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 experimental investigation was made of a preloaded spring-tab flutter model to determine the effects on flutter speed of aspect ratio, tab frequency, and preloaded spring constant. The rudder was mass-balanced, and the flutter mode studied was essentially one of three degrees of freedom (fin bending coupled with rudder and tab oscillations). Inasmuch as the spring was preloaded, the tab-spring system was a nonlinear one. Frequency of the tab was the most significant parameter in this study, and an increase in flutter speed with increasing frequency is indicated. At a given frequency, the tab of high aspect ratio is shown to have a slightly lower flutter speed than the one of low aspect ratio. Because the frequency of the preloaded spring tab was found to vary radically with amplitude, the flutter speed decreased with increase in initial displacement of the tab.

Description: Low-speed tests of a pilotless aircraft were conducted in the Langley propeller-research tunnel to provide information for the estimation of the longitudinal stability and. control, to measure the aileron effectiveness, and to calibrate the radome and the Machmeter pitot-static orifices. It was found that the model possessed a stEb.le variation of elevator angle required for trim throughout the speed range at the design angle of attack. A comparison of the airplane with and without JATO units and with an alternate rocket booster showed that a large loss in longitudinal stability and control resulting from the addition of the rocket booster to the aircraft was sufficient to make the rocket-booster assembly unsatisfactory as an alternate for the JATO units. Reversal of the aileron effectiveness was evident at positive deflections of the vertical wing flap indicating that the roll-stabilization system would produce roiling moments in a tight right turn contrary to its design purpose. Vertical-wing-flap deflections caused large errors in the static-pressure reading obtained by the original static-tube installation. A practical installation point on the fuselage was located which should yield reliable measurement of the free-stream static pressure.

Description: A wind-tunnel investigation has been made to determine the effects of unsymmetrical horizontal-tail arrangements on the power-on static longitudinal stability of a single-engine single-rotation airplane model. Although the tests and analyses showed that extreme asymmetry in the horizontal tail indicated a reduction in power effects on longitudinal stability for single-engine single-rotation airplanes, the particular "practical" arrangement tested did not show marked improvement. Differences in average downwash between the normal tail arrangement and various other tail arrangements estimated from computed values of propeller-slipstream rotation agreed with values estimated from pitching-moment test data for the flaps-up condition (low thrust and torque) and disagreed for the flaps-down condition (high thrust and torque). This disagreement indicated the necessity for continued research to determine the characteristics of the slip-stream behind various propeller-fuselage-wing combinations. Out-of-trim lateral forces and moments of the unsymmetrical tail arrangements that were best from consideration of longitudinal stability were no greater than those of the normal tail arrangement.

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: Several groups of new airfoil sections, designated as the NACA 8-series, are derived analytically to have lift characteristics at supercritical Mach numbers which are favorable in the sense that the abrupt loss of lift, characteristic of the usual airfoil section at Mach numbers above the critical, is avoided. Aerodynamic characteristics determined, from two-dimensional windtunnel tests at Mach numbers up to approximately 0.9 are presented for each of the derived airfoils. Comparisons are made between the characteristics of these airfoils and the corresponding characteristics of representative NPiCA 6-series airfoils. The experimental results confirm the design expectations in demonstrating for the NACA S-series airfoils either no variation, or an Increase from the low-speed design value, In the lift coefficient at a constant angle of attack with increasing Mach number above the critical. It was not found possible to improve the variation with Mach number of the slope of the lift curve for these airfoils above that for the NACA 6-series airfoils. The drag characteristics of the new airfoils are somewhat inferior to those of the NACA 6- series with respect to divergence with Mach number, but the pitching-moment characteristics are more favorable for the thinner new sections In demonstrating somewhat smaller variations of moment coefficient with both angle of attack and Mach number. The effect on the aero&ynamic characteristics at high Mach numbers of removing the cusp from the trailing-edge regions of two 10-percent-chord-thick NACA 6-series airfoils is determined to be negligible.

Description: Tests of a partial-span model of a large bomber-type air1ane were conducted to determine the. aerodynamic characteristics of the wing equipped with full-span flaps and a retractable spoiler end aileron lateral control system. The arrangement consisted of (1) a double slotted flap extending over aproximate1y 86 percent of the wing semispan, (2) a 20-percent constant-percentage-chord aileron extending from the outboard end of the flap to the wing tip, and (3) a retractable spoiler, located at the 65-percent wing-chord station and extending from approximately 63 percent of the wing semispan to the wing tip. In addition, tests were made of a wing vent (of 1 and 2 percent of the wing chord located directly behind the spoiler), perforations in the spoiler, a blot or cut-out along the lower edge of the spoiler and spoilers of various spans. With full-span flaps deflected and with the 2-percent vent open or closed the initial stalling of the wing occurred at the tips, but with the vents closed there probably would be no appreciable loss in lateral control until maximum lift was reached. The l-percent vent increased the rolling effectiveness of the spoiler at small spoi1er deflections, particularly at high angles of attack with flaps deflected. With flaps deflected the 2-percent vent caused a large reduction in both the wing lift and rolling effectiveness of the spoiler at large angles of attack. However, at small angle of attack the 2-percent vent increased the rolling effectiveness of the spoiler at small spoiler deflections. The simultaneous operation of the spoiler and vent (in contrast to a vent fixed in the wing) would result in a large increase in the effectiveness of the spoiler and would avoid any loss in wing lift as in a fixed vent arrangement. The tests of the spoiler modifications revealed that (1) the spoiler perforations reduced the rolling-moment and yawing-moment coefficients but caused the spoiler hinge-moment coefficients to become more positive; (2) the spoiler slot had no notable effect on the rolling-moment and yawing-moment characteristics but produced a positive increase in the spoiler hinge-moment coefficients at large spoiler deflections; (3) the effects produced by the individual modifications were additive when the various modifications were combined. In general, progressively decreasing the spoiler span by removing the segments from the inboard end of the spoiler caused a decrease in rolling effectiveness approximately proportional to the span of the segment.

Description: Variable charge-air flow, cooling-air pressure drop, and fuel-air ration investigations were conducted to determine the cooling characteristics of a full-scale air-cooled single cylinder on a CUE setup. The data are compared with similar data that were available for the same model multicylinder engine tested in flight in a four-engine airplane. The cylinder-head cooling correlations were the same for both the single-cylinder and the flight engine. The cooling correlations for the barrels differed slightly in that the barrel of the single-cylinder engine runs cooler than the barrel of te flight engine for the same head temperatures and engine conditions.

Description: Tests were conducted in the Langley 24-inch highspeed tunnel to ascertain the static-pressure and total-pressure losses through screens ranging in mesh from 3 to 12 wires per inch and in wire diameter from 0.023 to 0.041 inch. Data were obtained from a Mach number of approximately 0.20 up to the maximum (choking) Mach number obtainable for each screen. The results of this investigation indicate that the pressure losses increase with increasing Mach number until the choking Mach number, which can be computed, is reached. Since choking imposes a restriction on the mass rate of flow and maximum losses are incurred at this condition, great care must be taken in selecting the screen mesh and wire dimmeter for an installation so that the choking Mach number is

Description: Small high-speed single-cylinder compression-ignition engines were tested to determine their performance characteristics under high supercharging. Calculations were made on the energy available in the exhaust gas of the compression-ignition engines. The maximum power at any given maximum cylinder pressure was obtained when the compression pressure was equal to the maximum cylinder pressure. Constant-pressure combustion was found possible at an engine speed of 2200 rpm. Exhaust pressures and temperatures were determined from an analysis of indicator cards. The analysis showed that, at rich mixtures with the exhaust back pressure equal to the inlet-air pressure, there is excess energy available for driving a turbine over that required for supercharging. The presence of this excess energy indicates that a highly supercharged compression-ignition engine might be desirable as a compressor and combustion chamber for a turbine.

Description: A continuous 50-hour test was conducted to determine the effect of maximum cruise-power operation at ultra-lean fuel-air mixture and increased spark advance on the mechanical conditions of cylinder components. The test was conducted on a nine-cylinder air-cooled radial engine at the following conditions:brake horsepower, 750; engine speed, 1900 rpm; brake mean effective pressure, 172 pounds per square inch; fuel-air ratio, 0.052; spark advance, 30 deg B.T.C.; and maximum rear-spark-plug-bushing temperature, 400 F. In addition to the data on corrosion and wear, data are presented and briefly discussed on the effect of engine operation at the conditions of this test on economy, knock, preignition, and mixture distribution. Cylinder, piston, and piston-ring wear was small and all cylinder component were in good condition at the conclusion of the 50-hour test except that all exhaust-valve guides were bellmouthed beyond the Army's specified limit and one exhaust-valve face was lightly burned. It is improbable that the light burning in one spot of the valve face would have progressed further because the burn was filled with a hard deposit so that the valve face formed an unbroken seal and the mating seat showed no evidence of burning. The bellmouthing of the exhaust-valve guides is believed to have been a result of the heavy carbon and lead-oxide deposits, which were present on the head end of the guided length of the exhaust-valve stem. Engine operational the conditions of this test was shown to result In a fuel saving of 16.8 percent on a cooled-power basis as compared with operation at the conditions recommended for this engine by the Army Air Forces for the same power.

Description: A method is developed of easily determining the performance of a compressor system relative to that of the power section for a given altitude. Because compressors, reciprocating engines, and turbines are essentially flow devices, the performance of each of these power-plant components is presented in terms of similar dimensionless ratios. The pressure and temperature changes resulting from restrictions of the charge-air flow and from heat transfer in the ducts connecting the components of the power plant are also expressed by the same dimensionless ratios and the losses are included in the performance of the compressor. The performance of a mechanically driven, single-stage compressor in relation to the performance of a conventional air-cooled engine operating at sea-level conditions is presented as an example of the application of the method.

Description: A comparison has been made in flight of the antiknock characteristics of 33-R fuel with that of 28-R and a triptane blent. The knock-limited performance of the three fuels - 33-R, a blend of 80 percent 28-R plus 20 percent triptane (leaded to 4.5 ml TEL/gal), and 28-R - was investigated in two modified 14-cylinder double-row radial air-cooled engines. Tests were conducted on the engines as installed in the left inboard nacelle of an airplane. A carburetor-air temperature of approximately 85 deg F was maintained. The conditions covered at an engine speed of 2250 rpm were high and low blower ratios and spark advances of 25 deg and 32 deg B.T.C. For an engine speed of 1800 rpm only the high-blower condition was investigated for both 25 deg and 32 deg spark advances. For the conditions investigated the difference between 33-R and the triptane blend was found to be slight; the performance of 33-R fuel, however, was slightly higher than that of the triptane blend in the lean region. The knock-limited power obtained with the 33-R fuel was from 14 to 28 percent higher than that of the 28-R fuel for the entire range of test conditions; the greatest improvement was shown in the lean region.

Description: No abstract available

Description: No abstract available

Description: Efficiency tests have been conducted on a single-stage impulse engine having an 11-inch pitch-line diameter wheel with inserted buckets and a fabricated nozzle diaphragm. The tests were made to determine the effect of inlet pressure, Inlet temperature, speed, and pressure ratio on the turbine efficiency. An analysis is presented that relates the effect of inlet pressure and temperature to the Reynolds number of the flow. The agreement between the analysis and the experimental data indicates that the changes in turbine efficiency with Inlet pressure and temperature may be principally a Reynolds number effect.

Description: The results of a theoretical analysis of the hinge-moment characteristics of various sealed-internal-balance arrangements for control surfaces are presented. The analysis considered overhands sealed to various types of wing structure by flexible seals spanning gaps of various widths or sealed to the wing structure by a flexible system of linked plates. Leakage was not considered; the seal was assumed to extend the full spanwise length of the control surface. The effect of the developed width of the flexible seal and of the geometry of the structure to which the seal was anchored was investigated, as well as the effect of the gap width that is sealed. The results of the investigation indicated that the most nearly linear control-surface hinge-moment characteristics can probably be obtained from a flexible seal over a narrow gap (about 0.1 of the overhang chord), which is so installed that the motion of the seal is restricted to a region behind the point of attachment of the seal to the wing structure. Control-surface hinge moments that tend to be high at large deflections and low or overbalanced at small deflections will result if a very narrow seal is used.

Description: The laws of conservation of mass, momentum, and energy are applied to the compressible flow through a two-dimensional cascade of airfoils. A fundamental relation between the ultimate upstream and downstream flow angles, the inlet Mach number, and the pressure ratio across the cascade is derived. Comparison with the corresponding relation for incompressible flow shows large differences. The fundamental relation reveals two ranges of flow angles and inlet Mach numbers, for which no ideal pressure ratio exists. One of these nonideal operating ranges is analogous to a similar type in incompressible flow. The other is characteristic only of compressible flow. The effect of variable axial-flow area is treated. Some implications of the basic conservation laws in the case of nonideal flow through cascades are discussed.

Description: Efficiency tests have been conducted on a single-stage impulse turbine having a 13.2-inch pitch-line diameter wheel and a cast nozzle diaphram over a range of turbine speeds from 3000 to 17,000 rpm, pressure ratios from 1.5 to 5.0, inlet total temperatures from 1200 deg to 2000 deg R, and inlet total pressures from 18 to 59 inches of mercury absolute. The effect of inlet temperature and pressure on turbine efficiency for constant pressure ration and blade-to-jet speed ration is correlated against a factor derived from the equation for Reynolds number. The degree of correlation indicates that the change in turbine efficiency with inlet temperature and [ressure for constant pressure ration and blade-to-jet speed ration is principally a Reynolds number effect.

Description: No abstract available

Description: Tests were made of a model representative of a single-engine tractor-type airplane for the purpose of determining the stability and control effects of a propeller used as an aerodynamic brake. The tests were made with single-and dual-rotation propellers to show the effect of type of propeller rotation, and with positive thrust to provide basic data with which to compare the effects of negative thrust. Four configurations of the model were used to give the effects of tilting the propeller thrust axis down 5 deg., raising the horizontal tail, and combining both tilt and raised tail. Results of the tests are reported herein. The effects of negative thrust were found to be significant. The longitudinal stability was increased because of the loss of wing lift and increase of the angle of attack of the tail. Directional stability and both longitudinal and directional control were decreased because of the reduced velocity at the tail. These effects are moderate for moderate braking but become pronounced with full-power braking, particularly at high values of lift coefficient. The effects of model configuration changes were small when compared with the over-all effects of negative-thrust operation; however, improved stability and control characteristics were exhibited by the model with the tilted thrust axis. Raising the horizontal tail improved the longitudinal characteristics, but was detrimental to directional characteristics. The use of dual-rotation propeller reduced the directional trim charges resulting from the braking operation. A prototype airplane was assumed and handling qualities were computed and analyzed for normal (positive thrust) and braking operation with full and partial power. The results of these analyses are presented for the longitudinal characteristics in steady and accelerated flight, and for the directional characteristics in high- and low-speed flight. It was found that by limiting the power output of the engine (assuming the constant-speed propeller will function in the range of blade angles required for negative thrust) the stability and control characteristics may be held within the limits required for safe operation. Braking with full power, particularly at low speeds, is dangerous, but braking with very small power output is satisfactory from the standpoint of control. The amount of braking produced with zero power output is equal to or better than that produced by conventional spoiler-type brakes.

Description: A theory has been developed for resetting the blade angles of an axial-flow compressor in order to improve the performance at speeds and flows other than the design and thus extend the useful operating range of the compressor. The theory is readily applicable to the resetting of both rotor and stator blades or to the resetting of only the stator blades and is based on adjustment of the blade angles to obtain lift coefficients at which the blades will operate efficiently. Calculations were made for resetting the stator blades of the NACA eight-stage axial-flow compressor for 75 percent of design speed and a series of load coefficients ranging from 0.28 to 0.70 with rotor blades left at the design setting. The NACA compressor was investigated with three different blade settings: (1) the design blade setting, (2) the stator blades reset for 75 percent of design speed and a load coefficient of 0.48, and (3) the stator blades reset for 75 percent of design speed and a load coefficient of 0.65.

Description: A 1/5-scale model of the Republic x-84 airplane (Army Project MX-578) was tested in the Langley 300 MPH 7- by 10-foot tunnel. The primary object of the tests was twofold: to determine, a practicable method of increasing the longitudinal 3tability in the landing configuration, and to investigate the effects on longitudinal and lateral Stability of various external stored (fuel tanks, bombs, and rockets). The effects of the fuselage dive brakes were also determined, and the critical Mach numbers of certain of the airplane components were estimated. The use of the revised horizontal tail (of larger aspect ratio and area than the original) seemed to be the most feasible expedient for materially increasing the longitudinal stability in the landing configuration. The neutral-point shifts produced by the various external stores were unstable, the largest shift being about 2.5 percent mean aerodynamic chord. No appreciable aerodynamic trim changes were caused by the external stores. From the standpoint of range, maximum s peed, and rate of climb, the advantages of mounting the fuel tanks at the wing tips rather than inboard beneath the wings were clearly demonstrated by the tests. The effective dihedral parameter was the only static lateral-stability derivative appreciably affected by the external stores. At high lift coefficients, the tip-mounted tanks caused a large increase in the effective dihedral parameter (about 40 increase at a lift coefficient of 1.0). This increase was held undesirable, because the tendency toward oscillatory instability that it would cause would be heightened by the increased moments of inertia resulting from the weight of the tanks when carrying fuel. The fuselage dive brakes, when deflected, caused a change in trim tending to nose the airplane up; the neutral point also moved rearward upon deflecting the dive brakes. The amount of elevator required to overcome the change in trim was well within the available range of deflection. It was estimated that a drive-brake deflection of 900 would.decrease the terminal Mach number in a vertical dive by about 0.1. The estimated critical Mach number of the V-front canopy was about 0.04 greater than that of the original canopy. Pressure-distribution tests disclosed severe pressure peaks inside the nose of the jet entrance duct. These peaks, which would lead to separation and consequently poor pressure recovery at, the engine, could be reduced by, using a smaller nose,radius and: a modified internal lip shape

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: Low Mach number longitudinal-stability and control characteristics as predicted by use of wind tunnel data from a powered 3/16-scale model are compared with flight test measurements of a Navy BTD-1 airplane. The accuracy of the wind tunnel data and the discrepancies involved in attempting to correlate with flight data are discussed and analyzed. The comparison showed that wind tunnel predictions were, in general, in good agreement with flight test data. The predicted values were for the most part sufficiently accurate to show the satisfactory and unsatisfactory characteristics in the preliminary design stage and to indicate possible methods of improvement. The discrepancies which did occur were attributed principally to physical dissimilarities between model and airplane and the instability to determine accurately the flight power conditions. The effect of Mach number was considered negligible since the maximum flight test value was about 0.5. In order to simulate more closely the flight conditions and hence obtain more accurate data for predictions, it appears desirable to perform large-scale tests of unorthodox control surfaces such as the sealed vaned elevators with which the airplane was equipped.

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: A wind tunnel investigation was conducted to determine the performance of a 4000-pound-thrust axial-flow turbojet engine with a high flow compressor. Pressure altitudes included 5000 to 40000 feet with ram pressure ratios from 1.00 to 1.82. Altitudes included 20000 to 40000 feet and ram pressure ratios from 1.09 to 1.75. A comparison is made between engine performance with high flow and low flow compressors.

Description: A wind tunnel investigation was conducted to determine the performance of a turbine operating as an integral part of a turbojet engine. Data was obtained while the engine was running over full operable range of speeds at various altitudes and flight mach numbers, and with four nozzles of different outlet areas.A maximum turbine efficiency of 0.875 was obtained at altitude of 15 thousand feet, Mach number 0.53, and corrected turbine speed of 5900 rpm.

Description: The preignition characteristics of the R-2800 cylinder, as effected by fuel consumption, engine operating variables, and spark plug type and condition, were evaluated. The effects on preignition-limited performance of various percentages of aromatics (benzene, toluene, cumene, xylene) in a base fuel of triptane were investigated. Two paraffins (triptane and S + 6.0 ml TEL/gal) and two refinery blends (28-R and 33-R) were preignition rated. The effect of changes in the following engine operating variables on preignition limit was determined: inlet-air temperature, rear spark plug gasket temperature, engine speed, spark advance, tappet clearance, and oil consumption. Preignition limits of the R-2800 cylinder using Champion C34S and C35S and AC-LS86, LS87, and LS88 spark plugs were established and the effect of spark plug deterioration was investigated. No definite trends in preignition-limited indicated mean effective pressure were indicated for aromatics as a class when increased percentages of different aromatics were added to a base fuel of triptane. Three types of fuel (aromatics, paraffins, and refinery blends) showed a preignition range for this cylinder from 65 to 104 percent when based on the performance of S plus 6.0 ml TEL per gallon as 100 percent. The R-2800 cylinder is therefore relatively insensitive to fuel composition when compared to a CFR F-4 engine, which had a pre-ignition range from 72 to 100 percent for the same fuels. Six engine operating variables were investigated with the following results: preignition-limited indicated mean effective pressure decreased, with increases in engine speed, rear spark plug gasket temperature, inlet-air temperature, and spark advance beyond 20 F B.T.C. and was unaffected by rate of oil consumption or by tappet clearance. Spark plugs were rated over a range of preignition-limited indicated mean effective pressure from 200 to 390 pounds per square inch at a fuel-air ratio of 0.07 in the following order of increased resistance to preignition: AC-LS97, AC-LS88, Champion C358, AC-LS86, and Champion C34S. Spark plug deterioration in the form of cracks in the porcelain had been broken away from the center electrode and were retained in the spark plug cavity, the preignition limit was decreased as much as 57 percent. When the broken pieces had been removed, the preignition limit increased from that of the undamaged porcelain as the weight of removed porcelain was increased.

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: Temperature and pressure distributions for an original and modified 3000 pound thrust axial flow turbojet engine were investigated. Data are included for a range of simulated altitudes from 5000 to 45000 feet, Mach numbers from 0.24 to 1.08, and corrected engine speeds from 10,550 to 13,359 rpm.

Description: An investigation of a 1/7-scale powered model of the Kaiser Fleetwing all-wing airplane was made in the Langley full-scale tunnel to provide data for an estimation of the flying qualities of the airplane. The analysis of the stability and control characteristics of the airplane has been made as closely as possible in accordance with the requirements of the Bureau of Aeronautics, Navy Department's specifications, and a summary of the more significant conclusions is presented as follows. With the normal center of gravity located at 20 percent of the mean aerodynamic chord, the airplane will have adequate static longitudinal stability, elevator fixed, for all flight conditions except for low-power operation at low speeds where the stability will be about neutral. There will not be sufficient down-elevator deflection available for trim above speeds of about 130 miles per hour. It is probable that the reduction in the up-elevator deflections required for trim will be accompanied by reduced elevator hinge moments for low-power operation at low flight speeds. The static directional stability for this airplane will be low for all rudder-fixed or rudder-free flight conditions. The maximum rudder deflection of 30 deg will trim only about 15 deg yaw for most flight conditions and only 10 deg yaw for the condition with low power at low speeds. Also, at low powers and low speeds, it is estimated that the rudders will not trim the total adverse yaw resulting from an abrupt aileron roll using maximum aileron deflection. The airplane will meet the requirements for stability and control for asymmetric power operation with one outboard engine inoperative. The airplane would have no tendency for directional divergence but would probably be spirally unstable, with rudders fixed. The static lateral stability of the airplane will probably be about neutral for the high-speed flight conditions and will be only slightly increased for the low-power operation in low-speed flight. The airplane will not roll against the ailerons in a side-slip maneuver. Although the airplane would probably meet the minimum requirements of pb/2V of 0.07 at all speeds, there will be a loss in rolling ability of the airplane at high aileron deflections and at low flight speeds. It is estimated that the wing stall will be a gradual movement forward from the trailing edge and will be accompanied by no sudden pitching or rolling accelerations. Some stall warning may be indicated by reduction in the elevator and aileron force gradients and by the shaking of the controls caused by unsteady flow over the surfaces near the stall.

Description: The hydrodynamic characteristics of a 1/10-size powered dynamic model of the XP5Y-1 flying boat were determined in Langley tank no. 1. Stable take-offs were possible at all practicable positions of the center of gravity and flap deflections. An increase in gross load from 123.5 to 150.0 pounds (21.5 percent) had only a slight effect on the stable range for take-off. A decrease in forward acceleration from 3.0 to 1.0 feet per second per second had only a very small effect on the stable range for take-off. In general, the landings were free from skipping except at trims below 6 deg where one skip was encountered at an aft position of the center of gravity. The model porpoised during the landing runout at all positions of the center of gravity when landed at trims above 10 deg. Spray in the propellers was light at the design gross load, and was not considered excessive,at a gross load of 136.0 pounds.

Description: A model of the Consolidated Vultee Aircraft Corporation Skate 7 seaplane was tested in Langley tank no. 2. Presented without discussion in this paper are landing stability in smooth water, maximum normal accelerations occurring during rough-water landings, and take-off behavior in waves.

Description: An investigation has been conducted in the Langley 20-foot free-spinning tunnel to determine the effects of decreasing the rudder deflection, of decreasing the rudder span, and of differential rudder movements on the spin and recovery characteristics of a 0.057-scale model of the Chance Vought XF7U-1 airplane. The results indicated that decreasing the rudder span or the rudder deflections, individually or jointly, did not seriously alter the spin or recovery characteristics of the model; and recovery by normal use of controls (full rapid rudder reversal followed l/2 to 1 turn later by movement of the stick forward of neutral) remained satisfactory. Linking the original rudders so that the inboard rudder moves from full with the spin to neutral while the outboard rudder moves from neutral to full against the spin will also result in satisfactory spin and recovery characteristics. Calculations of rudder-pedal forces for recovery showed that the expected forces would probably be within the capabilities of a pilot but that it would be advisable to install some type of boost in the control system to insure easy and rapid movement of the rudders.

Description: The present report of Mr. Dupleich is the summary of a very extensive experimental study of the well-known mechanical phenomenon: the rotation in free fall (* air, for instance) of more or less elongated rectangles cut out of paper or pasteboard. This phenomenon, the conditions for existence of which depend chiefly on the elongated of the small plate and its weight per unit area, is essentially an aerodynamic phenomenon and as such, raises questions of a certain interest to our department.We believe that the modern concepts of the mechanics of fluids do not have the range attributed to them.

Description: A series of flight tests have been made at the Langley Flight Research Division at the request of the Bureau of Aeronautics, Department of the Navy, to determine the flying qualities of the Grumman F8F-1 air- plane. This paper presents the test results necessary to determine the longitudinal stability and control characteristics end the stalling characteristics. These tests were made between February and June of 1947- The range of Mach numbers covered in this investigation was approximately 0.10 to 0.62, and no attempt was made to investigate compressibility effects at higher Mach numbers. The lateral and directional stability and control characteristics of the subject airplane have already been reported (reference 1). Also presented in this paper is a discussion of the normal accelerations induced by yawing velocity and sideslip which were considered ob,jectionable by the pilot for this airplane. A discussion of the undesirable accelerations has been included with a view towards formulating some flying-qualities requirements limiting them.

Description: A spin investigation has been conducted in the Langley 20-foot free-spinning tunnel on a 1/24-scale model of the North American XP-86 airplane. The effects of control settings and movements upon the erect and inverted spin and recovery characteristics of the model were determined for the design gross weight loading. The long-range loading was also investigated and the effects of extending slats and dive flaps were determined. In addition, the investigation included the determination of the size of spin-recovery parachute required for emergency recovery from demonstration spins, the rudder force required to move the rudder for recovery, and the best method for the pilot to escape if it should become necessary to do so during a spin. The results of the investigation indicated that the XP-86 airplane will probably recover satisfactorily from erect and inverted spins for all possible loadings. It was found that fully extending both slats would be beneficial but that extending the dive brakes would cause unsatisfactory recoveries. It was determined that a 10.0-foot-diameter tail parachute with a drag coefficient of 0.7 and with a towline 30.0 feet long attached below the jet exit or a 6.0-foot-diameter wingtip parachute opened on the outer wing tip with a towline 6.0 feet long would insure recoveries from any spins obtainable. The rudder-pedal force necessary to move the rudder for satisfactory recovery was found to be within the physical capabilities of the pilot.

Description: This paper presents the results of measurements of longitudinal stability of a 1/50-scale model of the XP-88 airplane by the wing-flow method. Lift, rolling-moment, hinge-moment, and pitching-moment characteristics as well as the downwash at the tail were measured over a Mach number range from approximately 0.5 to 1.05 at Reynolds numbers below 1,000,000. No measurements of drag were obtained. No abrupt changes due to Mach number were noted in any of the parameters measured. The data indicated that the wing was subject to early tip stalling; that the tail effectiveness decreased gradually with increasing Mach number up to M = 0.9, but increased again at higher Mach numbers; that the variation of downwash with angle of attack did not change appreciably with Mach number except between 0.95 and 1.0 where d(epsilon)/d(alpha), decreased from 0.46 to 0.32; that at zero lift with a stabilizer setting of -1.5 deg there was a gradually increasing nosing-up tendency with increasing Mach number; and that the control-fixed stability in maneuvers at constant speed gradually increased with increasing Mach number.

Description: An investigation of the spin and recovery characteristics of a 0.057-scale model of the Chance Vought XF7U-1 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 were determined, as were also the effects of extending the wing slats, of center-of-gravity movement, and-of variation in the mass distribution. The investigation also included wing-tip spin-recovery-parachute tests, pilot-escape tests, and rudder-control-force tests. The investigation indicated that the spin and recovery characteristics of the airplane will be satisfactory for all conditions. It was found that a single 4.24-foot (full-scale) parachute when opened alone from the outboard wing tip or two 8.77-foot (full-scale) parachutes when opened simultaneously, one from each wing tip, would effect satisfactory emergency recoveries (the drag coefficients of the parachutes, based on the surface area of the parachute, were 0.83 and 0.70 for the 4.24- and 8.77-foot parachutes, respectively). The towline length in both cases was 25 feet (full scale). Tests results showed that, if the pilot should have to leave the airplane during a spin, he should jump from the outboard side (left side in a right spin) of the cockpit. The rudder-control force necessary for recovery from a spin was found to be rather high but appeared to be within the upper limits of a pilot's capabilities.

Description: Tests of a 1/20-scale dynamically similar model of the Northrop B-35 airplane were made to study its ditching characteristics. The model was ditched in calm water at the Langley tank no. 2 monorail. Various landing attitudes, speeds,and conditions of damage were simulated during the investigation. The ditching characteristics were determined by visual observation and from motion-picture records and time-history acceleration records. Both longitudinal and lateral accelerations were measured. Results are given in tabular form and time-history acceleration curves and sequence photographs are presented. Conclusions based on the model investigation are as follows: 1. The best ditching of the B-35 airplane probably can be made by contacting the water in a near normal landing attitude of about 9 deg with the landing flaps full down so as to have a low horizontal speed. 2. The airplane usually will turn or yaw but the motion will not be violent. The maximum lateral acceleration will be about 2g. 3. If the airplane does not turn or yaw immediately after landing, it probably will trim up and then make a smooth run or porpoise slightly. The maximum longitudinal decelerations that will be encountered are about 6g or 7g. 4. Although the decelerations are not indicated to be especially large, the construction of the airplane is such that extensive damage is to be expected, and it probably will be difficult to find ditching stations where crew members can adequately brace themselves and be reasonably sure of avoiding a large inrush of water.

Description: An investigation of the low-speed; power-off stability and control characteristics of a 1/20-scale model of the Consolidated Vultee XB-53 airplane equipped with full-span leading-edge slats has been conducted in the Langley free-flight tunnel. In this investigation it was found that the-full-span leading-edge slat gave about the same maximum lift coefficient as was obtained with the outboard single slotted flap and inboard slat. The stability and control characteristics were greatly improved except near the stall where the characteristics with the full-span slat were considered unsatisfactory because of a loss of directional stability and a slight nosing-up tendency.

Description: This report presents the results of the tests of a power-plant installation to improve the circumferential pressure-recovery distribution at the face of the engine. An underslung "C" cowling was tested with two propellers with full cuffs and with a modification to one set of cuffs. Little improvement was obtained because the base sections of the cuffs were stalled. A set of guide vanes boosted the over-all pressures and helped the pressure recoveries for a few of the cylinders. Making the underslung cowling into a symmetrical "C" cowling evened the pressure distribution; however, no increases in front pressures were obtained. The pressures at the top cylinders remained low and the high pressures at the bottom cylinders were reduced. At higher powers and engine speeds, the symmetrical cowling appeared best from the standpoint of over-all cooling characteristics.

Description: An investigation was conducted in the Cleveland altitude wind tunnel to determine the operational characteristics of an axial flow-type turbojet engine with a 4000-pound-thrust rating over a range of pressure altitudes from 5,000 to 50,OOO feet, ram pressure ratios from 1.00 to 1.86, and temperatures from 60 deg to -50 deg F. The low-flow (standard) compressor with which the engine was originally equipped was replaced by a high-flow compressor for part of the investigation. The effects of altitude and airspeed on such operating characteristics as operating range, stability of combustion, acceleration, starting, operation of fuel-control systems, and bearing cooling were investigated. With the low-flow compressor, the engine could be operated at full speed without serious burner unbalance at altitudes up to 50,000 feet. Increasing the altitude and airspeed greatly reduced the operable speed range of the engine by raising the minimum operating speed of the engine. In several runs with the high-flow compressor the maximum engine speed was limited to less than 7600 rpm by combustion blow-out, high tail-pipe temperatures, and compressor stall. Acceleration of the engine was relatively slow and the time required for acceleration increased with altitude. At maximum engine speed a sudden reduction in jet-nozzle area resulted in an immediate increase in thrust. The engine started normally and easily below 20,000 feet with each configuration. The use of a high-voltage ignition system made possible starts at a pressure altitude of 40,000 feet; but on these starts the tail-pipe temperatures were very high, a great deal of fuel burned in and behind the tail-pipe, and acceleration was very slow. Operation of the engine was similar with both fuel regulators except that the modified fuel regulator restricted the fuel flow in such a manner that the acceleration above 6000 rpm was very slow. The bearings did not cool properly at high altitudes and high engine speeds with a low-flow compressor, and bearing cooling was even poorer with a high-flow compressor.

Description: The problem of turbulence in aerodynamics is at present being attacked both theoretically and experimentally. In view of the fact however that purely theoretical considerations have not thus far led to satisfactory results the experimental treatment of the problem is of great importance. Among the different measuring procedures the hot wire methods are so far recognized as the most suitable for investigating the turbulence structure. The several disadvantages of these methods however, in particular those arising from the temperature lag of the wire can greatly impair the measurements and may easily render questionable the entire value of the experiment. The name turbulence is applied to that flow condition in which at any point of the stream the magnitude and direction of the velocity fluctuate arbitrarily about a well definable mean value. This fluctuation imparts a certain whirling characteristic to the flow.

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: The effect of rotor-blade length, inlet angle, and shrouding was investigated with four different nozzles in a single-stage modification of the Mark 25 aerial-torpedo power plant. The results obtained with the five special rotor configurations are compared with those of the standard first-stage rotor with each nozzle. Each nozzle-rotor combination was operated at nominal pressure ratios of 8, 15 (design), and 20 over a range of speeds from 6000 rpm to the design speed of 18,000 rpm. Inlet temperature and pressure conditions of 1OOOo F and 95 pounds per square inch gage, respectively, were maintained constant for all runs.

Description: A flight test was conducted at the Flight Test Station of the Pilotless Aircraft Research Division at Wallops Island, Va., to determine the longitudinal control and stability characteristics of a 0.5-scale model of the Fairchild Lerk Pilotless aircraft with the horizontal wing flaps deflected 15 degrees. The data were obtained by the use of a telemeter and also by radar tracking. The results show an increase of effectiveness of the longitudinal control in producing normal accelerations up to a Mach number of 0.75 where this effectiveness gradually decreased becoming negative at a Mach number of 0.89. Previous tests with wing flaps undeflected an increase in effectiveness up to Mach number of 0.93 where a sudden loss of control occurred. The model was dynamically stable throughout the speed range. The data confirmed the drag increase at the critical Mach number for large angles of attack is indicated in high-speed wind-tunnel tests.

Description: The NACA is conducting a general investigation of servo-mechanisms for use in powering aircraft control surfaces. This paper presents a theoretical analysis and the results of bench tests of a control-booster system which employs a variable displacement hydraulic pump. The booster is intended for use in a flight investigation to determine the effects of various booster parameters on the handling qualities of airplanes. Such a flight investigation would aid in formulating specific requirements concerning the design of control boosters in general. Results of the theoretical analysis and the bench tests indicate that the subject booster is representative of types which show promise of satisfactory performance. The bench tests showed that the following desirable features were inherent in this booster system: (1) No lost motion or play in any part of the system; (2) no detectable lag between motion of the contra1 stick and control surface; and (3) Good agreement between control displacements and stick-force variations with no hysteresis in the stick-force characteristics. The final design configuration of this booster system showed no tendency to oscillate, overshoot, or have other undesirable transient characteristics common to boosters.

Description: Flow-metering devices used by the NACA and by the manufacturer of the J33 turbojet engine were calibrated together to determine whether an observed discrepancy in weight flow of approximately 4 percent for the two separate investigations might be due to the different devices used to meter air flow. A commercial adjustable orifice and a square-edge flat-plate orifice used by the NACA and a flow nozzle used by the manufacturer were calibrated against surveys across the throat of the nozzle. It was determined that over a range of weight flows from 18 to 45 pounds per second the average weight flows measured by the metering device used for the compressor test would be 0.70 percent lower than those measured by the metering device used in the engine tests and the probable variation about this mean would be +/- 0.39 percent. The very close agreement of the metering devices shows that the greater part of the discrepancy in weight flow is attributable to the effect of inlet pressure.

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 the low-speed, power-off stability and control characteristics of a 1/20-scale model of the Consolidated Vultee XB-53 airplane has been conducted in the Langley free-flight tunnel. In the investigation it was found that with flaps neutral satisfactory flight behavior at low speeds was obtainable with an increase in height of the vertical tail and with the inboard slats opened. In the flap-down slat-open condition the longitudinal stability was satisfactory, but it was impossible to obtain satisfactory lateral-flight characteristics even with the increase in height of the vertical tail because of the negative effective dihedral, low directional stability, and large-adverse yawing moments of the ailerons.

Description: An investigation was conducted in the NACA Cleveland altitude wind tunnel to evaluate the performance characteristics of the X24C-4B turbojet engine over a range of simulated altitudes from 5000 to 45,000 feet,simulated flight Mach numbers from 0 to 1.08, and engine speeds from 4000 to 12,500 rpm. Performance data are presented to show graphically the effects of altitude at a flight Mach number of 0.25 and of flight Mach number at an altitude of 25,000 feet. The performance data are generalized to show the applicability of methods used to determine performance at any altitude from data obtained at a given altitude. A complete tabulation of performance data, as well as lubrication- and fuel- system data, is presented.

Description: Investigations were made of the turbine from a Mark 25 torpedo to determine the performance of the unit with three different turbine nozzles at various axial nozzle-wheel clearances. Turbine efficiency with a reamed nondivergent nozzle that uses the axial clearance space for gas expansion was little affected by increasing the axial running clearance from 0.030 to 0.150 inch. Turbine efficiency with cast nozzles that expanded the gas inside the nozzle passage was found to be sensitive to increased axial nozzle-wheel clearance. A cast nozzle giving a turbine brake efficiency of 0.525 at an axial running clearance of 0.035 inch gave a brake efficiency of 0.475 when the clearance was increased to 0.095 inch for the same inlet-gas conditions and blade-jet speed ratio. If the basis for computing the isentropic power available to the turbine is the temperature inside the nozzle rather then the temperature in the inlet-gas pipe, an increase in turbine efficiency of about 0.01 is indicated.

Description: A 1/7 scale semispan model of the XB-35 airplane was tested in the Langley 10 foot pressure tunnel, primarily for the purpose of investigating the effectiveness of a leading-edge slot for alleviation of stick-fixed longitudinal instability at high angles of attack caused by early tip stalling and a device for relief of stick-free instability caused by elevon up-floating tendencies at high angles of attack. Results indicated that the slot was not adequate to provide the desired improvement in stick-fixed stability. The tab-flipper device provided improvement in stick-free stability abd two of the linkage combinations tested gave satisfactory variations of control force with airspeed for all conditions except that in which the wing-tip "pitch-control" flap was fully deflected. However, the improvement in control force characteristics was accompanied by a detrimental effect on stick-fixed stability because of the pitching moments produced by the elevon tab deflection.

Description: The results obtained from gust and draft velocity measurements within thunderstorms for the period July 24, 1946 to August 6, 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. In two thunderstorm traverses, indications of ambient-air temperature were obtained from photo-observer records. These data are summarized in table III.

Description: An analysis of rim cooling, which cools the blade by condition alone, was conducted. Gas temperatures ranged from 1300 degrees to 1900 degrees F and rim temperatures from 0 degrees to 1000 degrees F below gas temperatures. Results show that gas temperature increases up to 200 degrees F are permissible provided that the blades are cooled by 400 degrees to 500 degrees F below the gas temperature. Relatively small amounts of blade cooling, at constant gas temperature, give large increases in blade life. Dependence of rim cooling on heat-transfer coefficient, blade dimensions, and thermal conductivity is determined by a single parameter.

Description: The spin and recovery characteristics of the Northrop XF-89 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 mainly estimated rather than determined by model tests because the XF-89 dimensional and mass characteristics were such as to make this airplane similar to several others, models of which have previously been tested. Brief tests were made on an available model of similar design to augment the estimation. The results indicate that the recovery characteristics will be satisfactory for all airplane loadings if recovery is attempted by use of rudder followed by moving the elevator down. The rudder pedal forces will be within the capabilities of the pilot but the elevator stick forces will be beyond the pilot's capabilities unless a trim tab, or a booster is used. A 9.5-foot-diameter flat-type tail parachute or a 5.0-foot-diameter flat-type wing-tip parachute with a drag coefficient of 0.7 will be a satisfactory emergency spin-recovery device for spin demonstrations and if it is necessary for the crew to abandon the spinning airplane, they should leave from the outboard side of the cockpit.

Description: An investigation was conducted to determine the effects of three design modifications of the original NACA injection impeller on the performance of an R-3350 engine. Different methods of injecting the fuel into the impeller air stream were studied and evaluated from the individual cylinder fuel-air ratios and the resulting cylinder temperatures. Each impeller was tested for a range of engine powers normally used in flight operation. The relatively simple design of the original injection impeller produced approximately the same mixture- and temperature-distribution characteristics as the modified impellers of more complex design. None of the modifications appreciably affected the manifold pressure, the combustion-air flow, nor the throttle angle required to maintain a given engine power,

Description: An altitude-test-chamber investigation was conducted to determine the operational and performance characteristics of a McDonnell afterburner with a fixed-area exhaust nozzle on a J34 engine. At rated engine speed, the altitude limit, as determined by combustion blow-out, occurred as a band of unstable operation of about 6000-foot altitude in width with minimum altitude limits from 31,000 feet at a simulated flight Mach number of 0.40 to about 45,500 feet at a simulated flight Mach number of 1.00. Considerable difficulty was experienced in attempting to establish or maintain balanced-cycle engine operation at altitudes above 36,000 feet. The fuel-air ratio for balanced-cycle operation and lean blowout of the afterburner, the augmented-thrust ratio, the total specific fuel consumption, and the afterburner combustion efficiency for balanced-cycle operation are summarized in a table. Satisfactory afterburner ignition was obtained over a range of flight Mach Numbers from 0.32 to 0.60 at altitudes from 10,000 to 30,000 and engine speeds from 10,000 to 12,500 rpm.

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: At the request of the Air Material Command, Arm Air Forces, an investigation was conducted at the NACA Cleveland laboratory to determine the performance characteristics of the XJ-41-V turbojet-engine compressor. The complete compressor was mounted on a collecting chamber having an annular air-flow passage simulating the burner annulus of the engine and was driven by an electric motor. The compressor was extensively instrumented to determine the overall performance of the compressor, the characteristic performance of each of the compressor components, the state of the air stream in the simulated burner annulus, and the operation of the compressor bearings. An initial investigation at an equivalent compressor speed of 8000 rpm was made to determine the performance of the compressor and the collecting chamber and to determine the similarity of the air stream at the entrance to the simulated burner annulus. The mechanical performance of the compressor over a range of actual compressors speeds from 3300 to 8000 rpm is reported.

Description: At the request of the Henschel Aircraft Works. A. G. Berlin. three models of the missile "Zitterrochen" were investigated at subsonic velocities.(open jet 215-millimeter diameter) and at supersonic velocities (open jet 110 by 130 millimeters) in order to determine the effect of various wing forms on the air forces and moments. Three-component measurements were taken, and one model was also investigated with deflected control plates.

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: Measurements on three tubes with flow regulated by suction at the trainling edge of the tube are described. It was possible to vary the mass of air flowing through the tube over a large range. Such tubes could be used for shrouded propellers.

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: This report contains the flight-test results of the longitudinal-stability and -control phase of a general flying qualities investigation of the Lockheed P-80A airplane (Army No. 44-85099). The tests were conducted at indicated airspeeds up to 530 miles per hour (0.76 Mach number) at low altitude and up to 350 miles per hour (0.82) Mach number) at high altitude. These tests showed that the flying qualities of the airplane were in accordance with the requirements of the Army Air Forces Stability and Control Specification except for excessive elevator control forces in maneuvering flight and the inadequacy of the longitudinal trimming control at low airspeeds.

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: Estimates of the static stick-fixed stability and control characteristics of the Consolidated Vultee model 240 airplane are presented in this report. The estimates are based on tests of a 0.092-scale powered model in the 10-foot wind tunnel of the Guggenheim Aeronautical Laboratory of the California Institute of Technology. Results of the analysis are evaluated in terms of the Army specifications for stability and control characteristics which are more specific and, in general, equal to or more rigid than the Civil Aeronautics Administration requirements. The stick-fixed stability and control characteristics of the Consolidated Vultee model 240 were found to be satisfactory except for the following: 1) Marginal longitudinal stability in the landing approach (flaps 30 deg, 50% minimum continuous power) with aft center of gravity (31% M.A.C.); 2) Marginal rudder control to hold zero sideslip in a climb after take-off with asymmetric power (flaps 30 deg, left engine inoperative, right engine delivering take-off power) with maximum rudder throw limited to +/- 18 deg; 3) Marginal dihedral effect with flaps 40 deg and engines delivering maximum continuous power.

Description: To determine the trim range in which a seaplane can take off without porpoising, stability tests were made of a Plexiglas model, composed of float, wing, and tailplane, which corresponded to a full-size research airplane. The model and full-size stability limits are in good agreement. After all structural parts pertaining to the air frame were removed gradually, the aerodynamic forces replaced by weight forces, and the moment of inertia and position of the center of gravity changed, no marked change of limits of the stable zone was noticeable. The latter, therefore, is for practical purposes affected only by hydrodynamic phenomena. The stability limits of the DVL family of floats were determined by a systematic investigation independent of any particular sea-plane design, thus a seaplane may be designed to give a run free from porpoising.

Description: A method for calculation of a counterrotating propeller which is similar to Walchner's method for calculation of the single propeller in the free air stream is developed and compared with measurements. Several dimensions which are important for the design are given end simple formulas for the gain in efficiency derived. Finally a survey of the behavior of the propeller for various operating conditions is presented.

Description: Results of measurements on a shrouded propeller are given. The propeller is designed for the high ratio of advance and high thrust loading. The effect of the shape of propeller and shroud upon the aerodynamic coefficients of the propulsion unit can be seen from the results. The highest efficiency measured is 0.71. The measurements permit the conclusion that the maximum efficiency can be essentially improved by shroud profiles of small chord and thickness. The largest static thrust factor of merit measured reaches according to Bendemann, a value of about zeta = 1.1. By the use of a nose split flap the static thrust for thin shroud profiles with small nose radius can be about doubled. In a separate section numerical investigations of the behavior of shrouded propellers for the ideal case and for the case with energy losses are carried out. The calculations are based on the assumption that the slipstream cross section depends solely on the shape of the shroud and not on the propeller loading. The reliability of this hypothesis is confirmed experimentally and by flow photographs for a shroud with small circulation. Calculation and test are also in good agreement concerning efficiency and static thrust factor of merit. The prospects of applicability for shrouded propellers and their essential advantages are discussed.

Description: The requirements on gas turbines for aircraft power units, namely, adequate efficiency, operation at high gas temperatures, low weight, and small dimensions, must be taken into consideration during the design of the blading. To secure good efficiency, it is necessary that the gas flow past the blades as smoothly as possible without separation. This is relatively easily obtainable in the accelerated flow of turbine blading, if the blade spacing is chosen small enough. A small blade spacing, however, is detrimental to the other requirements outlined above. Operation at high gas temperatures usually calls for blade cooling. This cooling is associated with a power input that lowers the turbine efficiency. Since the amount of heat that must be carried off for coding a blade can be influenced rather little, the gross power input for a turbine stage can be reduced by keeping the number of blades to a minimum, that is, with blades of high spacing ratio. But here also a limit is imposed, the exceeding of which is followed by separation of flow. Hence the requirement of finding blade forms on which the flow separates at rather high spacing ratios .

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: A preliminary investigation of an axial-flow gas turbine-propeller engine was conduxted. Performance data were obtained for engine speeds from 8000 to 13,000 rpm and altitudes from 5000 to 35,000 feet and compressor inlet ram pressure ratios from 1.00 to 1.17.

Description: Wind-tunnel measurements on projectiles are discussed. Tests at the Gottingen Tunnel are described. The tunnel operates on the Prandtl principle, that is, a brief stationary air stream produced in an evacuated tank by induction of atmospheric air.

Description: A 19XB-1 combustor was operated under conditions simulating zero-ram operation of the 19XB-1 turbojet engine at various altitudes and engine speeds. The combustion efficiencies and the altitude operational limits were determined; data were also obtained on the character of the combustion, the pressure drop through the combustor, and the combustor-outlet temperature and velocity profiles. At altitudes about 10,000 feet below the operational limits, the flames were yellow and steady and the temperature rise through the combustor increased with fuel-air ratio throughout the range of fuel-air ratios investigated. At altitudes near the operational limits, the flames were blue and flickering and the combustor was sluggish in its response to changes in fuel flow. At these high altitudes, the temperature rise through the combustor increased very slowly as the fuel flow was increased and attained a maximum at a fuel-air ratio much leaner than the over-all stoichiometric; further increases in fuel flow resulted in decreased values of combustor temperature rise and increased resonance until a rich-limit blow-out occurred. The approximate operational ceiling of the engine as determined by the combustor, using AN-F-28, Amendment-3, fuel, was 30,400 feet at a simulated engine speed of 7500 rpm and increased as the engine speed was increased. At an engine speed of 16,000 rpm, the operational ceiling was approximately 48,000 feet. Throughout the range of simulated altitudes and engine speeds investigated, the combustion efficiency increased with increasing engine speed and with decreasing altitude. The combustion efficiency varied from over 99 percent at operating conditions simulating high engine speed and low altitude operation to less than 50 percent at conditions simulating operation at altitudes near the operational limits. The isothermal total pressure drop through the combustor was 1.82 times as great as the inlet dynamic pressure. As expected from theoretical considerations, a straight-line correlation was obtained when the ratio of the combustor total pressure drop to the combustor-inlet dynamic pressure was plotted as a function of the ratio of the combustor-inlet air density to the combustor-outlet gas density. The combustor-outlet temperature profiles were, in general, more uniform for runs in which the temperature rise was low and the combustion efficiency was high. Inspection of the combustor basket after 36 hours of operation showed very little deterioration and no appreciable carbon deposits.

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: An investigation of a heated jet was conducted in conjunction with tests of an axial-flow jet-propulsion engine in the Cleveland altitude wind tunnel. Pressure and temperature surveys were made across the jet 10 and 15 feet behind the jet-nozzle outlet of the engine. Surveys were obtained at pressure altitudes of 10,000, 20,000, 30,000, and 40,000 feet with test-section velocities from 30 to 110 feet per second and test-section temperatures from 60 F to -50 F. From measurements taken throughout the operable range of engine speeds, tail-pipe outlet temperatures from 500 F to 1250 F and jet velocities from 400 to 2200 feet per second were obtained. The jet-survey data presented extend the work previously done with low-velocity and low-temperature jets to the region of high velocities and high temperatures. The results obtained agree with previously determined experimental data and with predicted theoretical expressions for the dimensionless transverse velocity and temperature profiles across a jet. The spread of both the temperature and the velocity profiles was very nearly linear. Dimensionless plots of temperature and velocity along the axis of a heated jet agree with experimental results of tests with a cold jet.

Description: This paper discussed the theory and design of dynamic "pressure augmentors" (diaphragms equal orifice plates and nozzles) and various forms of "pressure multipliers" (simple venturi tubes, Rateau-type multiple venturis, and a combination of shaped nozzle and simple venturi developed by the author). No complete theory of pressure multiplication is yet available; conditions of governing are discussed in relation to pressure-augmenting devices fitted either on the suction or the pressure side of the blower; fluctuations of output and power consumption caused by the presence of an augmentor are analyzed with the result that fitting on the pressure side appears generally preferable. Some considerations on the suitable design and selection of pressure-augmenting devices are appended.

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: Operating characteristics of the 11-stage 4000-pound-thrust axial-flow turbojet engine were determined. A standard compressor and a compressor with the blade angles of the rotor and stator blades increased 5 degrees to obtain greater air flow, were investigated.

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: Combustion chamber performance properties of a 3000-pound-thrust axial-flow turbojet engine were determined. Data are presented for a range of simulated altitudes from 15,000 to 45,0000 feet and a range of Mach numbers from 0.23 to 1.05 for various modifications of the engine.

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: The compression plane is intended for operation on or close to the surface of the water, and has a hull with a concave bottom which forms the upper surface of a tunnel into which air is forced under pressure to support part of the load. The results of the tests made in Langley tank no. 1 include values of the horizontal forces, trimming moment, and static pressure in the tunnel for a wide range of loads and speeds and two power conditions, and are presented in the form of curves against speed with load as a parameter. The results are scaled up to 10 times the model size for three conditions at which the model is self-propelled at a steady speed. Lift is obtained from the static pressure of air in the tunnel. In general, the ratio of the gross load to the total resistance increases with increase in load and decrease in speed. This ratio varies between l-7 and 5.7 at high speeds and has a maximum value of 7. The total resistance is nearly the same for both power conditions except at low speeds and heavy loads. No abrupt change in forces on the hull or flow around the hull occurs in. the region of zero draft. The centers of pressure are generally far aft. At the most efficient trim (1.2'), considerable bow-up moment would be required for practicable operation. There is no abrupt transition from the air-borne to the water- borne condition.

Description: An investigation was conducted in the NACA Cleveland altitude wind tunnel to determine the operational characteristics of the Westinghouse 19B-2, 19B-8, and 19XB-l jet-propulsion engines. The 19B engine is one af the earliest experimental Westinghouse axial flow engines. The 19XB-1 engine is an experimental prototype of the Westinghouse 15 series, having a rated thrust of 1400 pounds. Improvements in performance and operational characteristics have resulted in the 19XB-2B engine with a rated thrust of 1600 pounds. The operational characteristics were determined over a range of simulated altitudes from 5000 to 30,000 feet for the 19B engines and from 5000 to 35000 feet for the 19XB-l engine at airspeed from 20 to 380 miles per hour. The affects of altitude and airspeed on such operating characteristics as operating range, stability of combustion, starting, acceleration, and functioning of the fuel-control system are discussed. Damage to the engines that occurred during the investigation is also briefly discussed. The changes made in the combustion-chamber configuration to improve the operating we are described.

Description: Measurements were made, in dives to transonic speeds, of the static-pressure position error at a distance of one chord ahead of the McDonnell XF-88 airplane. The airplane incorporates a wing which is swept back 35 deg along the 0.22 chord line and utilizes a 65-series airfoil with a 9-percent-thick section perpendicular to the 0.25-chord line. The section in the stream direction is approximately 8-percent thick. Data up to a Mach number of about 0.97 were obtained within an airplane normal-force-coefficient range from about 0.05 to about 0.68. Data at Mach numbers above about 0.97 were obtained within an airplane normal-force-coefficient range from about 0.05 to about 0.68. Results of the measurements indicate that the static-pressure error, within the accuracy of measurement, is negligible from a Mach number of 0.65 to a Mach number of about 0.97. With a further increase in Mach number, the static-pressure error increases rapidly; at the highest Mach number attained in these tests (about M = 1.038), the error increases to about 8 percent of the impact pressure. Above a Mach number of about 0.975, the recorded Mach number remains substantially constant with increasing true Mach number; the installation is of no value between a Mach number of about 0.975 and at least 1.038, as the true Mach number cannot be obtained from the recorded Mach number in this range. Previously published data have shown that at 0.96 chord ahead of the wing tip of the straight-wing X-l airplanes, a rapid rise of position error started at a Mach number of about 0.8. In the case of the XF-88 airplane, this rise of position error was delayed, presumably by the sweep of the wing, to a Mach number of about 0.97.

Description: An investigation has been made in the Langley stability tunnel to determine the low-speed static stability and control characteristics of a model of the Bell MX-776. The results show the model to be longitudinally unstable in the angle-of-attack range around zero angle of attack and to become stable at moderate angles of attack. The results of the present investigation agree reasonably well with results obtained in other facilities at low speed. The present pitching-moment results at low Mach numbers also agree reasonably well with unpublished results of tests of the model at supersonic Mach numbers (up to Mach number 1.86). Unpublished results at moderate and high subsonic speeds, however, indicate considerably greater instability at low angles of attack than is indicated by low-speed results. The results of the present tests also showed that the pitching-moment coefficients for angles of attack up to 12deg remained fairly constant with sideslip angle up to 12deg. The elevators tested produced relatively large pitching moments at zero angle of attack but, as the angle of attack was increased, the elevator effectiveness decreased. The rate of decrease of elevator effectiveness with angle of attack was less for 8deg than for 20deg elevator deflection. Therefore although 8deg deflection caused an appreciable change in longitudinal trim angle and trim lift coefficient a deflection of 20deg caused only a small additional increase in trim angle and trim lift coefficient.