H..E COPY a. NO 3 J CONFID::TTIA1 RN ITo. L6H28c

NACA 1LEFp4LE CASEF 1 C,( , RI y- r RESEARCH MEMORAND)b141ON LANGLL AY

FOR AERONAUUS Langley Field Virginia

I1 74VE3TIGATIC OF DIVE BRAkES A.-'D A DIVE-RECOVERY FLA? ON A HIGH-ASPECT-RATIO WING IN THE LANGLEY 6-FOOT HIGH-SPEED TUNNEL By Axel T. Mattson OOGJ .r ON LOAN FROM TYP F?moria1 Aeronautical Laboratory Langley Field, Va. NATIONAL ADVISORY CO M MITTEE FOR AERONAUTICS LAN;LEY AFI' ' IJ rI c AL LA) LANGLEY FIELD, H AMPTON, VIR:rLA CLASSIFICATION CANCELLED ,r'pc in F APOVE A -. AUTHORITY R.L.DhYDN L-5-53 USTS FOR PUBUcATloI SHOULD BE ADDREssp" FoU.0WS: C}1ANGEDJ 1O7

NATiONAL ADV5osy COMMITTEE FOP *512 II sTRE r, N. W. WAS*49WTON 2E, 0. NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS WASHINGTON

St. 23. 191.6 CFIDENTIL NACA RH No. L6H28 CONFIDEGI

NATIONAL ADVISORY COMMITTEE FOR. 'MER 04UT IC S

RESEARCH M0RA14TD10 (./. p.I, INVSTiGATI0N OF DIVE BRAKES AND A DIVE-RECOVERY FIP'

ON A HIGH-ASPECT-RATIO WING IN THE

LANGLEY S-FOOT HIGH-SPEED TUNNEL By Axel T. Mattson

SUMMARY

The results of tests made to determine the aero-. dynamic characteristics of a solid brake, a slotted brake, and a dive-recovery flap mounted on a.high-aspect -ratio wing at high Mach numbers are presented. The data were obtainedin the Langle y 8-f0ot high-speed tunnel for cor- rected Mach numbers un to 0.9L0. 'The results have 'been analyzed with regard to the suitability of dive-control devices for a proposed high-speed air p lane in limiting the airplane terminal Mach number by the use of dive brakes and in achieving favorable dive-recovery charac- teristics by the use of a dive-recovery flap. The analysis of the results indicated that the slotted brake would limit the proposed airplane terminal Mach number to values below 0.880 for altitudes up to 5,000 feet and a wing loading of 80 pounds per square foot and the dive-recovery flap would produce trim changes required for controlled pull-outs at 25,000 feet for a Mach number range from 0.800 to 0.900. Basic changes in snanwiSe loading are presented to aid in the evaluation of the wing strength requirements.

INI1R ODUCT I ON

The investigation presented was undertaken because of the adverse compressibility phenomena that affect the longitudinal stability and control of high-speed airplanes and because of the vital need for safety in dives at high Mach numbers. The investigation was conducted in the Langley 8-foot high-speed tunnel in order to obtain data for use in evaluating the aerodynamic characteristics of dive-control devices at high Mach numbers. The aerody- namic characte 'qeuipped with a

C 0UF IDE NT I A L 2 CONFIDENTIAL NACA RM No L6H28c

solid brake, a slotted brake, and a dive-recovery flap for application as dive-control devices are presented. The dive brakes were investigated primarily to determine their application as speed-limiting devices at high Mach numbers. Some evaluation of their fli ght characteristics was also made. The dive-recovery flap was Investigated to dete'mine its effectiveness in producing a controlled pull-out from a high-speed dive.

The high-speed aerodynamic characteristics, which include the lift, span loading, pitching moment, drag, and wake widths for the high-aspect-ratio wing have been analyzed and are presented in reference 1. The data presented herein include the same aerodynamic character- istics as affected by a solid brake, a slotted brake, and a dive-recovery flap. Also included in the present inves- tigatiori are increments in average downwash resulting. from the addition of a dive-recovery flap. Data for Mach numbers up to O.9)O are presented.

SYMBOLS AND COEFFIC IENTS

a speed of sound in undisturbed stream, feet per second

b span of model, feet (.15)

c section chord of model, feet

C 1 mean aerodynamic chord (0.7 ft)

M Mach number in undisturbed stream (V/a)

P static pressure in undisturbed stream, pounds per O0 square foot

p local static pressure at a point on airfoil section, pounds per square foot P0 P pressure coefficient ,- \ q/

q undisturbed-stream dynamic pressure, pounds per square foot (.1 oil 2)

C 0NF IDE NT I A L NACA RM No. L6F128c C0I'DENTIAL 3

S area of complete model, square feet (1.10)' V velocity in undisturbed stream, feet per second x distance along chord from leading edge of section, feet

Y distance along semispan from wing center line, feet a angle of attack, degrees

P mass density in undisturbed stream, slugs per cubic foot estimated termtnal Mach number

R Reynolds number

AE increment in downwash angle, degrees

i t angle of incidence of stabilizer, degrees H loss of total pressure between free stream and station in wake w/s airplane wing loading, pounds per square foot

Subscripts: cr critical

L lower surface of airfoil section U upper surface of airfoil section The coefficients are defined as follows: e n section normal-force coefficient

e n = fi ( PI, - PU L

C ONF IDE I]IAL

CONFIDEI'TIAL NACA RM No. L6H23

Cm section pitching-moment coefficient about 25_percent& chord station

1 P Cm = - PT. IX - dx J0 C, wing normal-force coefficient

2 ccdy = .- J0

wing pitching-moment coefficient about 2. 5-percerit- 0 chord station C - 1'' 2 inn -- codym 0/ 1+ Sc' •Jo

wins profile-drag coefficient

cc- section loading

APPARATUS AND ETH aDS

The Langley 3-foot high-speed tunnel, in which the tests were conducted, is of the single-return closed- throat type. The Mach number at the throat is continu- ously controllable. The air-stream turbulence in the tunnel is small but slifthtly higher than in free air. A complete description of the basic model and tunnel setup is r-iven in reference 1 with the corrections for model constriction, wake constriction, lift-vortex interference, and model inaccuracies. All the data presented are for corrected Mach nuniers unless otherwise specified.

The high-aspect-ratio wing used in this investigation is the same wing used in the tests of reference 1 with dive-control devices added. The wing has an NACA 65-210 section, an aspect ratio of 9.0, a taper ratio of 2.5d.0, no sweepback, twist, or dihedral (reference 1). The effective span of the model is 37,8 inches, the root chord is 6 inches, and the tip chord is 2.4 inches. The model wing included 20 static-pressure orifices placed

C 0NF IDE NT IAL NACA RM No. L6R28c CONFIDENTIAL 5 at each of eight stations along the span. The spanwise locations of the stations in percent of the semispan are 11, 20, 30, 43, 56, 64, 80, and 95 . The four inboard stations were placed on the left half of the wing and the four outboard stations were placed on the right half of the wing. The wing and the locations of the orifices are shown in figure 1. In the investigation two types of dive brake having the same over-all dimensions and the same location were mounted on the model wing. (See fig. 2.) The solid brake consisted of a solid plate having a span approxi- mately 20 percent of the wing sernispan and a height the same as the v*ing thickness (10 percent of the local wing chord). The face of the slotted brake includes three banks of four slots inclined 10 0 to the air stream. Each slot was approximately 25 percent of the brake span and 10 percent of the brake height. (See fig. 3.) The dIve brakes were located on both the upper and lower surface of the model wing at 50 percent of the wing chord. The dive brakes were located spanwise, symmetrical about the wing center line; the center of the upper- surface brake was located at approximately 20 percent of the wing semispan; and the center of the lower-surface brake was located at a p proximately 12. percent of the wing semispan. (See fig. Lb .) The unsymmetrical spanwise location of the brakes was used because of a possible installation in which the brakes could be applied by rotating in a plane normal to the chord line about a common support. The inboard location was necessary to insure adequate model strength. The dive-recovery flap was located on the lower sur- face of the model wing. (See fig. 5.) The model flap was of the wedge type with a chord 10 percent of the local wing chord and a span 20 percent of the wing semi- span. (See fig. 6.) The leading edge of the flap was located at 30 percent of the wing chord and the center of the flap was approximately 20 percent of the wing semis- pan from the model center line. The flap was deflected 300 with reference to the wing chord line. Pressure orifices were installed on the dive-recovery flap at corresponding wing stations A, B, and C. (See fig. 6.) No force measurements were made during the tests. All the lift and moment data were obtained from

CONFIDENTIAL 6 CONFIDENTIAL NACA FM No. L6:28c

integrations of the pressure-distribution measurements, The pressures at the 164 orifices were recorded simul - taneously by photographing a multiple tube manometer. The pressure-distribution measurements were made for a corrected Mach number range from 0,400.to 0.90. The tunnel choked at an uncorrected Mach number of 0.945. The brake configurations were tested at angles of attack of 00, 2 0 , Lo, 70, and 100 for Mach numbers of o.Loo and 0.600 and at angles of attack of 0 0 , 2° and 40 for the complete Mach number range. The dive- recovery flaps were tested at angles of attack of -20, 0°, 2 0 , O, and 70 for the complete Mach number range.

The wake-survey drag measurements were made inde- pendently of the pressure measurements. These measure- ments weremade for uncorrected Mach numbers of O,Loo, Moo, 0.725, 0.760, 0.800, 0.850, and 0.850. As pointed out in reference 1, the tunnel choked with the wake-survey rake-supp.'rt strut installed when the uncor- rected Mach number in the region of the model was 0.880. A calibration of the tunnel with the wake-survey strut in place indicated that the results are essentially unaffected by choking effects up to the chokinig Mach number of the wake-survey strut. (See reference 1.) Wake-survey measurements were made at six spanwise sta- tions 1.4 root chords behind the 25-percent-chord line of the wing. These stations were 11, 20, 26.5, 30.2, 3.3, and 42.9 percent of the wing semispan from the model-support plate. Wake-survey measurements for out- board stations of 60, So, 95, and 102 percent semispan were obtained from the data of reference 1. These meas- urements were made for the complete Mach number and angle- of-attack ranges. In order to obtain wake-width measure- ments at a typical tail location, wake surveys were made at a station 2.82 root chords behind the 25-percent-chord line of the wing, 5 inches from the support plate or 26.4 percent wing semispan.

Downwash measurements were made at a station 2,87 root chords behind the 25-percent-chord line of the wing for the wing with dive-recovery flap and wing alone and 26,L percent of the wing semispan from the support plate. The measurements were made by a small calibrated yaw head. The yaw-head measurements were made at a point located 70 percent of the chord above the center line of the tunnel. The measurements were made for a Mach number range from 0. 400 to 0.880 at angles of attack of 0 0 , 20, 40, and 70,

CONFIDENTIAL NACA RM No. L6H28c CONFIDENTIAL 7

The variation of model Reynolds number in the Langley 8-foot high-speed tunnel, based on the mean aerodynamic chord of the model wing (0.37 ft), as ' a function of test Mach number is presented, in figure 7 . The variation of level-flight airplane lift coefficient with Mach number for wing loadings of 60 and 80 pounds per square foot are presented in figure 8. In order to aid in locating the various figures in this report, a figure index has been prepared giving the figure number, title, Mach 'number range, and angle-of- attack range. (See table I.)

RSULTS. Pressure Distribution and Span Loadings

Basic pressure distributions.- The basic chordwise pressure distributions for the solid brake, slotted brake, and dive-recovery flap for spanwise stations of 11, 20, 30, and L3 percent semispan for Mach numbers of approxi- mately 0. 00, 0.800, and 0.930 and angles of attack of 00, 40, and 100 are presented in figures 9 to 20. These data illustrate the section chordwise loadings that occur with the application of'the dive-control devices and the appearance of compressibility effects. The chordwise pressure-distribution measurements for all spanwise sta- tions have been integrated to determine the corresponding section normal-force and moment coefficients. The span- wise variations in section loading are presented in figures 21 to 23. When compared with the results of reference 1 figures 21 to 23 illustrate the changes in aerodynamic forces along the span of the model caused by the application of the dive-control devices for varia- tions in angle of attack and Mach number. Normal force and pitching _moments.- The spanwise variations in normal-force and pitching-moment coeffi- cients were integrated along the wing sernispan for each angle of attack and test Mach number to obtain total wing normal-force coefficients and total pitching-moment coef- ficients. The total wing pitching-moment coefficients are based on the mean aerodynamic chord of the wing (0.37 ft) and were calculated about a point located at 25 percent of the chord. These normal-force and pitching-moment coeffi- cients are presented for the solid brake, slotted brake,

CONFIDENTIAL 8 CONFIDENTIAL NACA RM No. L6H28c and dive-recovery flap at each angle of attack plotted against Mach number in figures 2L, 25, and 26, respec- tively. The results for the model wing alone are pre- sented in reference 1. The effects of the solid brake, slotted brake and dive-recovery flap on the static longitudinal stability chaacteristics of the wing are shown in figure 27. The variation of dC/da and dCm/da with Mach number for an airplane level-flight wing loading of 60 pounds per square foot for the wing alono and for the wing with solid brake, slotted brake, and dive-recovery flap are presented in figures 28 and 29. Ir. order to illustrate further the effects. of the dive-control devices on the aerodynamic characteristics of the wing, the increments in normal-force and moment coefficients due to the" solid brake, slotted brake, and diva-recovery flap are pre- sented against Mach number in figures 30 to 32.

Drag.

Section wake surveys at five spanwise stations have been reduced to total profile-drag coefficients by use of expressions including the effects of compressibility similar to those presented in reference 1. These results are presented for each dive-control configuration in figures 33 to 35 for constant angles of attack through a range of Maci-i number from o.Loo to 0.65. The wake- survey measurements taken at stations 11.0 and 20.2 per- cent.of the wing semispan - that is, directly behind the brakes - are subject to errors introduced by difficulty in obtaining measurements of the extreme total-pressure losses in the wake. The errors are believed to be a small p art of the total drag and become unimportant when the total drag is used to estimate the terminal Mach num- ber. The data and calculations presented are believed to he closely representative of the drag increases that can be expected.

From the drag results ohtaind an estimated varia- tionof airplane drag coefficient with Mach number for the solid and slotted brakes at wing loadings of 60 and 80 pounds per square foot are presented in figures 36 and 37, In these figures the variation of wing drag coefficient for 00 angle of attack obtained from results presented in reference 1 are presented as a check on the variation of estimated airplane drag coefficient with

CONFIDENTIAL NACA RM No. L61z28c CONFIDENTIAL 9

Mach number obtained from reference 2. The generalized drag curve obtained from unpublished, data was extrapolated up to a Mach number of 1.0 and the extrapolation are used only to illustrate the high terminal Mach numbers to be expected for this airplane with a wing loading of 60 and 80 pounds per square foot. The increments obtained for the solid and slotted brakes were added to the estimated airplane variation. From these data, the terminal Mach number estimated for the complete airplane with and with- out the solid and slotted brakes are presented in fig- ures 38 and 39 for wing loadings of 60 and 80 pounds per square foot for altitudes up to 35,000 feet.

Wake Profiles Vertical variations in total-pressure loss in the wake AH/q were measured at a station representing a typical horizontal-tail location and are presented for the wing with the dive-control devices for various angles of attack and Mach numbers in figures 40 to 42. This survey station is 2,82 root chords behind the 25-percent- chord line and 5 inches from the support plate (26.Lj-percent wing semispan), which represents a possible tail location. All the wake dimensions are given in terms of the wing.root chord. These results show the wake width in the region of the tail and the wake spread with increase in Mach number.

Downwash and Estimated Stability

Changes in downwash for model configurations with and without the dive-recovery flap were measured in a region representing a typical horizontal-tail location. These changes in downwash angles are representative of changes in tail loads resulting from the action of the dive-recovery flap. The variation in increments in down- wash with Mach number for constant angles of attack are presented in figure L3 An estimated qualitative comparison of elevator- fixed static longitudinal stability characteristics esti- mated for a proposed high-speed airplane with dive- recovery flaps is presented in figure 44.

CONFIDENT IAL 10 CONFIDENTIAL NACA RM No. L61322c

DISCUSSION Dive-Brake Characteristics

The primary purpose of a dive brake is to limit the terminal Mach number of a high-speed airplane without creating a serious stability problem, trim change, and flight hazard such as buffeting excited by low-frequency wake fluctuations. The slotted brake was designed pri- marily to decrease, with a minimum loss in braking effec- tiveness, some of the serious effects usually associated with a solid brake such as lift loss, moment changes, and buffeting tendencies.. A large increase in wing profile drag resulted from the application of the dive-control brakes. The results presented illustrate the drag coefficints that can be obtained when air brakes are introduced in flow fields induced by the wing and corresponding wing- interference drag (See figs. 33 and 34.) The drag variation with Mach number for both the solid and slotted brakes indicated rapid increases up to the highest test Mach number. The profile-drag coefficient for the slotted brake at a Mach number of 0.4 and an angle of attack of 00 is approximately 30 percent lower than the corresponding value obtained for the solid brake. The profiledrag coefficient for the slotted brake, however, increases morerapidly with Mach number so that for a Mach number of 0.85 the slotted brake presents a reduction in profile drag coefficient of approximately 8 percent lower than that of the total solid-brake drag. This greater rate in drag increase with Mach number can be accounted for by a choking condition for the slotted brake; that is, a con- stant mass flow through the brake slots, which with increasing Mach number will effectively cause the slotted brake to approach the characteristics of the solid brake. The variation of drag for the wing with dive brakes with angle of attack indicates a decrease in profile-drag coefficient with increase in angle of attack. This varia- tion is expected since at 0 0 angle of attack the brakes are normal to the free-stream air-flow direction, whereas with increases in wing angle of attack corresponding changes in brake angles occur. This change in brake angle may be visualized as a decrease in effective brake frontal area with increasing angles of attack.

CONFIDENTIAL NACA FM No. L6H23 CONFIDENTIAL 11

Estimated terminal Mach number,- The effectiveness of the solid and slotted dive brakes in limiting the terminal Mach number of a high-speed airplane for wing loadings of 60 and 80 pounds per square foot has been estimated by use of a generalized dra g curve from refer- ence 2 and the drag results presented herein. (See figs. 36 to 39.) The estimated terminal Mach number for the airplane with a wing loading of 60 pounds per square foot at 25,000 feet altitude is approximatel y 0.950. With the application of the slotted brake the estimated terminal Mach number for these same conditions is 0.o95 and with the application of the solid brake, 0.675. At 35,000 feet, the slotted brake will limit the estimated terminal Mach number to 0.805 and the solid brake will limit the estimated terminal Mach number to approxi- mately 0.795. For a wing loading of 30 pounds per square foot and 35,000 feet altitude, which represents the most severe condition considered, the slotted brake will limit the estimated terminal Mach number to 0.630 and the solid brake will limit the estimated terminal Mach number to 0.870.

Wake profiles.- The wake widths for the dive brakes at a station aximatjnc the probable tail location for varicus angles of attack and Mach numbers are presented in figures Lo and Ll. The solid brake produces a broad erratic wake that slightly increases in width with Mach number. The width of the wake ranges from approximately 70 percent of the wing root chord below the wing chord line to approximately 70 percent above the wing chord line for 00 angle of attack. At an angle of attack of L° and a Mach number of 0.$83, the wake boundary for the solid brake has increased to 1 wing chord above the wing chord line. The slotted brake provides a decrease in wake widths of a pproximately 20 percent compared with the solid brake. The wake boundaries for the slotted brake do not seem to be as erratic as those indicated for the solid brake, especially for the lower-surface brake, which suggests a possible reduction in magnitude of the wake vorticity. At an angle of attack of L° and a Mach number of 0.3bi, the wake boundary for the wing with the slotted brake is approximately i0 percent of the wing chord above the wing chord line as com pared with 1 chord length for the solid brake. The wake of the brakes if located in front of the horizontal tail will envelop the tail for a recommended tail location of 70 percent above the wing chord (references 1 and 3).

CONFIDENT IAL 12 CONFIDENTIAL NACA r?..M To. L61128c

Normal force and_Pitchin g moents.-m The variation of normal-force and pitching-moment coefficients, with Mach number for the solid and slotted brakes are presented in figures 2 L.1 and 25, respectively. The variation of normal-force coefficient with Mach number for the solid brake below the force break for angles of attack of 09 and 20 is small; however, a more ra p id increase is indi- cated for the larger angles of attack. At the higher Mach numbers, that is from 0.300 to 0.900, there is a decrease in normal-force coefficient and, consequently, an increase in angle for zero lift. For the slotted brake the variation of normal-force coefficlent with Mach number indicates a premature break in the normal-force- coefficient variation occurring at a Mach number of proximatel-y 0.600 for angles of attack of 0 0 and 20. (Sea fig. 25.) This premature force break is not con- sidered serious and examination of the soanwise section loadings indicates shifts that at the hib.er i.'iach numbers move inboard towarcl, the brake. (See fig. 22 ) At the higher Mach numbers, this shift in span loading increases the section loading at a station approximately 50 percent of the wing semis pan. It may be reasoned that, for a Macb. number range from 0,600 to 0.760, a choking condi- tion is occurring for the flow through the slotted brake, which demands a readjustment of the flow over the wing. Peyond. a Mach number of approximately 0,.e 0 0, the usual decrease in normal-force coefficient occurs; however, the force break is not so severe as that encountered for the wing alone in reference 1. (See figs. 2L and 25.) The data also indicate an increase in normal force beyond a Mach number of 0,900 for angles of attack of 0 0 and 20. The variation of pitching-moment coefficients about the wing 25-percent chord with Mach number for both the solid and slotted brakes indicates no unusual characteristics and both types of brake seemed to improve the wing-alone variations.

In order to evaluate more completely the effects of the dive brakes with regard to changes in . wing character- istics, the •varation of increments in normal force and pItching moments produced by the dive brakes with Mach number is considered. These data are presented In fig- ures 30 and 31. The results indicate a loss in normal- force coefficient for both the solid and slotted brakes; however, a greater loss is indicated for the solid brake.

Both the solid and slotted brake produce positive increments in wing pitching-moment coefficients. The

C 0NP IDE NT I A L NACA RM No. L6H28c CONFIDENTIAL 13

slotted brake, however, produces a slightly greater increment than the solid brake.

Stabilitv.- The static longitudinal stability char- acteristics for the wing with dive brakes compared with those for the wing alone are presented in figure 27. In general, the wing alone (reference 1) tends to become more stable with increases in Mach nuiifoer. For example, at a Mach number of0J00 for a normal-force-coefficient range from 0.1 to 0.2j. the wing alone is neutrally stable, becoming slightly unstable at the larger normal-force coefficients. As the ach number is increased, the wing alone becomes increasingly stable with large increases in stabilit y occurring at a Each number of 0.900. The application of the solid and slotted brakes increases to a small degree the stability of the wing except at high Mach numbers. The wing stability variations for both the solid and slotted brakes are displaced to a desirable positive or pull-out p itchng-moment coefficient range, which may be indicative of a positive trim change.

Application to complete air p lane.- The wing static longitudinal stability characteristics as defined by normal-force and moment coefficients for the slotted brake seem to be, in general, satisfactory with regard to application at high Mach numbers, and the effective- ness of the slotted brake in limiting the airplane ter- minal Mach number compares favorably with that of the solid brake.

It should be noted here that the inboard location of the dive brakes investigated was determined by model strength. Unpublished downwash measurement and fluctua- tions made behind the brakes at a tail location recom- mended in references 1 and 3 (70 percent above the wing chord line) indicated erratic flow produced by the dive brakes. Wake envelopment of the horizontal tail in the recommended location would be encountered. (See section entitled Wake profile".) These results indicate the undesirability of a brake location in front of the hori- zontal tail. An outboard location of the dive brakes is recommended. It is believed that an outboard location of the dive brakes would insure proper functioning of the horizontal tail and increase the trim lift coefficient for the airplane but an investigation of a specific air- plane configuration will be necessary to determine these effects.

CONFIDENTIAL l)i CONFIDENTIAL MACA RM No .. L6H28

Dive -Re cover y-Flap Characteristics

Normal force and pitching moments.- The variation of normal-force coefficient for the dive-recovery flap with Mach number indicates a rapid increase in normal- force coefficient up to a Mach nunber of 0.800 for 00 angle of attack. The normal-force coefficients decreased rapidly with further increases in Mach numbers The pitching-moment coefficient for constant angles of attack have small or nearly constant varIations for a Mach number rang e 0.11-00from to approximately 0.650. From a Mach number of 0.650 to 0.8O an increase in posi- tive or pull-out momentis indicated; however, at Mach numbers from o3oo to 0.925 a slight decrease occurs. At the highest Mach. number the pitching moments for all angles of attack assume a positive moment coefficient. (See fig. 26.)

Increments in normal-force and moment coefficients,- The dive-recovery flap, the fundamental purposes of which are to increase wing lift and to produce favorable pull- out moment changes at high Mach numbers, indicates an increase in normaL-force and positive moment coefficient through most of the test Mach number range for angles of attack of 00 9 20, and Lo. For anrrles of attack of 00, 20, and 140 , the normal-force coefficient increment for the dive-recovery flap increases rapidly with Mach number. For an angle of attack of 00 at a Mach number of 0.cIOC) the increment in normal-force coefficient produced by the dive-recovery flap is approximately 0.3. For the highest test angle of attack - that is, 70 - a loss in normal-force coefficient occurs for all Mach numbers. (See fig. 52.)

The basic chordwise pressure distributions and sec- tion loadings illustrate the variations discussed. (See figs. 17 to 20 and 23.) In p7,eneral, the dive-recovery. flap with its forward-chord location (30 percent of wing chord) increases the local pressures over the forward part of the lower surface of the airfoil with separated flow developed behind the flap and a corresponding decrease in pressures over the upper surface of the air- foil. At low Mach numbers this cnorctwise pressure dis- tribution does not represent an increase in over-all lift mainly because of the separated region induced behind the flap, which tends to nullify the increased hit experienced over the forward part of the. airfoil. The small increment in normal force that occurs for a Macli number of 0J400 can be seen in figure 32,

CONFIDENTIAL NACA RN I\o. L6H2c CCFIDENTIAL 15

With increases in Mach number, corresponding increases in local velocities distribute the' chordwise pressures so that rapid increases in lift occur over the forward part of the airfoil with the effect-.of the sepa- rated flow behind the flap tending to remain the same. At a Mach number of aproxiinately 0.929, however, the rearward Shift in upper-surface pressures seems to nullify some of the important increase in lift produced by the dive-recovery flap; a loss in flap effectiveness above a Mach number of 0. 0 00 is thus indicated. (See figs. 17 and 20.) The changes in wing moment produced by the dive- recovery flap is apparent from these corresponding changes in chordv.;ise loading. The increased loading for the for- ward part of the airfoil, which increases with increase in Mach number until tI-:e rearward shift in upper-surface loading occurs, will increase the pull-out moment for most conventional airplane center-of-gravity locations. The effects of the dive-recovery flaps on the variation in sp.anwise section loadings have been coiiipared with the wing alone for angles of attack 'of 00 and L°. (See fig. 2,.) For a Mach number of 0 .LU0 and an angle of attack of Lo the dive-recovery flap decreases the section loading; however, a slirht increase in section loading occurs outboard of the flap. Viith increase in Mach num- ber the dive-recovery flap for constant angle of attack up 'to approximately 0 produces increases in section loading over most of the span.

Stability. - The W:Lng with the dive -recovery flap for a Mach number range up to 0.850 follows closely the wing-alone static longitudinal stability variation, the curve being displaced to a positivO moment direction. (See fig. 27.) For Mach numbers of 0.900 and 0.92`3, how- ever, the wing and dive-recovery flap become slightly unstable with some decrease in the positive moment. At the higher normal-force coefficients and Mach numbers, a large increase in stability occurs. The effects of the stability changes for the dive-recovery flap will be discussed frther by use of downwash measurements and unpublished tail characteristic s to estimate over-all airplane static longitudinal stability char acterist ics at high LIa"chnumbers.

Dow,-Y,.---:ash.- The dive-recovery 'flap produces slight decrease in downwash, frQrn a Mach number of O.L00 to approximately 0,750. At a Mach number of 0.00 to a Mach number of 0.890 for angles of attack of 0, 20, and a rand increase in downwash occurs. (See fig. 43,) For CONFIDENTIAL

16 CONFIDENTIAL MACA RM No. L6H28c

00 angle of attack at a Mach number of 0.300 the downwash increment produced is essentially 00; however, at a Mach number of 0.890 the downwash increment has increased to approximately L°. The large increase in downwash du to the dive-recovery flap will produce a large pull-out moment for this range of Mach number. At an angle of attack of 7 0 .the dive-recovery flap does not contribute to an increased downash but shows a decrease in downwash of approximately 10 at a MaOh number of 0.850. (See fig. L3.) These downwash measurements were made at a station correstonding to 25.8 percent of the semispan. Examination of spanwise variation in section normal force for the dive-recovery flap compared with the wing alone f2) showsthat the variation in section loading at t.3-percent station with Mach number agrees quali- tatively with the measured increment in downwash.

Application to complete air p lane.- The general aero- dynamic merits of the dive-recovery f1a with regard to wing alone have been discussed; however, further analysis is needed to evaluate more com p letely the effectiveness of the dive-recovery flap in producing desirable airplane pull-out characteristics. By use of the wing results of reference 1, downwash measurements, unpublished tail char- acteristics, and the results presented herein, the longi- tudinal stability and trim changes were calculated for the airplane with and without the dive-recovery flap for Mach numbers of 0.800, 0.325, 0.850, and. 0.890. (See fig. 44.) These calculations are not considered quanti- tatively exact; however, they, illustrate the over-all elevator-fixed static longitudinal stability character- istics and trim changes that may be encountered with the application of the dve-recovery flap. These calculations indicate that a rather large decrease in trim lift coef- ficient will occur for the airplane without the dive- recovery flap for a Mach number increase from 0.300 to 0.890. The application of the dive-recovery flap, however, produces an increase in trim with elevator fixed for a Mach number of 0.300 and a similar but slightly greater increase in trim lift coefficient for a Mach num- ber of 0.890. These calculations represent a stick-fixed trim change that, for a w.ng loading of 60 vounds per square foot at .25,000 feet at Mach numbers of 0.600 and 0.890, will correspond to aoproximately a 3.6g pull- out; however, at lower altitudes the ncr:.cal acceleration will be somewhat greater. (See figs. 3 and )4I.) From the results and calculations of the present paper the dive-recovery, flap investigated may be concluded to

CONFIDENTIAL

NACA RM No. L6H.28c C0NFI DE, 'TIAL 17

provide satisfactory pull-outs for an airplane configura- tion in the range of Mach number from 0.800 to 0.900.

Wake profiles.- The dive-recovery flap produces a wake approximately 1 chord in width displaced below the wing chord line. At a Mach number of 0.d87 and. an angle of attack of L°, the Clap upper-wake boundary is approxi- mately 50 percent of the wing chord above the chord line; however, at this point a wake loss (H/q of approx. o.oL) was measured that is believed to be introduced by an upper-surface shock disturbance. This disturbance is apparent for all angles of attack and, for an angle of attack of 1i° at a Mach number of o827, extends to iLo per- cent of the wing chord above the wing chord. (See fig. 42.) The amplitude and fluctuations ifl the wake produced by this indicated shock disturbance should not be serious and is similar to that analyzed in reference 2,which may Five an indication of the structural stiffness needed for the fuselage and tail design.

Dr a.- The variation of drag coefficient with Mach number for the dive-recovery flap indicates a large increase in wing profile drag from a Mach number of 0.750 to the highest Mach number tested. At a Mach number of 0.850 for an angle of attack of L°, the dive --recovery flap re presents an increase of approximately 59 percent of the wing-alone drag. (See reference 1 and fig. 5.)

C OMCLTJDI ITC REMARKS

The present investigation of similar solid and slotted dive brakes mounted on a model of a high-aspect- ratio wing has shown that the slotted brakes are slightly less effective than the solid brakes in limiting the ter- minal Mach number of a proposed high-speed bomber. The slotted brake was estimated to limit an airplane terminal Mach number to values below 0.880 for altitudes up to 5,000 feet and wing loadings u to 80 }Dounds per square foot The slotted brake decreased slightly the serious effects usually associated With a solid brake such as lift loss, wing moment changes, and wake fluctuations. It is recommended that the slotted brake be located out- board on the wing to prevent wake envelopment of the horizontal tail.

CONFIDENTIAL

!S] CONFIDENTIAL NACA RN No. L6H28c

A dive-recovery flap of the same relative dimensions and location investigated herein installed on a high- s p eed airplane should produce trim changes required for pull-outs at 25,000feet for e. Mach number range from 0.300 to 0.900.

Langley Memorial Aeronautical Laboratory National Advisory Committee for Aeronautics Langley Field Va.

REFERENCES

1. Whitcomb, Richard T . Investigation of the Character- istics of a High-Aspect-Ratio Wing in the Langley 8-Foot High-Speed Tunnel. NACA EM No, L6HaBa, 19)46. 2. Bielatp Ralph P.: A Simple Method for Estimating Terminal Velocity Including Effect of Compressi- bility on Drag. MACA ACF No. L5G31, 19)45. 3. Fern, Antonio: Preliminary Investigation of Down- wash Fluctuations of a High-Aspect-Patio Wing in . Lan ley 8Foot High-Speed Tunnel. MACA EM No. L6H28b, 19)4b.

CONFIDENTIAL NACA RM No. L6H28c CONFIDENTIAL 19

OF TABLE I.- INDEX OF FIGURES

Range 7ieurel Title - Angle of Mach number attack (deg) Model details 1 Pressure-orifice loca- tions on 65-210 wing ------2 Solid and slotted dive brakes mounted on wing in Langley 8-foot high-speed tunnel

3 Slotted dive brakes------I

L. Solid and slotted dive brakes at 50-percent- chord location on 65-210 wing ------I 5 Dive-recovery flap mounted on wing

6 IT ocation of 300 dive-recovery flap at 30-percent-chord location on lower surface of 65-210 wing

Reynolds number

7 Variation of test Reynolds number with Mach number 10.400 to 0875 Level-flight lift- Variation of level- flight lift coef- ficient with Mach number o.Loo to O.900

CONFIDE NT TAL 20 COIDE'NTIAL NACA RM No. L6H26

TABLE I.- IND1X OF FIGURES - Continued

Range Figure I Title - Angleof Mach number attack (deg)

pressure distribution

9 Pressure distribution at the il-percent- semispan station for wing with solid brake 0.600 to 0.925 0 to 10.0

10 !Pressure dstrihution at the 20-percent- semispan station for wing with solid brake

11 Pressure distribution at the 30-percent- semispan station for wing with solid brake

12 Pressure distribution I at the 43- percent- semispan station for1 wing with solid brake

13 Pressure distribution at the 11-percent- semi-span station for wing with slotted brake -

1)4. Pressure distribution at the 20- percent- semis pan station for wing with slotted brake 15 Pressure distribution at the 30-percent- semispan station for wing with slotted brake

C CNF IDE T TAL

NACA EM No. L6H28'c COP'IDENTIAL 21

TABLE I.- INDEX. OF FIGURES - Continued

Range

Angle o-L' Fgure Title Mach number attack

Pressure distribution --Continued 16 Pressure distribution at the L3-percent- semispan station for wing with slotted brake o.Cüo to 0.925 0 to 10.0 17 Pressure distribution at 'the 11-percent semispan station for wing with diwe-. recovery flap 0 to ).O i8 Pressure distribution at the 20-percent- sernisran station for wing with dive recovery flap

19 Pressure distribution at the 30-per3ent- semispan station for wing with dive- recovery flap

20 Pressure distribution at the 43- oercent - semis-can station for wing with dive- recovery flap

S p anwise section loading

Spanwise variation in 21 I section loading for wing with solid ' brake . 0.LO0 to 0.9 1 0 0 to 10.0

CONIIDEi'JTAL

22 CONFIDENTIAL N A C A RM 7.o. L6H28e

TABLE I.- INDEX OF FIGURES - Continued

Range Angle of Figure! Title Mach number attack (deg)

Spanwise section loading - Continued

22 Spanwise variation in section loading for wing with slotted !0. 4 brake 00 to 0.937 0 to 10.0 23 Spanwise variation in section loading for wing with dive- recovery flap 10.4100 to 0.929 -2.0 to 7.0 Normal force and pitching moments

24 lVariation of normal- force and pitchIng- I moment coefficients with Mach number 1' or wing with solid brake 0.400 to 0.925 0 to 10.0 25 Variation of normal- force and pitching-

moment coefficients I with Mach number for wing with slotted brake

26 Variation of normal force and pitching- moment coefficients with Mach number for wing with dive- recovery flap 1-2.0 to 7.0 Stability 27 IStability characteris-I tics for various model configuration000 to_0.925 .2.0 to 10

C 0NF IDE NT I A L NACA RM No. L6H28c CONFIDENTIAL

TABLE I.- INDEX OF FIGURES - Continued

Range Ficure Title Angle of Mach number attack ____ (deg) Normal-force and pitching-moment slopes 28 ISlope of normal-force- coefficient curve for a loading of 60 nounds per square foot at an altitude of 35,000 feet 0.400 to 0.925 29 Slope of pitching- moment-coefficient curve for a loading of 60 pounds per squth'e foot at an altitude of 35,000 feet n

Normal-force and p it ching-moment increments

r jatjon of mere- m ental normal-force and pitching-moment coefficients with Mach number for

solid brake I OiLOO to 0. 0 25 1 0 to 10.0 31 Variation of mere:- mental normal-force and pitching-moment coefficients with Mach number for slotted brake 32 Variation of incre- mental normal-force and pitching-moment coefficients with Mach number for - dive-recovery flap 0 to 7.0

COIEiDENTIAL 2)4 CONFIDENTIAL T'TACA RM No. L628c

TABLE I.- INDEX (Y' FIGURES - Continued

Title Angle of Figure Mach number attack (deg)

Profile, drag - -

33 lVariation of profile- I drag coefficient I With Mach number for wing with solid brake OiiOO to 0.865 0 to 10.0 Variation of profile- drag coefficient with Mach number for wing with slotted brake 0.1400 to O.

35 Variation of profile I drag coefficient with Mach number for wing with dive- recovery flap o.L..00 to 0.650 1 -2 0 to 4.0 Terminal Mach number

36 sE timated variation of airplane drag coef- ficient with Mach -rum"- for terminal Mach number predic- tion for airplane wing loading of 60 pounds per square foot at various altitudes 0.LLOO to 0.900 ------

37 Estimated variation of airplane drag coef- ficient with Mach number for terminal Mach number predic- tion for airplane wing loading of 80 pounds per square1 foot at various altitudes

CONFIDEI'TTIAL

MACS,, Ri'!. No. L6H28c COMFI _DE NTIAL 2

TABLE I. - INDEX OF FIGURES - Continued

Range Figure Title - Angle of Mach number attack (deg)

Terminal Mach number - Continued

38 Estimated variation of teriinal Mach nuiiber With altitude for win loading, of 60 pounds per square foot O.00 to 0.90 39 Estimated variation of. terrinal Mach nu'rber -vvi`F' h altitude for wLng loading of So Dounds per square foot Wake profiles

ILO Vlake profiles for seera1 Mach numbers 2,82 root chords behind the 25-percent- chord line for wing with solid brake 0.70.0 to 0.8831 0 to 4. 0 Wake profiles for several Mach numbers 2.82 root chords behind the 25-percent- chord line for wing with slotted brake 1 0 - 7 /0 0 to 0.8841 42. ake profiles for several Mach numbers 2.82 root chords behind the 25-percent- chord line for wing with dive-recovery flap 10,76o to 0.887

C 0NFIDE MI1 IAL 26 CONFIDENTIAL MACA RM No. L6H28c!

TABLE I.-. INDEX OF FIGURES - Concluded

Range - Angle of Figure Title Mach number attack (deg)

Downwash

Variation of change ifl downwash angle with Mach number for the dive-recovery flap measured 70 percent of chord above the center line 9.400 to 0.9001 0 to 7.0 Estimated comparative stability

Estimated comparison of elevator-fixed static longitudinal stability character- istics for complete airplane with dive- recovery flap 0.8 to 0.89

C ONF IDE NT IAL NACA RN No. L6H28c -Fig. 1

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CONFIDENTIAL

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Figure 2.- Solid and slotted dive brakes mounted on wing in Langley 8-foot high-speed tunnel.

CONFIDENTIAL

NACA RM No. L6H28c Fig. 3

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IFNACA RN No. L6H28c Fig. 5 CONFIDENTIAL II

Figure 5.- Dive-recovery flap mounted on wing.

CONFIDENTIAL NACA RN No. L6H28c Fig. 6

CONFIDENTIAL JF

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CONFIDENTIAL -

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CONFIDENTIAL 30-percent-5emIspan 5ta f/on

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CONFIDENTIAL , I•l_

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CONFIDENTIAL

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NACA RM No. L6H28c Fig. 13b

CONFIDENTIAL - MEMMEMMEMEMEMEMMEMEM -/ •u•uuuuui•••••u•uiu U••UUURRU•Ul•UIUUIU MEMMEMMMEMEMEMEMEMM •u•••••••••••••••••• •uuui•u•i•i•i•uuua• -I. ••uuuum•iuuumuuu•ui MMMMMMMMMMMMMMMMMMM Ill MMMMMMMMMMMMMMMMMMMM MMMMEMMMM MEHMMMMMMM mMMMMMMMMMMwmwwHmMMM MEMEIMEMMMERNMEMEM 911111114, _m MMMMMMEMMMWMMMMMMIMM •usiimiiuiiiiuii F ••ru•R••u••••

ON 0 NEENNEEMEN -, MEMO BOMMEMEMEME MENEM ivauuuiiuuu•••iimu I••U UUMEN SEMEME90 mom iIUIlUINI1III]1l I. • U .E1JI ]4IlL1 . . iuuiii EMMEMEM mg d Percent chord Figure 13.— Continued. Fig. 13c .NACA RM No. L6H28c sommommmmmmmmmmmmmCONFIDENTIAL -/. -su•u•uu•uuu•uumi•u•mmmmmmmmmmmmmmmmmmmm -I. •mu•uuuu•uiuuuuuuuu mmmmmmmmmmmmmmmmmmmmmmmmmmmmmmmmmmmmmmmmimmmmmmmmmmmmmmmmmmm -I. •uu•muiuuuuiuu••u•mmmmmmmmmmmmmmmmmmm -t •mmmmmmmmmmmmmmmmmmmmummmmmmmmommmmmmmm .p1 -•iuuuuum••uuaiiiu•uiummmmmmmmmmmmumm"HIMMM, 1 mmmmmmmmmmmmmmmmmmmmi•uu••auuiii•ui••ummmmmmmmmmmmmmmmmmms EMMMMMMMMMIAMMMMMMMMMIIUUUN1IUUUUI•UUR UIEIIUUI•IUUUIR Emmmmmmmmmmmmmmmmmmm •u•uau•uriuuu•auiu uiiuuuiii••iuuu•iu offimma-mmmmmmmmmmmmmm mmmmmmmmmmm m1mmmmmm mmmmmmmmmmmmmmmmmmmm • uii•-._. JUIII EMEMMMMMMMMMMM ______• •NI•II U...... U...... Percent cflar1 Figure 13.- Concluded. NACA RM No. L6H28c Fig. 14a

/ CONFIDENTIAL

-4

—d

—)

PEIiEUUU —d •1U•UUUUUUI •• RMIuINNIMINIMMIMINIMEMINIMMINIMISIM

—d iui•u•uuuuuivauu ••u•••.•••u•...... m. 'p UUiiN•Uii•iUU• ••i•U Iu•iiuuuu••• iuuuiu• to SMINIMMIMIMIMEMINNUMMINIMMINIM U - MINIMMINIMMISMIummmommm 0 imuu•uuirriuuuiiu ME OMENS F lUIUU••IH1iIIl I•I MISIMMIRUNNIMERIMMINIMMIMMINIMIN iu••uuiuumiiM ME

rn••••u _ 'Cs d CONFIDENTIAL Ml r

------Percent cflor' MEMO -NE Fiure /4 Pi-e5.sure distribution at the 20-percent-

semfspan .s tat/on for w/h.q with s/otted broRe. Fig. 14b NACARM No. L6H28c

CONFIDENTIAL

iaauuuuu•uu••uu•iN •aiaIaIuauIuIl••••uu -i IMMMMMMMMMMMMMMMMMMMM MMMMMMMMMMMMMMMMMMMM -I. MMMMMMMMMMMMMMMMMM uuuuuusiiuu••iiii••u -I. MENMEMEMEMEMEMEMEME •uuu•iuuiuu•i•uu••u -I. uauu•uu•u•uamu••u !lUUUUUiiiUiUUU•i•ii iiiumuiiiu•i•iuuu•u•u !UUNUIII•IUUUUUU•UU• mMMMMMMMMMHmmMMMMMMlM mMMMMMMMMMmMm"HMMMMMl

MMMMMMMMMMMMMMMMMMME EMEMMMMMMMEMMMMMMMM - MMMMMMMMMMEMMMMMMMMM I••UUi•UNHU.UUII... • MMINUMMEIMMMINIMMMMMSE ,piMO MMMMMIEMNAM EMMMMM I..... iiuuuuusui•uuuauui•iu • i.uii.UU.aU.mIiii • _. s1I g I II :1 lIfl

4 - - -a - -a a # Percent chord Figure 14. Continued

il

/ NACA RM No. L6H28c Fig. 14c

- CONFIDENTIAL - - -I. •••.uuu•uu•iuui•a -1. uu•u•u•muu•i•i•••UU•R••U•RIUU•UUUII• -I ••u••uuuuiiuuuu•u•uu MMMMMSMMMMSMMMMMMM -I. MMMMMSMMMMMMEMMMMMMM ...... -I. • mMMMMMMMMMmQmMMMMMMM •••••u••u•uuu•••m• • ui•u•uuu•uiau•&iumuiuiu EMMMMMSMOMMMUNREMS - uu•uu•uuu•nu•uu•ii EMMMMMMMMMMMMMMMMM MMMMMMMMMMMMMMMMMMMM !1IUIR•I•U••I•UU• iuiiu••i•miiimuu••u•u MMMMMMMEMMIMMMMMMMM im iii•i•ui•••i• • - ONE ON

MENEM - iui•u••u• - U...... U...... • • • a a -.. I...... 'ercent cfl 'arar. F?qure 14.— Concluded. Fig. 15a - NACA RM No. L6H28c

F--CONFIDENTIAL -2 ,p. on

-1

-.1. Now u•ii•••iiu•uuuuiuuu MMEMENEENNEEMMMMMMM -ì !iIiliUURiUU•U•U•li •uuiu••i•iuu•i•••m -I •lLI••U•UiU••UliiUUi uiiiuuu•iui•••iui•uu mmmvkmmmmmmmmmmmmmmmm iu••iimuuuim••u•uuu•ii ommmmmmmmmmomommmmmm rniu•••uiiwamuu• F MMEMMMMMUMMMEMMUMMON

U • MMMEMMMMMMMMMMMMMMMM INN mmmmmmm • •uuiui• Emmmmmmmmm -' •••u• iiiiuiuuu :UIIlU • iuuuuuu• a...... a..... - - E E I W i.u••••• ......

0

CONFIDENTIAL Percent- chori Figure 15.- Pre5.sure d/strIbf/on at the 30-percent- set-n/span station for wing with s/otted brake.

NACA RN No. L6H28c - Fig. 15b

CONFIDENTIAL-- MENNEEMENEENNEEMOSOM • uuuuuuuiu•uuiuu•ii -I. mmmmmmmmmmmmmmmmmmmm mmmmmmmmmmmmmmmmmmmm -I. mmmmmmmmmmmmmmmmmmmm mmmmmmmmmmmmmmmmmmmm -I. ••u••muu•••uu•iuu••i •uiuu•uu•auuiu•uu -I. mmmmmmmmmmmmmmmmmmmm

-t mmmmmmmmmmmmmmmomm

p1 ••uu•u•u•u•uuui•i•uui •u•uuaiuuuuuis mmmmmmmmmmmmummmmmm •UU•I•U••W1UMUI•l NEENNOMEMEMEMBESOME

SMEMNIMMENWIMMENEEMERE • iIiiUVLUIllIl•ii

/ loommomosl4d", No - it NONE iim.iumiu immommommommommomom No RON ON ON sommommim MEMO i•u•u•uuuuuu• . . I •Ul•ii M•Ui•iMENmoomm • a • a a ..# a ê Mo. 4 I-ercent cfld- Figure /5 •- Continued.

- Fig. 15c NACA RM No. L6H28c

-4

i•u•au•uuuuui•uiii•u —/

iauiusuuiu•u•i•iauuu —/ iuu•••aiuumui••ui••• aui•uuuuiuii•uuuauuu ••u••••••••••••••••• iuu•uui•••uu•uu•imuu —ì MMmMMmMMMMmMmmmMMMMM jRlUlUUUNllV1.Li.U..N —I Iu••uuuu•riu•iuuui !RIiUlIUIUlII••i• Ll mmMMMmmMMMEMmMM Ikimwm I•IU•IUUUUIAUUUU•k1IL • uu•••uiuuuiiviu•u•u• •NEEMENEEMEMMENEMEMME

4 jriuuiiuuu•guuu•a•u 11. iUMMMLIMMNSMMMMEMMMM iui•uuiuusuuua•• umuuuu mommommomm ENO •uuutiOEMMOME MOOSE-_. uiva - u••uisuui•uaa••uu iuuu•uuumiu•umsu ______. ______ui...ii a ui•i••i Percent chcv-d Figure 15.– Concluded.

NACA RN No. L6H28c Fig. 16a

CONFIDENTIAL

V •0/) IW7ILMIi NOUN[ u•iuuaiiuiMENi••ai •.ur....ui...... uii. MESA MlMMMMMmMmMMMmM EMMEMMMMMMMMMMM mMMmMMmmMMMMMMmMMMMM IiiUUUU•UUi••UUll p1 PiiBiii•iuu••uiuiit mMmMmuMMMMMMMmmMMMMM mammmmoommmmmmmmmmmm p i•uiui rn•uauu I•I•lI _••..... mv,d_dMMMMMmMmmNmmMm EM MEMMEM 111IMMEMBROMMI'Malms • •iililNIiI•iUUui

- MENOMONEE •s•••••• iii... MENEM - vii•uuu 1III•I•I•UaURUIUIUUI _11hh11hlT1U0 I • ,;' S

!'!!! • a .,0 Percent cfl Fqure 16 - Pressure distribution. at the 43-percent- sern/span station for wing with slotted broke. -

- Fig. 16b NACA RM No. L6H28c

- -. CONFIDENTIAL -27 • MEMMOMMEMEMEMEMEM -i MMMMMMMMMMMMMMMMMMMM MMMEMMMMMMMMMMMMIMMMM MMMMMMMMMMMMMMMMMM - uiuiu•i•i•uumiiuiuu• u••aiiuua•uiuu•u•uiu -I. m•uuu••u•uiuuumuiuu• -I. MMMMMMMMMMMMMMMMMMMM MMMMMMMMMMMMMMMMMM -t iURUU•iRlRiIliUlIIN• MMMMMMMMSMMMMMMMMMMM ••uuuu•iva••ua•u•mmuu I uiuuau•uuiuuui•m uauuuurnii•m•auu - 0 SEMEMSEEMMOMMMEMEME

qi uuiiaumiuuuuuuu •

iRIIUIURU•URUR•UR - MENEM mmmmm^ - OMEN OEM 0 VM=^

OMEN ON MEN ONE • e • a a a a a . i-erce,,r cn'- F/qUre 16.—Conhnued.

NACA RM No. L6H28c Fig. 16c

a.' CONFIDENTIAL

•umuuusiaiuuuu•u• MMEMEMMEMEMEMEMEMEME -I. smoommmmmmmmmmmmmmmm uu•i•uu••uuu•iuuui• -I. mmmmmmmmmmmmmmmmmmmm mmmmmmmmmmmmmmmmmm Emmmmmmmmmmmmmmmmmmm Emmmmmmmmmmmmmmmmmmm -I. iu•••au•u•••uuu•• j•••••••••••.•.....m -I. i a..a•••••uauua•aa mMMMMMMMMMMMMMM1P0-MH2m ommmmmmmmmmmmmmummmm •••uu•••••u•••'••a• U. mmmmmmmmmmmMwmmmmm•• ••i•uaa•aiariau•aa IF ammonammmmmmmmmammm • mrRAMMMMMMMMMMMMMMMM •i•iiaami••u•uaauau• mmmmmmmmmmmmmmmmmmm mmmmmmmm U.....- IUUUUUUUU UUUUUU IUUUUUUU 4 UUUUUIUU

Emmmmmmmmmmmmmmmmmmm d U " CS Immmmmm IC 'erce,,t cflorI CU.tom Fiq.ire 16.- Concluded.

Pig. 17a NACA RM No. L6H28c

CONFIDENTIAL dZC aaui•uuuu•luuuuu IUIIU•IUIU••UR•UU -i iauu•u•auuuu•uii••u RIMME MMMMMMMMMIMMSMM RIMSEEMOMMEMMMOMM •u•LUIUUIUU•I•••UU MMMMMMMMMMMMMMMMMMM _z < •iiuiuuauuuiuu•••um•i p1 i•uiaaau•iuiius. je UIUU•UUNlUUUU

8 rAMEMEMMESOMMMM W004141mmm -. In MEN my—s-MMEMEMMURINNOME NMMMMMUMMMMMMMMMMMEM OWN

uiauuiu• NIMMMEMME -9 MENNEN No •u•uiui• IIUIUIRIII RURIUIIU WE ME M M ma...... • lUau.. 2 U••UU 4 ma.... U...... III_UIIUIUUUII

o /0 20 30 40 50 60 70 35 505 /00 Percent chand Figure /7.— Pressure distribution at the I/-percent-

s e /7)/.sp n station for wing with dive -recovery flap. NACA RM No. L6H28c

CONFIDENTIAL

p4 • I

• qi

. - Figure /7 ContThueá. - •

Fig. 17c NACA RM No. L6H28c

CONFIDENTIAL

-d Uu..IUI..UNuIU.IIu.R -I u•auu••u••uiiuiuu

-d

-d EU•II••IU•III•UUUII

-d PNII••II•ImU•UIU•• OMMMMMMMMMSMMMMMMMMM EMOMEMMMISMEMMEmmumm MMEMMMMMMEMS mmmm^ MMMMMMMMMF.dmmmmimm F. !U.UU.mIU.UuI.U.R.NMMMMMOMMMMMMMMMEMMMM •uiiuu••riu•iuuiui•uMEPAPEIRIMMEMEMME IMMENSE iNVAIUI•I1I•IU•ISIIUI•Nri•mu•muuu••uu•iuiu • •uii•ui SEEM MEMO • - . MEMMEMMEM *,mMMjMMMMM

I __ S WMMMMMMMIMMMIEMMMMMMMMwim m io 20 30 40 50 60 70 80 90 700 Percent chord Fiuré 17.-Concluded.

\

NACA RN No. L6H28c Fig. 18a

- CONFIDENTIAL

-,. iiiii•uiusi•iu•uiuu• umuau•uui•u•uuuuuu -I. EMMEMMMMMMMMMMMMMMMM iu•uu•uuu•u•uuuu -1 uauuuu•uiuu•ui•u•u• •ui vauum•..i• -I. ERNE MMMEMMMMMMEMEM NWMMMMMMEMMMMMMMMMMM -i • iiiuiuuu•uuuuiuuui•u

P, - uuiuu•uui•u••u•iiuu - iuuiu•uai•iuui•u•

uuuuuuii•iuuiu•s at • iva•auur•uu•u••u••i ••muuu•rnuuiuu••u•iu ExAMMEMUNMEMEMEMME •

ROMMEEMNE &W MEMEMEM - MMMMMMMMEMEMMMMMMME - IUU•SUT..... - ••iuuu•uu•uuuiumiiu • UI1.] IJ YNIN 'S qM o io 20 30 V43-23V 60 70 00 /00 Percent c/a'1 Figure 18.— Pressure distribution of the O-percenf- s emispan 5tclt/on for wing with dive-recovery flap.

- - Fig. 18b NACA RMNo. L6H28c

rniJrInENTIAL ãU•IuIIIIIUINuUMEMO MENMEMEMEMEMNIMMEMS —i - mmmmmmmmmmmmmmmmmmmmm - .....aI.u.I..I.IIuIuI -z iIUUlIU••U•IUUUURUUI —i MMMMMEMMMMMIMMMMMMMMM

—t •iRauI•u•lUSuUUUUIIUU —ì iSUiTlIUUUU p1 MMENSIMMENEMEMMISMENSIMMI iuhUIUUlliIUiilIUI MMMMMEMMMMMMOMMMMMMMM - mmmoommmmmmmmummmmmmm •au•siuiuiIu.0 I - - IUUUIUfiUUIUUlUISUI • 1UI$USUUII•UU•UUUUUUUU iUUUIUUI1UlUIUIU•UUl iuuuurni RuIauuIum IU•UINU -. u•u•uau - MEMO --MMMMMM - MENEM 000000 EMEMINEEMN • ______

'_, ,',# 'J'-. - Percent chord P7qLdre 18.- Cont/T)ud. -

NACA RM No. L6H28c Fig. 18c

—d - .u..••.•......

—d •ui••aum•iu•uiuuuuuu

uuiuuu•uuuimu•uuiuu

—d

- EMMNMMMWAMMMMMMMMMW EMMEMEMEMMMEMMEMEMEM NMMMEWMMMMMMMMEMMIM EEPSMIUMMEMEEMMEMMEEM Qi ••uu•uu•••uu••a•a H ft • rnuuuuauaiu•iiuiuimu

•uuia•i•uumauiu - uuu•••uuui••uau•a• •us•uuuu••uuuuauu•i • •IUIUIU••Ulli d _____ • mommommommmom iercent cflr1 Figure 18.-Concluded.

Fig. 19a Onfirinvu,riAl- - NACA RN No. L6H28c MEN IlUUUUl•uIUUU -I. IRIRIUUINUURU•SIIIU iIISUlNUIUUIIIUIU -I. - iiUlIlI•U••UUNIIIU IUUUIUUUUUUURIUUU•UIU PjIjUUUUUUUUIIUUNI -ì IIIiRUIUUIUIlUUUIUIUU - -I !I.URI.UUUUU•IUlIN MIMER MIMMMIMSMMIMIMIMIIMMIMIMMIMIMIMIMI WZWMIMIMMIMIIMIMIM MRSHIMMEMIMI iivaUIIRILUUU iISIUNUUUIUlISl UiIIIUtUIISIUUIl F • •p..IauumIaIUUIIaIR muIua,1Iu.aIm..auu ME - MINNIMMIMMIN ammommm MINIMUM... IMMENSE MINE MIMMIN URUUUUU*UUI•IUUIUU - HEINE IM NONE - SiUmUUIRRRRUulUa MEN.IIuINU.aI_t • iWki..1• • • •iuiuii ME MINNIE .• NOME Percent chord Figure 19. Pressure— distribution at the 30-percent- .sern/5pan station for wing with dive-recovery flap. Fig. 19c - NACA RN No. L6H28c -

-4 -i •iuuu•uu•uu•uuumu••u•iimuuuu•auuuii•iu -i -i•uiu•uuaiuuui••ii -i •••aiuiiuu•uuu••••a•uui•u•uuiiiau•uu. -I •••••••••••au•••••••• -I ••mu•uuuuuuuuu•uiuu imia•uiu••uiuuuia uuiu•uuuuuu•uauu•iu 1 •uiivauaiiu•uuuuuuu•u•u iu•uuium.•au•suuua••uuamiui• auui•iuu iiuuaui . MENOMONEEuuuiimu• - MENNEN....iUlIU• ••....u.....u...______i-ercenr cnorm- Figure -19.-Concluded - NACA RM No. L6H28c Fig. 19b

CONFIDENTIAL mmmMMMMMlMMMlMmMMMMMMm -I- NMIMMMMMMMMMMMMMMMMMIM

uu u u u u u i i • i u u u u • u u u

•. . u ' . . s ...... u . . .

IFW iu u a • i u u • • u • u • u u m u u

u• u a • u i • • u u u u u u u • •

iu • u i u u i • u • a i • • • u u MMMMMMMMMMMMMMMMMMMMI

iu u • i • i u u m • u u i a m m i a MENOMINEE

•

ri u u u u m u i u n u ni u u u • i i

ii u u u i u u • u • u mi • i i • •

i

ri • i i u u u r u u u i • u u u • • u •

• MEMO MEMEMEMMEMEME

uu • i s ii u i • a u • u u m u u

•i a u u n u • u i u • u u u u u

ru u u u u vi i i v a u m u u S -. EMEMEMEM I......

iuuuia

iiiuuu

• ______

/ercenr cflop' F?qure /9.- CorftintIed.

NACA RM No. L6H28c Fig. 20a

CONFIDENTIAL

-L

-I. !iiuuuuuu•uu•u••u•uu•u•

rniu•u••u•u•i••uu•iu••

-I.

-t iiuuumiuuuuuui•muuu ia•••uu•uu•uuu•uiiu P, !U•tUUIRiUI•Ul•IlUIl •R•U•UURI•U••II••U I ia•uu•umu••uuuau

MEMOWMlmmmmmmm ENOWL IN

Ii•.muiu•iu•uii

iiuumuuia• IMMINIMEMISMIN mMINIMMINIME - uuuuuiu•uuu•uuuu•u• MENNEN -

CONFIDENTIAL

• . /ercent chord Figure 20.— Pres5uredlstributlop, at the 43-percent- semispan sta f/on for wing with dive-rec o very flap. Fig. 20b - NACA RM No. L6H28c

- CONFIDENTIAL iauuu••uuuaium•i• Imuauuu•uu••iuiu••iuuu -/. •isiu••iiuu•uuu•uus - ••••u•••••••••••••••• -L muuuuuuauuuuuui•uuu -I. • P, i•uuuiuiuiivauuiu• :S - fill •iuuuiuuuui iiuiu•uuuuuu•iiuuiu •

MMEEMSEEM -. MEMEMEME IMMEMEMSM 4. I•I•IUI•• - EMMEMEMMEMMEMEMISMIM 111.11 •1•iU I..uI• immommomm. MEMEMMEMEMMISM 111111 __ --COMMITTEE FOR AERONAUTICS 10 111111MEMO 11111111 11O /0 20 30 40 50 60 70 80 90 /00 Percent chord qire O_ Con tinu,ed. Fig. 21a NACA RM No. L6H28c

CONFIDENTIAL •II•UUUUlUUlUUUR 1.8 (.7 iiuuuuuuiii•u•auu

0 .4 1 .1.)

.2

.1

0

-.1 0 FOE COMMITTEE FOR AERONAUTICS -2 NEI 0----70 U .iU 40 O 60 70 50 .90 /00 ."et-c en 1- .Sem Aspar (a)A4=0.400. F7cjure 2/ '—Spar7wise variation in sec-ion loading for wing with .50/id brake. NACA RM No. L6H28C Fig. 20c

CONFIDENTIAL -2

- MOOMMOMMMEMEMEMEM ME -I. iuui•u•uu•u•iuu•u••i ••umu•uuu•i•••u•uiiu• •u••••••••••••••••••• nommmmmmmmmmmmmmmmmmm -I. i••iu••uuuuu•uu•uuum uua•uuiuiu•uuiuuu•••• -i ommmmmmmmmmmmmmmmmmm 'p.1 mmmmmmommmmmonommumem mmmmmmmmmmosommmmm •muuuu•i•u•muu•u •au•iiaiuiu•uu•uu

•uuu•iai•uuiuuuuiu QI •vu•uuuu•••uuuuiu•u • MMENNOMMMMEMEMEMOMMEN su•uuuuuuu••i••uuuu riauu•ui••uu•uuuui•u I•UIUR••RUUUUU• •uiiu••uuuuuiuuiu• MENEM NONE - uuuui•uaiuiu•uuuuuuu • _____ MEN MEN •.. i'erce,,i- cfld Figure 20.-Concluded.

NACA RN No. L6H28c Fig. 21b

CONFIDENTIAL I e.I F. (._

...... la .8 EMMEMEMMUMEMSEMEMEN C.) •7 ILIIIIurd•IIiu•I•i••

C.) MNMMMMMMMv,J:MMMmMlMMMmM itimmmmmmmmmmmmusummmmm b5 MMMMMmMMmMMMMMMmmMmMM 0 lL1hiUi•WiiUUi•ULIi•l

-4- •uuismiuiuuu U Q) MEMMMMEMEr MMMEMMIMMMM Lo

.2 mmmmmmm mpopsoommmommm .1 Mmm mmmmMMMmMMMMMmwM lMmMMMmmMMMMMmMMMMMmM 0• MMwmMmmmMMMM -I MMMMEMMMMMMMMM

NATIONAL ADVISORY _____ -MMMMMlMMMM-_ a ercen/- .5,erni..5pan (b)A.4=O.OcO. Figure '21.-Continued.

Fig. 21c NACA RM No. L6H28

CONFIDENTIAL .8

.5 1J) N. (deg 0•

-.4

cz:.

0

L) / Q) (1)

0

-.1

NATIONAL ADVISORY COMMITTEE FCQ AIRONAUTICS -2 CONFIDENTIAL I I 0 10 20' 30 40 50 60 70 80 90 /00 Percent semi'.5 oar? (c) 1\4 0.700. Figure .21.-Continued. NACA RM No. L6H28c Fig. 21d

CONFIDENTIAL

(.) mi•IRu•uU•Ij•IiNONE •.m.uu.u1a....um...i

t.J

-

-.1

MMMEMMMEN NONE

cu IL' cDL' &Li IC/U Percent semispon (d)Iv?= a 808. Figure 21.- Con/-ihuea.

Ffg. 21e NACA RM No. L6H28c

CONFIDENTIAL

.7 ....u...isa•iumu•MENu •iii•uuuuui•m•iui•uu• 0.5 mmmmmmmmommmmmmmmmmmmm lU•NUIUSUNiUI•URU•iU• MMMMMMMMMmPMmmmMmMmMmmM mmmmmmmmmmmmmmmmm MMEMMMMMMMMMMMEMN-A 0 •uuuiuuuuuuiu () EMEMENEMNEEMEEMENNEENE Q). UlU•l•UA•UU FA NOWEEMM sommommom AMEMBEEMEENEM NNEWMEMENEEMMEMENEMEEM -I

NATIONAL ADVISORY mommmummommom COMMITTEE FOR R

!!!!I1U L JL/ ',-(J CJ(J .L.J L/ L1 IU!U!!L)L'UU (LI! C)L' / L/L r'ercen/ miipan (e) Iv! =0864. - Figure 21.-Con/-in ued. NACA RMN0. L6H28c Fig. 21f

CONFIDENTIAL .7 MEEMMEME MEMMEMMEMMOMMEMEMEME N ME MMMMMmMmmMMMMMMlMMM U 0 MMMMMMMMMMMMMM ME M SMEMEMMMEN MEN M NONE

.2 MEN MEMEM"EMEMMEENME

-k U Q) 'I MEN MMENEEMENEENEOMM (I) ENEMEMEM abb III' 0

-J MENEM a MEENEWEN ME mom

&M mmipm Ri 1 I C) -L) JU 4U :DO O'O 70 80. 90 . tOO Percent .sern ispar, - - (f)A-1=0.904. Coat/n Qed. NACA RM No. L6H28c

CONFIDENTIAL 'UI 'UI .U.' --- I') 0(1 eg,) C.) U

0

1 C.) .1) IUUU•UIUUUUUUI1UUUU lU'UU!iJ1UUU•U•VdUUUUtUU

l.UUUIUUUUUU.UUr;IU IUELUUUUUUUIUUUUUUUUU Percent sern Aspcn 0. 194-0. FYgure /.-ConcIudec/. •NACA RM No. L6H28c Fig. 22a

CONFIDENTIAL

I

.ii-z- -iii-.iii I-

U

0

.2

.4- Q)U .1

0

-.1• IUlU•UI•••lII;•U MOEN -2 MEMEMEMEMEM FEE 01 /0 20 30 40 SO 60 70 <30 Pcrcen*(c4M=Q.400. Semi,5pc7n

Figure Sparwie variation in section loading for wing with 1ot1ed bra/ce. -

C

/

Fig. 2ab NACA RM No. L6H28c

•uu•uu•uu•iuu•iii urn •muauum••uui•u

.8

.7

0 •aiiuiiiiii•u•iiu• . .4

lz •i•vu•uu•rnuiu 0• U (f)'

C U•UUlI•UUIU.;UU -I •••u•uiuiiuuuiuuu

i NATIONAL ADVISORY C6NFIIIEN^IAL •••uuuau•uiuuuuui•COMMITTEE FOR AERONAUTICS* C) /0 dO iO 40 ou 00 ((.1 CU .0 IOU Perceni- .semii'pn (b) A-1=O.00. Figure P-Con/-/nued.•

NACA RM No. L6H28c - Fig. 22c

CONFIDENTIAL

.8

(-) .5 U

iz 0• -k V.2

•1

C

••••uuuu•uu•i•iii MEMMUMMEMEMEME ME ONE - - Percen/- 5emi5pan (c),\4=O. TOO. F/gure &-Cop t I r u e d.

Fig. 22d - NATCA RM No. L6H28c

ONFIDENTIAL H .9

.8

.7

U. () —us

WIr)g O/OflG -(refererce I) i i.io.aoo

wwwo 0 'us' U EMN

0

-I 55•U•555U5.Ft;!V.. •uuuuuuuiuuuuuusssussu

Percent .sem/span (d)tA=0.8OO. F1 qure22.-00n7'inued. NACA RM No. L6H28c Fig. 22e

CONFIDENTIAL .8 UIU•l•UUlU•UI•IIUUij .7 I•UUUIU•IUllIUIIliii MMMMMMMMMMMMMMMMMMMM U 111111mll Immmmmmmm

rzh

1Z33 mmmmmmmam 0 mommo MEN ommmmum , w MEMEMEMMMEMEMEM ENDS om MERNEMAN MEMNIMMEMBEE

MEN tUiUUUIIUUauii

mmmmmmmmmmmmmwk wlir Ail .4, MEN lUUUl.lIUUIUIU...i...

fPerccnf Se rn/span (0)tvl=O.öO2. Figure 22-Cori*,nued.

/ S Fig. 22f NACA RM No. L6H28c

CONFIDENTIAL mommmommommommomommomm .7 IIal•IllIIIUUIUIIii I•lNi•IU•i•liiiuMmMMmmmMMMMmMmMMMMMmMM -Q U U mMMMMMMMMpwAdw2mMmMMMMMMM mmmmmmmmmmmmmmmmmmmmmm •IiIlINuuIul•u•IiU•uMMMMMMMMEMMMMMMMMMMMMm 0- MMMMmwmMMMM mmmmmmmmmmmmmmmmmmommm .0 o MMMMMEMMMMMMMMEMERNMEm NEEMEMEMEMEMEMEM MORON 0 MMEMENEEMMEMMEMEMMEEMM ••iupi••u•iuu•uuuu• iuiuu•uiiu•.0

-2 i!il•URRiiiUiU•IU Percent erv/span (f') 1\4 =0.9L2. Figure a2.-Cont,nued.

NACA RM No. L6H28c Fig. 22g

CONFIDENTIAL --- OMEN 0 .7

.6 MmMM M ME ME 0 mmmmmmm ME ME M : MOMMIMMIN SEE ME MORMINMEM ME 0 - 0 -.' MEIN HERE NNE= M U mqbqqhmmmm ME MEN am== M (1) MEN MEN 0 MOMEEM -I

-2 mommoommomm No ME MMEME No MEN ME c, IL) C) (1 .40 o W 70 80 .90 /00 icercen/ .sem/pern (g)A4=0.9

Fiqure 22.-Concluded 0

0

Fig. 23a NACA RM No. L6H28c

CONFIDENTIAL 1.0 uu•u•iiu ii•ii••uuu

.7 •uiuuiuurn•i•uuu•i

C)

() mummommommommomm

mmommoommmmmommmommmmm

mommommmmmmmanommummmm MMmmmMMMMMMMMMmmMMmWMM mommummmmmmumommmmewmm MMMMMMEMEMMENEEMEMENUM••uiuiui•uiiu MEMMEEMIM10101MEN1010101010: MOON -.1 ••uua•ui•uu•uu•uuuuu NATIONAL ADVISORY ME •muiiaiuum•aMEN uii o ic 20 .30 40 0 &) 70 Q 9U /00 Percer?* 5em ip an (a) Alv =0. 400. F/gore .- '.5pan wise var/aHon in sec-Hon load/hg for wing with dive-,recovery flap.

NACA RM No. L6H28c Fig. 23b

CONFIDENTIAL.

IL

.9 U1U$IUUUUUU••UU•! .8 ••u•••uuiuiu••uu•au .7

.6

(O5 MEEMMEMEMMEMMEN For,

0

MMEMME

/ ONNEMOSEM U MEMO ••iMENIk111III111ONE u MENlRIUUUUUruiUE L'I•i•iI••llUiUIUU• No U /U U . 0(1 /Ci - iU . .S(J IC/C) Percent semI.spar?. (b) f\/f=O.600. Figure 23.- Continued. Fi'g. 23c NACA RN No. L6H28c

CONFIDENTIAL cc (deg "I

Ic'

.9

4

U. (j

.4.

c.3.

1 0

.1

0 NATIONAL ADVISORY I COMMITTEE FOl AERONAUTICS U/ye-recovery-flap location—--H 20 130 JO 150 60 Perceni- serni.spari' (c) A-4=0. 700. Figure 3.- Continued,

NACA RM No. L6H28c Fig. 23d

CONFIDENTIAL IC

.9 ERNMENNEW mummulmoffilMqmmMMMmmMmMMk MENNIM MEN MmMMMmmmMMM MMMMMEMOMMMMEMMMMMMMMm •1) MMMMMMMMM Mmmmlml^et ,c. () EMMEIMMMMOMME11malummuLF .5 •••••iiuuiuuuuum ••uu•muuuuurn•ni•i• •uu•••uuuu•u•uuiiiuuMMMMMMMMMEMENEEMENEWMEM .2 ••••••••••••u•••• (•7 •u••uuiuu••uiva •um•uuuuuuu••••uiiiiu ••UU•IUUI•I••I•iIM 0 ••uuu•iuuiiusiiauaiuui MEN MEN •ui 11I1II1I111uiuNATIONAL ADVISORY

INNOMENEENEEMEMMMEMMENE G /0 za JO 40 C) 00 70 öU 9O /00 /Dercep71 aem/span (d) f\4=0.802. Figure 3.-Con/-,nued.

I • Fig. 23e NACA RM No. L6H28c

CONFIDENTIAL /0 UiUUiiI•I•IIII•UUIIU•ia••••m•••aINluIlIuI iuuuuu•IIuuuIlmu•U•airimuu•••i•iuiiiu H mmmElm mm=mmmmmmmmmmm

J.O 0 mmrAmmkiammmmmmommrommmummommmmmmmmmmmm mmmmmmmmmmmmmomommmkImmmmmmmmmmmmqmmmmmmmmmmom 0 mmmmmmmmmmillommommmmalmm H iUUlI5UUiIU 0 •UiiU1IlU•USiUUU NKNMMMH- "007m • /7 20 30 40 50 60 70 60 9O • /00 Percent 5eml,5pan (e) A-I 0.853. • NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS • Figure 2,3.-Con*/nued.

NACARM No. L6H28c . Fig. 23f

CONflDENTIAL . IC

.9

.7 (deg)_ ------4 ------7.------. .4 Z\ 2 Wl'n_q alone A7=0.9Q7 oc =00 7 ------

.1 1111111 ThY iiIIZIII

6

-.1

- ---- Dive-recOvery-f/op /Oca/O_ - - - '.111 1 . 11 I C) IC) ±0 .1U 40 U ou (U CiU U IC/U Percent sern/spcJr? CONFIDENTIAL NATIONAL ADVISORY (*') 1\4 = 0.904. COMMITTEE FOR AERONAUTICS. Figure 23.- ConlLir?ued. Fig. 23g • NACA RM No.. L6H28c

CONFIDENTIAL .6 •uiiuuuui•umi•u•uuii• .5

iumu•uuuauuuu•ui••u .Q.4 0

0 MEN MMMMMMMIEMMMM .1 0

Lo uiuuuuiu•uuui•um•m UIUI•IIU•I••IiEIU•U UUll••IUIUNRIlUUmom •uuuuu•••u•i•uuuiui C) /0 zo JO 40 JO OC) 10 dO C) IOU

Pe rc e nt Sern15,ban (9)1\4=0.929. Figure a3.-Concluded. -

NACA RM No. L6H28c Fig. 24

CONFIDENTIAL cj

1. IZ

0

IZ

0

!UUUUUUUUUUUUUUUUUUUIUI-U.-. UUUUUUUUUUUUUUUUUI Qz UUUURURUUUUUUUUIUUU•UUI UUUUUUUUUUUUUUUUUUUUUUI IZ UUUUUU•UIUUU••U•UUUUI 0 MMIMMMEMMIMEMEMEMEM I C 0 MEMEMEMINNEEMEMEN mil

0 •U•UUU l•lNUUUUUUUUUUU •u•uup•uuuuuu•u Mu mmmuumuiuuiuuiiiumuu WKWIS ID1211 .!iUIu UUUUUi

NATIONAL ADVISORY A/lac,") number, COMMITTEE FOR AERONAUTICS riqure 24.-Var/at/on of norrnci/- force and pitching - momeniL coeff/cieni with Mach number for wing with 'so/id bra,ke.

Fig. 25 NACA RM No. L6H28c

14

'S.

0- ci 45. SZ

'S

EMNIEEMMMM MEMMOMME MENNEN MEEMENEMMEN ••iuui••uuu••i•uiuiu •M••UUUUU••UI•N••U••

45. ••uuuumiiauui•uuu•u•u uuuu••ui•um•u••uau•u 'U 9' MENNOMMEENEEMEMEM 0 M cj EMEMMENEENNOMMEM W.

L. 0 •••UIi i.-hUiU -:1 •EU•UUUUUIWII ••5 U .i . . -4 .b .7 ' .& 1.0 - NATIONAL ADVISORY A-lad') numb er) '4 COMMITTEE FOR AERONAUTICS Figure 25-Variation of normal - force and p/t-chihq- moment coefficients with tvlach number for wing with .s/otteo' brake. - I

Fig. 27a,b NACA RM No. L6H28c

iii iii•iuuusauu•u• MMMMMMEMMMMME -Wing alone •uau•u•amuuuuu - JJLI IU!•UUUl•UUUl•.______Mmmmmmmmmmmmm-- —aJj 11 - ui•iuu•u••iu EMMMMMMMMMMMEMEMEME •uuuuuuiumuu••uiuu• ••uuuuuiiiuiuu••u•• mmmoommummmmommmmmmmmm MMMMMMMMMOMMMEMMOMMEN .0 •iiu•uiuumm•uuuuuuiu MEMEMEMEMENNEEMEMENNO .c) MOMEMEMENEEMEMEMOMEME

0 ci

c)

tv'orrna/-force coefficient, CN Figure 27.-Stability characteristics for various model configurations. NACA RM No. L6H28c Fig. 26

j

'CONFIDENTIAL

ILI

4-- -

(.7

4.6

0) .0 .5 '4- .4- 0i

0)

0 4-.

L 0

•OEMll• MEN ul••!uI1m!• A-la c/i n uc b e p /\4 NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS Fiqure26-Variafion of normal - force and pitchinq- morneni- coefficienl-s with A4cich number for wing with- dive - recovery flap. NACA RN No. .L6H28c Fig. 27c,d

M14

-. 'K 0 c) •1

Q)

I: -. iuu•uuuu•s•u•i rnu• .1 MEMO= 5 Normal force-cOefficient, Cv Figure 7-Cohtinued.

Fig. 27e,f NACA RM No. L6H28c

CONFIDENTIAL I Wing alone Wing with' ---5o/id brake -----iiotfed brake - - __--Q/ve-rec.overyflap -----=------

E 0

0)- /

0 0 (e)i\4=O.9OO. 4-'

0)

Norma/-force coerr,c,ent,cN Figure 27-Concluded.

NACA RN No.' L6H28c Fig. 28

L

U >.*-. 4—' q)

INt

1. LI)

cJ4 --- 0 'tJo OD

N LI

Qi

\

-j 5. 4 I—. 2 ('I I :::3 L&I N s4.J a • •'•-c L&. -5' z U0

'4J N (C) -5 'Si -5-5- 5':• 0 5Sj t 9210J - /OULIOU — AJ170 40 9d0/ ç'

Fig. 29 NACA RM No. L6H28c

Q-) MEMEMEME m ....__.w. lUllI1IUFU UUUITAIIUUILIIU .- 0 111111 IUUUWI Q-)0Q) OMEN 0 MEMEMEN MEMEMEN MEMEMEN ¼ mmmmmmom Emmmmmm -I......

-j mmmmommmmm Cr) OO)

I- z MMMMEMMMMM QOo 0 mmmmommmmm L&. 1 z , 0 mmmmommmmm •.- U EMMMEMMMMMI

0 0 0 I.

1ap°Pc AJflD LJOUJ_iLI/L-/,L!d jo ado

NACA RMNo. L6H28c Fig. 30

CONFIDENTIAL'

0

•0 LICN

MA

.1 m mmmmmmmmmm A Cm I 0W.0 m llmmmmmmmmwx-mm NATIONAL ADVISORY mmmmmmmmm 'COMMITTEE FOR AERONAUTICS

Immmi .6 .7 .8 .9 /0 A4ach nombe1'-, tvl Fig elre 30-VarithVon of /ricrernentci/ normal - • force and pitching -moment coefficients with A4qch nUmber for 50/id 'brake. Fig. 31 NACA RM No. L6H28c

CONFIDENT I AL .2 - -- (deg)

N0• -LC - - -7 - -4--

.2

.1 ,6 Cm m MOMEMEEMMMM

0 MMMMEMMEEMEME 0

CONFIDENTIAL L Emmmmmmmmm .4 .ö .7 .. .9 /0 Mach n urn be r, A4 Figure 3 h-Var/cit/on of incremental normal- force and pitching -moment coefficients with Mach, 'number for 51otted brake.

NACA RM No. L6H28c Fig. 32

• CONFIDENTIAL A

.2 IN MEMMEM0 MEOEMME LICN NEEMMINP-.dmmmmm - 0 M mummEMENEEMEM 0 HOMMEMMOMEMEM IN MMMMMMMmMmMMM MEMEMEEMMEMEMEM • -4 MMMEMEMMEMMEMMM ENEEMMEMEMMEMEMI MMEMMEMMOMMEMEM EMEEMMEMEMEMEME MMEMMENIMEMEMMIMM EMMEMEMMENEM

0 M EMEMEMEMMEMP—M NATIONAL ADVISORY E=C=MM MW ME im . .4 .) .b •.7 •a.. .9 10 •Mach number, /'vl Figure 3-Varia f/on of incremental normal- force and pitching' -moment coefficients with Iviach number for dive-recovery flop. Fig. 33 NACA RM NO. L6H28c

EMEMEMEMEN 0 MEMEMEMEME 3)

ill Ii _ _ MIII_LU_- mm - _ 'LQ) -J 4 Cr) I-z Id 0 LL 0z • Co . U C) .0 U

0 u/ ,./JJ & 0 c - / IJOJJ

NACA RM No. L6H28c Fig. 34

MENOMONEE 0 MEMMEMMEME M MEMENEEMEM M, IZ MEMEREMEME__ _ : 0 Ii LI MEMEMENEEM MEMEMEMEMEME .' -. MEMMEEMMEEME : Q) I MEMMEMEMESSM iiai MEMEMEMEMEME lz MEMEMEMMERME -0 MEMMEM M •-.- L 4 MEMMEME r) LQ) I- 2 w MEMEEMMEM a Li. MEMMEMEMEM 'st jz 0z U Ti L cr (\.

0c7 7 C aiIjjo 5DIp -//JO1d Fig. 35 NACA RMNO. L6H28c

.< MEEMEMEMMMMEMEMIN IME I MMUMMEMMIMIN EMMMOMMIN IME •Q) OMMMENUMI ME mommmmummmm MMMmMMmmMMMMEMIMMMMMEMIN _liii_ MEEMEMMINIMMMI MMMmme Em

NACA RM No. L6H28c Fig. 36

0 M No - - 4c:: a•u•uuai •ivauu•iuuwi•uuu•um•

•u•••uuS MEMEMMEMMEMM I..I • S - k1UP'4UUURIUlS

1-.

()

0

a •iuu• iuuiu•au• UuI•UUIRUlUUi1-ulNATIONAL CONFIDENT - IAL IN- • j .4 .D .b / .d . I.Q 1.1 l\AaCh number, A4

F-igure 36;- Esfirna fed variation of airplane drag' coefficient with 1V7o'ch number for terminal Iviach • number prediction for airplane wing loading of 60 pounds per square foot at various altitudes.

Fig. 37 - NACA RM No. L6H28c

jjj—ñiAiMil'Iar9r1 a aaaa.aauUURa iiuaaaaauIIaIIaIua aaaaaaaaaaaIaIIaaU aaaaaaaaaUUIlaIaaIIIaaI .4. aaauaIaaaaaaaaaIa aaaaaauurlIIauaaaUaIaaa .4 ivaa.iaa.UI1uI1IIaUaIaaU ME No .3 iaONEaa,auIIaIaaailUuIa .amom.II.Uaaa,aIIaaaUa .3 aaaaaa.IL1aIaIUaaaIaIU -- uMENaaIaIaUaRUaIaIU - 00 POP. No ME aaauaIaUaUUIaaNo -L vz-- auaaa,amaIaaau .Q) . u MENUa C') uiau.aaaaaa MMEM .1 0 • aaaaa.aa.auaSIaIS C') .aaaaiaauaIa la L .uaaaaaia.aaUaI,Uaaa1a '- aauauu..aaaI UaUUIUUI a__r.,MEIUauITaaiUrdUlaIaaU 'aaaiaaaaaa.aauuaauIUua aaaaauaai-Ia ..aaa.RaUaaaflaa!!!aU.' Mach number, I\.4

Figure 37.-Estimated variation of airplane drag coefficient with 'Iviach number for terminal /vlac/) number predict/on for airplane wing loading of 80 pounds per square foot at various altitudes.

NACA RM No. L6H28c Fig. 38

_•_•••.•••• fl) Q)Q) Mmmmmmmmmm- m

MmmmmMmmmmlm

MmmmmmMmmmmM III1U•E1•_UIMmmmmmwmmmmmI mmIIUiISIMMI (\I k0mUm sm LIMMM - -J 4 ism Ellm. I— z mm mmm ILl l L mmm 0 2 0 mm 0 II—. ••UU••UU.• I

O) N k? 4'v "quinu /.o/,-v 1D1i/UJJL

Fig. 39 NACA RM No. L6H28c

0 L0

'I) qj U

0 mmmmmmmmmMi!m0 MEMEMEMEME a fl) •1U1111_uuui URRIULI UU.UR MMMMMEMMMMEM mmmmmmmmmmom MEMMMMMEMMEM mmmmmmommmom o m1_mmmmolmomm_• I%, m m uii ui. I -j 4 m M I— z Li m mUt1-.

L& % 4 m mm mm 1._ z 8 m mm Em U)

MMMMMMMMEM-ml a-. 0 aqnu Fig. 41a NACA RM No. L6H28c

tO

LA '-I umi. U ••u•uuiii•uuuu •.U.U.UUIUUU•UUU-U B

0 •iu•iu••uu•ii

- '1-

3Q) I..... L

IuIuu•u•IUUU•w#Uua MENOMONEE•miuu•u•aumriri• J Q)3 MEMISIMENSIMMINMINESIMENSIMMIN m,MMbills P ••uuuu•i•uu•iuuast Q) (h_I U )r-. U•UUU•UUUUUUUUU•UU .Q Q) 2)

-j 4 ••UUUUUUUI•UU•IUUft qj I- z Lii 0 U- -z 0 0 ••auuuuuuiuii•u

NACA RM No. L6H28c Fig. 40c

U >.- CO

>

Q

\ Qi

Q)

II

(J Omni INN MOMMEM

4 I- z 0

z 0 0 S1 S

NACA RM No. L6H28c Fig. 4-Oa

S - U

hZ

5-'

•uiuuaui•••uuu••ii () •siiuuuuuiui•uu•r O'3 uuu•i••••uu.iiuu.ri MMMMMMMMMMMMMwEmmw (J MOMMOMMEMNSHOMM NMI b. auiumiuuu•iu• MEEMEMENEEMMEMMOMM MEMMEMENUUMENMEMEM q) QI UUUII•NII1U•S••• 5Qb •uamuuiriiu•ui•u MEMEMMESSOMEMOsamm IZ •uuuiurriviu••uuu

••u•uuii••um••u•i I, * MENNOMENEMEMEMEMME U MEMMEMMEMEMMEMEMOM I ('j •-4 •uuuiuui•iu•u•uu• Q) •.-4 '4Q) .....u...... ij •u•uuiiu•uuuuiuii• MEMEMOMEMEMENN I qj -I 4 I- z MENNOMMOMEMOMEN I -L

U. MMEMEMOMMEMEMMEME S_.S (I 8 tuauumuuuuuu•uuu• I S MEMEMEMEMMONENOME QJ 14 Fig. 40b - NACA RM No. L6H28c

0 MINEEMEMEMEMEME ME EMEMMINEMMEMESIM ME iuuui•iuuuutuu •u•uuiiuui•uuuu mmmmmmmmmm mmmmmmmmmmmmmmmmmrA 1- iuu•u•uiuu•iuui iz EMMOMMESSIMMImmmmumm o ) EMMMMMMMOSEMEMMMMM EMEMMMEEMMMOMMMMM NEIMMINMEROMMOMEMMM MIENNINIMMMEHASOMMEINIM 0 -- •

MINNININEEMMMEImmmmmilm MINIMMEMNIIIIINEMEMEMENSM .•••• mumuiui H •uu•uimm ommom uti J. Q)- •uuiiu•uuuiuiuuii (J MMMMMEMMMMMMEMMMMEI Q) MMMNMMMMMMMEMMMM

••••i••....i...i.. 0

4 EMMEMENIMEMMEMEM M. I- Uz Q) I' \ L z 0 •RUUUS••I•IIU•L1 U NOMMOMMEMOMEMEME 0) c N O 0

NACA RM No. L6H28c Fig. 41b

Uta

> .ui...i....i.-II. '.1 MEMEMMM MMMMMMMMME 0 mmmmmmmmmmmmmmmmmiw r\c mmmmmmmesommmmmmmmMENOMMMMMMMENEEMMM U MEMEMMEMMMMMMMMMMM Q) U

mmmmmmusommmmmmmm 'Q) mmmmmmmmmmmmmommMOMMEMMMMMMMMMM Lq MMMMMMMMMMMMMMEMMILki - iiiuu•u•miur•uuu•uu••uiu••ri 0 Q-) •UUIUU•UlIUU•NIL Po mmmmmmmmmmmmmmmmmm ) D mmmmmmmmmmmmmmmmmmmmmmmmmmmmmmmmmmm Qj mmmmmmmmmmmmmmmmmm 'I- Q)

mmmmmmmmmmmmmmmmmm 0 cj -J 4 mmmmmmmmmmmmmmmm I— z lmmmmmmmmmmmmmmmmm Id 0 Q) I. mmmmmmmmmmmmmmmmm z 0 U Immmmmmmmmmmmmmmmm a q tb-. (0rrr .

Fig. 41c NA-CA RM No. L6H28c

••••ONE ui•ui u• ••iiiuuuuu•uuu. I. •••u••uuuu•u..u: No MlmmmMmIMMMMMMmmM ii ••IUUUNURU•.•U: 1 MINSIMMENNEEMEMEM No MINEEMEENMEMEENNIMME MINEEMMENEEMENNIMME MIMMEMENNEEMEENEE muiu•u•uuiuuu• •U•1••UU••UUURI •uuuvavaiii u•• U momommmmmoommommes 0 •••iiiiiivauuuumii•u

No MlMMmmmmIMMMMm 0 MENpis MlMmmmmmmmMMml •uuuu•uuimuiuimui () IMMMIMMEEMEEMENEM NMI 1:3 •i•uu•usuu••iuiuuii 0 , •...... " :3

0 ci -J IMMISIMMEMEMEEMENE Ml 4 ••ui•m•auiu•uuauu z Id 0 Q) •••ii•uu••uuiuuuu L z :3 8 LI- •uuiau•uuu••u•u •Q "4 0 NACA RM No. L6H28c Fig. 42a

mmmmmmmmmmmmmmm •uu•••uuuuuuiu ffl,

•um•u•muiiuiiuiu' MMMMMMMMMMMMMMMME mmmmmmmmmmmmmmmmmm mmmmmmmmmmmmmmmmu MMMMMMMMMMMMMMOMEM MMMMMMMMMMMMOMMMEM mmmmmmmmmummmmmmm mmmmmmmmmmm OEM mmmmmmmmmm lmmmk%kl •..I.u.i.u::' •...il mmmmmmmmmmmmmmmmmm •UU•UUIlUIUURUIRI

mmmmmmmmmmmmm 0

mmmmmmmmmmmmmmmmm

-

• mmmmmmmmmmmmmim

mmmmmmmmmmmmmmmmm

V

Fig. 42b NACA RN No. L6H28c

cj

U

0

-

_J - 0 co , - - —I- - Ii I- (J

------FT ----y- (3 ------> -.-•--- - Q) ------7-----

L

- o

pc

Jill PQ) - U cQF__ - -- -

• - '2 • ------c

0 c -I - - - - 4 I- z -- --- z Ii - z 0 - • Zi U - 8

N (U - 0 /HV

NACA RM No. L6H28c Fig. 42c

MOMEMEMMEMMOMMEMSE •m•uuiuu••u••si•m •a•uuiu•uuiu•uian MMOMMEMEMEMEMEM Emmmmmmmmmmmmm mm ••••••a•a••••u•• •u•uuu•iiu•mu mm mmmmmmm It ui'

mmmmmmmmammmmmm Ii •iuuuuuuuuiiiiiuuuu w.

•uuu•uuuuu u ua Q) U •uu•uuuuu• uuiu. L 2) •uuuuu•uu riuiau

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uuuuuiuiuu uuuuuu

mmmmmmmmmmmmvwmmMm -tI 0 •..uu.uiuu•u•u u U - MMMMMMMMMMMMMMMO ON U ••uauuui•uuu•uuu Qj mmmmmmmmmmmmmmmmui 0 ••uuuuuuuu•uuuuuui ••u•••u•••uu•u•••u 2) (3 b •uuuu•uuu•ui•uuuui Q) •uuuuuuuuuuuuu•u•• uuuuuuiuiu•uuuuuuu uiuuuuuuuuuuuuuu• -J 4 p. z mmmmmmmmmmmmmmmmm 2) U. mmmmmmmmmmmmmmmmm L. z - 8 u•uuuuuuauuuuuuu• uiuuuuuuuuuuuuuuuu 1;: 0) q (Q rr) 0 - - Fig. 43 - NACA RN No. L6H28c

'I) q) H U

0Was MMEMEMEEMM 0 ,Q) MMMMOMMMEMIN co- EMIMEME_____ '4j1 0 MEMEMEM mold-MOME MEMMEME MEMMMMEMMMMM MMMMMMMINEMMM JZ mmmmMMMMEMMM —J 4 mmmmmmmmmmm fly I- z mmmmmmmmmmm Lii 0 c) Li. r z 0 0 mmmommmomom1 MMMEMMMMEMMM, cJ

bap 7 'L,'$DM U440 LII ./ &9UJJXIJ

NACA RN No. L6H28c Fig. 44

CO Restriction/Classification ONE, Cancelled OMNI ' --MMS" IlIllIllII:

MEMMMMMMM_ MIM _ _ • III....• •• MMMMMMIIMMMMMEMM MMMMMMMRMMMM•._ APAP Gowi 407pl .ete MMEMMMMIRIVIS, recovedive 4. SEMMMEMMMuiv EMEMMMMEMMM_IMMMM I 1111IMMIMMMMMMEMMMMIiuiii•-•• -:2 MMEMMEMMMMM I MOMMMM_MMMOM_•MM 0 -.3 •••••-•• I_ •_ Restriction/ Classification a. Cancelled EMSI ^MMEMPPME 111110 Cff c_I. .'' ..O . 1.0 Norm/ - force coeff/c/et t1 Figure44.-Est/rncifec/ comparison of elevator-fixed sfdtic /onqitud,na/ stability c1-)arac1-er15t1cs for complete airplane = 00. with dive-recovery flap. /.