ALBERT

Description: In the design of a combustor, information is necessary for the mixing of the fuel and air in order to determine the optimum combustor length. In scramjet combustors the mixing often takes place in a shear layer that is formed between the fuel and air. This research was an experimental study of shear layers in supersonic flows aimed at determining what mechanisms affect the shear layer so that the mixing could be better predicted. A second goal was to provide sufficient instream information for use in checking existing Computational Fluid Dynamic (CFD) codes. The shear layer between a supersonic two-dimensional air stream (M = 2 or M = 3) was mixed with a near sonic two-dimensional air stream (M = 1.2). Instream measurements of pitot pressure and cone static pressure were used to determine mean velocity profiles at various axial locations. These velocity profiles were used to determine the shear layer spreading rate and are compared with various predictions. Wall measurements of static pressure, temperature and skin friction were also taken and are presented. The instream measurements were also used for comparison with an existing CFD code. The upstream velocity, pressure and temperature profiles were used as a starting profile and the code was used to calculate downstream profiles for comparison with the experimental results. Reasonable agreement between the measured and calculated results was obtained.

Description: This paper presents a set of second-order closure models for low-Reynolds-number turbulence near the wall. Existing closure models for the Reynolds-stress equations were modified to show proper near-wall behavior. A dissipation-rate equation for the turbulent kinetic energy is also reformulated. The proposed models satisfy realizability and will not produce unphysical behavior. Fully developed channel flows are used for model testing. The equations are solved for the mean velocity, the Reynolds stresses, and the dissipation rate of the turbulent kinetic energy. The calculations are compared with both direct numerical simulations and with measurements. It is shown that the present models perform well in predicting the behavior of the turbulence near a wall. Significant improvements over previous models in predicting the components of the Reynolds stress tensor are obtained in the present models.

Description: The XV-15 Tilt-Rotor wing has six major aeroelastic modes that are close in frequency. To precisely excite individual modes during flight test, dual flaperon exciters with automatic frequency-sweep controls were installed. The resulting structural data were analyzed in the frequency domain (Fourier-transformed). Modal frequencies and damping were determined by performing curve fits to frequency-response magnitude and phase data. Results are given for the XV-15 with its original metal rotor blades. Frequency and damping values are also compared with new predictions by two different programs, CAMRAD and ASAP.

Description: The airplane pictured is the new Air Shark I, a four-place amphibian that makes extensive use of composite materials and cruises at close to 200 miles per hour under power from a 200-horsepower engine. Air Shark I is a "homebuilt" airplane, assembled from a kit of parts and components furnished by Freedom Master Corporation, Satellite Beach, Florida. The airplane incorporates considerable NASA technology and its construction benefited from research assistance provided by Kennedy Space Center (KSC) In designing the Shark, company president Arthur M. Lueck was able to draw on NASA's aeronautical technology bank through KSC's computerized "recon" library. As a result of his work at KSC, the wing of the Air Shark I is a new airfoil developed by Langley Research Center for light aircraft. In addition, Lueck opted for NASA-developed "winglets," vertical extensions of the wing that reduce drag by smoothing air turbulence at the wingtips. The NASA technology bank also contributed to the hull design. Lueck is considering application of NASA laminar flow technology-means of smoothing the airflow over wing and fuselage-to later models for further improvement of the Shark's aerodynamic efficiency. A materials engineer, Lueck employed his own expertise in designing and selecting the materials for the composite segments, which include all structural members, exposed surfaces and many control components. The materials are fiber reinforced plastics, or FRP They offer a high strength-to-weight ratio, with a nominal strength rating about one and a half times that of structural steel. They provide other advantages: the materials can be easily molded into finished shapes without expensive tooling or machining, and they are highly corrosion resistant. The first homebuilt to be offered by Freedom Master, Air Shark I completed air and water testing in mid-1985 and the company launched production of kits.

Description: Hi-Tech Inc., a company which manufactures water jetting equipment, needed a high pressure rotating swivel, but found that available hardware for the system was unsatisfactory. They were assisted by Marshall, which had developed water jetting technology to clean the Space Shuttles. The result was a completely automatic water jetting system which cuts rock and granite and removes concrete. Labor costs have been reduced; dust is suppressed and production has been increased.

Description: Composites are lighter and stronger than metals. Aramid fibers like Kevlar and Nomex were developed by DuPont Corporation and can be combined in a honeycomb structure which can give an airplane a light, tough structure. Composites can be molded into many aerodynamic shapes eliminating rivets and fasteners. Langley Research Center has tested composites for both aerospace and non-aerospace applications. They are also used in boat hulls, military shelters, etc.

Description: Stars and Stripes racing yacht brought the American's Cup back to the United States. Originating from NASA's drag reduction technology, the boats "secret weapon" was that the hull's underside was coated with riblets. Riblets are small, barely visible grooves on the surface of an airplane intended to reduce skin friction by smoothing the turbulent airflow next to the skin. Grooves are V-shaped with the angle pointing in the direction of the airflow. No deeper than a scratch, they have a pronounced beneficial influence on air turbulence. *No longer commercially available.

Description: Neico Aviation, Inc.'s tubular fuselage frame of Prescott Pusher is heliarc welded at the factory. Wings, tail, and control surfaces are of traditional aluminum. A rear mounted 180 horsepower engine with a four bladed pusher propeller allows speeds up to 200 miles per hour; an advanced rotary engine and other powerplants being tested are expected to boost cruise speed above 200 miles per hour.

Description: Macrodyne, Inc.'s laser velocimeter (LV) is a system used in wind tunnel testing of aircraft, missiles and spacecraft employing electro optical techniques to probe the flow field as the tunnel blows air over a model of flight vehicle and to determine velocity of air and its direction at many points around the model. However, current state-of-the-art minicomputers cannot handle the massive flow of real time data from several sources simultaneously. Langley developed instrument Laser Velocimeter Autocovariance Buffer Interface (LVABI). LVABI is interconnecting instrument between LV and computer. It acquires data from as many as six LV channels at high real time data rates, stores it in memory and sends it to computer on command. LVABI has application in variety of research, industrial and defense functions requiring precise flow measurement.

Description: Michael E. Fisher, President of AeroVisions International, has introduced the Culex light twin engine aircraft which offers economy of operation of a single engine plane, the ability to fly well on one engine, plus the capability of flying from short, unimproved fields of takeoff and landing distances less than 35 feet. Key element of design is an airfoil developed by Langley. Culex was originally intended to be factory built aircraft for special utility markets. However, it is now offered as a build-it-yourself kit plane.

Description: An approximate method for development of flow and thermal boundary layers in laminar regime on cylinders with arbitrary cross section and transpiration-cooled walls is obtained by use of Karman's integrated momentum equation and an analogous heat-flow equation. Incompressible flow with constant property values throughout boundary layer is assumed. Shape parameters for approximated velocity and temperature profiles and functions necessary for solution of boundary-layer equations are presented as charts, reducing calculations to a minimum. The method is applied to determine local heat-transfer coefficients and surface temperature-cooled turbine blades for a given flow rate. Coolant flow distributions necessary for maintaining uniform blade temperatures are also determined.

Description: The presence of radomes and instruments that are sensitive to water films or ice formations in the nose section of all-weather aircraft and missiles necessitates a knowledge of the droplet impingement characteristics of bodies of revolution. Because it is possible to approximate many of these bodies with an ellipsoid of revolution, droplet trajectories about an ellipsoid of revolution with a fineness ratio of 10 were computed for incompressible axisymmetric air flow. From the computed droplet trajectories, the following impingement characteristics of the ellipsoid surface were obtained and are presented in terms of dimensionless parameters: (1) total rate of water impingement, (2) extent of droplet impingement zone, and (3) local rate of water impingement. These impingement characteristics are compared briefly with those previously reported for an ellipsoid of revolution with a fineness ratio of 5.

Description: An investigation at a free-stream Mach number of 2.02 was made to determine the effects of a propulsive jet on a wing surface located in the vicinity of a choked convergent nozzle. Static-pressure surveys were made on a flat surface that was located in the vicinity of the propulsive jet. The nozzle was operated over a range of exit pressure ratios at different fixed vertical distances from the flat surface. Within the scope of this investigation, it was found that shock waves, formed in the external flow because of the presence of the propulsive jet, impinged on the flat surface and greatly altered the pressure distribution. An integration of this pressure distribution, with the location of the propulsive jet exit varied from 1.450 propulsive-jet exit diameters to 3.392 propulsive-jet exit diameters below the wing, resulted in an incremental lift for all jet locations that was equal to the gross thrust at an exit pressure ratio of 2.86. This incremental lift increased with increase in exit pressure ratio, but not so rapidly as the thrust increased, and was approximately constant at any given exit pressure ratio.

Description: Pressure-distribution measurements have been made on the fus elage of the Bell X- 1 research airplane. Data are presented for angles of attack from 2 deg. to 8 deg. during pull-ups at Mach numbers of about 0.78, 0.85, 0.88, and 1.02. The results of the investigation indicated that a large portion of the load carried by the fuselage was in the vicinity of the wing and may be attributed to wing-to-fuselage carryover. The presence of the wing from the 41 to 60 percent fuselage stations influenced the fuselage pressures from about 30 to 65 percent fuselage length at Mach numbers of approximat ely 0.78, 0.85, and 0.88, and from about 35 to 80 percent fuselage length at a Mach number of approximately 1.02. The fuselage contributed about 20 percent of the total airplane normal-force coefficient. The center of pressure of the fuselage load throughout the tests was located from 41 to 51 percent fuselage length, which corresponds to the forward half of the wing root-chord location.

Description: NACA model 11-C was tested according to the general method with the angle of afterbody keel set at five different angles from 2-1/2 degrees to 9 degrees, but without changing other features of the hull. The results of the tests are expressed in curves of test data and of non-dimensional coefficients. At the depth of step used in the tests, 3.3 percent beam, the smaller angles of afterbody keel give greater load-resistance ratios at the hump speed and smaller at high speed than the larger angles of afterbody keel. Comparisons are made of the load-resistance ratios at several other points in the speed range. The effect of variation of the angle of afterbody keel upon the take-off performance of a hypothetical flying boat of 15,000 pounds gross weight having a hull of model 11-C lines is calculated, and the calculations show that the craft with the largest of the angles of afterbody keel tested, 9 degrees, takes off in the least time and distance.

Description: Numerical solutions of the differential equation obtained from the momentum theorem for the development of a turbulent boundary layer along a thermally insulated surface in two-dimensional and in radial shock-free flow are presented in tabular form for a range of Mach numbers from 0.100 to 10. The solution can be used in a step-wise procedure with any given distribution of favorable pressure gradients and for zero pressure gradients. Solutions are also given for use with moderate adverse pressure gradients. The mean velocity in the boundary layer is approximated by a power-law profile. In view of the stepwise integration methods to be used, the exponent designated the profile shape can be varied along the surface between the integral fraction limits 1/5 and 1/11 through interpolation. Agreement obtained between theoretical and experimental boundary-layer development in a supersonic nozzle at a nominal Mach number of 2 indicates the general validity of the approximations used in the analysis - in particular, the method of extrapolating low-speed skin-friction relations to high Mach number flows. The extrapolation method used assumes that the skin-friction coefficient depend primarily on Reynolds number, provided that the density and the kinematic viscosity are evaluated at surface conditions.

Description: A method has been proposed for predicting the effect of a rapid blade-pitch increase on the thrust and induced-velocity response of a helicopter rotor. General equations have been derived for the ensuing motion of the helicopter. These equations yield time histories of thrust, induced velocity, and helicopter vertical velocity for given rates of blade-pitch-angle changes and given rotor-angular-velocity time histories. The results of the method have been compared with experimental results obtained with a rotor mounted on the Langley helicopter test tower. The calculated and experimental results are in good agreement, although, in general, the calculated thrust-coefficient overshoots are about 10 percent greater than those obtained experimentally.

Description: The condensation pressure of air was determined over the range of temperature from 60 to 85 K. The experimental results were slightly higher than the calculated values based on the ideal solution law. Heat of vaporization of oxygen was determined at four temperatures ranging from about 68 to 91 K and of nitrogen similarly at four temperatures ranging from 62 to 78 K.

Description: A cascade of 65-(12)10 compressor blades was tested at one geometric setting over a range of inlet Mach number from 0.12 to 0.89. Two groups of data are presented and compared: the first from the cascade operating conventionally with no boundary-layer control, and the second with the boundary layer controlled by a combination of upstream slot suction and porous-wall suction at the blade tips. A criterion for two-dimensionality was used to specify the degree of boundary-layer control by suction to be applied. The data are presented and an analysis is made to show the effect of Mach number on turning angle, blade wake, pressure distribution about the blade profile and static-pressure rise. The influence of boundary-layer control on these parameters as well as on the secondary losses is illustrated. A system of correlating the measured static-pressure rise through the cascade with the theoretical isentropic values is presented which gives good agreement with the data. The pressure distribution about the blade profile for an inlet Mach number of 0.21 is corrected with the Prandtl-Glauert, Karman-Tsien, and vector-mean velocity - contraction coefficient compressibility correction factors to inlet Mach numbers of 0.6 and 0.7. The resulting curves are compared with the experimental pressure distributions for inlet Mach numbers of 0.6 and 0.7 so that the validity of applying the three corrections can be evaluated.

Description: The empirical relation between the induced velocity, thrust, and rate of vertical descent of a helicopter rotor was calculated from wind tunnel force tests on four model rotors by the application of blade-element theory to the measured values of the thrust, torque, blade angle, and equivalent free-stream rate of descent. The model tests covered the useful range of C(sub t)/sigma(sub e) (where C(sub t) is the thrust coefficient and sigma(sub e) is the effective solidity) and the range of vertical descent from hovering to descent velocities slightly greater than those for autorotation. The three bladed models, each of which had an effective solidity of 0.05 and NACA 0015 blade airfoil sections, were as follows: (1) constant-chord, untwisted blades of 3-ft radius; (2) untwisted blades of 3-ft radius having a 3/1 taper; (3) constant-chord blades of 3-ft radius having a linear twist of 12 degrees (washout) from axis of rotation to tip; and (4) constant-chord, untwisted blades of 2-ft radius. Because of the incorporation of a correction for blade dynamic twist and the use of a method of measuring the approximate equivalent free-stream velocity, it is believed that the data obtained from this program are more applicable to free-flight calculations than the data from previous model tests.

Description: The autorotative performance of an assumed helicopter was studied to determine the effect of inoperative jet units located at the rotor-blade tip on the helicopter rate of descent. For a representative ramjet design, the effect of the jet drag is to increase the minimum rate of descent of the helicopter from about 1,OO feet per minute to 3,700 feet per minute when the rotor is operating at a tip speed of approximately 600 feet per second. The effect is less if the rotor operates at lower tip speeds, but the rotor kinetic energy and the stall margin available for the landing maneuver are then reduced. Power-off rates of descent of pulse-jet helicopters would be expected to be less than those of ramjet. helicopters because pulse jets of current design appear to have greater ratios of net power-on thrust to power-off, drag than currently designed rain jets. Iii order to obtain greater accuracy in studies of autorotative performance, calculations in'volving high power-off rates of descent should include the weight-supporting effect of the fuselage parasite-drag force and the fact that the rotor thrust does not equal the weight of the helicopter.

Description: The heat requirements for the icing protection of two radome configurations have been studied over a range of design icing conditions. Both the protection limits of a typical thermal protection system and the relative effects of the various icing variables have been determined. For full evaporation of all impinging water, an effective heat density of 14 watts per square inch was required. When a combination of the evaporation and running wet surface systems was employed, a heat requirement of 5 watts per square inch provided protection at severe icing and operating conditions.

Description: The Navier-Stokes equations of motion and the equation of continuity are transformed so as to apply to an orthogonal curvilinear coordinate system rotating with a uniform angular velocity about an arbitrary axis in space. A usual simplification of these equations as consistent with the accepted boundary-layer theory and an integration of these equations through the boundary layer result in boundary-layer momentum-integral equations for three-dimensional flows that are applicable to either rotating or nonrotating fluid boundaries. These equations are simplified and an approximate solution in closed integral form is obtained for a generalized boundary-layer momentum-loss thickness and flow deflection at the wall in the turbulent case. A numerical evaluation of this solution carried out for data obtained in a curving nonrotating duct shows a fair quantitative agreement with the measures values. The form in which the equations are presented is readily adaptable to cases of steady, three-dimensional, incompressible boundary-layer flow like that over curved ducts or yawed wings; and it also may be used to describe the boundary-layer flow over various rotating surfaces, thus applying to turbomachinery, propellers, and helicopter blades.

Description: An investigation was conducted on a 35 deg swept-wing fighter airplane to determine the effects of several blunt-trailing-edge modifications to the wing and tail on the high-speed stability and control characteristics and tracking performance. The results indicated significant improvement in the pitch-up characteristics for the blunt-aileron configuration at Mach numbers around 0.90. As a result of increased effectiveness of the blunt-trailing-edge aileron, the roll-off, customarily experienced with the unmodified airplane in wings-level flight between Mach numbers of about 0.9 and 1.0 was eliminated, The results also indicated that the increased effectiveness of the blunt aileron more than offset the large associated aileron hinge moment, resulting in significant improvement in the rolling performance at Mach numbers between 0.85 and 1.0. It appeared from these results that the tracking performance with the blunt-aileron configuration in the pitch-up and buffeting flight region at high Mach numbers was considerably improved over that of the unmodified airplane; however, the tracking errors of 8 to 15 mils were definitely unsatisfactory. A drag increment of about O.OOl5 due to the blunt ailerons was noted at Mach numbers to about 0.85. The drag increment was 0 at Mach numbers above 0.90.

Description: The trajectories of droplets in the air flowing past NACA 65(1)-208 airfoil and an NACA 65(1)-212 airfoil, both at an angle of attack of 4 degrees, were determined. The amount of water in droplet form impinging on the airfoils, the area of droplet impingement, and the rate of droplet impingement per unit area on the airfoil surface affected were calculated from the trajectories and are presented. The amount, extent, and rate of impingement of the NACA 65(1)-208 airfoil are compared with the results for the NACA 65(1)1-212 airfoil. Under similar conditions of operation, the NACA 65(1)-208 airfoil collects less water than the NACA 65(1)-212 airfoil. The extent of impingement on the upper surface of the NACA 65(1)-208 airfoil is much less than on the upper surface of the NACA 65(1)-212 airfoil, but on the lower surface the extents of impingement are about the same.

Description: A conservative zoning technique, wherein the flow field for a finite-difference calculation is divided into several regions to simplify grid generation, is discussed and is applied in the solution of a two-dimensional problem of complex topology. Calculations are performed on two zonal, or patched, grid systems for the supersonic flow over a double-airfoil configuration. The solution is smooth and continuous across the zonal interfaces, and shock waves pass through the boundaries without distortion. In addition, the time accuracy of the zonal boundary method is verified by a two-zone cylinder calculation with a stationary inner and a rotating outer mesh. The feasibility of the zonal approach for use in the solution of geometrically complex and unsteady problems is thus demonstrated.

Description: At the request of the Materiel Division, Wright Field, the National Advisory Committee for Aeronautics is conducting a program of flight tests on a Kellett YG-1B autogiro equipped with a new type of rotor blade. The new blades are tapered in both plan form. and thickness and are designed to avoid periodic blade twist. One phase of the investigation, involving determination of the moments of the resultant rotor force about the trunnions on which the hub is pivoted for control, has been completed. The results obtained are reported herein.

Description: The damping in roll and rolling effectiveness of two models of a missile having cruciform, triangular, interdigitated wings and tails have been determined through a Mach number range of 0.8 to 1.8 by utilizing rocket-propelled test vehicles. Results indicate that the damping in roll was relatively constant over the Mach umber range investigated. The rolling effectiveness was essentially constant at low supersonic speeds and increased with increasing mach numbers in excess of 1.4 over the Mach number range investigated. Aeroelastic effects increase the rolling-effectiveness parameters pb/2V divided by delta and decrease both the rolling-moment coefficient due to wing deflection and the damping-in-roll coefficient.

Description: An investigation of forced-convection heat transfer and associated pressure drops was conducted with air flowing through electrically heated Inconel tubes having various degrees of square-thread-type roughness, an inside diameter of 1/2 inch, and a length of 24 inches. were obtained for tubes having conventional roughness ratios (height of thread/radius of tube) of 0 (smooth tube), 0.016, 0.025, and 0.037 over ranges of bulk Reynolds numbers up to 350,000, average inside-tube-wall temperatures up to 1950deg R, and heat-flux densities up to 115,000 Btu per hour per square foot. Data The experimental data showed that both heat transfer and friction increased with increase in surface roughness, becoming more pronounced with increase in Reynolds number; for a given roughness, both heat transfer and friction were also influenced by the tube wall-to-bulk temperature ratio. Good correlation of the heat-transfer data for all the tubes investigated was obtained by use of a modification of the conventional Nusselt correlation parameters wherein the mass velocity in the Reynolds number was replaced by the product of air density evaluated at the average film temperature and the so-called friction velocity; in addition, the physical properties of air were evaluated at the average film temperature. The isothermal friction data for the rough tubes, when plotted in the conventional manner, resulted in curves similar to those obtained by other investigators; that is, the curve for a given roughness breaks away from the Blasius line (representing turbulent flow in smooth tubes) at some value of Reynolds number, which decreases with increase in surface roughness, and then becomes a horizontal line (friction coefficient independent of Reynolds number). A comparison of the friction data for the rough tubes used herein indicated that the conventional roughness ratio is not an adequate measure of relative roughness for tubes having a square-thread-type element. The present data, as well as those of other investigators, were used to isolate the influence of ratios of thread height to width, thread spacing to width, and the conventional roughness ratio on the friction coefficient. A fair correlation of the friction data was obtained for each tube with heat addition when the friction coefficient and Reynolds number were defined on the basis of film properties; however, the data for each tube retained the curve characteristic of that particular roughness. The friction data for all the rough tubes could be represented by a single line for the complete turbulence region by incorporating a roughness parameter in the film correlation. No correlation was obtained for the region of incomplete turbulence.

Description: Research was conducted to determine the effect of the electrode parameters of spacing, configuration, and material' on the energy required for ignition of a flowing propane-air mixture. In addition, the data were used to indicate the energy distribution along the spark length and to confirm previous observations concerning the effect of spark duration on ignition energy requirements. The data were obtained with a mixture at a fuel-air ratio of 0.0835 (by weight), a pressure of 3 inches of mercury absolute, a temperature of 80 F, and a mixture velocity of 5 feet per second. Results showed that the energy required for ignition decreased as the electrode spacing was increased; a minimum energy occurred at. a spacing of 0.65 inch for large electrodes. For small electrodes, the spacing for minimum energy was not sharply defined. Small-diameter electrodes required less energy than large-diameter electrodes if the spacing was less than the optimum distance of 0.65 inch; at a spacing equal to the optimum distance, no difference was noted. Significant effects of electrode material on ignition energy were ascribed to differences in the type of spark discharges produced; glow discharges required higher energy than the arc-glow discharges. With pure glow discharges, the ignition energy was substantially constant for lead, cadmium, brass, aluminum, and tungsten electrodes. A method is described for determining the energy distribution along a glow discharge. It was found that one-third to one-half of the energy in the spark was concentrated in a small region near the cathode electrode, and the remainder was uniformly distributed across the spark gap. It was impossible to ascertain the dependence of ignition on. this distribution. It was also observed that long-duration (600 microsec) sparks required much less energy for ignition than did short-duration (1 microsec) sparks.

Description: No abstract available

Description: No abstract available

Description: A vapor block resistant liquid artery structure for heat pipes. A solid tube artery with openings is encased in the sintered material of a heat pipe wick. The openings are limited to that side of the artery which is most remote from the heat source. The liquid in the artery can thus exit the artery through the openings and wet the sintered sheath, but vapor generated at the heat source is unlikely to move around the solid wall of the artery and reverse its direction in order to penetrate the artery through the openings. An alternate embodiment uses finer pore size wick material to resist vapor entry.

Description: A thermal microstructure measurement system (TMMS) operates autonomously h its own internal power supply and telemeters data to a platform. A thermal array is mounted on a cross-braced frame designed to orient itself normal to existing currents with fixed sensor positioning bars protruding from the cross bars. A plurality of matched thermistors, conductivity probes and inclinometers are mounted on the frame. A compass and pressure transducer are contained in an electronics package suspended below the array. The array is deployed on a taut mooring below a subsurface float. Data are digitized, transmitted via cable to a surface buoy and then telemetered to the platform where the data is processed via a computer, recorded and/or displayed. The platform computer also sends commands to the array via telemetry.

Description: Measurement of average skin-friction coefficients have been made on six rocket-powered free-flight models by using the boundary-layer rake technique. The model configuration was the NACA RM-10, a 12.2-fineness-ratio parabolic body of revolution with a flat base. Measurements were made over a Mach number range from 1 to 3.7, a Reynolds number range 40 x 10(exp 6) to 170 x 10(exp 6) based on length to the measurement station, and with aerodynamic heating conditions varying from strong skin heating to strong skin cooling. The measurements show the same trends over the test ranges as Van Driest's theory for turbulent boundary layer on a flat plate. The measured values are approximately 7 percent higher than the values of the flat-plate theory. A comparison which takes into account the differences in Reynolds number is made between the present results and skin-friction measurements obtained on NACA RM-10 scale models in the Langley 4- by 4-foot supersonic pressure tunnel, the Lewis 8- by 6-foot supersonic tunnel, and the Langley 9-inch supersonic tunnel. Good agreement is shown at all but the lowest tunnel Reynolds number conditions. A simple empirical equation is developed which represents the measurements over the range of the tests.

Description: A tank investigation has been conducted on a 1/8-size powered dynamic model of the Grumman JRF-5 airplane equipped with twin hydro-skis. The results of tests using two types of skis are presented: one had vertical sides joining the top surface to the chine; the other had the top surface faired to the chine to eliminate the vertical sides. Both configurations had satisfactory longitudinal stability although the model had a slightly greater stable elevator range available when the skis without the vertical sides were attached. Free model tests indicated no instability present when one ski emerged before the other. Considerable excess thrust was available at all speeds with either type of skis. A hump gross load-resistance ratio of 3.37 was obtained with the skis with the vertical sides and 3.53 with the other skis. Landing behavior in smooth water with yaw up to 15deg and roll up to 15deg in opposite directions was satisfactory with either type of skis.

Description: An investigation was made to determine the static longitudinal and lateral stability and control characteristics of a l/6-scale model of the revised Republic XF-84H airplane with and without the propeller operating. The model had a 40deg swept wing of aspect ratio 3.45 and was equipped with a thin, three-blade supersonic-type propeller. Modifications incorporated in the revised model included a raised horizontal tail, increased rudder size, wing fences at 65 percent semispan, and a modified wing leading edge outboard of the fences. The test results for flap-retracted and flap-deflected conditions indicated that the revised configuration should be satisfactory for most normal flight conditions provided the angle of attack does not exceed the angle for pitch-up. An abrupt pitch-up tendency of the model was evident for the zero thrust condition above approximately 15' angle of attack. Although the effects of power were destabilizing, power-on longitudinal stability was satisfactory through the angle-of-attack range for which the model was stable with zero thrust. Above the angle of attack for pitch-up, an uncontrollable left roll-off tendency would be expected with power on and slats retracted. Projection of wing slats or use of leading-edge chord-extensions with only the left extension drooped were found beneficial in controlling the roll-off tendency with power on; however the most effective means found was projection of only the left slat.

Description: Tests of a 1/5 scale model of a proposed 153-foot high-speed submarine have been conducted in the Langley full-scale tunnel at the request of the Bureau of Ships, Department of the Navy. The test program included: (1) force tests to determine the drag, control effectiveness, and static stability characteristics for a number of model configurations, both in pitch and in yaw, (2) pressure measurements to determine the boundary-layer conditions and flow characteristics in the region of the propeller, and (3) an investigation of the effects of propeller operation on the model aerodynamic characteristics. In response to oral requests from the Bureau of Ships representatives t hat the basic data obtained in these tests be made available to them as rapidly as possible, this data report has been prepared to present some of the more pertinent results. All test results given in the present paper are for the propeller-removed condition and were obtained at a Reynolds number of approximately 22,300,000 based on model length.

Description: Free-flight tests have been made to determine the zero-lift drag of several configurations of the XAAM-N-2 pilotless aircraft. Base-pressure measurements were also obtained for some of the configurations. The results show that increasing the wing-thickness ratio from 4 to 6 percent increased the wing drag by about 100 percent at M = 1.3 and by about 30 percent at M = 1.8. Increasing the nose fineness ratio from 5.00 to 6.25 reduced the drag coefficient of the wingless models a maximum of about 0.030 (10 percent) at M = 2.0. A corresponding change in nose shape for the winged models decreased the drag coefficient by about 0.05 in the Mach number range from 1.1 to 1.4; at Mach numbers greater than 1.6 no measurable reduction in drag coefficient was obtained. The drag of the present Sparrow fuselage is less than that of a parabolic fuselage which could contain the same equipment.

Description: A limited investigation of a 1/24-scale dynamically similar model of the Navy Bureau of Aeronautics DR-77 design was conducted in Langley tank no. 2 to determine the calm-water take-off and the rough-water landing characteristics of the design with particular regard to the take-off resistance and the landing accelerations. During the take-off tests, resistance, trim, and rise were measured and photographs were taken to study spray. During the landing tests, motion-picture records and normal-acceleration records were obtained. A ratio of gross load to maximum resistance of 3.2 was obtained with a 30 deg. dead-rise hydro-ski installation. The maximum normal accelerations obtained with a 30 deg. dead-rise hydro-ski installation were of the order of 8g to log in waves 8 feet high (full scale). A yawing instability that occurred just prior to hydro-ski emergence was improved by adding an afterbody extension, but adding the extension reduced the ratio of gross load to maximum resistance to 2.9.

Description: Transfer functions descriptive of the response of most engine variables were determined from transient data that were obtained from approximate step inputs in fuel flow and in exhaust-nozzle area. The speed responses of both spools to fuel flow and to turbine-inlet temperature appeared as identical first-order lags. Response to exhaust-nozzle area was characterized by a first-order lag response of the outer-spool speed, accompanied by virtually no change in inner-spool speed.

Description: The static lateral- and directional-stability characteristics of a high-speed fighter-type airplane, obtained from wind-tunnel tests of a model, are presented. The model consisted of a thin, unswept wing of aspect ratio 2.3 and taper ratio 0.385, a body, and a horizontal tail mounted in a high position on a vertical tail. Rolling-moment, yawing moment, and cross-wind-force coefficients are presented for a range of sideslip angles of -5 deg. to +5 deg, for Mach numbers of 0.90, 1.45, and 1.90. Data are presented which show the effects on the lateral and directional stability of: (1) component parts of the complete model, (2) modification of the empennage so as to provide different heights of the horizontal tail above the wing plane, (3) angle of attack, and (4) dihedral of the wing.

Description: The stator-blade angles in the twelfth through fifteenth stages of a 16-stage axial-flow compressor were increased 3O. The over-all performance of this modified compressor is compared to the performance of the compressor with original blade angles. The matching characteristics of the modified compressor and a two-stage turbine were obtained and compared to those of the compressor with original blade angles and the same turbine.

Description: An investigation has been conducted to determine the static stability and control and damping in roll and yaw of a 0.13-scale model of the Convair XFY-1 airplane with propellers off from 0 deg to 90 deg angle of attack. The tests showed that a slightly unstable pitch-up tendency occurred simultaneously with a break in the normal-force curve in the angle-of-attack range from about 27 deg to 36 deg. The top vertical tail contributed positive values of static directional stability and effective dihedral up to an angle of attack of about 35 deg. The bottom tail contributed positive values of static directional stability but negative values of effective dihedral throughout the angle-of-attack range. Effectiveness of the control surfaces decreased to very low values at the high angles of attack, The model had positive damping in yaw and damping in roll about the body axes over the angle-of-attack range but the damping in yaw decreased to about zero at 90 deg angle of attack.

Description: Altitude performance characteristics of the J65-B3 turbojet engine and its components were obtained at engine-inlet conditions corresponding to Reynolds number indices from 0.2 to 0.8 over a range of corrected engine speeds from 70 to 110 percent of rated speed. Engine operational limits up to an altitude of 75,000 feet together with ignition and windmilling characteristics were also obtained. The engine and component data are presented both in graphical and in tabulated form. The operational characteristics are presented in graphical form.

Description: An improved efficiency arcjet thruster has a constrictor and electrically-conductive nozzle anode defining an arc chamber, and an electrically-conductive rod having a tip spaced upstream from the constrictor and defining a cathode spaced from the anode by a gap generally coextensive with the arc chamber. An electrical potential is applied to the anode and cathode to generate an electrical arc in the arc chamber from the cathode to anode. Catalytically decomposed hydrazine is supplied to the arc chamber with generation of the arc so as to produce thermal heating and expansion thereof through the nozzle. The constrictor can have a electrically insulative portion disposed between the cathode tip and the nozzle anode, and an electrically-conductive anode extension disposed along the insulative portion so as to define an auxiliary gap with the cathode tip substantially smaller than the gap defined between the cathode and nozzle anode for facilitating startup of arc generation. The constrictor can also include an electrically-conductive electrode with a variable electrical potential to vary the shape of the arc generated in the arc chamber. Also, the cathode is mounted for axial movement such that the gap between its tip and the nozzle anode can be varied to facilitate a generally nonerosive generation of the electrical arc at startup and reliable steady state operation. Further, the arc chamber can have a nonparallel subsonic-to-supersonic transition configuration, or alternatively solely a nonparallel supersonic configuration, for improved arc attachment.

Description: A vapor block resistant liquid artery structure for heat pipes. A solid tube artery with openings is encased in the sintered material of a heat pipe wick. The openings are limited to that side of the artery which is most remote from the heat source. The liquid in the artery can thus exit the artery through the openings and wet the sintered sheath, but vapor generated at the heat source is unlikely to move around the solid wall of the artery and reverse its direction in order to penetrate the artery through the openings. An alternate embodiment uses finer pore size wick material to resist vapor entry.

Description: A compression pylon for an aircraft with a wing-mounted engine, that does not cause supersonic airflow to occur within the fuselage-wing-pylon-nacelle channel is presented. The chord length of the pylon is greater than the local chord length of the wing to which it is attached. The maximum thickness of the pylon occurs at a point corresponding to the local trailing edge of the wing. As a result, the airflow through the channel never reaches supersonic velocities.

Description: The study of the hydrodynamic properties of planing bottom of flying boats and seaplane floats is at the present time based exclusively on the curves of towing tests conducted in tanks. In order to provide a rational basis for the test procedure in tanks and practical design data, a theoretical study must be made of the flow at the step and relations derived that show not only qualitatively but quantitatively the inter-relations of the various factors involved. The general solution of the problem of the development of hydrodynamic forces during the motion of the seaplane float or flying boat is very difficult for it is necessary to give a three-dimensional solution, which does not always permit reducing the analysis to the form of workable computation formulas. On the other had, the problem is complicated by the fact that the object of the analysis is concerned with two fluid mediums, namely, air and water, which have a surface of density discontinuity between them. The theoretical and experimental investigations on the hydrodynamics of a ship cannot be completely carried over to the design of floats and flying-boat hulls, because of the difference in the shape of the contour lines of the bodies, and, because of the entirely different flow conditions from the hydrodynamic viewpoint.

Description: An impulse-momentum method for determining impact conditions for landing gears in eccentric landings is presented. The analysis is primarily concerned with the determination of contact velocities for impacts subsequent to initial touchdown in eccentric landings and with the determination of the effective mass acting on each landing gear. These parameters determine the energy-absorption requirements for the landing gear and, in conjunction with the particular characteristics of the landing gear, govern the magnitude of the ground loads. Changes in airplane angular and linear velocities and the magnitude of landing-gear vertical, drag, and side impulses resulting from a landing impact are determined by means of impulse-momentum relationships without the necessity for considering detailed force-time variations. The effective mass acting on each gear is also determined from the calculated landing-gear impulses. General equations applicable to any type of eccentric landing are written and solutions are obtained for the particular cases of an impact on one gear, a simultaneous impact on any two gears, and a symmetrical impact. In addition a solution is presented for a simplified two-degree-of-freedom system which allows rapid qualitative evaluation of the effects of certain principal parameters. The general analysis permits evaluation of the importance of such initial conditions at ground contact as vertical, horizontal, and side drift velocities, wing lift, roll and pitch angles, and rolling and pitching velocities, as well as the effects of such factors as landing gear location, airplane inertia, landing-gear length, energy-absorption efficiency, and wheel angular inertia on the severity of landing impacts. -A brief supplementary study which permits a limited evaluation of variable aerodynamic effects neglected in the analysis is presented in the appendix. Application of the analysis indicates that landing-gear impacts in eccentric landings can be appreciably more severe than impacts in symmetrical landings with the same sinking speed. The results also indicate the effects of landing-gear location, airplane inertia, initial wing lift, side drift velocity, attitude, and initial rolling velocity on the severity of both initial and subsequent landing-gear impacts. A comparison of the severity of impacts on auxiliary gears for tricycle and quadricycle configurations is also presented.

Description: Investigations were conducted of a 12 degree 21-inch conical diffuser of 2:l area ratio to determine the interrelation of boundary layer growth and performance characteristics. surveys were made of inlet and exit from, longitudinal static pressures were recorded, and velocity profiles were obtained through an inlet Reynolds number range, determined From mass flows and based on inlet diameter of 1.45 x 10(exp 6) to 7.45 x 10(exp 6) and a Mach number range of 0.11 to approximately choking. These investigations were made to two thicknesses of inlet boundary layer. The mean value, over the entire range of inlet velocities, of the displacement thickness of the thinner inlet boundary layer was approximately 0.035 inch and that of the thicker inlet boundary layer was approximately six times this value. The loss coefficient in the case of the thinner inlet boundary layer had a value between 2 to 3 percent of the inlet impact pressure over most of the air-flow range. The loss coefficient with the thicker inlet boundary layer was of the order of twice that of the thinner inlet boundary layer at low speeds and approximately three times at high speeds. In both cases the values were substantially less than those given in the literature for fully developed pipe flow. The static-pressure rise for the thinner inlet boundary layer was of the order of 95 percent of that theoretically possible over the entire speed range. For the thicker inlet boundary layer the static pressure rise, as a percentage of that theoretically possible, ranged from 82 percent at low speeds to 68 percent at high speeds.

Description: Performance and boundary-layer data were taken in a 12 degree 10-inch inlet-diameter conical diffuser of 2:1 exit- to inlet-area ratio. These data were taken for two inlet-boundary-layer conditions. The first condition was that of a thinner inlet boundary later (boundary-layer displacement thickness, delta* approximately equal to 0.034) produced by an inlet section approximately 1 inlet diameter in length between the entrance bell and the diffuser. The second condition was a thicker inlet boundary layer (delta* approximately equal to 0.120) produced by an additional inlet section length of approximately 6 diameters. Longitudinal static-pressure distributions were measured fro wall static orifices. Transverse total- and static-pressure surveys were made at the inlet and exit stations. Boundary-layer velocity distributions were measured at seven stations between the inlet and exit. These data were obtained for a Reynolds number (based on inlet diameter) range of 1 x 10(exp 6) to 3.9 x 10(exp 6). The corresponding Mach number range was from M = 0.2 to choking. At the maximum-power-available condition supersonic flow was obtained as far as 4.5 inches downstream from the diffuser inlet with a maximum Mach number of M approximately equal to 1.5. The total-pressure loss through the diffuser in percentage of inlet dynamic pressure was approximately 2.5 percent for the thinner inlet boundary later and 5.5 percent for the thicker inlet boundary later over the lower subsonic range. These valued increased with increasing flow rate- the values for the thicker inlet boundary later more than those for the thinner inlet boundary layer. The diffuser effectiveness, expressed as the ratio of the actual static-pressure rise to the ideal static-pressure rise, was about 85 percent for the thinner inlet boundary layer and about 67 percent for the thicker inlet boundary later in the lower subsonic range. These values decrease with increasing flow rate. Separated flow was observed for both inlet-boundary-layer conditions in the region of adverse pressure gradient just downstream of the transition curvature from inlet section to diffuser. The flow for the thinner-inlet-boundary-layer condition did not fully re-establish itself along the diffuser walls. The thicker inlet-boundary-layer flow, while not completely re-establishing the normal flow pattern downstream of the separated region, did re-establish more successfully than the thinner inlet boundary layer.

Description: A wind-tunnel investigation has been conducted to determine the stability and control characteristics of a full-size model of the Hughes MX-904 missile. Aerodynamic characteristics of the complete model through moderate ranges of angles of attack and yaw, with an additional test made through an angle of attack of 180 degrees, are presented. The effects of horizontal tail deflection are also included.

Description: A free-flight 0.12-scale rocket-boosted model of the North American MX-770 (X-10) missile has been tested in flight by the Pilotless Aircraft Research Division of the Langley Aeronautical Laboratory. Drag, longitudinal stability, and duct performance data were obtained at Mach numbers from 0.8 to 1.7 covering a Reynolds number range of about 9 x 10(exp 6) to 24 x 10(exp 6) based on wing mean aerodynamic chord. The lift-curve slope, static stability, and damping-in-pitch derivatives showed similar variations with Mach number, the parameters increasing from subsonic values in the transonic region and decreasing in the supersonic region. The variations were for the most part fairly smooth. The aerodynamic center of the configuration shifted rearward in the transonic region and moved forward gradually in the supersonic region. The pitching effectiveness of the canard control surfaces was maintained throughout the flight speed range, the supersonic values being somewhat greater than the subsonic. Trim values of angle of attack and lift coefficient changed abruptly in the transonic region, the change being associated with variations in the out-of-trim pitching moment, control effectiveness, and aerodynamic-center travel in this speed range. Duct total-pressure recovery decreased with increase in free-stream Mach number and the values were somewhat less than normal-shock recovery. Minimum drag data indicated a supersonic drag coefficient about twice the subsonic drag coefficient and a drag-rise Mach number of approximately 0.90. Base drag was small subsonically but was about 25 percent of the minimum drag of the configuration supersonically.

Description: An investigation of a 1/24- scale dynamically similar model of the Douglas C-124 airplane was made to determine the ditching characteristics and proper technique for ditching the airplane. Various conditions of damage, landing attitude, flap setting, and speed were investigated. The behavior of the model was determined from visual observations, motion- picture records, and time-history deceleration records. The results of the investigation are presented in table form, photographs, and curves. It was concluded on the basis of results from model tests with scale-strength bottoms (equivalent to 1150 pounds per square foot, full scale) that the airplane should be ditched at a medium nose-high landing attitude (near 7deg) with flaps full down. The airplane will probably make a smooth run with considerable damage resulting to the fuselage bottom just forward of the wing, but it is not likely that the water inflow will be overwhelming to personnel provided they are not in the belly compartment. Longitudinal decelerations in calm water will be about 2 1/2g and the landing run will be about four fuselage lengths.

Description: An investigation was conducted in the Langley 20-foot free-spinning tunnel on a 1/23-scale model of the McDonnell F3H-1N airplane. The effects of control settings and movements upon the erect and inverted spin and recovery characteristics of the model were determined for the clean condition. Spin-recovery parachute tests were also performed. The results indicated that erect spins obtained on the airplane for the take-off or combat loadings should be satisfactorily terminated if full rudder reversal is accompanied by moving the ailerons to full with the spin (stick full right in a right spin). The spins obtained should be oscillatory in pitch, roll, and yaw. Recoveries from inverted spins should be satisfactory by full reversal of the rudder. A 16.7-foot- diameter tail parachute with a towline length of 30 feet and a drag coefficient of 0.734 should be adequate for emergency recovery from demonstration spins.

Description: An application of airfoil design methods was used to design series of related turbine-blade profiles to satisfy the conditions of inlet flow angle and turning angle encountered in the usual range of turbine operation. A series of blade profiles applicable to most turbine blading requirements and a secondary series with particular reference to impulse conditions were designed. Five blade sections from these series ranging in mean-line turning angles from 63 deg. to 120 deg. were tested in low-speed cascade tunnels. From low-speed test results optimum blade angles of attack were selected at each test condition. The induced angle and the deviation angle of the flow were determined from the low-speed data. If these angles are known for the solidity and inlet angle of an application, the necessary camber is specified. A method of predicting high-speed pressure distributions from low-speed cascade test results is presented to extend the usefulness of the low-speed data. Sample high-speed tests of two of the five blade sections were made at Mach numbers up to the critical value. The results indicated satisfactory flow conditions in all of the blade passages tested.

Description: A ditching investigation of a model of the Convair-Liner airplane was made to observe the behavior and determine the safest procedure for making an emergency water landing. The ditching model was designed and constructed by the National Advisory Committee for Aeronautics. Design information on the airplane was furnished by the Consolidated Vultee Aircraft Corporation. A three-view drawing of the airplane is shown. The investigation was made in calm water at the Langley tank no. 2 monorail.

Description: An investigation has been made in the Langley gust tunnel with two identical airplane models approximating 1/40-scale models of the B-29, coupled in tandem with a boom so that the individual centers of gravity were equidistant from the single coupling joint at the tail of the lead airplane. Time histories of the boom joint load were obtained as the models were flown through a gust. The results indicate that on a similar configuration involving airplanes the size of B-29 airplanes a load on the boom joint of 10,000 to 14,000 pounds could be induced by encountering a gust of 50 feet per second and having a gradient distance of 17 chords, at a forward speed of 380 feet per second and that the total load is extremely sensitive to the steadiness of flight that can be maintained with or without a gust. It is felt that the results are probably satisfactory to show order of magnitude, but it does not appear possible that a precise determination of the joint load that would be applicable to the full-scale airplanes can be obtained by gust-tunnel tests.

Description: Flight tests have been conducted on rocket-propelled models of an airplane configuration incorporating a sweptback wing with inverse taper to investigate the drag, stability, and control characteristics at transonic and supersonic speeds. The models were tested with a conventional tail arrangement in the Mach number range from 0.55 to 1.2. In addition to the various aerodynamic parameters obtained, the flying qualities were computed for a full-scale airplane with the center-of-gravity location at 18 percent of the mean aerodynamic chord. Also, included in this investigation are drag measurements made on relatively simple fixed-control models tested with both conventional and V-tail arrangements.

Description: An elementary type of analysis has been used to determine the amount of wing tip that must be severed to produce irrevocable loss of control of a B-29 airplane. The remaining inboard structure of the Boeing B-29 wing has then been analyzed and curves are presented for the estimated reduction in structural strength due to four general types of damage produced by rod-type warhead fragments. The curves indicate the extent of structural damage required to produce a kill of the aircraft within 10 seconds.

Description: Method and apparatus for obtaining dynamic calibrations of pressure transducers. A calibration head (15), a flexible tubing (23) and a bellows (20) enclose a volume of air at atmospheric pressure with a transducer (11) to be calibrated subject to the pressure inside the volume. All of the other apparatus in the drawing apply oscillations to bellows (20) causing the volume to change thereby applying oscillating pressures to transducer (11) whereby transducer (11) can be calibrated.

Description: The invention does away with the necessity of moving parts such as a check valve in a nuclear reactor cooling system. Instead, a jet pump, in combination with a TEMP, is employed to assure safe cooling of a nuclear reactor after shutdown. A main flow exists for a reactor coolant. A point of withdrawal is provided for a secondary flow. A TEMP, responsive to the heat from said coolant in the secondary flow path, automatically pumps said withdrawn coolant to a higher pressure and thus higher velocity compared to the main flow. The high velocity coolant is applied as a driver flow for the jet pump which has a main flow chamber located in the main flow circulation pump. Upon nuclear shutdown and loss of power for the main reactor pumping system, the TEMP/jet pump combination continues to boost the coolant flow in the direction it is already circulating. During the decay time for the nuclear reactor, the jet pump keeps running until the coolant temperature drops to a lower and safe temperature where the heat is no longer a problem. At this lower temperature, the TEMP/jet pump combination ceases its circulation boosting operation. When the nuclear reactor is restarted and the coolant again exceeds the lower temperature setting, the TEMP/jet pump automatically resumes operation. The TEMP/jet pump combination is thus automatic, self-regulating and provides an emergency pumping system free of moving parts.

Description: A wake shield space processing facility (10) for maintaining ultra-high levels of vacuum is described. The wake shield (12) is a truncated hemispherical section having a convex side (14) and a concave side (24). Material samples (68) to be processed are located on the convex side of the shield, which faces in the wake direction in operation in orbit. Necessary processing fixtures (20) and (22) are also located on the convex side. Support equipment including power supplies (40, 42), CMG package (46) and electronic control package (44) are located on the convex side (24) of the shield facing the ram direction. Prior to operation in orbit the wake shield is oriented in reverse with the convex side facing the ram direction to provide cleaning by exposure to ambient atomic oxygen. The shield is then baked-out by being pointed directed at the sun to obtain heating for a suitable period.

Description: A systematic research program is being carried out in the Langley high-speed 7- by 10-foot tunnel to determine the aerodynamic characteristics of various arrangements of the component parts of research-type airplane models, including some complete model configurations. Data are being obtained on characteristics in pitch, sideslip, and during steady roll at Mach numbers from 0.40 to about 0.95. This paper presents results which show the effect of taper ratio on the aerodynamic characteristics in sideslip of wing-fuselage combinations having wings with a sweep of 45 degrees at the quarter-chord line, an aspect ratio of 4, and a NACA 65A006 airfoil section.

Description: A space vehicle thermal heat rejection system 10 utilizing separate optimized heat pipe components for the functions of heat acquisition, heat transport, and heat rejection. A honeycomb panel heat pipe evaporator section 20 performs the function of heat acquisition, and forms a closed thermodynamic system with a dual channel heat pipe transport section 30, which performs the function of heat transport. A plurality of truss or channel core heat pipe rejection fins 41 form the condenser section 40, which performs the function of heat rejection. A common wall 32 separates the condenser section 40 from the transport section 30. Using the above heat pipe components and having efficient interfacing between them results in high performance factors for the overall system.

Description: A fluid leak indicator (30) for detecting and indicating leaks in visually inaccessible fluid tubing joints (20, 21), such as those obstructed by insulation (24), includes a bag system (25) and a wicking system (30) surrounding or wrapping the joints (20, 21) under the visual obstructing material (24). Leaking fluid is collected in the bag (25) or on the wicking material (34) where it is conducted along the wicking material (34) to a visibly accessible capturing transparent indicator bulb (35) for providing a visual indication of the leak without requiring a chemical change in the capturing indicator bulb (35).

Description: The lift, pitching-moment, and drag characteristics of a missile configuration having a body of fineness ratio 9.33 and a cruciform triangular wing and tail of aspect ratio 4 were measured at a Mach number of 1.99 and a Reynolds number of 6.0 million, based on the body length. The tests were performed through an angle-of-attack range of -5 deg to 28 deg to investigate the effects on the aerodynamic characteristics of roll angle, wing-tail interdigitation, wing deflection, and interference among the components (body, wing, and tail). Theoretical lift and moment characteristics of the configuration and its components were calculated by the use of existing theoretical methods which have been modified for application to high angles of attack, and these characteristics are compared with experiment. The lift and drag characteristics of all combinations of the body, wing, and tail were independent of roll angle throughout the angle-of-attack range. The pitching-moment characteristics of the body-wing and body-wing- tail combinations, however, were influenced significantly by the roll angle at large angles of attack (greater than 10 deg). A roll from 0 deg (one pair of wing panels horizontal) to 45 deg caused a forward shift in the center of pressure which was of the same magnitude for both of these combinations, indicating that this shift originated from body-wing interference effects. A favorable lift - interference effect (lift of the combination greater than the sum of the lifts of the components) and a rearward shift in the center of pressure from a position corresponding to that for the components occurred at small angles of attack when the body was combined with either the exposed wing or tail surfaces. These lift and center-of-pressure interference effects were gradually reduced to zero as the angle of attack was increased to large values. The effect of wing-tail interference, which influenced primarily the pitching-moment characteristics, is dependent on the distance between the wing trailing vortex wake and the tail surfaces and thus was a function of angle of attack, angle of roll, and wing- tail interdigitation. Although the configuration at zero roll with the wing and tail in line exhibited the least center-of-pressure travel, the configuration with the wing and tail interdigitated had the least change in wing- tail interference over the angle - of-attack range. The lift effectiveness of the variable-incidence wing was reduced by more than 70 percent as a result of an increase in the combined angle of attack and wing incidence from 0 deg to 40 deg center dot The wing- tail interference (effective downwash at the tail) due to wing deflection was nearly zero as a result of a region of negative vorticity shed from the inboard portion of the wing. The lift characteristics of the configuration and its components were satisfactorily predicted by the calculated results, but the pitching moments at large angles of attack were not because of the influence of factors for which no adequate theory is available, such as the variation of the cross flow drag coefficient along the body and the effect of the wing downwash field on the after body loading.

Description: Preliminary information on the complex subject of the fatigue strength of fabricated structural members for aircraft is presented in the test results obtained on several different types of airship girders subjected to axial tension and compression in a resonance fatigue machine. A description of this machine as well as numerous photographs of the fatigue failures are given. There is also presented an extended bibliography on the subject of fatigue strength.

Description: This document presents equations for the two-dimensional stationary problem of gas dynamics, and uses them to derive other equations, including equations for vorticity.

Description: The vortices forming in flowing water behind solid bodies are not represented correctly by the solution of the potential theory nor by Helmholtz's jets. Potential theory is unable to satisfy the condition that the water adheres at the wetted bodies, and its solutions of the fundamental hydrodynamic equations are at variance with the observation that the flow separates from the body at a certain point and sends forth a highly turbulent boundary layer into the free flow. Helmholtz's theory attempts to imitate the latter effect in such a way that it joins two potential flows, jet and still water, nonanalytical along a stream curve. The admissibility of this method is based on the fact that, at zero pressure, which is to prevail at the cited stream curve, the connection of the fluid, and with it the effect of adjacent parts on each other, is canceled. In reality, however, the pressure at these boundaries is definitely not zero, but can even be varied arbitrarily. Besides, Helmholtz's theory with its potential flows does not satisfy the condition of adherence nor explain the origin of the vortices, for in all of these problems, the friction must be taken into account on principle, according to the vortex theorem.

Description: The use of the linearized equations of Chaplygin to calculate the subsonic flow of a gas permits solving the problem of the flow about a wing profile for absence and presence of circulation. The solution is obtained in a practical convenient form that permits finding all the required magnitudes for the gas flow (lift, lift moment velocity distribution over the profile, and critical Mach number). This solution is not expressed in simple closed form; for a certain simplifying assumption, however, the equations of Chaplygin can be reduced to equations with constant coefficients, and solutions are obtained by using only the mathematical apparatus of the theory of functions of a complex variable. The method for simplifying the equations was pointed out by Chaplygin himself. These applied similar equations to the solution of the flow problem and obtained a solution for the case of the absence of circulation.

Description: In the flow about a body with large subsonic velocity if the velocity of the approaching flow is sufficiently large, regions of local supersonic velocities are formed about the body. It is known from experiment that these regions downstream of the flow are always bounded by shock waves; a continuous transition of the supersonic velocity to the subsonic under the conditions indicated has never been observed. A similar phenomenon occurs in pipes. If at two cross sections of the pipe the velocity is subsonic and between these sections regions of local supersonic velocity are formed without completely occupying a single cross section, these regions are always bounded by shock waves.

Description: Frequency-domain methods are used to identify the open-loop dynamics of the XV-15 tilt-rotor aircraft from flight tests. Piloting and data analysis techniques are presented to determine frequency response plots and equivalent transfer function models. The open-loop pitch and roll dynamics for the hover flight condition exhibit unstable low-frequency oscillations, whereas the dynamics in the remaining degrees of freedom are lightly damped and generally decoupled. Comparisons of XV-15 flight-test and simulator data are more favorable for high-frequency inputs (omega greater than 1.0 rad/sec) than low-frequency inputs. Time-domain comparisons of the extracted transfer functions with step response flight data are very favorable, even for large amplitude motions. The results presented in this paper demonstrate the utility of the frequency-domain techniques for dynamics identification and simulator fidelity studies.

Description: A comprehensive rotorcraft analysis model, CAMRAD, was used to predict blade aerodynamic and structural loads and performance for comparison with flight test data. The data were obtained from an SA349/2 helicopter with an advanced geometry rotor. Sensitivity of the correlation to wake geometry, blade dynamics, and blade aerodynamic effects was investigated. Blade chordwise pressure coefficients were predicted for the transonic regimes near the blade tip using the model coupled with two finite-difference codes.

Description: A steady incompressible three-dimensional viscous flow analysis has been conducted for the Space Shuttle External Tank/Orbiter propellant feed line disconnect flapper valves with upstream elbows. The full Navier-Stokes code, INS3D, is modified to handle interior obstacles. Grids are generated by SVTGD3D code. Two dimensional initial grids in the flow cross section with and without the flappers are improved by elliptic smoothing to provide better orthogonality, clustering and smoothness to the three dimensional grid. The flow solver is tested for stability and convergence in the presence of interior flappers. An under-relaxation scheme has been incorporated to improve the solution stability. Important flow characteristics such as secondary flows, recirculation, vortex and wake regions, and separated flows are observed. Computed values for forces, moments, and pressure drop are in satisfactory agreement with water flow test data covering a maximum tube Reynolds number of 3.5 x 10(exp 6). The results will serve as a guide to improved design and enhanced testing of the disconnect.

Description: Combustion testing at total enthalpy conditions corresponding to flight Math numbers in excess of 12 requires the use of impulse facilities. The expansion tube is the only operational facility of its size which can provide these conditions without excessive oxygen dissociation or driver gas contamination. Expansio tube operation is described herein and the operational parameters having the largest impact on its performance are determined. These are: driver-to-intermediate chamber pressure ratio, driver gas molecular weight and specific heat ratio, and driver gas temperature. Increases in the lase named parameter will markedly affect the test section static pressure. Preliminary calibration tests are discussed and test gas conditions which have been achieved are presented. Calculated and experimental test times are compared and the parameters affecting test time are discussed. The direction of future work using this important experimental tool is indicated.

Description: An experimental investigation of the variation of aileron rolling effectiveness and total drag with Mach number has been made using 1/6-scale rocket-propelled models of the Bell MX-776. Three models having constant-chordwise-thickness full-span aileron at approximate deflections of 2 deg, 5 deg, and 15 deg have been flown. Positive control effectiveness over the Mach number range between approximately 0.5 and 1.2 was obtained from the models and no indication of reversal of effectiveness was encountered. The ratio of tip helix angle to aileron deflection indicated a decrease in proportional rolling effectiveness with increasing deflections in the Mach number range from approximately 0.7 to 1.0. A drag rise of about 125 percent in the transonic region between Mach numbers of 0.85 and 1.02 followed by a gradual decrease at higher speeds was revealed.

Description: As part of a program to determine the feasibility of using a fighter airplane as a parasite in combination with a Consolidated Vultee RB-36 for long-range reconnaissance missions (project FICON), an experimental investigation has been made in the Langley free-flight tunnel to determine the dynamic stability and control characteristics of a 1/17.5-scale model of a Chance Vought F7U-3 airplane in several tow configurations. The investigation consisted of flight tests in which the model was towed from a strut in the tunnel by a towline and by a direct coupling which provided complete angular freedom. The tests with the direct coupling also included a study of the effect of spring restraint in roll in order to simulate approximately the proposed full-scale arrangement in which the only freedom is that permitted by the flexibility of the launching and retrieving trapeze carried by the-bomber. For the tow configurations in which a towline was used (15 and 38 feet full scale), the model had a very unstable lateral oscillation which could not be controlled. The stability was also unsatisfactory for the tow configuration in Which the model was coupled directly to the strut with complete angular freedom. When spring restraint in roll was added, however, the stability was satisfactory. The use of the yaw damper which increased the damping in yaw to about six times the normal value of the model appeared to have no appreciable effect on the lateral oscillations in the towline configurations, but produced a slight improvement in the case of the direct coupling configurations. The longitudinal stability was satisfactory for those cases in which the lateral stability was good enough to permit study of longitudinal motions.

Description: An investigation of a 1/24-scale dynamically similar model of the Boeing B-47 airplane was made to determine the ditching characteristics and proper ditching technique for the airplane. Various conditions of damage, landing attitude, flap setting, and speed were investigated. The behavior of the model was determined from visual observations, motion-picture records, and time-history deceleration records. The results of the investigation are presented in table form, photographs, and curves. The airplane should be ditched at the lowest speed and highest attitude consistent with adequate control; the flaps should be full down. The airplane will probably make a deep but fairly smooth run. The fuselage bottom will be damaged and partially filled with water; consequently, crew members should be assigned ditching stations near an exit in the upper or forward part of the fuselage. The nacelles may be expected to be torn away from the wing. In calm water the maximum decelerations will be about 3g and the landing run will be about 6 fuselage lengths.

Description: At the request of the Air Materiel Command, an investigation was made in the Langley free-flight tunnel to determine the longitudinal stability and control characteristics of models coupled together in a tandem configuration for aerial refueling similar to one proposed by the Douglas Aircraft Company, Inc. Static force tests were made with 1/20-scale models of the B-29 and F-80 airplanes to determine the effects of rigidly coupling the airplanes together. The Douglas configuration differs from the rigid configuration tested in that it provides for some freedom in pitch and vertical displacement. The force tests showed that, for the bomber alone, the aerodynamic center was 0.21 mean aerodynamic chord behind the center of gravity (stable) but that for the tandem configuration with rigid coupling the aerodynamic center was 0.28 mean aerodynamic chord forward of the center of gravity of the combination (unstable). This reduction in stability was caused by the downwash of the bomber on the fighter. The pitching moment produced by elevator deflection of the bomber was reduced approximately 50 percent by addition of the fighter. Some recent flight tests made in the free-flight tunnel on models in a similar tandem configuration indicated that, with a hinged coupling permitting freedom in pitch, the stability of the combination was better than that obtained with a rigid coupling and was about the same as that for the bomber alone.

Description: At the request of the Air Materiel Command an investigation was made in the Langley free-flight tunnel to determine the static longitudinal stability and control characteristics of models coupled together in a tandem configuration proposed by All American Airways, Inc. Force tests were made using 1/20-scale models of B-29 end F-80 airplanes to determine the effects of coupling the fighter to the tail of the bomber. The results of the investigation showed that for the bomber alone the aerodynamic center was 0.21 mean aerodynamic chord behind the center of gravity (stable) but that for the tandem configuration the aerodynamic center was 0.09 mean aerodynamic chord forward of the center of gravity, of the combination (unstable). The elevator effectiveness of the bomber was reduced approximately 50 percent by addition of the fighter. Some recent flight tests made in the free-flight tunnel with models simulating the proposed configuration indicate that the reduction in stability may be minimized by incorporating a hinged coupling permitting freedom in pitch.

Description: An investigation has been made in the Langley 9- by 12-inch supersonic blowdown tunnel to determine the effects of external-store location on the lift, drag, and pitching-moment characteristics of a 45 degree sweptback wing at Mach numbers of 1.41, 1.62, and 1.96. The spanwise, chordwise, and vertical location of a Douglas-Aircraft Company, Inc., store of fineness ratio 8.58 was systematically varied over the outer 60 percent of the wing semispan. A brief investigation of strut sweep angle was also made. The test Reynolds number based on the wing mean aerodynamic chord ranged from 1.3 x 10(exp 6) to 1.5 x 10(exp 6).

Description: An investigation was made by the NACA wing-flow method to determine the drag, pitching-moment, lift, and angle-of-attack characteristics at transonic speeds of various configurations of a semispan model of an early configuration of the XF7U-1 tailless airplane. The results of the tests indicated that for the basic configuration with undeflected ailavator, the zero-lift drag rise occurred at a Mach number of about 0.85 and that about a five-fold increase in drag occurred through the transonic speed range. The results of the tests also indicated that the drag increment produced by -8.0 degrees deflection of the ailavator increased with increase in normal-force coefficient and was smaller at speeds above than at speeds below the drag rise. The drag increment produced by 35 degree deflection of the speed brakes varied from 0.040 to 0.074 depending on the normal-force coefficient and Mach number. These values correspond to drag coefficients of about 0.40 and 0.75 based on speed-brake frontal area. Removal of the fin produced a small positive drag increment at a given normal-force coefficient at speeds during the drag rise. A large forward shift of the neutral-point location occurred at Mach numbers above about 0.90 upon removal of the fin, and also a considerable forward shift throughout the Mach number range occurred upon deflection of the speed brakes. Ailavator ineffectiveness or reversal at low deflections, similar to that determined in previous tests of the basic configuration of the model in the Mach number range from about 0.93 to 1.0, was found for the fin-off configuration and for the model equipped with skewed (more highly sweptback) hinge-line ailavators. With the speed brakes deflected, little or no loss in the incremental pitching moment produced by deflection of the ailavator from O degrees to -8.00 degrees occurred in the Mach number range from 0.85 to 1.0 in contrast to a considerable loss found in previous tests with the speed brakes off.

Description: An investigation has been conducted in the Langley 20-foot free-spinning tunnel to determine the spin and recovery characteristics of a 0.057-scale model of the modified Chance Vought XF7U-1 airplane. The primary change in the design from that previously tested was a revision of the twin vertical tails. Tests were also made to determine the effect of installation of external wing tanks. The results indicated that the revision in the vertical tails did not greatly alter the spin and recovery characteristics of the model and recovery by normal use of controls (fill rapid rudder reversal followed approximately one-half turn later by movement of the stick forward of neutral) was satisfactory. Adding the external wing tanks to cause the recovery characteristics to become critical and border on an unsatisfactory condition; however, it was shown that satisfactory recovery could be obtained by jettisoning the tanks, followed by normal recovery technique.

Description: Two theoretical procedures are developed for designing asymmetric supersonic nozzles for which the calculated exit flow is nearly uniform over a range of Mach numbers. One procedure is applicable at Mach numbers less than approximately 3. This approach yields, without iteration, a nozzle for which the calculated exit flow is uniform at two Mach numbers and, with proper design, is nearly uniform at Mach numbers between, slightly above, and slightly below these two. The use of an inclined and curved sonic line is an essential feature of this approach, The second procedure requires iteration and is used far designs at Mach numbers exceeding 3. Although it is not a necessary feature, an inclined and curved sonic line is also used in this procedure. In both approaches the flow field dawn stream of the sonic line is determined using the method of characteristics.

Description: Preliminary results of one phase of a control-motion study program involving several jet fighter-type airplanes are presented in time-history form and are summarized as maximum measured quantities plotted against indicated airspeed. The results pertain to approximately 1,000 maneuvers performed by a Republic F-84G jet-fighter airplane during squadron operational training. The data include most tactical maneuvers of which the F-84G airplane is capable. Maneuvers were performed at pressure altitudes of 0 to 30,000 feet with indicated airspeeds ranging from the stalling speed to approximately 515 knots.

Description: A flight test was made to determine the servoplane effectiveness and stability characteristics of the free-floating horizontal stabilizer to be used on the XF10F airplane. The results of this test indicate that servoplane effectiveness is practically constant through the speed range up to a Mach number of 1.15, and the stabilizer static stability is satisfactory. A loss of damping occurs over a narrow Mach number range near M = 1.0, resulting in dynamic instability of the stabilizer in this narrow range. Above M = 1.0 there is a gradual positive trim change of the stabilizer.

Description: An investigation of the 1XP excitation of inclined single-rotation propellers has indicated a new concept for determining propeller shaft forces and moments of an inclined propeller. This report presents preliminary results, in particular to the counterrotating propeller.

Description: A flight investigation of a 1/7-scale rocket-powered model of the XF10F Grumman XFl0F airplane in the swept-wing configuration has been made. The purpose of this test was to determine the static longitudinal stability, damping in pitch, and longitudinal control effectiveness of the airplane with the center of gravity at 20 percent of the wing mean aerodynamic chord. Only a small amount of data was obtained from the test because, immediately after booster separation at a Mach number of 0.88, the configuration was directionally unstable and diverged in sideslip. Simultaneous with the sideslip divergence, the model became longitudinally unstable at 3 degree angle of attack and -6 degree sideslip and diverged in pitch to a high angle of attack. During the pitch-up the free-floating horizontal tail became unstable at 5 degree angle of attack and the tail drifted against its positive deflection limit.

Description: A tank investigation has been conducted of a 1/10-size powered-dynamic model of the Edo model 142 hydra-ski research airplane. The results of tests of two configurations are presented: One included a large ski and a ski well; the other, a small ski without a well. Water take-offs would be possible with the available thrust for either configuration: however, the configuration with the large ski emerged sooner and had less resistance from ski emergence until take-off. Longitudinal stability and landing behavior in smooth water were satisfactory for both configurations. Some alteration to the design of the tail would be desirable in order to reduce the spray loads.

Description: Tests have been made at the Langley Aeronautical Laboratory on a 6000-horsepower propeller dynamometer installed at a ground test facility to determine the effect of a half-scale model of the Wright Aeronautical Development Center 30,000-horsepower whirl rig upon the aerodynamic characteristics of a three-blade NACA 10-(3)(062)-045 propeller. The model of the whirl rig was mounted in front of the 6000-horsepower propeller dynamometer. Static propeller tests were made for 0deg, 5deg, 10deg, 15deg, and 20deg blade angles over a range of rotational speeds from 600 to 2200 rpm in 100-rpm increments. Measurements were made of propeller thrust and torque, stresses in the propeller blades, and static and total pressures over the surface of the model. Propeller thrust and torque were increased up to 33 percent by the presence of the model of the whirl rig, but the average increase was from 5 to 10 percent. Blade vibratory stresses were small.

Description: The aerodynamic characteristics in pitch of the Army Ordnance Corps T205 3.5-inch HEAT rocket with various head designs and one fin modification have been determined at velocities of 500, 700 and 900 feet per second in the Langley high-speed 7- by 10-foot tunnel. The results presented are those of the full-scale model. Comparison of results obtained at 500 feet per second shows, in general, that for changes on the forward portion of the head the missile configurations having the greatest stability - most rearward center-of-loads location - were those having the highest drag. However, very limited comparisons indicate that the shape of the rear position of the head may be an important factor in reducing the drag and increasing the restoring moments. Generally, large increases in drag were noted for the various head designs with an increase in Mach number from 0.62 to 0.82. Pitching-moment-curve slopes increased with Mach number on all models except those having reasonably well-faired forward sections. These models showed a decrease in stability with increases in Mach number.

Description: An investigation of the static and dynamic longitudinal stability characteristics of 1/3.7 scale rocket-powered model of the Bell MX-776A has been made for a Mach number range from 0.8 to 1.6. Two models were tested with all control surfaces at 0 degree deflection and centers of gravity located 1/4 and 1/2 body diameters, respectively, ahead of the equivalent design location. Both models were stable about the trim conditions but did not trim at 0 degree angle of attack because of slight constructional asymmetries. The results indicated that the variation of lift and pitching moment was not linear with angle of attack. Both lift-curve slope and pitching-moment-curve slope were of the smallest magnitude near 0 degree angle of attack. In general, an increase in angle of attack was accompanied by a rearward movement of the aerodynamic center as the rear wing moved out of the downwash from the forward surfaces. This characteristic was more pronounced in the transonic region. The dynamic stability in the form of total damping factor varied with normal-force coefficient but was greatest for both models at a Mach number of approximately 1.25. The damping factor was greater at the lower trim normal-force coefficients except at a Mach number of 1.0. At that speed the damping factor was of about the same magnitude for both models. The drag coefficient increased with trim normal-force coefficient and was largest in the transonic region.

Description: An experimental investigation has been conducted to determine the dynamic stability and control characteristics of a 0.13-scale free-flight model of the Convair XFY-1 airplane in test setups representing the setup proposed for use in the first flight tests of the full-scale airplane in the Moffett Field airship hangar. The investigation was conducted in two parts: first, tests with the model flying freely in an enclosure simulating the hangar, and second, tests with the model partially restrained by an overhead line attached to the propeller spinner and ground lines attached to the wing and tail tips. The results of the tests indicated that the airplane can be flown without difficulty in the Moffett Field airship hangar if it does not approach too close to the hangar walls. If it does approach too close to the walls, the recirculation of the propeller slipstream might cause sudden trim changes which would make smooth flight difficult for the pilot to accomplish. It appeared that the tethering system proposed by Convair could provide generally satisfactory restraint of large-amplitude motions caused by control failure or pilot error without interfering with normal flying or causing any serious instability or violent jerking motions as the tethering lines restrained the model.

Description: An experimental investigation has been made in the Langley stability tunnel at low speed to determine the static longitudinal and lateral stability characteristics of a l/9-scale powered model of the Convair XFY-1 vertically rising airplane. Effects of thrust coefficient were investigated for the complete model and for certain components of the model. Effects of control deflections and of propeller-blade angle were investigated briefly for the complete model. Most of the tests were made through an angle-of-attack range from about -4 deg. to 29 deg, and the thrust-coefficient range was from 0 t o 0.7. In order to expedite distribution of these data to interested persons, no analysis of the data has been prepared for this report,

Description: Investigations have been conducted to determine by means of total-pressure surveys the boundaries of single and twin jets discharging through convergent nozzles into quiescent air. The jet boundaries for the region from the nozzle outlets to a station 6 nozzle diameters downstream are presented for nozzle pressure ratios ranging from 2.5 t o 16.0 and for twin-Jet nozzle center-line spacings ranging from 1.42 to 2.50 nozzle diameters. The effects of these parameters on the interaction of twin Jets are discussed. In order to ascertain the utility of the results for other than the test conditions, the effects of jet temperature, Reynolds number, and humidity on the pressure boundaries have been briefly investigated. The result indicate that for a jet of 2.6 the pressure boundaries are slightly smaller than those of corresponding unheated jets and that the effects of Reynolds number and humidity are negligible.

Description: The two-dimensional motion of an incompressible fluid about a closed contour with a definite velocity in magnitude and direction at infinity is considered. If, without changing the direction of the velocity at infinity, the magnitude is increased, the configuration of the streamlines remains unchanged and only the numbering of the stream function changes. There exists only one family of curves that can serve as streamlines in the incompressible flow about a given contour (at a given angle of attack); for example, the contour of an airplane wing. The case is quite different with a compressible fluid.

Description: The hydrodynamic characteristics of a preliminary design of the Martin XP6M-1 flying boat have been determined. Longitudinal stability during take-off and landing, resistance of the complete model, and behavior during taxiing and landing in rough water are presented.

Description: Calculations have been made to find range? attainable by bombers of gross weights from l40,000 to 300,000 pounds powered by turbine-propeller power plants. Only conventional configurations were considered and emphasis was placed upon using data for structural and aerodynamic characteristics which are typical of modern military airplanes. An effort was made to limit the various parameters invoked in the airplane configuration to practical values. Therefore, extremely high wing loadings, large amounts of sweepback, and very high aspect ratios have not been considered. Power-plant performance was based upon the performance of a typical turbine-propeller engine equipped with propellers designed to maintain high efficiencies at high-subsonic speeds. Results indicated, in general, that the greatest range, for a given gross weight, is obtained by airplanes of high wing loading, unless the higher cruising speeds associated with the high-wing-loading airplanes require-the use of thinner wing sections. Further results showed the effect of cruising at-high speeds, of operation at very high altitudes, and of carrying large bomb loads.