PDF/ocr_dataset_1019.jsonl

{"citation_id": "19930082496", "source_url": "https://ntrs.nasa.gov/api/citations/19930082496/downloads/19930082496.pdf", "page_number": 12, "total_pages": 50, "image_filename": "19930082496_p12.jpg", "text": "NACA TN No. 1836\n\nextensive, as indicated in figure 9. Apparently, a difference in coefficients of expansion of the film and of the base material together with the more severe thermal gradients at the higher temperatures cracked the film and exposed new surfaces to oxidation, resulting in this appearance. The film thickness at this time was determined as approximately 0.027 inch.\n\nOn the basis of this evaluation, it was concluded that the thermal-shock resistance of this ceramal is excellent, relative to currently available ceramics.\n\n### Quasi-Service Turbine-Blade Evaluation\n\n**Phase 1.** - The results obtained in phase 1 are presented in table III, which shows that during this phase of the evaluation, one ceramal blade failed after 10 minutes of operation at 17,500 rpm and an indicated inlet gas temperature of $2000^\\circ$ F. The total time of operation was 4 hours and 40 minutes under the various conditions noted in table III. The remaining two ceramal blades used in phase 1 fractured within an additional 5 minutes of operation at the same conditions. Examination of the failures indicated that the blades fractured in the neck-roll junction at a point of possible stress concentration caused by the presence of both peening and an abrupt change in root configuration (fig. 10). During overhaul, after failure of the first ceramal blade, minute cracking was noted on the back side of one of the other ceramal blades at the neck-roll junction. This occurrence might reasonably have been the result of stress concentration. These three failures may also have been largely the result of\n\n(1) operation at a shaft speed (17,500 rpm) that may have caused resonant vibration of the blade to occur\n\n(2) peening stresses, in addition to the stresses existing at operating conditions at time of fracture, exceeding the strength of the material.\n\n**Phase 2.** - The roots of a second set of three ceramal blades were copper-plated and then so installed as to permit redistribution of load and minimizing of stress-concentration effects. Copper was selected as a plating in order to preserve the full advantage of the excellent thermal conductivity of the ceramal ($20.56 \\text{ Btu/(hr)(sq ft)}$) ($^\\circ\\text{F/ft}$) at room temperature, reference 2). The metal blades had a substantially lower thermal conductivity ($8.38 \\text{ Btu/(hr)(sq ft)}$) ($^\\circ\\text{F/ft}$) at $392^\\circ$ F). The high thermal conductivity of the ceramal blade material was expected to result in lower operating temperatures for these blades and would increase operating temperatures of the disk rim by heat conduction to the disk. Of the 139 metal blades, 127 were original blades from phase 1, and 12 were new blades installed as control blades. The operation program was modified to eliminate wheel operation at or near 17,500 rpm.", "timestamp": "2026-07-22T04:47:01.164566+00:00"}
{"citation_id": "19930082542", "source_url": "https://ntrs.nasa.gov/api/citations/19930082542/downloads/19930082542.pdf", "page_number": 6, "total_pages": 53, "image_filename": "19930082542_p6.jpg", "text": "NACA TN No. 1867\n\n(2) Solution-treated stock\n\n(a) Temperature of solution-treating  \n(b) Time of solution treatment and cooling rate  \n(c) Aging temperature  \n(d) Aging time  \n(e) Hot-cold-work after various solution treatments  \n(f) Temperature of hot-cold-working  \n(g) Amount of hot-cold-work  \n(h) Aging prior to hot-cold-work  \n(i) Aging after hot-cold-work  \n(j) Effect of holding time at $1200^\\circ$ F before testing  \n\nSolution treatments were varied from $1800^\\circ$ to $2300^\\circ$ F, aging treatments from $1350^\\circ$ to $1750^\\circ$ F, and hot-cold-working from room temperature to $1800^\\circ$ F. A diagram which shows all the 67 treatments used is given in figure 2.\n\nThe individual treatments were made on bars $6\\frac{1}{2}$ inches long cut from 7/8-inch-square bar stock. Solution treatments were carried out in a gas-fired furnace. Aging was done in an electric resistance furnace. Hot-cold-working was accomplished by rolling in a 5-inch, two-high rolling mill.\n\nThe general procedure for rolling was as follows: The bar was heated for 1 hour at a temperature $20^\\circ$ F higher than the desired rolling temperature. In two passes the bar was reduced 10 percent. Reductions of 5 and 10 percent were made with no reheat, 15 and 20 percent with one reheat, and 25 percent with two reheats. A reheat consisted in placing the partially rolled bar in the furnace to bring it back to $20^\\circ$ F above rolling temperature. All hot-cold-rolled bars were given a final stress relief of 1 hour at $1200^\\circ$ F.\n\nThe treated bars were sectioned to give a $5\\frac{1}{2}$-inch-long by 0.505-inch-diameter tensile specimen from one end, four $2\\frac{1}{2}$-inch-long by 0.160-inch-diameter rupture specimens from the other end, and hardness and metallographic specimens from the small piece of stock remaining.\n\nThe room-temperature tensile tests were conducted in a 60,000-pound hydraulic testing machine. The modified Martin type extensometer system used had a sensitivity of 0.000003-inch per inch.\n\nRupture tests were run in individual stationary units applying the stress through a simple-beam and knife-edge system. Approximately 24 hours was allowed for temperature adjustments prior to application of the stress. Only the minimum number of tests needed to indicate the 100- and 1000-hour rupture strengths were run. The effect of time at temperature prior to loading on rupture properties was studied by", "timestamp": "2026-07-22T04:47:03.293204+00:00"}
{"citation_id": "19930093773", "source_url": "https://ntrs.nasa.gov/api/citations/19930093773/downloads/19930093773.pdf", "page_number": 25, "total_pages": 47, "image_filename": "19930093773_p25.jpg", "text": "24\nNACA RM E9G09\n\nAltitude\n(ft)\nO 5,000\n□ 15,000\n◇ 25,000\n△ 35,000\n▽ 45,000\n◁ 50,000\n\nAir flow, $W_a$, lb/sec\n\nEngine speed, N, rpm\n\n(b) Air flow.\n\nFigure 4. - Continued. Effect of altitude on variation of engine performance with engine speed at flight Mach number of 0.21.", "timestamp": "2026-07-22T04:47:04.879221+00:00"}
{"citation_id": "19930082447", "source_url": "https://ntrs.nasa.gov/api/citations/19930082447/downloads/19930082447.pdf", "page_number": 22, "total_pages": 24, "image_filename": "19930082447_p22.jpg", "text": "20\n\nVertical-velocity ratio, $\\frac{\\dot{y}}{\\dot{y}_0}$\n\nAt maximum acceleration\n\nAt rebound\n\nApproach parameter, $\\xi$\n\nFigure 6.-Variation of vertical velocity with approach parameter.\n\nNACA\n\nNACA TN NO. 1775", "timestamp": "2026-07-22T04:47:06.032928+00:00"}
{"citation_id": "19930085965", "source_url": "https://ntrs.nasa.gov/api/citations/19930085965/downloads/19930085965.pdf", "page_number": 62, "total_pages": 67, "image_filename": "19930085965_p62.jpg", "text": "1125\n\nNACA RM E9E06\n\n[Figure: Oscillogram (a) showing a waveform with 99 root-mean-square ampere turns.]\n\n(a) 99 root-mean-square ampere turns.\n\n[Figure: Oscillogram (b) showing a waveform with 89 root-mean-square ampere turns; 139 direct-current ampere turns.]\n\n(b) 89 root-mean-square ampere turns; 139 direct-current ampere turns.\n\nNACA\nC-23405\n5-6-49\n\nFigure 17. - Search-coil voltage oscillograms for comparing wave shapes with and without superimposed direct-current ampere turns. Armco Magnetic Ingot Iron; air gap, 0.04 inch.\n\n61", "timestamp": "2026-07-22T04:47:13.190367+00:00"}
{"citation_id": "19930082498", "source_url": "https://ntrs.nasa.gov/api/citations/19930082498/downloads/19930082498.pdf", "page_number": 11, "total_pages": 49, "image_filename": "19930082498_p11.jpg", "text": "```markdown\n10\nNACA TN No. 1838\n\nmuffler. In addition, stiffeners were added to the flat sides of the muffler. The vibration of the sides was so severe, however, that the stiffeners broke loose during the test. Mufflers of this type, with large flat sides, were therefore discarded as impractical.\n\nExpansion-chamber types.- Several expansion-chamber types of mufflers were investigated and some were very effective, although in general the back pressures were higher than for the straight-through type of mufflers due to energy losses in the expansion process. Not only with these expansion-chamber but also with the resonant-chamber types, the use of an oval cross section instead of a circular cross section for the outside shell has no measurable effect on the over-all sound-pressure level, provided the cross-sectional area is the same in both cases. This result is a check with the theory, which indicates that if only plane pressure waves need be considered the attenuation will not be a function of the cross-sectional shape of an expansion chamber. An oval shape is sometimes preferred because of the particular space requirements in a specific installation. Tests of a 3-inch by 12-inch by 24-inch oval muffler with two different tail-pipe lengths (configurations 64 and 65, table II) show how the effectiveness of such a muffler can be increased by the proper choice of tail-pipe length. The over-all sound-pressure level is reduced from 91.6 to 87 decibels at 2000 rpm by making the tail pipe long enough to bring the resonant frequency of the muffler and tail-pipe combination, considered as a Helmholtz resonator, well below the lowest frequency present in the exhaust. The longer tail pipe reduces the intensity of the fundamental frequency component at 2000 rpm from 91 to 80 decibels.\n\nEven with the proper choice of tail-pipe configuration, the over-all sound-pressure level is 5.5 to 6.5 decibels higher for muffler 65 than for muffler 62, which has a chamber volume about four times higher than that of muffler 65. Thus, for expansion-chamber mufflers, as was also discovered in the case of resonant-chamber mufflers, large chamber volumes are required for the reduction of low-frequency noise.\n\nCombinations.- A muffler consisting of a single expansion chamber and a single resonant chamber in combination as a unit (muffler 67, table II) produced excellent attenuation with a reasonable value of back pressure, yet it was only 30 inches long, which is considerably shorter than other mufflers of similar performance. This result shows that it is possible to build a much smaller muffler than the one used in the flight demonstration with little sacrifice in performance. Inasmuch as no concerted effort was made to obtain a muffler of minimum size, it is quite possible that a muffler of equal performance even smaller than muffler 67 could be designed for the same speed range. Another combination consisting of two of the best mufflers (mufflers 62 and 45) placed in series was tested (fig. 2(e) and muffler 68, table II). The resultant sound-pressure levels are regarded as the lowest which can be obtained with this engine by means of exhaust muffling alone. For example, the over-all sound-pressure level at 2000 rpm is reduced to 79.5 decibels. The remaining sound is believed to consist mostly of engine air intake and clatter noises.\n```", "timestamp": "2026-07-22T04:47:14.140673+00:00"}
{"citation_id": "19930085838", "source_url": "https://ntrs.nasa.gov/api/citations/19930085838/downloads/19930085838.pdf", "page_number": 108, "total_pages": 118, "image_filename": "19930085838_p108.jpg", "text": "106\nNACA RM No. L9B23\n\n<!-- Image (145, 110, 823, 929) -->\n\nFigure 16.- Variation of section lift characteristics with deflection of straight-sided Frise aileron on two NACA 7-series-type airfoils with double slotted flap and flap. $\\alpha_0 = 0^\\circ$; $\\delta_f = 0^\\circ$; $R = 6.0 \\times 10^6$ (approx.).", "timestamp": "2026-07-22T04:47:21.050065+00:00"}
{"citation_id": "19930082245", "source_url": "https://ntrs.nasa.gov/api/citations/19930082245/downloads/19930082245.pdf", "page_number": 29, "total_pages": 66, "image_filename": "19930082245_p29.jpg", "text": "28\n\n| Section angle of attack, $\\alpha$, deg | $\\delta_a$ (deg) |\n| :--- | :--- |\n| 8 | -12 |\n| 6 | -6 |\n| 4 | -4 |\n| 2 | -2 |\n| 0 | 0 |\n| -2 | 2 |\n| -4 | 4 |\n| -6 | 12 |\n| -8 | 18 |\n| -10 | 90 |\n| -12 | |\n| -14 | |\n\nMach number, M\n.1 .2 .3 .4 .5 .6 .7 .8 .9\n\n| Section pitching-moment coefficient, $c_m$ | $\\delta_a$ (deg) |\n| :--- | :--- |\n| .16 | -12 |\n| .12 | -6 |\n| .08 | -4 |\n| .04 | -2 |\n| 0 | 0 |\n| -.04 | 2 |\n| -.08 | 4 |\n| -.12 | 12 |\n| -.16 | 18 |\n| -.20 | |\n| -.24 | 30 |\n| -.28 | |\n\nMach number, M\n.1 .2 .3 .4 .5 .6 .7 .8 .9\n\n(d) $c_n=0.2$.\nFigure 6.—Continued.\n\nNACA TN NO. 1596", "timestamp": "2026-07-22T04:47:23.082657+00:00"}
{"citation_id": "19930082617", "source_url": "https://ntrs.nasa.gov/api/citations/19930082617/downloads/19930082617.pdf", "page_number": 4, "total_pages": 58, "image_filename": "19930082617_p4.jpg", "text": "NACA TN 1962\n3\n\nThe test rig and the attachment of the cylinder to it were very much the same as those used in the tests described in reference 1. Differences worth mentioning are a heavy stiffening grid of steel channels added to the end stand and a lever arrangement operated by a mechanical jack which permitted the application of higher loads than the previous interlinked frame system. In addition a calibrated load link was inserted between the loading head and the cable supporting the counterweight for it. Throughout the test, the tension in the cable was kept constant by adding or removing weights from the counterweight pan. At each increase in the jack load the tension in the counterweight cable was checked.\n\nThe load was measured by means of pairs of Baldwin Southwark SR-4 type A-1 electric strain gages cemented to opposite sides of a calibrated load link. Either type A-1 or A-11 SR-4 electric strain gages were used to measure the strain in the stringers.\n\nEach cylinder had three bands of gages with the exception of cylinder 77 which had four bands. A band is a series of gages arranged in a cross section of the cylinder around the entire circumference. In all specimens the second field from the loading head and a field in the center of the specimen had gages on every stringer. On the second field from the fixed end, gages were placed on every other stringer. With cylinder 77 the fourth band of gages was placed at 25 percent of the length of the specimen from the loading end. The position of the strain gages is shown in figure 1. The error in the strain-gage readings is believed to be less than a strain of $\\pm 10 \\times 10^{-6}$.\n\nFrom five to seven load increments of approximately 500 to 1000 pounds were applied to each test specimen. At each stage of loading, readings of all gages were taken and then checked by a second observer. This loading procedure was continued up to the buckling load, which was characterized by a sudden drop of the applied moment and a corresponding jump in the distorted shape. An exception to this type of failure was found with cylinder 74, in which a gradual rather than a sudden decrease of the applied load was encountered.\n\nPRESENTATION OF TEST RESULTS\n\nResults of the strain measurements in the stringers are presented for one end band near each end and for one band in the middle of each specimen. In the case of cylinder 77, measurements made in an additional intermediate band are also shown. The presentation is in the form of diagrams in which the strain is plotted against the distance of the stringer from the horizontal diameter of the cylinder. These basic data are contained in figures 2 to 26.", "timestamp": "2026-07-22T04:47:23.915053+00:00"}
{"citation_id": "19930085842", "source_url": "https://ntrs.nasa.gov/api/citations/19930085842/downloads/19930085842.pdf", "page_number": 96, "total_pages": 104, "image_filename": "19930085842_p96.jpg", "text": "92\nNACA RM L9C29\n\n[Figure: A graph plotting three curves against a common x-axis. The top curve is labeled $V/nD$ and corresponds to the left y-axis labeled \"Propeller advance-diameter ratio, $V/nD$\". The middle curve is labeled $C_{DR}$ and corresponds to the right y-axis labeled \"Resultant-drag coefficient, $C_{DR}$\". The bottom curve is labeled $C_L$ and corresponds to the left y-axis labeled \"Lift coefficient, $C_L$\". The common x-axis is labeled \"Torque coefficient, $Q_c$\". A box in the lower right corner of the graph contains the text \"NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\".]\n\n(f) $\\alpha_u = 44^\\circ$.\nFigure 44.- Concluded.", "timestamp": "2026-07-22T04:47:25.054688+00:00"}
{"citation_id": "19930082614", "source_url": "https://ntrs.nasa.gov/api/citations/19930082614/downloads/19930082614.pdf", "page_number": 5, "total_pages": 36, "image_filename": "19930082614_p5.jpg", "text": "NACA TN 1939\n3\n\nV true airspeed, feet per second\n$V_v$ vertical component of speed, feet per minute\n$\\Delta V$ speed change in time interval $\\Delta t$, feet per second\n$\\bar{V}$ average speed during time interval $\\Delta t$, feet per second\nW airplane weight, pounds\nW/S wing loading, pounds per square foot\na longitudinal acceleration, feet per second squared\n$\\bar{a}$ average longitudinal acceleration during time interval $\\Delta t$, feet per second squared\nb wing span, feet\ng acceleration due to gravity, feet per second squared\nh altitude, feet\n$\\Delta h$ altitude change in time interval $\\Delta t$, feet\nn indicated normal acceleration factor $\\left( \\frac{\\text{indicated acceleration normal to flight path}}{g} \\right)$\nq dynamic pressure $\\left( \\frac{1}{2}\\rho V^2 \\right)$, pounds per square foot\nr radius of curvature of flight path, feet\nt time, seconds\n$\\Delta t$ time increment, seconds\n$\\alpha$ angle of attack, degrees\n$\\gamma$ flight-path angle from the horizontal (positive for a climb), degrees\n$\\Delta \\gamma$ change in flight-path angle in the time interval $\\Delta t$, degrees\n$\\bar{\\gamma}$ average flight-path angle during the time interval $\\Delta t$, degrees\n$\\rho$ air density, slugs per cubic foot", "timestamp": "2026-07-22T04:47:34.282428+00:00"}
{"citation_id": "19930082585", "source_url": "https://ntrs.nasa.gov/api/citations/19930082585/downloads/19930082585.pdf", "page_number": 6, "total_pages": 30, "image_filename": "19930082585_p6.jpg", "text": "NACA TN 1907\n\nANALYSIS\n\nIt is assumed that the initial steady state of hovering flight is entirely determined, and the physical quantities of this hovering condition are denoted by the subscript 0. In the final state of steady autorotative descent, the subscript f is used to denote the physical quantities.\n\nThe simultaneous differential equations governing the transition period are six in number:\n\n(1) For the vertical acceleration of the craft, with the rotor thrust a function of $\\Omega$, $V$, $v$, $\\beta$, $\\dot{\\beta}$, $\\theta$, and so forth.\n\n(2) For the angular acceleration of the rotor, with the rotor torque a function of $\\Omega$, $V$, $v$, and so forth.\n\n(3) For the flapping motion of the blades.\n\n(4) For the hunting motion of the blades.\n\n(5) For the induced velocity $v$ as a function of $T$, $V$, $\\dot{V}$, and so forth (unsteady flow).\n\n(6) For the angle of incidence $\\theta$ as a function of the flapping and lag angles.\n\nThere is as yet neither theory nor empirical data on which to base the fifth of the above equations. For convenience, it is assumed that the induced velocity $v$ varies with time after power failure according to\n\n$$\nv = v_f + (v_0 - v_f)e^{-kt} \\tag{1}\n$$\n\nwhere the value of the decay coefficient $k$ is arbitrary and may be assigned different values in order to investigate its importance. It may be anticipated that the effect of the approximation represented by equation (1) is small, and that the value of $k$ chosen is not critical, when the initial and final values of induced velocity are not widely different, as in the usual case. The induced velocity is considered to be constant over the rotor disk. While exponential variation of induced velocity (equation (1)) represents a convenient assumption, it is not unlikely that the actual variation would be somewhat different. Verification of the supposition that the effect is small, must, failing a theoretical approach, await experimental evidence.", "timestamp": "2026-07-22T04:47:41.017026+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 14, "total_pages": 99, "image_filename": "19930082511_p14.jpg", "text": "12\nNACA TN No. 1826\n\nsymmetry, the entrance-lip condition may be satisfied without an additional electrode. If the lifting element is not in the horizontal plane of symmetry, an upstream electrode will be needed, with the potential of each strip adjusted relative to this electrode. For the unsymmetrical closed-open-closed analogy, the requirement that upstream and downstream velocities be equal necessitates further measurements with a setup corresponding to that of figure 8(c).\n\nAcceleration-Potential Analogies\n\nBasic concepts of the analogies.- The pressure has the properties of a potential - designated acceleration potential - in a field consisting of a small perturbation flow superposed on a uniform stream. If the pressure in the undisturbed stream is taken as zero, then the perturbation velocities are related to the pressure by the following equations:\n\n$$u = \\frac{1}{\\rho} \\int_{-\\infty}^{t} - \\frac{\\partial p}{\\partial x} dt = - \\frac{1}{\\rho U} \\int_{-\\infty}^{x} \\frac{\\partial p}{\\partial x} dx = - \\frac{p}{\\rho U}$$\n\n$$v = \\frac{1}{\\rho} \\int_{-\\infty}^{t} - \\frac{\\partial p}{\\partial y} dt = - \\frac{1}{\\rho U} \\int_{-\\infty}^{x} \\frac{\\partial p}{\\partial y} dx$$\n\n$$w = \\frac{1}{\\rho} \\int_{-\\infty}^{t} - \\frac{\\partial p}{\\partial z} dt = - \\frac{1}{\\rho U} \\int_{-\\infty}^{x} \\frac{\\partial p}{\\partial z} dx$$\n\nwhere\n\n| | |\n| :--- | :--- |\n| $\\rho$ | density |\n| $t$ | time |\n| $p$ | pressure |\n\nSince, by the first equation, $u$ is proportional to $p$, it is simpler merely to consider the perturbation velocity $u$ itself as the potential, with $v$ and $w$ given by the following equations:\n\n$$v = \\int_{-\\infty}^{x} \\frac{\\partial u}{\\partial y} dx$$\n\nand\n\n$$w = \\int_{-\\infty}^{x} \\frac{\\partial u}{\\partial z} dx$$", "timestamp": "2026-07-22T04:47:42.690202+00:00"}
{"citation_id": "19930082485", "source_url": "https://ntrs.nasa.gov/api/citations/19930082485/downloads/19930082485.pdf", "page_number": 22, "total_pages": 62, "image_filename": "19930082485_p22.jpg", "text": "NACA TN No. 1810\n\n$\\sqrt{\\frac{Z_{1}}{Z_{m}}} = \\exp \\left[ -\\frac{n_{o}}{2 \\Delta C} \\left( c_{1}^{2} - c_{m}^{2} \\right) \\right]$\n\n$= \\exp \\left\\{ -\\frac{n_{o}}{2} \\frac{c_{1}}{\\Delta C} \\left[ 1 - \\left( \\frac{c_{m}}{c_{1}} \\right)^{2} \\right] \\right\\}$\n\n$= \\exp \\left[ \\frac{n_{o}}{2} \\frac{c_{1}}{\\Delta C} \\left( 1 + \\frac{\\Delta C}{4 c_{1}} \\right) \\right]$\n\n$= \\exp \\left[ a \\left( 1 - \\frac{1}{4 b} \\right) \\right]$\n\nand similarly,\n\n$\\sqrt{\\frac{Z_{2}}{Z_{m}}} = \\exp \\left[ -a \\left( 1 - \\frac{3}{4 b} \\right) \\right]$\n\nThe velocities at the channel wall, that is, at the ends of the velocity-potential line considered, are computed from the value of $Z_{m}$ by use of equations (39) and (40).\n\nThe simple stream-filament theory presented may be used to compute the velocity distribution on the surface of any channel to which the assumptions used apply.", "timestamp": "2026-07-22T04:47:43.970364+00:00"}
{"citation_id": "19930090382", "source_url": "https://ntrs.nasa.gov/api/citations/19930090382/downloads/19930090382.pdf", "page_number": 36, "total_pages": 37, "image_filename": "19930090382_p36.jpg", "text": "38\nNACA RM L9I07\n\nMaximum efficiency, $\\eta_{max}$\nAdvance ratio, J\n\nCONFIDENTIAL\nCONFIDENTIAL\n\n| | | | | | | | | | | | | | | | | |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| | | | | | | | | | | | | | | | | |\n| | | | | | | | | | | | | | | | | |\n| | | | | | | | | | | | | | | | | |\n| | | | | | | | | | | | | | | | | |\n| | | | | | | | | | | | | | | | | |\n| | | | | | | | | | | | | | | | | |\n| | | | | | | | | | | | | | | | | |\n| | | | | | | | | | | | | | | | | |\n| | | | | | | | | | | | | | | | | |\n| | | | | | | | | | | | | | | | | |\n| | | | | | | | | | | | | | | | | |\n| | | | | | | | | | | | | | | | | |\n| | | | | | | | | | | | | | | | | |\n| | | | | | | | | | | | | | | | | |\n| | | | 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{"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 3, "total_pages": 78, "image_filename": "19930082618_p3.jpg", "text": "TECH LIBRARY KAFB, NM\n0065341\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nTECHNICAL NOTE 1945\n\nAERODYNAMIC CHARACTERISTICS OF 15 NACA AIRFOIL\nSECTIONS AT SEVEN REYNOLDS NUMBERS\nFROM $0.7 \\times 10^6$ TO $9.0 \\times 10^6$\n\nBy Laurence K. Loftin, Jr. and Hamilton A. Smith\n\nSUMMARY\n\nAn investigation has been made of the two-dimensional aerodynamic characteristics of 15 NACA airfoils at four Reynolds numbers from $2.0 \\times 10^6$ to $0.7 \\times 10^6$. These data, together with those from previous NACA papers for the same airfoils at three Reynolds numbers from $3.0 \\times 10^6$ to $9.0 \\times 10^6$, are presented and analyzed in the present paper. The airfoils investigated consisted of 10 systematically varied NACA 6-series airfoils and 5 airfoils of the NACA 4- and 5-digit series. The NACA 6-series airfoils had thickness ratios varying from 9 to 18 percent of the chord, design lift coefficients varying from 0 to 0.6, and positions of minimum pressure on the basic thickness form at zero lift varying from 30 to 60 percent of the chord. The NACA 4- and 5-digit-series sections investigated consisted of the NACA 0012, and the NACA 44- and 230-series sections of 12-percent and 15-percent thickness. The tests were made for both smooth and rough surface conditions and also included the determination of the effectiveness of the different airfoils at various Reynolds numbers when equipped with split flaps.\n\nThe results of the investigation indicate that the drag coefficient at the design lift coefficient and the maximum lift coefficient are the important aerodynamic characteristics which are most affected by variations in the Reynolds number between $9.0 \\times 10^6$ and $0.7 \\times 10^6$. For each of the 15 airfoils in both the smooth and rough surface conditions, the drag coefficient at design lift increased as the Reynolds number was lowered from $9.0 \\times 10^6$ to $0.7 \\times 10^6$. For the smooth NACA 6-series airfoils the magnitude of this increase became larger with increasing airfoil thickness and with rearward movement of the position of minimum pressure on the basic thickness form at zero lift. In the rough surface condition and at the lower Reynolds numbers in the smooth surface", "timestamp": "2026-07-22T04:47:44.310588+00:00"}
{"citation_id": "19930092013", "source_url": "https://ntrs.nasa.gov/api/citations/19930092013/downloads/19930092013.pdf", "page_number": 21, "total_pages": 21, "image_filename": "19930092013_p21.jpg", "text": "Positive directions of axes and angles (forces and moments) are shown by arrows\n\n| Axis | | Force (parallel to axis) symbol | Moment about axis | | | Angle | | Velocities | |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| Designation | Symbol | | Designation | Symbol | Positive direction | Designation | Symbol | Linear (component along axis) | Angular |\n| Longitudinal | $X$ | $X$ | Rolling | $L$ | $Y \\longrightarrow Z$ | Roll | $\\phi$ | $u$ | $p$ |\n| Lateral | $Y$ | $Y$ | Pitching | $M$ | $Z \\longrightarrow X$ | Pitch | $\\theta$ | $v$ | $q$ |\n| Normal | $Z$ | $Z$ | Yawing | $N$ | $X \\longrightarrow Y$ | Yaw | $\\psi$ | $w$ | $r$ |\n\nAbsolute coefficients of moment\n$$C_l = \\frac{L}{qbS} \\quad \\text{(rolling)}$$\n$$C_m = \\frac{M}{qcS} \\quad \\text{(pitching)}$$\n$$C_n = \\frac{N}{qbS} \\quad \\text{(yawing)}$$\n\nAngle of set of control surface (relative to neutral position), $\\delta$. (Indicate surface by proper subscript.)\n\n### 4. PROPELLER SYMBOLS\n\n$D$ Diameter\n$p$ Geometric pitch\n$p/D$ Pitch ratio\n$V'$ Inflow velocity\n$V_s$ Slipstream velocity\n$T$ Thrust, absolute coefficient $C_T = \\frac{T}{\\rho n^2 D^4}$\n$Q$ Torque, absolute coefficient $C_Q = \\frac{Q}{\\rho n^2 D^5}$\n\n$P$ Power, absolute coefficient $C_P = \\frac{P}{\\rho n^3 D^5}$\n$C_s$ Speed-power coefficient $= \\sqrt[5]{\\frac{\\rho V^5}{P n^2}}$\n$\\eta$ Efficiency\n$n$ Revolutions per second, rps\n$\\Phi$ Effective helix angle $= \\tan^{-1} \\left( \\frac{V}{2\\pi r n} \\right)$\n\n### 5. NUMERICAL RELATIONS\n\n1 hp = 76.04 kg-m/s = 550 ft-lb/sec\n1 metric horsepower = 0.9863 hp\n1 mph = 0.4470 mps\n1 mps = 2.2369 mph\n\n1 lb = 0.4536 kg\n1 kg = 2.2046 lb\n1 mi = 1,609.35 m = 5,280 ft\n1 m = 3.2808 ft", "timestamp": "2026-07-22T04:47:45.793492+00:00"}
{"citation_id": "19930082496", "source_url": "https://ntrs.nasa.gov/api/citations/19930082496/downloads/19930082496.pdf", "page_number": 13, "total_pages": 50, "image_filename": "19930082496_p13.jpg", "text": "12 NACA TN No. 1836\n\nThe results obtained in phase 2 of this evaluation with the second set of ceramal blades are indicated in table IV, which shows that 12 hours and 13 minutes of operation were completed. At the end of this time, service failure had taken place in 96 metal blades from phase 1 of this investigation, 8 control blades (a ninth blade had begun to crack), 4 replacement, and 1 ceramal blade. Of the three ceramal blades, only one can be considered a service failure. One ceramal blade, which was inadvertently fractured during overhaul after 9 hours and 42 minutes of operation, is shown in figure 11 in which a two-layered oxide film is apparent. The fracture of this blade during routine wheel handling is indicative of the need for special care in the handling of bodies fabricated of materials of this type.\n\nA second ceramal blade (fig. 12(a)) was destroyed after 12 hours and 13 minutes of operation because of wheel-dovetail failure that allowed the blade to fly loose. A study of figure 12(b) indicates distortion of disk material about the failed dovetail 33 minutes before failure. The ceramal blades were not removed to other positions because they were held very tightly and blade injury might result through moving. Enlargement of the dovetail roll, peening, and local heating of the disk because of the high thermal conductivity of the ceramal blade material undoubtedly resulted in a stress-temperature condition that caused fracture of the dovetail. The third ceramal blade apparently fractured as a result of factors normally operating to induce blade failure (fig. 13(a)), although the blade fractured simultaneously with the disk-dovetail failure.\n\nDuring the course of operation, the copper flowed from the ceramal blade roots, which is also indicated in figure 12(b). This condition was noted shortly after the wheel was put into operation. The greater part of phase 2 of this investigation was conducted with the ceramal blades retained in the manner indicated. A thin layer of copper remained fused to the blades and to the wheel dovetail, retaining the blades in place. Stress at the root neck-roll junction was relieved. A ceramal blade (fig. 12(b)) was nicked; this nick occurred during run 3 and was caused either by a flying blade fragment or by foreign matter in the inlet gas.\n\nThe results of the quasi-service evaluation of the ceramal blades show that all three ceramal blades survived 9 hours and 42 minutes of operation at which time six of the original twelve control blades survived. Two of the three ceramal blades survived to 12 hours and 13 minutes of operation at which time a total of four control blades survived. In addition, four metal blades installed to replace original metal blades had fractured.", "timestamp": "2026-07-22T04:47:48.844940+00:00"}
{"citation_id": "19930093773", "source_url": "https://ntrs.nasa.gov/api/citations/19930093773/downloads/19930093773.pdf", "page_number": 26, "total_pages": 47, "image_filename": "19930093773_p26.jpg", "text": "NACA RM E9G09\n25\n\n1159\n\n| Altitude (ft) | |\n| :--- | :--- |\n| 5,000 | O |\n| 15,000 | □ |\n| 25,000 | ◇ |\n| 35,000 | △ |\n| 45,000 | ▽ |\n| 50,000 | ◁ |\n\nFuel flow, $W_f$, lb/hr\n\nEngine speed, $N$, rpm\n\n(c) Fuel flow.\n\nFigure 4. - Continued. Effect of altitude on variation of engine performance with engine speed at flight Mach number of 0.21.", "timestamp": "2026-07-22T04:47:49.313959+00:00"}
{"citation_id": "19930082487", "source_url": "https://ntrs.nasa.gov/api/citations/19930082487/downloads/19930082487.pdf", "page_number": 22, "total_pages": 33, "image_filename": "19930082487_p22.jpg", "text": "```markdown\n20\n\nPressure coefficient, P\n\nCrest\n\nM, 1.0 1.4\n\nx/c\n0.025\n.100\n.183\n.333\n\nx/c\n0.025\n.100\n.157\n.257\n\nx/c\n0.050\n.150\n.215\n.365\n\nM, 1.0 1.4\n\nMcr Md Ms\n\nMcr Md Ms\n\nMcr Md Ms\n\nFree-stream Mach number, M_o\n\n(d) $\\alpha, 2^\\circ$; upper surface. (e) $\\alpha, 4^\\circ$; upper surface. (f) $\\alpha, 6^\\circ$; upper surface.\n\nFigure 4.- Concluded.\n\nNACA\n\nNACA TM NO. 1813\n```", "timestamp": "2026-07-22T04:47:50.301746+00:00"}
{"citation_id": "19930082542", "source_url": "https://ntrs.nasa.gov/api/citations/19930082542/downloads/19930082542.pdf", "page_number": 7, "total_pages": 53, "image_filename": "19930082542_p7.jpg", "text": "6\nNACA TN No. 1867\n\nholding duplicate stressed specimens for 1 and 24 hours at 1200° F\nbefore applying the stress. The materials used for this study were\nsolution-treated or solution-treated and aged.\n\nBrinell hardness tests were run on all the bars. Original\nmetallographic samples were prepared for observation, and photo-\nmicrographs were made of representative samples.\n\nRESULTS\n\nThe detailed test data from the tensile tests at room temperature\nand rupture tests at 1200° F are given in tables I and II. The\nreported rupture strengths are based on the best logarithmic curves of\nstress against rupture time which could be drawn through the available\ndata. Graphical presentation of the data has been used to show the\nfindings from the various specific treatments studied.\n\nTreatment of Hot-Rolled Stock\n\nThe effects of aging and of hot-cold-working the particular\nas-rolled stock used in this investigation, and summarized in figure 3,\nwere:\n\nAging.- The properties of the as-rolled stock were reduced at room\ntemperature by aging in the temperature range from 1350° to 1750° F.\nThese same treatments had very little effect on the stress for rupture\nin 1000 hours. Aging at 1500° to 1750° F did reduce the stress for\nrupture in 100 hours and increase the ductility in the rupture test.\n\nHot-cold-work.- Strength and hardness at room temperature were\nincreased and ductility reduced in proportion to the amount of\nreduction during hot-cold-rolling at 1200° F. A reduction of only\n10 percent produced a yield strength at 0.02-percent offset of\n100,000 psi.\n\nA reduction of 15 percent developed the maximum rupture strength.\nLower rupture strengths were found after larger amounts of hot-cold-work.\nThe reduction of 15 percent increased the stress for rupture of the hot-\nrolled material from 49,000 to 63,000 psi at 100 hours and from 37,500\nto 49,000 psi at 1000 hours. The hot-cold-work had no consistent effect\non ductility in the rupture test.\n\nThe only difference between a reduction of 10 percent at room\ntemperature and at 1200° F was somewhat lower ductility in the rupture\ntest due to working at room temperature.", "timestamp": "2026-07-22T04:47:52.574535+00:00"}
{"citation_id": "19930085965", "source_url": "https://ntrs.nasa.gov/api/citations/19930085965/downloads/19930085965.pdf", "page_number": 63, "total_pages": 67, "image_filename": "19930085965_p63.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T04:47:54.703656+00:00"}
{"citation_id": "19930082498", "source_url": "https://ntrs.nasa.gov/api/citations/19930082498/downloads/19930082498.pdf", "page_number": 12, "total_pages": 49, "image_filename": "19930082498_p12.jpg", "text": "NACA TN No. 1838\n11\n\nCONCLUSIONS\n\nAn experimental investigation has been made of the noise of a typical six-cylinder light-airplane engine and propeller combination and of the exhaust noise of the same engine without the propeller with a series of 68 muffler and tail-pipe configurations. The following conclusions have been drawn from the results of this investigation:\n\n1. Since engine and propeller noise are about equal for this power-plant installation, both engine-exhaust and propeller noises must be reduced to obtain a sizable reduction in the over-all sound-pressure level.\n\n2. Most of the exhaust sound energy is concentrated at low frequencies; therefore, sound-reduction methods applicable only to high-frequency sound are of little value in reducing the over-all sound-pressure level. The loudest component of engine-exhaust noise is at the fundamental firing frequency of the engine.\n\n3. The types of present-day commercial airplane mufflers investigated produced very little reduction in the over-all sound-pressure level.\n\n4. From the results of this exploratory investigation the following conclusions regarding the details of muffler design for light aircraft may be drawn:\n\n    a. Both resonant-chamber and expansion-chamber mufflers require large chamber volumes to reduce low-frequency noise.\n\n    b. If the low-frequency cut-off point is chosen too close to the fundamental firing frequency of the engine the sound-pressure reduction will be low.\n\n    c. Baffles between muffler chambers must be well sealed to avoid leakage or the muffler characteristics will be altered.\n\n    d. Mufflers of a given cross-sectional area of either circular or oval cross section, with all other dimensions the same, appear to give equal results.\n\n    e. The lower the required back pressure, the larger the muffler must be.\n\n    f. The straight-through or resonant-chamber mufflers have lower back pressure, in general, than the expansion-chamber mufflers.", "timestamp": "2026-07-22T04:48:01.371623+00:00"}
{"citation_id": "19930085838", "source_url": "https://ntrs.nasa.gov/api/citations/19930085838/downloads/19930085838.pdf", "page_number": 109, "total_pages": 118, "image_filename": "19930085838_p109.jpg", "text": "NACA RM No. L9B23\n107\n\nSection lift coefficient, $c_l$\nFlap deflection, $\\delta_f$\n(a) 0.177c thick airfoil\n\n$\\delta_f$ $c_b$\n$\\square$ 25° 0.351$c_a$\n$\\square$ 25° .406$c_a$\n$\\diamond$ 40° .351$c_a$\n$\\diamond$ 40° .406$c_a$\n\nSection lift coefficient, $c_l$\nFlap deflection, $\\delta_f$\n(b) 0.154c thick airfoil\n\nFigure 17.- Variation of section lift characteristics with flap deflection on two NACA 7-series-type airfoils with double slotted flap and straight-sided Frise aileron.\n$\\alpha_a = 0^\\circ$; $\\delta_a = 0^\\circ$; R = 6.0 x $10^6$ (approx.)", "timestamp": "2026-07-22T04:48:01.943333+00:00"}
{"citation_id": "19930082245", "source_url": "https://ntrs.nasa.gov/api/citations/19930082245/downloads/19930082245.pdf", "page_number": 30, "total_pages": 66, "image_filename": "19930082245_p30.jpg", "text": "```markdown\nNACA TN No. 1596\n\nSection angle of attack, $\\alpha$, deg\n$\\delta_a$ (deg)\n-12\n-6\n-4\n-2\n0\n2\n4\n12\n18\n30\nMach number, M\n.1 .2 .3 .4 .5 .6 .7 .8 .9\n\nSection pitching-moment coefficient, $c_m$\n$\\delta_a$ (deg)\n-12\n-6\n-4\n-2\n0\n2\n4\n12\n18\n30\nMach number, M\n.1 .2 .3 .4 .5 .6 .7 .8 .9\n\n[Figure: Two graphs side-by-side. Left graph plots Section angle of attack vs. Mach number for various $\\delta_a$ values. Right graph plots Section pitching-moment coefficient vs. Mach number for various $\\delta_a$ values. Both graphs have grid lines and labeled axes. NACA logo appears in bottom right of right graph.]\n\n(e) $c_n = 0.4$.\n\nFigure 6 —Continued.\n\n29\n```", "timestamp": "2026-07-22T04:48:06.993491+00:00"}
{"citation_id": "19930085842", "source_url": "https://ntrs.nasa.gov/api/citations/19930085842/downloads/19930085842.pdf", "page_number": 97, "total_pages": 104, "image_filename": "19930085842_p97.jpg", "text": "NACA RM L9C29\n93\n\n[Figure: Graph showing Angle of attack, $\\alpha$, deg vs. Lift coefficient, $C_L$ and Propeller advance-diameter ratio, $V/nD$ vs. Lift coefficient, $C_L$. Curves are labeled with $\\beta$, deg values of 30, 20, 14, and 11.5. Ticks indicate points of simulation of flight propeller-operating conditions. An Airplane curve is shown for 1200 bhp at 1085 rpm, normal gross weight.]\n\nFigure 45.- Curves used for the determination of flight propeller-operating lift coefficients from model data. Model curves duplicate airplane $Q_c$ against $C_L$ for full-power operation; basic model configuration; all controls neutral.", "timestamp": "2026-07-22T04:48:11.457700+00:00"}
{"citation_id": "19930082617", "source_url": "https://ntrs.nasa.gov/api/citations/19930082617/downloads/19930082617.pdf", "page_number": 5, "total_pages": 58, "image_filename": "19930082617_p5.jpg", "text": "4\nNACA TN 1962\n\nThe variation of the strain along the stringer at the edge of the cutout of cylinder 77 is shown in figure 27 for various loads. The length of the cutout in this cylinder was approximately 2.6 times the diameter of the cylinder, or approximately 50 percent greater than that of the cutouts of cylinders 72, 73, 74, and 75.\n\nA comparison of the strains in the middle bands of cylinders 75 and 77 is given for two applied loads in figures 28 and 29. These cylinders were identical except that their cutouts were approximately 1.56 and 2.6 times their diameter, respectively.\n\nThe maximum moment and strain attained by each cylinder are listed in table II, together with a comment on the type of failure. The behavior after collapse is described in table III.\n\nPhotographs of the test specimens after failure are presented in figures 30 to 38.\n\nDISCUSSION OF TEST RESULTS\n\nStrain Distribution\n\nAt low loads for cylinders with bottom cutouts good straight lines were obtained in the plots of strain against distance from the horizontal diameter except near the edge of the cutout. Near the edge stringer the strain always decreases in the full bands, increases in the middle bands of cylinders 74 and 75, and decreases in the middle bands of cylinders 72 and 73.\n\nIn general, for side cutouts the linearity is not so good as for bottom cutouts especially on the compression side. However, the deviations from linearity appear to be small.\n\nAs a rule, good agreement was obtained between measurements taken on locations symmetrically situated with respect to the vertical planes of symmetry of the cylinders. Exceptions are band H of cylinder 76 and bands B and V of cylinder 74. Moreover, considerable deviations from symmetry occurred at high loads when random displacements were induced by the approach of buckling.\n\nIn the edge stringer of cylinder 77 the strains in the middle band were considerably higher than in the full portions of the cylinder. The strain in the cutout region is highest at about the 25-percent mark along the cutout and is lower in the center. Close to failure, however, the pattern changes entirely since the edge stringer takes less than its share while the stringers farther away from the cutout take more. This is true for cylinder 77, the specimen with the longest cutout, but not necessarily for all the others.", "timestamp": "2026-07-22T04:48:14.297904+00:00"}
{"citation_id": "19930082447", "source_url": "https://ntrs.nasa.gov/api/citations/19930082447/downloads/19930082447.pdf", "page_number": 23, "total_pages": 24, "image_filename": "19930082447_p23.jpg", "text": "NACA TN No. 1775\n21\n\nPercent of total number\n40\n35\n30\n25\n20\n15\n10\n5\n0\n1.00 1.03 1.06 1.09 1.12 1.15 1.18 1.21 1.24 1.27 1.30\n\nTheoretical\n$[F(R)]_{M=1.036}$\n$[F(R)]_{M=1.085}$\n\nAverage of experimental data\n\nFigure 7.—Distribution of experimental values of the dead-rise function.\n\nNACA", "timestamp": "2026-07-22T04:48:23.498059+00:00"}
{"citation_id": "19930082614", "source_url": "https://ntrs.nasa.gov/api/citations/19930082614/downloads/19930082614.pdf", "page_number": 6, "total_pages": 36, "image_filename": "19930082614_p6.jpg", "text": "4\nNACA TN 1939\n\nSubscripts\n\ne estimated\n\no value at t = 0\n\n1 value at beginning of time interval $\\Delta t$\n\nCONTROL OF AIRPLANE SPEED\n\nFlight experience has indicated that the control of forward speed, which is possible through the use of aerodynamic brakes, is of considerable value in the operation of airplanes of all types. The required action of brakes in controlling the speed differs among various airplanes, and for any one airplane the requirements may differ with the type of maneuver that is to be performed. For the purpose of the present discussion, brakes are considered according to their use in producing longitudinal deceleration at approximately constant altitude, in permitting greater angles of dive at moderately low speeds, and in avoiding dangerously high speeds.\n\nDeceleration at Constant Altitude\n\nIt is to be expected that a device that controls the longitudinal deceleration would find especial applications in the operation of combat airplanes. A sudden deceleration would be required in order for a fighter airplane which was overtaking its target to slow down so as to have a maximum amount of time for firing. Rapid decelerations may also be called for in traffic-control zones during poor visibility in order to prevent collision.\n\nAerodynamic brakes may assist the pilot in performing various maneuvers. If the maximum normal acceleration is fixed by the maximum to which the pilot may be subjected or by structural limitations of the airplane, the minimum radius of curvature of the flight path varies with the square of the speed. Thus, a reduction in speed by the use of air brakes prior to and during a maneuver would effect a substantial decrease in the minimum turning radius.\n\nThe performance of aerodynamic brakes used primarily for speed reduction is indicated by the magnitude of the longitudinal deceleration in level flight. Calculations showing the variation of speed with time can be used to measure the comparative suitability of different brakes on the same airplane in producing needed changes of speed.", "timestamp": "2026-07-22T04:48:25.566356+00:00"}
{"citation_id": "19930090382", "source_url": "https://ntrs.nasa.gov/api/citations/19930090382/downloads/19930090382.pdf", "page_number": 37, "total_pages": 37, "image_filename": "19930090382_p37.jpg", "text": "NACA RM L9I07\n39\n\nCONFIDENTIAL\nNACA 4-(5)(08)-03\nNACA 4-(4)(06)-04\n\nEnvelope of maximum efficiency\n.4 .5 .6 .7 .8 .9 1.0\n\nCONFIDENTIAL\nForward Mach number, M\nFigure 8.- Effect of forward Mach number on envelope efficiency.\n\nNACA-Langley - 10-28-49 - 300", "timestamp": "2026-07-22T04:48:26.298445+00:00"}
{"citation_id": "19930082476", "source_url": "https://ntrs.nasa.gov/api/citations/19930082476/downloads/19930082476.pdf", "page_number": 22, "total_pages": 41, "image_filename": "19930082476_p22.jpg", "text": "20\n\nCHART 7.- EFFECT OF COMBINED AILERON DEFLECTIONS ON THE SPIN CHARACTERISTICS OF MODEL (RUDDERS AND AILERONS UNLINKED)\n[Right erect spins; elevator set to 1½° up, rudders set to neutral, ailerons set as indicated]\n\nA. Loading 1 ($\\frac{I_x - I_y}{mb^2} = -3 \\times 10^{-4}$; $\\mu = 5.04$; loading 1 in table II and point 1 in fig. 4)\nB. Loading 2 ($\\frac{I_x - I_y}{mb^2} = -49 \\times 10^{-4}$; $\\mu = 7.89$; loading 2 in table II and point 2 in fig. 4)\n\n| Right aileron up setting, deg | | | | | | Right aileron up setting, deg | | | | | |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| | | | | | | | | | | | |\n| | | | | | | | | | | | |\n| | | | | | | | | | | | |\n| | | | | | | | | | | | |\n| | | | | | | | | | | | |\n| | | | | | | | | | | | |\n| | | | | | | | | | | | |\n| | | | | | | | | | | | |\n| | | | | | | | | | | | |\n| | | | | | | | | | | | |\n| | | | | | | | | | | | |\n| | | | | | | | | | | | |\n| | | | | | | | | | | | |\n| | | | | | | | | | | | |\n| | | | | | | | | | | | |\n| | | | | | | | | | | | |\n| | | | | | | | | | | | |\n| | | | | | | | | | | | |\n| | | | | | | | | | | | |\n| | | | | | | | | | | | |\n| | | | | | | | | | | | |\n| | | | | | | | | | | | |\n| | | | | | | | | | | | |\n| | | | | | | | | | | | |\n| | | | | | | | | | | | |\n| | | | | | | | | | | | |\n| 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{"citation_id": "19930082712", "source_url": "https://ntrs.nasa.gov/api/citations/19930082712/downloads/19930082712.pdf", "page_number": 1, "total_pages": 14, "image_filename": "19930082712_p1.jpg", "text": "TN-1998\nNACA TN 1998\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\n\nTECHNICAL NOTE 1998\n\n[Stamp: RECEIVED DEC 27 1948 A.G.H.-W.D LOS ANGELES]\n\nAERODYNAMIC CHARACTERISTICS OF THE\nNACA 8-H-12 AIRFOIL SECTION AT SIX REYNOLDS NUMBERS\nFROM $1.8 \\times 10^6$ TO $11.0 \\times 10^6$\n\nBy Raymond F. Schaefer and Hamilton A. Smith\n\nLangley Aeronautical Laboratory\nLangley Air Force Base, Va.\n\nTECHNICAL LIBRARY\nAIRESEARCH MANUFACTURING CO.\n9851-9861 SEPULVEDA BLVD.\nLOS ANGELES 45,\nCALIFORNIA\n\n[Figure: NACA logo]\n\nWashington\nDecember 1949", "timestamp": "2026-07-22T04:48:31.862490+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 15, "total_pages": 99, "image_filename": "19930082511_p15.jpg", "text": "NACA TN No. 1826\n13\n\nThe necessity of performing an integration in order to determine v or w is a basic disadvantage of the acceleration-potential analogy compared with the velocity-potential analogy in which v and w are measurable directly.\n\nIn the analogies to be discussed in the present section, the perturbation velocity u is considered analogous to the electrical potential in a dilute electrolyte solution. A metal serves as a boundary along which u is constant, and the local intensity of current flowing into it gives $\\frac{\\partial u}{\\partial n}$; an insulator serves as a boundary where $\\frac{\\partial u}{\\partial n}$ is zero. The lifting element in either two or three dimensions is represented by a pair of short metal plates separated by an insulator; when the upper plate is maintained at a higher potential than the lower plate, the arrangement represents a thin airfoil with suction (large u) on its upper surface and pressure (small or negative u) on its lower surface. The current at each of the two plates should be the same in order that the slope of the airfoil surface (proportional to v) be the same on both upper and lower sides. In order always to satisfy this condition the voltage source activating the lifting element should not be tapped to any other electrode in the field.\n\nTwo-dimensional closed-open tunnel.- The setup for the two-dimensional closed-open tunnel with the lifting element on the center line is shown in figure 9(a). The walls of the upstream closed region are represented by insulators, which establish that $\\frac{\\partial u}{\\partial y} = 0$ at every point; hence, the condition that $v = \\int_{-\\infty}^{x} \\frac{\\partial u}{\\partial y} dx = 0$ at every point on the closed boundary is satisfied. The two free boundaries are represented by metal; and electrically connecting them, as shown, satisfies the further condition that they have the same potential (the same u). The flow of current into the lower boundary then equals the flow of current out of the upper boundary, so that the ultimate downstream value of v will be the same on both upper and lower boundaries, as is desired. In fact, for the symmetrical case illustrated, the value of v will be the same at all pairs of opposite points on the two free boundaries; so the boundaries will be everywhere parallel. No special attention need be paid to the entrance lips - the entrance lip condition is automatically satisfied since the potential u is continuous at these points (although the potential gradients at these points are infinite).\n\nFor the off-center position of the lifting element (fig. 9(b)) no modification of the circuits is needed. The difference between the potential in the upstream closed region and the potential of the free boundaries, which is the upstream perturbation velocity u, is measured with the aid of the probe P. As in the symmetrical case, the ultimate downstream value of v will be the same for both the upper and the lower boundaries; however, it is no longer true that the two boundaries will be everywhere parallel, and the ultimate width of the jet will be different from the width of the closed part.", "timestamp": "2026-07-22T04:48:32.837782+00:00"}
{"citation_id": "19930082585", "source_url": "https://ntrs.nasa.gov/api/citations/19930082585/downloads/19930082585.pdf", "page_number": 7, "total_pages": 30, "image_filename": "19930082585_p7.jpg", "text": "6\nNACA TN 1907\n\nIt is assumed that the angle of incidence varies with time exponentially, and according to\n\n$$\n\\theta = \\theta_F + A_1 e^{-a_1 t} + A_2 e^{-a_2 t} \\tag{2}\n$$\n\nwhere $A_1$, $A_2$, $a_1$, and $a_2$ are arbitrary, except that\n\n$$\nA_1 + A_2 = \\theta_0 - \\theta_F \\tag{2a}\n$$\n\nThe changes in incidence due to flapping or hunting are thus combined with whatever changes the pilot may make manually. By suitably choosing the constants, $A_1$, $A_2$, $a_1$, and $a_2$, rapid or slow pitch reductions after power failure may be investigated. Thus, the necessity of solving the fourth, fifth, and sixth of the basic equations is obviated, by the assumptions represented by equations (1) and (2), and the problem simplified enormously.\n\nIt may further be anticipated that the variation of $\\beta$ with time will be of minor importance both to the designer and in its effect on the variations of $V$ and $\\Omega$ with time. For the analysis of flapping angle it is therefore assumed that, in the transient period, $\\Omega$ is a constant and has the value of the average between $\\Omega_0$ and $\\Omega_F$. This average value is denoted by $\\Omega_a$.\n\nIn the following analysis only untwisted and untapered blades are considered. The equations developed on this basis can probably be applied to any blades, with fair accuracy, by using \"equivalent\" chord and angle of incidence.\n\nSolution for $\\beta$\n\nThe equation for the vertical acceleration of the helicopter may then be written as\n\n$$\n\\ddot{V} = g - \\frac{\\rho b a c R^3 \\Omega_a^2}{2 \\frac{W}{g}} \\left( \\frac{\\bar{V} - V}{2 \\Omega_a R} - \\frac{\\dot{\\beta}}{3 \\Omega_a} + \\frac{\\theta}{3} \\right) \\tag{3}\n$$", "timestamp": "2026-07-22T04:48:33.049933+00:00"}
{"citation_id": "19930093773", "source_url": "https://ntrs.nasa.gov/api/citations/19930093773/downloads/19930093773.pdf", "page_number": 27, "total_pages": 47, "image_filename": "19930093773_p27.jpg", "text": "26\nNACA RM E9G09\n\nSpecific fuel consumption based on net thrust, $W_f/F_n$\nlb/(hr)(lb thrust)\n\nAltitude\n(ft)\nO 5,000\n□ 15,000\n◇ 25,000\n△ 35,000\n▽ 45,000\n△ 50,000\n\nEngine speed, N, rpm\n\n[NACA logo]\n\n(d) Specific fuel consumption\nFigure 4. - Continued. Effect of altitude on variation of engine\nperformance with engine speed at flight Mach number of 0.21.\n\n1159", "timestamp": "2026-07-22T04:48:36.327364+00:00"}
{"citation_id": "19930082485", "source_url": "https://ntrs.nasa.gov/api/citations/19930082485/downloads/19930082485.pdf", "page_number": 23, "total_pages": 62, "image_filename": "19930082485_p23.jpg", "text": "22\nNACA TN No. 1810\n\nAPPENDIX B\n\nBLADE-DESIGN PROCEDURE\n\nThe velocity distribution for cascades of high solidity may be determined by use of the stream-filament theory presented in appendix A. Blades may be drawn by inspection to fit the velocity vectors specified by the vector diagrams. Starting the blade design by first drawing the camber line seems to be practically desirable. The ends of the camber line are drawn tangent to a line at an angle of $\\alpha_e + \\nu_e$ from the horizontal at the trailing edge and to a line at an angle $\\alpha_i + \\nu_i$ at the blade entrance (fig. 19). The camber line may be symmetrically drawn about line X, vertically drawn through the camber line where its slope is zero.\n\nA symmetrical airfoil is superimposed upon the camber line, thus a first approximation of the blade shape is formed. The maximum thickness of the symmetrical airfoil is placed at line X (fig. 20). A circle forming the leading edge is drawn with a diameter equal to about 30 percent of the maximum thickness of the foil. The trailing edge is drawn with an included angle of about $12^\\circ$. The maximum thickness of the foil will, of course, depend on the total turning angle $\\alpha_i + \\alpha_e$ and the desired surface velocities.\n\nAn orthogonal network of streamlines and velocity-potential lines are drawn in the channel formed by two airfoils. The network may be extended ahead of and behind the cascade so that the surface velocities may be computed at any point on the blade (fig. 19). The computed surface velocities in regions from $\\phi_A$ to the rear stagnation point and from $\\phi_B$ to the front stagnation point are approximate, inasmuch as the channel is bounded by an assumed streamline in these regions. The velocities at both ends of each velocity-potential line are computed by use of the methods of appendix A.\n\nIn general, the value of $\\frac{\\mu \\ n_o}{T}$ at the channel entrance will not be equal to that at the exit because of deviations from two-dimensional flow. For convenience, this term can be considered to vary linearly with the axial depth of the blade, as shown in figure 21. The following relations, based on the general energy equation and the isentropic relations for a perfect gas are used to evaluate $\\frac{\\mu \\ n_o}{T}$", "timestamp": "2026-07-22T04:48:36.521593+00:00"}
{"citation_id": "19930082487", "source_url": "https://ntrs.nasa.gov/api/citations/19930082487/downloads/19930082487.pdf", "page_number": 23, "total_pages": 33, "image_filename": "19930082487_p23.jpg", "text": "NACA TN No. 1813\n21\n\n[Figure: Graph plotting Free-stream Mach number, $M_\\infty$ (y-axis, 0.4 to 0.9) against Angle of attack, $\\alpha$, deg (x-axis, -6 to 8). The graph contains multiple curves labeled $M_s$, $M_d$, $(M_p)_{exp}$, $(M_p)_{calc}$, and $M_{cr}$. A NACA logo is present in the bottom right corner of the plot area.]\n\nFigure 5.—Variation of critical, drag-divergence, and shock-stall Mach numbers with angle of attack for the NACA 23015 airfoil section.", "timestamp": "2026-07-22T04:48:38.311078+00:00"}
{"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 4, "total_pages": 78, "image_filename": "19930082618_p4.jpg", "text": "2\nNACA TN 1945\n\ncondition, the saving in minimum drag to be derived from the use of\nNACA 6-series as compared with NACA 5-digit-series airfoil sections\ndisappears.\n\nDecreasing the Reynolds number from $9.0 \\times 10^6$ to $0.7 \\times 10^6$ caused\nreductions in the maximum lift of all the airfoils in both the smooth\nand rough surface conditions. The magnitude and character of this\nreduction varied rather inconsistently with airfoil design and surface\ncondition, however, so that the comparative merits of a group of airfoils\nchanged markedly and in a rather unpredictable manner with Reynolds\nnumber and surface condition.\n\nINTRODUCTION\n\nTwo-dimensional aerodynamic data corresponding to Reynolds numbers\nof $3.0 \\times 10^6$, $6.0 \\times 10^6$, and $9.0 \\times 10^6$ are now generally available\n(reference 1) for a rather large number of systematically derived\nNACA 6-series and 4-digit- and 5-digit-series airfoil sections. Although\nthe range of Reynolds number covered by the investigations reported in\nreference 1 is reasonably wide, engineering design problems such as\nmay be encountered in the selection of wing sections for small, personal-\ntype airplanes may require data for a range of Reynolds number extending\nbelow $3.0 \\times 10^6$.\n\nWith a view toward providing a basis upon which to choose airfoils\nfor such applications, the two-dimensional aerodynamic characteristics\nof 15 NACA airfoil sections have been determined at Reynolds numbers\nof $0.7 \\times 10^6$, $1.0 \\times 10^6$, $1.5 \\times 10^6$, and $2.0 \\times 10^6$. The results of this\ninvestigation are given in the present paper. In order to give a more\ncomprehensive picture of the manner in which the aerodynamic\ncharacteristics of the 15 airfoils vary with Reynolds number, data\nobtained for these airfoils at Reynolds numbers of $3.0 \\times 10^6$, $6.0 \\times 10^6$,\nand $9.0 \\times 10^6$ (references 1 to 3 and previously unpublished data) are\nalso presented.\n\nThe airfoils investigated consisted of 10 NACA 6-series sections\nand 5 airfoils of the NACA 4- and 5-digit-series groups. The airfoils\nwere chosen to show the effect upon the resultant aerodynamic\ncharacteristics at different Reynolds numbers of systematic variations\nin airfoil thickness, camber, and thickness distribution. Lift,\ndrag, and pitching-moment data are presented for each of the plain, smooth\nairfoils at the seven Reynolds numbers from $0.7 \\times 10^6$ to $9.0 \\times 10^6$. A", "timestamp": "2026-07-22T04:48:40.140877+00:00"}
{"citation_id": "19930082496", "source_url": "https://ntrs.nasa.gov/api/citations/19930082496/downloads/19930082496.pdf", "page_number": 14, "total_pages": 50, "image_filename": "19930082496_p14.jpg", "text": "NACA TN No. 1836\n13\n\nThe replacement blade fractured during run 7 was one of 19 blades installed after run 5. The replacement blades that fractured during run 8 represented one blade each of samples of 6, 17, and 8 blades installed after runs 1, 3, and 6, respectively. Statistically, each of these particular blades is the poorest of its respective sample, and this fact must be realized in order to evaluate the significance of the failure of these four replacement blades. During phase 2 there were eight starts and eight shutdowns. During starting, the inlet-gas temperature was brought from room temperature to slightly under evaluation temperature within 30 seconds. Upon shutting down, combustion was abruptly terminated and air flow reduced slightly. Both operations produced thermal shock, which had no noticeable effect upon the ceramal blades.\n\nThe fracture surface of the service-failed ceramal blade after 12 hours and 13 minutes of operation is shown in figure 13(b). A two-layered oxide film is in evidence. This film was found to be of different thickness on the leading and trailing edges of the blade, which might be indicative of the temperatures at those areas. The film thicknesses for this blade and for the blade inadvertently fractured after 9 hours and 42 minutes of operation are given in the following table:\n\n| Time to fracture | | Leading-edge thickness (in.) | Trailing-edge thickness (in.) |\n| :--- | :--- | :--- | :--- |\n| (hr) | (min) | | |\n| 9 | 42 | 0.0134 | 0.0035 |\n| 12 | 13 | .0223 | .0165 |\n\nThese measurements were taken by use of a microscope fitted with a filar eyepiece, with the fractured surface of a blade viewed through the eyepiece. The scale was tenacious and showed no tendency to flake off during turbine operation. This type scale would tend to preserve aerodynamic shape and to minimize the risk of flying scale damaging sound blades. Expedients such as ceramic coating may be found necessary to minimize oxidation for long periods of operation. No appreciable elongation of the ceramal blades occurred.\n\nX-Ray-Diffraction Study\n\nThe two layers of scale of a fractured blade specimen were examined. The outer layer was composed mainly of titanium dioxide $TiO_2$ (rutile). The inner layer was composed mainly of cobalt titanate $CoTiO_3$. The base material of the ceramal exhibited no significant change.", "timestamp": "2026-07-22T04:48:40.517158+00:00"}
{"citation_id": "19930085965", "source_url": "https://ntrs.nasa.gov/api/citations/19930085965/downloads/19930085965.pdf", "page_number": 64, "total_pages": 67, "image_filename": "19930085965_p64.jpg", "text": "1125\n\nNACA RM E9E06\n\n[Figure: Oscillogram (c) showing a waveform with multiple peaks and troughs]\n\n(c) 383 root-mean-square ampere turns.\n\n[Figure: Oscillogram (d) showing a waveform with fewer peaks and troughs compared to (c)]\n\n(d) 281 root-mean-square ampere turns; 414 direct-current ampere turns.\n\nNACA\nC-23406\n5-6-49\n\nFigure 17. - Concluded. Search-coil voltage oscillograms for comparing wave shapes with and without superimposed direct-current ampere turns. Armco Magnetic Ingot Iron; air gap, 0.04 inch.\n\n63", "timestamp": "2026-07-22T04:48:40.899891+00:00"}
{"citation_id": "19930082498", "source_url": "https://ntrs.nasa.gov/api/citations/19930082498/downloads/19930082498.pdf", "page_number": 13, "total_pages": 49, "image_filename": "19930082498_p13.jpg", "text": "12\nNACA TN No. 1838\n\ng. The tail-pipe configuration chosen may have a large effect\non the exhaust-system characteristics.\n\nLangley Aeronautical Laboratory\nNational Advisory Committee for Aeronautics\nLangley Air Force Base, Va., December 29, 1948\n\nREFERENCES\n\n1. Czarnecki, K. R., and Davis, Don D., Jr.: Dynamometer-Stand Investigation\nof the Muffler Used in the Demonstration of Light-Airplane Noise\nReduction. NACA TN No. 1688, 1948.\n\n2. Vogeley, A. W.: Sound-Level Measurements of a Light Airplane Modified\nto Reduce Noise Reaching the Ground. NACA TN No. 1647, 1948.\n\n3. Morley, A. W.: Progress of Experiments in Aero-Engine Exhaust\nSilencing. R. & M. No. 1760, British A.R.C., 1937.", "timestamp": "2026-07-22T04:48:44.314781+00:00"}
{"citation_id": "19930082450", "source_url": "https://ntrs.nasa.gov/api/citations/19930082450/downloads/19930082450.pdf", "page_number": 16, "total_pages": 37, "image_filename": "19930082450_p16.jpg", "text": "NACA TN No. 1778\n15\n\nTABLE 4.- N-FAIRL PROPERTIES - Concluded ($\\frac{t_w}{t_f} = 0.79$; $\\frac{b_w}{t_f} = 9.8$; $\\frac{b_f}{t_f} = 0.4$; $\\frac{t_w}{t_f} = 3$; $\\frac{D_w}{t_f} = 4.1$; $\\frac{d}{t_f} = 1.93$; $\\frac{D}{t_f} = 18.3$)\n\n| $\\frac{t_w}{t_f}$ | 33 | 34 | 35 | 36 | 37 | 38 | 39 | 40 | 41 | 42 | 43 | 44 | 45 |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| 25 | 2.212 | 2.247 | 2.282 | 2.317 | 2.352 | 2.387 | 2.422 | 2.457 | 2.492 | 2.527 | 2.562 | 2.597 | 2.632 |\n| 26 | 2.202 | 2.237 | 2.272 | 2.307 | 2.342 | 2.377 | 2.412 | 2.447 | 2.482 | 2.517 | 2.552 | 2.587 | 2.622 |\n| 27 | 2.216 | 2.247 | 2.280 | 2.312 | 2.345 | 2.377 | 2.410 | 2.442 | 2.474 | 2.506 | 2.539 | 2.571 | 2.604 |\n| 28 | 2.171 | 2.203 | 2.234 | 2.265 | 2.297 | 2.328 | 2.359 | 2.390 | 2.422 | 2.453 | 2.484 | 2.515 | 2.546 |\n| 29 | 2.131 | 2.162 | 2.193 | 2.224 | 2.255 | 2.286 | 2.317 | 2.348 | 2.379 | 2.410 | 2.441 | 2.472 | 2.503 |\n| 30 | 2.093 | 2.123 | 2.152 | 2.181 | 2.210 | 2.239 | 2.268 | 2.297 | 2.327 | 2.356 | 2.385 | 2.414 | 2.443 |\n| 31 | 2.058 | 2.087 | 2.116 | 2.145 | 2.174 | 2.203 | 2.232 | 2.261 | 2.290 | 2.319 | 2.348 | 2.377 | 2.406 |\n| 32 | 2.026 | 2.053 | 2.080 | 2.107 | 2.135 | 2.162 | 2.189 | 2.216 | 2.243 | 2.271 | 2.299 | 2.326 | 2.353 |\n| 33 | 1.995 | 2.022 | 2.049 | 2.076 | 2.103 | 2.130 | 2.157 | 2.184 | 2.211 | 2.238 | 2.265 | 2.292 | 2.319 |\n| 34 | 1.965 | 1.991 | 2.017 | 2.042 | 2.068 | 2.093 | 2.119 | 2.145 | 2.171 | 2.196 | 2.222 | 2.248 | 2.274 |\n| 35 | 1.937 | 1.962 | 1.987 | 2.012 | 2.037 | 2.062 | 2.087 | 2.112 | 2.137 | 2.162 | 2.187 | 2.212 | 2.237 |\n| 36 | 1.912 | 1.936 | 1.960 | 1.984 | 2.009 | 2.033 | 2.057 | 2.081 | 2.106 | 2.130 | 2.154 | 2.178 | 2.203 |\n| 37 | 1.888 | 1.912 | 1.936 | 1.959 | 1.983 | 2.006 | 2.029 | 2.052 | 2.076 | 2.099 | 2.123 | 2.146 | 2.170 |\n| 38 | 1.864 | 1.888 | 1.909 | 1.932 | 1.955 | 1.978 | 2.001 | 2.024 | 2.047 | 2.070 | 2.093 | 2.116 | 2.139 |\n| 39 | 1.842 | 1.864 | 1.886 | 1.908 | 1.931 | 1.953 | 1.976 | 1.998 | 2.021 | 2.043 | 2.066 | 2.088 | 2.110 |\n| 40 | 1.820 | 1.842 | 1.864 | 1.886 | 1.907 | 1.929 | 1.951 | 1.972 | 1.994 | 2.015 | 2.037 | 2.058 | 2.080 |\n| 41 | 1.781 | 1.802 | 1.823 | 1.844 | 1.865 | 1.885 | 1.906 | 1.927 | 1.948 | 1.968 | 1.989 | 2.010 | 2.031 |\n| 42 | 1.746 | 1.765 | 1.785 | 1.805 | 1.825 | 1.845 | 1.865 | 1.885 | 1.905 | 1.925 | 1.944 | 1.964 | 1.984 |\n| 43 | 1.713 | 1.732 | 1.751 | 1.770 | 1.789 | 1.808 | 1.827 | 1.846 | 1.865 | 1.884 | 1.903 | 1.922 | 1.941 |\n| 44 | 1.683 | 1.702 | 1.720 | 1.738 | 1.757 | 1.775 | 1.794 | 1.812 | 1.831 | 1.850 | 1.868 | 1.886 | 1.905 |\n| 45 | 1.657 | 1.674 | 1.692 | 1.709 | 1.726 | 1.743 | 1.761 | 1.778 | 1.796 | 1.813 | 1.831 | 1.848 | 1.866 |\n| 46 | 1.631 | 1.648 | 1.665 | 1.681 | 1.698 | 1.715 | 1.732 | 1.748 | 1.765 | 1.782 | 1.799 | 1.816 | 1.833 |\n| 47 | 1.607 | 1.624 | 1.640 | 1.656 | 1.672 | 1.688 | 1.705 | 1.721 | 1.737 | 1.753 | 1.770 | 1.786 | 1.802 |\n| 48 | 1.586 | 1.601 | 1.617 | 1.633 | 1.649 | 1.664 | 1.680 | 1.695 | 1.711 | 1.726 | 1.742 | 1.757 | 1.773 |\n| 49 | 1.567 | 1.581 | 1.596 | 1.611 | 1.626 | 1.641 | 1.656 | 1.671 | 1.686 | 1.701 | 1.716 | 1.731 | 1.746 |\n| 50 | 1.547 | 1.561 | 1.576 | 1.590 | 1.605 | 1.620 | 1.635 | 1.649 | 1.664 | 1.678 | 1.693 | 1.707 | 1.722 |\n| 55 | 1.495 | 1.508 | 1.521 | 1.534 | 1.547 | 1.560 | 1.573 | 1.586 | 1.599 | 1.612 | 1.625 | 1.638 | 1.651 |\n| 60 | 1.466 | 1.481 | 1.494 | 1.506 | 1.519 | 1.531 | 1.544 | 1.556 | 1.569 | 1.581 | 1.594 | 1.606 | 1.619 |\n| 65 | 1.458 | 1.470 | 1.483 | 1.495 | 1.507 | 1.519 | 1.531 | 1.543 | 1.555 | 1.567 | 1.579 | 1.591 | 1.603 |\n| 70 | 1.458 | 1.470 | 1.483 | 1.495 | 1.507 | 1.519 | 1.531 | 1.543 | 1.555 | 1.567 | 1.579 | 1.591 | 1.603 |\n| 75 | 1.458 | 1.470 | 1.483 | 1.495 | 1.507 | 1.519 | 1.531 | 1.543 | 1.555 | 1.567 | 1.579 | 1.591 | 1.603 |\n\n| $\\frac{t_w}{t_f}$ | 33 | 34 | 35 | 36 | 37 | 38 | 39 | 40 | 41 | 42 | 43 | 44 | 45 |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| 25 | 8.845 | 9.195 | 9.578 | 9.966 | 10.36 | 10.75 | 11.15 | 11.55 | 11.95 | 12.35 | 12.75 | 13.15 | 13.55 |\n| 26 | 8.672 | 9.050 | 9.448 | 9.809 | 10.19 | 10.59 | 10.98 | 11.38 | 11.77 | 12.18 | 12.58 | 12.99 | 13.40 |\n| 27 | 8.532 | 9.009 | 9.282 | 9.662 | 10.04 | 10.43 | 10.82 | 11.21 | 11.61 | 12.01 | 12.40 | 12.81 | 13.21 |\n| 28 | 8.406 | 8.770 | 9.142 | 9.518 | 9.893 | 10.28 | 10.66 | 11.05 | 11.44 | 11.84 | 12.23 | 12.63 | 13.03 |\n| 29 | 8.277 | 8.641 | 9.008 | 9.376 | 9.751 | 10.13 | 10.51 | 10.90 | 11.29 | 11.68 | 12.07 | 12.46 | 12.86 |\n| 30 | 8.154 | 8.510 | 8.873 | 9.244 | 9.612 | 9.997 | 10.37 | 10.75 | 11.14 | 11.52 | 11.91 | 12.30 | 12.69 |\n| 31 | 8.033 | 8.389 | 8.745 | 9.109 | 9.477 | 9.848 | 10.22 | 10.60 | 10.98 | 11.36 | 11.75 | 12.14 | 12.53 |\n| 32 | 7.913 | 8.265 | 8.622 | 8.989 | 9.353 | 9.721 | 10.08", "timestamp": "2026-07-22T04:48:46.592885+00:00"}
{"citation_id": "19930085838", "source_url": "https://ntrs.nasa.gov/api/citations/19930085838/downloads/19930085838.pdf", "page_number": 110, "total_pages": 118, "image_filename": "19930085838_p110.jpg", "text": "108\n\nIncrement of section angle of attack, $\\Delta\\alpha_o$, deg\n\n$\\sigma_b = 0.351c_a$\n\n$\\delta_t$\n(deg)\n$\\circ$ 0\n$\\diamond$ -5\n$\\square$ -10\n$\\triangle$ -15\n$\\nabla$ -20\n\n[Figure: Graph plotting Increment of section angle of attack vs Aileron deflection with data points and a trend line. A NACA logo is present in the bottom right corner of the plot area.]\n\nAileron deflection, $\\delta_a$, deg\n\n(a) 0.177c thick airfoil; $\\delta_f = 0^\\circ$; $c_l = 0.4$.\n\nFigure 18.- Lift effectiveness of the straight-sided Frise ailerons and flaps on the NACA 7-series-type airfoils with double slotted flaps. $R = 6 \\times 10^6$ (approx.).\n\nNACA RM No. L9E63", "timestamp": "2026-07-22T04:48:48.814178+00:00"}
{"citation_id": "19930085842", "source_url": "https://ntrs.nasa.gov/api/citations/19930085842/downloads/19930085842.pdf", "page_number": 98, "total_pages": 104, "image_filename": "19930085842_p98.jpg", "text": "94\nNACA RM L9C29\n\nTicks indicate points of simulation of flight\npropeller-operating conditions\n\n$\\alpha$, deg\n60\n40\n20\n\nSimulations of $Q_c$ vs $C_L$ for full-power operation\nSimulations of $Q_c$ vs $C_L$ for 118 percent full-power operation\nAirplane $V_{HD}$ vs $C_L$ curve for 14.2 percent decrease\nin weight\nAirplane $V_{HD}$ vs $C_L$ curve for 23.6 percent decrease\nin weight\n\n6\n4\n$V_{HD}$\n2\n0 2 3 4 5 6 7 8 9 10\nLift coefficient, $C_L$\n\nNATIONAL ADVISORY\nCOMMITTEE FOR AERONAUTICS\n\nFigure 46.- Curves used for the determination of flight propeller-\noperating lift coefficients from model data at $\\beta = 11.5^\\circ$. Basic\nmodel configuration; all controls neutral.", "timestamp": "2026-07-22T04:48:51.728395+00:00"}
{"citation_id": "19930082245", "source_url": "https://ntrs.nasa.gov/api/citations/19930082245/downloads/19930082245.pdf", "page_number": 31, "total_pages": 66, "image_filename": "19930082245_p31.jpg", "text": "```markdown\n16\n14\n12\n10\n8\n6\n4\n2\n0\n-2\n-4\n-6\n.1 .2 .3 .4 .5 .6 .7 .8 .9\nMach number, M\n\nSection angle of attack, $\\alpha$, deg\n\n$\\delta_a$ (deg)\n-12\n-6\n-4\n-2\n0\n2\n4\n12\n18\n30\n\n.16\n.12\n.08\n.04\n0\n-.04\n-.08\n-.12\n-.16\n-.20\n-.24\n-.28\n.1 .2 .3 .4 .5 .6 .7 .8 .9\nMach number, M\n\nSection pitching-moment coefficient, $C_m$\n\n$\\delta_a$ (deg)\n-12\n-6\n-4\n-2\n0\n2\n4\n12\n18\n30\n\n(F) $C_n=0.6$.\nFigure 6 .-Continued.\n\nNACA\nNACA TN No. 1596\n30\n```", "timestamp": "2026-07-22T04:48:55.292916+00:00"}
{"citation_id": "19930082617", "source_url": "https://ntrs.nasa.gov/api/citations/19930082617/downloads/19930082617.pdf", "page_number": 6, "total_pages": 58, "image_filename": "19930082617_p6.jpg", "text": "NACA TN 1962\n\nCylinders 75 and 77 were identical except that the lengths of their cutouts were 1.56 and 2.6 times their diameter, respectively. The strain distribution for these two specimens was almost the same except very close to the cutout. For example, at low loads the edge stringer was more highly stressed with the short cutout than with the long cutout.\n\nAt high loads, usually above 75 percent of the critical moment, the variations from linearity are considerable for all cylinders. This is because the strain readings reflect random variations in the displacements at loads close to failure.\n\nFailure of Cylinders with Bottom Cutouts\n\nThe five cylinders which had bottom cutouts were cylinders 72, 73, 74, 75, and 77. Of these, the first three failed in a definite general-instability pattern. In cylinder 75 failure occurred first by the shearing of two bolts at ring-stringer intersections. The load was then completely removed and the bolts replaced. When the load was reapplied, the specimen failed at a higher moment by the shearing of one of the new bolts and the cracking of an edge stringer at the bolt failure. In cylinder 77, which had the longest cutout, buckling was precipitated by a tension diagonal failure of one of the sheet panels adjacent to the cutout. Except in one side of cylinder 75 where the bolt failure renders the results doubtful, all the buckling shapes of the edge stringers resembled that of a fixed-ended column. The length of the distortion of the edge stringer was slightly longer than the cutout in cylinder 72, slightly shorter than the length of the cutout for cylinders 73, 74, and 75, and about two-thirds the length of the cutout for cylinder 77. The side of cylinder 75 on which the bolts failed was the only edge stringer to fail in an S-shaped curve.\n\nFailure of Cylinders with Side Cutouts\n\nThree cylinders with side cutouts were tested. Two of these, cylinders 76 and 79, failed in a general-instability pattern and the third, cylinder 78, failed in tension.\n\nOf the three specimens, the one that failed in tension required the highest moment, namely, 451,225 inch-pounds. Several stringers on the upper side of the cylinder tore close to the loading head.\n\nCylinders 76 and 79 failed at 324,000 and 370,000 inch-pounds, respectively. Both buckled in a general-instability pattern in which the bottom stringer was S-shaped and the sides were pushed in where the bottom stringer buckled out.", "timestamp": "2026-07-22T04:48:57.277737+00:00"}
{"citation_id": "19930082542", "source_url": "https://ntrs.nasa.gov/api/citations/19930082542/downloads/19930082542.pdf", "page_number": 8, "total_pages": 53, "image_filename": "19930082542_p8.jpg", "text": "NACA TN No. 1867\n\nAging at $1400^\\circ$ F after 15 percent hot-cold-work sharply reduced all properties over those of the hot-cold-worked material except hardness, ductility, and rupture strength at 1000 hours.\n\nAgglomeration of excess constituents.— Severe working and annealing in the temperature range from $1800^\\circ$ to $1400^\\circ$ F to agglomerate excess constituents resulted in rupture properties similar to those obtained by aging at $1500^\\circ$ to $1750^\\circ$ F. It had been expected that such treatments would reduce rupture strengths to very low values.\n\nProperties after Solution Treatment\n\nThe major part of the work was conducted on stock which had been solution-treated. The effects of various treatments should be more generally applicable after solution treatments than for as-rolled stock which can be a quite variable material depending on rolling conditions. Graphical treatment has been used to show the findings from the various experiments:\n\nTime of solution treatment and cooling rate.— No variation of properties with solution time was observed at either $2050^\\circ$ or $2200^\\circ$ F solution temperatures. (See fig. 4.) Rupture properties were somewhat reduced with air-cooling when compared with those obtained by water-quenching.\n\nTemperature of solution treatment.— The major effects of solution treatments alone were to reduce strength and increase ductility at room temperature and to reduce the stress for rupture in 100 hours and the ductility in the rupture test. (See fig. 5.) The highest rupture strengths resulted from solution-treating at $2100^\\circ$ F, although there was very little difference over the temperature range from $1950^\\circ$ to $2100^\\circ$ F. Higher-temperature treatments resulted in lower strengths. Ductility in the rupture test fell off rapidly with increasing solution temperature up to $2100^\\circ$ F.\n\nSolution treatments decrease the slope of the curves of stress against rupture time. This was indicated by most of the observed stresses for fracture in 1000 hours being equal to or higher than those for the hot-rolled stock and the 100-hour strengths being sharply reduced.\n\nSome susceptibility to brittleness at stress concentrations is to be expected when the solution-treatment temperature is high enough to cause low elongation in the rupture test. This was evidenced by a tendency for fracture to occur in gage marks or in fillets after the solution temperature was raised above $2100^\\circ$ F.\n\nThe microstructures changed gradually with increasing temperature of solution treatment. (See fig. 6.) Partial resolution of the precipitates which formed during heating up occurred at $1950^\\circ$ F. Insofar", "timestamp": "2026-07-22T04:49:00.425926+00:00"}
{"citation_id": "19930082914", "source_url": "https://ntrs.nasa.gov/api/citations/19930082914/downloads/19930082914.pdf", "page_number": 1, "total_pages": 66, "image_filename": "19930082914_p1.jpg", "text": "NATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\n\nTECHNICAL NOTE\nNo. 1857\n\nINVESTIGATION WITH AN INTERFEROMETER OF THE TURBULENT\nMIXING OF A FREE SUPERSONIC JET\n\nBy Paul B. Gooderum, George P. Wood,\nand Maurice J. Brevoort\n\nLangley Aeronautical Laboratory\nLangley Air Force Base, Va.\n\n[Figure: NACA logo]\n\nWashington\nApril 1949", "timestamp": "2026-07-22T04:49:07.485596+00:00"}
{"citation_id": "19930082712", "source_url": "https://ntrs.nasa.gov/api/citations/19930082712/downloads/19930082712.pdf", "page_number": 2, "total_pages": 14, "image_filename": "19930082712_p2.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T04:49:09.630718+00:00"}
{"citation_id": "19930082614", "source_url": "https://ntrs.nasa.gov/api/citations/19930082614/downloads/19930082614.pdf", "page_number": 7, "total_pages": 36, "image_filename": "19930082614_p7.jpg", "text": "NACA TN 1939\n\nSpeed Control in Dives\n\nAerodynamic brakes which keep the speed moderately low during dives are needed for the performance of some tactical maneuvers. For the dive bomber, low speed is required as a precaution against excessive normal accelerations during pull-outs at low altitude. In rapid descent of large airplanes, the desirability of avoiding high forward speeds is apparent. The braking which would be sufficient to prevent increases in speed may be used to gauge the performance of brakes for airplanes in controlled-speed dives. Aerodynamic data for brakes to be used on airplanes in such dives may be evaluated by a comparison of the measured brake drag with the brake drag which would be required to reduce the longitudinal acceleration to zero.\n\nSpeed Control at High Speed\n\nThe maximum safe diving speed is commonly specified as the maximum indicated airspeed upon which calculations of loads are based in the structural analysis. The rapid change in altitude resulting from a steep dive at high speed means that the true speed would have to be reduced in order to keep from exceeding the allowable indicated speed. Calculations which show the variation with time of the speed of an airplane with brakes indicate whether the drag of the brakes is sufficient to produce the required reductions in speed. Methods are presented in this report to aid in making such calculations.\n\nOne of the most important examples of the employment of aerodynamic brakes is their use in avoiding dangerous compressibility effects. These effects take the form of changes in longitudinal stability, such as described in reference 1, or as other erratic behavior of airplanes or their controls. Without brakes, pilots of modern high-speed combat airplanes are not always able to avoid speeds at which such compressibility effects appear.\n\nA method of calculation which shows the variation with time of speed and altitude of an airplane throughout various maneuvers can be used to determine whether the aerodynamic brakes are adequate to keep the maximum Mach number below a certain critical value.\n\nANALYSIS\n\nThe foregoing discussion has indicated various instances wherein a need for aerodynamic brakes has been observed. The drag increments due to air brakes do not by themselves afford a complete measure of the degree to which brakes meet these needs. Brake designs can be evaluated by studies which demonstrate the effects of the brakes upon the forward speed of airplanes. These effects are analyzed in this", "timestamp": "2026-07-22T04:49:11.297894+00:00"}