National Advisory Committee for Aeronautics (NACA) - 740 Matching Results

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High-speed aerodynamic characteristics of four thin NACA 63-series airfoils
No Description Available.
Preliminary Results of an Altitude-Wind-Tunnel Investigation of a TG-100A Gas Turbine-Propeller Engine, V, Combustion-Chamber Characteristics
An investigation to determine the performance and operational characteristics of the TG-1OOA gas turbine-propeller engine was conducted in the Cleveland altitude wind tunnel. As part of this investigation, the combustion-chamber performance was determined at pressure altitudes from 5000 to 35,000 feet, compressor-inlet rm-pressure ratios of 1.00 and 1.09, and engine speeds from 8000 to 13,000 rpm. Combustion-chamber performance is presented as a function of corrected engine speed and.correcte& horsepower. For the range of corrected engine speeds investigated, over-all total-pressure-loss ratio, cycle efficiency, ana the frac%ional loss in cycle efficiency resulting from pressure losses in the combustion chambers were unaffected by a change in altitude or compressor-inlet ram-pressure ratio. The scatter of combustion- efficiency data tended to obscure any effect of altitude or ram-pressure ratio. For the range of corrected horse-powers investigated, the total-pressure-loss ratio an& the fractional loss in cycle efficiency resulting from pressure losses in the combustion chambers decreased with an increase in corrected horsepower at a constant corrected engine speed. The combustion efficiency remained constant for the range of corrected horse-powers investigated at all corrected engine speeds.
Supersonic-tunnel tests of two supersonic airplane model configurations
Report presenting supersonic-tunnel tests of two models of similar supersonic airplane configurations at Mach numbers of 1.55, 1.90, and 2.32 to determine values of the drag, lift, pitching moment, yawing moment, and side force. The models were similar except for the vertical wing location relative to the body axis and horizontal tail; one had a high wing and one had a low wing. Results regarding the precision of data, Reynolds numbers of tests, results at the different Mach numbers, and Schileren photographs are provided.
Data Obtained in the Flight Measurements to Determine the Stability and Control Characteristics of a C-54D Airplane (AAF No. 42-72713) and a Summary of the Test Program
The flight investigation of the C-54D airplane was initiated to determine the necessity of changes or additions to existing handling-qualities requirements to cove the case of instrument approaches with large airplanes. This paper gives a brief synopsis of the results and presents the measured data of tests to determine the stability and control characteristics. It was found that no new requirements were necessary to cover the problems of instrument approaches. The C-54D airplane tested met the Amy and Navy stability and control requirements except for the following items. The control-system friction with autopilot installed vas double that allowed by the requirements. The amount of friction was found to impair the controllability of the airplane in precision flying. The lateral and directional characteristics were good except that the maximum pb/2V was slightly below the minimum required, and the elevator-control forces to obtain the maximum pb/2V at low speeds were above the Army and Navy requirements. The longitudinal stability and control characteristics were good except that the elevator-control forces exceeded the limits of the Army and Navy requirements in turns and in landings. The stalling characteristics were considered good in all conditions with the stall warning in the form of tail buffeting occurring at speeds approximately 5 miles per hour above the stall.
Preliminary Results of Altitude-Wind-Tunnel Investigation of X24C-4B Turbojet Engine. I - Pressure and Temperature Distributions
Pressures and temperatures throughout the X24C-4B turbojet engine are presented in both tabular and graphical forms to show the effect of altitude, flight Mach number, and engine speed on the internal operation of the engine. These data were obtained in the NACA Cleveland altitude wind tunnel at simulated altitudes from 5000 to 45,000 feet, simulated flight Mach numbers from 0.25 to 1.08, and engine speeds from 4000 to 12,500 rpm. Location and detail drawings of the instrumentation installed at seven survey stations in the engine are shown. Application of generalization factors to pressures and temperatures at each measuring station for the range of altitudes investigated showed that the data did not generalize above an altitude of 25,000 feet. Total-pressure distribution at the compressor outlet varied only with change in engine speed. At altitudes above 35,000 feet and engine speeds above 11,000 rpm, the peak temperature at the turbine-outlet annulus moved inward toward the root of the blade, which is undesirable from blade-stress considerations. The temperature levels at the turbine outlet and the exhaust-nozzle outlet were lowered as the Mach number was increased. The static-pressure measurements obtained at each stator stage of the compressor showed a pressure drop through the inlet guide vanes and the first-stage rotor at high engine speeds. The average values measured by the manufacturer's instrumentation werein close agreement with the average values obtained with NACA instrumentation.
Preliminary Results of Altitude-Wind-Tunnel Investigation of X24C-4B Turbojet Engine. IV - Performance of Modified Compressor, Part 4, Performance of Modified Compressor
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.
Tests of the AN/SPS-1 Radar Antenna in the Langley Full-Scale Tunnel
Tests have been conducted to determine the drive-motor torque and the static force and moment characteristics of the AN/SPS-1 radar antenna. Shifting the longitudinal position of the antenna had very little effect on the drive-motor torque, which reached a maximum value expressed in terms of dynamic pressure (T/q)(sub max) of 1.15 at an azimuth angle of 245. The maximum observed values of rolling, pitching, and yawing moments in terms of dynamic pressure are -29.0, 66.6, and 13.4, respectively.
High-Speed Wind-Tunnel Investigation of the Lateral Control Characteristics of Plain Ailerons on a Wing with Various Amounts of Sweep
A three-dimensional investigation of straight-sided-profile plain ailerons on a wing with 30 degrees and 45 degrees of sweepback and sweepforward was made in a high-speed wind tunnel for aileron deflections from -10 degrees to 10 degrees and at Mach numbers from 0.60 to 0.96. Wing configurations of 30 degrees generally reduced the severity of the large changes in rolling-moment and aileron hinge-moment coefficients experienced by the upswept wing configurations as the result of compression shock and extended to higher Mach numbers the speeds at which such changes occurred.
Investigation of Pressure Distribution over an Extended Leading-Edge Flap on a 42 Degrees Sweptback Wing
Pressure distribution over an extended leading-edge flap on a 42 degree swept-back wing was investigated. Results indicate that the flap normal-force coefficient increased almost linearly with the angle of attack to a maximum value of 3.25. The maximum section normal-force coefficient was located about 30 percent of the flap span outboard of the inboard end and had a value of 3.75. Peak negative pressures built up at the flap leading edge as the angle of attack was increased and caused the chordwise location of the flap center of pressure to be move forward.
Investigation of wing characteristics at a Mach number of 1.53 I : triangular wings of aspect ratio 2
No Description Available.
Preliminary results of a flight investigation to determine the effect of negative flap deflection on high-speed longitudinal-control characteristics
Report presenting flight testing on two propeller-driven airplanes with wings of NACA 66-series and NACA 230-series airfoil sections to determine the effect of deflecting the landing flaps upward on the high-speed longitudinal-control characteristics.
A suggested method of analyzing for transonic flutter of control surfaces based on available experimental evidence
No Description Available.
Experimental Investigation of a Preloaded Spring-Tab Flutter Model
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.
Wind-Tunnel Investigation of the Low-Speed Characteristics of a 1/8-Scale Model of the Republic XP-91 Airplane with a Vee and a Conventional Tail
Low-speed wind-tunnel tests of a l/8 scale model of the Republic XP-91 airplane were made to determine its low-speed characteristics and the relative merits of a vee and a conventional tail on the model. The results of the tests showed that for the same amount of longitudinal and directional stability the conventional tail gave less roll due to sideslip than did the vee tail. The directional stability of the model was considered inadequate for both the vee and conventional tails; however, increasing the area and aspect ratio of the conventional vertical tail provided adequate directional stability. It was possible with negative wing dihedral and open main landing gear doors to reduce the excessive roll due to sideslip for the landing configuration (flaps and gear down) to a more reasonable value commensurate with the aileron power. The use of variable wing incidence to adjust the longitudinal balance was sufficiently effective to reduce the predicted up-elevator required for landing by approximately 5 deg.
Damping in pitch and roll of triangular wings at supersonic speeds
A method is derived for calculating the damping coefficients in pitch and roll for a series of triangular wings and a restricted series of sweptback wings at supersonic speeds. The elementary "supersonic source" solution of the linearized equation of motion is used to find the potential function of a line of doublets, and the flows are obtained by surface distributions of these doublet lines. The damping derivatives for triangular wings are found to be a function of the ratio of the tangent of the apex angle to the tangent of the Mach angle. As this ratio becomes equal to and greater than 1.0 for triangular wings, the damping derivatives, in pitch and in roll, become constant. The damping derivative in roll becomes equal to one-half the value calculated for an infinite rectangular wing, and the damping derivative in pitch for pitching about the apex becomes equal to 3.375 times that of an infinite rectangular wing.
Preliminary Results of Altitude-Wind-Tunnel Investigation of X34C-4B Turbojet Engine. III - Compressor Performance, 3, Compressor Performance
The performance of the 11-stage axial-flow compressor in the X24C-4B turbojet engine was analyzed on the basis of results obtained from an investigation of the complete engine in the NACA Cleveland altitude wind tunnel. The engine was operated with four, exhaust nozzles of different outlet area over a range of engine speeds from 6000 to 12,500 rpm, corrected engine speeds from approximately 6100 to 13,600 rpm, and compressor Mach numbers from 0.45 to 1.00. Data are presented for engine operation over a range of simulated altitudes from 15,000 to 45,000 feet and simulated flight Mach numbers from 0.24 to 1.08.
Calculation of Wing Bending Moments and Tail Loads Resulting from the Jettison of Wing Tips During a Symmetrical Pull-Up
A preliminary analytical investigation was made to determine the feasibility of the basic idea of controlled failure points as safety valves for the primary airplane structure. The present analysis considers the possibilities of the breakable wing tip which, in failing as a weak link, would relieve the bending moments on the wing structure. The analysis was carried out by computing the time histories of the wing and stabilizer angle of attack in a 10g pull-up for an XF8F airplane with tips fixed and comparing the results with those for the same maneuver, that is, elevator motion but with tips jettisoned at 8g. The calculations indicate that the increased stability accompanying the loss of the wing tips reduces the bending moment an additional amount above that which would be expected from the initial loss in lift and the inboard shift in load. The vortex shed when the tips are lost may induce a transient load requiring that the tail be made stronger than otherwise.
A comparison with flight data of vertical-tail loads in various maneuvers estimated from sideslip angles and rudder deflections
Report presenting a comparison of the vertical-tail loads determined from pressure-distribution measurements in flight in various maneuvers with the corresponding vertical-tail loads. Some of the maneuvers investigated included slow rolls, steady sideslips, fishtails, and rolling pull-outs.
Preliminary tests to determine the maximum lift of wings at supersonic speeds
Report presenting the results of a test program in a supersonic tunnel to determine the maximum lift of wings operating at supersonic speeds. A variety of wing plan forms of several thickness distributions were tested at a range of Mach numbers, Reynolds numbers, and angles of attack. The lift results, drag results, and Schileren photographs are described.
Tests of submerged duct installation on a modified airplane in the Ames 40- by 80-foot wind tunnel
No Description Available.
Effect of Air-Flow Distribution and Total-Pressure Loss on Performance of One-Sixth Segment of Turbojet Combuster
An investigation has been conducted on a one-sixth segment of an annular turbojet combustor to determine the effects of modification in air-flow distribution and total-pressure loss on the performance of the segment. The performance features investigated during this series of determinations were the altitude operational limits and the temperature-rise efficiency. Altitude operational limits of the combustor segment, for the 19XB engine using the original combustor-basket design were approximately 38,000 feet at 17,000 rpm and 26,000 feet at 10,000 rpm. The altitude operational limits were approximately 50,000 feet at 17,000 rpm and 38,000 feet at 10,000 rpm for a combustor-basket design in which the air-passage area in the basket was redistributed so as to admit gradually no more than 20 percent of the air along the first half of the basket. In this case the total pressure loss through the combustor segment was not appreciably changed from the total-pressure loss for the original combustor basket design. Altitude operational limits of the combustor segment for the 19XB engine were above 52,000 feet at 17,000 rpm and were approximately 23,000 feet at 10,000 rpm for a combustor-basket design in which the distribution of the air-passage area in the basket was that of the original design but where the total-pressure loss was increased to 19 times the inlet reference kinetic pressure at an inlet-to-outlet density ratio of 2.4. The total-pressure loss for the original design was 14 times the inlet kinetic reference pressure at an inlet-to-outlet density ratio of 2.4.
An investigation at low speed of a large-scale triangular wing of aspect ratio two II: the effect of airfoil section modifications and the determination of the wake downwash
Report presenting a study of the characteristics of a large-scale triangular wing to include the effects of section modifications. The wing in this report is the same as the one in the previous report but features various degrees of rounding of the wing leading edge and wing maximum thickness rather than having sharp edges. Results regarding the effects of airfoil section modifications, visible trailing vortices, and surveys in the extended chord plane are provided.
Yaw characteristics and sidewash angles of a 42 degree sweptback circular-arc wing with a fuselage and with leading-edge and split flaps at a Reynolds number of 5,300,000
Report presenting testing of the low-speed aerodynamic characteristics in yaw of a 42 degree sweptback wing of circular-arc airfoil sections in the pressure tunnel. The wing had an aspect ratio of 3.94, taper ratio of 0.625, and no dihedral or twist. Results regarding lateral-stability parameters of plain wing, effect of wing flaps on lateral-stability parameters, effect of fuselage on lateral-stability parameters, a comparison with the NACA 64(sub 1)-112 wing, characteristics in the extended yaw range, and airflow characteristics in the region of a vertical tail are provided.
Tests of the landing on water of a model of a high-speed airplane - Langley tank model 229
No Description Available.
Tests of the landing on water of a model of a high-speed airplane - langley tank model 229
No Description Available.
Comparison of the transonic drag characteristics of two wing-body combinations differing only in the location of the 45 degree sweptback wing
Report presenting the drag of a series of wing-body combinations by the free-fall method in order to provide information on the drag characteristics of promising transonic and supersonic airplane arrangements. Time histories, Mach number variations, and drag coefficients for several areas of the body are provided.
Evaluation of Gust and Draft Velocities from Flights of P-61C Airplanes within Thunderstorms July 12, 1947 to July 18, 1947 at Clinton County Army Air Field, Ohio
The gust and draft velocities from records of NACA instruments installed in P-61c airplanes participating in thunderstorm flights at Clinton County Army Air Field, Ohio, from July 12, to July 18, 1947 are presented.
Investigation of Single-Stage Modified Turbine of Mark 25 Torpedo Power Plant
Efficiency investigations have been made on a single-stage modification of the turbine of a Mark 25 aerial torpedo to determine the performance of the unit with five different turbine nozzles. The output of the turbine blades was computed by analyzing the windage and mechanical-friction losses of the unit. The turbine was faund to be most efficient with a cast nozzle having sharp-edged inlets to the nine nozzle ports. An analysis af the effectiveness af the first and second stages of the standard Mark 25 torpedo turbine indicates that the first- stage turbine contributes nearly all the brake power produced at blade-jet speed ratios above 0.26.
An experimental investigation of NACA submerged air inlets on a 1/5-scale model of a fighter airplane
No Description Available.
Flight Measurements of the Flying Qualities of a Lockheed P-80A Airplane (Army No. 44-85099) - Stalling Characteristics
This report contains the flight-test results of the stalling characteristics measured during the flying-qualities investigation of the Lockheed P-8OA airplane (Army No. 44-85099). The tests were conducted in straight and turning flight with and without wing-tip tanks. These tests showed satisfactory stalling characteristics and adequate stall warning for all configurations and conditions tested.
Analysis, Verification, and Application of Equations and Procedures for Design of Exhaust-Pipe Shrouds
Investigations were made to develop a simplified method for designing exhaust-pipe shrouds to provide desired or maximum cooling of exhaust installations. Analysis of heat exchange and pressure drop of an adequate exhaust-pipe shroud system requires equations for predicting design temperatures and pressure drop on cooling air side of system. Present experiments derive such equations for usual straight annular exhaust-pipe shroud systems for both parallel flow and counter flow. Equations and methods presented are believed to be applicable under certain conditions to the design of shrouds for tail pipes of jet engines.
The calculation of the heat required for wing thermal ice prevention in specified icing conditions
No Description Available.
The calculation of the heat required for wing thermal ice prevention in specified icing conditions
Flight tests were made in natural icing conditions with two 8-ft-chord heated airfoils of different sections. Measurements of meteorological variables conducive to ice formation were made simultaneously with the procurement of airfoil thermal data. The extent of knowledge on the meteorology of icing, the impingement of water drops on airfoil surfaces, and the processes of heat transfer and evaporation from a wetted airfoil surface have been increased to a point where the design of heated wings on a fundamental, wet-air basis now can be undertaken with reasonable certainty.
Determination of plate compressive strengths
From Summary: "Results of local-instability tests of H-, Z-, and C-section plate assemblies of four extruded aluminum alloys and two magnesium alloys, obtained in an extensive investigation to determine plate compressive strengths of aircraft structural materials, are summarized. On the basis of the general relationships found between the plate compressive strengths and the compressive stress-strain curves, methods applicable to flat plates and based upon the use of the compressive stress-strain curve are suggested for determining the critical compressive stress and the average stress at maximum load."
Effect of changing manifold pressure, exhaust pressure, and valve timing on the air capacity and output of a four-stroke engine operated with inlet valves of various diameters and lifts
Report presenting a series of tests made with a CFR engine to determine the effect of inlet-valve capacity, inlet and exhaust pressure, and valve timing on the volumetric efficiency at various speeds. Three combinations of inlet and exhaust pressures and seven valve-timing arrangements were used.
The Elasto-Plastic Stability of Plates
This article explains results developed from the following research: 'The Stability of Plates and Shells beyond the Elastic Limit.' A significant improvement is found in the derivation of the relations between the stress factors and the strains resulting from the instability of plates and shells. In a strict analysis, the problem reduces to the solution of two simultaneous nonlinear partial differential equations of the fourth order in the deflection and stress function, and in the approximate analysis to a single linear equation of the Bryan type. Solutions are given for the special cases of a rectangular plate buckling into a cylindrical form, and of an arbitrarily shaped plate under uniform compression. These solutions indicate that the accuracy obtained by the approximate method is satisfactory.
Flight Tests of Rocket-Powered "Tin-Can" Models of AAF Project MX-800
Flight tests were made of six noninstrumented rocket-powered "Tin Can" models of AAF Project MX-800. Velocity and drag data were obtained by use of CU Doppler radar. The existence of stability and adequate structural strength for flight near zero lift was checked by visual and photographic observation. Drag data obtained during the tests agreed reasonably well with estimates based on experimental data from NACA RM-2 rocket-powered drag research models.
Fundamental Aerodynamic Investigations for Development of Arrow-Stabilized Projectiles
The numerous patent applications on arrow-stabilized projectiles indicate that the idea of projectiles without spin is not new, but has appeared in various proposals throughout the last decades. As far as projectiles for subsonic speeds are concerned, suitable shapes have been developed for sometime, for example, numerous grenades. Most of the patent applications, though, are not practicable particularly for projectiles with supersonic speed. This is because the inventor usually does not have any knowledge of aerodynamic flow around the projectile nor any particular understanding of the practical solution. The lack of wind tunnels for the development of projectiles made it necessary to use firing tests for development. These are obviously extremely tedious or expensive and lead almost always to failures. The often expressed opinion that arrow-stabilized projectiles cannot fly supersonically can be traced to this condition. That this is not the case has been shown for the first time by Roechling on long projectiles with foldable fins. Since no aerodynamic investigations were made for the development of these projectiles, only tedious series of firing tests with systematic variation of the fins could lead to satisfactory results. These particular projectiles though have a disadvantage which lies in the nature cf foldable fins. They occasionally do not open uniformly in flight, thus causing unsymmetry in flow and greater scatter. The junctions of fins and body are very bad aerodynamically and increase the drag. It must be possible to develop high-performance arrow-stabilized projectiles based on the aerodynamic research conducted during the last few years at Peenemuende and new construction ideas. Thus the final shape, ready for operational use, could be developed in the wind tunnel without loss of expensive time in firing tests. The principle of arrow-stabilized performance has been applied to a large number of caliburs which were stabilized by various means ...
Further investigation of preheating and postheating in spot-welding 0.040-inch Alclad 24S-T
No Description Available.
Instability of outstanding flanges simply supported at one edge and reinforced by bulbs at other edge
No Description Available.
Introduction to the problem of rocket-powered aircraft performance
An introduction to the problem of determining the fundamental limitations on the performance possibilities of rocket-powered aircraft is presented. Previous material on the subject is reviewed and given in condensed form along with supplementary analyses. Some of the problems discussed are: 1) limiting velocity of a rocket projectile; 2) limiting velocity of a rocket jet; 3) jet efficiency; 4) nozzle characteristics; 5) maximum attainable altitudes; 6) ranges. Formulas are presented relating the performance of a rocket-powered aircraft to basic weight and nozzle dimensional parameters. The use of these formulas is illustrated by their application to the special case of a nonlifting rocket projectile.
Investigation of the fuselage interference on a Pitot-static tube extending forward from the nose of the fuselage
No Description Available.
A method for estimating heat requirements for ice prevention gas heated hollow propeller blades
Report presenting a method for determining the temperature and flow of heated gas necessary for ice prevention of hollow propeller blades in flight and icing conditions. A variety of conditions are taken into consideration and suggested simplifications and short methods are provided in order to not overcomplicate the design modifications.
A method for estimating heat requirements for ice prevention on gas-heated hollow propeller blades
Report presenting a detailed method for determining the temperature and flow of heated gas necessary for ice prevention of hollow propeller blades in flight and icing conditions. Expressions for the total external and internal heat transfer are combined to determine the surface temperatures of each segment.
Notes and tables for use in the analysis of supersonic flow
No Description Available.
Theoretical additional span loading characteristics of wings with arbitrary sweep, aspect ratio, and taper ratio
The Weissinger method for determining additional span loading has been used to find the lift-curve slope, spanwise center of pressure, aerodynamic center location, and span loading coefficients of untwisted and uncambered wings having a wide range of plan forms characterized by various combinations of sweep, aspect ratio, and taper ratio. The results are presented as variations of the aerodynamic characteristics with sweep angle for various values of aspect ratio and taper ratio. Methods are also included for determining induced drag and the approximate effects of compressibility. Despite the limitations of a lifting line method such as Weissinger's, the good agreement found between experimentally and theoretically determined characteristics warrants confidence in the method. In particular, it is believed that trends observed in results of the Weissinger method should be reliable. One of the most significant results showed that for each angle of sweep there is a taper ratio for which aspect ratio has little effect on the span loading and for which the loading is practically elliptical. This elliptic loading is approached at a taper ratio of 1.39 for 30 degree of sweepforward, 0.45 for zero degree of sweepback. (author).
Free-Spinning-Tunnel Tests of a 1/24-Scale Model of the McDonnell XP-88 Airplane with a VEE Tail
An investigation of the spin and recovery characteristics of a 1/24-scale model of the McDonnell XP-88 airplane has been conducted in the Langley 20-ft free-spinning tunnel. Results of tests with a conventional tail have been previously reported; the results presented herein are for the model with a vee tail installed. The effects of control settings and movements on the erect and inverted spin and recovery characteristics of the model. In the normal loading were determined. Tests of the model in the long-range loading also were made. The investigation included leading-edge-flap, spin-recovery-parachute, and rudder-pedal-force tests. The recovery characteristics of the model were satisfactory for the normal loading. Deflecting the leading-edge flaps improved recoveries. The results indicated that with the external wing tanks installed (long-range loading) recoveries may be poor and, therefore, if a spin is inadvertently entered in this condition the tanks should be jettisoned if recovery does not appear imminent immediately after it is attempted. A 10-foot spin-recovery tail parachute with a towline 40 feet long and a drag coefficient of 0.63 was found to be effective for spin recovery. The rudder pedal force required for spin recovery was indicated to be within the capabilities of the pilot.
Analysis of effect of basic design variables on subsonic axial-flow-compressor performance
From Summary: "A blade-element theory for axial-flow compressors has been developed and applied to the analysis of the effects of basic design variables such as Mach number, blade loading, and velocity distribution on compressor performance. The relations among several efficiencies useful in compressor design are derived and discussed. The possible gains in useful operating range obtainable by the use of adjustable stator blades are discussed and a rapid approximate method of calculating blade-angle resettings is shown by an example."
Comparison between flight-measured and calculated span load distribution at high Mach numbers
No Description Available.
Preliminary tank tests of NACA hydro-skis for high-speed airplanes
Report presenting the results from tank landing and take-off tests with a dynamic model of a hypothetical jet-propelled airplane equipped with NACA hydro-skis. Results regarding landing tests, take-off tests, and practical considerations for the creation of hydro-ski configurations are provided. The hydro-skis suitable for flush retraction into streamline fuselages do appear to offer a practical means of water take-offs and landings.