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CURRENT NACA REPORTS NACA TN 2955 ESTIMATION OF FORCES AND MOMENTS DUE TO ROLLING FOR SEVERAL SLENDERTAIL CON FIGURATIONS AT SUPERSONIC SPEEDS. Percy J. Bobbitt and Frank S. Malvestuto, Jr. July 1953. 71p. diagrs. (NACA TN 2955) A method, based on conformaltransformation techniques, for solving twodimensional boundary value problems has been used to evaluate the velocity potentials, span loadings, pressure distributions, and associated stability derivatives for several slender tail arrangements performing a steady rolling mo tion. Illustrative variations of the rolling stability derivatives for several series of tail shapes, as well as sample span loadings and pressure distributions, are included. NACA TN 2967 AN ANALYSIS OF THE POWEROFF LANDING MANEUVER IN TERMS OF THE CAPABILITIES OF THE PILOT AND THE AERODYNAMIC CHARAC TERISTICS OF THE AIRPLANE. Albert E. von Doenhoff and George W. Jones, Jr. August 1953. 42p. diagrs. (NACA TN 2967) A theoretical analysis of the poweroff landing maneuver is presented In which an attempt is made to consider both the aerodynamic characteristics of the airplane and the human capabilities of the pilot. Certain factors are derived which allow for the var intions in pilor judgment in a manner which is di rectly connected to pilot capabilities; however, fu ture research Is needed for the actual numerical evaluation of the factors. An analysis including these factors is set up by means of which the mimi mum length of landing field required for landing an airplane safely could be found. NACA TN 2976 A STUDY OF THE STABILITY OF THE INCOM PRESSIBLE LAMINAR BOUNDARY LAYER ON IN FINITE WEDGES. Neal Telervin. August 1953. 41p. diagrs., 4 tabs. (NACA TN 2976) Hartree's numerical solutions of the boundary layer equations for the flow over infinite wedges are used to confirm a result prenlously obtained by the use of Schlichting's approximate method for the cal culation of the laminar boundary layer; namely, that in a region of falling pressure a thick velocity pro file can be more stable than a thin profile although the velocity at the edge of the boundary layer and the pressure grad~ent are the same for both profiles. The Investlgation also leads to the inference that the calculated effects of a change in boundarylayer thickness on the stabusity and on thees Reynolds number should beeseai by replacing the Schlichting as qi ameter fiahqqy of velocity profiles by the Har d ~singleparameter  family of velocity profiles. NACATNBBO ( DEC y 355 THE AERODYNAMIC CHARA E IC O A ASPECTRATIO20 WING HA N FOL SECTIONS AND EMPLOY BOUNDARY.* LAYER CONTROL BY SUCTION. B~me C ocke, Jr., Marvin P. Fink and Stanley M. Gottlieb. August 1953. 63p. diagrs., photos., 2 tabs. (NACA TN 2980) An investigation has been conducted to study the aerodynamic characterishecs of an aspectratio20 wing employing thiek airfoil sections and boundary layer control by suction. Data from models tested In the Langley fullscale tunnel and the Langley low turbulence pressure.tunnel are included in this re port The results indicate the effects of varylng suction flow rate, suctionslot configuration, wing surface condition, fap deflection, and Mach number. NACA TN 2982 SUPERSONIC FLOW PAST OSCILLATING AIRFOLS INCLUDING NONLINEAR THICKNESS EF FECTS. Milton D. Van Dyke. July 1953. 41p. diagrs. (NACA TN 2982) A solution to second order In thickness as derived for harmonically ascillating twodimensional airfoils in supersonic flow. For slow osnilations of an arbitrary profile, the result rs found as a series in cludang the third power of frequency For arbitrary frequencies, the method of solution for any specific profile Is indi~cated, and the exp~licit solution derived for a single wedge. Nonianear thickness effects are found generally to reduce the torsional damping, and so enlarge the range of Mach numbers within which torsional Instabdlity Is possible. NACA TN 2985 A FLIGHT INVESTIGATION OF THE EF FECT OF STEADY ROLLING ON THE NATURAL FREQUEN CIES OF A BODYTAIL COMBINATION. Norman R. Bergrun and Paul A. Nickel. August 1953 27p. diagrs., photo 2 tabs. (NACA TN 2985) Flight data have been obtained with a freely falling bodytail combination to show the effects of steady * AVAILABLE ON LOAN ONLY. ADDRESS REQUESTS FOR DOCUMENTS TO NACA, 17P4 F ST, NW. WASHINGTON 9s. D C, CITING CODE NUMBER ABOVE EACH TITLE; THE REPORT TITLE AND AUTHOR. National Advisory Committee For Aeronautics Research Abstracts NO.47 AUGUST 19, 1953 RESEARCH; ABSTRACTS NO. 47 NAA TN 2992 APPLICATION OF A CHARACTERISTIC BLADE TOBLAD)E SOLUTION TO FLOW IN A SUPER SONIC ROTOR WITH VARYING STREAMFILAMENT THIICKNESS. Eleanor L.Costlow. Auguatl 95. 36p. diagrs., 5 tabs. (NACA TN 2992) An analysis of the circumferential bladetoblade flow properties in a supersonic impeller has been made using the method of characteristics on an arbi trary stream surface of revolution. The method takes into account variable streamfilament thick ness and curvatrire along the flow surface. The re sults of an example considering streamfilament thickness variation alone indicate an appreciable dif  ference between the flow properties calculated with the planeflow characteristic equations and those de termined with the characteristics method demon strated here. The effect of a streamfilament thickness reduction from bladepassage inlet to exit was to reduce the relative velocity and the absolute value of the flow angle in the blade passage. NACA TN 2993 CALIBRATION OF ST'RAINGAGE INSTALLATIONS IN AIRCRA FT ST RUCT URES FOR THE MEASURE  MENT OF FLIGHTI LOADS. T. R. Skopinslu, William S. Aiken, Jr. and Wilber B. Huston. August 1953. 70p. diagrs., 10 tabs. NACA TN 2993. Formerly RM L52G31) A general procedure is developed for calibrating straingage installations in aircraft st ructures for application to flight measurements of loads. The basic procedure can be modified as necessary to adit the requirements of any particular structure. The application of the procedure is illustrated by results for two typical structures. NACA RM E53E28 EXPERIMENTAL, INVESTIGATION OF THE EF FECTS OF SUPPORT INTERFERENCE ON THE PRESSURE DISTRIBUTION OF A BODIY OF REVO LUTION AT A MACH NUMBER QF 3. 12 AND) REY NOLDS NUMBERS FROM 2 x 10" to 14 x 106. L. Eugene Baughman and John R. Jack. August 1953. 18p. diagrs., photo., tab. (NACA RM E53E28) An experimental investigation wasaeperformed to de termine the effect on base and forebody pressures of using a sting Imodified wEith varying length splitter plates and fins instead of a conventional sting to sup port a conecylinder body of revolution. The in vestigation was conducted at a Mach number of 3. 12 for a Reynolds number range of 2 x 106 to 14 x 106 and for an angle of attack range of Oo to So. For Reynolds numbers of 8 x 106 and 14 x 106 there was a negligible effect of the splitter plate modification on the base pressure, and at a Reynolds number of 2 x 106 tere was asmall effect. Positioning the leading edge of the splitter plate at or ahead of the base made no appreciable change in the influence of the modifications pn base pressure at a Reynolds number of 14 x 10 With the fintype modification there was a small increase in base pressure. rolling on the response in the pitching and yawng motions. Observed effects are compared with those predicted by theory, and reasonably good agreement is obtained. Charts are also presented to provide a rapid means for estimating the effective values of the saii~yety deivatve Cm, and C"B under steady rolling conditions. NACA T`N 2988 TEMPERATURES, THERML STRESS, AND SHOCK IN HEATGENERATING PLATES OF CONSTANT CONDUCTIVITY AND OF CONDUCTIVITY THAT VARIES LINEARY WITH TEMPERATURE. S. V. Manson. July 1953. 62p. diagrs. (NACA TN 2988) Workig formulas are presented for the steadystate temperatures and thermal stress in heatgenerating infinite plates of constant conductivity, and of con ductivity that decreases linearly with temperature as the temperature increases, for the case in which the heat is generated uniformly throughout the plate thickness and both faces of the plate are equally cooled. A criterion is indicated for determining the surface cooling conditions under which the thermal shocks at the surface and midiplane will be smaller than, equal to, or greater than the steadystate thermal stresses at those planes. The dimension less parameters governing the transient tempera tures and thermal stresses in materials of linearly varying conductivity are derived by a similarity study of the conduction equation and boundary con ditions of the transient state. A numerical tech nique for solving the transientstate equations is indicated in detail. N~ACA TN 2989 COMPARISON OF SECONDARY FLOWS AND BOUNDARYLAYER ACCUMULATIONS IN SEV ERAL TUJRBINE NOZZLES. Milton G. Kofakey, Hubert W. Allen and Howard Z. Herzig. Augus' 1953. 58p. diagrs., photos., 3 tabs. (NACA TN 2989) An investigation was made of losses and secondary flows in three different turbine nozzle configura tions in annular cascades. Appreciable outer shroud loss cores (passage vortices) were found to exist at the discharge of blades which had thickened suction surface boundary layers near the outer shroud. Blade designs having thinner boundary layers did not show such outer shroud lose cores but indicated greater inward radial flow of low mo mentum air in the blade wake which resulted in greater contribution to inner shroud loss regions. The blade wake was a combination of profile loss and low momnentumn air from the outer shroud, and the magnitude of the wake loss is to this extent an indi cation of the presence or absence of radial flow. At a higher Mach number, shockboundarylayer thick ening on the blade suction surfaces provided an additional radial flow path for low mnomentumn air which resulted in large inner shroud loss regions accompanied by large deviations from design values of discharge angle, NACA RESEARCH ABSTRACTS NO.47 NACA RM E53 FO2 PRELIMINARY RESULTS OF HEAT TRANSFER FROM A STATIONARY AND) ROTATING ELLIPSOI DAL SPINNER. U. von Glahn. August 1953. 35P diagrs., photo., 2 tabs. (NACA RM E53FO2) The convective heat transfer coefficients were de termined for an ellipsoidal spinner of 30inch maxi mum diameter for both stationary and rotating oper ation. The range of conditions studied included air speeds up to 275 miles per hour, rotational speeds up to 1200 rpm, and angles of attack of zero and 4o The results rndicate that a higher heat tranfer oc curred with rotation of the spinner. Transition from laminar to turbulent flow occurred over a large range of Reynolds numbers primarily because of sur face roughness of the spinner. BIRITISH REPORTS N23901 Royal Aircraft Establishment (Gt. Brit.) FUNCTION GENERATORS BASED ON LINEAR IN TERPOLATION, WITH APPLICATIONS TO ANA LOGUE COMPUTING. E. G. C. Burt and O. Hi. Lange. April 1953. 2p. diagrs. (RAE Tech. Note GWV 244) The use of function generators in electronic analogue comrputing and simulation greatly extends the range at problems which can be solved by these methods. This paper presents a technique in which diode units are used to approximate to the functions by linear in terpolation. It is shown that the method can be ex tended to deal wiith a wide class of functions, in cluding multvaiate functions. Analogue multiplica tion and division are discussed as particular cases of funtion generators, and formulas for the general function are developed. The results are presented of an experimental generator for sin x in the range  p Ex p, in which the error is about 1 percent of the umamu output* N23902* Royal Aircraft Establishment (Gt. Brit. ) MOVING; PARTIAL RANGE SMOOTHING WVITH CENTRALDIFFERENCES USING A NATIONAL ACCOUNTING MACHIN~E. J. D. Downes. April 1953. 33p. diagrs., 12 tabs. (RAE Tech.Note GW 250) Based on a suggestion by Dr. Th. W. Schmidt of G. W. Trials Division, a method of partial range smoohing of experimental data by central differences which greatly reduces the normal amount of labor and permits simple checking, is described. The method is particularly suited for use where a National Accounting M~achine as well as a desk cal culator is available. The observed quantities are converted to linear combinations of their difrences and used in an adaptation of the 'leastsquares" method of curvefitting. Additional, relevant appli cations of the National Machine are included in nuerical exape. N23929e Aeronautical Research Council (Gt. Brit. ) THE APPLICATION OF THE: POLY'GON METHO TO THE CALCULATION OF THE COMPRESSIBLE: SUBSONIC FLOW ROUND TWODIMENSIONAL PROFILES. L.C. Woods. 1953. 31p. diagrs., 7 tabs. (ARC CP 115) This paper describes the method now used by the author of applying the polygon method to the calcula tion of the compressible subsonic flow round two dimensional airfoils. Tables have been constructed which can be used for all airfoil shapes, putting the polygon method on the same footing numerically with Goldstein's aApproximation IDI for incompressible flow. A method of calculating lift and moment coef , ficients, and their rates of change with incidence is also given in the paper. N23930* Aeronautical Research Council (Gt. Brit. ) WAKE SURVEY AND STRAINGAUGE MEASURIE MENTS ON AN INCLINED PROPELLER IN THE R. A. E. 24FI'. TUNNEL. PART I. WAKE SUR VEY. J. G. Russell. 1953. 61p. diagrs., photos., 4 tabs. (ARCCP 117) This report describes tests carried out in the RAE 24foot wind tunnel with a 16 foot diameter, 4 bladed propeller, during October and November 1949. Wake survey and blade straingage measurements were made at tunnel speeds of 100 and 170 fps with the propeller axis inclined at angles of Oo, 5o, 100, and 150 to the air flow. The blade angles and propeller rotational speeds were also varied within the limits imposed by the 1, 500 hp electric motor, The lift grading curves at the points of maximum. and minimum loading, derived from total head meas urements made in the slipstream by means of a pilot comb, have been compared with estimated values, and estimated power absorption figures compared with measured values. N23931* Aeronautical Research Council (Gt. Brit. ) A COMPANION OF CALCULATED AND MEAS URED BASE, PRESSURES OF CYLINDRICALLY BASED PROJECTILES. W. F. Cope. Appendix: CALCULATION OF REYNOLDS NUMBER EFFECT ON PROJECTILES AT SUPERSONIC SPEEDS. W. F. Cope. 1953. 12p. diagrs., photo. (ARC CP 118) In 1946 at the VI International Congress for Applied Mechanics a paper wras read on a method of calcula ting the base pressure of a cylindrically based pro jectile. At that time there were few written base pressure determinations and therefore it was not possible to check completely the theory put forward. Since that time measurements have been made. In this report these measurements are compared with results of calculations according to the theory. The original report, "Calculation of Reynolds Number Effect on Projectiles at Supersonic Speeds, is in cluded as an appendi. NACA RESEARCH ABSTRACTS NO.47 This is an attempt to extend the usual blockage for mulas to high subsonic speeds. It la based on an assumption which perhaps can only be fully justified by results. The conclusion reached seems edIfl eiently reasonable to suggest that the method might be worth taking further either analytically or ex perimentally. N23936* ARECTU U VRE Sa he Mdathe~m ties Division. 1953. 18p. (ARC R &M2461. Former ly ARC 10, 754; Perf. 384; & C '2138) Critical tables are given from which the complete downwash due to a unit rectangular vortex can be read to three places of decimals, on selected lines coplanar with the vortex. These lines are at speci fled integral multiples of the semlwidth of the vortex from the center line of the vortex, and are of imr portance in the solution of wing loading problems by vortex lattice theory. N23937* Aeronautical Research Council (Gt. Brit.) THE EFFECT OF CURVATURE OF SURFACE ANDI THICKNESS OF TRAILING EDGE ON AILERO HINGE MOMENTS. Part I. MODIFICATION TO AILERON ON UPPER SURFACE. A. S. Batson, C. H. Burge and J. R. Greening. Par II. MODIFI CATION TO AILRON ON LOWER SURFACE A. S. Batson, C. H. Burge and W. C. Skelton. Part II. MODIFICATION TO AILERON ON BOT SUR FACES. A. S. Batson, C. H. Burge and W. C. S Ilton. 15350 48p~od Ar.,ht. 018tabs.127 ARC 5490; & C 1200; ARC 6188; 86 C 1426) This investigation was made to provide additional data on the effect on hinge moment of rounding the upper surface and of thickening the trailing edge of an aileron. Hinge moments were measured on an aileron (24.6 percent balance), and on ain aileron (271/2 percent balance) with various modifications to the upper surface, lower surface, and both sur faces. Strips (approximately 1 percent of aileron chord) were also fitted for a few cases near the trailing edges. N23938* Aeronautical Research Council (Gt. Brit.) THE THEORETICAL ESTIMATION OF POWER REQUIRMENTS FOR SLOTSUCTION AEROFOILB, WITH NUMERICAL RESULTS FOR TWO THICK GRIFFITH TYPE SECTIONS. J. H. Preston, (AR g M 25d7A{ orel 11@ 01 2.8027 gra. 276; FM 1052; ARC 11,610; Perf. 466; FM 1263) This report describes a method for assessing the performance of slotsuction airfoils in terms of an effective drag coefficient, which takes into account the power requirements of the suction pump neglect ing slot entry and duct losses. When the suctionalot is located at a velocity discontinuity the suction flow N239324 Aeronautical Research Council (Gt. Brit.) EFFECT OF MEAN STRESS ON THE FATIGUE STRENGTH OF D. T.D. 364 ROUND BARS WITH AND WITHOUT TRANSVERSE HOLES. G. M. Norris. 1953. 15p. photos., diagrs., 3 tabs. (ARC CP 120) Endurance curves for zero mean stress and for ten sile mean stresses of 5 tons/in2 and 10 tons/in2 wr eermined tor round a eimennc em hined from dettermined also for similr specimens with a trans verse hole. Variations in surface roughness caused some scatter in results and an attempt is made to correlate degree of surface finish with fatigue strength. N23933+ Aeronautical Research Council (Gt. Brit. ) THE LOW SPEED PERFORMANCE OF A HELI COPTER. A. L.Oliver. 1953. 12p. diagrs. (ARC CP 122) The analysis and estimation of helicopter perform ance is dependent upon the accurate assessment of rotor induced velocity. An empirical curve relating the flow through the rotor to the flight speed is used for vertical flight and the momentum theory is suf  ficiently accurate for a tip speed ratio greater than about 0. 1, but no simple method has been generally available for the intermediary speed range. Sets of empirical curves covering this speed range and based on the analysis of lowspeed flight perform ance are given in this report. The charts give values of the rotor induced velocity varying smooth_ ly from the vertical flight state to the forward flight region in which the momentum theory becomes ac curate. The charts are presented in forms suitable for determining steady flight performance and also for estimating rotor thrust during accelerated motion in, for exaple, takeoff flight. . N23934* Aeronautical Research Council (Gt. Brit. ) THE USE OF INFLUENCE FACTORS IN PROBLEMS OF FLUID FLOW. K. H.V. Britten. 1952. 13p. diagrs., 2 tabs. (ARC R& M 2441. Formerly ARC 10, 500; FM 1094) A vigorous mathematical analysis has been applied to some empirical reanlts, obtained by Thom, on the use of the "squares method" in problems of fluid flow. The problem considered is that of determining the incompressible and compressible flow past an arbitrarily shaped body placed in a wind tunnel. N23935* Aeronautical Research Council (Gt. Brit. ) TUNNEL WALL ]EFFECT FROM MASS FLOW CON SIDERAIONS. A. Thom. 1953. 10p. diagrs., 2 tabs. (ARC R &M2442. Formerly AC11, 004; FIM 11731 TP 211) NACA RESEARCH ABSTRACTS NO.4~7 required to prevent separation can be calculated, using the elementary theory suggested by Sir Geoffrey Taylor. The method is applied to two Griffith type airfoils and the drags are compred with those of normal thin airfoils 20 percent thick N2393g" Aeronautical Research Council (Gt. Brit.) ELECTRONICS APPLIED TO THE MEASURMENT OF PHYSICAL QUAN~T~ITIES G. E. Bennett, G. R. Richards and E. C.Voss. 1952. 1214p. diagrs., photos. (ARC R M 2627. Formerly RAE Instn. 1) The report describes the application of electro mechanical and elect ronic principles to the design of instruments for the measurement of physical quan tities such as movement, strain, pressure, accel eration, and vibratory motion, writh particular refer ence to the special requirements of aeronautical en gineering. The dynamic characteristics of pickups are considered, and subdivided on an electrical basis into electromagnetic, capacitance and resist ance types, a detailed description of each type being given. An account is given of the circuits used for the conversion of the electrical variation produced in each type of packup into a corresponding voltage or current, particular mention being made of bridge circuits and resonance circuit methods. The spec ial requirements of amplifiers, and the best basic circuits for satisfying them, are considered and il lustrated by detailed reference to a number of par ticular amplifier designs; in particular, direct coupled and carrier amplifiers are considered. The requirements of recording equipment and the vari ous recording methods are discussed, and a de talled account given of photographic recording and various oscillograph cameras, their optical arrange ments, components and timing devices. N23940 Aeronautical Research Council (Gt. Brit.) TOWING TANK TESTS ON A LARGE SIXENGINE FrLYING BOAT SEAPLAE, TO SPECIFICATION 10/46 PRINCESS. PART I. GENERAL PORPOISING STABILITY, TRIM AND SPRAY CLEARANCE* A. Gi. Smith, G. L. Fletcher, T. B. Owen and D). ~F. Wright. 1953. 29p. diagrs., photos., 6 tabs. ( RR &M. 2641; 11, 462. Formerly RAE Aero This report gives the results of the first series of towing tank tests made at the Royal Aircraft Estab lishment Towing Tank (up to May 1947) on a powered dynaic model of a sixengine transport flying boat,, later naed the Princess class, and designed to specification 10/46, on the basis of whch fullscale hull construction was started; latter tests have been made to further improve the hull step and afterbody and test the effect of modifications to the aerodynam ic superstructure and power units. N26032* National Gas Turbine Establishment (Gt. Brit.) THE THERMIODYNAMYICS OF FRICTIONAL RE SISTED ADIABAIC FLOW OF GASES THROUGH DUCTS O]F CONST~ANT AND VARYING CROSS SEC TION. W. R. Thomson. September 1952. 45p. diagram. (NGTE R.110) The report presents an analytical study dealing with the adiabatic flow of gases with frictional losses through ducts of constant and varying cross section. The thermodynamic treatment is along lines pub lished by other workers such as Bailey and Fabri and is essentially onedimensional in character insofar that frictional effects are assumed to be uniformly distributed over the total crosssectional area of flow. With this simplifying assumption, relation ships are deduced connecting the pressure, tempera ture, velocity and flow area of the gas at any one plane wcith those at any other plane in a duct. The main relationships are unusable for quantitative es timation except through graphs and the main value of the report lies in the presentation of these graphs, the use of which should facilitate the solution of duct flow problems. N26067* Aeronautical Research Council (Gt. Brit. ) ON BOUNDARY LAYERS AND UPSTREAM IN FLUENCE. I. A COMPARISON BETWEEN SUB SONIC AND SUPERSONIC FLOIWS. M. J. Lighthill. October 23, 1952. 17p. diagrs., 3 tabs. (ARC 15, 297; FM 1805) It is pointed out that there are two separate mechan isms for upstream influence through the boundary layer in supersonic flow, and that one of these (that involving separation) operates also in subsonic flow. A quantitative theory of subsonic flow up a step is given to illustrate this. The main differences be tween the subsonic and supersonic flows are as follows. (1) The boundaries of dead air regions are nearly straight in supersonic flow but are usually highly curved in subsonic flow. (2) Separation (whether of the laminar or turbulent layer) occurs at a much lower pressure coefficient in supersonic flow; this is only slightly due to the fact that the fluid near est the wall is then lighter and so more easily brought to rest; it is due much more to the relative sudden ness of the pressure rise ahead of the dead air region. (3) However, for a given pressure coefficient in the dead air region, the distance of upstream influence is somewhat greater in the subsonic flow, except at the higher pressures. N26085* National Gas Turbine Establishment (Gt. Brit. ) A METHOD FOR MEASURING THE REACTION FORCE OF A HIGH PRESSURE GAS STREAML. P. J. Fletcher. March 1953. 9p. diagrs. (NGTE Memo. M. 184) A method has been developed for measuring the re action force produced by a high pressure gas stream. Briefly the method is to utilize the jet reaction to produce a bending moment in the gas supply pipe by arranging for the jet to discharge in a direction at right angles to that of the entering air. The strain due to bendig is indicated by resistance strain gages and the reaction force deduced from direct calibra tion. The method thus obviates the need for flexible joints and hence eliminates the problem of ensuring flexibility and sensitivity at high pressure. A de tailed description of the method of installing the gages is given as well as the precautions necessary NACA RESEARCH ABSTRACTS NO.47 to separate the required bending strain from that due to pressure forces. With the particular apparatus described a thrust of up to 150 lb produced by a gas stream at 10 atm pressure can be measured to writh in +0. 21b. NACALangley 81953 4000 
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