NASA TECHNICAL NOTE
STATIC LONGITUDINAL AERODYNAMIC CHARACTERISTICS OF A HAMMERHEAD-SHAPED LITTLE JOE I1 - LUNAR MODULE MODEL AT MACH 0.30 TO 1.20
by Arvo A. Lzloma and Charles D. Harris LHgZey Research Center Langley Station, Hampton, vu.
NATIONAL AERONAUTICS AND SPACE ADMINISTRATION e WASHINGTON, TECH LIBRARY KAFB, NM I11111111111 11111 Ill11 lllll lllll1IlU 1111 Ill 0330753 NASA TN D-4139
STATIC LONGITUDINAL AERODYNAMIC CHARACTERISTICS OF A
HAMMERHEAD-SHAPED LITTLE JOE II - LUNAR MODULE
MODEL AT MACH 0.30 TO 1.20
By Arvo A. Luoma and Charles D. Harris
Langley Research Center Langley Station, Hampton, Va.
NATIONAL AERONAUTICS AND SPACE ADMINISTRATION .. . ~ .. . For sale by the Clearinghouse for Federal Scientific and Technical Information Springfield, Virginia 22151 - CFSTl price $3.00 STATIC LONGITUDINAL AERODYNAMIC CHARACTERISTICS OF A HAMMERHEAD-SHAPED LITTLE JOE 11 - LUNAR MODULE MODEL AT MACH 0.30 TO 1.20
By Arvo A. Luoma and Charles D. Harris Langley Research Center
SUMMARY
An investigation of the static longitudinal aerodynamic characteristics of a 0.03-scale model of a proposed Little Joe 11 - lunar module suborbital space vehicle was made in the Langley %foot transonic pressure tunnel at Mach numbers from 0.30 to 1.20 and at angles of attack from approximately -12O to 12'. This configuration was of the hammerhead type, in that the diameter of the lunar module shroud was larger than the diameter of the Little Joe I1 launch vehicle. Three sizes of stabilizing fins, three nose shapes on the lunar module aerodynamic shroud, and two shroud skirts, which extended over the converging juncture between the shroud and the launch vehicle, were investigated. Flap-type trailing-edge controls on the stabilizing fins were uniformly deflected for pitch control. Control hinge moments were measured on the small fins. Limited transition-strip studies were made on one of the configurations. All configurations investigated had static longitudinal instability at all test condi- tions about a moment reference center 1.3 base diameters forward of the base. Severe model buffeting occurred at a Mach number of 0.90 in two narrow ranges of angles of attack near -7O and 7O for the configuration with a shroud nose consisting of a blunt- nosed cone of 30° half-angle and a frustum of a cone of 150 half-angle. Associated with the severe model buffeting were abrupt changes in the magnitudes of the aerodynamic forces and moments as the angle of attack was increased through the critical range. . The addition of transition strips extended the Mach number range in which model buffeting and abrupt changes in aerodynamic forces and moments occurred to a Mach number of 1.00. The additim of a boattailed shroud skirt to the smooth configuration substantially reduced the model buffeting and the extent of the abrupt changes in forces and moments at a Mach number of 0.90. The trailing-edge controls on all three sizes of stabilizing fins were effective for pitch control at all test conditions, although reductions in control effective- ness occurred at high control deflections and also at the highest angles of attack at Mach numbers of 0.95 and 1.00. INTRODUCTION The National Aeronautics and Space Administration at the Langley Research Center has investigated several configurations of the Little Joe II launch vehicle in combination with an aerodynamic shroud which will house the lunar module (LM) during the boost phase of suborbital flights for LM systems testing. The main objective of the investiga- tions was to determine the longitudinal stability and aerodynamic control characteristics of the Little Joe II - LM space vehicle and the effect of fin size and shape of the LM aerodynamic shroud on these characteristics. The attitude control system of the Little Joe II launch vehicle consists of a combination of aerodynamic control from trailing-edge flap-type controls on the four stabilizing fins and reaction control with fixed nozzles at the roots of the fins. The results of an investigation in the Langley 8-foot transonic pressure tunnel of the static longitudinal stability, longitudinal aerodynamic control, and trailing-edge con- trol hinge moments of a 0.03-scale model of a proposed Little Joe 11 - LM suborbital space vehicle are presented herein. Three sizes of stabilizing fins (with full-scale plan- form areas per fin of 4.65, 7.33, and 9.30 square meters), three nose shapes on the LM aerodynamic shroud, and two LM shroud skirts, which extended over the converging juncture between the shroud and the launch vehicle, were tested. The investigation was made at Mach numbers from 0.30 to 1.20, at angles of attack from approximately -12' to 12O, and with the trailing-edge controls uniformly deflected from Oo to -30° for pitch control. Control-surface hinge moments were investigated for deflections of 00 and -loo on one of the small fins. Test Reynolds numbers based on the diameter of the base of the model varied over the Mach number range from approximately 0.70 x 106 to 1.60 x lo6. The most promising of the three nose shapes on the LM aerodynamic shroud has been investigated with the intermediate and the large stabilizing fins at supersonic speeds also; these results are reported in reference 1. References 2 to 4 report the results of similar tests of the Little Joe II launch vehicle in combination with the Apollo spacecraft at Mach numbers from low subsonic to supersonic.
SYMBOLS
The aerodynamic force and moment data are referred to both the body and wind axes, the origin being on the model reference line and 1.3 reference diameters forward of the model base as shown in figure 1. The hinge-moment data (obtained only with 4.65-square-meter fins) are referred to the hinge line of the trailing-edge control.
2 CA axial-force coefficient corrected for base pressure, force uncorrected for base pressure qs
CA,b base axial-force coefficient, - bp,b,lfT) + @p,b72)(y]
CD drag coefficient corrected for base pressure, (CA cos a + CN sin a)
ch,2 hinge-moment coefficient of trailing-edge control on fin 2 (see fig. l(a)), Hinge moment qsccc
cL lift coefficient corrected for base pressure, (CN cos a - CA sin a) Pitching moment Cm pitching-moment coefficient, qsdref
Normal force CN normal-force coefficient, qs
pb,l pressure coefficient in balance chamber, - CP7b7 1 q
pressure coefficient at rim of model base, pb,2 - CP,b,2 cl
Ab,l area of sting hole in booster airframe
Ab,2 area of rim at base of booster airframe - CC mean aerodynamic chord (rearward of hinge line) of single trailing-edge cont r o 1
d diameter
reference length (base diameter)
lift-drag ratio, CL/C~
M Mach number of undisturbed stream 3
I static pressure of undisturbed stream
static pressure in balance chamber
static pressure at rim of model base
dynamic pressure of undisturbed stream
Reynolds number, based on reference length dref
radius
reference area, :(drefy
area rearward of hinge line of single trailing-edge control
longitudinal distance of center of pressure (point on model reference line through which line of action of resultant of normal and axial forces passes) from model base; positive direction forward
center-of-pressure parameter
angle of attack, based on model reference line
deflection of trailing-edge control measured in plane perpendicular to hinge line; positive direction when trailing edge is down; all four controls had same deflection ch, - da,dCh72, per deg
4 CN,! = -dCN, per deg da
- dCN, per deg cN6 -d6
APPARATUS
Tunne 1
The investigation was made in t le Langley 8-foot transonic pressure tunnel. The test section in this tunnel is square in cross section with the upper and lower walls axially slotted to permit changing the test section Mach number continuously from 0 to over 1.20 with negligible effects of choking and blockage. The total pressure of the tun- nel air can be varied from a minimum value of about 0.25 atmosphere (0.25 X 105.N/m2) at all test Mach numbers to a maximum value of about 1.5 atmospheres (1.52 X lo5 N/m2) at transonic Mach number and about 2.0 atmospheres (2.03 X lo5 N/m2) at Mach numbers of 0.40 or less. The stagnation temperature of the tunnel air is automatically controlled and is usually held constant at 120° F (49O C). The tunnel air is dried until the dewpoint temperature in the test section is reduced sufficiently to avoid condensation effects.
Model The model used in the present investigation was a 0.03-scale model of a proposed Little Joe 11 - LM suborbital space vehicle which consisted of the Little Joe I1 launch vehicle and an aerodynamic shroud to house the LM. Four stabilizing fins with trailing- edge controls were spaced 90° apart at the base of the launch vehicle. The model was supported in the tunnel by a sting which had a ratio of sting diameter to model base diameter of 0.30. A drawing of the model is given in figure 1, and photographs of several of the configurations are shown as figure 2. Peripheral corrugations, evident in the photographs of the Little Joe II launch vehicle, have been left off the drawing in figure l(a) for simplicity; the dimensions of the corrugations are given in reference 3. Geometric characteristics of the model are given in table I.
5 Three sizes of stabilizing fin with planform areas (full scale) per fin of 4.65, 7.33, and 9.30 square meters (identified herein as small, intermediate, and large, respec- tively) were investigated. The fins had wedge-shaped sections with blunt trailing edges and 45' sweepback of the leading edge. Aerodynamic control on the Little Joe I1 launch vehicle is achieved by the use of a flap-type control surface at the trailing edge of each fin. Various deflections of the trailing-edge controls were obtained by the use of deflection-setting plates bent to the proper angles. Details of the fins are given in fig- ure 3. For part of the investigation, a trailing-edge control with a hinge-moment beam for hinge-moment measurements was used on fin 2 of the small fins. (See fig. 3(b).) Hydraulic actuator fairings near the midspan of the fins and reaction control fairings at the roots of the fins as shown in figures l(a), 2, and 3 were included on the configurations. Three nose shapes on the LM aerodynamic shroud, identified herein as shroud noses 1, 2, and 3, were investigated. Photographs and drawings of the shroud noses are given in figures 2 and 4, respectively. Shroud nose 1 was a blunt-nosed cone of 15' half-angle. Shroud nose 2 was a combination of a blunt-nosed cone of 30° half-angle and a frustum of a cone of 15O half-angle. Shroud nose 3 was a combination of a blunt-nosed cone of 60' half-angle and a frustum of a cone of 150 half-angle. All shroud noses were attached to the cylindrical portion of the LM shroud as shown by the drawings. The cylindrical portion of the LM shroud was extended rearward 1.905 meters (full scale) over the launch vehicle in several tests by two types of shroud skirt in an attempt to improve the static longitudinal stability characteristics: a 4O boattailed shroud skirt and a cylindrical (Oo boattail) shroud skirt as shown in figure 5.
Instrumentation Aerodynamic forces and moments were measured with a six-component internal strain-gage balance housed in the model launch vehicle airframe. The model and balance were supported by a 3.49-centimeter-diameter sting which, in turn, was attached to the remotely operated tunnel central support system. A .one-component strain gage mounted in the trailing-edge control of fin 2 of the small fins (see figs. l(a) and 3(b)) was used for determining the hinge moment of the trailing-edge control. A static-pressure orifice, located within the chamber surrounding the six- component strain-gage balance, and eight static-pressure orifices, located around the rim of the model base and manifolded to a single tube, were connected to pressure- sensitive electrical pickups. These static pressures were used in the base-pressure corrections. The overall forces and moments on the model, the hinge moment of the trailing- edge control of fin 2 of the small fins, the angle of attack, and the static pressure in the
6 chamber surrounding the strain-gage balance and that on the rim of the model base were recorded electronically on punch cards.
TESTS, CORRECTIONS, AND ACCURACY
Tests The model investigated consisted of the Little Joe 11 launch vehicle in combination with the LM aerodynamic shroud. Three nose shapes on the LM aerodynamic shroud, two shroud skirts, and three sizes of stabilizing fins were investigated. The model was also tested with the fins off. The configurations with the fins on always included the reaction control fairings and, except for one test, always included the hydraulic actuator fairings. Usually, the tests were made with the shroud skirts off. The various configurations of the Little Joe 11 - LM space vehicle were investi- gated at Mach numbers from 0.30 to 1.20 and at angles of attack from approximately -12O to 12O. The stagnation temperature was 120' F (49O C) except at the two lowest Mach numbers where it was usually a few degrees less. The total pressure of the investiga- tion was 1 atmosphere at Mach numbers up to 0.95 for most of the configurations; how- ever, at high subsonic Mach numbers for configurations with large control-surface deflections, and at Mach numbers of 1.00 and 1.20 for all configurations, the total pres- sure was maintained at less than 1 atmosphere in order to stay within balance load limits. Table I1 gives the variation of Reynolds number with Mach number for the various configurations. All configurations were investigated in pitch at a sideslip angle of Oo. Uniform deflections only of the four trailing-edge controls were investigated. The small fins were investigated at control deflections of Oo and -100, the intermediate fins at control deflections of Oo, -loo, and -20°, and the large fins at control deflections of Oo, -so, -loo, -20°, and -3OO. The hinge moment of the trailing-edge control on fin 2 of the small fins was measured at deflections of Oo and -loo on the configurations with shroud noses 1 and 2 and shroud skirt off. All configurations were tested with natural transition on the model. The configura- tion with shroud nose 2, shroud skirt off, and fins off was also tested with three 0.16-centimeter-wide strips of No. 60 carborundum grains located immediately aft of the surface junctures on the LM shroud as indicated in figure l(b).
Correcti ons
The axial-force coefficient CA includes the correction for the base axial-force coefficient CA,b; the coefficients CD and CL, which are based on CA and CN,
7 are therefore also corrected for base pressure. The aerodynamic force and moment data presented herein are considered to be free of tunnel-boundary interference. No sting-interference corrections have been made to the data except to the extent of the partial correction for sting interference inherent in the base-pressure correction. The angle of attack has been corrected for the deflection of the balance and sting support under aerodynamic load.
Accuracy The accuracy of the data, based primarily on the static calibrations and the repeatability of the data, is estimated to be as follows: a,deg...... rtO.1 M ...... *0.005 At Mach numbers of 0.70 and above:
CN ...... rt0.02 CA ...... rtO.008 cm ...... *0.01 CL ...... rto.02 CD ...... rtO.008 ch,2 ...... ho.007
At Mach numbers less than 0.70, where the dynamic pressures were substantially less, the accuracy of the data expressed in aerodynamic-coefficient form was corre- spondingly poorer than that listed. During the tests, the aerodynamic loading on the trailing-edge controls modified the no-load (M = 0) values of the control deflection. The actual values of the control deflection of the control instrumented for hinge-moment measurements (on fin 2 of small fins) are given by the following empirical equations for the two control deflections investigated: 6 = 0' - 0K0o (rM2 6 = -10' + (-K ) (12' - a)M2 where a is in degrees, Oo and -loo are the no-load (M = 0) values of the control deflec- tion, and the empirical constants K0o and K-lOo are as follows:
8 The trailing-edge controls which were not instrumented for hinge-moment measurements had a deflection flexibility which was estimated to be about one-half that of the instru- mented trailing-edge control. In several of the figures presented herein, the no-load values of control deflection are used in identifying the data.
PRESENTATION OF RESULTS
The basic longitudinal aerodynamic data are presented in figures 6 to 14 as plots of aerodynamic coefficients against angle of attack. Hinge-moment characteristics are included in figures 11 and 14. The results have been arranged into groupings which make possible a more direct comparison of the effects of model components and control deflec- tions; this arrangement necessitated the presentation of the results for some of the con- figurations more than once. Center-of -pressure results are presented herein for the configurations with fins on and controls undeflected and with fins off. The value of the center-of-pressure parameter Xcp/dref shown herein at an angle of attack of Oo was taken as the value of the average slope ACm/ACN between angles of attack from approximately -loto lo; at angles of attack other than Oo the value of Xcp/dref shown was obtained by using the following equation:
In this equation, Cm,eLyl and CN,ea1 are the values of Cm and CN, respectively, corresponding to a given value a1 of angle of attack, and Cm,cN=o is the value of Cm when CN = 0. The center-of-pressure location is referenced to the base of the model. Summary longitudinal aerodynamic characteristics are shown plotted against Mach number in figures 15 to 24. Schlieren photographs are presented in figure 25. The slopes CN~,Cma, CmcN, and ch, shown herein are average values between angles
9 of attack from approximately -lLo to 1-lo . The slopes Ch6. CN~.and Cm6 shown 2 2 herein are average values between control deflections from approximately Oo to approx- imately .loo . The results of this investigation are presented in the following figures: Figure Basic longitudinal aerodynamic characteristics: Effect of shroud nose (fins off; shroud skirt off) ...... 6 Effect of shroud skirts (shroud nose 1; fins off) ...... 7 Effect of boattailed shroud skirt (shroud nose 1; small fins. 6 = Oo) ...... 8 Effect of boattailed shroud skirt (shroud nose 1; large fins. 6 = Oo) ...... 9 Effect of boattailed shroud skirt (shroud nose 2; small fins. 6 = Oo) ...... 10 Effect of control deflection (shroud nose 1; small fins) ...... 11 Effect of control deflection (shroud nose 1; intermediate fins) ...... 12 Effect of control deflection (shroud nose 1; large fins) ...... 13 Effect of control deflection and transition strips (shroud nose 2; small fins) ...... 14 Summary longitudinal aerodynamic characteristics: Effect of shroud nose (fins off; shroud skirt off) ...... 15 Effect of shroud skirts (shroud nose 1; fins off) ...... 16 Effect of boattailed shroud skirt (shroud nose 1; small fins. 6 = 0') ...... 17 Effect of boattailed shroud skirt (shroud nose 1; large fins. 6 = Oo) ...... 18 Effect of boattailed shroud skirt (shroud nose 2; small fins. 6 = Oo) ...... 19 Effect of fin size (shroud nose 1; 6 = Oo) ...... 20 Effect of control deflection (shroud nose 1; small fins) ...... 21 Effect of control deflection (shroud nose 1; intermediate fins) ...... 22 Effect of control deflection (shroud nose 1; large fins) ...... 23 Effect of control deflection and transition strips (shroud nose 2; small fins) ...... 24 Schlieren photographs of flow about LM shroud with shroud noses 1 and 2. shroud skirt off ...... 25
DISCUSSION
Longitudinal St abi lit y All configurations tested had static longitudinal instability at all test conditions about a moment reference center 1.3 base diameters forward of the base . (For example. see fig. 13(d). ) The longitudinal-stability parameter increased (configuration cmcY became more unstable) as the Mach number was increased to about 0.95, then decreased
10
. (configuration became less unstable) with further increase in Mach number to 1.00, and then either remained essentially constant (configuration instability remained same) or decreased (configuration became less unstable), depending on configuration, with further increase in Mach number above 1.00. (See figs. 20 and 24.) Fin size.- Adding stabilizing fins to the configuration or increasing the size of the fins reduced the amount of static longitudinal instability as might be expected, but not enough to make the configuration stable. (See figs. ll(d), 12(d), 13(d), and 20.) The center-of-pressure location varied somewhat with change in angle of attack and Mach number, the location being farthest forward of the base usually at angles of attack near 0' and at Mach numbers near 0.95. For configurations with shroud nose 1, the center of pressure was forward of. the model base by 3.5 to 5.3 base diameters for the fins-off configuration (fig. ll(h)), by 2.0 to 2.6 base diameters for the configuration with the small fins (fig. ll(h)), by 1.6 to 2.2 base diameters for the configuration with the intermediate fins (fig. 12(h)), and by 1.4 to 1.9 base diameters for the configuration with the large fins (fig. 13(h)). Shroud-nose modifications. - The effect on the static longitudinal stability of changing the shape of the shroud nose was usually quite small. Shroud nose 2 reduced the static longitudinal instability at Mach numbers less than approximately 0.86 and at a Mach number of 1.20. Shroud nose 3 reduced somewhat more the static longitudinal instability at Mach numbers less than approximately 0.92. At the other Mach numbers, however, shroud noses 2 and 3 generally increased the static longitudinal instability. (See fig. 15 and compare figs. 17 and 19.) Shroud nose 2 caused a large shift in the pitching-moment curves and, in general, a reduction in the amount of static longitudinal instability at the higher angles of attack at a Mach number of 0.90 for the configuration with the transition strips both on and off, and at Mach numbers of 0.95 and 1.00 for the configuration with the transition strips on (fig. 14(d)). The effect of shroud nose 2 on the aerodynamic characteristics at these test conditions is discussed more fully in the section "Unsteady Flow.'' Shroud skirts.- The effect of the shroud skirts on the static longitudinal stability was generally small, usually reducing the amount of static longitudinal instability at Mach numbers near 0.90, but increasing the amount of static longitudinal instability at the other Mach numbers. (See figs. 16 to 19.) However, the shroud skirts had a favor- able effect on the undesirable aerodynamic characteristics observed for the configura- tion with shroud nose 2 at the higher angles of attack at a Mach number of 0.90 (fig. 10(d)); this effect is discussed further in the section "Unsteady Flow."
11 Unsteady Flow
Hammerhead-shaped__ - - _____ configurations.-- As figure 1 shows, the Little Joe 11 - LM space vehicle was of the hammerhead type, in that the diameter of the LM aerodynamic shroud was larger than the diameter of the Little Joe I1 launch vehicle. The ratio of these diameters was 1.38. As noted in references 5 to 7, pressure fluctuations exist in two types of flow regions on hammerhead-shaped configurations. First, at high subsonic free-stream Mach numbers there are the regions of flow expansions followed by shock waves (for example, at cone-cylinder junctures) in which there are extremely local but relatively high pressure fluctuations. Second, there are the regions of separated flow in which large areas may be subjected to unsteady flow and, hence, to pressure fluctuations. Flow separation due to the hammerhead configuration, particularly at subsonic Mach numbers, can result in the significant pressure fluctuations associated with buffeting. References 6 and 7 also point out that an abrupt body convergence angle between the shroud and the launch vehicle can extend the separation effects and pressure fluctuations to supersonic Mach numbers. In the present investigation, the body convergence was a frustum of a cone of 58.9' half-angle. This body convergence angle was abrupt, and schlieren photographs indicate that separation off the shroud occurred at all Mach num- bers including 1.20.
Model buffeting and abrupt changes in forces__ and moments.- Severe model buffeting (visually observed) occurred at a Mach number of 0.90 for the configuration with shroud nose 2 and shroud skirt off in two narrow ranges of angles of attack near 70 and -7'. This buffeting occurred with the fins both on and off. The configurations with shroud noses 1 and 3, however, did not exhibit this severe buffeting. Associated with the severe model buffeting were abrupt changes in the magnitudes of the aerodynamic forces and moments as the angle of attack was increased through the critical range (fig. 14). These abrupt changes included a large decrease in the magnitude of the pitching moment, a rearward movement of the center of pressure of about 1.2dref when the fins were on and about 2.5dref when the fins were off, an increase in normal force of about 25 percent when the fins were on and an insignificant amount when the fins were off, and an increase in axial force of about 25 percent. Effect of transition strips.- Three transition strips of No. 60 carborundum grains were added to the configuration with shroud nose 2, fins off, and shroud skirt off, as indicated in figure l(b). Brief tests of this configuration were made at Mach numbers from 0.80 to 1.00 to determine primarily the effect of the transition strips on the model buffeting and on the abrupt changes in the forces and moments observed at a Mach num- ber of 0.90 for the configuration without the transition strips. Direct comparisons of the effect of the transition strips are available at test Mach numbers of 0.80, 0.90, and 1.00, where tests were made both with and without the transition strips. Tests were also
12 made at Mach numbers of 0.85 and 0.95 of the configuration with the transition strips on. (See fig. 14.) The transition strips had no effect on the force and moment characteristics at a Mach number of 0.80. At a Mach number of 0.90, the transition strips had only a small effect on the magnitudes of the abrupt changes in pitching moment, center of pressure, and axial force, but did lower (by somewhat less than lo)the angle-of-attack range in which the abrupt changes occurred. At a Mach number of 1.00, the transition strips caused flow changes at angles of attack near go which resulted in model buffeting and abrupt changes in pitching moment, center of pressure, normal force, and axial force. These characteristics were not observed at a Mach number of 1.00 for the configuration with the transition strips off. Addition of the transition strips also reduced the axial force and the base axial force at all angles of attack at a Mach number of 1.00. At a Mach number of 0.95, the configuration with the transition strips experienced severe buffeting and the characteristic abrupt changes in pitching moment and center of pressure associated with the buffeting. These characteristics occurred at angles of attack near 6O at a Mach number of 0.95. At a Mach number of 0.85, the configuration with the transition strips was not affected by buffeting or abrupt changes in forces and moments. The results of reference 8 show that at Mach numbers less than 0.95 the flow.rear- ward of the juncture on a 30° half-angle cone-cylinder configuration was markedly affected by the presence of a transition strip located immediately aft of the juncture (the flow was attached when the transition strip was on, separated when the transition strip was off). A relative scarcity of roughness particles in a part of the transition strip due either to erosion or a lack of care in the application of the transition strip also was found to have an important effect on the separation characteristics. The flow on a 15O half- angle cone-cylinder configuration, however, was essentially the same (that is, attached) whether the transition strip was on or off. In the present investigation, shroud nose 2 consisted of a blunt-nosed cone of 30° half-angle and a frustum of a cone of 15O half- angle followed by a cylinder. The sensitivity of the flow on the configuration with shroud nose 2 to transition-strip effects would be expected to be somewhat less than that of the flow on a 30° half-angle cone-cylinder configuration but more than that of the flow on a 15O half-angle cone-cylinder configuration. Effect of shroud skirt.- Addition of the boattailed shroud skirt to the smooth con- figuration with shroud nose 2 and fins on (fig. 10) substantially reduced the model buffeting and the extent of the large abrupt changes in aerodynamic forces and moments at a Mach number of 0.90, and delayed the occurrence of these abrupt changes to some- what higher angles of attack. At a Mach number of 0.95, the configuration with the boat- tailed shroud skirt did not buffet or show the abrupt changes in forces and moments
13 .. , evident at a Mach number of 0.90. The cylindrical shroud skirt was not tested with shroud nose 2.
Schlieren photographs..- -*___ - Schlieren photographs are shown in figure 25(a) for the configuration with shroud nose 1 and shroud skirt off at Mach numbers of 0.80 and 0.90 for angles of attack from 00 to 4' and in figure 25(b) for the configuration with shroud nose 2 and shroud skirt off at a Mach number of 0.90 for angles of attack from -7O to 8O. Typical regions of flow expansions followed by shock waves are evident at the sphere- cone juncture of shroud nose 1 (fig. 25(a)), the cone-cone juncture of shroud nose 2 (fig. 25(b)), and the cone-cylinder junctures of both shroud noses 1 and 2. At a Mach number of 0.90, a shock-wave system was located in the region of the converging juncture between the shroud and the launch vehicle for the configurations with both shroud noses 1 and 2. The schlieren photographs show that for the configuration with shroud nose 2 (fig. 25(b)) the shock-wave system on the windward surface of the shroud moved upstream and away from the converging juncture as the angle of attack was increased through the critical range. (Compare the schlieren photographs for angles of attack of -6O and -7O and for 6O and 8O.) Also evident in the schlieren photographs at an angle of attack of 8' is a small region of flow expansion at the beginning of the converging juncture. This region of flow expansion is evident to a lesser extent at an angle of attack of -7O, but not at lower angles of attack. Schlieren photographs were not obtained for the configuration with shroud nose 1 at these higher angles of attack or for the configurations with shroud skirt. An explanation for the severe model buffeting at a Mach number of 0.90 is suggested by the observed facts that the buffeting was eliminated by the addition of the boattailed shroud skirt and that the buffeting did not occur for the configuration with shroud nose 1, and by a study of the schlieren photographs. The buffeting is possibly explained by the unsteady flow in the region of the converging juncture, the unsteadiness of this flow having been aggravated or intensified by the adverse effects of the interactions with the boundary layer of the shock waves in the regions of the converging juncture, the cone- cylinder juncture, and the cone-cone juncture. The strength and type of the shock waves in the regions of the cone-cylinder and the cone-cone junctures are dependent, of course, on the extent and degree of the flow expansions and the nature of the boundary layer pre- ceding these shock waves.
Longitudinal Control Longitudinal aerodynamic control on the Little Joe I1 - LM space vehicle is obtained from flap-type trailing-edge controls on the stabilizing fins. As figure 3 shows, the planforms of the controls used on the three sizes of fins tested in this investigation were dissimilar and, whereas full-span controls were used on the small and intermediate fins, only partial-span controls were used on the large fins. 14 The intermediate fins were approximately 50 percent larger than the small fins in both the total area of the fins and in the area of the control. The large fins were 100 per- cent larger than the small fins in the total area of the fins, but the area of the control on the large fins was actually less than the area of the control on the small fins. The controls on all three sizes of stabilizing fins were found to be effective for pitch control at all test conditions. (See figs. ll(d), 12(d), 13(d), and 14(d).) However, there were reductions in control effectiveness at control deflections greater than - 100 for the controls on the intermediate fins (fig. 12(d)) and on the large fins (fig. 13(d)) at all angles of attack and Mach numbers. The controls on the small fins were not investi- gated at deflections greater than -loo (figs. ll(d) and 14(d)). There were also losses in control effectiveness at all control deflections at the highest angles of attack at Cm6 Mach numbers of 0.95 and 1.00 for the controls on the intermediate fins (fig. 12(d)) and on the large fins (fig. 13(d)), but not for the controls on the small fins (figs. ll(d) and 14(d)). The variation with Mach number of the control-effectiveness parameter Cm6 is shown in figure 20(b) for the controls on the three sizes of fins, and this variation is seen to be similar for all three sizes of fins. There was first a slight increase in con- trol effectiveness as the Mach number was increased to about 0.85, then a more rapid increase in control effectiveness as the Mach number was further increased to about 0.90, and then a uniform reduction in control effectiveness as the Mach number was increased above 0.90, so that at a Mach number of 1.20 the control effectiveness was less than that at the lowest Mach numbers. Figure 20(b) shows that the control effectiveness was greatest for the controls on the intermediate fins. This greater control effectiveness was due, of course, primarily to the fact that the area of the control on the intermediate fins was substantially larger than the area of the control on the large fins (56 percent larger) or the area of the con- trol on the small fins (50 percent larger).
Control Hinge Moments Hinge-moment data were obtained only for the control on fin 2 (fig. l(a))of the small fins. The variation of hinge-moment coefficient ch 2 with angle of attack was 9 approximately linear at subsonic Mach numbers for the two control deflections of Oo and -loo investigated. (See figs. ll(i) and 14(i).) The divergence from a linear variation of ch,2 with angle of attack was greater at a Mach number of 1.00 than at the lower Mach numbers, and this divergence was still greater at a Mach number of 1.20. The hinge-moment coefficient C and the hinge-moment parameters cha! h,2 and ch6 are shown plotted against Mach number at an angle of attack of Oo in
15 figures 21(b), 21(c), 24(b), and 24(c). The magnitude of the parameter ch6 is seen to be substantially greater at Mach numbers of 0.90 and above than at the lower Mach numbers. The magnitudes of the hinge-moment results of the present investigation and the trends with angle of attack, control deflection, and Mach number check well with the corresponding results obtained in the investigation of reference 3 on the same fin-control configuration on the Little Joe I1 launch vehicle when tested in combination with the Apollo spacecraft. In the investigation of reference 3 the diameter of the Apollo space- craft was the same as that of the Little Joe II launch vehicle, and the Apollo spacecraft, which was essentially conical of 33O half-angle, included the launch escape system. The control-effectiveness hinge-moment parameter shown plotted cm6ph6 against Mach number in figures 21(c) and 24(c) is a measure of the effectiveness of the control in producing a pitching moment in terms of the hinge moment developed on the control while producing this pitching moment. A large value of this parameter is desir- able since such a value would tend to indicate good pitch effectiveness and low control hinge moments. It is seen in figures 21(c) and 24(c) that the control effectiveness parameter cm6/ch6 was, on the average, essentially invariant as the Mach number was increased to 0.80, but then decreased substantially with further increase in Mach number to 1.20.
CONCLUSIONS
An investigation of the static longitudinal aerodynamic characteristics of a 0.03-scale model of a proposed Little Joe I1 - lunar module suborbital space vehicle was made in the Langley 8-foot transonic pressure tunnel at Mach numbers from 0.30 to 1.20 and at angles of attack from approximately -12O to 12O. This configuration was of the hammerhead type, in that the diameter of the lunar module aerodynamic shroud was larger than the diameter of the Little Joe 11 launch vehicle. Three sizes of stabi- lizing fins, three nose shapes on the lunar module aerodynamic shroud, and two shroud skirts, which extended over the converging juncture between the shroud and the launch vehicle, were investigated. Flap-type trailing-edge controls on the stabilizing fins were uniformly deflected for pitch control from Oo to -loo on the small fins, from Oo to -20° on the intermediate fins, and from Oo to -30° on the large fins. Control hinge moments were measured on the small fins. Limited transition-strip studies were made on one of the configurations. The following conclusions are indicated: 1. All configurations investigated had static longitudinal instability at all test con- ditions about a moment reference center 1.3 base diameters forward of the base.
16 2. Depending on angle of attack and Mach number, the center of pressure was for- ward of the model base by 3.5 to 5.3 base diameters for the fins-off configuration, by 2.0 to 2.6 base diameters for the configuration with the small fins, by 1.6 to 2.2 base diam- eters for the configuration with the intermediate fins, and by 1.4 to 1.9 base diameters for the configuration with the large fins. 3. Shroud-nose changes or the addition of shroud skirts to the configuration usually had a small and variable effect on the static longitudinal stability. At a Mach number of 0.90, however, the flow on the configuration with one of the shroud noses was very criti- cal to model attitude. 4. Severe model buffeting occurred at a Mach number of 0.90 in two narrow ranges of angles of attack near -7O and 7O for the configuration with a shroud nose consisting of a blunt-nosed cone of 30° half-angle and a frustum of a cone of 15O half-angle. Associ- ated with the severe model buffeting were abrupt changes in the magnitudes of the aero- dynamic forces and moments as the angle of attack was increased through the critical range. 5. The addition of transition strips to the surface junctures on the lunar module shroud had no beneficial effect on the severe model buffeting or the abrupt changes in forces and moments which occurred at a Mach number of 0.90, and actually precipitated model buffeting and abrupt changes in forces and moments at a Mach number of 1.00. Severe model buffeting and abrupt changes in forces and moments also occurred at a Mach number of 0.95 for the configuration with the transition strips on. 6. The addition of a boattailed shroud skirt to the smooth configuration substantially reduced the model buffeting and the extent of the abrupt changes in the aerodynamic forces and moments which occurred at a Mach number of 0.90 and delayed the occurrence of these abrupt changes to somewhat higher angles of attack. No model buffeting or abrupt changes in forces and moments occurred at a Mach number of 0.95 for the con- figuration with the shroud skirt. 7. The trailing-edge controls on all three sizes of stabilizing fins were effective for pitch control at all test conditions, although reductions in control effectiveness occurred at control deflections greater than -loo at all angles of attack and Mach num- bers and also occurred at all control deflections for the controls on the intermediate and the large fins at the highest angles of attack at Mach numbers of 0.95 and 1.00. 8. The magnitude of the rate of change of hinge-moment coefficient with control deflection was greater at Mach numbers of 0.90 and above than at the lower Mach numbers
17 9. The ratio of pitching moment produced by control deflection to hinge moment developed by control deflection was less at Mach numbers of 1.00 and 1.20 than at low subsonic Mach numbers.
Langley Research Center, National Aeronautics and Space Administration, Langley Station, Hampton, Va., May 2, 1967, 124-07-02-42-23.
REFERENCES
1. Fournier, Roger H.; and Wassum, Donald: Static Stability and Control Characteristics of a 0.030-Scale Model of a Little Joe II-LEM Space Vehicle at Mach 1.57 to 4.63. NASA TM X-1139, 1965. 2. Henderson, William P. : Subsonic Static Longitudinal Aerodynamic Characteristics at Angles of Attack From -2O to 95' of a Proposed Little Joe II - Apollo Space Vehi- cle. NASA TM X-751, 1963. 3. Luoma, Arvo A.: Longitudinal Stability and Control, Roll Control, and Hinge Moments of a Litt1.e Joe 11 - Apollo Model at Mach Numbers up to 1.20. NASA TM X-1002, 1964. 4. Fournier, Roger H.; and Silvers, H. Norman: Static Longitudinal Aerodynamic Char- acteristics of a 0.028-Scale Model of a Proposed Little Joe II - Apollo Space Vehicle at Mach Numbers From 1.50 to 2.16. NASA TM X-754, 1963. 5. Coe, Charles F.: Steady and Fluctuating Pressures at Transonic Speeds on Two Space-Vehicle Payload Shapes. NASA TM X-503, 1961. 6. Coe, Charles F.: The Effects of Some Variations in Launch-Vehicle Nose Shape on Steady and Fluctuating Pressures at Transonic Speeds. NASA TM X-646, 1962. 7. Coe, Charles F.; and Nute, James B.: Steady and Fluctuating Pressures at Transonic. Speeds on Hammerhead Launch Vehicle. NASA TM X-778, 1962. 8. Kelly, Thomas C.: A Transonic Investigation of Base Pressures Associated With Shallow Three-Dimensional Rearward-Facing Steps. NASA TN D-2927, 1965.
18 I
TABLE 1.- GEOMETRIC CHARACTERISTICS OF 0.03-SCALE MODEL OF PROPOSED LITTLE JOE II - LM SUBORBITAL SPACE VEHICLE
Little Joe 11 launch vehicle: Body: Cross-sectional area (reference area, S), m2 ...... 0.0108 Diameter (reference length, dref), cm ...... 11.73 Area of sting hole, Ab,l, m2 ...... 0.0054 Area of rim, ~b,2,m2 ...... 0.0055 Stabilizing fins (exposed, single-fin values given) : Small fins (4.65 m2 full scale): Airfoil section (parallel to root chord) ...... Wedge, 100 total angle Area (includes area of control), m2 ...... 0.0042 Span,cm...... 7.087 Root chord, cm ...... 6.688 Tip chord, cm ...... 3.360 Aspect ratio ...... 1.20 Sweepback of leading edge, deg ...... 45 Trailing-edge control: Type...... Flap Area, aft of hinge line, Sc: m2...... 0.0013 Percent of fin area ...... 30 - Mean aerodynamic chord, cc, cm ...... 2.428 Intermediate fins (7.33 m2 full scale): Airfoil section (parallel to root chord) ...... Wedge, 10' total angle Area (includes area of control), m2 ...... 0.0066 Span,cm...... 8.659 Root chord, cm...... 9.078 Tip chord, cm ...... 2.705 Aspect ratio ...... 1.141 Sweepback of leading edge, deg ...... 45 Trailing-edge control: Type...... Flap Area, aft of hinge line, Sc: m2...... 0.0019 Percent of fin area ...... 28.7 - Mean aerodynamic chord, cc, cm ...... 2.913
19 I I1 I I I II .,, , ......
TABLE 1.- GEOMETRIC CHARACTERISTICS OF 0.03-SCALE MODEL OF PROPOSED LITTLE JOE II - LM SUBORBITAL SPACE VEHICLE - Concluded
Large fins (9.30 m2 full scale): Airfoil section (parallel to root chord) ...... Wedge, 10' total angle Area (includes area of control), m2 ...... 0.0084 Span, cm ...... 9.91 Root chord, cm ...... 9.868 Tipchord, cm...... 4.57 Aspect ratio ...... 1.17 Sweepback of leading edge, deg ...... 45 Trailing-edge control: Type...... Flap Area, aft of hinge line, Sc: m2 ...... 0.0012 Percent of fin area ...... 14.3 - Mean aerodynamic chord, cc, cm ...... 2.743 LM aerodynamic shroud: Cylindrical section : Cross-sectional area, m2 ...... 0.021 Diameter, cm ...... 16.15 Shroud nose length: Shroud nose 1, cm ...... 12.83 Shroud nose 2, cm ...... 10.43 Shroud nose 3, cm ...... 9.91 Overall length from nose of LM shroud to base of launch vehicle: Shroud nose 1, cm ...... 52.98 Shroud nose 2, cm ...... 50.60 Shroud nose 3, cm ...... 50.06
20 TABLE 11.- VARIATION OF REYNOLDS NUMBER WITH MACH NUMBER
Configuration Reynolds number at -
Shroud I Fins 6, deg M = 0.30 M = 0.50 M = 0.70 M = 0.80 M = 0.85 M = 0.90 M = 0.93 M = 0.95 M = 0.97 M = 1.00 M = 1.20 Nose Skirt
1.53~10~1.54~10~ 1.55~10~ 1.56~10~ 1.27x106 1.05~10~ 1 Off Small -10 .71 1.09 1.05 1 Off Off .71 1.07 1 1.06 1 Off Intermediate 0 .70 1.08 1 .96 1 Off Intermediate -10 ------1.08 1 .96 1 Off Intermediate -20 ------1.08 1 .82 1 Off Large 0 .71 1.08 1 .97 1 Off Large -5 .70 1.08 1 .96 1 Off Large -10 .69 1.04 1 .96 1 Off Large -20 .70 1.07 1 .81 1 Off Large -30 .70 1.07 1 .77 2 Off Small 0 .70 1.07 1.05 2 Off Small -10 .72 1.08 1.05 2 Off Off .71 1.05 a2 Off Off ------1 ------a2 Off Off ------3 Off Off .69 1 1.05 1 Cylindrical Off -____-- 1 1.06 1 Boattailed 1 1.05 1 Boattailed Small i -:-L 1 .96 1 Boattailed Large 1 .96 2 Boattailed Small 1 .96 I - aWith transition strips. h3 N
d.16.15
Reaction control fairings
LM shroud nose LM shroud body Little JoelI launch vehicle -12.83 14.63
LM aerodynamic shroud
~ 27.46-
(a) Drawing. Shroud nose 1; large fins; 6 = 00.
Figure 1.- Drawing of 0.03-scale model of Little Joe I1 - LM space vehicle and location of transition strips on LM shroud. Shroud skirt off. All dimensions are in centimeters unless otherwise noted. -23.72 IO 43 -: II 6.6 I
Tangency
UTransit ion st rips
(b) Location of transition strips. Shroud nose 2; fins off.
Figure 1.- Concluded. (a) Shroud nose 1; small fins; 6 = -100. L-63-9533
(b) Shroud nose 2; small fins; 6 = -loo. L-63-9532 Figure 2.- Photographs of Little Joe I I - LM model showing various configurations and model components.
24 (c) Shroud nose 3; small fins; 6 = Oo. L-64-2703
(d) Shroud nose 1; large fins; 6 = 00. L-64-2704 Figure 2.- Continued.
25 (e) Shroud nose 1; cylindrical shroud skirt; small fins; 6 = Oo. L-64-2702
(f) LM shroud noses. L-64-2701
Figure 2.- Concluded.
26 Hydraulic actuator fairing- 1
.03' gap' I Section A-A
I
(a) Small (4.65 m2) fin with uninstrumented trailing-edge control.
Figure 3.- Details of stabilizing fins. All dimensions are in cm unless otherwise noted. .I7 * -
- Sectlo" c-c
e rubber flexlbln filler d:0.16 L f-
Sectlon E-E / /," F - '
Set %rei no. 1-64 NC X 0.32 long
' --Hinge line 1'
Section A-A
(b) Small (4.65 m2) fin with instrumented trailing-edge control (fin 2).
Figure 3.- Continued.
L Hydraulic actuator fairing--\ IO0
2h2 1
7.51
I
(c) Intermediate (7.33 m2) fin.
Figure 3.- Continued. W 0 Hydraulic actuator fairing7
I Section A-Ab I------
(d) Large (9.30 m2) fin.
Figure 3.- Concluded. - d= 16.1 5 I Shroud nose I
Shroud nose 2
r= 11.687 I I t -- --d= 16.15 Shroud nose 3
Figure 4.- Details of LM shroud noses. All dimensions are in cm unless otherwise noted.
31 Boattailed shroud skirt
Cylindrical shroud skirt
Shroud skirt off
Figure 5.- . Details of LM shroud skirts. All dimensions are in cm unless otherwise noted.
32
- ...... _ . 1.6 - I----
Angle of attack,a,deg
(a) CN against a.
Figure 6.- Effect of shroud nose on longitudinal aerodynamic characteristics. Fins off; shroud skirt off. (Flagged symbols indicate points from repeat run.) w w .4 I I .2 .2 I.o A 4 'I iil* Va0 0, E .8 T" f .6 0, .-aJ .- 0 .-u .- Lc + Lc c .4 g .6 0 0 I aJ0 aJ0 I I p .2 p .4 I f I - I - .-0 .-0 I I ZO Z .2 T I
.6 0 I I ~ I.4 .4 I I.2 .2 1.0 0 z .8 .8 I N 7pc I .6 _14 .6 0 .4 .4 I .2 .2 I
0 I. 0 __ -16 -12 -8 -4 048I ! 16 -16 -12 -8 -4 04 8 12 16 Angle of attack,a ,deg Angle of attackp ,deg
(b) CA against a.
Figure 6.- Continued.
34 Shroud nose-- .4 r 0 1 U I1 L .2 0 3 .2 0 .p, a nIii I 3- .2 t C .-e, 0 -- 0
.2
0 .2 .2 0 0
.2 .2
0- -16 -12 -8 -4 0 4 8 12 16 116 -12 -8 -4 0 4 8 12 16 Angle of attack,a ,deg Angle of attack,o,deg
(c) CA,b against a.
Figure 6.- Continued. w ua
I I ! -2.8 -___- M.0.30 ---~M-0.50 ____ M =0.70 -__ M-0.80 M.0.85 ____ -16 -12 -8 -4 0 0 0 0 0 4 8 12 16 Angle of attack,a,deg
(d) Cm against a.
Figure 6.- Continued. -1.2-
-16 -12 -8 -4 0 0 0 0 4 8 12 16 Angle of attack,a ,deg
(d) Cm against a. Concluded. w Figure 6.- Continued. 4 w 03
Angle of attack,a ,deg
Angle of attack,a,deg
(e) CL against a.
Figure 6.- Continued. -16 -12 -8 -4 0 4 8 12 16 Angle of attack,a,deg Angle of attack,a,deg
(f) CD against a.
Figure 6.- Continued.
39
I 2
I Shroud nose
0
-I
-2 2
I 2
0 I
0 q -I >O .-0- .-2 t t e -2 m-1 e 82 P U I ._.I-+ AI
0
-I
-2 I 2 0 I -I 0 I
-I 0
a -2 -1 -16 -12 -8 -4 0 4 8 12 16 Angle of attack,a ,deg Angle of attock,a,deg
(g) VD against a.
Figure 6.- Continued.
40
I tIroud nose 4 1 7 3 6
21 5
M.0.90
'4 M.0.95 4
I I M=1.00 0- 0-
45L
31 2
I M.0.85 0- 0- 0 -16 -12 -8 -4 O 4 ' 16 -16 -12 -8 -4 8 12 16 Angle of attack,a ,deg Angle of attack,a,deg
h) Xcp/d,ef against a.
Figure 6.- Concluded.
41 -1.2 -----M.0.30 _--__ M=0.50 M.0.70 M=0.80 -12 -8 -4 0 0 0 0 4 8 12 16
__----
Angle of attack,a,deg
(a) CN against a.
Figure 7.- Effect of cylindrical and boattailed shroud skirts on longitudinal aerodynamic characteristics. Shroud nose 1; fins off. (Flagged symbols indicate points from repeat run.) Shroud skirt 0 Off .4
.2
0 .4
.2 0 i 0 1.0 .4 8 a U' Y- .2 ?- .-c0, c .-V .$ .6 c .- c -zo al $4 al s2 I .2 .-0 x U 0 I.2 .4 I.o .2 Ill .8 .M=0.85
.6
.4
.2
0 -16 -12 -8 - 404 12 16 Angle of attaCk,a ,deg Angle of attack,a,deg
(b) CA against a. Figure 7.- Continued.
43 Shroud skirt 0 Off .4 .4
.2 .2
0 0 .4 9 a y- .2 C t .-a, .-a, .-0 _-0 'c =0a, aJ 0 0 a, 2. w-0 -I 0 .-x 0 aJ a, In In 0 Q .2 m
n U .4
.2 .2 1
0 M=1.20( ~ -'-I 0 I 111 I I 11 ' -16 -12 -8 -4 0 4 8 12 16 -16 -12 -8 -4 0 4 8 12 16 Angle of attack,Q ,deg Angle of attackp,deg
(c) CA,b against a.
Figure 7.- Continued. --
E-
al - .-0
-16 -12 -8 -4 0 0 0 0 0 4 8 12 16 Angle of attack,a,deg
(d) Cm against a.
Figure 7.- Continued. Angle of attack,a,deg
(d) Cm against a. Concluded.
Figure 7.- Continued. . ._
.8
-1 :-.4 .-0) gou 0 + r -.4 A
Angle of attack,o ,deg
L." I I I I I
I .2
J y- .8 .-0) ._0 =48' + 50
-.4
-.8
Angle of attack,a,deg
(e) CL against a.
Figure 7.- Continued. Shroud skirt .4 dl 1.0 .2 .8
0 I/ .6 .4 .4 .2 .2 0 0 I.4 1 1 I.2 1 1.0
.8
n v .6 + .-cW u 'G .4 r W 8 e* .2 n 0 1.6
I.4
I.2 3
I.o
.8 .8
.6 .6
.4 Q .4
.2 .2
0 0 -16 -12 -8 -,4048 12 16 -16 -12 - 8 12 16 Angle of attack,a,deg Angle of attack,a ,deg
(f) CD against a.
Figure 7.- Continued. 3~ I 21 Shroud skirt I 0 Off a L
I/ ~ I O1 I -7 C -2 I z -I
I -2 2 PI .-0- 1 ce 80 P f 'c -I -.I
2 E t I I 0 0 -I -I
-2 I 2
I
0
--I
-2 -I I -16 -12 -8 -4 0 4 8 12 16 Angle of attack,a ,deg Angle of attack,a,deg
(g) VD against a.
Figure 7.- Continued.
49 S iud skirt 0 IricaI iiled 6
5
4
3
2
I
0 6
5
5 4
4 3
3 2 -XCP -XCP dref dref 2 I
I 0
5 5 i E 4 4
3 3
2 2
I I
0 0 -16 -12 -8 - 48112 16 -16 -12 -8 - 0 8 Angle of attack,a,deg Angle of attack,a ,deg
(h) xcp/dref against a.
Figure 7.- Concluded.
50 (a) CN against a. Figure 8.- Effect of boattailed shroud skirt on longitudinal aerodynamic characteristics. Shroud nose 1; small fins; 6 = 00. cn N
-2.0 M.0.93 M.0.95 M.0.97 M.1.00 M= 1.20 ___--- -16 -12 -8 -4 0 0 0 0 0 4 8 12 16 Angle of attack,a, deg
(a) CN against a. Concluded.
Figure 8.- Continued. Shroud skirt 0 Off 0 Boattailed
c t - 0 1 b
h 0 1.4
c 1.2 c / i
1.0
.8
.6
.4
-8 -4 0 4 4o1 M.1.20 -16 -12 12 16 -16 -12 -8 4 8 12 16 Angle of attack,a,deg Angle of attack,a,deg
mr (b) CA against a. Figure 8.- Continued.
, 53 'I Shroud skirt 0 Off Ir 0 Boattailed
.4
.2 4 a " ;-0 aJ al 0 g .4 .-'5 .4 .c cI c - W .-0 0 :: .2 :.2 Win 0 c mO .-Go 1 x 0 .4 inW m0 .2
0 .4 .4
.2 .2
0 1 0 -16 -12 -8 -4 0 4 8 12 16 -16 -12 -8 -4 0 4 8 12 16 Angle of attack,a,deg Angle of attackp, deg
(c) CA,b against a.
Figure 8.- Continued.
54 (d) Cm against a.
Figure 8.- Continued. C
-2.0 -16 -12 -8 Angle of attack,a,deg
(d) Cm against a. Concluded.
Figure 8.- Continued. c
Angle of ottackp,deq
(e) CL against a.
Figure 8.- Continued. I
Angle of attack,a ,de¶
(e) CL against a. Concluded.
Figure 8.- Continued. .6 0 .4
.2
.6 Shroud skirt 0 Off
.4 . I.o n .2 .8 0 \ +-c I ‘c .o, 0 M.0.50 .-V .6 Lc 00, v .6 .4 e01 n .4 e 2 c 0)K .2 .- M-0.83 go0 1.c ,M=0.70 0 010 .8 & .8 \ i d .6 .6 k I .4 1 f!I .4 I .2 I ll .2 I M=0.80- M=0.90 C C I - -I 2 -8 -4 0 4 8 12 16 - i -12 -a -4 0 4 8 12 16 Angle of attack,a,deg Angle of attack,a ,deg
(f) CD against a.
Figure 8.- Continued.
59 I.2
1.0
.a
.6 Shroud skirt 0 Off .4 Boattailed 1.6 .2 I.4 0 1.4 I.2
I.2 1.0
1.0 .8
0" .8 .6 c .-5 0 .6 .4 8 u0 g .4 z- .2 D ._0, .-0 1 .2 %O 8 CT, 0 g 1.6
I.4 I.4
I.2 1.2
I.o 1.0
.8 .8
.6 .6
.4 .4
.2 .2
n 26 -12 -8 -4 0 4 8 12 -46 -12 -8 -4 0 4 8 12 16 Angle of attack,a ,deg Angle of attack,a,deg
(f) CD against a. Concluded.
Figure 8.- Continued.
60 Shroud skirt 0 Off 0 Boatfoiled
jlI
01
-I 2 I -2 I 2
I
-IO1 0
-I -21 .J e-2 A .-0- 2 ce gI
PL
-IOI I;
-IO/
I
-IOI
-16 -12 -8 -4 0 4 8 12 16 -16 -12 -8 -4 0 4 8 12 16 Angle of attack,a ,deg Angle of ottack,a,deg
(g) VD against a.
Figure 8.- Continued.
61 Shroud skirl 0 Off Boattailed
- 2 I --- -~ - .- I
0
2
I
u ~~ ~ -16 -12 -8 -4 0 4 8 12 16 -16 -12 -8 -4 0 4 8 12 16 Angle of attack,a,deg Angle of attack,a ,deg
(h) xcp/dref against a.
Figure 8.- Concluded.
62 Shroud skirt -
Boattailed - 2.4
C
W
--
Angle of attackp, deg (a) CN against u. Figure 9.- Effect of boattailed shroud skirt on longitudinal aerodynamic characteristics. Shroud nose 1; large fins; 6 = 0'. Q, w I I I I .oo I .2' M =c I I I I I
i
b
d1
i
ii I I/ I i M=I.O d 51 I -16 -12 -8 -4 0 c 04 12 16 Angle of attack,a,deg
(a) CN against a. Concluded.
Figure 9.- Continued.
64 Shroud skirt 0 Off 0 Boattailed
M=0.95
.8
.6
.4
II .2
M- 1.20 .o C -16 -12 - -4 0 8 12 16 - , -12 -8 -4 0 4 8 16 Angle of attack,a,deg Angle of attack,a ,deg
(b) CA against a.
Figure 9.- Continued.
65 .4
.2 Shroud skirt 0 Off Boattailed
.4 .4
.2 .2
0 ---- 0 -16 -12 -8 -4 0 4 8 12 16 -16 -12 -8 -4 0 4 8 12 16 Angle of attack,a,deg Angle of attackla ,deg
(c) CA,b against a.
Figure 9.- Continued. E v .8- Shroud skirt - .- .-0 Boattailed - rc
.- 8 .z -.a CL
-1.2 -16 -12 -8 -4 0 0 0 \O 4 8 12 16 Angle of OttaCk,a,deg - -_ 1.6 __- -7 Fv-----
(d) Cm against a.
Figure 9.- Continued. M.0.30 - M.0.50 M~0.70 __ M~0.80 h4.0.85 ____-- . -2.8-16 -12 -8 -4 0 0 0 0 0 4 8 12 16 Angle of attack,a,deg
(e) CL against a.
Figure 9.- Continued. IIII I Shroud skirt 1 Off Boottoiled / I13 M=O.1 / 1 ‘I I II i
i d I1, i J
-1.21 d I
-1.61 f
I -2.8 6=1.00I -16 -12 I 0 0 12 16 Angle of attack,a,deg 1
(e) CL against a. Concluded.
Figure 9.- Continued.
69 Shroud skirt 0 Off Boattailed I.4
I.2
I.o
.8
.6
.4
.2
0
1.2
1.0
.8
.6
.4
.2
n" -16 -12 -8 -4 0 4 8 12 16 -16 -12 -8 -4 0 4 12 Angle of attack,a,deg Angle of attack,a,deg
(f) CD against a.
Figure 9.- Continued.
70 1.6 c I.4
Shroud skirl I .2 0 Off Boattailed 1.6 1.0
I .4
I .2
I .o
.8 0
c 1.00 .6 - $ 2.4 .4 B 2.2 0 .2 P *. 2.0 gcM-0.90 E I.€ 1.8 8 1.E 1.6 i n f I .4 1.4 s / 3 I .2 I .2
1.c 1.0
.8 .8
.6 .6
.4 .4
.2 .2
I c .20 C - C - - - -12 -8 -4 04 8 12 16 - -12 -8 -4 4 8 12 16 Angle of attack,a,deg Angle of attack,a,deg
(f) CD against a. Concluded.
Figure 9.- Continued.
71
I 3
2 3
I 2
0 I I -I I- 0 T I -2 -I .30 -3 I -2 _I 3 I/ 2 2
I I 0 t 0 -I -I
e -2 -2 J J ._0- .--0- fp -3 g3 P c s - 52 I . 0
-I 2 -2
~~ M.0.70 I -3 3 0 i
2 -I
I -2
0 I
-I 0 I I -2 'I -I I -2 I -216 -12 -8t -4 0 -16 -12 -8 -4 04812 16 Angle of attack,a ,deg Angle of attack,a,deg
(g) VD against a.
Figure 9.- Continued.
72 ! Shroud skirt M.0.30 0 0 Off 0 Boattailed
, I llllll M.0.50
I I
- " -16 -12 -8 -4 0 4 8 12 16 -16 -12 -8 -4 0 4 8 12 16 Angle of attack,a,deg Angle of attack,a ,deg
(h) xcp/dref against a.
Figure 9.- Concluded. w4 - ~~ I Ill -2.0 M.0.30 M-b.50 M.0.70 M =0.80 M.0.85 -16 -12 -8 -4 0 0 0 0 0 4 8 12 16 Angle of attack,a,deg
(a) CN against a. Figure 10.- Effect of boattailed shroud skirt on longitudinal aerodynamic characteristics. Shroud nose 2; small fins; 6 = 00. Angle of attack,a ,deg
(a) CN against a. Concluded.
Figure 10.- Continued. Shroud skirt 0 Off I .4 .o I I Boattailed
3 .2 .8
0 .6 .4 .4 .2 .2 .O
i1 _L
I.4
I.2 I 1.0 1.0 I I .8 .8 1 .6 .6 4 .4 I 1 .2 .2 I M=1.20 n 0 I 1 v - ; -12 -8 -4 4 8 12 16 -16 -12 -8 -4 0 4 8 12 16 0 Angle of attock,a,deg Angle of attack,a,deg
(b) CA against a.
Figure 10.- Continued.
76 .2 Shroud skirt 0 Off .4' ' I ' 1 ' I 0 Boattailed 0 .2 .2 0 0
C -0, .2 .-0 .2 rc.c aJ 80 0 Q) 2 .4 I
.2
0
$2 .2 n- 0 -16 -12 -8 -4 0 4 8 I2 I6 -I6 -12 -8 -4 0 4 8 12 16 Angle of attack,a ,deg Angle of attack,a,deg
(c) CA,b against a.
Figure 10.- Continued. -1.2 M.0.30 M.b.50 - , , , M.0.70 1 , , , , -16 -12 -8 -4 0 0 0 0 4 8 12 16 Angle of attack,a ,deg
Angle of attack,a,deg
(d) Cm against a.
Figure 10.- Continued. " / / / / -1.2 d kJ E -/ L
(e) CL against a.
Figure 10.- Continued. 00 0
---- -16 -12 -8 -4 0 0 0 0 4 8 12 16 Angle of attack,a,deg
(e) CL against a. Concluded.
Figure 10.- Continued. 1
b
1.6 1
I.4
1.2
1.0
08 U -c
0
2.0
I .8
16
14
I2 1.2
1.c 1.0
.8 .8
.6 .6
.4 .4
1 .2
M=0.90 M.1.20 C I -12 -8 -4 0 4 8 12 16 o1 4 -16 -12 -8 - 16 Angle of attack,. ,deg Angle of attack,a,deg
.I (f) CD against a. ; Figure 10.- Continued.
81 II II I I I I1 I1 I1 I Im-Im I1111 111-1-
3
2
I Shroud skirt 0 Off o Boattoiled 0 2
-I I
-2 0
-I
-2
B c 'c JI
0
-I
-2
I . LL -I mh=e f_l i t.Ll. 1 -16 -12 -8 -4 0 4 8 I 12 16 Angle of ottack,a ,deg Angle of attack,a,deg
(g) VD against a.
Figure 10.- Continued.
82 Shroud skirt 0 Off 0 Boottoiled
3LL2 U
I/ 0 M!O 2 41 I 3 0 2 lo111
I 2 M-0.95 dref O
3
2
I I IIIII 0 1 0 It
U -16 -12 -8 -4 0 4 8 12 16 :I6 -12 -8 -4 0 4 8 12 16 Angle of ottock,a ,deg Angle of ottack,a,deg
(h) Xcp/dref against a.
Figure 10.- Concluded.
83 (a) CN against a.
Figure 11.- Effect of deflection of trailing-edge controls on longitudinal aerodynamic characteristics and on hinge moments of control on fin 2. Shroud nose 1; small fins; shroud skirt off. (Flagged symbols indicate points from repeat run.) u
I
Angle of attack,a,deg
(a) CN against a. Continued.
Figure 11.- Continued. Ill Ill lIllllllllI 11l111111lllll1l I I I I1
oc 1.2 8, deg 0 P P -10 /” Fins off P p A 7 >./i
/6 r- F d
M= -4 0 0 4 8 12 16 Angle of attack,a,deg
(a) CN against a. Concluded.
Figure 11.- Continued.
86 1.2 I
I.o
.8
.6
A
.2
oh I.o
.8 i
.6
.4
7
.6
A
.2
ol M=1.00 I.4
I.2 z 1.0
.8
.6
.4 .4
.2 -4 04 12 16 o1 M=1.20 I 0 -16 -12 -16 -12 -8 -4 4 8 12 16 II"Angle of Iattack,a ,deg 1J i_ Angle of attack,a,deg (b) CA against a. t Figure 11.- Continued.
87 .4 .4
.2 .2
0 0
4 4
.2 .2
0 0 .4 .4
.2 .2
0 0 -16 -12 -8 -4 0 4 8 12 16 -16 -12 -8 -4 0 4 8 12 16 Angle of attack,o,deg Angle of attack,a ,deg
(c) CA,b against a.
Figure 11.- Continued. I 2.4 ___ -
-16 -12 -8 -4 0 0 0 0 0 0 4 8 12 16 Angle of ottock,a,deg
(d) Cm against a. Figure 11.- Continued.
03 W co 0
I -2.8 , ,M=O.90 M.0.95 , ,M=b.95 M~0.97. M.1.00 M='1.20 - -16 -12 -8 -4 0 0 0 0 0 0 4 8 12 16 Angle of attack,a ,deg
(d) Cm against a. Concluded.
Figure 11.- Continued. -.- Ill I I I
(e) CL against a.
Figure 11.- Continued. Ms0.90 M.b.93 Ms0.95 M =6.97 M='I.00 , ,M=1.20 __ -2.0 ______-__ -16 -12 -8 -4 0 0 0 0 0 0 4 8 12 16 Angle of attack,a,deg
(e) CL against a. Concluded.
Figure 11.- Continued. 8, deg 00 [7 -10 0 Fins off .8
.6
.4 J
.2
M=0.85 0 .;: n I 0 .o f .2 .-a, 0 .8 E M.0.70 g 0 n B V e0 .8 0 .6
.4
M=O. .2 C I .2
1.0 1.0 c \ .8 .8 c .6 .6
.4 .4
.2
-16 M-0.83-12 -8 -4 0 0 url=0.93 4 4812 16 216 -12 -8 -4 12 16 IllAngle 1ottac k,a, deg 2Angle of attack,a,2 deg (f) CD against a. Figure 11.- Continued.
93 I.4 8, de9 00 0 -10 I.2 0 Fins off C C 1.0 c / .8
.6 / .4
.2
oj M=I.(
m P 1.4 I n , , , 3, 'r
C
.4
III.2
.20~ o] M-0.97 - lrvl -16 -12 -4 0 4 8 12 16 36 -12 - 04 8 12 Angle of attack,a,deg Angle of attack,a ,deg
(f) CD against a. Concluded.
Figure 11.- Continued.
94 Ill 8, deg 00 0 -10 0 Fins off
-2 -'i c
2
I
01
-I
-2
2
I 2
LID ,I L/D I
-I
-IOI I
2
I
0
-I
-2 2
I
-2 -8 -4 0 4 8 12 I6 -I 6 -I 2 A$ -16 -12 -8 -4 0 4 8 12 16 Angle of attock,a,deg Angle of attock,a,deg
(g) VD against a.
Figure 11.- Continued.
95 .. , ...... -
dref 5
4
3
2
I
0 5
4 3 .$ 2
I
12 0 16 -16 -12 -8 -4 0 4 16 Angle of attack,a, deg I Angle of attack,a,deg
(h) xCp,bref against a.
Figure 11.- Continued.
96 .2 .2
0 0
-.2 -.2
b 0 .2 9 0
L E- 0
72
.2
0 0
-3 -.2 -. L -16 -12 -8 -4 0 4 8 12 16 -16 -12 -8' -4 0 4 8 12 16 Angle of attack,a,deg Angle of attack,a,deg
(i) ch,2 against a.
Figure 11.- Concluded. I , -2.8 ~- M80.30 M.0.50 M-0.70 M-0.80 M80.85 -16 -12 -8 -4 0 0 0 0 0 4 8 12 16 Angle of attack,a,deg
(a) CN against a.
Figure 12.- Effect of deflection of trailing-edge controls on longitudinal aerodynamic characteristics. Shroud nose 1; intermediate fins; shroud skirt off. (Flagged symbols indicate points from repeat run.) 2.8
2.4
2.0
I .6
z 1.2
2-C 0, .-'3 .8 *- u-0, 8 0) .4 Y P 10 E 8 -.4
-.8
-1.2
-1.6
-2.0 -2.48 6o,, 11 I I I/ , 1 , I I 2 // 1 i I I/ I IIIII -2.8 I I I M.0.90 d M.0.95 6 M~1.00 M=1.20 -16 -12 -8 -4 0 0 0 0 4 8 12 16 Angle of attack,a, deg
(a) CN against a. Concluded.
(0 Figure 12.- Continued. (0 .. .6 1.2 .4 1.0 .2 .8 0 .6
.4
.2
0
2.0
I.8
1.6
C I I.4 - i
0 12 16 -16 -12 -8 -4 4 6 Angle of aiiack,o,deg
Ib) CA against a.
Figure 12.- Continued.
100 4 deg 00 0 -10 0 -20 .4 ' A Fins off I
W v) 0 m m .2 .2
0 0
.1 I I
Ill I .2
nv " -I6 -12 -8 -4 0 4 8 12 16 -16 -12 -8 -4 0 4 8 12 16 Angle of attack,a,deg Angle of attack,a,deg
(c) CA,b against a.
Figure 12.- Continued. id ,,,,I /, ,'! I I I' I I
I I I
-2.8 . M.b.30 Md.50 , , , .M-6.70 M.0.00 -16 -12 -8 -4 0 0 0 0 4 8 12 16 Angle of attack,a ,deg
(d) Cm against a.
Figure 12.- Continued. --
E
-16 -12 -8 -4 0 0 0 0 0 4 8 12 16 Angle of attack,a,deg
(d) Cm against a. Concluded.
Figure 12.- Continued. I I I I I -2.8 - ,M-0.30 , , , .M=0.50 ,M=0.70 M.0.80 M.0.05 -16 -12 -8 -4 0 0 0 0 0 4 8 12 16 Angle of ottack,a,deg
le) CL against a.
Figure 12.- Continued. --
L.U Ill I I I I
Angle of attock,u,deg
(e) CL against a. Concluded.
Figure 12.- Continued. .-(u
/,,ll,lll,!I !',;,!I1 , I! OiM=0.50 ' I I I M.0.80 1 1 11 1 -16 -12 -8 -4 0 4 8 12 16 36 -12 '-8 -4 ' b ' 4 ' 8 ' I2 16 Angle of attack,a ,deg
(f) CD against a.
Figure 12.- Continued. 2.2
2.0
I .8
1.6
M.0.85 1.4
1.6 1.2 < i 1.4 :L 1.0
I.2 .8 \ 1.0 .6 \
.8 .4
1.8 I .t
1.6 1.Z
I .4 I./
1.2 I.i
1.0 1.c
.8 .e
.6 .6
I .4 .4I 0 .2 0.95
OLN ~ 8 12 16 0 -16 -12 iitl3 -16 --12 -8 -4 0 12 16 IAngle of attack,a,deg :i Angle of attock,a,deg (f) CD against a. Concluded,
Figure 12.- Continued.
107 3
2 3
I 2
0 I
-I 0
-2 -I
-3 -2 3 2 :_1 2 I :: I 0 0 -I -I -2 0 :-2 e2 .- .-0- 2- c + e -3 210 "0 2 73 U 22 io.- J 1
I -I
0 -2 2 -I I -2 0 -3 -I 2 -2 I 2
0 I
-I 0
-2 -I
-3 -2 -16 -12 -8 -4 0 48 16 -16 -12 -8 -4 0 4 8 12 16 Angle of attack,a ,deg Angle of attock,a,deg lg)I VD against a. Figure 12.- Continued.
108 45l 8, deg 00 4 Finsoff 3I21
4 0‘t I 3
2I
I
1 0- 6
5
413
dref
5
4
3
2
‘I ,I M-
5~ 413l
I ,I ,I 2l M= -16 -12 -8 -4 0 4 8 16 -16 -12 - -4 0 4 8 12 16 Angle of attack,a ,deg Angle of attack,a,deg
(h) kp/dref against a.
Figure 12.- Concluded.
109 3.2
2.8
2A
2.0
I .6
I .2
2 y. .8 .-c .-0 r ..4 W 8 "0 .+ - 2 -.4 L z0 -.8
-1.2
-1.6
-2.0
-2.4
-2.8
-3.2
-3.6 -I Angle of attack,o,deg
(a) CN against a.
Figure 13.- Effect of deflection of trailing-edge controls on longitudinal aerodynamic characteristics. Shroud nose 1; large fins; shroud skirt off. (Flagged symbols indicate points from repeat run.) 3.2
2.8
2.4
2.0
I.6
I .2
z v- .8 + .-C $ .4 aJ 8 aJ0 ? L - -.4 6 z -. 8
-1.2
-1.6
-2.0
-2.4
-2.8
-3.2
-3.6 -16 -12 -8 -4 0 0 0 0 4 8 12 16 Angle of attack,a ,deg
(a) CN against a. Concluded.
Figure 13.- Continued. 4 deg 50 0 -5 0 -10 A -20 n -30 D Fins off 1.2
I.o
.8
.6
A
.2
0
.8 1.8 .6 1.6 .4 I.4 .2 I.2 0 I.o 1.0
.8 .8
.6 .6
.4 .4
.2 .2
0 0 -16 -12 -8 -4 0 4 8 12 16 -16 -12 -8 -4 0 4 8 12 16 Angle of attack,= ,deg Angle of attock,a,deg
(b) CA against a.
Figure 13.- Continued.
112
-...... _...... I n -50 b Fins off .2 -7
0 .2
0 2
0 .. I I
.2
0 a, I-1 v) .7 I cz
.2 .2
0 0 .. .. I
.2 .2 n- -16 -12 -8 -4 0 4 8 12 16 Anale of attack,a .dea
(c) CA,b against a.
Figure 13.- Continued. -2.0 - ,-
-2.4 c_
I -2.8 M4.30 - M.b.50 M.0.70 -16 -12 -8 -4 0 0 0 4 8 12 16 Angle of attack,a,deg
(d) C, against a.
Figure 13.- Continued. .-+ Q
-16 -12 -8 -4 0 0 0 4 8 12 16 Angle of ottock,a ,deg
(d) Cm against a. Continued.
Figure 13.- Continued. (d) Cm against a. Concluded.
Figure 13.- Continued. (e) CL against a.
Figure 13.- Continued. -3.2 ,M=0.05 - Ma0.90 ,M=0.95 , , I , I , I ML1.00 , 1 , , , , M-1.20 -_- -16 -12 -8 -4 0 0 0 0 0 4 8 12 16 Angle of attack,a,deg
(e) CL against a. Concluded.
Figure 13.- Continued. 4 deg 00 1.2 0 -5 .." 4 0 -10 A -20 .8
.6
.A-1 h I
.2
n 00 +- c .a, 1.4
aJ aJ 8 1.2 I s 1.2 0 0 0 i L3 nE! 1.0 1.0
.8 .8
.6 .6
.4 .4
.2 .2 ' lllll! ll!l I1 0 M.0.50) ) 1 I " -16 -12 -8 -4 0 4 8 12 16 -16 -12 -8 -4 0 4 8 12 16 Angle of attaCk,a,deg Angle of attack,a,deg
(f) CD against a.
Figure 13.- Continued. I.4
I.2 0 -10
1.0 2.0
.8 I.8
.6 1.6
.4 I.4
.2 1.2
0 1.0
1.4 .8
1.2 .6
1.0 .4
0 0 O. .8 :.2 ._ 5 00) { .6 %o 2.4 e .4 a 2.2 .2 2.0 0 I.8
1.6 I I 1.4 B 1.2 1 1.0 I .8
.6
.4
.2
12 0 -16 -12 -8 -4 0 4 8 12 16 -16 -12 - -4 0 4 Angle of attack,a ,deg Angle of attack,a,deg 1 (f) CD against a. Concluded.
Figure 13.- Continued.
120 OlI I _II I -I C
-2 -I I -3. -2
2
I
0
-I c
:n -2 P 9- 2 t 53 e e m +I .- .-Lr zJ -I 0
-I 0Ii -2 - -I 2
-2 I
-3 0
3~ -I 2 -2 I "i -I
-2
-3-I 6 -12 -8 -4 0 4 8 12 16 36 -12 -8 -4 0 4 8 12 16 Angle of attock,a,deg Angle of attack,a ,deg
(g) VD against a.
Figure 13.- Continued.
121 IIIII
5
4-
3-
2-
I-
0- 5 5
4 4
3 3
2 2
I I
0 0 -16 -12 -8 -4 0 4 8 12 16 -16 -12 -8 -4 0 4 8 12 16 Angle of attack,a,deg Angle of attack,a,deg
(h) xcp/dref against a.
Figure 13.- Concluded.
122 -2.0 , I lllll I Ill I -2.4 I 1 M=0.30 1 /Ill1 M=0.50 I I 1 1 I M.0.70 M.C.80 1 -16 -12 -8 -4 0 0 0 0 4 8 12 16 Angle of attack,a,deg
(a) CN against a.
Figure 14.- Effect of deflection of trailing-edge controls and transition strips on longitudinal aerodynamic characteristics and on hinge moments of controls on fin 2. Shroud nose 2; small fins; shroud skirt off. I 1 -2.4 - M-0.851 , M=0.9a M.d.951 , , ' M=[.OCJ I.,, M-1.20 -16 -12 -8 -4 0 0 0 0 0 4 8 12 16 Angle of attack,a,deg
(a) CN against a. Concluded.
Figure 14.- Continued. 8..- des Transition 00 Off 0 -10 Off Off -1 I 1 .4
.2
3+- 0-
a I I .4
1.2
1.0 0 .8
.6 I .4 I .4 I I .2 I b=0.90 0- 0L! -16 -12 -8 -4 0 4 I 12 16 -16 - -4 16 Angle of attack,a ,deg Angle of attack,a,deg
(b) against a. CASI - Figure 14.- Continued. ::
125
I- I
6, deg Transition 00 Off I17 -10 Off .4
.2
0
a- ", tc .-a, .-V + W 0 0 a,
w- -I 0 .-x Q W .4 v) m0 .2 .4 0 .2 .2 IIIIIIIIIIIIIIIII M=1.20 0 n -16 -12 -8 -4 0 4 8 12 16 -16 -12 -8 -4 0 4 8 12 16 Angle of attack,a.dea Angle of attack,a,deg
(c) CA,b against a. Figure 14.- Continued. IIIIII 6, deg Transition 2.4- - 0 -10 Off ___ -______g--_p 0 Fins off Off ,c> / off 2.0- Fins On 0 Y /
- / / P -d V / -1.6- d / d
-2.0, , i , I" I I I,,, IIIIII lIII1,l i/i M.0.70 1 1 1 I I 1 M.0.80 I I I 1 I I I M.0.85 I 1 j I 1 1 I
(d) Cm against a.
Figure 14.- Continued. Angle of attack,a,deg
(d) Cm against a. Concluded.
Figure 14.- Continued, (e) CL against a.
Figure 14.- Continued. w w 0
I I -2.0 M4.05 M-0.90 M.b.95 MJ1.00 , M.1.20' __-- -16 -12 -8 -4 0 0 0 0 0 4 El 12 16 Angle of attack,a,deg
(e) CL against a. Concluded.
Figure 14.- Continued. 8, deg Transition 1.0 00 0 -10 :if I 0 Fins off Off 6 Fins off On
.4
.2 :f0.95 I .e 1 1.6 c
C I.4
1.2 k
1.0
.8
n y- .6 .-aJ ._0 .&r .4 8 m e .2 n = 1.00 0- 0- M 2.c
I.€
I.€
I.4
I.i
1.c
.E
.6 .E
.4
.2 cl M.0.90 1.20 C ____M= -16 -12 - -4048 16 - 1 -12 -8 -4 0 4 8 Angle of attack,a,deg Angle of atfack,a,deg
(f) CD against a.
Figure 14.- Continued.
131 3
2
I
0 6, deg Transition 00 Off 0 -10 Off -I 0 Fins off Off dFins off On -2
2 -I 0 I tl-l
0
-I -2 0Ell -1 pJ I 2 -2 L/D I A 0 -I “H -I
-2 I
2 0
I -I
0 I
-I 0
-2 - -16 -12 -8 -4 0 4 8 12 16 116 -12 -8 -4 0 4 8 12 16 Angle of attock,a ,deg Angle of attack,a ,deg
(g) VD against a. Figure 14.- Continued.
132 t
8, deg Transition 5 00 Off 0 Fins off Off dFiffi off On 4 7~ 3
2 56i
I 4 N 0- 3 Q 4 !I -i 2 0:: 3 I
2 0- ow
I
M.0.50
4 i 3 I: I-- -XCP dref
I
M=O. 5 5 4 4 3 3 2 2 I I M=O
5
4 4
3 3
2 2
I II M.0.85 n/ M=I. 0 - -16 -12 -8 -4 0 4 8 12 16 -16 -12 -8 -4 0 4 8 Angle of attack,a,deg Angle of attock,a ,deg
(h) Xcp/dr,f against a.
Figure 14.- Continued.
133 6, deg Transition 00 0 -10 .4
.2
0
.L .L .4 .4 .2 2 0 .2 t-!= 0--I .-W .-u I .+.I- 0 - M.0.50 W '1 0 0 L + W c W -.2 .2 c E5 W 0 .c_ 0 I
-2 M.0.70
.2
0
72 ~ ~~ -16 -12 -8 -4 0 4 8 12 16 :I6 -12 -8 -4 0 4 8 12 16 Angle of attack,a,deg Angle of attackla ,deg
(i) ch,2 against a.
Figure 14.- Concluded. OIII Shroud nose 0 1 2 8 3 #.I6
-12 I If .2 .3 .4 -5 -6 .7 .8 .9 1.0 I. I 1.2 1.3 Mach number,M i I
3' ' I .o I II 1.3 .2 -3 .4 .5 -6 .7 .8 .9 1.1 I .2 Mach number, M \ i Figure 15.- Summary information on effect of shroud nose on longitudinal aerodynamic parameters. Shroud skirt off; fins off; a = Oo. (Flagged symbols indicate points from repeat run.)
13 5 Mach number, M
Figure 16.- Summary information on effect of cylindrical and boattailed shroud skirts on longitudinal aerodynamic parameters. Shroud nose 1; fins off; a = Oo. . (Flagged symbols indicate points from repeat run.)
I Shroud skirt =It= I I .7 .8 -9 1.0 I.I 1.2 1,3 Mach number, M
i
.7 .8 .9 I .o 1.1 1.2 1.3 Mach number ,M xcp d ref
.6 .7 -8 -9 I.o I .I 1.2 1.3 Much mmber,M
Figure 17.- Summary information on effect of boattailed shroud skirt on longitudinal aerodynamic parameters. Shroud nose 1; small fins; 6 = 00; a = 00.
137 -24
.20 Boattailed
.I6
.I21.2 .3 .4 .5 .6 .7 .8 .9 1.0 1.1 1.2 1.3 Mach number, M
.I 2
.08 c"a .04
1 1 1 1 1 1 1 1 1
0 .2 .3 .4 .5 .6 .7 .8 .9 1.0 1.1 1.2 1.3 Moch number M
Mach number M
Figure 18.- Summary information on effect of boattailed shroud skirt on longitudinal aerodynamic parameters. Shroud nose 1; large fins; 6 = Oo; a = Oo.
138 .8 .9 I -0 1.1 1.2 1.3 Mach number ,M
.6 .7 .8 -9 1.0 1.1 1.2 1.3 Mach wmber,M
xcp dref
12 .3 .4 .5 .6 -7 -8 .9 1.0 1.1 1.2 1.3 Mach number,M
Figure 19.- Summary information on effect of boattailed shroud skirt on longitudinal aerodynamic parameters. Shroud nose 2; small fins; 6 = 00; a = @.
139 Mach number,M
6
5
4 xCP d ref 3
2
I -2 .3 .4 .5 .6 Mach number,M
(a) Cmcl and xcp/dref against M. 6 = 00.
Figure 20.- Summary information on effect of fin size on longitudinal aerodynamic parameters. Shroud nose 1; shroud skirt off; a = 00. (Flagged symbols indicate points from repeat run.)
140 Fins 0 Small 0 intermediate -.o 2
-.03
‘,8 -.04
-.O 5
-.06 Mach number, M
.05
.04
‘Ng -03
.02
.o I 3 Mach number, M
Ibl Cm6 and CN~against M.
Figure 20.- Concluded.
141
I, .% .9 1.0 1.1 1.2 1.3 Mach number ,M
.3 .4 .5 .6 .7 -8 .9 1.0 1.1 1.2 1.3 Mach number,M
(a) Cma and Xcp/dref against M.
Figure 21.- Summary information on effect of deflection of trailing-edge controls on longitudinal aerodynamic parameters. Shroud nose 1; small fins; shroud skirt off; a = 00. (Flagged symbols indicate points from repeat run.)
142 (N r" 0 --. --. t" t 0 0 ------L-l- .-0) 0 rc.- Y- Q) 0 0 t t Q) E 0 E I Q) 0 C .3
Mach number ,M
Mach number ,M
(b) ch.2 and chn against M.
Figure 21.- Continued. -. 02 I ‘mg -.03 -4
704 .2 .3 .4 .5 .6 .7 .8 -9 1.0 Ill I.2 3 Mach number,M
I.3 Mach number, M
.5 .6 .7 .8 .9 1.0 1.1 I .2 I.3 Mach number, M
1.3 Mach number ,M
(c) chb. cN~Cmg, and cm6/chg against M.
Figure 21.- Concluded.
144 .24~ 4deg 0 0 i .2oF 0 -10 0 -20 IA Fins off --fHA -12 T- I -08 i -04 +' TI .3 .4 .5 .6 -7 .8 .9 1.0 1.1 1.2 1.3 Mach number, M
4- -,
i
.5 I .7 -8 .9 1.0 1.1 1.2 1.3 Mach number, M
(a) Cm, and Xcp/dref against M. Figure 22.- Summary information on effect of deflection of trailing-edge controls on longitudinal aerodynamic parameters. Shroud nose 1; intermediate fins; shroud skirt off; u = OO. (Flagged symbols indicate points from repeat run.)
145 -.O 4
-.O 5
- 06 .2 .4 .5 .6 .7 .8 .9 I.o 1.1 1.2 1.3 Mach number, M
(b) Cmg and CN~against M.
Figure 22.- Concluded. weg -24 - 00 - 0 -5 0 -10 .20 -- A -20 i .- -30 b Fins off "1
4
OL- .2 .4 .5 .9 1.0 Mach number,M
E ._
F c
/- /.I 4 f- xCP dref t
.. 2 - 1, 7-I.o .9 3 Mach number,M
(a) Cma and X,p/dref against M.
Figure 23.- Summary information on effect of deflection of trailing-edge controls on longitudinal aerodynamic parameters. Shroud nose 1; large fins; shroud skirt off; a = Oo. (Flagged symbols indicate points from repeat run.)
147
I Mach number,M
(b) cmb and CNb against M.
Figure 23.- Concluded. -26 ... 8 ,deg Transit ion 1, 0 0 Off r .22. 0 -10 Off & 0 Fins off Off TI\ '. l- 6 Fins off On .I 8 d> II % cmc, \ .I4 II i I I -44- \ i I d d c1 - 1
.3 .4 .5 .7 .8 I .o I.I I1.2 Mach number,M
f \ \\
<
\ \
i
-3 .4 -5 .6 -7 .8 -9 I 1.1 1.2 1.3 Mach number,M
(a) Chand Xcp/dref against M.
Figure 24.- Summary information on effect of deflection of trailing-edge controls on longitudinal aerodynamic parameters. Shroud nose 2; small fins; shroud skirt off; a = Oo.
149 cu r" 0 t"c .-a, .-0 ct r(- a, 0 0 tc E 0 E I va, .-c I Mach number, M
C ha
~~
(b) ch,2 and ch, against M.
Figure 24.- Continued. .3 .4 .5 .6 .7 -8 -9 I .o I. I 1.2 1.3 Mach number,M 9-
!I .5 .8 -9 Mach number, M L-.I rI I
.5 .6 1.0 1.1 1.2 1.3 Mach number, M i I I -A- I I I +I I I I I .4 .5 .6 1.0 Mach number, M
(c) chp CN~Cmp and Cmb/chh against M.
Figure 24.- Concluded.
151 M.0.80 ; Q =O" M=0.90;a=Oo
M=0.80;~=2" M=0.90; a = 2"
M.0.80; 0,-4"
(a) With shroud nose 1. L-67-964
Figure 25.- Schlieren photographs of airflow about LM aerodynamic shroud with shroud noses 1 and 2. Transition strips off.
152 M=0.90;a= 4"
M.0.90 :a = 6O
M=O.gO;a =o0 M.0.90; a =8'
(b) With shroud nose 2. L-67-965
Figure 25.- Concluded.
i NASA-Langley, 1967 ~ 31 L-4344 153 “The aeronautical and space activities of the United States shall be conducted 50 a.r to contribute . . . to the expansion of human Knowl- edge of phenomena in the atmoJphere and space. The Administration shall provide f or the widest practicable and appropriate dissemination of information concerning its activities and the results thereof .”
-NATIONALAERONAUTICS AND SPACBAn OF 1958
NASA SCIENTIFIC AND TECHNICAL PUBLICATIONS
TECHNICAL REPORTS: Scientific and technical information considered important, complete, and a lasting contribution to existing knowledge. TECHNICAL NOTES: Information less broad in scope but nevertheless of importance as a contribution to existing knowledge. TECHNICAL MEMORANDUMS: Information receiving limited distribu- tion because of preliminary data, security classification, or other reasons. CONTRACTOR REPORTS: Scientific and technical information generated under a NASA contract or grant,and considered an important contribution to existing knowledge. TECHNICAL TRANSLATIONS: Information published in a foreign language considered to merit NASA distribution in English. SPECIAL PUBLICATIONS: Information derived from or of value to NASA activities. Publications include conference proceedings, monographs, data compilations, handbooks, sourcebooks, and special bibliographies.
TECHNOLOGY UTILIZATION PUBLICATIONS: Information on tech- nology used by NASA that may be of particular interest in commercial and other non-aerospace applications. Publications include Tech Briefs, Technology Utilization Reports and Notes, and Technology Surveys.
QI,I+,,
Details on the availability of these publications may be obtained from:
SCIENTIFIC AND TECHNICAL INFORMATION DIVISION NATIONAL AERONAUTICS AND SPACE ADMINISTRATION
Washington, D.C. 20546