ALBERT

Description: Charts are presented for computing the thrust, fuel consumption, and other performance values of a turbojet engine for any given set of operating conditions and component efficiencies. The effects of the pressure losses in the inlet duct and combustion chamber, the variation in the physical properties of the gas as it passes through the cycle, and the change in mass flow by the addition of fuel are included. The principle performance charts show the effects of the primary variables and correction charts provide the effects of the secondary variables.

Description: The performance of hypothetical turbojet systems, without thrust augmentation, as power plants for supersonic airplanes has been calculated. The thrust, thrust power, air-fuel ratio, 1 specific fuel consumption, cross-sectional area, and thrust coefficient are shown for free-stream Mach numbers from 1.2 to 3. For comparison, the performance of ram-jet systems over the same Mach number range has also been calculated. For Mach numbers between 1.2 and 2 the calculated thrust coefficient of the turbojet system was found to be larger than the estimated drag coefficient, and the specific fuel consumption was calculated to be considerably less than the specific fuel consumption of the ram-jet system. The turbojet system therefore appears to merit consideration as a propulsion method for free-stream Mach numbers between approximately 1.2 and 2.

Description: An equation is presented for calculating the heat flow required from the surface of an internally heated windshield in order to prevent the formation of ice accretions during flight in specified icing conditions. To ascertain the validity of the equation, comparison is made between calculated values of the heat required and measured values obtained for test windshields in actual flights in icing conditions. The test windshields were internally heated and provided data applicable to two common types of windshield configurations; namely the V-type and the type installed flush with the fuselage contours. These windshields were installed on a twin-engine cargo airplane and the icing flights were conducted over a large area of the United States during the winters of 1945-46 and 1946-47. In addition to the internally heated windshield investigation, some test data were obtained for a windshield ice-prevention system in which heated air was discharged into the windshield boundary layer. The general conclusions resulting from this investigation are as follows: 1) The amount of heat required for the prevention of ice accretions on both flush- and V-type windshields during flight in specified icing conditions can be calculated with a degree of accuracy suitable for design purposes. 2) A heat flow of 2000 to 2500 Btu per hour per square foot is required for complete and continuous protection of a V-type windshield in fight at speeds up to 300 miles per hour in a moderate cumulus icing condition. For the same degree of protection and the same speed range, a value of 1000 Btu per hour per square foot suffices in a moderate stratus icing condition. 3) A heat supply of 1000 Btu per hour per square foot is adequate for a flush windshield located well aft of the fuselage stagnation region, at speeds up to 300 miles per hour, for flight in both stratus and moderate cumulus icing conditions. 4) The external air discharge system of windshield thermal ice prevention is thermally inefficient and requires a heat supply approximately 20 times that required for an internal system having the same performance.

Description: Convenient charts are presented for computing the thrust, fuel consumption, and other performance values of a turbojet system. These charts take into account the effects of ram pressure, compressor pressure ratio, ratio of combustion-chamber-outlet temperature to atmospheric temperature, compressor efficiency, turbine efficiency, combustion efficiency, discharge-nozzle coefficient, losses in total pressure in the inlet to the jet-propulsion unit and in the combustion chamber, and variation in specific heats with temperature. The principal performance charts show clearly the effects of the primary variables and correction charts provide the effects of the secondary variables. The performance of illustrative cases of turbojet systems is given. It is shown that maximum thrust per unit mass rate of air flow occurs at a lower compressor pressure ratio than minimum specific fuel consumption. The thrust per unit mass rate of air flow increases as the combustion-chamber discharge temperature increases. For minimum specific fuel consumption, however, an optimum combustion-chamber discharge temperature exists, which in some cases may be less than the limiting temperature imposed by the strength temperature characteristics of present materials.

Description: The icing characteristics, the de-icing rate with hot air, and the effect of impact ice on fuel metering and mixture distribution have been determined in a laboratory investigation of that part of the engine induction system consisting of a three-barrel injection-type carburetor and a supercharger housing with spinner-type fuel injection from an 18-cylinder radial engine used on a large twin-engine cargo airplane. The induction system remained ice-free at carburetor-air temperatures above 36 F regardless of the moisture content of the air. Between carburetor-air temperatures of 32 F and 36 F with humidity ratios in excess of saturation, serious throttling ice formed in the carburetor because of expansion cooling of the air; at carburetor-air temperatures below 32 F with humidity ratios in excess of saturation, serious impact-ice formations occurred, Spinner-type fuel injection at the entrance to the supercharger and heating of the supercharger-inlet elbow and the guide vanes by the warn oil in the rear engine housing are design features that proved effective in eliminating fuel-evaporation icing and minimized the formation of throttling ice below the carburetor. Air-flow recovery time with fixed throttle was rapidly reduced as the inlet -air wet -bulb temperature was increased to 55 F; further temperature increase produced negligible improvement in recovery time. Larger ice formations and lower icing temperatures increased the time required to restore proper air flow at a given wet-bulb temperature. Impact-ice formations on the entrance screen and the top of the carburetor reduced the over-all fuel-air ratio and increased the spread between the over-all ratio and the fuel-air ratio of the individual cylinders. The normal spread of fuel-air ratio was increased from 0.020 to 0.028 when the left quarter of the entrance screen was blocked in a manner simulating the blocking resulting from ice formations released from upstream duct walls during hot-air de-icing.

Description: Flight tests were made in natural icing conditions with two 8-ft-chord heated airfoils of different sections. Measurements of meteorological variables conducive to ice formation were made simultaneously with the procurement of airfoil thermal data. The extent of knowledge on the meteorology of icing, the impingement of water drops on airfoil surfaces, and the processes of heat transfer and evaporation from a wetted airfoil surface have been increased to a point where the design of heated wings on a fundamental, wet-air basis now can be undertaken with reasonable certainty.

Description: A theoretical analysis of the radial temperature distribution through the rotor and constant cross sectional area blades near the coolant passages of liquid cooled gas turbines was made. The analysis was applied to obtain the rotor and blade temperatures of a specific turbine using a gas flow of 55 pounds per second, a coolant flow of 6.42 pounds per second, and an average coolant temperature of 200 degrees F. The effect of using kerosene, water, and ethylene glycol was determined. The effect of varying blade length and coolant passage lengths with water as the coolant was also determined. The effective gas temperature was varied from 2000 degrees to 5000 degrees F in each investigation.

Description: A theoretical analysis of the cross-sectional temperature distribution of a water-cooled turbine blade was made using the relaxation method to solve the differential equation derived from the analysis. The analysis was applied to specific turbine blade and the studies icluded investigations of the accuracy of simple methods to determine the temperature distribution along the mean line of the rear part of the blade, of the possible effect of varying the perimetric distribution of the hot gas-to -metal heat transfer coefficient, and of the effect of changing the thermal conductivity of the blade metal for a constant cross sectional area blade with two quarter inch diameter coolant passages.

Description: The performance at inlet pressure of 21 inches mercury absolute and inlet temperature of 538 R for the 10-stage axial-flow X24C-2 compressor from the X24C-2 turbojet engine was investigated. the peak adiabatic temperature-rise efficiency for a given speed generally occurred at values of pressure coefficient fairly close to 0.35.For this compressor, the efficiency data at various speeds could be correlated on two converging curves by the use of a polytropic loss factor derived.

Description: On the basis of the investigations so far completed on the behavior of PTL power plants under various operating conditions, in which the influence of the propeller characteristics is of considerable importance, the most important aspects of a control system for turbine-propeller jet power plants are deduced. A simple possible means for its concrete realization, which is also applicable to TL [NACA comment: TL, jet] power plants, is presented by means of examples. A control device of this kind is now being developed.

Description: A theoretical analysis of the temperature distribution through the trailing portion of a blade near the coolant passages of liquid cooled gas turbines was made. The analysis was applied to obtain the hot spot temperatures at the trailing edge and influence of design variables. The effective gas temperature was varied from 2000 degrees to 5000 degrees F in each investigation.

Description: Axial blowers are gaining importance as aircraft engine superchargers. However, the pressure head obtainable per stage is small. Due to the necessary great number of stages, the physical length of the blower becomes too great for an airworthy device. This report discusses several types of construction that permit a reduction in the length of the blower.

Description: An altitude-wind-tunnel investigation of a TG-100A gas turbine-propeller engine was performed. Pressure and temperature data were obtained at altitudes from 5000 to 35000 feet, compressor inlet ram-pressure ratios from 1.00 to 1.17, and engine speeds from 800 to 13000 rpm. The effect of engine speed, shaft horsepower, and compressor-inlet ram-pressure ratio on pressure and temperature distribution at each measuring station are presented graphically.

Description: Rim cracking in turbine wheels with welded blades was evaluated. The problem is explained on the basis of the occurrence of plastic flow in the rim during transient starting conditions when thermal compressive stresses resulting from high-temperature gradients exceed the proportional elastic limit of the material.

Description: An investigation has been made in the Langley two-dimensional low-turbulence pressure tunnel to develop the optimum configuration of a 0.35-chord slotted flap on an NACA 65(sub 1120)-111 airfoil section modified by removing the trailing-edge cusp. The section pitching-moment characteristics and the effects of standard roughness on the section characteristics were determined for the flap retracted at Reynolds numbers ranging from 3.0 x 10(exp 6) to 9.0 x 10(exp 6).

Description: Powered models of three different flying boats were landed in oncoming wave of various heights and lengths. The resulting motions and acceleration were recorded to survey the effects of varying the trim at landing, the deceleration after landing, and the size of the waves. One of the models had an unusually long afterbody. The data for landing with normal rates of deceleration indicated that the most severe motions and accelerations were likely to occur at some period of the landing run subsequent to the initial impact. Landings made at abnormally low trims led to unusually severe bounces during the runout. The least severe landing occurred after a small lending when the model was rapidly decelerated at about 0.4 g in a simulation of the proposed use of braking devices. The severity of the landings increased with wave height and was at a maximum when the wave length was of the order of from one and one-half to twice the over-all length of the model. The models with afterbodies of moderate length frequently bounced clear of the water into a stalled attitude at speeds below flying speed. The model with the long afterbody had less tendency to bounce from the waves and consequently showed less severe accelerations during the landing run than the models with moderate lengths of afterbody.

Description: Theoretical pressure distributions and measured lift, drag, and pitching moment characteristics at three values of Reynolds number are presented for a group of NACA four-digit-series airfoil sections modified for high-speed applications. The effectiveness of flaps applied to these airfoils and the effect of standard leading-edge roughness were also investigated at one value of Reynolds number. Results are also presented of tests of three conventional NACA four-digit-series airfoil sections.

Description: While the gas turbine by itself has been applied in particular cases for power generation and is in a state of promising development in this field, it has already met with considerable success in two cases when used as an exhaust turbine in connection with a centrifugal compressor, namely, in the supercharging of combustion engines and in the Velox process, which is of particular application for furnaces. In the present paper the most important possibilities of combining a combustion engine with a gas turbine are considered. These "combination engines " are compared with the simple gas turbine on whose state of development a brief review will first be given. The critical evaluation of the possibilities of development and fields of application of the various combustion engine systems, wherever it is not clearly expressed in the publications referred to, represents the opinion of the author. The state of development of the internal-combustion engine is in its main features generally known. It is used predominantly at the present time for the propulsion of aircraft and road vehicles and, except for certain restrictions due to war conditions, has been used to an increasing extent in ships and rail cars and in some fields applied as stationary power generators. In the Diesel engine a most economical heat engine with a useful efficiency of about 40 percent exists and in the Otto aircraft engine a heat engine of greatest power per unit weight of about 0.5 kilogram per horsepower.

Description: After defining the aims and requirements to be set for a control system of gas-turbine power plants for aircraft, the report will deal with devices that prevent the quantity of fuel supplied per unit of time from exceeding the value permissible at a given moment. The general principles of the actuation of the adjustable parts of the power plant are also discussed.

Description: An investigation of the spin and recovery characteristics of a 1/24-scale model of the McDonnell XP-88 airplane has been conducted in the Langley 20-foot free-spinning tunnel. The effects of control settings and movements on the erect and inverted spin and recovery characteristics of the model in the normal loading were determined. Tests of the model in the long-range loading also were made. The investigation included tail-modification, spin-recovery parachute, pilot-escape, and rudder-pedal-force tests. Recoveries were generally satisfactory for spins in the normal loading provided the ailerons were not held against the spin. Satisfactory recoveries were obtained regardless of the aileron setting when the leading-edge flaps were deflected and normal recovery technique was used or when the horizontal tail was raised 70 inches, full scale. Recoveries were rapid from all inverted spins obtained. In the long-range loading with tanks on, it may be necessary to jettison the tanks in order to obtain recovery. A 12.0-foot spin-recovery parachute at the tail or a 4.0-foot parachute opened on the outer wing tip (drag coefficient of 0.66) was found to be effective for recoveries from demonstration spins. Test results showed that in an emergency the pilot should attempt to escape from the outboard side of the spinning airplane. The rudder-pedal forces in a spin were indicated to be within the capabilities of the pilot.

Description: From flight tests of 0.5-scale models of the Fairchild Lark pilotless aircraft conducted at the flight test station of the Pilotless Aircraft Research Division at Wallops Island, Va., some evaluations of the static longitudinal stability were obtained by analysis of the short-period oscillations induced by the abrupt movement of the rudder elevators. The analysis shows that for the Lark configuration with wing flap deflections of 0 degrees and 15 degrees the static longitudinal stability decreases slightly up to the critical Mach number and than as the Mach number increases further the stability increases greatly.

Description: At the request of the Air Materiel Command, Army Air Forces an investigation of the low-speed, power-off stability and control characteristics of the McDonnell XP-85 airplane is being conducted in the Langley free-flight tunnel. The XP-85 airplane is a parasite fighter carried in a bomb bay of the B-36 airplane. As a part of the investigation a few force tests were made of a 1/5 scale model of the XP-85 with a conventional tail assembly installed in place of the original design five-unit tail assembly. The total area of the conventional assembly was approximately 80 percent of the area of the five-unit assembly. The results of this investigation showed that the conventional tail assembly gave about the same longitudinal stability characteristics as the original configuration and improved the directional and lateral stability.

Description: Flight tests to determine propeller performance have been made of a Curtiss No. 838-102-18 three-blade propeller having trailing-edge extensions on a Republic P-47D-28 airplane in climb and high speed. These tests are a part of a general propeller flight-test program at the Langley Laboratory of the National Advisory Committee for Aeronautics. Results of climb tests indicate that when power is changed from approximately 1475 horsepower at 2550 rpm (roughly normal power) to 2400 horsepower at 2700 rpm (approximately military power) there is a loss in propeller efficiency of 3 percent at an altitude of 7000 feet, and 4 percent at 21,000 feet. At an airplane Mach number of 0.7 there is a gain of 9 percent in propeller efficiency when the power coefficient per blade is increased from 0.06 to 0.09. Optimum power coefficient per blade at this Mach number is estimated to be approximately 0.12. An analysis to determine the effect of the addition of extensions on the performance of the basic propeller blades indicates that climb performance was increased but high-speed performance was reduced. Both effects, however, were small.

Description: The performance of the 11-stage axial-flow compressor, modified to improve the compressor-outlet velocity, in a revised X24C-4B turbojet engine is presented and compared with the performance of the compressor in the original engine. Performance data were obtained from an investigation of the revised engine in the MACA Cleveland altitude wind tunnel. Compressor performance data were obtained for engine operation with four exhaust nozzles of different outlet area at simulated altitudes from 15,OOO to 45,000 feet, simulated flight Mach numbers from 0.24 to 1.07, and engine speeds from 4000 to 12,500 rpm. The data cover a range of corrected engine speeds from 4100 to 13,500 rpm, which correspond to compressor Mach numbers from 0.30 to 1.00.

Description: An investigation of two 1/14 scale model configurations of an outboard nacelle for the XB-36 airplane was made in the Langley two-dimensional low-turbulence tunnels over a range of airplane lift coefficients (C (sub L) = 0.409 to C(sub L) = 0.943) for three representative flow conditions. The purpose of the investigation was to develop a low-drag wing-nacelle pusher combination which incorporated an internal air-flow system. The present investigation has led to the development of a nacelle which had external drag coefficients of similar order of magnitude to those obtained previously from tests of an inboard nacelle configuration at the corresponding operating lift coefficients and from approximately one-third to one-half of those of conventional tractor designs having the same ratio of wing thickness to nacelle diameter.

Description: Tests have been conducted in the Langley high-speed 7- by 10-foot tunnel over a Mach number range from 0.40 to 0.91 to determine the stability and control characteristics of an 0.08-scale model of the Chance Vought XF7U-1 airplane. The basic lateral stability characteristics of the complete model with undeflected control surfaces are presented in the present report with a very limited analysis of the results.

Description: This report contains the flight-test results of the stalling characteristics measured during the flying-qualities investigation of the Lockheed P-8OA airplane (Army No. 44-85099). The tests were conducted in straight and turning flight with and without wing-tip tanks. These tests showed satisfactory stalling characteristics and adequate stall warning for all configurations and conditions tested.

Description: Requirements of an automatic engine control, as affected by engine characteristics, have been analyzed for a direct-coupled turbojet engine. Control parameters for various conditions of engine operation are discussed. A hypothetical engine control is presented to illustrate the use of these parameters. An adjustable speed governor was found to offer a desirable method of over-all engine control. The selection of a minimum value of fuel flow was found to offer a means of preventing unstable burner operation during steady-state operation. Until satisfactory high-temperature-measuring devices are developed, air-fuel ratio is considered to be a satisfactory acceleration-control parameter for the attainment of the maximum acceleration rates consistent with safe turbine temperatures. No danger of unstable burner operation exists during acceleration if a temperature-limiting acceleration control is assumed to be effective. Deceleration was found to be accompanied by the possibility of burner blow-out even if a minimum fuel-flow control that prevents burner blow-out during steady-state operation is assumed to be effective. Burner blow-out during deceleration may be eliminated by varying the value of minimum fuel flow as a function of compressor-discharge pressure, but in no case should the fuel flow be allowed to fall below the value required for steady-state burner operation.

Description: The present treatise reports on theoretical investigations and test-stand measurements which were carried out in the BMW Flugmotoren GMbH in developing the hollow blade for exhaust gas turbines. As an introduction the temperature variation and the stress on a turbine blade for a gas temperature of 900 degrees and circumferential velocities of 600 meters per second are discussed. The assumptions onthe heat transfer coefficients at the blade profile are supported by tests on an electrically heated blade model. The temperature distribution in the cross section of a blade Is thoroughly investigated and the temperature field determined for a special case. A method for calculation of the thermal stresses in turbine blades for a given temperature distribution is indicated. The effect of the heat radiation on the blade temperature also is dealt with. Test-stand experiments on turbine blades are evaluated, particularly with respect to temperature distribution in the cross section; maximum and minimum temperature in the cross section are ascertained. Finally, the application of the hollow blade for a stationary gas turbine is investigated. Starting from a setup for 550 C gas temperature the improvement of the thermal efficiency and the fuel consumption are considered as well as the increase of the useful power by use of high temperatures. The power required for blade cooling is taken into account.

Description: Computations were made to determine the temperature distribution and cooling of solid gas-turbine blades.A range of temperatures was used from 1500 degrees to 2500 degrees F, blade-root temperatures from 100 degrees to 1000 degrees F, blade thermal conductivity from 8 to 220 BTU/(hr)(sq ft)(degrees F/ft), and net gas to metal heat transfer coefficients from 75 to 250 BTU/(hr)(sq ft)(degrees F).

Description: Effect of inlet-air pressure and temperature on the performance of the X24-2 10-Stage Axial-Flow Compressor from the X24C-2 turbojet engine was evaluated. Speeds of 80, 89, and 100 percent of equivalent design speed with inlet-air pressures of 6 and 12 inches of mercury absolute and inlet-air temperaures of approximately 538 degrees, 459 degrees,and 419 degrees R ( 79 degrees, 0 degrees, and minus 40 degrees F). Results were compared with prior investigations.

Description: A calulation of the flow in turbine blading is reported that includes the calculation of effect of centrifugal force. Frictional losses on the stator blades and rotor blades are allowed.

Description: An investigation of the antiknock effectiveness of various additive-water solutions when used as internal coolants has been conducted at the NACA Cleveland laboratory. Nine compounds have been previously run in a CFR engine and the results are presented. In an effort to find a good anti-knock-coolant additive with more desirable physical properties than those of the nine compounds previously investigated, water solutions of four alkyl amines, three alkanolamines, six amides, and eight heterocyclic compounds were investigated and the results are presented.

Description: Qualitative investigations have shown that use of the NACA injection impeller with the R-3350 engine increases the inertia of the fuel-injection system and, when the standard fuel-metering system is used, this increase in inertia results in poor engine acceleration characteristics. This investigation was therefore undertaken to determine whether satisfactory acceleration characteristics of the engine equipped with the injection impeller could be obtained by simple modifications to the fuel-monitoring system. The engine was operated with two types of carburetor; namely, a hydraulic-metering carburetor incorporating a vacuum-operated accelerating pump and a direct-metering carburetor having a throttle-actuated accelerating pump. The vacuum-operated accelerating pump of the hydraulic-metering carburetor was modified to produce satisfactory accelerations by supplementing the standard air chamber with an additional 75-cubic spring. The throttle-actuated accelerating pump of the direct-metering carburetor was modified to produce satisfactory accelerations by replacing the standard 0.028-inch-diameter bleed in the load-compensator balance line with a smaller bleed of 0.0225-inch diameter. The results of this investigation indicated that both carburetors can be easily modified to produce satisfactory acceleration characteristics of the engine and no definite choice between the types of carburetor and accelerating pump can be made. Use of the direct-metering carburetor, however, probably resulted in better fuel distribution to the cylinders during the acceleration period and reduced the backfire hazard because all the fuel is introduced through the injection impeller.

Description: A Douglas C-74 airplane, during a test dive at about 0.525 Mach number, experienced uncontrollable longitudinal oscillations sufficient to cause shedding of the outer wing panels and the subsequent crash of the airplane. Tests of a section of the horizontal tail plane from a C-74 airplane were conducted in the Ames 16-foot high-speed wind tunnel to investigate the possibility of the tail as a contributing factor to the accident. The results of the investigations of fabric-covered elevators in various conditions of surface deformation are presented in this report.

Description: Attempts were made to alleviate the buffeting of external fuel tanks mounted under the wings of a twin-engine Navy fighter plane. The Mach number at which the buffeting began was increased from 0.529 to 0.640 by streamlining the sway braces and increasing the lateral rigidity of the sway brace system. Further increases of the Mach number, at which buffeting began to 0.725, was obtained by moving the external fuel tank to a position under the fuselage.

Description: This report contains the flight-test results of the lateral and directional-stability and control phase (including tests with wing-tip tanks) of a general flying-qualities investigation of the Lockheed P-80A airplane (Army No. 44-85099). These tests were conducted at indicated airspeeds up to 494 miles per hour (0.691 Mach number) at low altitude and up to 378 miles per hour (0.816 Mach number) at high altitude. These tests showed that the flying qualities of the airplane were for the most part in accordance with the requirements of the Army Air Forces Stability and Control Specifications. The only major deficiency noted was the negative lateral stability with the wing-tip tanks installed.

Description: An investigation was conducted to determine the operational and performance characteristics of the TG-100A gas turbine-propeller engine II. Windmilling characteristics were deterined for a range of altitudes from 5000 to 35,000 feet, true airspeeds from 100 to 273 miles per hour, and propeller blade angles from 4 degrees to 46 degrees.

Description: The NACA 23012-4 airfoil was investigated for the purpose of increasing lift by means of blowing out air from the wing, in conjunction with the effect of plain flap of variable contour and slotted flap of 25 percent chord length. The wing also was provided with a hinged nose, to be deflected at will. Air was blown out frcm the wing immediately in front of the flap; also at the opening between wing and hinged nose,tangentially to the surface of the wing. Another device employed to increase maximum lift was a movable slat, to be opened to form a clot. Lift was measured in relation to the volume of blown-out air and considerable increases were observed with increasing volume.

Description: The sea-level performance of I-16 turbojet engine at zero ram was investigated to determine the effects of an intake duct, shroud, and tail pipe intended for installation in an XFR-1 airplane. Engine speeds ranged from 8000 to 16,500 rpm for several variations of the intake duct and tail pipes.

Description: Flight and ground investigations have been made to compare an exhaust-ejector installation with a standard exhaust-collector-ring installation on air-cooled aircraft engines in a twin-engine airplane. The ground investigation allowed that, whereas the standard engine would have overheated above 600 horsepower, the engine with exhaust ejectors cooled at take-off operating conditions at zero ram. The exhaust ejectors provided as much cooling with cowl flaps closed as the conventional cowl flaps induced when full open at low airspeeds. The propulsive thrust of the exhaust-ejector installation was calculated to be slightly less than the thrust of the collector-ring-installation.

Description: The performance of a mixed-flow impeller in combination with a semivaneless diffuser were experimentally investigated. The diameter of the impeller was 11.0 inches and a maximum tip diameter of 14.74 inches. The semivaneless diffuser had an overall diameter of 28.00 inches. The performance properties of the mixed-flow impeller were also investigated with a 34.00 inch vane loss diffuser having a transition section of the same geometry as the semivaneless diffuser.

Description: The calculation of infinitesimal conical supersonic flow has been applied first to the simplest examples that have also been calculated in another way. Except for the discovery of a miscalculation in an older report, there was found the expected conformity. The new method of calculation is limited more definitely to the conical case.

Description: The Russian AM 35 and AM 38 aircraft engines have superchargers with a swirl throttle, which appears to be a purely Russian development. This paper gives the results of test runs of the two engines, including the effects of the swirl throttle on engine performance.

Description: A study was made of heat transfer in turbine blades and the effects on blade temperature of cooling the blade root and tip, changing the dimensions of the blades, raising the cycle temperatures, insulating with ceramics, and cooling by circulation of air or water through hollow blades.

Description: Four methods of boundary-layer control were tried during an investigation to improve the flow in the impeller passages of a V-1710-93 engine-stage supercharger. The boundary layer along the impeller front shroud was removed by suction. In one method the removal was accomplished by recirculation of the air to the impeller inlet; in another method, by external removal. In the other methods, slots were cut through the impeller-blade faces first at 30 percent and then at 30 and 70 percent of the mean-flow-path length measured from leading edges of the rotating inlet guide vanes to introduce air from the high-pressure side of the blades into the region where stagnation and separation were suspected. A slight improvement in performance was obtained when the boundary layer was removed through the impeller front shroud. In general, this improvement become more pronounced as the amount of air removed was increased even though the excessive impeller frontal clearance maintained for these tests, together with an exaggerated negative pressure gradient, apparently induced flow separation on the diffuser front and rear walls as well as on the impeller front shroud. The use of slots in the impellers at the locations selected had a detrimental effect on the supercharger performance characteristics.

Description: An investigation was conducted on a multicylinder aircraft engine on a dynamometer stand to determine the effect of induction-system icing on engine operating characteristics and to compare the results with those of a previous laboratory investigation in which only the carburetor and the engine-stage supercharger assembly from the engine were used. The experiments were conducted at simulated glide power, low cruise power, and normal rated power through a range of humidity ratios and air temperatures at approximately sea-level pressure. Induction-system icing was found to occur within approximately the same limits as those established by the previous laboratory investigation after making suitable allowances for the difference in fuel volatility and throttle angles. Rough operation of the engine was experienced when ice caused a marked reduction in the air flow. Photographs of typical ice formations from this investigation indicate close similarity to icing previously observed in the laboratory.

Description: The basic principles of the control of TL ongincs are developed on .the basis of a quantitative investigation of the behavior of these behavior under various operating conditions with particular consideration of the simplifications pormissible in each case. Various possible means of control of jet engines are suggested and are illustrated by schematic designs.

Description: Tests of a 1/20-scale model of the Fleetwings XBTK-1 airplane have been performed in the Langley 15-foot free-spinning tunnel to determine the trim tendencies of the airplane at attitudes above the stall. The results of the tests indicated that the model would trim longitudinally only in the normal range of angles of attack and that the yaw trim tendencies for such longitudinal trim conditions were normal. Although wide oscillations in yaw were noted for some conditions, they occurred at angles of attack larger than those indicated as possible for longitudinal trim and spin equilibrium. It appears, therefore, that the oscillatory motions reported for the airplane may have been the direct result of control movements rather than the result of inherent oscillatory tendencies.

Description: It will be shown that by the use of the concept of similarity a simple representation of the characteristic curves of a compressor operating in combination with a turbine may be obtained with correct allowance for the effect of temperature. Furthermore, it bec~mes possible to simplify considerably the rather tedious investigations of the behavior of gas-turbine power plants under different operating conditions. Characteristic values will be derived for the most important elements of operating behavior of the power plant, which will be independent of the absolute valu:s of pressure and temperature. At the same time, the investigations provide the basis for scale-model tests on compressors and turbines.

Description: The spin and recovery characteristics of the Curtiss-Wright XP-87 airplane, as well as the spin-recovery parachute requirements, the control forces that would be encountered in the spin, and the best method for the crew to attempt an emergency escape, are presented in this report. The characteristics were estimated rather than determined by model tests because the XP-87 dimensional and mass characteristics were considered to be noncritical and because data were available from model tests of several similar airplanes. The study indicated that the recovery characteristics of the airplane will be satisfactory for all loadings if the controls are reversed fully and rapidly. The control forces, however, will probably be beyond the capabilities of the pilot unless some additional balance or a booster is used. A 6-foot tail parachute or a 3.5-foot wing-tip parachute with a drag coefficient of 0.7 will be a satisfactory, emergency spin-recovery device for spin demonstrations. If it is necessary for the crew to abandon the spinning airplane, they should leave from the outboard side of the cockpit.

Description: The results obtained from measurements of gust and draft velocities within thunderstorms for the period July 22, 1946 to July 23, 1946 at Orlando, Florida, are presented herein. These data are summarized in tables I and II, respectively, and are of the type presented in reference 1 for previous flights. Inspection of photo-observer records for the flights indicated that no data on ambient air temperature variations within thunderstorms were obtained.

Description: Tests with a dynamically similar model of the Navy XP4M-1 airplane were made to determine the best way to land the airplane in calm and rough water, to determine its probable ditching performance, and to determine practicable modifications which could be incorporated in the design of the airplane that would improve its ditching characteristics. The results were obtained by making visual observations, by recording longitudinal decelerations,a nd by taking motion pictures of the landings. A list of conclusions from the test results is included.

Description: An investigation has been made in the Langley two-dimensional low-turbulence tunnel to develop the optimum configuration of a 0.35-chord slotted flap on an NACA 65 (112)-111 airfoil section modified by removing the trailing-edge cusp. The results of the investigation indicate that for the optimum configuration at a Reynolds number of 2.4 x 10(exp 6), the flap deflection was 45 degrees and the flap leading-edge radius center was 0.73 percent-chord behind and 4.46 percent-chord below the slot lip. The maximum section lift coefficient for the optimum configuration at a Reynolds number of 2.4 x 10(exp 6) was 2.46 or 0.12 higher than that obtained for an NACA 65-210 airfoil section with a 0.250-chord slotted flap.

Description: This report presents the result of tests of a 0.35-scale model of the Bell P-39N-l airplane. Included are the longitudinal-stability and - control characteristics of the airplane as indicated by tests of the model equipped with each of two different sets of elevators. The results indicate good longitudinal stability and control throughout the speed range encounterable in flight. The variation of estimated stick force with speed was less when the model was equipped with elevators constructed to the theoretical design dimensions than when equipped with elevators as built to scale from measurements of the corresponding-parts of the actual airplane. The predicted stick forces required to produce the normal accelerations attainable in flight are within the limits specified by the Army Air Forces.

Description: Spin tests have been performed in the Langley 20-foot free-spinning tunnel on a 1/35-scale model of the Douglas XB-43 airplane. The spin and recovery characteristics were determined for several loading conditions of the airplane. The effects of installing a dorsal fin and of installing a ventral fin were investigated. Emergency escape of the crew was simulated and the stick and rudder pedal forces necessary to effect recoveries on the airplane were determined.

Description: A preliminary evaluation of the spin and recovery characteristics of the XF3D-1 airplane has been made, based primarily on the results of the free-spinning tunnel tests of a model which closely simulated the XF3D-1 in tail design, tail length, and mass loading. Estimates have been made of the rudder-pedal force that may be encountered in effecting recovery from a spin and of the spin recovery parachute requirements of the airplane for demonstration spins. The method of bail-out which should be used if it becomes necessary for the crew to abandon the airplane during a spin is indicated. It was indicated that the recovery characteristics of the XF3D-1 airplane in the clean condition for erect and inverted spins would be satisfactory for all loadings specified by the contractor as possible on the airplane. However, if a spin is inadvertently entered while the landing flaps are down, recovery may be slow. The slow-down brakes and the landing flaps should be retracted immediately upon the inception of a spinning condition, after which recovery from the spin should be attempted. The pedal force necessary to reverse the rudder during a spin will be within the physical capabilities of the pilot. Opening a 10-foot diameter parachute attached to the tail (laid-out-flat diameter, drag coefficient 0.7) or a 4.5-foot diameter parachute attached to the outboard wing tip will insure satisfactory spin recovery from demonstration spins. If it becomes necessary for the crew to abandon the airplane during a spin, they should leave from the outboard side of the cockpit.

Description: Tests of a 1/7 size model of the Grumman XJR2F-1 amphibian were made in Langley tank no.1 to examine the landing behavior in rough water and to measure the normal and angular accelerations experienced by the model during these landings. All landings were made normal to the direction of wave advance, a condition assumed to produce the greatest accelerations. Wave heights of 4.4 and 8.0 inches (2.5 and 4.7 ft, full size) were used in the tests and the wave lengths were varied between 10 and 50 feet (70 and 350 ft, full size). Maximum normal accelerations of about 6.5g were obtained in 4.4 inch waves and 8.5g were obtained in 8.0 inch waves. A maximum angular acceleration corresponding to 16 radians per second per second, full size, was obtained in the higher waves. The data indicate that the airplane will experience its greatest accelerations when landing in waves of about 20 feet (140 ft, full size) in length.

Description: In the following, high-speed measurements on a swept-back wing are reported. The curves of lift, moment, and drag have been determined up to Mach numbers of M = 0.87, and they are compared to a rectangular wing. Through measurements of the total-head loss behind the wing and through schlieren pictures, an insight into the formation of the compression shock at high Mach numbers has been obtained.

Description: Tests were conducted in calm water in Langley tank no. 2 and in calm and rough water at an outdoor catapult in order to determine the best way to make a forced landing of an Army A-26 airplane and to determine its probable ditching behavior. These tests were requested by the Air Materiel Command, Army Air Forces, in their letter of March 26, 1943, WEL:AW:50.

Description: A series of investigations of several 1/14-scale models of an inboard nacelle for the XB-36 airplane was made in the Langley two-dimensional low-turbulence tunnels. The purpose of these investigations was to develop a low-drag wing-nacelle pusher combination which incorporated an internal air-flow system. As a result of these investigations, a nacelle was developed which had external drag coefficients considerably lower than the original basic form with the external nacelle drag approximately one-half to two-thirds of those of conventional tractor designs. The largest reductions in drag resulted from sealing the gaps between the wing flaps and nacelle, reducing the thickness of the nacelle training-edge lip, and bringing the under-wing air inlet to the wing leading edge. It was found that without the engine cooling fan adequate cooling air would be available for all conditions of flight except for cruise and climb at 40,000 feet. Sufficient oil cooling at an altitude of 40,000 feet may be obtained by the use of flap-type exit doors.

Description: A 1/5.5-size powered dynamic model of the Columbia XJL-1 amphibian was landed in Langley tank no. 1 in smooth water and in oncoming waves of heights from 2.1 feet to 6.4 feet (full-size) and lengths from 50 feet to 264 feet (full-size). The motions and the vertical accelerations of the model were continuously recorded. The greatest vertical acceleration measured during the smooth-water landings was 3.1g. During landings in rough water the greatest vertical acceleration measured was 15.4g, for a landing in 6.4-foot by 165-foot waves. The impact accelerations increased with increase in wave height and, in general, decreased with increase in wave length. During the landings in waves the model bounced into the air at stalled attitudes at speeds below flying speed. The model trimmed up to the mechanical trim stop (20 deg) during landings in waves of heights greater than 2.0 feet. Solid water came over the bow and damaged the propeller during one landing in 6.4-foot waves. The vertical acceleration coefficients at first impact from the tank tests of a 1/5.5-size model were in fair agreement with data obtained at the Langley impact basin during tests of a 1/2-size model of the hull.

Description: Tests were made with a 1/16 size dynamically similar model of the Navy XP2V-1 airplane to study its performance when ditched. The model was ditched in calm water at the Langley tank no. 2 monorail. Various landing attitudes, speeds, and conditions of damage were simulated. The performance of the node1 was determined and recorded from visual observations, by recording time histories of the longitudinal decelerations, and by taking motion pictures of the ditchings From the results of the tests with the model the following conclusions were drawn: 1. The airplane should be ditched at the normal landing attitude. The flaps should be fully extended to obtain the lowest possible landing speed; 2. Extensive damage will occur in a ditching and the airplane probably will dive violently after a run of about 2 fuselage lengths. Maximum longitudinal decelerations up to about 4g will be encountered; and 3. If a trapezoidal hydroflap 4 feet by 2 feet by 1 foot is attached to the airplane at station 192.4, diving will be prevented and the airplane will probably porpoise in a run of about 4 fuselage lengths with a maximum longitudinal deceleration of less than 3.5g.

Description: An investigation has been conducted in the NACA Cleveland altitude wind tunnel to evaluate the performance characteristics of a modified X24C-4B turbojet engine over a range of simulated altitudes from 5000 to 45,000 feet, simulated flight Mach numbers from 0.25 to 1.07, and engine speeds from 4000 to 12,500 rpm. The engine was modified by the manufacturer to improve the velocity and temperature profiles within the engine. Performance data are graphically presented to show the effect of altitude at a flight Mach number of 0.25 and the effect of flight Mach number at an altitude of 25,000 feet. Original and modified engine performances for several specific operating conditions are compared. A complete tabulation of average pressures and temperatures throughout the engine, performance data, and lubrication and fuel-system data is presented.

Description: An investigation of the air-stream fluctuations at the tail of the D-558-1 airplane has been made at high speed for the purpose of determining the vertical region in which the horizontal tail may be placed without becoming subject to tail buffeting. The investigation was made for a range of Mach numbers from 0.775 to 0.907, and a range of vertical positions at the tall to include two proposed horizontal-tail positions. The tests were made at two angles of attack, 0,2 deg. and 4.2 deg., representative, of the angles of attack for high-speed level flight and a pull-out condition.

Description: This report contains the results of tests of a 1/3-scale model of the Lockheed YP-90A "Shooting Star" airplane and a comparison of drag, maximum lift coefficient, and elevator angle required for level flight as measured in the wind tunnel and in flight. Included in the report are the general aerodynamic characteristics of the model and of two types of dive-recovery flaps, one at several positions along the chord on the lower surface of the wing and the other on the lower surface of the fuselage. The results show good agreement between the flight and wind-tunnel measurements at all Mach numbers. The results indicate that the YP-80A is controllable in pitch by the elevators to a Mach number of at least 0.85. The fuselage dive-recovery flaps are effective for producing a climbing moment and increasing the drag at Mach numbers up to at least 0.8. The wing dive-recovery flaps are most effective for producing a climbing moment at 0.75 Mach number. At 0.85 Mach number, their effectiveness is approximately 50 percent of the maximum. The optimum position for the wing dive-recovery flaps to produce a climbing moment is at approximately 35 percent of the chord.

Description: Spin tests of a 1/16-scale model of the Chance Vought XF5U-1 airplane have been performed in the Langley 20-foot free-spinning tunnel. The effect of control position and movement upon the erect and inverted spin and recovery characteristics ae well as the effects of propellers, of stability flaps, and of various revisions to the design configuration have been determined for the normal fighter loading. The investigation also included spin recovery parachute, tumbling, and pilot-escape tests. For the original design configuration, with or without windmilling propellers, the recovery characteristics of the model were considered unsatisfactory. Increasing the maximum upward deflection of the ailavators from 45 deg to 65 deg resulted in greatly improved recovery characteristics. Dimensional revisions to the original airplane configuration, which satisfactorily improved the general spin and recovery characteristics of the model, consisted of: (1) a supplementary vertical tail 34 inches by 59 inches (full-scale) attached to a boom 80 inches aft of the trailing edge of the airplane in the plane of symmetry, (2) a large semispan undersurface spoiler placed along the airplane quarter-chord line and opened on the outboard side in a spin, or (3) two additional vertical tails 64 inches by 52 inches (full-scale) located at the tips of the ailavators. A satisfactory parachute arrangement for emergency spin recovery from demonstration spins was found to be an arrangement consisting of a 13.3-foot parachute attached by a 30-foot towline to the arresting gear mast on the airplane and opened simultaneously with an 8-foot parachute on the outboard end of the wing attached by a 3-foot towline. Tests indicated that pilot escape from a spin would be extremely hazardous unless the pilot is mechanically ejected from the cockpit. Model tumbling tests indicated that the airplane would not tumble.

Description: Spin tests have been conducted in the Langley free-spinning tunnel on a 1/16-scale model of the McDonnell XP-85 airplane with the normal X-tail replaced with a short-coupled conventional tail arrangement. The effect of the conventional tail arrangement and the effects of various modifications upon the spin and recovery characteristics of the model were determined. The results of the tests indicated that installation of the conventional tail arrangement wil not provide satisfactory recoveries from spins of the airplane. Satisfactory recoveries will be obtainable, however, either by installing in addition a very large ventral fin (17.94 sq ft, full-scale) below the tail or by decreasing the width of the fuselage and making it flat sided rearward of the wing trailing edge.

Description: This report presents the results of wind-tunnel tests of a 0.22-scale model of the North American XP-82 airplane with several modifications designed to reduce the buffeting of the airplane. The effects of various modifications on the air flow over the model are shown by means of photographs of tufts. The drag, lift, and pitching-moment coefficients of the model with several of the modifications are shown. The result indicate that, by reflexing the trailing edge of the center section of the wing and modifying the radiator air-scoop gutter and the inboard lower-surface wing fillets, the start of buffeting can be delayed from a Mach number of 0.70 to 0.775, and that the diving tendency of the airplane would be eliminated up to a Mach number of 0.80.

Description: Measurements are described which were taken in the large wind tunnel of the AVA on a rectangular wing "Mustang 2" with nose flap of a chord of 10 percent. Besides force measurements the results of pressure-distribution measurements are given and compared with those on the same profile "without" nose flap.

Description: The results obtained from measurements of gust velocities, draft velocities, and ambient-air temperature within thunderstorms for the period from September 11, 1946 to September 16, 1946 at Orlando, Florida are presented herein. These data are summarized in.and presented.

Description: This report presents comprehensive pressure-distribution measurements on four (4) swept-back wings (phi = 0 deg, 15 deg, 30 deg, and 45 deg) of constant chord and over a large range of angles of attack with symmetrical air flow. The distributions, experimentally obtained, were compared with theoretical ones calculated by the methods of Weissinger and Multhopp.

Description: A low-speed investigation in the Langley propeller-research tunnel of annular air inlets designed to avoid compression shocks and attendant boundary-layer separation on the fuselage ahead of the inlets at transonic flight speeds by maintaining substream flow velocities on the fuselage nose was reported in NACA RM No. L6J04. In the present investigation, one of the original annular inlets was converted by the installation of a canopy and a nose-wheel fairing into a twin side inlet in order to study problems involved in applying such an inlet to a fighter-type airplane. Extensive measurements of pressures on the surface of the model and surveys of the internal flow were conducted at angles of attack of 0 degrees, 3 degrees, and 6 degrees over a wide range of inlet-velocity ratio.

Description: As a means of preparing for high-altitude flight with spark-ignition engines in conjunction with exhaust-gas turbosuperchargers, various methods of modifying the exhaust-gas temperatures, which are initially higher than a turbine can withstand are mathematically compared. The thermodynamic results first obtained are then examined with respect to the effect on flight speed, climbing speed, ceiling, economy, and cruising range. The results are so presented in a generalized form that they may be applied to every appropriate type of aircraft design and a comparison with the supercharged engine without exhaust-gas turbine can be made.

Description: Tests of two-blade propellers having the NACA 4-(3)(06.3)-06 and NACA 4-(3)(06.4)-09 blade designs (blade activity factors of 179 and 263, respectively) have been made in the Langley 8-foot high-speed tunnel through a range of blade angle from 20 degrees to 70 degrees for free-stream Mach numbers from 0.165 to 0.725 to determine the effects of high solidity and compressibility on propeller characteristics. The tests are part of a general investigation of propellers at high forward speeds. Results previously reported for similar tests of two-blade propellers having the NACA 4-308-03 and NACA 4-308-045 blade designs (blade activity factors of 87 and 133, respectively) are included for comparison. The results showed that the 0.06- and 0.09-solidity blades, although producing efficiencies of the order of 90 percent, were less efficient than blades of conventional solidity. The variation in average blade lift coefficient with solidity at a constant blade angle and advance-diameter ratio through the speed range of these tests was found to be analogous to the variation of wing lift coefficient with aspect ratio, indicating that high-solidity blades may be desirable at very high speeds. Because of power limitations of the test equipment, conclusive evidence of the possible favorable effects of increased blade solidity at high speeds was not obtained. Further tests are desirable.

Description: In accordance with a request of the Bureau of Aeronautics, Navy Department, tests were performed in the Langley 20-foot free-spinning tunnel to determine the spin and recovery characteristics of a 1/24 scale model of the Grumman XTB3F-1 airplane. The airplane is a two-place, midwing torpedo bomber equipped with a tractor propeller and an auxiliary jet engine. The effect of control setting and movement on the erect and inverted spin and recovery characteristics of the model were determined for the normal loading. Brief tests with mass extended slightly along the fuselage were also made, however, in order to determine the effect of such a mass variation on elevator effectiveness. Tests were performed to determine the size of emergency spin-recovery tail and wing-tip parachutes required for satisfactory recovery by parachute action alone. The investigation also included emergency pilot-escape tests and tests to determine the rudder pedal and elevator stick forces necessary to move the rudder and elevator for recovery.

Description: The probable spin and recovery characteristics of the XSN2J-1 air-plane have been estimated on the basis of the results of brief test a performed on a model of an airplane of somewhat similar design. The spin-recovery tail-parachute requirements for the airplane were also determined end, in addition, an analysis was made to determine the best method of emergency pilot escape during a spin. The results of the investigation indicate that the recovery characteristics of the airplane will be satisfactory for all probable loading conditions of the airplane. A 6-foot-diameter tall parachute attached to a 30-foot tow-line will be satisfactory as a spin-recovery device for emergency recovery from demonstration spins. If the occupants of the airplane decide to abandon the airplane in a spin, they should leave the airplane from the outboard side of the cockpit and as far rearward as possible.

Description: Tests have been conducted in the Langley high-speed 7- by 10-foot tunnel over a Mach number range from 0.40 to 0.91 to determine the stability and control characteristics of an 0.08-scale model of the Chance Vought XF7U-1 airplane. The aileron characteristics of the complete model are presented in the present report with a very limited analysis of the results.

Description: As one phase of a comprehensive canopy load investigation, conventional front and rear sliding canopies which are typified by installation on the SB2C-4E airplane, were tested in the Langley full-scale tunnel to determine the pressure distributions and the aerodynamic loads on the canopies. A preliminary analysis of the results of these tests is presented in this report. Plots are presented that show the distribution of pressure at four longitudinal stations through each canopy for a range of conditions selected to determine the effects of varying canopy position, yaw, lift coefficient, and power. The results indicate that the maximum loads, based on the external-internal pressure differential, for the front and rear canopies were obtained with the airplane simulating the high speed flight condition. The highest loading on the front canopy was in the exploding direction for the configuration with the front and rear canopies closed. The highest loads on the rear canopy were in the crushing direction with the front canopy open and the rear canopy closed. For most of the simulated flight conditions, the highest loads on the front canopy, per unit area, were over twice as great as the highest loads on the rear canopy when the comparison was made for the most critical canopy configuration in each case. The external pressure distribution over the front and rear canopies, which were fairly symmetrical to 0 degree angle of yaw, were greatly distorted at other yaw attitudes, particularly for the propeller operating conditions. These distorted pressure distributions resulted in local exploding and crushing loads on both canopies which were often considerably higher than the average canopy loads.

Description: A total of 197 hours of time histories of airspeed and altitude have been obtained on Lockheed Constellation airplanes flying between New York, N.Y. and San Francisco, Calif. during May 1946 and June 1946. Data for 130 hours were previously analyzed to determine the probability of attaining excessive airspeeds and Mach numbers and the results have been published. After the publication, data for additional 67 hours became available. All the data were obtained at altitudes less than 20,000 feet and under approximately the same conditions. The combined data have been analyzed to obtain the results given in the tables and figures contained herein. Based on the combined data, the probability of exceeding a given airspeed or Mach number is of the same order of magnitude as that given. The conclusions are, therefore, unchanged.

Description: An investigation to determine the performance and operational characteristics of the TG-1OOA gas turbine-propeller engine was conducted in the Cleveland altitude wind tunnel. As part of this investigation, the combustion-chamber performance was determined at pressure altitudes from 5000 to 35,000 feet, compressor-inlet rm-pressure ratios of 1.00 and 1.09, and engine speeds from 8000 to 13,000 rpm. Combustion-chamber performance is presented as a function of corrected engine speed and.correcte& horsepower. For the range of corrected engine speeds investigated, over-all total-pressure-loss ratio, cycle efficiency, ana the frac%ional loss in cycle efficiency resulting from pressure losses in the combustion chambers were unaffected by a change in altitude or compressor-inlet ram-pressure ratio. The scatter of combustion- efficiency data tended to obscure any effect of altitude or ram-pressure ratio. For the range of corrected horse-powers investigated, the total-pressure-loss ratio an& the fractional loss in cycle efficiency resulting from pressure losses in the combustion chambers decreased with an increase in corrected horsepower at a constant corrected engine speed. The combustion efficiency remained constant for the range of corrected horse-powers investigated at all corrected engine speeds.

Description: Based upon a simplified representation of the mode of operation of the pulse-jet tube, the effect of the influences mentioned in the title were investigated and it will be shown that, for a jet tube with a fccmndesigned to be aerodynamically favorable, the ability to operate is at least questionable. By taking into account the course of the development of pressure by combustion, a new insight has been obtained into the processes of motion within the jet tube, an insight that explains a number of empirical observations, namely: certain particulars of the sequence of pressure variations; the existence of an optimum valve-opening ratio; the occurrence of an intrusion of air; and the existence of a flight speed above lrhichthe jet tube ceases to operate. At too great an opening ratio or at too great a flight s-peed, the continuous flow through the tube is too predominant over the oscilla~ory process to perinitthe occurrence of an explosion powerful enough to maintain continuous operation. Certain possible means of making the operation of the jet tube more independent of the flight speed and of reducing the flow losses were proposed and discussed.

Description: An analysis has been made of airfoil data taken on several NACA 16-series propeller airfoils from tests of 5-inch-chord models in the Langley 24 inch high-speed tunnel and l2-inch-chord models in the Langley 8 foot high-speed tunnel, This analysis has shown that the combined effects of Reynolds number changes and variations in airfoil characteristics resulting from differences in models and tunnels are such that when 5 inch-chord and l2-inch-chord data are applied to full-scale propeller design at or near the design condition, differences of less than 1 percent in efficiency will be involved.

Description: Free-spinning-tunnel tests have been made on a 1/27-scale model of the Douglas XF3D-1 airplane to confirm a preliminary evaluation made of the airplane spin and recovery characteristics and previously reported. Recovery characteristics were satisfactory for erect and inverted spins when the model was in the clean condition. When the slow-down brakes were open, recoveries were slow. The pedal force necessary to reverse the airplane rudder during a spin will be within the physical capabilities of the pilot. A 10-foot-diameter parachute attached to the tail of the airplane (laid-out-flat diameter, drag coefficient 0.7) or a 4.5-foot-diameter parachute attached to the outboard wing tip will be satisfactory for emergency spin recovery from demonstration spins. If it becomes necessary for the crew to abandon the airplane during a spin, they should leave from the outboard side of the cockpit. The test results indicated spin and recovery characteristics generally similar to those indicated in the preliminary evaluation.

Description: The teat results showed that with either of the three tail arrangements, the model usually spun in flat attitudes with oscillations about the lateral and longitudinal axes. In general, full reversal of the rudder pedals did not stop the spinning rotation. To make the model satisfactorily meet-the spin-recovery requirements it was found that installation of either a very large ventral fin (l7.9 square feet, full scale) below the tail or a somewhat smaller ventral fin and rudder (12.4 square feet, total . full-scale area) with a rudder throw of at least +/-22deg was required. Either a 21.3-foot tail parachute or a 6.4-foot wing-tip parachute (drag coefficient approximately 0.70) appears necessary as an emergency spin-recovery device during demonstration spins.

Description: A preliminary analytical investigation was made to determine the feasibility of the basic idea of controlled failure points as safety valves for the primary airplane structure. The present analysis considers the possibilities of the breakable wing tip which, in failing as a weak link, would relieve the bending moments on the wing structure. The analysis was carried out by computing the time histories of the wing and stabilizer angle of attack in a 10g pull-up for an XF8F airplane with tips fixed and comparing the results with those for the same maneuver, that is, elevator motion but with tips jettisoned at 8g. The calculations indicate that the increased stability accompanying the loss of the wing tips reduces the bending moment an additional amount above that which would be expected from the initial loss in lift and the inboard shift in load. The vortex shed when the tips are lost may induce a transient load requiring that the tail be made stronger than otherwise.

Description: Efficiency investigations have been made on a single-stage modification of the turbine of a Mark 25 aerial torpedo to determine the performance of the unit with five different turbine nozzles. The output of the turbine blades was computed by analyzing the windage and mechanical-friction losses of the unit. The turbine was faund to be most efficient with a cast nozzle having sharp-edged inlets to the nine nozzle ports. An analysis af the effectiveness af the first and second stages of the standard Mark 25 torpedo turbine indicates that the first- stage turbine contributes nearly all the brake power produced at blade-jet speed ratios above 0.26.

Description: Results obtained from gust and draft velocity measurements within thunderstorms for the period August 17, 1946 to August 19, 1946 at Orlando, Florida are presented herein. These data are summarized in tables I and II and are of the type presented in reference 1 for previous flights. Inspection of photo-observer records taken on the present flights indicated that mo ambient-air temperature data were obtained.

Description: Wind-tunnel tests were made of a 1/25 scale model of the Martin JRM-1 airplane to determine: (1) The longitudinal stability and control characteristics of the JRM-1 model near the water and lateral and directional stability characteristics with power while moving on the surface of the water, the latter being useful for the design of tip floats; (2) The stability and stalling characteristics of the wing with a modified airfoil contour; (3) Stability characteristics of a hull of larger design gross weight; The test results indicated that the elevator was powerful enough to trim the original model in a landing configuration at any lift coefficient within the specified range of centers of gravity. The ground-board tests for evaluating the aerodynamic forces and moments on an airplane in a simulated cross wind indicate a high dihedral effect in the presence of the ground board and, consequently, during low-speed taxying and take-off, large overturning moments would result which would have to be overcome by the tip floats.

Description: The results obtained from measurements of gust velocities, draft velocities, and ambient-air temperature within thunderstorms for the period September 17, 1946 to September 18, 1946 at Orlando, Fla. are presented herein. These data are summarized in tables I, II, and III, respectively, and are of the type presented in reference 1 for previous flights.

Description: The XF-12 airplane was designed by Republic Aviation Corporation to provide the Army Air Forces with a high performance, photo reconnaissance aircraft. A series of air-stream surveys were made n the vicinity of the empennage of a 1/8.33-scale powered model of the XF-12 airplane in the Langley 19-foot pressure tunnel. Surveys of the vortical-tail region were made through a range of yaw angles of plus or minus 20 degrees at a high and low angle of attack. The horizontal-tail surveys were made over a fairly wide range of angles of attack at zero degrees yaw. Several power and flap conditions were investigated. The results are presented in the form a dynamic pressure ratios, sidewash angles, and downwash angles plotted against vertical distance from the fuselage center line. The results of the investigation indicate that a vertical tail located in a conventional position would be in a field of flow where the dynamic pressure ratios at the horizontal tail to be increased; for equal lift coefficients, the effect of power or flap deflection on the direction of flow at any particular point in the region of the horizontal tail is small.

Description: The performance of the 11-stage axial-flow compressor in the X24C-4B turbojet engine was analyzed on the basis of results obtained from an investigation of the complete engine in the NACA Cleveland altitude wind tunnel. The engine was operated with four, exhaust nozzles of different outlet area over a range of engine speeds from 6000 to 12,500 rpm, corrected engine speeds from approximately 6100 to 13,600 rpm, and compressor Mach numbers from 0.45 to 1.00. Data are presented for engine operation over a range of simulated altitudes from 15,000 to 45,000 feet and simulated flight Mach numbers from 0.24 to 1.08.

Description: Three methods for calculating span load distribution, those developed by V.M Falkner, Wm. Mutterperl, and J. Weissinger, have been applied to five swept wings. The angles of sweep ranged from -45 degrees to +45 degrees. These methods were examined to establish their relative accuracy and case of application. Experimentally determined loadings were used as a basis for judging accuracy. For the convenience of the readers the computing forms and all information requisite to their application are included in appendixes. From the analysis it was found that the Weissinger method would be best suited to an over-all study of the effects of plan form on the span loading and associated characteristics of wings. The method gave good, but not best, accuracy and involved by far the least computing effort. The Falkner method gave the best accuracy but at a considerable expanse in computing effort and hence appeared to be most useful for a detailed study of a specific wing. The Mutterperl method offered no advantages in accuracy of facility over either of the other methods and hence is not recommended for use.

Description: Preliminary tests have been made of a small burner to meet the requirements for application to supersonic ram jets. The principal requirements were taken as: (1) efficient combustion in a high-velocity air stream, (2) utilization for combustion of only a small fraction of the air passing through the unit, (3) low resistance to air flow, (4) simple construction, and (5) light weight. Tests of a small burner were carried to stream velocities of nearly 150 feet per second and fuel rates such that one-eighth to one-fourth of the total air was involved in combustion. Commercial propane was selected as the fuel since its low boiling point facilitated vaporization. Combustion which was 80 percent complete along with low aerodynamic losses was obtained by injecting the fuel evenly, prior to ignition, and allowing it to mix with the air without appreciably disturbing the stream. The pressure drop due to frictional losses around the burner and to the adjacent inside walls of the ram jet is small compared with the pressure drop due to combustion.

Description: A 1/12-scale model of a horizontal tail of a fighter airplane was tested through the transonic speeds in the high-speed flow over an airplane wing, the surface of which served as a reflection plane for the model. Measurements of lift, elevator-hinge moment, angle of attack, and elevator angle were made in the Mach number range from 0.75 to 1.04 for elevator deflections ranging from 10 degrees to minus 10 degrees, and for angles of attack of minus 1.2 degrees, 0.4 degrees, and 3.4 degrees. The equipment used to measure the hinge moments of the model proved to be unsatisfactory, and for this reason the hinge-moment data are considered to be only qualitative.

Description: Tests to evaluate the effect of a conical windshield on the drag of a bluff body at supersonic speeds were performed for the following configurations: a sharp nose fuselage with stabilizing fins,a blunt nose fuselage with a hemispherical shape, and a blunt nose fuselage with a conical point. Results of the drag coeeficient are described at Mach 1.0 and the greatest Mach number of 1.37.

Description: Pressures and temperatures throughout the X24C-4B turbojet engine are presented in both tabular and graphical forms to show the effect of altitude, flight Mach number, and engine speed on the internal operation of the engine. These data were obtained in the NACA Cleveland altitude wind tunnel at simulated altitudes from 5000 to 45,000 feet, simulated flight Mach numbers from 0.25 to 1.08, and engine speeds from 4000 to 12,500 rpm. Location and detail drawings of the instrumentation installed at seven survey stations in the engine are shown. Application of generalization factors to pressures and temperatures at each measuring station for the range of altitudes investigated showed that the data did not generalize above an altitude of 25,000 feet. Total-pressure distribution at the compressor outlet varied only with change in engine speed. At altitudes above 35,000 feet and engine speeds above 11,000 rpm, the peak temperature at the turbine-outlet annulus moved inward toward the root of the blade, which is undesirable from blade-stress considerations. The temperature levels at the turbine outlet and the exhaust-nozzle outlet were lowered as the Mach number was increased. The static-pressure measurements obtained at each stator stage of the compressor showed a pressure drop through the inlet guide vanes and the first-stage rotor at high engine speeds. The average values measured by the manufacturer's instrumentation werein close agreement with the average values obtained with NACA instrumentation.

Description: A preliminary investigation of the over-all performance of a simply constructed, short-life, turbojet engine was conducted. The unit was operated at a pressure altitude of 15,000 feet for ram-pressure ratios of 1.2 t o 1.8. The corrected engine speed was varied from the minimum for good combustion to about 17,000 rpm, which is approximately 75 percent of rated speed. The performance is given by generalized parameters that permit the calculation of performance at any altitude. The corrected net thrust of the turbojet engine increased with ram-pressure ratio for a given corrected engine speed above 14,500 rpm and reached a maximum of 425 pounds at a ram-pressure ratio of 1.8 and a corrected engine speed of 16,650 rpm, The corrected thrust specific fuel consumption decreased with flight speed for corrected engine speeds higher than 13,600 rpm, The minimum corrected thrust specific fuel consumption of 1.48 was obtained at a ram-pressure ratio of 1,8 and a corrected engine speed of 15,000 rpm. For all ram-pressure ratios, choking occurred in the engine for corrected engine speeds greater than 14,500 rpm.

Description: An investigation has been conducted in the Cleveland altitude wind tunnel to determine the operational characteristics of the I-40 jet-propulsion engine over a range of pressure altitudes from 10,000 to 50,000 feet and ram-pressure ratios from 1.00 to 1.76. Engine operational data were obtained with the engine in the standard configuration and with various modifications of the fuel system, the electrical system, and the combustion chambers. The effects of altitude and airspeed on operating speed range, starting, windmilli.ng, acceleration, speed regulation, cooling, and vibration of the standard and modified engines were determined, and damage to parts was noted. Maximum engine speed was obtainable at all altitudes and airspeeds wi th each fuel-control system investigated. The minimum idling speed was raised by increases in altitude and airspeed. The lowest minimum stable speeds were obtained with the standard configuration using 40-gallon nozzles with individual metering plugs. The engine was started normally at altitudes as high as 20,000 feet with all of the fuel systems and ignition combinations except one. Ignition at 70,000 feet was difficult and, although successful ignition occurred, acceleration was slow and usually characterized by excessive tail-pipe temperature. During windmilling investigations of the engine equipped with the standard fuel system, the engine could not be started at ram-pressure ratios of 1.1 to 1.7 at altitudes of 10,000, 20,000 and 30,000 feet. When equipped with the production barometric and Monarch 40-gallon nozzles, the engine accelerated in 12 seconds from an engine speed of 6000 rpm to 11,000 rpm at 20,000 feet and an average tail-pipe temperature of 11000 F. At the same altitude and temperature, all the engine configurations had approximately the same rate of acceleration. The Woodward governor produced the safest accelerations, inasmuch as it could be adjusted to automatically prevent acceleration blow out. The engine speed was held constant by the Woodward governor and the Edwards regulator during simulated dives and climbs at constant throttle position. The bearing cooling system was satisfactory at all altitudes and airspeeds. The engines operated without serious failure, although the exhaust cone, the tail pipe, and the airplane fuselage were damaged during altitude starts.

Description: Performance characteristics of the turbine in the 19B-8 jet propulsion engine were determined from an investigation of the complete engine in the Cleveland altitude wind tunnel. The investigation covered a range of simulated altitudes from 5000 to 30,000 feet and flight Mach numbers from 0.05 to 0.46 for various tail-cone positions over the entire operable range of engine speeds. The characteristics of the turbine are presented as functions of the total-pressure ratio across the turbine and the turbine speed and the gas flow corrected to NACA standard atmospheric conditions at sea level. The effect of changes in altitude, flight Mach number, and tail-cone position on turbine performance is discussed. The turbine efficiency with the tail cone in varied from a maximum of 80.5 percent to minimum of 75 percent over a range of engine speeds from 7500 to 17,500 rpm at a flight Mach number of 0.055. Turbine efficiency was unaffected by changes in altitude up to 15,000 feet but was a function of tail-cone position and flight Mach number. Decreasing the tail-pipe-nozzle outlet area 21 percent reduced the turbine efficiency between 2 and 4.5 percent. The turbine efficiency increased between 1.5 and 3 percent as the flight Mach number changed from 0.055 to 0.297.