PDF/ocr_dataset_1056.jsonl

{"citation_id": "19930085548", "source_url": "https://ntrs.nasa.gov/api/citations/19930085548/downloads/19930085548.pdf", "page_number": 2, "total_pages": 46, "image_filename": "19930085548_p2.jpg", "text": "NACA RM No. E8L30 RESTRICTED\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nRESEARCH MEMORANDUM\n\nEXPERIMENTAL STUDY OF LOOP-SCAVENGED COMPRESSION-IGNITION CYLINDER FOR GAS-GENERATOR USE\n\nBy Hampton H. Foster, F. Ralph Schuricht and Max J. Tauschek\n\nSUMMARY\n\nA preliminary experimental investigation was made of the performance and general operating characteristics of a small ($3\\frac{1}{4}$ by $4\\frac{1}{2}$ in.) single-cylinder, two-stroke-cycle, loop-scavenged engine using compression ignition at low compression ratios, high inlet-manifold temperatures, and high inlet-manifold and exhaust-gas pressures. The investigation was conducted to determine experimentally the performance characteristics of a ported cylinder for gas-generator use, to compare the results with those obtained by an analysis of a piston-type gas-generator engine, and to indicate the practicability of operating an engine cylinder at the required conditions.\n\nThe experimental results, in general, are in reasonable agreement with the performance values analytically obtained for the piston-type burner. Scavenging was unsatisfactory at rich fuel-air mixtures; consequently, the charging efficiency and the power output were somewhat lower than anticipated. The thermal efficiency experimentally determined checked well with analytical results at low fuel-air ratios. Heat losses from the cylinder were inordinately high; these high losses were partly attributed to the high surface-volume ratio of the cylinder and to the low coolant temperature used to expedite the recording of data in this initial investigation. When the heat-rejection rate was considered, the calculated and the measured exhaust-gas temperatures agreed very closely.\n\nOperation of the cylinder at low compression ratios, high inlet-manifold temperatures, high inlet-manifold and exhaust-gas pressures, and high maximum cylinder pressures presented no new problems nor difficulties. Operation was quite smooth because of the low rate of pressure rise in the cylinder.\n\nRESTRICTED", "timestamp": "2026-07-22T06:06:12.902787+00:00"}
{"citation_id": "19930093769", "source_url": "https://ntrs.nasa.gov/api/citations/19930093769/downloads/19930093769.pdf", "page_number": 27, "total_pages": 39, "image_filename": "19930093769_p27.jpg", "text": "26\nCONFIDENTIAL\nNACA RM No. E8L10a\n\nFuel\nAN-F-58\nGasoline\n\nCorrected specific fuel consumption based on net thrust, lb/(hr)(lb thrust)\n\n1.5\n.5\n(a) Altitude, 5000 feet; flight Mach number, 0.\n\n3.5\n2.5\n1.5\n(b) Altitude, 20,000 feet; flight Mach number, 0.60.\n\n4.5\n3.5\n2.5\n1.5\n9,000 10,000 11,000 12,000 13,000 14,000\nCorrected engine speed, rpm\n(c) Altitude, 20,000 feet; flight Mach number, 0.85.\n\nFigure 6. - Comparison of corrected specific fuel consumption based on net thrust for AN-F-58 fuel and gasoline at various altitudes and flight Mach numbers.\n\nCONFIDENTIAL\n1070", "timestamp": "2026-07-22T06:06:17.395222+00:00"}
{"citation_id": "19930085485", "source_url": "https://ntrs.nasa.gov/api/citations/19930085485/downloads/19930085485.pdf", "page_number": 5, "total_pages": 26, "image_filename": "19930085485_p5.jpg", "text": "NACA RM No. L8K02 CONFIDENTIAL 3\n\n$\\rho$ mass density of air, slugs per cubic foot\n\nV air velocity, feet per second\n\nM average Mach number over span of model\n\n$M_T$ tunnel reference Mach number\n\n$M_l$ local Mach number\n\nR Reynolds number\n\nThe rolling-moment data have been corrected in accordance with the method of reference 2 for reflection-plane models. This correction is for extremely low Mach numbers. No correction has been made for Mach number effect. The correction applied was as follows:\n\n$$C_{l_a} = 0.897 C_{l_{\\text{measured}}}$$\n\nAll data are presented about the wind axes.\n\nMODEL\n\nThe semispan wing model for these tests had a leading-edge sweepback of $42.7^\\circ$, a taper ratio of 0.50, and an aspect ratio of 4.0; other geometric characteristics are shown in figure 1. The wing, made of steel with a polished surface, had a 10-percent-thick circular-arc section normal to the 50-percent-chord line, had no dihedral, and was mounted as a midwing (fig. 1). The polished-brass fuselage was semi-circular in cross section and was bent to the contour of the bump. The fuselage for the tests on the side wall of the tunnel was made of hardwood.\n\nThe 50-percent-span outboard aileron was attached to the wing with a $\\frac{1}{32}$-inch-thick copper insert (fig. 1). This insert was bent to obtain the required aileron deflection. The deflection was checked before and after each test. The various aileron profiles investigated are shown in figure 2. The aileron chord was 20 percent of the wing chord.\n\nTEST TECHNIQUE\n\nThe tests were performed in the Langley high-speed 7- by 10-foot tunnel which is capable of reaching the choking Mach number. In order to obtain transonic speeds in the tunnel, an application of the NACA wing-flow method of testing was made (reference 3). This method of testing at transonic speeds involves placing the model in the high-velocity\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:06:21.186946+00:00"}
{"citation_id": "19930082703", "source_url": "https://ntrs.nasa.gov/api/citations/19930082703/downloads/19930082703.pdf", "page_number": 7, "total_pages": 28, "image_filename": "19930082703_p7.jpg", "text": "NACA TN 1983\n\npeak acceleration was reached. Much less control deflection was needed for recovery from this pull-up than was needed for helicopter A.\n\nPilot's comments.- The pilot's comments on this maneuver in helicopter B, as compared with the corresponding one in helicopter A, was that as a result of removal of the divergent tendency, the feeling of apprehension was greatly reduced. Although this removal of the divergent tendency was believed to be more important than any further improvements could be, the pull-up characteristics were still considered to be by no means satisfactory because of the difficulty in anticipating, during the early phase of the maneuver, the acceleration (and change in flight path and attitude angle) that would be reached later.\n\nStick-force gradient.- Consideration of experience with airplane flying qualities suggested that the introduction of a stick-force gradient might provide the pilot with a means for anticipating the final results by providing a continuous indication of the magnitude of the control deflection from trim. Three different values of force gradient were accordingly tried, by use of suitable springs attached to the control stick. The largest gradient (8 lb/in.) only aggravated the pilot's impressions. The smallest gradient (2 lb/in.) had no noticeable effect; control friction, although approximately overcome by control vibration, may have been responsible for this result. The intermediate value ($\\frac{1}{2}$ lb/in.) was reported by the pilot to produce definite improvement but still to leave much to be desired. This intermediate value was sufficient to return the control promptly to trim when the stick was deflected and released, in spite of the friction present.\n\nStick-fixed oscillations.- With helicopter B, longitudinal disturbances in level flight and moderate climbs at 80 miles per hour, stick fixed, were slowly damped out. Lateral oscillations were noticeable during these trials. In some cases, these lateral motions were checked by use of lateral control. For the climbs, longitudinal stick motions were easily avoided during this process because the stick trimmed against the forward stop.\n\nThe fact that the pilot was willing to fly helicopter B with the stick against the forward stop gives a further indication of the difference between helicopters B and A. With helicopter A, a sizeable margin of control had to be maintained in order that the divergent tendencies could be successfully checked. This margin was needed in normal flight as well as in maneuvers.", "timestamp": "2026-07-22T06:06:21.839821+00:00"}
{"citation_id": "19930086076", "source_url": "https://ntrs.nasa.gov/api/citations/19930086076/downloads/19930086076.pdf", "page_number": 45, "total_pages": 50, "image_filename": "19930086076_p45.jpg", "text": "NACA RM E9F09\n\n[Figure: Flame holder L3 showing installation of molybdenum plates before operation.]\n\nFigure 17. - Flame holder L3 showing installation of molybdenum plates before operation.\n\nNACA\nC-22903\n2-5-49\n\n43", "timestamp": "2026-07-22T06:06:22.037713+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 26, "total_pages": 78, "image_filename": "19930082483_p26.jpg", "text": "24\nNACA TN No. 1807\n\ncalculated by use of equation (4). The driving-fluid losses for other degrees of admission are therefore established by equation (28), which becomes\n\n$$(driving-fluid\\ loss)_F = \\frac{3}{2} (1-F) (driving-fluid\\ loss)_{120^\\circ} \\quad (41)$$\n\nThis relation has been verified at 180° admission.\n\nPresentation Methods\n\nPerformance presentation. - Turbine performance for full and partial admission is presented in the form of a \"carpet\" plot (fig. 8). The procedure is an adaption of the methods of presentation of experimental results outlined in reference 7. Absence of an abscissa in the figure is to be noted. This composite or carpet plot is obtained by first plotting the corrected power output as the ordinate against total-pressure ratio as the abscissa with rotor speed as a parameter. (See diagram (a).)\n\n[Figure: (a) Graph showing Turbine power output corrected to sea level vs. Total-pressure ratio, $P_1'/P_e'$, with curves for $N/\\sqrt{\\theta_1}$]\n\n[Figure: (b) Graph showing Turbine power output corrected to sea level vs. $P_1'/P_e'$, with curves for $N/\\sqrt{\\theta_1}$ and a schematic below showing overlapping ranges of $P_1'/P_e'$]", "timestamp": "2026-07-22T06:06:23.080131+00:00"}
{"citation_id": "19930086097", "source_url": "https://ntrs.nasa.gov/api/citations/19930086097/downloads/19930086097.pdf", "page_number": 33, "total_pages": 36, "image_filename": "19930086097_p33.jpg", "text": "NACA RM A9H11 CONFIDENTIAL 31\n\n[Figure: Drag polars for wings 5, 6, and 7; $M_\\infty=1.5$, $Re = 0.5 \\times 10^6$.]\n\nFigure 13.—Drag polars for wings 5, 6, and 7; $M_\\infty=1.5$, $Re = 0.5 \\times 10^6$.\n\nCONFIDENTIAL\n\nNACA-Langley - 11-1-49 - 400", "timestamp": "2026-07-22T06:06:25.719046+00:00"}
{"citation_id": "19930082546", "source_url": "https://ntrs.nasa.gov/api/citations/19930082546/downloads/19930082546.pdf", "page_number": 16, "total_pages": 65, "image_filename": "19930082546_p16.jpg", "text": "NACA TN No. 1870\n\nFor conditions of figure 14 a variation in equation (2) of $R_{\\theta}$ resulted in a nearly uniform change in pressure amplitude for the fundamental frequency of a two-blade propeller throughout the given tip Mach number range. Figure 15 shows the amount of this variation for three values of $R_{\\theta}$ at $\\frac{x}{D} = -\\frac{1}{8}$. For these conditions calculations using $R_{\\theta} = 0.7R$ most nearly duplicated the experimental results.\n\nThus it may be seen that the maximum pressures which usually occur at $\\frac{x}{D} = -\\frac{1}{8}$ may be predicted by using an effective radius varying from $0.7R$ to $0.8R$ for the propeller in these tests. This propeller is believed to be representative of high-speed propellers. Since propellers are normally operated through a wide range of loading conditions, a value of $R_{\\theta}$ which will be valid for the extreme case is considered most useful. For this particular propeller $R_{\\theta} = 0.8R$ is recommended to give conservative calculated pressures.\n\nFigure 7 shows that the ratio of pressure coefficient to power coefficient is lower for the lightly loaded and the stalled propeller than for the heavily loaded propeller. Thus, since the value of $R_{\\theta} = 0.8R$ will adequately predict the pressures for a heavily loaded propeller, it will tend to overestimate the pressures at other operating conditions.\n\nDeming in reference 2 shows that for a propeller at a given blade angle the sound pressures at a distance vary approximately as the powers of the tip speed of 5, 6.5, and 8 for mB values of 2, 4, and 6, respectively. Since the power varies approximately as the cube of the tip speed, the sound pressure at constant power may be seen to vary as the powers of the tip speed of 2, 3.5, and 5 for mB = 2, 4, and 6, respectively. At a distance then, an increase in tip speed at constant power results in an increase of sound pressure for all harmonics. This does not apply for all harmonics, however, in the region near the propeller. Figure 14(a) shows that for a given blade angle the pressures varied considerably less with tip speed than was observed in reference 2. In figure 16 the experimental data of figure 14(a) is replotted to show the effect of tip Mach number at constant power on the free-space pressures of each harmonic. For these conditions the pressure per unit power is decreased as the tip Mach number is increased for mB = 2 whereas for mB = 6 the trend seems to reverse. The pressure amplitude of mB = 4 seems to be essentially independent of tip Mach number.\n\nCalculations in the plane of rotation for the pressure amplitude of the fundamental of a two-blade propeller have been made by means of Gutin's simplified equation and also by equation (2) of the present paper. The results obtained by using the two methods are plotted as a ratio against d/D in figure 17 for tip Mach numbers of 0.75 and 1.00. The Gutin equation is seen to underestimate the pressures at low d/D values.", "timestamp": "2026-07-22T06:06:27.347186+00:00"}
{"citation_id": "19930082613", "source_url": "https://ntrs.nasa.gov/api/citations/19930082613/downloads/19930082613.pdf", "page_number": 11, "total_pages": 46, "image_filename": "19930082613_p11.jpg", "text": "10\nNACA TN 1938\n\nHeat treatment and mechanical finishing. - The results of the thermal treatments of type-B liners indicated that the heat treatments selected were ineffective in preventing cracking. A slight improvement was noted in some of the reamed, sanded, and vapor-blasted liners that had been heat-treated. The mechanical finishing of seven as-fabricated type-B liners, however, materially reduced cracking in the accelerated-life determinations as shown in table III.\n\nDISCUSSION OF RESULTS\n\nThe location of buckling in the zone in which temperatures and thermal gradients were apparently greatest indicate that thermal stresses of a large magnitude were induced in the metal. These stresses were a result of the following conditions:\n\n1. Over-all temperature differentials produced in a liner during combustion. The hottest zone is between the first and the third row of louvers from the intake end of the liners.\n\n2. Temperature gradients produced at individual louvers by secondary combustion air, which enters the louvers and cools the metal immediately downstream. An abrupt line of demarcation between hot and cold areas occurs at or near the stress-relieving holes of the louvers.\n\nBuckling occurs between the stress-relieving holes of the louvers and the air-intake holes because of the geometry of the holes and louvers and the large temperature gradients (fig. 16). Operational conditions such as starting, stopping, and power surging, which produce thermal changes in the liner, are believed to raise and lower the buckle to some extent. The fluctuating thermal stresses thereby produced are believed to fatigue the metal thermally and to produce a crack in the buckle of the type shown in figure 3(c).\n\nBy alternately heating and cooling the louver flap, an outward and inward movement of the flap was produced. The operational conditions previously described therefore probably caused a similar motion of the flaps producing a thermal fatigue. Any bending of the louver flap would apply maximum bending stresses at the inner edge of the stress-relieving holes (fig. 3(a), point 1). Because the lower side of a flap projects into the path of the combustion gases and thus becomes hot and the upper side is cooled by the flow of secondary combustion air, tensile and compressive stresses are produced in upper and lower portions, respectively, as a result of differences in thermal expansion (fig. 17). These stresses produce the upward bending upon heating.", "timestamp": "2026-07-22T06:06:33.371238+00:00"}
{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 2, "total_pages": 72, "image_filename": "19930085491_p2.jpg", "text": "NACA RM No. A8J04 CONFIDENTIAL\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nRESEARCH MEMORANDUM\n\nAERODYNAMIC STUDY OF A WING-FUSELAGE COMBINATION EMPLOYING  \nA WING SWEPT BACK $63^\\circ$.— CHARACTERISTICS AT A MACH NUMBER  \nOF 1.53 INCLUDING EFFECT OF SMALL VARIATIONS OF SWEEP  \n\nBy Robert T. Madden\n\nSUMMARY\n\nWind-tunnel tests have been performed at a Mach number of 1.53 to determine experimentally the longitudinal characteristics of a wing-fuselage combination which theory indicates should be capable of attaining maximum lift-drag ratios greater than 10 to 1 at moderate supersonic speeds. The wing had a leading-edge sweep of $63^\\circ$, an aspect ratio of 3.42, a taper ratio of 0.25, and an NACA 64A006 section parallel to the plane of symmetry. The primary objectives of the investigation were to determine to what extent the theoretical maximum lift-drag ratio could be realized experimentally and to determine the static longitudinal stability characteristics. Secondary objectives included the evaluation of the effects of Reynolds number and small variations of sweep at a constant Mach number. To determine this latter effect, the wing panels were rotated about the midpoint of the root chord to obtain a variation of leading-edge sweep angle from $57.0^\\circ$ to $69.9^\\circ$. In addition to the force tests, liquid-film studies were made to determine the nature of the boundary-layer flow.\n\nAt a Reynolds number of 0.62 million, the $63^\\circ$ wing configuration had a maximum lift-drag ratio of 6.7; whereas theory indicated a value of 10.1. Liquid-film studies revealed that the difference between experiment and theory was primarily due to separation of the laminar boundary layer which occurred even at zero lift. Although the linear theory indicated a fixed center-of-pressure position, the experimental results showed that the center of pressure varied with lift coefficient over approximately 20 percent of the mean aerodynamic chord. This difference was also attributed to the effects of separation. As might be expected, increased Reynolds number had a marked influence on the extent of separation and\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:06:34.515374+00:00"}
{"citation_id": "19930093769", "source_url": "https://ntrs.nasa.gov/api/citations/19930093769/downloads/19930093769.pdf", "page_number": 28, "total_pages": 39, "image_filename": "19930093769_p28.jpg", "text": "NACA RM No. E8L10a CONFIDENTIAL 27\n\nCorrected specific fuel consumption based on net thrust, lb/(hr)(lb thrust)\n\nFuel\nAN-F-58\nGasoline\n\n(d) Altitude, 20,000 feet; flight Mach number, 1.00.\n\n(e) Altitude, 35,000 feet; flight Mach number, 1.00.\n\n(f) Altitude, 50,000 feet; flight Mach number, 0.85.\n\nCorrected engine speed, rpm\n\nFigure 6. - Concluded. Comparison of corrected specific fuel consumption based on net thrust for AN-F-58 fuel and gasoline at various altitudes and flight Mach numbers.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:06:41.286357+00:00"}
{"citation_id": "19930085485", "source_url": "https://ntrs.nasa.gov/api/citations/19930085485/downloads/19930085485.pdf", "page_number": 6, "total_pages": 26, "image_filename": "19930085485_p6.jpg", "text": "```markdown\n4\nCONFIDENTIAL\nNACA RM No. L8K02\n\nflow field generated over the curved surface of a bump on the tunnel\nfloor (fig. 3). A sketch showing the location of the model on the bump\nis given in figure 4. An electrical strain-gage balance was mounted in\na chamber in the bump to measure the aerodynamic forces and moments of\nthe model. The chamber is sealed except for a hole through which the\nbutt of the wing passes. The fuselage which was approximately $\\frac{1}{32}$ inch\nabove the bump surface covered this hole.\n\nThe chordwise variation of Mach number along the surface of the\nbump is shown in figure 5. This figure also presents the vertical varia-\ntion of Mach number at a chordwise station 12 inches from the leading edge\nof the bump. It should be noted that at a given tunnel Mach number the\nlocal Mach number obtained from surface static pressure measurements at\nstation 12 is somewhat higher than the maximum value indicated from the\nvertical survey (fig. 5). This difference in Mach number is brought about\nby not taking into account the total pressure loss in the boundary layer\nfor the surface survey. Extrapolation of the vertical survey to the\nsurface of the bump gives nearly the same Mach number as is obtained from\nthe surface survey. The test Mach number was the average Mach number over\nthe span of the model. The average Mach number over the span of the\nmodel is higher than the average Mach number over the span of the aileron\nby approximately 0.01 at the lowest Mach number and 0.03 at the highest\nMach number tested (fig. 5). No attempt has been made to evaluate the\neffect of the variation in Mach number along the chord and span of the\nmodel.\n\nMechanical difficulty with the balance used in the bump made it\nnecessary to conduct part of the investigation at subsonic speeds on the\nwall of the Langley high-speed 7- by 10-foot tunnel by means of the setup\nshown in figure 6. The reflection plane was spaced out from the tunnel\nwall to situate the model out of the tunnel boundary layer. The Mach\nnumber did not vary over 1 percent over the chord and span of the model\nfor these tests.\n\nThe variation of Reynolds number of the model with Mach number for\naverage conditions is presented in figure 7.\n\nRESULTS AND DISCUSSIONS\n\nWing-Fuselage Aerodynamic Characteristics\n\nThe lift and drag characteristics of the model obtained from the\ntransonic bump are presented in figures 8 and 9, respectively. The sub-\nsonic lift, drag, and pitching-moment characteristics for the model\nobtained from the wall mount are presented in figures 10 to 14. In\ngeneral, the results from the two methods are in good agreement.\n\nCONFIDENTIAL\n```", "timestamp": "2026-07-22T06:06:49.999631+00:00"}
{"citation_id": "19930085548", "source_url": "https://ntrs.nasa.gov/api/citations/19930085548/downloads/19930085548.pdf", "page_number": 3, "total_pages": 46, "image_filename": "19930085548_p3.jpg", "text": "2\nNACA RM No. E8L30\n\nINTRODUCTION\n\nThe potentialities of a gas-generator engine comprising a two-stroke-cycle compression-ignition engine, a compressor, and a turbine are presented in reference 1. In this type of power plant, the piston engine drives its own supercharging compressor and the exhaust gases from the engine are utilized in a turbine, which produces the net useful work of the cycle. A diagrammatic sketch of a gas-generator power plant is shown in figure 1.\n\nAside from the external operating conditions, the performance of the gas-generator engine is determined by (1) a maximum allowable cylinder pressure, (2) a maximum allowable turbine-inlet temperature, and (3) the necessity that the work output of the piston component of the engine must equal the work requirements of the compressor. In order to satisfy these three conditions simultaneously, compression ratio, manifold pressure, and fuel-air ratio must be adjusted to the proper values. Calculations in reference 1 indicate that compression ratios from 4 to 7, manifold pressures of approximately 80 to 160 pounds per square inch absolute, and over-all fuel-air ratios of approximately 0.03 may be used. The high inlet density results in high air capacity for the gas-generator engine and leads to a low specific engine weight, and the high expansion ratio results in good fuel economy.\n\nThe operation of the principal component of this power plant, the two-stroke-cycle compression-ignition engine, is certainly unique and unusual as compared with conventional compression-ignition-engine practice. Consequently, experimental data must be obtained pertaining to the performance of this component and the theory relative to the influence of the rate of combustion-pressure rise on the performance of a pressure- and temperature-limited cylinder; confirmation is also necessary of the theoretical expressions for such items as engine efficiency, heat rejection, compression and combustion pressures, and charging characteristics. An investigation was therefore conducted at the NACA Lewis laboratory to determine experimentally the performance characteristics of a ported cylinder for gas-generator use and to compare the results with those obtained in reference 1.\n\nThe unusual operating conditions imposed upon the engine may lead to questions about the mechanical practicability of such an engine. Although a limiting maximum cylinder pressure and exhaust temperature have been maintained, compression ratio, charging pressure and temperature, and exhaust pressure are so far removed from conventional practice that unforeseen mechanical and thermal loads may result in premature engine failure. A careful study of the practicability of operating engine cylinders at these conditions is therefore warranted.", "timestamp": "2026-07-22T06:06:54.149328+00:00"}
{"citation_id": "19930086076", "source_url": "https://ntrs.nasa.gov/api/citations/19930086076/downloads/19930086076.pdf", "page_number": 46, "total_pages": 50, "image_filename": "19930086076_p46.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:06:55.348561+00:00"}
{"citation_id": "19930082546", "source_url": "https://ntrs.nasa.gov/api/citations/19930082546/downloads/19930082546.pdf", "page_number": 17, "total_pages": 65, "image_filename": "19930082546_p17.jpg", "text": "16\nNACA TN No. 1870\n\nAt a given d/D value the order of agreement of the two methods is seen to change with tip Mach number and also may be different for each harmonic and at other points in space. This would preclude the use of Gutin's simplified equation with a convenient adjustment factor since the adjustment factor would probably be different in every case.\n\nPhase Relations\n\nThe fuselage-wall designer should know not only the relative amplitudes of the harmonics of pressure produced by the propeller but also something of the phase relations. Equation (1) will predict the phase between the impinging pressures of any given harmonic at two different points in space. The phase may also be predicted by use of equation (2). For given conditions equation (2) gives the pressure at a point in space as the product of a constant term and the square root of the sum of the squares of the real and imaginary components which are, respectively, the first and last terms within the large parentheses. If the algebraic values of each of these terms are known, the phase relations may be easily determined.\n\nBy this method calculations of the pressures produced simultaneously by the fundamental frequency at two points in space, equidistant ahead of and behind the propeller plane and for a tip Mach number of 0.75, gave a phase difference of 165°. Comparative measurements at these same operating conditions gave a corresponding value of 155°; thus the validity of equation (2) is further verified. Similar calculations for the same propeller at the same tip speed but for a larger blade-angle setting gave a phase difference of 125°. A comparison of these results indicates that the phase angle between the pressures ahead of and behind the propeller plane tends to decrease in magnitude as $C_Q$ increases with respect to $C_T$.\n\nFigure 18 shows the total pressure wave forms as recorded at three different points in space for five different tip Mach numbers. These are Du Mont dual-beam cathode-ray oscillograph pictures of the microphone voltage output, which is the upper trace, and a timing line of 300 cycles per second. The small vertical line on the timing line indicates the time at which the propeller blade passes through the xy-plane and is closest to the microphone. The line tracing the pressure indicates positive pressure when it moves downward and negative pressure when it moves upward, and time increases from left to right. The photographs taken at a tip Mach number of 1.00 indicate a relatively large contribution by the higher harmonics, whereas at the lower tip Mach numbers the low harmonics are clearly predominant. Figure 18 is included primarily for information in case a more detailed analysis of these wave forms is desired.", "timestamp": "2026-07-22T06:06:55.558987+00:00"}
{"citation_id": "19930086097", "source_url": "https://ntrs.nasa.gov/api/citations/19930086097/downloads/19930086097.pdf", "page_number": 34, "total_pages": 36, "image_filename": "19930086097_p34.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:06:57.003913+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 78, "total_pages": 96, "image_filename": "19930085880_p78.jpg", "text": "76\nNACA RM No. L9C03\n\n[Figure: A graph plotting Load, lb against Wetted area, sq ft. The y-axis ranges from 0 to 32. The x-axis ranges from 0 to .35. There are five curves representing different speeds (10, 15, 20, 25, 30 fps). A symbol of an inverted triangle is shown in the upper left corner.]\n\nSpeed\n(fps)\n30\n25\n20\n15\n10\n\nLoad, lb\nWetted area, sq ft\n\n(b) $\\tau = 8^\\circ$.\n\nNACA\n\nFigure 22.- Continued.", "timestamp": "2026-07-22T06:06:58.669423+00:00"}
{"citation_id": "19930082592", "source_url": "https://ntrs.nasa.gov/api/citations/19930082592/downloads/19930082592.pdf", "page_number": 13, "total_pages": 50, "image_filename": "19930082592_p13.jpg", "text": "12\nNACA TN 1914\n\nThe maximum percentage of cobalt that may be added without a drastic increase in oxidation penetration at 1625° F is apparently between 20 and 30 percent. At temperatures in the vicinity of 2000° F, the use of 50-percent cobalt will not result in much more oxidation than would be encountered with 20 percent of cobalt. The critical composition of tungsten ceramal at 1625° F is between 10 and 20 percent of tungsten with the 20-percent-tungsten ceramal showing a marked increase in oxidation penetration. At higher temperatures, no sudden change in oxidation-rate constant with change in temperature is noted. The oxidation-rate constants of the 20- and 30-percent-tungsten ceramals at 2000° F are approximately equal.\n\nThe resistance to oxidation of ceramals within the range of compositions and temperatures investigated may be determined by using figures 1(a), 4(d), and 6(d). By interpolation, the oxidation penetration to be expected as a result of exposure at temperatures from 1625° to 2000° F and with a variation in metallic composition up to 30 percent can be estimated.\n\nThese results make it clear that before the high melting points and the high-temperature properties of metallic molybdenum, tungsten, and cobalt can be used to advantage in air atmospheres, protection of these metals will be necessary either by use of coatings applied to the exterior or by use of additional components in the ceramal itself that will combine chemically with the oxygen or with the oxides formed into a dense material highly resistant to the diffusion of oxygen. The cobalt ceramal shows promise because of the type of oxide coating formed. This coating forms in two layers with the inner layer occurring as a composite of the body and oxides, with a gradual change in composition moving outward from the inner interface.\n\nSUMMARY OF RESULTS\n\nThe following results were obtained from an investigation conducted to determine the oxidation-penetration characteristics of tungsten carbide base ceramals containing 5, 10, 20, and 30 percent of molybdenum, tungsten, or cobalt at various temperatures and exposure periods:\n\n1. The oxidation resistance in air atmospheres at 1625°, 1785°, and 2000° F of molybdenum, tungsten, and cobalt - titanium carbide base ceramals was found to vary in the following manner:\n\n(a) The depth of oxide penetration for the cobalt and tungsten ceramals was approximately the same through the time", "timestamp": "2026-07-22T06:07:01.505439+00:00"}
{"citation_id": "19930082613", "source_url": "https://ntrs.nasa.gov/api/citations/19930082613/downloads/19930082613.pdf", "page_number": 12, "total_pages": 46, "image_filename": "19930082613_p12.jpg", "text": "NACA TN 1938\n\nMechanically produced fatigue stresses are also believed to contribute to the cracking mechanism and may result from vibrations of the engine, air flowing over the flaps, and pulsations that occur during combustion. Air flow would tend to produce a motion in the flap similar to that of a vibrating reed.\n\nTranscrystalline characteristics of the cracks and the brittle nature of the failure are also indicative of fatigue failures, particularly because the longer, more rapidly propagated cracks have few intergranular characteristics except in the branches.\n\nStress raisers may be classified into two groups: those produced by fabrication processes and those formed after the liner is put into operation.\n\nThe investigation has shown the importance of stress raisers produced by the punching operations, namely, that by removing torn and worked metal from punched edges cracking is greatly retarded and liner life thereby increased. Because most cracks originate in the upper bend of the lower flaps at the inner portion of stress-relieving holes, it is possible that the working of the grains in the bend is also harmful. Surface scales, which formed during engine operation are believed not only to lengthen and widen the cracks, but to act as stress raisers at the tips. Subsurface scales, which are also found at most crack edges, are inherently harmful because by expanding the metallic lattice they act as stress raisers and lower resistance to fatigue.\n\nThe intercrystalline characteristics of many cracks or branches of cracks and the accompanying scale formations indirectly indicate that some cracks first start at grain boundaries weakened or stressed by oxide penetrations. Similarly, the improvement in resistance to cracking of mechanically finished liners indicates that other cracks originate from small fissures produced by punching operations. Because the vast majority of cracks were shown to originate in the stress-relieving holes of the louvers, however, and because bending of the flap produces maximum stresses at the stress-relieving holes, it is believed that the cracks are predominantly the result of louver-flap motion. Thermal and mechanical fatigue stresses could therefore cause cracking at the stress-relieving holes even if stress raisers from punching and scaling were not present.", "timestamp": "2026-07-22T06:07:01.644256+00:00"}
{"citation_id": "19930082918", "source_url": "https://ntrs.nasa.gov/api/citations/19930082918/downloads/19930082918.pdf", "page_number": 5, "total_pages": 62, "image_filename": "19930082918_p5.jpg", "text": "4\n\nNACA TN 1940\n\nPrior to use, the stock for this investigation was solution-treated 10 hours at $2200^\\circ$ F and water-quenched. The solution treatment was made unusually long for the purpose of distributing the precipitant atoms randomly in the matrix and so that internal strain, due to the prior working of the material, would be reduced to a very low level.\n\nModerate grain growth took place over the major portion of the bar cross section during the 10-hour treatment, but on two diagonally opposite corners pronounced growth took place. A hardness survey of the as-rolled bar stock showed the average hardness across one diagonal to be higher than across the other; from this the conclusion may be drawn that the rolling operation had worked the bar across one diagonal preferentially. All mechanical testing and physical measurements were restricted to the fine-grained section of the bar stock. Figure 1 shows representative structures of the cross section of the bar stock as-rolled and after the 10-hour solution treatment.\n\nEXPERIMENTAL PROCEDURES\n\nThe general procedure was to age the solution-treated stock at selected temperatures and time periods and then to carry out the microstructural studies and X-ray diffraction measurements in order to establish the structural characteristics resulting from the aging treatments. The strength properties resulting from the aging treatments were measured for short time periods at $1200^\\circ$ F. These experiments were intended to establish the relationship between short-time creep and rupture properties and nucleation, precipitation, and precipitate particle size and distribution during aging.\n\nThe details of the experimental procedures are described in the following sections.\n\nAging\n\nAging treatments were carried out at $1200^\\circ$, $1400^\\circ$, and $1600^\\circ$ F for time periods of 1, 10, 100, and 1000 hours and such other intermediate times as became necessary. The samples were heated in small automatically controlled muffle furnaces in an air atmosphere. These furnaces were at temperature when samples were placed in them and the time period of heating was considered started after the specimens had been in the furnace for 1/4 hour. After aging, all samples were air-cooled. Sufficient stock was aged at each condition for the microstructural, X-ray, and mechanical tests.", "timestamp": "2026-07-22T06:07:01.935390+00:00"}
{"citation_id": "19930093769", "source_url": "https://ntrs.nasa.gov/api/citations/19930093769/downloads/19930093769.pdf", "page_number": 29, "total_pages": 39, "image_filename": "19930093769_p29.jpg", "text": "28\nCONFIDENTIAL\nNACA RM No. EBL10a\n\nFuel\n-Δ- AN-F-58\n--□-- Gasoline\n\nCombustion efficiency, percent\n(a) Altitude, 5000 feet; flight Mach number, 0.\n\nCombustion efficiency, percent\n(b) Altitude, 20,000 feet; flight Mach number, 0.60.\nCorrected engine speed, rpm\nNACA\n\nFigure 7. - Comparison of combustion efficiency for AN-F-58 fuel and gasoline at various altitudes and flight Mach numbers.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:07:03.728083+00:00"}
{"citation_id": "19930082566", "source_url": "https://ntrs.nasa.gov/api/citations/19930082566/downloads/19930082566.pdf", "page_number": 14, "total_pages": 44, "image_filename": "19930082566_p14.jpg", "text": "12\n\nTABLE 1.- FATIGUE TEST DATA FOR STRESS RATIO $R = \\sigma_2/\\sigma_1 = 0$\n\n| Specimen | Load, P' (lb) | Load, P''' (lb) | Diameter, d (in.) | Average wall thickness, t (in.) | Stress, $\\sigma_1'$ (psi) | Stress, $\\sigma_2'$ (psi) | Stress, $\\sigma_1'''$ (psi) | Stress, $\\sigma_2'''$ (psi) | Stress ratio, R $\\sigma_2'/\\sigma_1'$ | Number of cycles, N |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| G-4 | $16.0 \\times 10^3$ | $1.00 \\times 10^3$ | 2.00 | 0.0534 | $48.0 \\times 10^3$ | $0 \\times 10^3$ | $3.0 \\times 10^3$ | $0 \\times 10^3$ | 0 | $0.0599 \\times 10^6$ |\n| G-3 | 15.0 | 1.50 | 2.00 | .0551 | 13.0 | 0 | 4.0 | 0 | 0 | .0653 |\n| F-3 | 22.0 | 1.50 | 2.00 | .0766 | 46.0 | 0 | 3.0 | 0 | 0 | .0733 |\n| G-8 | 11.0 | 1.00 | 2.00 | .0530 | 33.0 | 0 | 3.0 | 0 | 0 | .0889 |\n| G-7 | 21.0 | 1.50 | 2.00 | .0776 | 43.0 | 0 | 3.0 | 0 | 0 | .0950 |\n| G-5 | 14.5 | 2.00 | 2.00 | .0530 | 44.0 | 0 | 6.0 | 0 | 0 | .1070 |\n| G-6 | 11.0 | .75 | 2.00 | .0526 | 33.0 | 0 | 2.0 | 0 | 0 | .1342 |\n| H-2 | 9.5 | 1.00 | 2.00 | .0527 | 29.0 | 0 | 3.0 | 0 | 0 | .3105 |\n| H-7 | 7.2 | .70 | 2.00 | .0499 | 23.0 | 0 | 2.0 | 0 | 0 | 2.0654 |\n\nTABLE 2.- FATIGUE TEST DATA FOR STRESS RATIO $R = \\sigma_2/\\sigma_1 = 2.0$\n\n| Specimen | Pressure, P' (psi) | Pressure, P''' (psi) | Diameter, d (in.) | Average wall thickness, t (in.) | Stress, $\\sigma_1'$ (psi) | Stress, $\\sigma_2'$ (psi) | Stress, $\\sigma_1'''$ (psi) | Stress, $\\sigma_2'''$ (psi) | Stress ratio, R $\\sigma_2'/\\sigma_1'$ | Number of cycles, N |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| H-3 | $1.750 \\times 10^3$ | $0.200 \\times 10^3$ | 2.00 | 0.0507 | $17.25 \\times 10^3$ | $34.50 \\times 10^3$ | $2.00 \\times 10^3$ | $4.00 \\times 10^3$ | 2.0 | $0.0438 \\times 10^6$ |\n| H-5 | 1,400 | .200 | 2.00 | .0514 | 13.50 | 27.00 | 2.00 | 4.00 | 2.0 | .0513 |\n| H-4 | 1,400 | .200 | 2.00 | .0526 | 13.25 | 26.50 | 2.00 | 4.00 | 2.0 | .1074 |\n| H-6 | 1,050 | .200 | 2.00 | .0533 | 9.75 | 19.50 | 2.00 | 4.00 | 2.0 | .2185 |\n| F-1 | .750 | .200 | 2.00 | .0471 | 8.00 | 16.00 | 2.00 | 4.00 | 2.0 | .2395 |\n| H-8 | .750 | .100 | 2.00 | .0525 | 7.15 | 14.30 | .95 | 1.90 | 2.0 | .2481 |\n| H-5 | .750 | .125 | 2.00 | .0505 | 7.41 | 14.82 | 1.24 | 2.48 | 2.0 | .5194 |\n| H-10 | .825 | .200 | 2.00 | .0533 | 7.75 | 15.50 | 2.00 | 4.00 | 2.0 | .5643 |\n| G-10 | .625 | .125 | 2.00 | .0515 | 6.06 | 12.12 | 1.21 | 2.42 | 2.0 | 8.2253 |\n\nNACA TN No. 1889", "timestamp": "2026-07-22T06:07:16.146369+00:00"}
{"citation_id": "19930085548", "source_url": "https://ntrs.nasa.gov/api/citations/19930085548/downloads/19930085548.pdf", "page_number": 4, "total_pages": 46, "image_filename": "19930085548_p4.jpg", "text": "NACA RM No. E8L30\n\nFor this work, a small-scale, loop-scavenged, two-stroke-cycle, compression-ignition cylinder was selected as the simplest type of cylinder that was expected to satisfy the gas-generator requirements. This cylinder was operated with compression ratios of 4 to 7, inlet-manifold temperatures ranging from $300^\\circ$ to $600^\\circ$ F, manifold pressures of 80 to 135 pounds per square inch absolute, and over-all fuel-air ratios up to 0.060. The engine speed was held constant at 1800 rpm, and the charge air flow limited to 1 cylinder volume per cycle. Although this engine speed and this rate of flow are not necessarily optimum, the values were selected as the mean between possible limits of the variables (reference 1) in order to limit the number of variables under investigation.\n\nAPPARATUS\n\nA ported cylinder with a $3\\frac{1}{4}$-inch bore and a $4\\frac{1}{2}$-inch stroke was fabricated from steel and the bore was chrome-plated to prevent rapid wear. A detachable cylinder head with various spacers afforded a means of obtaining a change in compression ratio. The inlet- and exhaust-port arrangements were similar to those used by Rogowski and Bouchard (reference 2, fig. 4, section D-D). In this design, two of the eight inlet ports were inclined at an angle of $60^\\circ$ with the base, and the horizontal inlet angles of the other six inlet ports were so arranged as to direct the incoming air upward and toward the inlet side of the cylinder (fig. 2). Four cast-iron piston rings (wedge-shaped cross section) were used above the piston pin; two rings (rectangular cross section) were used below the pin to seal the manifold pressure from the crankcase. A four-plunger pump driven at one-tenth engine speed provided metered lubrication to the cylinder bore at four equally spaced points just above the top of the ports. An oil jet from the small end of the connecting rod was directed at the under side of the crown of the aluminum-alloy piston to cool the piston. The cylinder was mounted on a CFR crankcase. A 100-horsepower dynamometer equipped with the necessary accessories and instrumentation was used to start the engine, to absorb the power, and to motor the engine in order to obtain friction data and compression pressures. The dynamometer torque was indicated by scales. Figures 3 and 4 show general views of the setup.\n\nThe fuel-injection pump had a 10-millimeter plunger and a 10-millimeter stroke. The maximum plunger velocity was 0.0125 inch per degree of cam rotation. The spring-loaded injection valve had an opening pressure of 3300 pounds per square inch. Cross sections of the combustion chamber at various compression ratios, the location of the injection valve, and a sketch of the spray pattern are shown", "timestamp": "2026-07-22T06:07:17.632139+00:00"}
{"citation_id": "19930085485", "source_url": "https://ntrs.nasa.gov/api/citations/19930085485/downloads/19930085485.pdf", "page_number": 7, "total_pages": 26, "image_filename": "19930085485_p7.jpg", "text": "NACA RM No. L8K02 CONFIDENTIAL 5\n\nThe variation of lift coefficient with Mach number at several angles of attack for the aileron with t = 0 (circular arc), t = 0.50 (flat sides), and t = 1.00 (flat sides) is presented in figure 8. From these data it appears that the variation of lift coefficient with Mach number is practically unaffected by the various aileron profiles tested. The variation of lift coefficient with angle of attack for several subsonic Mach numbers is presented in figures 10 to 12. At a Mach number of 0.607 the lift-curve slope near an angle of attack of 0° is less than the slope at larger angles of attack; however, at higher Mach numbers the slope becomes more nearly linear through the test angle-of-attack range. The low Reynolds number at which this investigation was made is probably responsible for part of the nonlinearity of the lift curves near zero lift.\n\nThe drag characteristics for the model with circular-arc aileron (t = 0) are shown in figures 10 and 13. At subcritical speeds the drag coefficient at an angle of attack of 0° is nearly constant at a value of about 0.011. The drag rise for this model configuration comes at a Mach number of approximately 0.90. Increments of drag coefficients $\\Delta C_D$ resulting from changes in aileron contour are presented in figure 9, for Mach numbers between 0.50 and 1.15. The flat-sided aileron, (t = 0.50) shows a small increase in drag coefficient (0.002) over the circular-arc contour aileron below a Mach number of 1.05. Above a Mach number of 1.05 the values of $\\Delta C_D$ become negative. As might be expected, the flat-sided aileron (t = 1.00) gave greater drag coefficients, approximately 0.006 at subsonic Mach numbers, than the circular-arc aileron (t = 0) through the Mach number range tested.\n\nThe pitching-moment characteristics for the various aileron modifications investigated are presented in figures 10 to 14. These results are summarized in the following table:\n\n| Aileron contour | M | Aerodynamic-center location (percent $\\bar{c}$) | | |\n| :--- | :--- | :--- | :--- | :--- |\n| | | $C_L = -0.2$ | $C_L = 0$ | $C_L = 0.2$ |\n| Circular-arc | 0.607 | 18 | -7 | 18 |\n| Circular-arc | .941 | -- | 41 | 25 |\n| Flat-sided, t = 0.50 | .607 | 24 | 9 | 20 |\n| Flat-sided, t = 0.50 | .934 | -- | 30 | 30 |\n| Flat-sided, t = 1.00 | .597 | 29 | 21 | 24 |\n| Flat-sided, t = 1.00 | .965 | -- | 38 | 38 |\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:07:19.506727+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 79, "total_pages": 96, "image_filename": "19930085880_p79.jpg", "text": "NACA RM No. L9C03\n77\n\n[Figure: A graph plotting Load (lb) on the y-axis against Wetted area (sq ft) on the x-axis. The y-axis ranges from 0 to 32. The x-axis ranges from 0 to .35. There are five curves representing different speeds (10, 15, 20, 25, 30 fps). The curves use different markers: circles for 10, squares for 15, diamonds for 20, triangles for 25, and inverted triangles for 30. There is a small inset diagram in the upper left corner showing a triangle shape.]\n\nSpeed\n(fps)\n30\n25\n20\n15\n10\n\nLoad, lb\n32\n28\n24\n20\n16\n12\n8\n4\n0\n\nWetted area, sq ft\n0\n.05\n.10\n.15\n.20\n.25\n.30\n.35\n\nNACA\n\n(c) $\\tau = 12^\\circ$.\n\nFigure 22.- Continued.", "timestamp": "2026-07-22T06:07:21.764681+00:00"}
{"citation_id": "19930086076", "source_url": "https://ntrs.nasa.gov/api/citations/19930086076/downloads/19930086076.pdf", "page_number": 47, "total_pages": 50, "image_filename": "19930086076_p47.jpg", "text": "NACA RM E9F09\n45\n\n[Figure: A damaged metal flame holder assembly with zigzag weld patterns, mounted on a base plate with bolt holes. A scale ruler labeled \"INCHES\" is visible at the bottom right. A NACA stamp reads \"C-22919 2-4-49\".]\n\nFigure 18. - Flame holder 13 after 5 minutes of operation.", "timestamp": "2026-07-22T06:07:23.534114+00:00"}
{"citation_id": "19930093769", "source_url": "https://ntrs.nasa.gov/api/citations/19930093769/downloads/19930093769.pdf", "page_number": 30, "total_pages": 39, "image_filename": "19930093769_p30.jpg", "text": "NACA RM No. EBL10a CONFIDENTIAL 29\n\nCombustion efficiency, percent\n\nFuel\n-△- AN-F-58\n--□-- Gasoline\n\n(c) Altitude, 20,000 feet; flight Mach number, 0.85.\n\n(d) Altitude, 20,000 feet; flight Mach number, 1.00.\n\nCorrected engine speed, rpm\n\n[NACA logo]\n\nFigure 7. - Continued. Comparison of combustion efficiency for AN-F-58 fuel and gasoline at various altitudes and flight Mach numbers.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:07:31.533933+00:00"}
{"citation_id": "19930086097", "source_url": "https://ntrs.nasa.gov/api/citations/19930086097/downloads/19930086097.pdf", "page_number": 35, "total_pages": 36, "image_filename": "19930086097_p35.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:07:33.672995+00:00"}
{"citation_id": "19930082592", "source_url": "https://ntrs.nasa.gov/api/citations/19930082592/downloads/19930082592.pdf", "page_number": 14, "total_pages": 50, "image_filename": "19930082592_p14.jpg", "text": "NACA TN 1914\n\nand temperature range investigated. The cobalt ceramals were considered better with regard to over-all oxidation because of the type of oxide coating formed.\n\n(b) At $1625^\\circ$ F, the 20- and 30-percent-tungsten ceramals oxidized much more rapidly than the 5- and 10-percent-tungsten ceramals. At $2000^\\circ$ F, the oxidation-penetration rates of the tungsten ceramals were of the same order of magnitude.\n\n(c) The molybdenum ceramals were inferior to both the cobalt and the tungsten ceramals in oxidation resistance, showing no self-protective properties. The oxide coatings formed on the molybdenum ceramals did not inhibit further oxidation. Their oxidation rates increased greatly as temperature increased and as the percentage of molybdenum was increased. At $1625^\\circ$ F, the oxidation rates of the 20- and 30-percent-molybdenum ceramals were lower than would be expected from the rates of the other compositions.\n\n2. The oxidation-penetration - time curves may be used to estimate the expected depths of oxidation penetration through the temperature range of $1625^\\circ$ to $2000^\\circ$ F and 5- to 30-percent metallic element.\n\nCONCLUSION\n\nCombined with data on the effect of the metallic components on the strength properties of the ceramals, the oxidation-penetration curves and the oxidation-rate constant plotted against the reciprocal of absolute temperature can be used in deciding the optimum percentage of alloying metallic element to add to a titanium carbide body for the highest strength within the allowable oxidation-rate limit.\n\nLewis Flight Propulsion Laboratory,  \nNational Advisory Committee for Aeronautics,  \nCleveland, Ohio, February 11, 1949.\n\nREFERENCES\n\n1. Deutsch, George C., Repko, Andrew J., and Lidman, William G.: Elevated-Temperature Strength of Several Titanium-Carbide Ceramals. NACA TN 1915, 1949.", "timestamp": "2026-07-22T06:07:33.876054+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 37, "total_pages": 114, "image_filename": "19930086061_p37.jpg", "text": "```markdown\nNACA RM L9207\n\nVertical distance from tunnel floor, feet\n10\n9\n8\n7\n6\n5\n-4 -3 -2 -1 0 1 2 3 4\nLeft Horizontal distance from model mount, feet Right\n\n$\\Psi = -35^\\circ$\n$\\alpha = 40^\\circ$\n\n$\\Psi = 0^\\circ$\n$\\alpha = 40^\\circ$\n\n$\\Psi = 35^\\circ$\n$\\alpha = 40^\\circ$\n\n-2.0°\n$\\Psi = -35^\\circ$\n$\\alpha = 0^\\circ$\n\n-1.5°\n\n-1.0°\n$\\Psi = 0^\\circ$\n$\\alpha = 0^\\circ$\n\n-0.5°\n\n0.5°\n\n0°\n\n0.5°\n\n$\\Psi = 35^\\circ$\n$\\alpha = 0^\\circ$\n\nTunnel\n$\\ell$\nPositive\nangle\n\nTunnel\n$\\ell$\n\nNACA\n\n(b) Yaw angularity.\n\nFigure 5.- Concluded.\n\n33\n```", "timestamp": "2026-07-22T06:07:35.676143+00:00"}
{"citation_id": "19930082546", "source_url": "https://ntrs.nasa.gov/api/citations/19930082546/downloads/19930082546.pdf", "page_number": 18, "total_pages": 65, "image_filename": "19930082546_p18.jpg", "text": "NACA TN No. 1870\n\nCHARTS FOR ESTIMATING FREE-SPACE PRESSURES\n\nThe theory given in this paper is adequate for predicting free-space oscillating pressures for any static condition. The complexity of the method, however, makes it desirable to provide a more convenient means of estimating these pressures. Therefore the charts of figure 19 are presented. In contrast to the analytical method these charts do not predict the pressures at a given point but instead give a first approximation of the maximum free-space pressure coefficients of a given harmonic near the plane of rotation of the propeller. This information may be determined easily from the appropriate chart, provided that the power coefficient, tip Mach number, and tip clearance are known for a given propeller.\n\nThe charts are based on data for uninstalled conditions and the pressures involved were determined by averaging the maximum values measured in front of and behind the plane of rotation at each test condition. These maximum values usually occurred at $\\frac{x}{D} = \\pm \\frac{1}{8}$. The free-space pressure coefficients thus obtained were found to vary approximately linearly with power coefficient as do those measured in the plane of rotation. (See fig. 11.) Thus the thrust terms are neglected and the charts are based on power coefficients of the tests. The charts may be used, however, for power coefficients larger than those for which data were taken. The charts are based primarily on experimental measurements at $\\frac{d}{D} = 0.083$ and on a sufficient number of measurements at other $d/D$ values to establish the attenuation curve in figure 20. This curve was faired from a composite plot of data which were adjusted to equal magnitudes at $\\frac{d}{D} = 0.083$.\n\nCharts for mB values of 2, 3, 4, 5, 6, and 8 were all determined by faired data from two-blade and four-blade propellers. In equation (2) where m and B always appear as a product, the second harmonic of a two-blade propeller has the same strength as the fundamental of a four-blade propeller for the same operating conditions. Because of this fact, which has also been confirmed experimentally, and because the fundamental frequency has been found to be predominant in this critical region of maximum pressures, the charts are useful for estimating pressures produced by the fundamental frequencies of propellers which have from 2 to 8 blades; they may also be used to predict the pressures of harmonics in the range of values of mB from 2 to 8.\n\nAs first illustrated in figure 12, the charts show in general that at tip Mach number 1.00 all harmonics have very nearly the same maximum amplitude for comparable operating conditions whereas at the lower tip Mach numbers the lower-order harmonics are predominant.", "timestamp": "2026-07-22T06:07:36.100606+00:00"}
{"citation_id": "19930083221", "source_url": "https://ntrs.nasa.gov/api/citations/19930083221/downloads/19930083221.pdf", "page_number": 3, "total_pages": 47, "image_filename": "19930083221_p3.jpg", "text": "NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nTECHNICAL NOTE NO. 1824\n\nLINEARIZED COMPRESSIBLE-FLOW THEORY\n\nFOR SONIC FLIGHT SPEEDS\n\nBy Max. A. Heaslet, Harvard Lomax, and John R. Spreiter\n\nSUMMARY\n\nThe partial differential equation for the perturbation velocity potential is examined for free-stream Mach numbers close to and equal to one. It is found that, under the assumptions of linearized theory, solutions can be found consistent with the theory for lifting-surface problems both in stationary three-dimensional flow and in unsteady two-dimensional flow. Several examples are solved including a three-dimensional swept-back wing and a two-dimensional harmonically oscillating wing, both for a free-stream Mach number equal to one.\n\nINTRODUCTION\n\nMuch of the recent progress in the theoretical analysis of compressible-flow fields is attributable to the successful application of linearization methods. Although the basic assumptions used in conventional linearized theory appear at first glance to be highly restrictive, it has been found that, just as in the analogous case of thin-airfoil theory for incompressible flow, the methods have many fields of utilization adequate for most engineering purposes. Since the basic methods are so well known and depend on such relatively simple mathematical tools, it appears obvious that the range of applicability of the theory should be explored completely. Such is the purpose of the present report. It has been more or less tacitly presumed in the past that such applications cannot treat cases for which the flight velocity is near the speed of sound. In the study of two-dimensional steady-state problems in airfoil theory, this presumption is certainly true. The Prandtl-Glauert and Ackeret rules for variation of pressure coefficient with free-stream Mach number in the subsonic and supersonic regimes, respectively, are clearly invalid for Mach numbers near one, since perturbation velocities", "timestamp": "2026-07-22T06:07:36.235928+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 27, "total_pages": 78, "image_filename": "19930082483_p27.jpg", "text": "NACA TN No. 1807\n25\n\nThe curves (each representing constant speeds) may then be separated by displacing each curve, together with its abscissa, to the right a convenient distance $X$ proportioned to the rotor speed. The result is a series of vertically oriented plots of corrected power output against pressure ratio with the curves of lower speed at the left and the speeds increasing towards the right. (See diagram (b).)\n\nAll these curves, each representing a constant corrected rotor speed, cover the same range of pressure ratios and so permit construction of horizontal curves connecting points of equal values of pressure ratio on the constant speed curves. (See diagram (c).)\n\n[Figure: Diagram (c) and Diagram (d) showing turbine power output corrected to sea level vs. $p_1'/p_e'$ and $N/\\sqrt{\\theta_1}$]\n\nThese curves of constant pressure ratio perform the same function as do the series of pressure-ratio abscissas associated with the constant-speed curves, so that the abscissas may be removed. In this manner a performance grid can be readily constructed upon which may be superimposed contours of constant over-all efficiency. (See diagram (d).)\n\nThe utility of this form of performance presentation becomes apparent if consideration is given to the information regarding", "timestamp": "2026-07-22T06:07:42.939932+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 3, "total_pages": 149, "image_filename": "19930083192_p3.jpg", "text": "Page\n\nTRANSIENT AERODYNAMICS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 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. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .", "timestamp": "2026-07-22T06:07:48.318686+00:00"}
{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 3, "total_pages": 72, "image_filename": "19930085491_p3.jpg", "text": "2\nCONFIDENTIAL\nNACA RM No. A8J04\n\nconsequently on the measured aerodynamic characteristics. Increasing\nthe Reynolds number to 0.84 million increased the maximum lift-drag\nratio to 7.2 and reduced the total center-of-pressure travel to\napproximately 12 percent of the mean aerodynamic chord.\n\nIn the determination of the effects of sweep, it was found that\nthe sweep angle for maximum lift-drag ratio was 67° for this general\ntype of configuration at a Mach number of 1.53. Values of maximum\nlift-drag ratio of 7.1 and 7.4 were obtained at Reynolds numbers of\n0.62 and 0.95 million, respectively. The optimum sweep angle resulted\nfrom the decrease in minimum drag coefficient and the increase in\ndrag due to lift as the sweep angle was increased. The total center-\nof-pressure travel with lift coefficient increased with increasing\nangles of sweep.\n\nThe results of these tests indicate that further improvements\nin maximum lift-drag ratio and longitudinal stability may be expected\nat full-scale Reynolds numbers. However, since the large adverse\nlifting-pressure gradients may cause leading-edge separation even at\nhigh Reynolds numbers, the theoretical value of maximum lift-drag\nratio may never be obtained with this wing. Therefore, the use of\ncamber and twist to reduce the adverse gradient is indicated as a\nmeans of improving the boundary-layer flow characteristics and maxi-\nmum lift-drag ratio.\n\nINTRODUCTION\n\nThe possibility of attaining supersonic flight speeds without a\nlarge increase in fuel consumption per mile of flight over that\nrequired for level subsonic flight depends largely upon obtaining\nhigh lift-drag ratios at the desired flight Mach number. The theo-\nretical aspects of efficient supersonic flight have been considered\nby Jones in reference 1. As a result of this theoretical study, it\nhas been indicated that lift-drag ratios greater than 10 to 1 may be\nobtained up to a Mach number of approximately 1.5 by using large\nangles of sweepback and relatively high aspect ratios. Thus the\nthrust required and the fuel consumption for level supersonic flight\nnear a Mach number of 1.5 should be considerably less than that\nnecessary for straight-wing configurations which develop lift-drag\nratios of approximately 6 to 1.\n\nThe most effective gains resulting from the use of sweepback\nat supersonic flight speeds are realized when the wing leading edge\nis swept behind the Mach lines originating at the apex of the wing\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:07:50.437183+00:00"}
{"citation_id": "19930085548", "source_url": "https://ntrs.nasa.gov/api/citations/19930085548/downloads/19930085548.pdf", "page_number": 5, "total_pages": 46, "image_filename": "19930085548_p5.jpg", "text": "4\nNACA RM No. E8L30\n\nin figure 5. The nozzle used was selected on the basis of a brief\npreliminary investigation. The fuel had a cetane number of 50, a\nspecific gravity of 0.835 at 60° F, and a hydrogen-carbon ratio of\n0.149. Fuel flow was measured with a rotameter.\n\nHigh-pressure combustion-scavenging air was obtained from the\nlaboratory air system. Weight flow was controlled by suitable valves\nin the inlet and exhaust systems and was measured by a thin-plate\norifice installed according to A.S.M.E. specifications. Surge tanks\n(figs. 4 and 6) located before and after the engine were equipped\nwith pressure taps to measure the inlet-manifold and exhaust-gas\npressures. Maximum cylinder pressures were measured with a balanced-\ndiaphragm valve and a pressure gage. A mercury manometer connected\nbetween the inlet manifold and the exhaust tank was used to indicate\nthe pressure drop across the cylinder during operation. The readings\nwere in close agreement with the differences between inlet-manifold\nand exhaust-gas pressures as indicated by calibrated Bourdon gages.\n\nAn electrically operated gas-sampling valve was connected to\nthe combustion chamber from which gas samples were directly piped\nto a mixture analyzer (reference 3). Samples of gas for a 20°\ncrank-angle period could be obtained for any desired part of the\nstroke.\n\nDuring part of this investigation, the engine was operated on\na four-stroke cycle, that is, with fuel injection at the end of\nevery second compression stroke in order to insure the removal of\nunburned fuel from the cylinder prior to the air-charging process.\nThis operation was accomplished by the use of 2:1 reduction gear\nbetween the engine and the fuel-injection pump.\n\nIndicator-card (pressure-time diagrams) data were obtained with\na modified Farnboro electric indicator (reference 4).\n\nPROCEDURE\n\nVariable fuel-air-ratio runs were made over a range of inlet-\nmanifold pressures and corresponding exhaust-gas pressures, so that\nthe chosen scavenging ratio (ratio of volume of air flowing through\nthe cylinder per cycle measured at inlet-manifold conditions to the\nvolume of the cylinder at the time of port closing) was constant.\nThe inlet-air flow was controlled by throttling the flow of exhaust\ngases. The engine speed was held constant by varying the load on\nthe dynamometer for changes in fuel-air ratio; the fuel-injection\nadvance angle was also held constant.", "timestamp": "2026-07-22T06:07:54.231440+00:00"}
{"citation_id": "19930093769", "source_url": "https://ntrs.nasa.gov/api/citations/19930093769/downloads/19930093769.pdf", "page_number": 31, "total_pages": 39, "image_filename": "19930093769_p31.jpg", "text": "30\nCONFIDENTIAL\nNACA RM No. E8L10a\n\n<!-- Image (138, 122, 839, 866) -->\n\n(e) Altitude, 35,000 feet; flight Mach number, 1.00.\n(f) Altitude, 50,000 feet; flight Mach number, 0.85.\nFigure 7. - Concluded. Comparison of combustion efficiency for\nAN-F-58 fuel and gasoline at various altitudes and flight Mach\nnumbers.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:07:56.539390+00:00"}
{"citation_id": "19930082703", "source_url": "https://ntrs.nasa.gov/api/citations/19930082703/downloads/19930082703.pdf", "page_number": 8, "total_pages": 28, "image_filename": "19930082703_p8.jpg", "text": "6\nNACA TN 1983\n\nHelicopter C\n\nA time history of a pull-and-hold maneuver for helicopter C is given in figure 8.\n\nPitching velocity.- The angular-velocity curve differs from that for helicopter A in a manner similar to the differences discussed in the comparison of helicopter B with helicopter A, except that the changes are more pronounced. That is, a greater downward concavity is shown during the first second after control deflection, and a more definite peak value is evident.\n\nNormal acceleration.- The normal-acceleration curve shows an initial jump, similar to those for helicopters A and B, followed by a comparatively short and much less definite pause (the slope never dropping all the way to zero as before). The time to the peak value is not appreciably different from that for helicopter B.\n\nPilot's comments.- The pilot's opinion of the pull-up characteristics of helicopter C was that they were satisfactory. The apprehension associated with the divergent tendency for helicopter A was absent and, in addition, the difficulty of anticipating the eventual result, which remained in helicopter B, was also absent. Normal flying was found to be correspondingly simpler. Furthermore, this helicopter could be flown for comparatively long periods in moderately rough air with the cyclic control stick held fixed by the friction clamp provided.\n\nThe control friction for this helicopter was moderate. No longitudinal stick-force gradients were apparent. Although the longitudinal characteristics in the pull-and-hold maneuver were considered relatively satisfactory without force gradients, the pilot believed that stable force gradients would be necessary for completely satisfactory pull-up characteristics. Stable force gradients are necessary for readily returning to trim conditions following maneuvers, as well as for assisting the pilot in judging and controlling the maneuvers.\n\nStick-fixed oscillations.- The longitudinal oscillations, following a disturbance at 80 miles per hour in level flight, were almost deadbeat for this configuration.", "timestamp": "2026-07-22T06:08:03.036664+00:00"}
{"citation_id": "19930082546", "source_url": "https://ntrs.nasa.gov/api/citations/19930082546/downloads/19930082546.pdf", "page_number": 19, "total_pages": 65, "image_filename": "19930082546_p19.jpg", "text": "18\nNACA TN No. 1870\n\nThe effect of tip Mach number on the oscillating pressures for a propeller operating at constant power may be estimated from the relation of $C_p$, $C_P$, and $M_t$ in the following manner. Since $C_p = \\frac{p}{\\rho n^2 D^2}$, $C_P = \\frac{P}{\\rho n^3 D^5}$, and $M_t = \\frac{\\pi n D}{c}$,\n\n$$ \\frac{p}{P} = \\frac{\\pi}{c} \\frac{C_p}{C_P M_t D^2} $$\n\nor\n\n$$ \\frac{p}{P / \\left( \\frac{\\pi D^2}{4} \\right)} = \\frac{\\pi^2}{4c} \\frac{C_p}{C_P M_t} $$\n\nThus in the charts of figure 19, lines of constant oscillating pressure per unit propeller power are straight radial lines through the origin. If the slope of the $C_p/C_P$ curve at a given point is greater than the slope of a straight line from that point to the origin as at point B in figure 19(c), the oscillating pressure will increase with an increase in tip Mach number for a constant power. If on the other hand the slope of the $C_p/C_P$ curve at a given point is less than the slope of the straight line to the origin as at point A in figure 19(c), the free-space pressure will decrease with increasing tip Mach number.\n\nIn general the charts of figure 19 show that at the low values of mB, the $C_p/C_P$ curves are relatively flat and the oscillating pressures will decrease with increasing tip Mach number at constant power. For the higher mB values the reverse is true. This effect has already been indicated in figure 16 and is further shown in figure 21 where the ratio $C_p/C_P M_t$, which is proportional to the oscillating pressures per unit propeller power, is plotted for various values of mB as a function of tip Mach number. Data in figure 21 is faired data taken from the charts of figure 19.\n\nFigure 21 shows that for values of mB less than 4 the oscillating pressure per unit power decreases with increased tip Mach number. The conclusion may be drawn that the pressure due to the fundamental mode of excitation for a four-blade propeller is essentially independent of tip Mach number when the power is held constant. Hence changing the tip Mach number will not materially affect the primary modes of fuselage vibration. It may be noted, however, that the large increase in pressure amplitude of the higher harmonics with increase in tip Mach number will greatly increase the noise levels in the fuselage.", "timestamp": "2026-07-22T06:08:03.637637+00:00"}
{"citation_id": "19930082566", "source_url": "https://ntrs.nasa.gov/api/citations/19930082566/downloads/19930082566.pdf", "page_number": 15, "total_pages": 44, "image_filename": "19930082566_p15.jpg", "text": "```markdown\nTABLE 3.- FATIGUE TEST DATA FOR STRESS RATIO $R = \\sigma_2/\\sigma_1 = 1.0$\n\n| Specimen | Load, $P^I$ (lb) | Load, $P^{III}$ (lb) | Pressure, $p^I$ (psi) | Pressure, $p^{III}$ (psi) | Diameter, d (in.) | Average wall thickness, t (in.) | Stress, $\\sigma_1^I$ (psi) | Stress, $\\sigma_2^I$ (psi) | Stress, $\\sigma_1^{III}$ (psi) | Stress, $\\sigma_2^{III}$ (psi) | Stress ratio, R $\\sigma_2'/\\sigma_1'$ | Number of cycles, N |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| F-10 | $4.00 \\times 10^3$ | $0.50 \\times 10^3$ | $1.200 \\times 10^3$ | $0.200 \\times 10^3$ | 2.00 | 0.0539 | $23.0 \\times 10^3$ | $22.5 \\times 10^3$ | $3.5 \\times 10^3$ | $3.5 \\times 10^3$ | 0.98 | $0.0966 \\times 10^6$ |\n| F-8 | 2.75 | .50 | .800 | .100 | 2.00 | .0540 | 15.5 | 15.0 | 2.5 | 2.0 | .97 | .1049 |\n| F-5 | 3.50 | 1.00 | 1.050 | .200 | 2.00 | .0545 | 20.0 | 19.5 | 4.5 | 3.5 | .98 | .1609 |\n| F-9 | 2.50 | .50 | .750 | .100 | 2.00 | .0497 | 15.5 | 15.0 | 2.5 | 2.0 | .97 | .1916 |\n| F-6 | 2.50 | .50 | .800 | .125 | 2.00 | .0534 | 15.0 | 15.0 | 2.5 | 2.5 | 1.00 | .2158 |\n| F-4 | 2.30 | .50 | .650 | .100 | 2.00 | .0545 | 12.5 | 12.0 | 2.5 | 2.0 | .96 | $1.8679$ |\n| E-6 | 1.60 | .60 | .575 | .125 | 2.00 | .0505 | 10.8 | 11.4 | 3.1 | 2.5 | 1.06 | 5.1044 |\n| E-7 | 2.40 | 1.58 | .750 | .150 | 2.00 | .0518 | 14.6 | 14.5 | 3.0 | 2.9 | .99 | 5.1362 |\n\nTABLE 4.- FATIGUE TEST DATA FOR STRESS RATIO $R = \\sigma_2/\\sigma_1 = 0.5$\n\n| Specimen | Load, $P^I$ (lb) | Load, $P^{III}$ (lb) | Pressure, $p^I$ (psi) | Pressure, $p^{III}$ (psi) | Diameter, d (in.) | Average wall thickness, t (in.) | Stress, $\\sigma_1^I$ (psi) | Stress, $\\sigma_2^I$ (psi) | Stress, $\\sigma_1^{III}$ (psi) | Stress, $\\sigma_2^{III}$ (psi) | Stress ratio, R $\\sigma_2'/\\sigma_1'$ | Number of cycles, N |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| F-11 | $7.75 \\times 10^3$ | $0.75 \\times 10^3$ | $0.850 \\times 10^3$ | $0.200 \\times 10^3$ | 2.00 | 0.0543 | $30.50 \\times 10^3$ | $15.50 \\times 10^3$ | $4.00 \\times 10^3$ | $3.50 \\times 10^3$ | 0.51 | $0.1185 \\times 10^6$ |\n| F-2 | 8.00 | .75 | .825 | .200 | 2.00 | .0534 | 32.50 | 15.50 | 4.00 | 3.50 | .48 | .1320 |\n| E-10 | 10.60 | 1.00 | 1.000 | .125 | 2.00 | .0545 | 40.12 | 18.35 | 4.07 | 2.30 | .46 | .1403 |\n| H-11 | 9.50 | .90 | .975 | .200 | 2.00 | .0527 | 38.00 | 18.50 | 4.50 | 4.00 | .49 | .1596 |\n| E-3 | 6.50 | 1.10 | .710 | .125 | 2.00 | .0523 | 26.56 | 13.58 | 4.44 | 2.20 | .51 | .2450 |\n| F-7 | 4.40 | .75 | .480 | .150 | 2.00 | .0539 | 17.44 | 8.91 | 3.60 | 2.78 | .51 | .2600 |\n| G-9 | 5.50 | 1.10 | .600 | .125 | 2.00 | .0542 | 21.66 | 11.07 | 4.38 | 2.30 | .51 | .2695 |\n| E-1 | 4.10 | .50 | .425 | .125 | 2.00 | .0513 | 16.86 | 8.29 | 2.77 | 2.44 | .49 | 5.1953 |\n\nNACA TN NO. 1889\n\nNACA\n\n13\n```", "timestamp": "2026-07-22T06:08:05.885537+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 38, "total_pages": 114, "image_filename": "19930086061_p38.jpg", "text": "```markdown\n34\n\n$\\alpha = 4.1^\\circ$\n$C_L = 0.15$\n\n$\\alpha = 8.1^\\circ$\n$C_L = 0.33$\n\n$\\alpha = 14.1^\\circ$\n$C_L = 0.60$\n\n$\\alpha = 24.1^\\circ$\n$C_L = 0.82$\n\n-4\n-3\n-2\n-1\nP\n0\n1\n\nStation 1\n$\\frac{y}{b/2} = 0$\n\n-3\n-2\n-1\nP\n0\n1\n\nUpper\nLower\nTwo dimensional\n(calculated at\nequal $C_L$)\n\nStation 2\n$\\frac{y}{b/2} = 0.167$\n\n-3\n-2\n-1\nP\n0\n1\n\nStation 3\n$\\frac{y}{b/2} = 0.333$\n\n0 2 4 6 8 10\nx/c\n\n0 2 4 6 8 10\nx/c\n\n0 2 4 6 8 10\nx/c\n\n0 2 4 6 8 10\nx/c\n\n(a) Stations: 1, 2, 3.\n\nFigure 6.- Chordwise pressure distribution about wing 1 at angles of attack of $4.1^\\circ$, $8.1^\\circ$, $14.1^\\circ$, and $24.1^\\circ$.\n\nNACA RM L9D07\n```", "timestamp": "2026-07-22T06:08:08.190730+00:00"}
{"citation_id": "19930082918", "source_url": "https://ntrs.nasa.gov/api/citations/19930082918/downloads/19930082918.pdf", "page_number": 6, "total_pages": 62, "image_filename": "19930082918_p6.jpg", "text": "NACA TN 1940\n5\n\nOptical-Microscope Studies\n\nAfter aging, the individual samples were ground on a cross-sectional face, polished with No. 1 emery cloth, Nos. 1, 1/0, 2/0, and 3/0 emery papers, transferred to a cloth disc, polished using a commercial chromium buffing compound, and finished on a Gamal wet wheel; they were then electrolytically etched for 5 seconds in 10 percent chromic acid at 1 ampere per square inch. After study under a microscope, representative photomicrographs were taken of each sample at 1000 diameters under slightly oblique illumination.\n\nElectron-Microscope Studies\n\nIn order to obtain electron micrographs from metallographic samples, replicas which will transmit the electron stream must be prepared of the surfaces. For the investigations reported herein, Formvar replicas were prepared according to the technique outlined in reference 2. The metallographic surfaces from which these replicas were prepared were the same polished and etched surfaces photographed optically. The replicas (shadow cast with chromium) were then mounted in an RCA Model B electron microscope and photographs taken at approximately 3000X. Enlargement to 8500X was done photographically.\n\nX-Ray Studies\n\nSample preparation.— A great deal of difficulty with variable and unreproducible diffraction data was initially encountered. The difficulty was found to be due to both mechanically disturbed metal surfaces and unrandom grain orientation. Eventually the development of the following technique produced results sufficiently free from these difficulties for this investigation:\n\n(1) A layer of metal 0.020 inch thick was electrolytically removed from the surface of all samples which were used for X-ray analysis. The amount of metal which had to be removed in order to get below the artificially strained surfaces produced by a cut-off wheel, by grinding, or as a result of metallographic polishing was determined by X-ray diffraction patterns of the type shown by figure 2 and made by the apparatus illustrated in figure 3. The patterns in figure 2, taken from a surface initially ground, show the emergence of the reflections of the individual grains and finally resolution of the $\\alpha_1\\alpha_2$ doublet of the molybdenum K radiation with increasing depth of metal removal. The amount of surface removal shown in figure 2 as necessary to obtain a strain-free surface was typical for all samples used.", "timestamp": "2026-07-22T06:08:08.601030+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 42, "total_pages": 98, "image_filename": "19930086073_p42.jpg", "text": "1.4\n1.2\n1.0\n.8\n.6\n.4\n.2\n0\n-.2\n\nLift coefficient, $C_L$\n\n| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | |", "timestamp": "2026-07-22T06:08:10.890399+00:00"}
{"citation_id": "19930085572", "source_url": "https://ntrs.nasa.gov/api/citations/19930085572/downloads/19930085572.pdf", "page_number": 3, "total_pages": 17, "image_filename": "19930085572_p3.jpg", "text": "NACA RM No. E5L02 RESTRICTED\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nRESEARCH MEMORANDUM\n\nCOMPARISON OF FLIGHT PERFORMANCE OF AN-F-58\nAND AN-F-32 FUELS IN J35 TURBOJET ENGINE\n\nBy Loren W. Acker and Kenneth S. Kleinknecht\n\nSUMMARY\n\nA flight investigation was conducted to determine the comparative performance of AN-F-58 and AN-F-32 fuels in a 4000-pound-thrust turbojet engine.\n\nThe results indicate that the performance of AN-F-58 fuel was equivalent to that of AN-F-32 fuel over the range of conditions investigated. The investigation of AN-F-58 fuel, compared with that of AN-F-32 fuel, indicated a 3-percent-higher net thrust and fuel consumption (same specific fuel consumption) at the high engine speeds; a slightly inferior blow-out limit (250 rpm higher); equally successful starts at altitudes between 5000 and 30,000 feet but somewhat longer acceleration time; and similarly small carbon deposits after $7\\frac{1}{2}$ hours of operation. These small differences, however, are attributable to the normal reproducibility of test conditions and the scatter of data for this type of investigation.\n\nINTRODUCTION\n\nThe need of the armed forces for a turbojet-engine fuel available in great quantities led to the development of the new specification fuel designated AN-F-58, which has much wider limits than the AN-F-32 specification fuel that is currently used.\n\nAs part of an extensive program undertaken at the NACA Lewis laboratory to investigate the performance of AN-F-58 fuel in several turbojet engines and single combustors from these engines, a flight investigation has been conducted to determine comparative performance of AN-F-58 and AN-F-32 fuels in a 4000-pound-thrust turbojet engine. Data presented compare the jet-engine performance parameters and operating characteristics of both fuels under similar operating conditions. The performance data were reduced to standard sea-level conditions using standard reduction parameters (reference 1).\n\nRESTRICTED", "timestamp": "2026-07-22T06:08:11.443075+00:00"}
{"citation_id": "19930082592", "source_url": "https://ntrs.nasa.gov/api/citations/19930082592/downloads/19930082592.pdf", "page_number": 15, "total_pages": 50, "image_filename": "19930082592_p15.jpg", "text": "14\nNACA TN 1914\n\n2. Hoffman, Charles A., Ault, G. Mervin, and Gangler, James J.:\nInitial Investigation of Carbide-Type Ceramal of 80-Percent\nTitanium Carbide Plus 20-Percent Cobalt for Use as Gas-Turbine-\nBlade Material. NACA TN 1836, 1949.\n\n3. Pilling, N. B., and Bedworth, R. E.: The Oxidation of Metals at\nHigh Temperatures. Jour. Inst. Metals (London), vol. XXIX,\nno. 1, 1923, pp. 529-582; discussion, pp. 583-584; correspondence,\npp. 584-591.\n\n4. Rhodin, T. N.: Thin Oxide Films on Tungsten and Thin Oxide Films\non Molybdenum. TP No. 2398, Inst. Metals Div., Metals Tech.,\nvol. 15, no. 4, June 1948, pp. 25-26. (Discussion of papers by\nE. A. Gulbransen and W. S. Wysong in Metals Tech., Sept. 1947.)\n\n5. Anon.: Metals Handbook, 1948 Edition. Am. Soc. Metals (Cleveland),\n1948, p. 223.\n\n6. Scheil, Erich: Über das Zundern von Metallen und Legierungen.\nZeitschr. f. Metallkunde, Jahrg. 29, Heft 7, Juli 1937, S. 209-\n214.\n\n7. Li, K. C., and Wang, Chung Yu: Tungsten. Reinhold Pub. Corp.\n(New York), 1943, p. 208.\n\n8. Dunn, John Stanley: The Oxidation of Tungsten: Evidence for the\nComplexity of Tungstic Oxide, $WO_3$. Jour. Chem. Soc. (London),\nPt. I, 1929, pp. 1149-1150.", "timestamp": "2026-07-22T06:08:24.832227+00:00"}
{"citation_id": "19930082703", "source_url": "https://ntrs.nasa.gov/api/citations/19930082703/downloads/19930082703.pdf", "page_number": 9, "total_pages": 28, "image_filename": "19930082703_p9.jpg", "text": "NACA TN 1983\n7\n\nDISCUSSION\n\nCharacteristics Appreciated\n\nFrom consideration of the results presented herein and those of\nreference 1, it is concluded that, for the helicopters represented, the\nmost important factor in the longitudinal characteristics in both pull-\nups and steady flight is whether or not a prolonged stick-fixed\ndivergence will occur. Further improvement is concluded to relate to\nthe continuous development of the normal acceleration in contrast with\na pause in development of acceleration during the first second following\nabrupt control deflection.\n\nIn order to arrive at these conclusions and to formulate tentative\nrequirements therefrom, it is necessary to show that normal flying is\nproperly represented by pull-ups and also that other characteristics\nwhich might be expected to be important can be considered subordinate\nto acceleration characteristics.\n\nFor all cases, the degree of pilot satisfaction with the character-\nistics in an abrupt pull-and-hold maneuver correlated with his satis-\nfaction with the normal-flying characteristics. The stipulation of\nsatisfactory pull-up characteristics must, of course, be taken as a\nnecessary rather than a sufficient condition; for example, if (at 80 mph)\nhelicopter C had exhibited an unstable variation of stick position with\nspeed, it would not have been considered satisfactory in normal flight\nregardless of pull-up characteristics. As another example, the stick-\nforce characteristics were actually considered to be in need of improve-\nment.\n\nAs was discussed at length in reference 1, improvement in stick\nforces or provision of stick-free stability does not appear to be the\nprimary need for these helicopters, although the desirability of good\nstick-force characteristics cannot be too strongly emphasized. Even\nwith complete stick-fixed stability, the forces should be such that the\nstick will tend to return to the original trim position when deflected.\nFurthermore, the greater the stick-fixed instability, the greater will\nbe the improvement achieved by incorporating stick-free stability,\ninasmuch as it can partially mask the difficulties imposed by the stick-\nfixed instability.\n\nAnother alternate possibility requiring discussion is the use of\npitching velocity rather than normal-acceleration characteristics as a\ncriterion. For the cases under consideration, improvements in one of\nthese characteristics are accompanied by improvements in the other, but\nlogical pitching-velocity characteristics are reached more readily than\nlogical normal-acceleration characteristics and are not sufficient as a", "timestamp": "2026-07-22T06:08:26.265750+00:00"}
{"citation_id": "19930082613", "source_url": "https://ntrs.nasa.gov/api/citations/19930082613/downloads/19930082613.pdf", "page_number": 13, "total_pages": 46, "image_filename": "19930082613_p13.jpg", "text": "12\nNACA TN 1938\n\nSUMMARY OF RESULTS\n\nThe investigation of turbojet combustion-chamber liners to determine the factors contributing to failure by cracking yielded the following results:\n\n1. Most cracks originated at the inner portion of the stress-relieving holes of the louvers in the upper bend of the louver flaps. They were propagated at first almost perpendicularly from the edge of the hole but then turned either parallel to or away from the center line of the louver.\n\n2. Buckling was found at or near almost every crack and usually extended from the stress-relieving holes of the louver to the nearest air-intake hole.\n\n3. Some of the cracks probably began in an intercrystalline manner; others are believed to originate in fissures produced by punching operations. As the cracks enlarged, they tended to become more and more transcrystalline. Most of the cracks were partly intercrystalline or had intercrystalline branches.\n\n4. Surfaces exposed to hot gases were covered with scale. Subsurface scale (internal oxidation) occurred at many of these surfaces, particularly the edges of cracks.\n\n5. Neither carburization nor decarburization was detected.\n\n6. No conclusions or correlations in regard to the mechanisms of cracking could be drawn from the grain-size measurements, the presence of intermetallic compounds, or solid nonmetallic inclusions.\n\n7. Carbonaceous deposits absorbed a relatively large percentage (0.3 percent) of sulfur from combustion gases. The sulfur picked up by these deposits, however, did not contribute appreciably to the failure by cracking.\n\n8. Heat treatments investigated were ineffectual in preventing cracking.\n\nCONCLUDING REMARKS\n\nThe following observations may be made from this investigation:\n\n1. The most common types of crack that extend from the stress-relieving holes of the louvers probably are caused by thermal and mechanical fatigue of the louver flaps.", "timestamp": "2026-07-22T06:08:27.480101+00:00"}
{"citation_id": "19930082566", "source_url": "https://ntrs.nasa.gov/api/citations/19930082566/downloads/19930082566.pdf", "page_number": 16, "total_pages": 44, "image_filename": "19930082566_p16.jpg", "text": "14\nNACA TN No. 1889\n\n<!-- Image (106, 120, 799, 375) -->\n\nFigure 1.- Nomenclature for fluctuating stresses.\n\n<!-- Image (106, 524, 761, 862) -->\n\nFigure 2.- Typical S-N diagram.", "timestamp": "2026-07-22T06:08:29.611978+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 39, "total_pages": 114, "image_filename": "19930086061_p39.jpg", "text": "NACA RM L9J07\n\n$\\alpha=4.1^\\circ$\n$C_L=0.15$\n\n$\\alpha=8.1^\\circ$\n$C_L=0.33$\n\n$\\alpha=14.1^\\circ$\n$C_L=0.60$\n\n$\\alpha=24.1^\\circ$\n$C_L=0.82$\n\n-3\n-2\nP -1\n0\n1\n\n-3\n-2\nP -1\n0\n1\n\n-2\nP -1\n0\n1\n\n-2\nP -1\n0\n1\n\nStation 4\n$\\frac{y}{b/2}, 0.500$\n\nStation 5\n$\\frac{y}{b/2}, 0.667$\n\nStation 6\n$\\frac{y}{b/2}, 0.833$\n\nStation 7\n$\\frac{y}{b/2}, 0.916$\n\nUpper\nLower\nTwo dimensional\n(calculated at\nequal $C_l$)\n\n0 2 4 6 8 10\nx/c\n\n0 2 4 6 8 10\nx/c\n\n0 2 4 6 8 10\nx/c\n\n0 2 4 6 8 10\nx/c\n\n(b) Stations 4,5,6,7.\n\nFigure 6.- Concluded.\n\nNACA\n\n35", "timestamp": "2026-07-22T06:08:36.841286+00:00"}
{"citation_id": "19930085485", "source_url": "https://ntrs.nasa.gov/api/citations/19930085485/downloads/19930085485.pdf", "page_number": 8, "total_pages": 26, "image_filename": "19930085485_p8.jpg", "text": "6 CONFIDENTIAL NACA RM No. I8K02\n\nIt can be seen from the preceding table that aileron contour had a considerable effect on the aerodynamic-center location of the model at subsonic Mach numbers. Thickening the aileron trailing edge shifted the aerodynamic center rearward by approximately 2 to 10 percent at a subcritical Mach number of approximately 0.60 at lift coefficients of 0.20 and -0.2. The jog in the pitching-moment-coefficient curve near zero lift for the circular-arc contour aileron (fig. 10) is equivalent to about a 25-percent forward shift in the aerodynamic-center location. This forward shift of the aerodynamic-center location for the flat-sided aileron (fig. 11) near zero lift is approximately 11 percent. These jogs may have been accentuated by the relatively low test Reynolds numbers. However, inasmuch as this forward shift in aerodynamic center occurs within a very small lift range (about 0.05C_L), it is thought that it will result in no serious stability or control problems.\n\nAbove the critical Mach number for this model there is a considerable rearward shift of the aerodynamic-center position. However, the limited amount of data obtained does not justify any conclusions concerning the effect of aileron contour.\n\n### Aileron Control Characteristics\n\nThe results of the investigation of the aileron control characteristics for the various aileron contours are shown in figures 15 to 19. The rolling-moment coefficients of the aileron with circular-arc (t = 0) and flat-sided (t = 1.00) contour presented in reference 1 are reproduced in figures 15 and 16 for comparison with additional data on the aileron with other trailing-edge thicknesses (t = 0.37 and 0.50) in figures 17 and 18.\n\nFrom the rolling-moment-coefficient data presented in figure 17 it appears that the aileron with a value of t = 0.50 will give control throughout the Mach number range tested. However, at an angle of attack of 3.1° (fig. 17(b)) the aileron effectiveness is very low. The rolling-moment characteristics of the aileron with a value of t = 0.37 (fig. 18) appear to be less satisfactory than either of the other ailerons with thickened trailing edges because of very low effectiveness and reversal for δ_a = -5° and α = 3.5° at a Mach number of 1.15 (fig. 18(b)).\n\nA summary of the effects of aileron profile on the rolling-moment coefficients is presented in figure 19. The aileron with thick trailing edges gave greater values of ∂C_l_a/∂δ_a than the circular-arc aileron. The aileron with a ratio of trailing-edge thickness to hinge-line thickness of 0.50 shows the most linear variation of C_l_a with δ_a of the four aileron-contour configurations investigated and, in addition, the effectiveness (∂C_l_a/∂δ_a) is very near the maximum value obtained for any of the aileron contours investigated.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:08:37.042510+00:00"}