ALBERT

Description: Average spanwise blade temperatures and cooling-air pressure losses through a small (1.4-in, span, 0.7-in, chord) air-cooled turbine blade were calculated and are compared with experimental nonrotating cascade data. Two methods of calculating the blade spanwise metal temperature distributions are presented. The method which considered the effect of the length-to-diameter ratio of the coolant passage on the blade-to-coolant heat-transfer coefficient and assumed constant coolant properties based on the coolant bulk temperature gave the best agreement with experimental data. The agreement obtained was within 3 percent at the midspan and tip regions of the blade. At the root region of the blade, the agreement was within 3 percent for coolant flows within the turbulent flow regime and within 10 percent for coolant flows in the laminar regime. The calculated and measured cooling-air pressure losses through the blade agreed within 5 percent. Calculated spanwise blade temperatures for assumed turboprop engine operating conditions of 2000 F turbine-inlet gas temperature and flight conditions of 300 knots at a 30,000-foot altitude agreed well with those obtained by the extrapolation of correlated experimental data of a static cascade investigation of these blades.

Description: Two short turbojet combustors designed for use with vaporized hydrocarbon fuels were tested in a one-quarter annular duct. The experimental combustors consisted of many small "swirl-can" combustor elements manifolded together. This design approach allowed the secondary mixing zone to be considerably reduced over that of conventional combustors. The over-all combustion lengths, for the two configurations were 13.5 and 11.0 inches, approximately one-half the length of the shortest conventional combustors. These short combustors did not provide combustion efficiencies as high as those for conventional combustors at low pressures. However, over the range of combustor-inlet total-pressures expected in aircraft capable of flight at Mach numbers of 2.5 and above, these short combustors gave very high efficiencies. A combustion efficiency of 97 percent was obtained at a combustor-inlet total-pressure of 25.0 inches of mercury absolute, reference velocity of 120 feet per second, and inlet-air total temperature of 1160 deg R. By proportioning the fuel flow between the manifold rows of can combustor elements, control of the combustor-outlet radial total-temperature profile was demonstrated. Combustor totalpressure loss varied from 0.75 percent of the inlet total pressure at isothermal conditions and a reference velocity of 75 feet per second to 5.5 percent at a total-temperature ratio of 1.8 and a reference velocity of 180 feet per second.

Description: This analysis investigates the application of gas turbine engines at a cruise Mach number of 4.

Description: A program was conducted in an altitude facility at the NACA Lewis laboratory to investigate the effects of rapid inlet pressure oscillations on the operation of a current turbo jet engine. These pressure oscillations were approximately sinusoidal in form and were generated to cover a frequency range of 2 to 75 cycles per second and an amplitude range of 10 to 70 percent of the free-stream total pressure. As the oscillation progressed through the compressor, the amplitude was attenuated considerably and a relatively large phase shift (lag) occurred. Engine stall limits obtained during pressure oscillations differed from quasi-steady-state stall limits as defined by over-all compressor pressure ratio.

Description: A lightweight turbine rotor assembly was devised, and components were evaluated in a full-scale jet engine. Thin sheet-metal airfoils were brazed to radial fingers that were an integral part of a number of thin disks composing the turbine rotor. Passages were provided between the disks and in the blades for air cooling. The computed weight of the assembly was 50 percent less than that of a similar turbine of normal construction used in a conventional turbojet engine. Two configurations of sheet-metal test blades simulating the manner of attachment were fabricated and tested in a turbojet engine at rated speed and temperature. After 8-1/2 hours of operation pieces broke loose from the tip sections of the better blades. Severe cracking produced by vibration was determined as the cause of failure. Several methods of overcoming the vibration problem are suggested.

Description: An investigation has been conducted in the Langley 4- by 4-foot supersonic pressure tunnel to determine the aerodynamic characteristics in pitch and sideslip of a generalized canard airplane model. Two wings of equal area but differing in plan form were investigated. The model was equipped with a trapezoidal canard surface with an area 12 percent of the wing area, a low-aspect-ratio vertical tail, and twin ventral fins. The interference effects of the canard wake on the wing result in little or no gain in the total lift at a Mach number of 1.41 but at a Mach number of 2.01 a substantial portion of the canard lift is retained with a resultant increase in total lift. Because these interference effects of the canard wake appear to be concentrated near the leading edge of the wing, the proper location of the wing leading edge with respect to the center of moments may result in a substantial increase in the moment increment provided by a canard surface even though the total lift provided by the canard is small. For these configurations the trapezoidal wing retained the most lift and had the largest favorable moment increment produced by the canards. The canard configurations have the same characteristic decrease in directional-stability with angle of attack as most conventional high-fineness-ratio supersonic configurations. Although the presence of the canard surface caused a small increase in the directional stability at a Mach number of 1.41 for the delta-wing configuration, the presence of the canards resulted in small decreases in the directional-stability level at a Mach number of 2.01 for both wing configurations. A canard deflection of 15 deg provides an increase in the positive effective dihedral approximately as large as that provided by the presence of the vertical tail. This effect of canard deflection might complicate the lateral-control problem in the case of a rolling pull-up maneuver.

Description: A theoretical method is presented for predicting the dynamic lateral stability characteristics of an airplane towed in tandem by a much larger airplane. Values of period and time to damp to one-half amplitude and rolling motions calculated by an analog computer have been correlated with results of two experimental investigations conducted in the Langley free-flight tunnel which were part of a U.S. Air Force program (Project FICON) to develop a satisfactory arrangement by which a bomber could tow a parasite fighter. In general, the theoretical results agree with the experimental results.

Description: The performance and operational characteristics of two afterburner configurations for the Iroquois turbojet engine were evaluated in an altitude test chamber over a range of afterburner equivalence ratios at afterburner-inlet pressures from 733 to 3186 pounds per square foot absolute. These conditions correspond to an altitude range from 38,700 to 66,800 feet at a flight Mach number of 1.5. The only difference between the two afterburner configurations was in the pattern of afterburner fuel injection. At an afterburner-inlet pressure of approximately 3100 pounds per square foot absolute, corresponding to an altitude of 38,700 feet and a_ flight Mach number of 1.5, the combustion efficiency of both configurations reached peak values of 0.80 to 0.85 at equivalence ratios of 0.35 to 0.40. However, further reduction in the afterburner-inlet pressure severely affected combustion efficiency. For example, at an afterburner inlet pressure level of 700 to 1000 pounds per square foot absolute, the efficiency for both configurations was 0.20 to 0.40.

Description: An investigation of the low-speed static stability and control characteristics of 1/4-scale models of two configurations suitable for lifting reentry from satellite orbit has been made in the Langley free- flight tunnel. One of the models was a thick, all-wing configuration having a delta plan form and the other was a flat delta wing with a half-cone fuselage. The investigation showed that, in general, the all-wing configuration had better longitudinal and lateral stability characteristics than the flat delta configuration.

Description: A flight investigation was made to obtain experimental information on the handling qualities of a normal-acceleration type of automatic longitudinal control system. The control system was installed in a subsonic fighter-type airplane. In hands-off (stick-free) flight the normal-acceleration control system attempted to regulate the normal acceleration to a constant value which is dependent on the automatic-control-system trim setting. In maneuvering flight a given pilot's stick deflection produced a proportional change in normal acceleration, the change in acceleration being independent of flight condition. A small side-located controller stick was used by the pilot to introduce signals into the automatic control system. In the flight program emphasis was placed on the acceleration-limiting capabilities of the control system. The handling qualities were investigated in maneuvers such as slow and rapid pull-ups and turns and also in flight operations such as cruising, stalls, landings, aerobatics, and air-to-air tracking. Good acceleration limiting was obtained with the normal-acceleration control system by limiting the magnitude of the input signal that the pilot could introduce into the control system. The same values of control-system gain settings could be used from an acceleration-limiting stand-point at both 10,000 and 30,000 feet for the complete speed range of the airplane. The response characteristics of the airplane-control system combination were also satisfactory at both high and low altitude with these same values of control-system gain setting. In the pilot's opinion, the normal-acceleration control system provided good stability and control characteristics in flight operations such as cruising, stalls, landings, aerobatics, and air-to-air tracking.

Description: Methods are presented for estimating the directional stability derivative increments contributed by the stabilizing surfaces of subsonic and supersonic aircraft. These methods are strictly applicable at zero angle of attack and small angles of sideslip. The procedure of totaling the incremental coefficients to obtain an estimation of the total empennage side-force and yawing-moment coefficient derivatives is also shown, together with numerical examples. A correlation is presented between estimated and experimental incremental coefficients which indicates that the methods of this report generally estimate the increment of side force gained by the addition of a panel to within +/-10 percent of the experimental value while the yawing-moment increment is generally estimated to within +/-20 percent. This is true for both subsonic and supersonic Mach numbers. An example application of the methods to one of the problems in directional stability, that of minimizing the effect of Mach number on the side-force coefficient derivative of the empennage, is discussed.

Description: A literature survey was conducted to determine the relation between aircraft ignition sources and inflammables. Available literature applicable to the problem of aircraft fire hazards is analyzed and, discussed herein. Data pertaining to the effect of many variables on ignition temperatures, minimum ignition pressures, and minimum spark-ignition energies of inflammables, quenching distances of electrode configurations, and size of openings incapable of flame propagation are presented and discussed. The ignition temperatures and the limits of inflammability of gasoline in air in different test environments, and the minimum ignition pressure and the minimum size of openings for flame propagation of gasoline - air mixtures are included. Inerting of gasoline - air mixtures is discussed.

Description: As part of a general investigation of propellers at high forward speeds, tests of two 2-blade propellers having the NACA 4-(3)(8)-03 and NACA 4-(3)(8)-45 blade designs have been made in the Langley 8-foot high-speed tunnel through a range of blade angle from 20 degrees to 60 degrees for forward Mach numbers from 0.165 to 0.725 to establish in detail the changes in propeller characteristics due to compressibility effects. These propellers differed primarily only in blade solidity, one propeller having 50 percent and more solidity than the other. Serious losses in propeller efficiency were found as the propeller tip Mach number exceeded 0.91, irrespective of forward speed or blade angle. The magnitude of the efficiency losses varied from 9 percent to 22 percent per 0.1 increase in tip Mach number above the critical value. The range of advance ratio for peak efficiency decreased markedly with increase of forward speed. The general form of the changes in thrust and power coefficients was found to be similar to the changes in airfoil lift coefficient with changes in Mach number. Efficiency losses due to compressibility effects decreased with increase of blade width. The results indicated that the high level of propeller efficiency obtained at low speeds could be maintained to forward sea-level speeds exceeding 500 miles per hour.

Description: An investigation has been made to explore the possibilities of axial-flow compressors operating with supersonic velocities into the blade rows. Preliminary calculations showed that very high pressure ratios across a stage, together with somewhat increased mass flows, were apparently possible with compressors which decelerated air through the speed of sound in their blading. The first phase of the investigation was the development of efficient supersonic diffusers to decelerate air through the speed of sound. The present report is largely a general discussion of some of the essential aerodynamics of single-stage supersonic axial-flow compressors. As an approach to the study of supersonic compressors, three possible velocity diagrams are discussed briefly. Because of the encouraging results of this study, an experimental single-stage supersonic compressor has been constructed and tested in Freon-12. In this compressor, air decelerates through the speed of sound in the rotor blading and enters the stators at subsonic speeds. A pressure ratio of about 1.8 at an efficiency of about 80 percent has been obtained.

Description: NACA instrumentation has been installed ii the X-J4 airplanes to obtain stability and control data during the acceptance tests conducted by the Northrop Aircraft Corporation. This report presents data obtained on the stalling characteristics of the airplane in the clean and gear- down configurations. The center of gravity was located at approximately 18 percent of the mean aerodynamic chord during the tests. The results indicated that the airplane was not completely stalled when stall was gradually approached during nominally U accelerated flight but that it was completely stalled during a more abruptly approached stall in accelerated flight. The stall in accelerated flight was relatively mild, and this was attributed to the nature of the variation of lift with angle of attack for the 001-614 airfoil section, the plan form of the wing, and to the fact that the initial sideslip at the stall produced (as shown by wind-tunnel tests of a model of the airplane) a more symmetrical stall pattern.

Description: Contents: Preliminary notes on the efficiency of propulsion systems; Part I: Propulsion systems with direct axial reaction rockets and rockets with thrust augmentation; Part II: Helicoidal reaction propulsion systems; Appendix I: Steady flow of viscous gases; Appendix II: On the theory of viscous fluids in nozzles; and Appendix III: On the thrusts augmenters, and particularly of gas augmenters

Description: An investigation was conducted to determine the effects of water injection on the over-all performance of a modified J33-A-27 turbojet-engine compressor at the design equivalent speed of 11,800 rpm. The water-air ratio by weight was 0.05. With water injection the peak pressure ratio increased 9.0 per- cent, the maximum efficiency decreased 15 percent (actual numerical difference 0.12), and. the maximum total weight flow increased 9.3 percent.

Description: The compressor from the XT-46 turbine-propeller engine was revised by removing the last two rows of stator blades and by eliminating the interstage leakage paths described in a previous report. With the revised compressor, the flow choking point shifted upstream into the last rotor-blade row but the maximum weight flow was not increased over that of the original compressor. The flow range of the revised compressor was reduced to about two-thirds that obtained with the original compressor. The later stages of the compressor did not produce the design static-pressure increase probably because of excessive boundary-layer build-up in this region. Measurements obtained in the ninth-stage stator showed that the performance up to this station was promising but that the last three stages of the compressor were limiting the useful operating range of the preceding stages. Some modifications in flow-passage geometry and blade settings are believed to be necessary, however, before any major improvements in over-all compressor performance can be obtained.

Description: The power plant from a Mark 25 aerial torpedo was investigated both as a two-stage turbine and as a single-stage modified turbine to determine the effect on overall performance of nozzle size and shape, first-stage rotor-blade configuration, and axial nozzle-rotor running clearance. Performance was evaluated in terms of brake, rotor, and blade efficiencies. All the performance data were obtained for inlet total to outlet static pressure ratios of 8, 15 (design), and 20 with inlet conditions maintained constant at 95 pounds per square inch gage and 1000 F for rotor speeds from approximately 6000 to 18,000 rpm.

Description: This report presents the results of wind-tunnel force tests which were conducted to determine the low-speed stability and control characteristics of a full-scale Northrop XSSM-A-3 missile. Tests were made through a range of angles of attack, sideslip, and control deflection, and at various Reynolds numbers. Characteristics of the complete missile are compared with the characteristics of the missile with the landing skids extended, with the vertical tail removed, and with the fuselage alone. No analysis of the data has been made in order to make the results available as soon as possible.

Description: A theoretical investigation has been made to determine the effect on the lateral stability of the Douglas D-58-II airplane of an autopilot sensitive to yawing velocity. The effects of inclination of the gyro spin axis to the flight path and of tire lag in the autopilot were also determined. The flight conditions investigated included landing at sea level, approach condition at 12,000 feet, and cruising at 50,000 feet at Mach numbers of 0.80 and 1.2. The results of the investigation indicated that the lateral stability characteristics of the D-558-II airplane for the flight condition discussed should satisfy the Air Force - Navy period-damping criterion when the proposed autopilot is installed. Airplane motions in sideslip subsequent to a disturbance in sideslip are presented for several representative flight conditions in which a time lag in the autopilot of 0.10 second was assumed.

Description: The investigation of the lateral stability of an automatically controlled glide bomb led also to the attempt of clarifying the influence of a phugoid oscillation or of any general longitudinal oscillation on the lateral stability of a glide bomb. Under the assumption that its period of oscillation considerably exceeds the rolling and yawing oscillation and that c(sub a) is, at least in sections, practically constant, the result of this test is quite simple. It becomes clear that the influence of the phugoid oscillation may be replaced by suitable variation of the rolling-yawing moment on a rectilinear flight path instead of the phugoid oscillation. If the flying weight of the glide bomb of unchanged dimensions is increased, an increase of the flight velocity will be more favorable than an increase of the lift coefficient. The arrangement of the control permits lateral stability to be achieved in every case; a minimum rolling moment due to sideslip proves of great help.

Description: Performance data obtained with recording oscillographs are presented to show the transient response of the General Electric Integrated Electronic Control operating on the J47 RXl-3 turbo-Jet engine over a range of altitudes from 10,000 to 45,000 feet and at ram pressure ratios of 1.03 and 1.4. These data represent the performance of the final control configuration developed after an investigation of the engine transient behavior in the NACA altitude wind tunnel. Oscillograph traces of controlled accelerations (throttle bursts),oontrolled decelerations (throttle chops), and controlled altitude starts are presented.

Description: An investigation of the longitudinal stability and of the all-movable horizontal tail and aileron control of a 1/80-scale reflection-plane model of the Consolidated Vultee Skate 9 seaplane has been made through a Mach number range of 0.6 to 1.16 on the transonic bump of the Langley high-speed 7- by 10-foot tunnel. At moderate lift coefficients (0.4 to 0.8) and below a Mach number of 1.0 the model was statically unstable longitudinally. The static longitudinal stability of the model at low lift coefficients increased with Mach number corresponding to a shift in aerodynamic center from 37 percent mean aerodynamic chord at a Mach number of 0.60 to 64 percent at a Mach number of 1.10. Estimates indicate that the tail deflection angle required for steady flight and for accelerated maneuvers of the Skate 9 airplane would probably not vary greatly with Mach number at sea level, but for accelerated maneuvers at altitude the tail deflection angle would probably vary erratically with Mach number. The variation of rolling-moment coefficient with aileron deflection angle was approximately linear, agreed well with theory, and held for the range of aileron deflections tested (-17.1 deg to 16.6 deg). At low lift coefficients the drag rise occurred at Mach numbers of 0.95 and 0.90 for the wing alone and the complete model, respectively. The effects of the canopy on the model were small. For the ranges investigated, angle-of-attack and Mach number changes caused no large pressure drops in the jet-engine duct.

Description: A model of the Convair Y2-2 airplane was tested in Langley tank no. 2 to determine whether satisfactory stability in yawed landings was possible with a certain ventral fin. Free-body landings were made in smooth and rough water at two speeds and two rates of descent with the model yawed 15deg. The behavior of the model was determined by visual observations and from motion-picture re.cords. It was concluded that satisfactory stability was possible with the ventral fin as tested but that the characteristics of the model shock absorbers and the settings of the elevon control surfaces had an appreciable influence on behavior.

Description: A modified J33-A-27 compressor was operated over a range of equivalent impeller speeds from 6100 to 13,250 rpm in order to obtain the over-all compressor performance. At the equivalent design speed of 11,800 rpm, the maximum efficiency of 0.764 and peak pressure ratio of 4.56 occurred at an equivalent weight flow of 104.07 pounds per second. At the highest equivalent speed (13,250 rpm) a maximum efficiency of 0.711, a maximum equivalent weight flow of 123.00 pounds per second, and a peak pressure ratio of 5.76 were obtained.

Description: The data obtained from the flight of a simplified (dummy) rocket-propelled model of the MX-656 have been analyzed to determine the booster-model characteristics and the model-alone characteristics up to a Mach number of 1.3. The data indicate that the model-booster combination is satisfactory. The model alone is longitudinally stable i n the Mach number range covered by the test (0.9 to 1.3) with the center of gravity at -15 percent of the mean aerodynamic chord. With the stabilizer setting at 0 deg. the variation of normal-force coefficient with Mach number is not large. The total-drag-coefficient variation with Mach number is not unusual. About 12 percent of the total drag at a Mach number of 1.3 can be attributed to body base drag.

Description: An investigation is being conducted to determine the performance of the 12-stage axial-flow compressor of the XT-46 turbine-propeller engine. This compressor was designed to produce a pressure ratio of 9 at an adiabatic efficiency of 0.86. The design pressure ratios per stage were considerably greater than any employed in current aircraft gas-turbine engines using this type of compressor. The compressor performance was evaluated at two stations. The station near the entrance section of the combustors indicated a peak pressure ratio of 6.3 at an adiabatic efficiency of 0.63 for a corrected weight flow of 23.1 pounds per second. The other, located one blade-chord downstream of the last stator row, indicated a peak pressure ratio of 6.97 at an adiabatic efficiency of 0.81 for a corrected weight flow of 30.4 pounds per second. The difference in performance obtained at the two stations is attributed to shock waves in the vicinity of the last stator row. These shock waves and the accompanying flow choking, together with interstage circulatory flows, shift the compressor operating curves into the region where surge would normally occur. The inability of the compressor to meet design pressure ratio is probably due to boundary-layer buildup in the last stages, which cause axial velocities greater than design values that, in turn, adversely affect the angles of attack and turning angles in these blade rows.