ALBERT

Description: An investigation was conducted to compare the performance of two 25-ft-diam rotors which had identical dimensions and were similar in construction but different in blade airfoil-sections. Tests were conducted at indicated blade pitch angles from 3 degrees to 11.5 degrees and rotor speeds of 200, 290, and 371 rpm. The 23012.6 rotor required 2 percent less power to hover than the 0012.6. At thrust coefficients above design, the performance of the 23012.6 became better than the 0012.6 rotor.

Description: A relatively simple equation has been found to express with fair accuracy, variation in manifold-charge temperature with charge in engine operating conditions. This equation and associated curves have been checked by multi cylinder-engine data, both test stand and flight, over a wide range of operating conditions. Average mixture temperatures, predicted by the equations of this report, agree reasonably well with results within the same range of carburetor-air temperatures from laboratories and test stands other than the NACA.

Description: An investigation has been conducted in the Langley rectangular high-speed tunnel to determine the effect of compressibility on the pressure distribution for a modified NACA 65,3-019 airfoil having a 0.20-chord flap. The investigation was made for an angle-of-attack range extending from -2 to 12 deg at .20 flap deflections from 0 to -12 deg. Test data were obtained for Mach numbers from 0.28 to approximately 0.74. The results show that the effectiveness of the trailing-edge-type control surface rapidly decreased and approached zero as the Mach number increased above the critical value.

Description: In conjunction with a program of research on the general problem of stability of airplanes in the climbing condition, tests have been made of a spring-loaded tb which. is referred to as a ?springy tab,? installed on the elevator of a low-wing scout bomber. The tab was arranged to deflect upward with decrease in speed which caused an increase in the pull force required to trim at low speeds and thereby increased the stick-free static longitudinal stability of the airplane. It was found that the springy tab would increase the stick-free stability in all flight conditions, would reduce the danger of inadvertent stalling because of the definite pull force required to stall the airplane with power on, would reduce the effect of center-of-gravity position on stick-free static stability, and would have little effect on the elevator stick forces in accelerated f11ght. Another advantage of the springy tab is that it might be used to provide almost any desired variation of elevator stick force with speed by adjusting the tab hinge-moment characteristics and the variation of spring moment with tab deflection. Unlike the bungee and the bobweight, the springy tab would provide stick-free static stability without requiring a pull force to hold the stick back while taxying. A device similar to the springy tab may be used on the rudder or ailerons to eliminate undesirable trim-force variations with speed.

Description: As part of an investigation to increase the power output of the V-1710-93 engine at altitude, the engine-stage supercharger was combined with a constant-area vaneless diffuser designed to improve the performance of the engine-stage supercharger at the rated engine operating point. The performance of the modified supercharger was investigated in a variable-component supercharger test rig and compared with that of the standard supercharger with an 8-vaned diffuser. A separate evaluation of the component efficiencies and a study of the flow characteristics of the modified supercharger was made possible by internal diffuser instrumentation. At the volume flow required by the engine for rated operating conditions, the modified supercharger increased the over-all adiabatic efficiency 0.05 and the over-all pressure coefficient 0.035. Furthermore, the capacity of the engine-stage supercharger was increased by replacing the standard 8-vaned diffuser with the vaneless diffuser. The peak over-all adiabatic efficiency for the modified supercharger, however, was 0.05 to 0.07 lower than that of the standard unit over the range of tip speeds investigated. The improved performance of the modified supercharger at rated engine operating conditions resulted from a shift of the point of peak adiabatic efficiency and pressure coefficient of the standard supercharger to a higher flow. The energy loss through the vaneless diffuser was found to be small. Because of the restricted diffuser diameter, however, diffusion was inadequate, which resulted in a relatively small static-pressure rise through the diffuser, high diffuser-exit velocities, and excessive collector-case losses.

Description: A dynamically similar model of the Army P-38 airplane was tested to determine the best way to land this airplane on the water and to determine its probable ditching performance. The tests consisted of ditching the model at various landing attitudes, flap settings, speeds, weights, and conditions of simulated damage. The model was ditched in calm water from the tank towing carriage and a few ditching were made in both calm and rough water at the outdoor catapult. The performance of the model was determined by making visual observations, by recording lengths of run and time histories of decelerations, and by taking motion pictures of the ditchings.

Description: In the present report the true weight distribution law of the wing structure along the span is investigated. It is shown that the triangular distribution and that based on the proportionality to the chords do not correspond to the actual weight distribution, On the basis of extensive data on wings of the CAHI type airplane formulas are obtained from which it is possible to determine the true diagram of the structural weight distribution along the span from a knowledge of only the geometrical dimensions of the wing. At the end of the paper data are presented showing how the structural weight is distributed between the straight center portion and the tapered portion as a function of their areas.

Description: An investigation was conducted in the Langley impact basin of the water loads on a half scale model of the XJL-1 hull whose forebody has a vee bottom with exaggerated chine flare. The impact loads, moments, and pressures were determined for a range of landing conditions. A normal full-scale landing speed of 86 miles per hour was represented with effective flight paths ranging from 0.6deg to 11.6deg. Landings were made with both fixed trim and free-to-trim mounting of the float over a trim range of -15deg to 12deg into smooth water and into waves having equivalent full-scale length. of 120 feet and heights ranging from 1 to 4 feet. All data and results presented in this report are given in terms of equivalent full-scale values. Summary tables and illustrative plots are used in presenting the material. The following maximum values of load and pressure are those which are apropos for effective flight paths less than 6.5deg which was the maximum value obtained in tests with the XJL-1 hull model representing full-scale landings with vertical velocity of 4.5 feet per second into 4-foot waves. The maximum local pressure on the flat portion of the bottom is 130 pounds per square inch which was measured on a 2-inch-diameter circular area near the step. The maximum local pressure obtained in the curved area near the chines is 200 pounds per square inch. This pressure was also measured near the step.

Description: Tables I and II of this report summarize the gust and draft velocity data for thunderstorm flights 25 and 26 of August 21, 1946 and August 22, 1946, respectively. These dta were evaluated from records of NACA instruments installed in P-61C airplanes and are of the type presented in reference 1 for previous flights. Table III summarizes the readings of a milliammeter which was used in conjunction with other equipment to indicate ambient air temperature during thunderstorm surveys. These data were read from motion-picture records of the instrument and include all cases in which variations in the instrument indications were noted during the present flights.

Description: An investigation was made in the Cleveland Altitude wind tunnel to determine the performance of an Aeroproducts H20C-162-X11M2 four-blade propeller on a YP-47M airplane at high blade loadings and high engine powers. The propeller characteristics were obtained for a range of power coefficients from 0.30 to 1.00 at free-stream Mach numbers of 0.40 and 0.50. The results of the force measurements are indicative only of trends in propeller efficiency with changes in power coefficient and advance-diameter ratio because unknown interference effects existed during the investigation. At a free-stream Mach number of 0.40, the envelopes of the efficiency curves decreased about 11% between advance-diameter ratios of 2.40 and 4.40. An increase in power coefficient from 0.30 to 0.80 at an advance-diameter ratio of 2.40 had little effect on the propeller efficiency. A change in power coefficient from 0.40 to 1.00 at an advance-diameter ratio of 4.40 increased the propeller efficiency by about 40%. For conditions below the stall the thrust loading on the outboard blade sections increased more rapidly than on the inboard sections as the power coefficient was increased or as the advance-diameter ratio was decreased. For conditions beyond the stall, the thrust loading decreased on the outboard sections and increased on the inboard sections.

Description: Three modifications of the auxiliary-stage supercharger for the V-1710-93 engine were designed and tested as part of an investigation to improve the power output and the altitude performance of the engine. A 12-vane diffuser was substituted for the standard 11-vane diffuser, and a vaneless discharge passage and a modified scroll were designed to increase the flow capacity of the supercharger and thereby to increase the performance at the high volume flows required by the engine. With the 12-vane diffuser installed and the carburetor replaced by an adapter, the equivalent volume flow at the peak efficiency point was increased 25 percent at the lowest speed investigated and 9.5 percent at the highest speed. When the carburetor was used, any increase in equivalent volume flow was masked by choking in the carburetor. A small decrease in the peak adiabatic efficiency resulted from using the 12-vane diffuser. At the high volume flows where the supercharger is required to operate, the performance was improved by the 12-vane diffuser. With the vaneless discharge passage, the surge-free range of the supercharger was increased 35 percent at the lowest tip speed investigated by increasing the maximum air flow. The maximum air flow at high tip speeds was again limited by choking in the carburetor, which masked the effect of the vaneless discharge passage on the maximum air flow. At the high volume flows near the operating point of the supercharger, the performance with the vaneless discharge passage was better than that with the standard 11-vane diffuser. At the low volume flows when the standard 11-vane diffuser gave better performance. The modified scroll gave performance characteristics that were practically the same as those of the standard scroll except at high tip speeds, where the peak performance was improved.

Description: An altitude-wind-tunnel investigation has been made to determine the performance of Hamilton Standard 6507A-2 four-blade and three-blade propellers on a YP-47M airplane at high blade loadings and high engine powers. Characteristics of the four-blase propeller were obtained for a range of power coefficients from 0.10 to 1.00 at free-stream Mach numbers of 0.20, 0.30, 0.40. Characteristics of the three-blade propeller were obtained for a range of power coefficients from 0.30 to 1.00 at a free-stream Mach number of 0.40. Results of the force measurements indicate primarily the trend of propeller efficiency for changes in power coefficient or advance-diameter ratio because no corrections for the effects of tunnel-wall constriction on the installation were applied. Slipstream surveys are presented to illustrate blade thrust load distribution for certain operating conditions. Within the range of advance-diameter ratios investigated at each free-stream Mach number, the efficiency of the four-blade propeller decreased as the power coefficient was increased from 0.10 to 1.00. For the three-blade propeller, nearly constant maximum efficiencies were obtained for power coefficients from 0.32 to 0.63 at advance-diameter ratios between 1.90 and 3.00. In general, for conditions below the stall and critical tip Mach number, the maximum thrust load shifted from the inboard sections toward the tip sections as the power coefficient was increased or as the advance-diameter ratio was decreased. For conditions beyond the stall or critical tip Mach number, losses in thrust occurred on the outboard blade sections owing to flow break-down; the thrust load increased slightly on the inboard sections.

Description: Tests have been made in Langley tank no. I of a dynamic model of the Consolidated Vultee PB2Y-3 airplane. These tests were made using an alternate hull form, the purpose of which was to reduce the bow spray and eliminate the landing instability which are objectionable features of the production design. The major differences from the PB2Y-3 hull included a deeper step to improve the landing stability , and a lengthened forebody and increased beam to reduce the sway in the propellers and on the flaps. The tests showed that the spray characteristics of the revised hull form were much better than that ot the production design. In addition the take-off and landing stability of the model with the alternate hull were satisfactory.

Description: Tests of a powered dynamic model of the Columbia XJL-1 amphibian were made in Langley tank no.1 to determine the hydrodynamic stability and spray characteristics of the basic hull and to investigate the effects of modifications on these characteristics. Modifications to the forebody chime flare, the step, and the afterbody, and an increase in the angle of incidence of the wing were included in the test program. The seaworthiness and spray characteristics were studied from simulated taxi runs in smooth and rough water. The trim limits of stability, the range of stable positions of the enter of gravity for take-off, and the landing stability were determined in smooth water. The aerodynamic lift, pitching moment, and thrust were determined at speeds up to take-off speed.

Description: The XF-12 airplane is a high performance, photo-reconnaissance aircraft designed by the Republic Aviation Corporation for Army Air Forces. A series of tests of a 1/8.33-scale powered model was conducted in the Langley 9-foot pressure tunnel to obtain information relative to the aerodynamic design of the airplane. This report presents the results of tests to determine the static longitudinal stability and stalling characteristics of the model. From this investigation it was indicated that the airplane will possess a positive static margin for all probable flight conditions. The stalling characteristics are considered satisfactory in that the stall initiates near the root section and progresses toward the tips. Early root section stalling occurs, with the flaps retracted and may cause undesirable tail buffeting and erratic elevator control in the normal flight range. From considerations of sinking speed landing flap deflections of 40 degrees may be preferable to 55 degrees of 65 degrees.

Description: Tests were performed on a partial span of the wing of a McDonnell XFD-1 airplane to determine a combination of sealed internal balance and spring-tab stiffness for the aileron that would give satisfactory stick-force characteristics for the airplane. Two sealed internal balances were tested in combination with spring tabs of various stiffnesses. One of the combinations was tested at several speeds to determine the variation of stick force with speed. Estimates, based on the results of the tests, indicate that for this airplane any reduction of stick force by use of the spring tab reduces the helix angle pb/2V below the required value of 0.09. The estimates show that, of the configurations tested, the most satisfactory combination for obtaining a stick force of 30 pounds at 300 miles per hour indicated airspeed is a 0.48-chord internal balance in combination with a spring-tab stiffness of 500 pounds per inch. With this combination, a wing-tip helix angle of 0.078 is estimated. Stick-force curves for all configurations show a rapid increase in stick force above approximately 20 deg. total aileron deflection.

Description: An investigation was conducted in the Cleveland altitude wind tunnel to determine the aerodynamic characteristics and the oil delivery critical altitude of the oil-cooler installation of an XTB2D-1 airplane. The investigation was made with the propeller removed end with the engine operating at 1800 brake horsepower, an altitude of 15,000 feet (except for tests of oil-delivery critical altitude), oil-cooler flap deflections from -20 degrees to 20 degrees and inclinations of the thrust axis of 0 degrees, 1.5 degrees, and 6 degrees. At an inclination of the thrust axis of 0 degrees and with the propeller operating, the total-pressure recovery coefficient at the face of the oil cooler varied from 0.84 to 1.10 depending on the flap deflection. With the propeller removed, the best pressure recovery at the face of the oil cooler was obtained at an inclination of the thrust axis of 1.5 degrees. Air-flow separation occurred on the inner surface of the upper lip of the oil-cooler duct inlet at an inclination of the thrust axis of 0 degrees and on the inner surface of the lower lip at 6 degrees. Static pressure coefficients over the duct lips were sufficiently low that no trouble from compressibility would be encountered in level flight. The oil-delivery critical altitude at cruising power (2230 rpm, 1675 bhp) was approximately 18,500 feet for the oil system tested.

Description: This paper presents the results of the aileron investigation and includes rolling-moment, yawing-moment, and aileron hinge-moment coefficients and pressure coefficients across the aileron-balance seal through a range of angle of attack, tab deflection, and aileron deflection with flaps neutral and deflected 20 degrees and 55 degrees. Some of the effects of wing roughness and balance seal leakage on the aileron and tab characteristics are also presented.

Description: By the method of images the horizontal steady motion of a wing at small heights above the ground was investigated in the wind tunnel, A rectangular wing with Clark Y-H profile was tested with and without flaps. The distance from the trailing edge of the wing to the ground was varied within the limits 0.75 less than or = s/c less than or = 0.25. Measurements were made of the lift, the drag, the pitching moment, and the pressure distribution at one section. For a wing without flaps and one with flaps a considereble decrease in the lift force and a,drop in the drag was obtained at angles of attack below stalling. The flow separation near the ground occurs at smaller angles of attack than is the case for a great height above the ground. At horizontal steady flight for practical values of the height above the ground the maximum lift coefficient for the wing without flaps changes little, but markedly decreases for the wing with flaps. Analysis of these phenomena involves the investigation of the pressure distribution. The pressure distribution curves showed that the changes occurring near the ground are not equivalent to a change in the angle of attack. At the lower surface of the section a very strong increase in the pressures is observed. The pressure changes on the upper surface at angles of attack below stalling are insignificant and lead mainly to an increase in the unfavorable pressure gradient, resulting in the earlier occurrence of separation. For a wing with flaps at large angles of attack for distances from the trailing edge of the flap to the ground less than 0.5 chord, the flow between the wing end the ground is retarded so greatly that the pressure coefficient at the lower surface of the section is very near its limiting value (P = 1), and any further possibility of increase in the pressure is very small. In the application an approximate computation procedure is given of the change of certain aerodynamic characteristics for horizontal steady flight near the ground.

Description: The results obtained from an evaluation for gust and draft velocities of acceleration and airspeed-altitude records taken by NACA recording instruments installed in P-61c airplanes participating in thunderstorm flights 6, 7, and 8 of July 9, 1946, July 10, 1946, and July 11, 1946, respectively, are presented herein. These data are summarized in tables I and II. In accordance with a recent discussion with a member of the U.S. Weather Bureau staff, the tabulated results for the present flight include in addition to data of the type presented in reference 1, the initial heading of the airplane for each traverse, the pressure altitude at the start of each traverse in increments of 500 feet, and the gust gradient distance when it could be evaluated. The cloud entry and exit times for the present data were taken from motion-picture records of the pilot's instrument panels whenever such records were available while the length fo the traverses in seconds and feet was taken from the airspeed-altitude records. In many cases, however, poor agreement is indicated between the duration of the cloud traverses as obtained from the motion-picture records and from the airspeed-altitude records. This result is believed to be due to camera stoppages, inaccurate spring mechanisms of the clocks, and loss of motion-picture record in exposure or development. With reference to the evaluation of gust data, the nominal threshold was about 2 feet per second. In making gust counts to this threshold, some gusts below that threshold have been included due to limitations of the procedure used. Thus, it will be noted that in some instances gust counts are given in table I although now corresponding gust velocities are listed.

Description: The results of an investigation of a 1/3-scale model of the Chance Vought XF5U-1 airplane in the Langley full-scale tunnel are presented in this report. The maximum lift and stalling characteristics of several model configurations, the longitudinal stability characteristics of the model, and the effectiveness of the control surfaces were determined with the propellers removed. The propulsive characteristics, the effect of propeller operation on the lift, and the static thrust of the model propellers were determined at several propeller-blade angles. The results with the propellers removed showed that the maximum lift coefficient of the complete model configuration was only 0.97 was compared with the value of 1.31 for the model configuration in which the engine-air ducts and canopy are removed. The model with the propellers removed (normal center-of-gravity position) has a positive static margin, stick fixed, varying from 5 to 13 percent of the mean aerodynamic chord throughout the unstalled range of lift coefficients. The unit horizontal tail is sufficiently powerful to trim the airplane with the propellers removed throughout the unstalled range of lift coefficients. The peak propulsive efficiencies for beta = 20 degrees and beta = 30 degrees were increased 7 percent at C(sub L) congruent to 0.67 and 20 percent at C(sub L) congruent to 0.74, respectively, with the propellers rotating upward in the center than with the propellers rotating downward in the center. Indications are that the minimum forward-flight speed of the airplane for full-power operation at sea level will be about 90 miles per hour. Decreasing the weight and increasing the power reduced this value of minimum speed and there were no indications from the results of a lower limit to the minimum speed.

Description: During the first flight tests of the Republic XP-84 airplane it was discovered that there was a complete lack of stall warning. A short series of development tests of a suitable stall-warning device for the airplane was therefore made on a 1/5-scale model in the Langley 300 MPH 7- by 10-foot tunnel. Two similar stall-warning devices, each designed to produce early root stall which would provide a buffet warning, were tested. It appeared that either device would give a satisfactory buffet warning in the flap-up configuration, at the cost of an increase of 8 or 10 miles per hour in minimum speed. Although neither device seemed to give a true buffet warning in the flaps-down configuration, it appeared that either device would improve the flaps-down stalling characteristics by lessening the severity of the stall and by maintaining better control at the stall. The flaps-down minimum-speed increase caused by the devices was only 1 or 2 miles per hour.

Description: The gust and draft velocities evaluated from acceleration and airspeed-altitude records taken by NACA instruments installed in P-61c airplanes participating in thunderstorm flights 9, 10, and 11 of July 12, 1946, July 17, 1946, and July 18, 1946, respectively, are presented in references 1 and 2 for previous flights. In accordance with a recent discussion with a member of the U.S. Weather Bureau staff, motion-picture records of the pilots' instrument panels for the present flights were inspected to note variations in the readings of a milliammeter used in conjunction with other equipment to indicate ambient air temperature. The inspection indicated that the instrument read zero throughout all traverses.

Description: At the request of the Air Materiel Command, Army Air Forces, flight tests were conducted on a P-5lD-20-NA (AAF No. 44-63826) airplane equipped with a horn-balanced rudder. This rudder was fitted with an unbalancing tab as was the original product fan rudder. Tests were made both with the unbalancing tab in operation and with the unbalancing tab locked. The modification to the original vertical tail consisted of removing the cap from the top of the fin and adding 1.91 square feet of area to the rudder as the horn balance and 0.82 square foot of area to the top of the rudder aft of the hinge line. A comparison of the directional stability and control characteristics of the P-51D airplane with three different vertical-tail configurations - the horn balance, original production, and extended tail (with unbalancing tab locked) configurations are presented herein. The tests of the horn-balanced rudder were conducted at the Langley Laboratory in 1945. The tests of the original-tail configuration were previously conducted at the Ames Aeronautical Laboratory. Tests of the extended-tail configuration were conducted at the Langley Laboratory and are reported.

Description: Summaries of the gust and draft velocities evaluated from acceleration and airspeed-altitude records taken by NACA instruments installed n P-61c airplanes participating in thunderstorm flights 12 and 13 of July 19, 1946, and July 20, 1946, respectively, are presented in tables I and II herein. These data are of the type presented in reference 1 for previous flights. Inspection of the motion picture records of the pilots' instrument panels for the present flights indicated that the milliameter connected to equipment for measuring ambient air temperature read zero throughout all traverses.

Description: Two modified fuel-injection systems, a drilled-inducer type and a spinner type, that prevent serious fuel-evaporation icing were installed on a V-type, liquid-cooled aircraft engine and a preliminary investigation was conducted to determine the effect on engine operating characteristics. The spinner system was also ground - and flight tested on a twin-engine fighter airplane. Flight measurements of cylinder-head temperature over a range of fuel-air ratios and engine power conditions were made at an altitude of approximately 10,000 feet. Starting and accelerating of the engine on the ground were unaffected by the fuel-injection modifications. During the flight investigation, no appreciable variation occurred between the maximum and minimum cylinder-head temperature with the standard and modified system for the same power condition and no irregularity of mixture distribution could be detected throughout the power range of the engine. Normal mixture distribution was also indicated by a similar response of cylinder-head temperature for variations of fuel-air ratio at manifold pressures of 25 and 35 inches of mercury absolute. Both modified fuel-injection systems required less fuel-nozzle pressure than the standard system to obtain the desired fuel-air ratio for given air-flow condition.

Description: The results of tests made to determine the aerodynamic characteristics of a solid brake, a slotted brake, and a dive-recovery flap mounted on a high aspect ratio wing at high Mach numbers are presented. The data were obtained in the Langley 8-foot high-speed tunnel for corrected Mach numbers up to 0.940. The results have been analyzed with regard to the suitability of dive-control devices for a proposed high-speed airplane in limiting the airplane terminal Mach number by the use of dive brakes and in achieving favorable dive-recovery characteristics by the use of a dive-recovery flap. The analysis of the results indicated that the slotted brake would limit the proposed airplane terminal Mach number to values below 0.880 for altitudes up to 35,000 feet and a wing loading of 80 pounds per square foot and the dive-recovery flap would produce trim changes required for controlled pull-outs at 25,000 feet for a Mach number range from 0.800 to 0.900. Basic changes in spanwise loading are presented to aid in the evaluation of the wing strength requirements.

Description: An investigation of the 19XB-2A gas turbine is being conducted at the Cleveland laboratory to determine the effect on turbine performance of various inlet pressures, inlet temperatures, pressure ratios, and wheel speeds. The engine of which this turbine is a component is designed to operate at an air flow of 30 pounds per second at a compressor rotor speed of 17,000 rpm at sea-level conditions. At these conditions the total-pressure ratio is 2.08 across the turbine and the turbine inlet total temperature is 2000 degrees R. Runs have been made with turbine inlet total pressures of 20, 30, 40, and 45 inches of mercury absolute for a constant total pressure ratio across the turbine of 2.40, the maximum value that could be obtained. Additional runs have been made with total pressure ratios of 1.50 and 2.00 at an inlet total pressure of 45 inches of mercury absolute. All runs were made with an inlet total temperature of 800 degrees R over a range of corrected turbine wheel speeds from 40 to 150 percent of the corrected speed at the design point. The turbine efficiencies at these conditions are presented.

Description: Because the results of preliminary flight tests had indicated. the P-63A-1 airplane possessed insufficient directional stability, the NACA and the manufacturer (Bell Aircraft Corporation) suggested three vertical-tail modifications to remedy the deficiencies in the directional characteristics. These modifications included an enlarged vertical tail formed by adding a tip extension to the original vertical tail, a large sharp-edge ventral fin, and a small dorsal fin. The enlarged vertical tail involved only a slight increase in total vertical-tail area from 23.73 to 26.58 square feet but a relatively much larger increase in geometric aspect ratio from 1.24 to 1.73 based on height and area above the horizontal tail. At the request of the Air Material Command, Army Air Forces, flight tests were made to determine the effect of these modifications and of some combinations of these modifications on the directional stability and control characteristics of the airplane, In all, six different vertical-tail. configurations were investigated to determine the lateral and directional oscillation characteristics of the airplane, the sideslip characteristics, the yaw due to ailerons in rudder-fixed rolls from turns and pull-outs, the trim changes due to speed changes; and the trim changes due to power changes. Results of the tests showed that the enlarged vertical tail approximately doubled the directional stability of the airplane and that the pilots considered the directional stability provided by the enlarged vertical tail to be satisfactory. Calculations based on sideslip data obtained at an indicated airspeed of 300 miles per hour showed that the directional stability of the airplane with the original vertical tail corresponded to a value of 0(sub n beta) of -0.00056 whereas for the enlarged vertical tail the estimated va1ue of C(sub n beta) was -0.00130, The ventral fin was found to increase by a moderate amount the directional stability of the airplane with the original vertical tail for smal1 sides1ip angles at low speeds but little consistent change in directional stability was effected by the ventral fin at higher speeds, The effectiveness of the ventral fin was generally much less when used with the enlarged vertical tail than when used with the original vertical tail. The ventral and dorsal fins were found to be very effective in eliminating rudder-force reversals which occurred in low-speed, high-engine-power, sideslipped conditions of flight . Sideslip tests at two altitudes for approximately the sane engine power and indicated airspeed showed that a small decrease in static directional stability occurred with increasing altitude and this decrease in stability was attributed to the increased propeller blade angles required at high altitudes. The variations of rudder pedal force with indicated airspeed using normal rated power and a constant rudder tab setting through the speed range were desirably small for all the configurations tested. The rudder pedal force changed by about 50 pounds for a power change from engine idling power, to normal rated power and this pedal force change was largely independent of airspeed or of vertical-tail configuration for the various configurations tested.

Description: An investigation of the performance of several propellers on the YP-47M airplane at high blade loadings has been conducted in the Cleveland altitude wind tunnel at the request of the Air Materiel Command, Army Air Forces. As part of the program, a study was made of a Curtiss 836-14C2-18R1 four-blade propeller. The investigation was made for a range of power coefficients from 0.10 to 1.00 at free-stream Mach numbers of 0.30, 0.40, and 0.50 for density altitudes from 10,000 to 45,000 feet, engine powers from 150 to 2500 brake horsepower, and for engine speeds from 1000 to 2900 rpm. The propeller efficiencies were obtained from force measurements and the blade thrust load distribution was obtained by two diametrically opposed slipstream survey rakes shown in this paper.

Description: An investigation was conducted in the Cleveland altitude wind tunnel to determine the performance of a Curtiss propeller with four 838-lC2-lSRl blades on a YP-47M airplane at high blade loadings and engine powers. The study was made for a range of power coefficients between 0.30 and 1.00 at free-stream Mach numbers of 0.40 and 0.50. The results of the force measurements indicate primarily the trend of propeller efficiency for changes in power coefficient or advance-diameter ratio, inasmuch as corrections for the effects of tunnel-wall constriction on the installation have not been applied. Slip-stream pressure surveys across the propeller disk are presented to illustrate blade thrust load distribution for several operating conditions. At a free-stream Mach number of 0.40, nearly constant peak efficiencies were obtained at power coefficients from 0.30 to 0.70. A change in power coefficient from 0.70 to 0.90 reduced the peak efficiency about 5 percent. Blade stall at the tip sections became evident for a power coefficient of 0.91 when the advance-diameter ratio was reduced to 1.87. At a free-stream Mach number of 0.50, the highest propeller efficiencies were obtained for power coefficients from 0.80 to 1.00 at advance-diameter ratios above 2.90. At advance-diameter ratios below 2.90, the highest efficiencies were obtained for power coefficients of 0.60 and 0.70. The envelope of the efficiency curves decreased about 12 percent between advance-diameter ratios of 2.60 and 4.20. Local compressibility effects became evident for a power coefficient of 0.40 when the advance-diameter ratio was decreased to 1.75.

Description: Knock-limited performance data were obtained for three fuels on an R-1830-75 engine in a B-24D airplane at engine speeds of 1800, 2250, and 2600 rpm, a spark advance of 25 degrees B.T.C., and carburetor-air temperatures of 85 F for 1800 and 2250 rpm and 100 F for 2600 rpm. The test fuels were a blend of 80 percent 28-R plus 20 percent triptane (leaded to 4.5 ml TEL/gal), a blend of 80 percent 28-R plus 15 percent toluene (leaded to 4.5 ml TEL / gal), and 28-R fuel. The knock-limited manifold pressure of the toluene blend depreciated more in the lean region than the triptane blend or 28-R fuel. The knock-limited brake horsepower for the triptane blend varied from 16 to 25 percent higher than 28-R in the lean region and 18 to 30 percent higher in the rich region. The knock-limited brake horsepower of the toluene blend was approximately 15 percent higher than that of 28-R in the rich region and varied from 2 to 10 percent higher in the lean region. Knock limits of the triptane blend and 28-R fuel tested in the R-1830-75 engine agreed with limits for the same fuels determined with the R-1830-94 engine for engine speeds of 1800 and 2250 rpm.

Description: In the calculation of the dimensions of modern machines and building constructions, account is taken of the frequency of the occurrence of the anticipated loads. It is generally assumed that these loads will be repeated an infinite number, or at any rate some millions, of times during the total working life of the construction, When calculating the dimensions of the structural parts of aircraft, on the contrary, a consideration only of those frequencies in the appearance of the loads which actually come into play in the various states of stress is allowable. This is because in aircraft construction it is absolutely essential not only to ensure adequate structural strength but also to keep down the structural weight to the lowest possible limit, Strength tests in which this requirement is directly taken into account have recently been carried out by the DVL Material Strength Department.

Description: When flying in a turn or pulling out of a dive, the airscrew exerts a gyroscopic moment on the aircraft, In the case of airscrews with three or more blades, arranged symmetrically, the value of the gyroscopic moment is J(sub x) omega(sub x) omega(sub y), where J(sub x) denotes the axial moment of inertia about the axis of rotation of the airscrew, omega(sub x) the angular upeed of the airscrew about its axis, and omega (sub Y) the rotary speed of the whole aircraft about an axis parallel to the plane of the airscrew (e.g., when pulling up, the transverse axis of the aircraft). The gyroscopic moment then tends to rotate the aircraft about an axis perpendicular to those of the two angular speeds and, in the came of airscrews with three or more blades, is constant during a revolution of the airscrew. With two-bladed airscrews, on the contrary, although the calculate gyroscopic moment represents the mean value in time, it fluctuates about this value with a frequency equal to twice the revolutions per minute. In addition, pulsating moments likewise occur about the other two axes. This fact is known from the theory of the asymmetrical gyro; the calculations that have been carried out for the determination of the various gyroscopic moments, however, mostly require an exact knowledge of the gyro theory. The problem will therefore be approached in another manner based on quite elementary considerations. The considerations are of importance, not only in connection with the gyroscopic moments exerted by the two-bladed airscrew on the aircraft, but also with the stressing of the blades of airscrews with an arbitrary number of blades.

Description: The present investigation was intended to ascertain how far a tail unit is subject to disturbance by the jet of a propulsion unit. The parameters upon which this disturbing influence depends, and the values it reaches, had to be determined.