investigating the response as the craft penetrates gusts of various sharpness. Though there is no difficulty in principle in assuming many degrees of freedom, the computational labor is great, and most investigators have consideredonly a small number. Typical.examples, of these calculations are given in Reference 20,, in which the authors have:alculated the response of a rigid airplane to gusts of various degrees of.ha.snes, considering the plunging degree of freedom only. Calculations have been.made using both steady- and unsteady-flow theory, In eference 21, the response of an airplane entering a sinusoidal gust, with the WADC. TR 57-401 2

lplunging and wing-bending degrees of freedom considered, has been calculated. Calculations on the influence of unsteady flow and. structural flexibility on the rigidbody oscillations of an aircraft are numerous) typical examples are given in References 22 through 25. A general treatment of dynamics calculations is given in Reference 26. Finally, the nature of atmospheric turbulence, as deduced from meteorological measurements or the response of an airplane, is reported in the literature, i-nd-ypical- examples are given in References 27 through 31. EXPERIMENTAL TECHNIQUES To devise an experimental technique to check theoretical predictions and to -provid:dedsign data, the following qualifications or criteria were established. REPEATABILITY The gust pattern and response must be repeatable under similar circumstances:and these conditions must be capable of controlled variation. GU:T TESTING:g-'' " The mechanism of the gust should be such that the model is either entering or:merging from the gust. It is in this transitional period that the lift lag.occurs, GUST STRUCTURE The gust structure should preferably be a monotonic and rapid increase or decrease in angle of attack. A monotonic variation will in general be more reproducible, and the calculations necessary for a comparison with theory will be less involved than for more complicated gust structures. The more rapid the variation of:agle of attack, the greater will be the lift lag, and hence greater accuracy will be obtainable in its measurement. DYNAMIC VELOCITY The mean velocity in the wind-stream direction should remain as constant as Po.5sible, making the gust effect one of an angle-of-attack change that can be reproduied in a quasi-steady manner.: QUASI-TEADY EFFECT The mechanism producing the gust should be such that the gust pattern can be locked across the model~ Quasi-steady measurements can then be made and the lift lag determined experimentally. WADC TR 57-401 3

It was felt that along with the above conditions, and if at all possible, the model should be fixed in the tunnel and the gust passed over it, rather than the reverse. The instrumentation for the measurement of forces and for control of angle of attack is much simpler for a fixed than for a moving model. The passage of a gust across a fixed model may be achieved in two ways: (1) a disturbance placed in the flow upstream of the model will be carried past the model by the tunnel flow; and (2) a continuous variation of the boundary conditions at the tunnel wall will cause an angle-of-attack change to sweep past the model. Descriptions of the devices used for generating gusts are given below, followed by a detailed description of the experimental equipment and measurement technique. For the first device, that of generating a disturbance and allowing the flow to carry the disturbance, the following system was used. A "Venetian blind" arrangement of airfoils (slats), referred to hereafter as a Vortex Gust Generator, was placed in the tunnel upstream of the model. A spring-loaded mechanism permitted a rapid change of angle of attack of the slats up to 10~ and return to O~ in about 1/25 second. The vortex sheets shed by the slats during their motion were carried downstream and simulated a severe gust when they passed over the model. A schematic diagram is shown in Figures la and lb. The other technique was that of changing the wall boundary conditions. This is done by moving a bump along the floor of the tunnel. The flow field which passes over the model, shown schematically in Figure 2, simulates the emergence of the model from a gust, whose severity depends upon the rate of movement of the bump and the airspeed. To carry out this program, two wind tunnels were used. Initial development of the Vortex and Moving Bump Gust Generator models was carried out in a small openreturn tunnel having a nominal test-section dimension of 21 inches x 29 inches. A sketch of the tunnel with pertinent dimensions is given in Figure 3. Details of the configuration are given in Appendix I. A second tunnel was used to extend the results of the Moving Bump Gust Generator model by increasing the Reynolds number. A sketch of this tunnel is shown in Figures 4a and 4b, and details of its configuration are given in Appendix I. Shown in Figure 5 are mechanical details of the Vortex Gust Generator. Figure 6 shows details of the Moving Bump Gust Generator for its installation in the first tunnel and its installation in the larger tunnel. To measure the gust effect, the following measurements are made: position of bump in test section; angle of flow; and wind speed and normal force on wing model. For the moving-bump tests, the position of the bump is measured by coupling the bump to a potentiometer. The bump position can be determined within an accuracy of + 005 ic D il f th st n iri r in inc Appendix II, part A. In Figure 7 are shown bump histories for a series of tests. It may be seen that the bump reaches its speed within approximately 1/2 chord and remains constant thereafter, The angle of flow was measured for both the Vortex and the Moving Bump Gust WIADC TR 57-401 4

Generators by means of a hot-wire anemometer with a % probe. An accuracy of ~ 0.1 degree was limited by the accuracy of probe setting. A sensitivity of 0.02 degree was obtained by the electrical equipment. Methods of mounting and the circuitry are given in Appendix II, part B. Typical angle-of-attack plots for the Vortex and Moving Bump Gust Generators are given in Figures 8 and 9. Dynamic data were initially recorded on a dual beam DuMont oscilloscope with a Polaroid-Land camera that photographed the face of the scope. In the later tests, particularly those in the 5-foot x 7-foot tunnel, a Consolidated recording oscillograph was used. The normal forces were measured by a single-component balance. The balance converted the normal force into either a rotational or translational displacement that was amplified and measured by means of a Schaevitz Linear Variable Differential Transformer. Details of the balance are given in Appendix II, part C. Prior to and immediately after a series of dynamic runs, a quasi-steady run and calibration run were made to be certain that the data were consistent. LIFT-LAG THEORY A gust that sweeps across a wing causes a lift history that is different from the response that would occur if the gust pattern could be treated as a series of steady-state distributions across the wing. Part of this difference is caused by the fact that, as the lift changes, vorticity is shed from the trailing edge of the wing. These vortices are carried downstream by the flow and influence the normal velocity distribution across the wing chord. The other factor contributing to this difference is due to the force required to accelerate the air particles from one steady state to another. The various contributions are expressed by the equation: L = La +LQ + L. (1) Reference will continually be made to these components, whose symbols and definitions are given below. Dynamic lift: L'. The complete lift experienced by the wing. Quasi-steady lift: LQ. The steady-state lift due to the normal velocity distribution of the gust. Apparent mass lift: Lao Lift acting on the wing due to the force required to accelerate the air particles from one steady state to another. Wake lift: Lw. Lift on the wing due to a normal velocity distribution across the wing, imposed by vortices in the wake. Lift lag: L'-LQo. The difference between the dynamic and quasisteady lift at any instant of time. WADC TR 57-401 5

Reference 8 develops the dynamic lift in terms of the three components listed above. The equation is: 2b 2b' ap s2b YQ XdX + pV, rQ + PV0 -J/2b) dS (2) where rQ is the quasi-steady circulation, 7Q is the quasi-steady vorticity distribution across the wing, and y7() is the vorticity distribution in the wake. The terms in the above equation are identified below: La P fa 2b Q XdX TLQ = p VWQ pV = 0 PO 2b QUASI-STEADY AND APPARENT MASS LIFTS Using the linearized theory for an airfoil of arbitrary camber distribution as given in Reference 34, the quasi-steady vorticity distribution: Q and circulation rQ for a normal velocity distribution g(x) are: f n ln cos n\ nl b:l1 sin 9 00 FQ _ 2icb Bn n=l bbn+l nBn 2 B n n n= 2- g(x) sinn9 nsin GdG X = 2b cos G.g(x) is the normal velocity distribution of the gust (see Figure 10). LQ is solved for directly upon knowing the g(x) distribution. rQ can also be solved directly) the differentiation with respect to time is related to the speed at which the gust is passing over the wing. WADC TR 57-401 6

WAKE LIFT:T1 vortlcity distribution in the wake is related tothe quasi-steady circulatio-by the Eelmholtz theremy and the.equation is:.+Vot.1/2 2b+Vt / rQ(t) + 2b Ot (2): = O (3) ^2)2bd = 0 Thu 7 () may b obtained for any function of r with time. The wake lift relltig frcom a unit step in. (Wagner proble) has 1een computed (see Reference 8), i.; ay be written as: LW(o^) = - pV; (cr) - ( 4) The:lfunction (:)- can be approximatedby: (.-a) O.,165e-0455 + o 355e-0O3:a (see Reference 9) For an arbitrary q distribution with time, the wake lift is found by using -e. superposition principle. The equations given above are based upon the following assumptions: 1) Flow is two-dimensional 2) Thin airfoil theory can be applied. ) Thewake remains flat - 4) The gust effect is that of a change of angle of attack across the airfoil,. It should be noted that all three terms on the right-hand side of Equation 2 A-re Beterined from qua.sisteady values. These values are fixed by the normal vel-lty distribution across the wing. TEST RESULTS MOVING BUMP GUST GENEATOR -The first of the mlving-b.ump tests were conducted in the small open-return tunlnl:d's ribed in Appendix I, The lift-sensitive wing had a chord of 6 inches a sa of 2 ihes two-dimens:ioal flow was maintaied by use of end plates, _T.-: gleo.f-attaek.distribution for the gust passing over the wing is shown in igiure 81 This.distribution passes over the wing at the speed with which the bump " movyed. Figure 7 showthat the bump is rapidly accelerated to a constant speed. WADC R 57401l 7

As the bump moves with a velocity VB, the normal velocity distribution as determined by quasi-steady measurements requires correction because the motion of the bump causes the streamlines to distort relative to the quasi-steady flow conditions (see Figure 11)4 The dynamic streamlines are flatter, and therefore the gust would be less intense than indicated by quasi-steady tests. Typical lift responses for dynamic and quasi-steady runs are shown in Figure 12. It is seen that the effect indicated in Figure 11 is not discernible for the blum speed used The experimental lift lag and theoretical lift lag are compared in Figure 13, It should be noted that the experimental ad theoretical lift lags agree well up to their maximum values. However, the experimental results fall off faster than those computed by theory. No explanation is apparent at this time for this lack of agreement. The maximum Reynolds number tested with the small tunnel was 148,000. A larger version that was identical in the gust mechanism was built to extend the Reynolds number. The tunnel and equipment are described in Appendix I. In Figure 7 the bump velocity is shown as a functionof displacement and time. From this curve it can be seen that the bump has reached nearly uniform speed within the first 1/2 chord of travel (12-inch chord). The gust input, expressed as angle of attack, is how as a function of bump position in Figure 8. These data include both quasisteady and dynamic runs. There is no correction necessary for the effect of the bump motion upon the streamlines, but it was necessary to make a correction to the dynamic position of the bump because of the stretch of the position-indicator cable. Tests were conducted at tunnel speeds of 72.5, 55, and 20 feet per second. Bump speeds of 1443, 12.5, and 10.1 feet per second were used. At lower speeds a superimposed oscillation was picked up by the wing. The fluctuation of this oscillation was such that the results were not conclusive. Due to insufficient time and money, only the results of the test at the speed of 72.5 feet per second are analyzed and presented herein. In Figure 14 are shown the quasi-steady and dynamic force responses as a function of time for a Reynolds number of 465,000, From these data the theoretical lift lag. is computed and the experimental lift lag obtained, as shown in Figure 1l It should be noted that the experimental and theoretical lift lag are very similar to those'meured in the 21-inch x 29-inch tunnel. After completion of lift-lag tests, tests were made of lift overshoot at stall The wing was set at a negative angle of attack sufficiently large to cause the wing to stall when the bump flow field was removed by moving the bump downstream (see Figure 16). The lift response is shown in Figure 17 for two bump speeds. In Figure 18, the ratio of CN-Ma (Dynamic Test)/ CN0Max (Static Test) is plotted as a function of rate of change of angle of attack with time. The data def initely show that there is a measurable lift overshoot and that the technique used is capable of measuring flow details of the phenomena. The data of Figure 18 indicate that the ratio CNx (Dynomic Test)/CNMax (Static Test) plotted against do/dt givesa simple straight line ucuve, within the limits of the tests. The data given in Figure 18 demonstrate that loads of the order of 35 percent greater than the steady-state values can be reached when an aircraft enters a severe iWADC TR 57-401 8

tt a high angle of attack. These results are preliminary, but they demonstrate afe ld.ofalpplication of the gust-generating equipment. Two- and three-dimensional should be studied in detail to determine design criteria for flight through rough air at high angles of attack.:X GUST GENERATOR -- ecause there is no way of making the flow pattern stand still with the vortex e or gust-simulating device, quasi-steady values of the lift cannot be obtained. A-des-cribed in References 32 and 33, the flow pattern and dynamic lift are measured Isiultaneously for each run. This makes repeatability unnecessary. l-'T- e sts were made at wind speeds of 15, 30, and 60 feet per second. Shown in Fgure 19 are representative oscillograms for the above tunnel-speed conditions. I Figures 20a, 20b, and 20c are shown the reduced data for these tunnel speeds. I-'In Figure 20a are shown six runs at a tunnel speed of 59.7 feet per second. Th variations in lift measurements follow the variations in flow as shown by the onitor ire. I-, 0 The hot-wire probe is located -2-1/2 inches ahead of the leading edge of the ifoil. Velocity measurements made at this station indicate that the same velocitfcurst there as occurs at the reference Pitot-tube point for the tunnel. Thus, f-ra tunne velocity,of 59.7 feet per second, 0.0035 second is required for the:Ilw pattern to travel from the probe to the, leading edge of the lifting surface. laly, it requires 0.0119 second for the pattern to travel from the probe to t-htrailing edge of the lifting surface. Referring to Figure 20a, it is observed.azero inclination of the stream occurs at 0.015 second while the lift crosses tezero axis at 0.042 second giving a time lag of 0.027 second. With reference o the mid-chord of the airfoil, the time delay is (0.0035 + 0.0119)/2 = 0.0077 0eond. The difference between the measured time lag.of 0.042 second and the time elay of 0.0077 second is attributable to the aerodynamic lift lag.;: -The maximum dynamic lift calculated from theory is about twice the experimenal -value shown in Figure 20a. Whether the theory is incorrect for velocity gradients-of the magnitude indicated here, or the instrumentation is inadequate, is un-'certain. CONCLUSIONS The feasibility of generating reproducible gusts of sufficient intensity and sc:ae at Reynolds numbers high enough to permit the extrapolations of measured unlift increments to full scale has been demonstrated. In a wind tunnel of Xlo turbulence, reasonable accuracy in the determination.of lift lag is demonstrated o1 1a.ai:rfoil of 1-foot chord at a Reynolds.s number of 450,000. No difficulties are anticipated in increasing the Reynolds number to several times this value. The "lift overshoot" at angles of attack above the steady-state stall was measWDC TR 57-401 9

Uired for two gust intensities. The maximum lift coefficient attained.: a value 35 percent above the steady-state value before stalling. A detailedj investigation of this phenomenon represents an important potential use of the facility. WADC TR T57401 10

REFERENCES 1. JlDonely, P., An Experimental Investigation of the Nomal Acceleration of an 12 Dtoneiy, P., Summary of Information Relating to Gust Loads on Airplanes, NACA TN..1976, _ 1949.'Britt, J. -B. and Scrtont, C, Measurements of Pithg Moment Derivatives for an Airfoil Oscillating About the Half Chord Axis, Br ARC RAM, 1921 jlaing Airfoil in Two-Dimensional Fo NACA T l 46, 1951 ~~ P~r~yr nes tigand Neilson instigzat C1ai s eady.Aer..... ic n, General Teory Aerodynamic Instaity and the Mechanism of lutter iNACA TR 996,6 193.:... v. ra, THrris and ears meail, "asirfments of Z wfor ao nirsimating Arfoil, Co cli ege 5of rt, 9-. Cranfield 0-... Tmne R. L, Epe Unsteady tl roynaic Deivatives of a Sinus idally 019il10. Luke, Y L., Tableof Coefficients for Compressible Flutter Calculations5 R-' -68 0......., Pitching Airfoil at Subs-onic ch lumbers from Oswoillating Coefficients with -" F~jutter, NACA TR 496', 935. 7.11 Garriek, I.,'E., anderubInw, S. I., T he oretical Study of Air Forces,on an Os-8...v.,KarCm. Th._ and.Sears. W. R., "Airfoil Theory for Non-Uniform Motion," J. Aro. Sci,. 5, 379-390, 19380:9: oes o R: T The Unst Lift of a Finit in tream NACA TR 681, 1940. 1-esR E, Effet o finite -pan on the Alrload Distribution for Oscillating i5^l:Biot 1311 Boektnle. Aerodynamiac; Theory of Oscillating Wings o Finite Span, 1i. Luke, Y aL, Table of Coeffiient Cretion ressible Flutter C aloulations s USAF WAD -. TR 68000. selsky, B, D-rin.tion of diial Lift and Moment of a Two-Dimensional Pitehi: Airfoil at. Sub nic eh Nbers from Ose illating Coefiients with /r~ilt~l Cald.ulation:s for M = 0.7, NACA TN 2613, 1952.. "t~: Biot, M. A., I"Loads on a Supersonic Wing Striking a Szp Edged Gust" J. Aero. 00, 6-' 296,300o 1949.!'Garriek I.'E, and. Ru binow, S. I.. Theoretieal t Study of Air Forces on an Oseillllert~.Oted Thin, in'a rSuoni-e Min:'Stream, NACA TN 1383, 19. >-,'ReSry:E. E lfer ot: Finite.pon the Airlod Distribution for Oscillating — s NACA "TN 11 - 1947; TN 1!i95, - 1947- (with Stevei, JO E)5 -Bio:t.ad Boeheia:AerOd.m,r y of Oscllagin Wings of Finite SWin, ~ci" t e ort No* o. -5.'i6o Wase: S,-.':.et Rio corrections in Flutter CaI'eulations, U~AF MR MCEXA:-:TR 57'7401 11

i7 bWaiker, Pe B., Grwth of Circulation About a Wing and Apparatus for Measuring Fluid Motin, Br ARC R and M 1402, 1931. 18. -Farren W. S., Reaction on a Wing Whose Angle of Incidence is Changing Rapidly, Br ARC R and M 1648, 1935. 19. Kue-the, A. M., Circulation Measurements About the Tip of an Airfoil During Flight hrough a Gust, NACA TN 685, 1939..2Q0. Bisplinghoff, R.- L.,'Isakson, G., and O:Brien, T. F., "Gust Loads on a Rigid Airpla.e with Pitehing Neglected," J. Aero.. Sci, 18,- 33-42, 1951. 21:KordesE., and Houbolt, J. C., Evaluation of Gust' Response Characteristics -ofd::e Existing Aircraft with Wing Bending Flexibility Included, NACA TN 2897, 9, 53. _:2.- Go-land M., M al.,. Effects of Airplane Elasticity and Unsteady Flow on Longi-:;:tdinal Stabilityy, dw"est Re-search Institute Project No. R108E-108. 4: -.: 2-5. St attler, I..'C.o., "Dynamic Stability at High Speed from Unsteady Flow Theory,''J. Aero Sci., 17 232-242, 1950. 2,4.,, Lyon,- H. M., and Ripley, J.., A General Survrey of the Effects of Flexibility of t-he Fuselage, Tail Un t, and Control Systems on Longitudinal Stability and Con-. trol: Br ARC R and M 2415, 1950. 2..~, J. 5-, Pai, S. I., and Sears, W.'Ro, Some Aeroelastic Properties of Swept Wings," J. AeroSci, e 1~6; 1.05-115, 1949 o -26f. S.hetzer, J o D., Damics for Aerodynamicis ts, Douglas Aircraft Company, S M,.. 41077y, 1951o 27" Donely, P., Effeetive Gust Structure at Low Altitude as Determined from the Reactions of an Airplane, NACA TR 692, 1940.::..- - -........ 28" Cleme-tson, G., A' IiVestigation of the Power Spectral Density of Atmospheric Turbulence MIT eSsis, 1950., 29-. Sherlock, R. H., Storm Loading and Strength of Wood Pole Lines and Study of Wind Gusts, Edison Electric Institute, 1936. 30. Br:RAE Report, A.o 2341. High Altitude Gust Investigatigation. 31 Hus: and' Portman Study of Natural Winds and Computation of the Austausch Tur-: b'.ulence Consant, Guggen eim Airship Institute, Report No. 149,:2.' Kethe A M, SchetZer,.J. D., and Garby, L., C., Research Design. Problems Re.'la to Facilities for.:Simatig the Ae rodynami Effects of Atmospherici Guss_ onAireraft Comonnents, The University of Michigan, Engineering Research Ititute Projeet-2099, Progress Report No. 4, 1953. WAflC:TR 57-401 12

3 Ku:ethe, A. M.,.Schetzer- JO D., and Garby, Lo C., Research Design fro~blems e-lain to IFacilities for' imulatin g the Aerodyaic EffetG of At ospheric Gts on Aircraft CopoBnents, The University of Michigan, Engineering Research Institute Project 2099, Progress Report No. 5, 1954.. v. Karman, Thl, and Burgers, J. M., -General Aerodynamic Theory of Perfect Thid s, Aeroaatic Theory, ed. W. 0. Durad., Vol. II, Calif. Inst. Tech., Paesadena 19430:'Af R 574701 13

APPENDIX I DESCRIPTION OF TEST TUNNELS A.: OPEN-RETURN TUNNEL Initial test work for both the Vortex and Moving Bump Gust Generators was done 1ina Sall open-return tunnel. A sketch of the tunnel is shown in Figure 3. Data 3tin5ent to this tunnel are listed below. Speed range: 0 - 65 feet per second Turbulence level: 0.03 percent (at) V = 50 feet per second Contraction ratio: 16 Nominal test section size:'21 inches high 29 inches wide 100 inches long I- i The tunnel was constructed so that test sections are interchangeable. One st sectiion contained the Vortex Gust Generator, and another contained a Moving p Gust Generatorn i Tu-nnel speed was controlled by changing the revolutions per minute of a fixed thfan. A Reeves Vari-speed drive is the mechanism of speed change. i CLOSED RETUDRN 5-FOOTx 7-FOOT TUNNEL WWork was continued with the Moving Bump Gust Generator on a larger scale. For tis, a tunnel was constructed as a joint effort between the Air Force and the Unieityo Previous experienc e had indicated the importance of keeping the turbulence l as low as possible. Figure shows the plan view and dimensions at critical stations Data pertinent to this tunnel are: peed range: 0 - 250 feet per second Turbulence level: 0o02 percent (at) V 100 feet per second 0.035 percent (at) V 200 feet per second Contraction ratio: 15 Test section size: feet high 7 feet wide 25 feet long Power plant: Ward-Leonard drive, 1250 horsepower available 300 horsepower required at maximum condition Fan: Fixed pitch, 10 blades of laminated wood 3Maximum diffuser angle: 505 degrees except in rapid expansion section Settling chamber: 5 screens, 30-mesh wire of 0,075-inch diameter spaced 30 inches apart. After the last screen there is a distance of 10 feet before the nozzle starts* C TR 57-401 14

An artist's sketch of the tunnel is shown in Figure 4b. Rails for the moving blmp to travel on are covered by the bottom corner fillets when gust tests are not being made. WADC TR 57-401 15

APPENDIX II INSTRUMENTATION DETAILS A. BUMP-POSITION INDICATOR To indicate the position of the bump in the test section for both dynamic and quasi-steady tests, a system as shown in Figure 6a was used. A cable attached to the nose of the bump is wrapped around the drum of a 10-turn potentiometer. Travel of the bump rotates the potentiometer. The electrical circuit used is shown in Figure 21. To position the bump for quasi-steady measurements, holes were drilled through the floor of the test section at given distances downstream. A tight pin inserted through the hole would allow the bump to rest at a known position. These holes were located within + 0.05 inch. B. ANGLE-OF-ATTACK MEASUREMENTS Angle-of-attack measurements were made using a hot-wire anemometer with a XC probe*. The circuit used is shown in Figure 22. This arrangement was necessary because steady-state angles of attack are necessary for the flow measurement. At a given tunnel speed the hot wire is calibrated by rotating it through known angles. Tests are then conducted at the same tunnel speed and wire-heating currents. C. SPEED MEASUREMENTS A check was made of the influence of the bump upon the local wind speed in the vicinity of the wing trailing edge. With the bump in a forward positionthere is a slight speed increase. This falls off rapidly as the bump is moved back. A comparison made between the quasi-steady and dynamic speed effect shows no measurable difference. The measurement was made by using a hot wire mounted in the horizontal plane. A direct-current amplifier was used to amplify the effect of steady-state velocity changes. D. BALANCE SYSTEM The balance arrangements used operated on similar principles, those of carrying the lift load into a flexure beam and measuring its deflection by means of a Schaevitz Transformer. Two-dimensional flow was maintained for the Vortex Gust Generator balance by using wing stubs that extended from the ends of the lift-sensitive wing to the floor and ceiling of the tunnel. These stubs were not in direct contact with the lift-sensitive wing, but the gap was kept very small. For the moving-bump system, end plates were used to keep the flow two-dimensional. Figure 23a shows details of a balance system for the Moving Bump Gust Generator. The arrangement of the stiff flexure beams was such that the model was sensitive only to the normal force. WADC TR 57-401 16

To11gie a good. dynamie response, the frequency of the wing-balance combination.l ke s:pt:s high as poassble The frequency response was.checked by using a mechan-:-shker and, for a given force input, measuring the balance output. The freuencyT response was found to be flat out to 40 cycles. A Fourier analysis of the gut function indicated that components beyond 10 cycles per second. are negligible. WADC TTR 57-401 17

U I l Co \J1 Fi TOP VIEW OF TEST SECTICN.~ll /~GUST GENERATOR TEST MODEL MECHANIVSM Figure la. Schematic diagram of Vortex Gust Generator.

o Vortex shed as slats move from a zero JobI~ ~ \ X position to a highangle |Q~~~~ >>^~~~~~ \ ^^~ ~of attack }\ ~ ^ ^ ^ \ Vortex shed as slats are returned to neutral Slots Figure lb. Sketch showing vortex pattern formed by action of the slats.

I-3 START FINAL POSITION igure 2. Normal velocity distribution for three positions of the Figure 2. Normal velocity distribution for three positions of the bump

-'' - 1. — - ~' I I l-: L. 8 a lp''446 Ix I"x 1x 0" HONEYCOMB o / 18"TYR 28 MESH SCREENS k^~~~~~~~~~ a l ~~~~~~~~~~~~~~~ rt / ~4 BLADE FAN I,_I- I IF ~ ~r^^~~~~~~~~~~~~~~~~~~~- 55 H.P REVES VARI-SPEED DRIVE Figure 3. Open-return tunnel. The Vortex Gust Generator is shown. The Moving Bump Gust Generator is built into another test section that is interchangeable with the above.

::,:ii^4:i10 _.i 5X 28Mesh: 0.a0075 Dia.^^^ or) 00 ro 1-28 Mesh Screen ~ T Sec t i I 1 0' 1 _"5-28 Msh, 0.0075' Dia. ^ ^... 1", I~ ~ r. Sc eens Lul. Sta. (3)a(14) Sta (15)8(1) Sta. (Q)ua( Sta.(3) Sta.-ethru~0 Sta. (1)8( ______________ ^~ "^^^ GROUND LEVEL Fgr. I v o 6 - 30-54 Figure 4a. Plan view of 5-foot x 7-foot gust-generator model.

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Photograph showing bench calibration of Vortex Gust Generator and the associated electrical equipment.

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0 F-J. -,DIRE F DIRECTION ro A. TST MODEL.. POTENTIOMETER TO DETERMINE POSITIOr/N OF B UMP. C. MOVING BUMP. D. SHOCK ABSORBING CYLINDER. E. ACCELERATING MECHANISM (AIRCRAFT SHOCK CORD e WINDING DRUM. P. f rc -/ Figure 6a. Schematic diagram of Moving Bump Gust Generator.

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Figure 6c. Photograph of moving bump in 5-foot x 7-foot tunnel.' 60" 22 " 1-~< 86" -i Figure 6d. sketch of moving bump in 5-foot x 7-foot tunnel., WADC TR 57-401 30

3.5 Symbol Chord Length Gust Generator Model v F^ -12" 5'x5 7' +77 0 6" 21"x 29" ^ 2.5 i o: f t I.-F~~~~L Ii''.': "; ffi. 0.10 0.15 0.20 0.25 0. TIME (SEC) Figure 7. Bump position in chord lengths versus time. WADC TR 57-401 31

0 wx b a: 7 ~^^ o 5'x 7' Bump, dynamic and W o6't / quasi- steady runs -jO 5 LLa 4- ^^0 W^5 3 21 "x 29 Bump tests ox.%,, 3 2I | | 0^ 0 0 1 2 3 4 BUMP POSITION -CHORD LENGTHS Figure 8. Flow inclination in degrees versus bump position in wing chords.

toFr~~ a rA o 0 SYMBOL X STATION H-_.06.08 10.12 ^.14/ 16 18.20 20 /I1 --—..30 40 -8 _____ ____________50 ~~~~~ C') CDfor various stations along tunnel axis. LL. -.. Figure 9. Flow deflection for Vortex Gust Generator versus time for various stations along tunnel axis.

Model of the wake - - -2b I I -b +2b x= 2b Cos. Model of gust impact Figure 10. Sketch showing mathematical model of lift-lag system. WADC TR 57-401 34

4 — 0 f f i - --:.c.. Quasi - Steady Flow Pattern - __-___~ - Dynamic Flow Pattern for Bump at same position Figure 11. Quantitative effect of bump position upon flow-streamline distribution.

H|^~~~~~~ 1 ~~21"x 29'TUNNEL ~,kn^~~~~~ 1::TEST-SECTION VELOCITY -44.4 FPS -,^ 1.^~2 ^REYNOLDS NUMBER-148,000 o' MAXIMUM BUMP SPEED-10.1 FPS 40 o o DYNAMIC RUNS Z1.08 —-' t,. - 1:; QUASI-STEADY LIFT —uo 1 35 Z 8:::o ^ BUMP POSITION - n.6 Im 4 25.2. 20 0.05.10.15.20.25.30.35 TIME- SECONDS Figure 12. Measured dynamic lift, quasi-steady lift, and bump position versus time for a test-section velocity of 44.4 feet per second.

.12,v7I' / \21"x 29 TUNNEL o.I I TEST-SECTION VELOCITY- 44.4 FPS'I/ \ ~REYNOLDS NUMBER -148,000.10 EXPERIMENTAL.09__ / I\. THEORETICAL | / \ s BASED ON MEASURED X~OQ_.08^~~ > | / \ \ STREAM INCLINATION W.08 O f \ 0 *BBASED ON MEASURED LQ Z'Q 0.06.04 ^/I,// \%.03 1 I -'.02 1.0.05.10.15.20.25.30 TIME -SECONDS Figure 13. Comparison of experimental and theoretical lift lag versus time for test-section velocity of 44.4 feet per second.

5'x 7' Tunnel Tunnel Speed 72.5 fps 7 \ Bump Speed 14.3 fps Reynolds Number 465,000.zX1 \ \ Dynamic Quasi-Steady 114 —- 40 3 30 0??-I.: W -Bump Position ~ N20 __ _ _ 10 ^^X~~~ = 0.05.1.15.2 TIME IN SEC Figure 14. Measured dynamic lift, quasi-steady lift, and bump position versus time for test-section velocity of 72.5 feet per second. WADC TR 57-401 38

5'x7' Tunnel; Tunnel Speed 72.5 fps; Bump Speed 14.3 fpss,.:Reynolds Number 465,000 - lH.8.7 Theoretical based on measured LQ z.6 1 -.5.4 C) ('3 Experimental I-..I 0.04.08.12.16.20.24.28 TIME- SEC Figure 15. Comparison of theoretical and experimental lift lag for Moving Bump Gust Generator.

IF Bump in mid position Bump in aft position Figure 16. Sketchshowing flow patterns at three bump positions for lift-overshoot experiment. WADC TR 57-401 40

12.60 12,60 130 ). 10.50 Nom o Bump speed A -II*/Sec z 0 J.1.2.3.4.5 - T IME IN SECONDS -t Figure 17. Lift response at two bump speeds in region of wing stall. 0.. a: 0..I <Nominal tunnel speed 40' /Sec 0 Bump speed a -I1' /Sec 1.4 is."~~~~0 7I Se I./0 0.20 30 40 50: -:ATRATIE OF CHANGE OF ANGLE OF ATTACK ON WNG-DEG/SEC Figure 18. Effect of rate of change of angle of attack upon maximum normal force coefficient. WAC TR.57-401 41 z 1.2 10 20 41

C) 0:-I a b c Tunnel speed = 59.7 ft/sec. Tunnel speed = 30 ft/sec. Tunnel speed = 15 ft/sec. Dashed line is a 50-cps time pulse. Dashed line is a 50-cps time pulse. Dashed line is a O5-cps time pulse. Monitor probe trace is first trace Monitor probe is upper trace in Monitor probe is upper curve in to cross time-pulse line in both both pictures. both pictures. pictures. Lifting surface response is lower Lifting-surface response is lower Lifting surface trace is second curve in both pictures. curve in both pictures. trace to cross time-pulse line in both pictures. Figure 19. Oscillograms of the lifting-surface and monitor-probe esponses to the Vortex Gust Generator.

0 4 TEST VELOCITY 59.7 f.p.s \ > \ Z0. I 10,4 ~1 z _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ (r0 ^. ~ ^ ~ ~ 02.06.08 ~.10, 12.14 TIME IN SECONDS -.208 TEST VELOCITY 30 f.p.s.06 ~~6.04 b~ / 2.01~ ~ 02 \.O' j / TIME IN SECONDS z_.3' 1 _ _.o _.1 _14 TIMIN SECON'D15 C.p.: I-.02 0.01 W.0.02. *.... D Figure oo WC5 0.10 0.15.20 025 Q30 038 0.40

Hg~~~~~~~~~~~~ ~~~~~Bump Pin Inserted to position-bump for quasi-steady runs. Pin remove for dynamic tests I 0 Turn Cable -, —potentiometer 24 v Recording oscillograph or oscilloscope Figure 21. Sketch of bump-position mechanism.

^-1 1 ~~ Potentiometer used to ^~~~~~~~~f~~~~~~~~ I y~reduce unbalance 4::0 / ~* ^ * ~ Hot wires I-0 V1~~~~~1 ^ H'1I11H D.C. Amp Oscilloscope with Polaroid- Land Camera or oscillograph Heating circuit of Thile Wright hot wire anemometer Figure 22. Electrical circuit used to measure flow angles.

i ~ ANGLE OF ATTACK PLATE o BALANCE BLOCK 0 Js ~END PLATES ROD THROUGH CENTER OF WING, NOT IN CONTACT WITH WING, TO HOLD END PLATES APART — 0N WING ^'/ ANGLE OF ATTACK ADJUSTING MECHANISM Figure 23a. Exploded view of balance as used in Moving Bump Gust Generator.

BALANCE BLOCK -'t^ ^^^-^^ ~"LIFT PLATE,,^", -e:,.:,l3 ^WING ATTACHED TO THIS PLATE 1^ YtL\ x- MOUNTING HOLES TO g W~ANGLE OF ATTACK PLATE ATTACHED TO BB^^/ /.~; ^ll^< BALANCE BLOCK FLEXIVE BEAM ATTACHED TO LIFT PLATE SCHAEVITZ TRANSFORMER PICKUP MECHANISM D.R.NEWTON SCHAEVITZ TRANSFORMER AND ZEROING MECHANISM Figure 23b. View showing flexure-beam block and Schaevitz Transformer mounting. WADC TR 57-401 47