PDF/ocr_dataset_1017.jsonl

{"citation_id": "19930082476", "source_url": "https://ntrs.nasa.gov/api/citations/19930082476/downloads/19930082476.pdf", "page_number": 16, "total_pages": 41, "image_filename": "19930082476_p16.jpg", "text": "14\nNACA TN No. 1801\n\nTABLE II.- MASS CHARACTERISTICS AND INERTIA PARAMETERS\nFOR LOADINGS TESTED ON THE MODEL\n\n[Model values converted to corresponding full-scale values]\n\n| Loading | Loading condition | Weight (lb) | Wing loading (lb/sq ft) | Relative density | | Center of gravity | |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| | | | | Sea level | 5000 feet | $x/\\bar{c}$ | $z/\\bar{c}$ |\n| 1 | Normal | 1424 | 9.99 | 4.35 | 5.04 | 0.182 | 0.088 |\n| 2 | Mass extended along fuselage | 1491 | 10.46 | 4.55 | 5.29 | .173 | .088 |\n| 3 | Mass extended along wings | 1499 | 10.51 | 4.57 | 5.32 | .199 | .101 |\n| 4 | Relative density approximately doubled from normal loading | 2929 | 20.54 | 8.93 | 10.39 | .187 | .025 |\n\n| | Moments of inertia (slug-ft$^2$) | | | Inertia parameters | | |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| Loading | $I_X$ | $I_Y$ | $I_Z$ | $\\frac{I_X - I_Y}{mb^2}$ | $\\frac{I_Y - I_Z}{mb^2}$ | $\\frac{I_Z - I_X}{mb^2}$ |\n| 1 | 701 | 712 | 1347 | $-3 \\times 10^{-4}$ | $-160 \\times 10^{-4}$ | $163 \\times 10^{-4}$ |\n| 2 | 731 | 921 | 1583 | -49 | -154 | 203 |\n| 3 | 1481 | 790 | 2127 | 165 | -319 | 154 |\n| 4 | 1289 | 1440 | 2588 | -18 | -140 | 158 |\n\nNACA", "timestamp": "2026-07-22T04:43:10.124227+00:00"}
{"citation_id": "19930090382", "source_url": "https://ntrs.nasa.gov/api/citations/19930090382/downloads/19930090382.pdf", "page_number": 32, "total_pages": 37, "image_filename": "19930090382_p32.jpg", "text": "34\nNACA RM L9I07\n\nCONFIDENTIAL\n\nThrust coefficient, $C_T$\nPower coefficient, $C_P$\nTip Mach number, $M_t$\nEfficiency, $\\eta$\n\n<!-- Image (77, 109, 883, 999) -->\n\nCONFIDENTIAL\nAdvance ratio, J\n(1) M=0.90.\nFigure 5 - Continued.", "timestamp": "2026-07-22T04:43:12.866796+00:00"}
{"citation_id": "19930086015", "source_url": "https://ntrs.nasa.gov/api/citations/19930086015/downloads/19930086015.pdf", "page_number": 50, "total_pages": 54, "image_filename": "19930086015_p50.jpg", "text": "NACA RM A59E24\n\nCONFIDENTIAL\n\nTheory\nExperiment : $\\circ \\Delta \\alpha = 5.35^\\circ$ (Model Vertical)\n$\\triangle \\Delta \\alpha = 3.70^\\circ$ (Model Horizontal)\n$\\square \\Delta \\alpha = 5.74^\\circ$ (Model Horizontal)\n\nLoading coefficient\nper unit angle of attack, $P_\\alpha$, per deg\n\nPercent of local chord\n\n(b) M = 1.40.\nFigure 14.- continued.\n\nNACA\n\nCONFIDENTIAL\n\n49", "timestamp": "2026-07-22T04:43:21.432154+00:00"}
{"citation_id": "19930086081", "source_url": "https://ntrs.nasa.gov/api/citations/19930086081/downloads/19930086081.pdf", "page_number": 42, "total_pages": 44, "image_filename": "19930086081_p42.jpg", "text": "40\nNACA RM L9H05\n\nCONFIDENTIAL\n\n<!-- Image (89, 110, 868, 842) -->\n\n(c) Hinge moment plotted against $\\alpha$.\nFigure 17.- Concluded.", "timestamp": "2026-07-22T04:43:26.279581+00:00"}
{"citation_id": "19930082485", "source_url": "https://ntrs.nasa.gov/api/citations/19930082485/downloads/19930082485.pdf", "page_number": 16, "total_pages": 62, "image_filename": "19930082485_p16.jpg", "text": "NACA TN No. 1810\n\n$$\ng = \\frac{z_{\\mathrm{m}}^{3/2}}{\\gamma - 1} \\left( 1 - z_{\\mathrm{m}} \\right)^{\\frac{2 - \\gamma}{\\gamma - 1}}\n$$\n\nThen,\n\n$$\n(1 - z)^{\\frac{1}{\\gamma - 1}} = \\frac{f}{\\sqrt{z_{\\mathrm{m}}}} - g \\frac{z}{z_{\\mathrm{m}}^{3/2}}\n$$\n\nThen equation (11) becomes\n\n$$\n\\frac{\\mathrm{d}W}{\\rho_{\\mathrm{t}} \\, n_{\\mathrm{o}} \\sqrt{2 \\, c_{\\mathrm{p}} \\, T_{\\mathrm{t}}}} = \\left[ f \\sqrt{\\frac{z}{z_{\\mathrm{m}}}} - g \\left( \\frac{z}{z_{\\mathrm{m}}} \\right)^{3/2} \\right] \\frac{\\mathrm{d}n}{n_{\\mathrm{o}}}\n$$\n\nLet\n\n$$\n\\mu = \\frac{1}{\\rho_{\\mathrm{t}} \\, n_{\\mathrm{o}} \\sqrt{2 \\, c_{\\mathrm{p}} \\, T_{\\mathrm{t}}}} \\int_{0}^{W} \\mathrm{d}W\n$$\n\n$$\n= \\frac{W}{\\rho_{\\mathrm{t}} \\, n_{\\mathrm{o}} \\sqrt{2 \\, c_{\\mathrm{p}} \\, T_{\\mathrm{t}}}}\n$$\n\nIf equation (18) is integrated,\n\n$$\n\\mu = \\int_{0}^{n_{\\mathrm{o}}} f \\sqrt{\\frac{z}{z_{\\mathrm{m}}}} \\frac{\\mathrm{d}n}{n_{\\mathrm{o}}} - \\int_{0}^{n_{\\mathrm{o}}} g \\left( \\frac{z}{z_{\\mathrm{m}}} \\right)^{3/2} \\frac{\\mathrm{d}n}{n_{\\mathrm{o}}}\n$$\n\nbut\n\n$$\n\\sqrt{\\frac{z}{z_{\\mathrm{m}}}} = \\frac{V}{V_{\\mathrm{m}}}\n$$\n\nand from equations (7) and (4)", "timestamp": "2026-07-22T04:43:28.946648+00:00"}
{"citation_id": "19930082487", "source_url": "https://ntrs.nasa.gov/api/citations/19930082487/downloads/19930082487.pdf", "page_number": 15, "total_pages": 33, "image_filename": "19930082487_p15.jpg", "text": "NACA UW No. 1813\n\n(a) $M_o$, 0.60\n\n(b) $M_o$, 0.65\n\nFigure 2.— Simultaneously obtained pressure distributions and schlieren photographs for NACA 23015 airfoil section; $\\alpha, 2^\\circ$.", "timestamp": "2026-07-22T04:43:29.566385+00:00"}
{"citation_id": "19930093773", "source_url": "https://ntrs.nasa.gov/api/citations/19930093773/downloads/19930093773.pdf", "page_number": 19, "total_pages": 47, "image_filename": "19930093773_p19.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T04:43:31.299229+00:00"}
{"citation_id": "19930085842", "source_url": "https://ntrs.nasa.gov/api/citations/19930085842/downloads/19930085842.pdf", "page_number": 90, "total_pages": 104, "image_filename": "19930085842_p90.jpg", "text": "86\nNACA RM L9029\n\nPropeller advance\ndiameter ratio, V/nD\n6\n4\n2\n0\n\nV/nD\n\n$\\sqrt{C_{DR}}$\n\nResultant-drag coefficient, $C_{DR}$\n2\n1\n0\n\nLift coefficient, $C_L$\n14\n12\n10\n8\n6\n4\n2\n0\n\n$C_L$\n\n$\\beta$, deg\n11.5\n\nNATIONAL ADVISORY\nCOMMITTEE FOR AERONAUTICS\n\n0\n0.04\n0.08\n0.12\n0.16\n0.20\n0.24\n0.28\n0.32\n0.36\nTorque coefficient, $Q_c$\n\n(m) $\\alpha_u = 84^\\circ$.\nFigure 43.- Concluded.", "timestamp": "2026-07-22T04:43:31.501495+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 10, "total_pages": 99, "image_filename": "19930082511_p10.jpg", "text": "8\nNACA TN No. 1826\n\nof reference 8 provides an angle correction factor $\\delta$ of 0.25 for the flow far upstream of the wing, whereas $\\delta$ should be zero far upstream; a correction of -0.25 should therefore be added to all values of $\\delta$ computed by this image system for points within the tunnel.\n\nSummary of boundary conditions.- A basic physical characteristic of the flow is provided by the condition that the velocity be continuous at the entrance lip, which also helps to provide uniqueness. The velocity on the free surface is not necessarily the velocity far upstream in the closed portion; in fact, for the two-dimensional case, it is even possible for the pressures on the two free surfaces to be different from each other. Equality of the velocities in the upstream and downstream closed portions has been recognized as an additional condition. Neglecting the upper portion of the closed exit may be desirable if the flow is so depressed that it does not make contact with the upper part of the exit. Neglecting the entire closed exit region may appreciably simplify the problem without introducing excessive inaccuracy if the region of interest is much closer to the entrance than to the exit. In general, adequate treatment of the exit (for large lift on the body in the tunnel) seems very unlikely.\n\nThe discussion in the preceding sections has concerned mainly the physical flow conditions, and relatively little interpretation in terms of boundary conditions on the perturbation potential has been given, although such formal interpretation would appear a trivial task. The reason that this extension has not been made is that, in a number of instances, as will appear subsequently, slight modifications of the basic viewpoint, leading to somewhat modified boundary conditions, are desirable for convenience of solution. Accordingly, the statements of the boundary conditions on the perturbation potentials will be given when the solutions are discussed.\n\nSUGGESTED ELECTRICAL ANALOGIES\n\nVelocity-Potential Analogies\n\nBasic concepts of the analogies.- In the analogies to be discussed in the present section (none of which have yet been constructed), the perturbation velocity potential in the space within the wind tunnel is considered analogous to the electrical potential in a dilute electrolyte solution contained in a vessel of the same shape. An insulating material such as Bakelite, the conductivity of which is negligible compared with that of the solution, provides a boundary where the normal potential gradient $\\frac{\\partial \\phi}{\\partial n}$ is zero; and a metal, the conductivity of which is practically infinite relative to that of the solution, serves as a constant-potential boundary along which the longitudinal gradient $\\frac{\\partial \\phi}{\\partial x}$ is zero. In such a setup, current is analogous to velocity except for a difference in sign (in the usual convention, current flows down a voltage gradient", "timestamp": "2026-07-22T04:43:31.713530+00:00"}
{"citation_id": "19930082585", "source_url": "https://ntrs.nasa.gov/api/citations/19930082585/downloads/19930082585.pdf", "page_number": 1, "total_pages": 30, "image_filename": "19930082585_p1.jpg", "text": "TN 1907\nNACA TN 1907\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\n\nTECHNICAL NOTE 1907\n\n[Stamp: RECEIVED JUL 5 1949 N.A.C.A. - W. LOS ANGELES]\n\nAN ANALYSIS OF THE TRANSITION OF A HELICOPTER FROM\nHOVERING TO STEADY AUTOROTATIVE VERTICAL DESCENT\n\nBy A. A. Nikolsky and Edward Seckel\n\nPrinceton University\n\nTECHNICAL LIBRARY\nAIRESEARCH MANUFACTURING CO.\n9851-9951 SEPULVEDA BLVD.\nINGLEWOOD,\nCALIFORNIA\n\n[NACA Logo]\n\nWashington\nJune 1949", "timestamp": "2026-07-22T04:43:32.172182+00:00"}
{"citation_id": "19930082614", "source_url": "https://ntrs.nasa.gov/api/citations/19930082614/downloads/19930082614.pdf", "page_number": 1, "total_pages": 36, "image_filename": "19930082614_p1.jpg", "text": "NACA TN 1939\n8375\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\n\nTECHNICAL NOTE 1939\n\nTHE EFFECTS OF AERODYNAMIC BRAKES UPON THE\nSPEED CHARACTERISTICS OF AIRPLANES\n\nBy Jack D. Stephenson\n\nAmes Aeronautical Laboratory\nMoffett Field, Calif.\n\n[Figure: NACA logo]\n\nWashington\nSeptember 1949\n\n319.98/41\n\nAFMDC\nTECHNICAL LIBRARY\nAFL 2811\n\nTECH LIBRARY KAFB, NM\n0065356", "timestamp": "2026-07-22T04:43:32.390410+00:00"}
{"citation_id": "19930082542", "source_url": "https://ntrs.nasa.gov/api/citations/19930082542/downloads/19930082542.pdf", "page_number": 2, "total_pages": 53, "image_filename": "19930082542_p2.jpg", "text": "NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nTECHNICAL NOTE NO. 1867\n\nA STUDY OF EFFECTS OF HEAT TREATMENT AND HOT-COLD-WORK\nON PROPERTIES OF LOW-CARBON N-155 ALLOY\n\nBy J. W. Freeman, E. E. Reynolds, D. N. Frey, and A. E. White\n\nSUMMARY\n\nPhysical properties at room temperature and rupture test characteristics at $1200^\\circ$ F were used as criterions to evaluate the effects of systematic variations of solution treatments, aging treatments, and hot-cold-work on the properties of bar stock from one heat of low-carbon N-155 alloy. The range in yield strength for 0.02-percent offset at room temperature was from 30,000 to $13\\frac{1}{4}$,000 psi. Rupture strengths at $1200^\\circ$ F ranged from 40,000 to 66,000 psi at 100 hours and from 35,000 to 56,000 psi at 1000 hours. This rupture-strength range is equivalent to such extreme variation as 100 to approximately 600,000 hours for fracture at $1200^\\circ$ F under a stress of 40,000 psi for the same bar stock with different treatments.\n\nHot-cold-work in amount of 10- to 15-percent reduction at temperatures below $1400^\\circ$ F produced the highest strengths at both room temperature and $1200^\\circ$ F. Solution temperatures prior to hot-cold-work should be between $1950^\\circ$ and $2100^\\circ$ F. The best properties from solution treatments alone were obtained by quenching from $2100^\\circ$ F. After this solution treatment the alloy can be expected to have a yield strength of about 40,000 psi and rupture strengths at 100 and 1000 hours of about 51,000 and 40,000 psi, respectively. Aging treatments at $1350^\\circ$ to $1400^\\circ$ F developed yield strengths between 40,000 and 50,000 psi and rupture strengths of about 50,000 and 40,000 psi at 100 and 1000 hours, respectively. Incomplete data indicate that aging treatments after hot-cold-work may be quite beneficial. The hot-worked condition provides the best all-round properties provided the hot-working conditions can be properly controlled.\n\nSolution treatments alone result in the lowest strength for several hundred hours in the rupture test. Proper aging improves the shorter-time rupture strength but results in steep curves of stress against rupture time so that the aged condition is the weakest at the longer time periods. Hot-cold-work after the proper solution treatment produces the best rupture strength at both long and short time periods.\n\nSolution-treating at temperatures above $2100^\\circ$ F yields material with excessively low ductility in the rupture test. Such materials are also subject to brittleness at points of stress concentration during rupture testing. Hot-cold-work acutely magnifies these two shortcomings.", "timestamp": "2026-07-22T04:43:33.162204+00:00"}
{"citation_id": "19930085965", "source_url": "https://ntrs.nasa.gov/api/citations/19930085965/downloads/19930085965.pdf", "page_number": 56, "total_pages": 67, "image_filename": "19930085965_p56.jpg", "text": "NACA RM E9E06\n55\n\n1125\n51-1565\n\nBlade\nD\nd\nA\n4\n7\n8\n9\n3\n50°\n0\n1\n2\n6\n5\nA\nNACA\n\nFigure 12. - Geometry used in permeance calculations.", "timestamp": "2026-07-22T04:43:33.471659+00:00"}
{"citation_id": "19930082245", "source_url": "https://ntrs.nasa.gov/api/citations/19930082245/downloads/19930082245.pdf", "page_number": 23, "total_pages": 66, "image_filename": "19930082245_p23.jpg", "text": "22\nNACA TN No. 1596\n\n.2354c\n$c_a=0.20c$\n.00375c\nChord line\nAirfoil section contour\n.0354c radius\nHinge axis\n(a) True-contour aileron.\n\n.2354c\n$c_a=0.20c$\n.00375c\nChord line\n.208c radius\n.0354c radius\nStraight lines\nHinge axis\nStraight lines\n.04c\n(b) Beveled-trailing-edge aileron.\nNACA\n\nFigure 3.- Dimensions of 0.20c plain\nailerons used on NACA 66,1-115\nairfoil section. c=24 inches.", "timestamp": "2026-07-22T04:43:35.940052+00:00"}
{"citation_id": "19930082498", "source_url": "https://ntrs.nasa.gov/api/citations/19930082498/downloads/19930082498.pdf", "page_number": 7, "total_pages": 49, "image_filename": "19930082498_p7.jpg", "text": "```markdown\n6\nNACA TN No. 1838\n\nexpected from these muffler types. The automobile muffler, on the other hand, was designed for an eight-cylinder engine of lower horsepower and higher speed than the test engine. Because of these fundamental differences, the results cannot be considered representative of the performance that this muffler would give when used with the automobile engine for which it was designed.\n\nAirplane mufflers.- The commercial airplane mufflers produce very little reduction in the over-all sound-pressure level. These mufflers, however, are designed for and used on present-day light airplanes where, as has already been pointed out, the propeller noise is so high that no large reduction in exhaust noise is practical. These mufflers are shown in figure 2(a).\n\nAutomobile mufflers.- Most automobile mufflers contain internal baffles which force the exhaust-gas flow to reverse direction two or more times in passing through the mufflers. This type of muffler is unacceptable to airplane-engine manufacturers because of the high back pressures resulting from the flow reversals. Some automobile mufflers, however, are of the \"straight-through\" type, in which the stream of exhaust gas flows unobstructed through the muffler, and lower back pressure results. With an automobile muffler of the straight-through type (configuration 11, table II and fig. 1) attached to each of the two engine exhaust pipes many of the higher harmonics are reduced to below 60 decibels. The over-all level, however, which is 92 decibels at 2000 rpm, is only slightly lower than the level for the airplane mufflers because the fundamental note remains strong. A wye was attached to connect the original exhaust pipes and one of the automobile mufflers of configuration 11 was attached to the outlet of the wye. The results of this test (configuration 12, table II) show that this configuration produces a very marked reduction in the intensity of the fundamental note and reduces the over-all sound level from 92 decibels with two mufflers to 85 decibels with one muffler at 2000 rpm. For this installation one muffler attached to a collector pipe has proved to be more effective than two identical mufflers attached to separate cylinder banks. This result may be due, at least in part, to increased back pressure. The single automobile muffler, which was designed for a smaller engine, is unable to handle the large volume of exhaust-gas flow from the airplane engine without excessive back pressure; therefore this particular arrangement is considered unsatisfactory for the engine used in this investigation.\n\nSpecial Commercial Muffler\n\nA special muffler was designed for this engine by a commercial firm. This straight-through type of muffler, a three-dimensional cutaway view of which is shown as muffler 13 in figure 1, accomplished a large reduction in over-all noise. Unfortunately, however, the back pressure is high. This muffler (muffler 13, table II) had an over-all sound-pressure\n```", "timestamp": "2026-07-22T04:43:36.253192+00:00"}
{"citation_id": "19930082447", "source_url": "https://ntrs.nasa.gov/api/citations/19930082447/downloads/19930082447.pdf", "page_number": 18, "total_pages": 24, "image_filename": "19930082447_p18.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T04:43:42.580723+00:00"}
{"citation_id": "19930085838", "source_url": "https://ntrs.nasa.gov/api/citations/19930085838/downloads/19930085838.pdf", "page_number": 103, "total_pages": 118, "image_filename": "19930085838_p103.jpg", "text": "NACA RM No. L9B23\n101\n\n[Figure: A graph plotting \"Flap section hinge-moment coefficient, $c_{h1}$\" on the vertical axis against \"Section angle of attack, $\\alpha_o$, deg\" on the horizontal axis. The vertical axis ranges from $-.36$ to $.20$. The horizontal axis ranges from $-20$ to $20$. The graph contains multiple curves representing different values of $\\delta_f$ (deg) as indicated in the legend: $0, -7, -10, -15, -20$. A note in the upper right indicates $\\delta_a = -5^\\circ$. The NACA logo is present in the bottom right corner of the plot area.]\n\n(1) $\\delta_f = 40^\\circ$.\nFigure 12.- Continued.", "timestamp": "2026-07-22T04:43:42.902365+00:00"}
{"citation_id": "19930082496", "source_url": "https://ntrs.nasa.gov/api/citations/19930082496/downloads/19930082496.pdf", "page_number": 8, "total_pages": 50, "image_filename": "19930082496_p8.jpg", "text": "NACA TN No. 1836\n\nA typical ceramal gas-turbine blade is shown in figure 6. Metal blades obtained from the U. S. Air Force were used as control blades. These blades had the following nominal composition:\n\n| C | Ni | Fe | Cr | Mo | Co |\n| --- | --- | --- | --- | --- | --- |\n| 0.15-0.35 | 1.75-3.25 | 0.5-2.0 | 25.5-29.5 | 5.0-6.0 | Remainder |\n\nThe inspection of all blades upon receipt for internal and external flaws by radiographic and fluorescent-oil methods, respectively, did not reveal any flaws.\n\nThe metal blades originally had shrouds. These shrouds were ground and faired into the airfoil-section contour so that the original blade length was preserved. This grinding was done to make the configuration of the metal blades the same as that of the ceramal blades; the ceramal blades had been made without a shroud in order to simplify fabrication. The blades were installed in a turbine disk that had been previously inspected by radiographic and fluorescent-oil methods.\n\nThe investigation required two phases, each resulting in a particular setup procedure:\n\nPhase 1. - Three ceramal blades were installed at equal intervals about the disk and the remaining 139 dovetails were fitted with metal blades. The metal blades were installed in accordance with applicable U. S. Air Force technical orders. The ceramal blades were installed in a slightly different fashion; they had a root slightly narrower than the thickness of the disk, which allowed the disk material to be peened over the root, thereby locking the blade in place. Some peening was done on the roll-neck junction of the disk dovetail. The wheel was then so ground that all root protuberances of the metal blades were flush with the disk surfaces.\n\nPhase 2. - Three ceramal blades were installed to replace the ones that failed during phase 1. The root sections of these three blades were plated with copper of approximately 0.01-inch thickness. The corresponding disk dovetails were enlarged to receive these blades. In addition, the radius of the dovetail junction at the roll and neck was increased. The fit was such that the copper just became scratched when a blade was installed. The wheel was peened only at the bottom of the root roll. The wheel at this stage is shown in figure 7. Twelve new metal blades were installed at equal intervals", "timestamp": "2026-07-22T04:43:47.613998+00:00"}
{"citation_id": "19930090382", "source_url": "https://ntrs.nasa.gov/api/citations/19930090382/downloads/19930090382.pdf", "page_number": 33, "total_pages": 37, "image_filename": "19930090382_p33.jpg", "text": "NACA RM L9I07\n35\n\nCONFIDENTIAL\n\nPower coefficient, $C_P$\nThrust coefficient, $C_T$\n\nTip Mach number, $M_t$\nEfficiency, $\\eta$\n\nAdvance ratio, J\n(1) M=0.90 Concluded.\nFigure 5 - Continued.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T04:43:50.596946+00:00"}
{"citation_id": "19930082476", "source_url": "https://ntrs.nasa.gov/api/citations/19930082476/downloads/19930082476.pdf", "page_number": 17, "total_pages": 41, "image_filename": "19930082476_p17.jpg", "text": "NACA TN No. 1801\n15\n\nTABLE III.— MODEL TEST CONDITIONS\n[Erect spins to pilot's right]\n\n| Loading | Controls | Data presented in chart |\n| :--- | :--- | :--- |\n| 1 | Linked | 1 |\n| 2 | Linked | 2 |\n| 3 | Linked | 3 |\n| 4 | Linked | 4 |\n| 1 and 2 | Unlinked (effect of combined aileron deflections) | 5 |\n| 1, 2, 3, and 4 | Unlinked (effect of individual aileron deflections) | 6 |\n| 2, 3, and 4 | Unlinked (effect of individual and combined rudder deflections) | 7 |\n| 3 and 4 | Unlinked (effect of combined rudder deflections) | 8 |\n\n[Figure: NACA logo]", "timestamp": "2026-07-22T04:43:52.318185+00:00"}
{"citation_id": "19930086015", "source_url": "https://ntrs.nasa.gov/api/citations/19930086015/downloads/19930086015.pdf", "page_number": 51, "total_pages": 54, "image_filename": "19930086015_p51.jpg", "text": "```markdown\nCONFIDENTIAL\n\nTheory\nExperiment : $\\circ \\Delta \\alpha = 5.35^\\circ$ (Model Vertical)\n$\\triangle \\Delta \\alpha = 3.70^\\circ$ (Model Horizontal)\n$\\square \\Delta \\alpha_f = 5.74^\\circ$ (Model Horizontal)\n\nLoading coefficient\nper unit angle of attack, $P_\\beta$, per deg\n.18\n.16\n.14\n.12\n.10\n.08\n.06\n.04\n.02\n0\n\nPercent of local chord\n0 20 40 60 80 100\n\n(c) $M=1.50$.\n\nFigure 14.- continued.\n\nNACA\n\nCONFIDENTIAL\n\n50\n\nNACA RM A59E24\n```", "timestamp": "2026-07-22T04:43:57.392157+00:00"}
{"citation_id": "19930082487", "source_url": "https://ntrs.nasa.gov/api/citations/19930082487/downloads/19930082487.pdf", "page_number": 16, "total_pages": 33, "image_filename": "19930082487_p16.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T04:44:02.550380+00:00"}
{"citation_id": "19930086081", "source_url": "https://ntrs.nasa.gov/api/citations/19930086081/downloads/19930086081.pdf", "page_number": 43, "total_pages": 44, "image_filename": "19930086081_p43.jpg", "text": "NACA-Langley - 10-7-49 - 400\n\nCONFIDENTIAL\n\nNACA RM L9H05\n\n| $C_{N_f}$ | Theory | Experiment |\n| :--- | :--- | :--- |\n| | $a = 0^\\circ$ | $a = 0^\\circ$ |\n| | | $a = 2^\\circ$ |\n| | | $a = 4^\\circ$ |\n\n| $C_{C_f}$ | |\n| :--- | :--- |\n| | |\n\n| $C_{M_f}$ | |\n| :--- | :--- |\n| | |\n\n| $C_{BM_f}$ | |\n| :--- | :--- |\n| | |\n\n$\\delta, deg$\n\n.07t/c control\n\nCONFIDENTIAL\n\nNACA\n\nFigure 18.- Variation of the aerodynamic characteristics with deflection of two half-delta control surfaces tested in presence of a semispan delta wing.\n\n41", "timestamp": "2026-07-22T04:44:02.725417+00:00"}
{"citation_id": "19930085842", "source_url": "https://ntrs.nasa.gov/api/citations/19930085842/downloads/19930085842.pdf", "page_number": 91, "total_pages": 104, "image_filename": "19930085842_p91.jpg", "text": "NACA RM L9C29\n87\n\nPropeller advance-diameter ratio, V/nD\n10\n8\n6\nV/nD\n\nLift coefficient, $C_L$\n.4\n.3\n.2\n.1\n$C_L$\n\nTorque coefficient, $Q_c$\n0\n.004\n.008\n.012\n.016\n.020\n.024\n\nResultant drag coefficient, $C_{DR}$\n.2\n0\n-.2\n-.4\n$C_{DR}$\n\nNATIONAL ADVISORY\nCOMMITTEE FOR AERONAUTICS\n\n(a) $\\alpha_u = 5^\\circ$.\n\nFigure 44.- Variation of $C_L$, $C_{DR}$, and V/nD with $Q_c$ for $\\beta = 20^\\circ$.\nBasic model configuration; wing-tip support; all control surfaces neutral.", "timestamp": "2026-07-22T04:44:09.323608+00:00"}
{"citation_id": "19930082485", "source_url": "https://ntrs.nasa.gov/api/citations/19930082485/downloads/19930082485.pdf", "page_number": 17, "total_pages": 62, "image_filename": "19930082485_p17.jpg", "text": "16\nNACA TN No. 1810\n\n$$ \\sqrt{\\frac{Z}{Z_m}} = \\exp \\left[ - \\frac{n_o}{2\\Delta C} (C^2 - C_m^2) \\right] \\quad (21) $$\n\nlet\n\n$$ J = \\int_{C_1}^{C_2} \\exp \\left[ - \\frac{n_o}{2\\Delta C} (C^2 - C_m^2) \\right] \\frac{dC}{\\Delta C} \\quad (22) $$\n\nand\n\n$$ K = \\int_{C_1}^{C_2} \\exp \\left[ - \\frac{3n_o}{2\\Delta C} (C^2 - C_m^2) \\right] \\frac{dC}{\\Delta C} \\quad (23) $$\n\nThen equation (20) becomes\n\n$$ \\mu = fJ - gK $$\n\nor\n\n$$ f = g \\frac{K}{J} + \\frac{\\mu}{J} \\quad (24) $$\n\nEquations (22) and (23) may be evaluated for J and K for $C_1 > C_2$ by letting\n\n$$ t^2 = - \\frac{n_o C^2}{2\\Delta C} \\quad (25) $$\n\nThen,\n\n$$ C = t \\sqrt{- \\frac{2\\Delta C}{n_o}} $$\n\nand\n\n$$ \\frac{dC}{\\Delta C} = - dt \\sqrt{\\frac{-2}{\\Delta C \\ n_o}} $$", "timestamp": "2026-07-22T04:44:09.823478+00:00"}
{"citation_id": "19930093773", "source_url": "https://ntrs.nasa.gov/api/citations/19930093773/downloads/19930093773.pdf", "page_number": 20, "total_pages": 47, "image_filename": "19930093773_p20.jpg", "text": "NACA RM E9G09\n\n[Figure: View of J47 turbojet engine installed in test section of altitude wind tunnel.]\n\nFigure 1. - View of J47 turbojet engine installed in test section of altitude wind tunnel.\n\nNACA\nC-22509\n11-5-48\n\n19", "timestamp": "2026-07-22T04:44:10.122327+00:00"}
{"citation_id": "19930092013", "source_url": "https://ntrs.nasa.gov/api/citations/19930092013/downloads/19930092013.pdf", "page_number": 17, "total_pages": 21, "image_filename": "19930092013_p17.jpg", "text": "APPARATUS FOR VARYING EFFECTIVE DIHEDRAL IN FLIGHT\n13\n\n[Figure: Graph showing \"Time to double amplitude, sec\" and \"Period, sec\" vs. \"Effective dihedral, deg\". Includes a region labeled \"Region of approximately neutral stability\" and a note \"$T_n \\sim 38 \\text{ sec}$\". Legend: $\\circ$ Landing approach, $\\square$ Cruising, $\\triangle$ High speed.]\n\nFIGURE 15.—Variation of period and damping of the lateral oscillations with effective dihedral.\n\n[Figure: Time histories of typical rudder-fixed aileron rolls. Four subplots (a), (b), (c), (d) showing \"Pilot-applied aileron angle, deg\", \"Rolling velocity, deg/sec\", \"Sideslip angle, deg\", and \"Sideslip angle, deg\" vs. \"Time, sec\".]\n\n(a) $\\Gamma_e, 28.4^\\circ$.\n(b) $\\Gamma_e, 22.7^\\circ$.\n(c) $\\Gamma_e, 14.2^\\circ$.\n(d) $\\Gamma_e, 5.3^\\circ$ (normal airplane).\n\nFIGURE 16.—Time histories of typical rudder-fixed aileron rolls. Landing-approach condition.", "timestamp": "2026-07-22T04:44:11.225175+00:00"}
{"citation_id": "19930082450", "source_url": "https://ntrs.nasa.gov/api/citations/19930082450/downloads/19930082450.pdf", "page_number": 14, "total_pages": 37, "image_filename": "19930082450_p14.jpg", "text": "NACA TN No. 1778\n13\n\nTABLE 3.-Z-PLANE PROPERTIES - Concluded\n$\\frac{W}{t_B} = 0.63; \\frac{t_B}{t_W} = 10.9; \\frac{t_W}{t_F} = 0.8; \\frac{t_F}{t_B} = 3; \\frac{t_F}{t_W} = 4; \\frac{t_F}{t_B} = 1.88; \\frac{t_W}{t_B} = 10.3$\n\n| $\\frac{t_F}{t_B}$ | 33 | 34 | 35 | 36 | 37 | 38 | 39 | 40 | 41 | 42 | 43 | 44 | 45 |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| 25 | 1.853 | 1.874 | 1.897 | 1.919 | 1.941 | 1.963 | 1.986 | 2.008 | 2.030 | 2.052 | 2.075 | 2.097 | 2.119 |\n| 26 | 1.825 | 1.847 | 1.869 | 1.891 | 1.913 | 1.935 | 1.958 | 1.980 | 2.002 | 2.024 | 2.046 | 2.068 | 2.090 |\n| 27 | 1.799 | 1.820 | 1.841 | 1.861 | 1.871 | 1.892 | 1.913 | 1.935 | 1.954 | 1.974 | 1.995 | 2.015 | 2.036 |\n| 28 | 1.761 | 1.781 | 1.801 | 1.820 | 1.840 | 1.860 | 1.880 | 1.900 | 1.920 | 1.939 | 1.959 | 1.979 | 1.998 |\n| 29 | 1.734 | 1.754 | 1.772 | 1.792 | 1.811 | 1.830 | 1.850 | 1.869 | 1.888 | 1.908 | 1.927 | 1.945 | 1.964 |\n| 30 | 1.711 | 1.729 | 1.748 | 1.766 | 1.785 | 1.803 | 1.822 | 1.840 | 1.859 | 1.877 | 1.896 | 1.914 | 1.932 |\n| 31 | 1.687 | 1.706 | 1.723 | 1.741 | 1.759 | 1.777 | 1.795 | 1.813 | 1.831 | 1.849 | 1.867 | 1.884 | 1.902 |\n| 32 | 1.665 | 1.683 | 1.701 | 1.718 | 1.736 | 1.753 | 1.770 | 1.787 | 1.805 | 1.822 | 1.839 | 1.857 | 1.874 |\n| 33 | 1.644 | 1.662 | 1.679 | 1.696 | 1.713 | 1.730 | 1.747 | 1.764 | 1.781 | 1.797 | 1.814 | 1.831 | 1.847 |\n| 34 | 1.627 | 1.643 | 1.660 | 1.676 | 1.692 | 1.708 | 1.723 | 1.741 | 1.758 | 1.774 | 1.790 | 1.806 | 1.823 |\n| 35 | 1.606 | 1.623 | 1.639 | 1.655 | 1.671 | 1.687 | 1.702 | 1.718 | 1.734 | 1.750 | 1.766 | 1.782 | 1.798 |\n| 36 | 1.592 | 1.607 | 1.623 | 1.638 | 1.654 | 1.669 | 1.685 | 1.700 | 1.716 | 1.731 | 1.747 | 1.762 | 1.777 |\n| 37 | 1.575 | 1.591 | 1.606 | 1.621 | 1.636 | 1.651 | 1.666 | 1.681 | 1.696 | 1.711 | 1.726 | 1.741 | 1.756 |\n| 38 | 1.561 | 1.575 | 1.590 | 1.605 | 1.619 | 1.634 | 1.649 | 1.663 | 1.678 | 1.692 | 1.707 | 1.721 | 1.736 |\n| 39 | 1.547 | 1.560 | 1.575 | 1.589 | 1.603 | 1.617 | 1.632 | 1.646 | 1.660 | 1.674 | 1.688 | 1.702 | 1.716 |\n| 40 | 1.533 | 1.546 | 1.560 | 1.574 | 1.588 | 1.602 | 1.616 | 1.630 | 1.643 | 1.657 | 1.671 | 1.685 | 1.698 |\n| 41 | 1.505 | 1.520 | 1.534 | 1.547 | 1.560 | 1.573 | 1.587 | 1.600 | 1.613 | 1.626 | 1.640 | 1.653 | 1.666 |\n| 42 | 1.488 | 1.497 | 1.510 | 1.522 | 1.535 | 1.547 | 1.560 | 1.572 | 1.584 | 1.598 | 1.611 | 1.623 | 1.636 |\n| 43 | 1.463 | 1.475 | 1.487 | 1.499 | 1.511 | 1.523 | 1.536 | 1.548 | 1.560 | 1.572 | 1.584 | 1.596 | 1.608 |\n| 44 | 1.446 | 1.456 | 1.468 | 1.480 | 1.491 | 1.503 | 1.514 | 1.525 | 1.537 | 1.548 | 1.560 | 1.571 | 1.582 |\n| 45 | 1.426 | 1.537 | 1.448 | 1.459 | 1.471 | 1.482 | 1.493 | 1.504 | 1.515 | 1.526 | 1.537 | 1.548 | 1.560 |\n| 46 | 1.410 | 1.420 | 1.431 | 1.442 | 1.453 | 1.464 | 1.475 | 1.486 | 1.497 | 1.508 | 1.519 | 1.529 | 1.540 |\n| 47 | 1.394 | 1.404 | 1.415 | 1.425 | 1.436 | 1.446 | 1.457 | 1.467 | 1.477 | 1.488 | 1.498 | 1.508 | 1.518 |\n| 48 | 1.380 | 1.390 | 1.400 | 1.410 | 1.420 | 1.430 | 1.440 | 1.450 | 1.460 | 1.470 | 1.480 | 1.490 | 1.500 |\n| 49 | 1.368 | 1.377 | 1.387 | 1.396 | 1.406 | 1.415 | 1.425 | 1.434 | 1.444 | 1.453 | 1.463 | 1.472 | 1.482 |\n| 50 | 1.356 | 1.364 | 1.374 | 1.383 | 1.392 | 1.401 | 1.411 | 1.420 | 1.429 | 1.438 | 1.448 | 1.457 | 1.466 |\n| 55 | 1.298 | 1.304 | 1.311 | 1.318 | 1.325 | 1.332 | 1.339 | 1.346 | 1.353 | 1.360 | 1.367 | 1.374 | 1.381 |\n| 60 | 1.254 | 1.259 | 1.265 | 1.271 | 1.277 | 1.283 | 1.289 | 1.295 | 1.301 | 1.307 | 1.313 | 1.319 | 1.325 |\n| 65 | 1.225 | 1.229 | 1.234 | 1.238 | 1.243 | 1.247 | 1.252 | 1.256 | 1.261 | 1.265 | 1.270 | 1.274 | 1.279 |\n| 70 | 1.201 | 1.205 | 1.209 | 1.213 | 1.217 | 1.221 | 1.225 | 1.229 | 1.233 | 1.237 | 1.241 | 1.245 | 1.249 |\n| 75 | 1.185 | 1.189 | 1.193 | 1.196 | 1.200 | 1.204 | 1.207 | 1.211 | 1.215 | 1.218 | 1.222 | 1.226 | 1.229 |\n\n| $\\frac{t_F}{t_B}$ | 33 | 34 | 35 | 36 | 37 | 38 | 39 | 40 | 41 | 42 | 43 | 44 | 45 |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| 25 | 5.314 | 6.076 | 6.532 | 6.609 | 6.569 | 7.111 | 7.412 | 7.690 | 7.921 | 8.124 | 8.307 | 8.472 | 8.622 |\n| 26 | 5.702 | 6.537 | 7.047 | 7.140 | 7.088 | 7.676 | 8.011 | 8.321 | 8.574 | 8.798 | 9.000 | 9.182 | 9.350 |\n| 27 | 5.927 | 6.845 | 6.097 | 6.356 | 6.818 | 6.873 | 7.152 | 7.416 | 7.686 | 7.863 | 8.239 | 8.322 | 8.803 |\n| 28 | 5.493 | 5.736 | 5.957 | 6.244 | 6.499 | 6.752 | 7.000 | 7.285 | 7.532 | 7.827 | 8.076 | 8.316 | 8.548 |\n| 29 | 5.296 | 5.637 | 5.882 | 6.132 | 6.385 | 6.641 | 6.897 | 7.139 | 7.465 | 7.697 | 7.924 | 8.147 | 8.365 |\n| 30 | 5.121 | 5.350 | 5.581 | 5.823 | 6.068 | 6.317 | 6.567 | 6.812 | 7.058 | 7.304 | 7.550 | 7.796 | 8.042 |\n| 31 | 5.207 | 5.442 | 5.681 | 5.923 | 6.168 | 6.416 | 6.665 | 6.922 | 7.179 | 7.444 | 7.707 | 7.973 | 8.241 |\n| 32 | 5.121 | 5.350 | 5.581 | 5.823 | 6.068 | 6.317 | 6.567 | 6.812 |", "timestamp": "2026-07-22T04:44:13.627438+00:00"}
{"citation_id": "19930085965", "source_url": "https://ntrs.nasa.gov/api/citations/19930085965/downloads/19930085965.pdf", "page_number": 57, "total_pages": 67, "image_filename": "19930085965_p57.jpg", "text": "56\n\nMagnetic-flux density, gausses\n15,000\n10,000\n5,000\n0\n\nMagnetic-field intensity, oersteds\n0 10 20 30 40 50\n\n[Figure: Normal magnetization curve for Armco Magnetic Ingot Iron. The curve starts at the origin (0,0) and rises steeply before leveling off. Points labeled A, B, C, D, E, F, and G are marked on or near the curve.]\n\nFigure 13. - Normal magnetization curve for Armco Magnetic Ingot Iron.\n(Data from reference 7.)\n\nNACA\nNACA RM E9E06\n1125", "timestamp": "2026-07-22T04:44:13.836749+00:00"}
{"citation_id": "19930082447", "source_url": "https://ntrs.nasa.gov/api/citations/19930082447/downloads/19930082447.pdf", "page_number": 19, "total_pages": 24, "image_filename": "19930082447_p19.jpg", "text": "NACA TN No. 1775\n17\n\n<!-- Image (171, 212, 893, 774) -->\n\nFigure 3.-Variation of load-factor coefficient with approach parameter.", "timestamp": "2026-07-22T04:44:14.347910+00:00"}
{"citation_id": "19930082245", "source_url": "https://ntrs.nasa.gov/api/citations/19930082245/downloads/19930082245.pdf", "page_number": 24, "total_pages": 66, "image_filename": "19930082245_p24.jpg", "text": "NACA TN No. 1596\n\n$10 \\times 10^6$\n\nReynolds number, R\n\n[Figure: Graph showing variation of Reynolds number with Mach number, with a curve rising from approximately (0.25, 3.5) to (0.75, 7.5), and a NACA logo in the lower right corner of the plot area.]\n\nMach number, M\n\nFigure 4.—Variation of Reynolds number with Mach number in the wind-tunnel tests of an NACA 66,1-115 airfoil section equipped with 0.20c plain ailerons.\n\n23", "timestamp": "2026-07-22T04:44:14.707050+00:00"}
{"citation_id": "19930090382", "source_url": "https://ntrs.nasa.gov/api/citations/19930090382/downloads/19930090382.pdf", "page_number": 34, "total_pages": 37, "image_filename": "19930090382_p34.jpg", "text": "36\nNACA RM L9I07\n\nCONFIDENTIAL\n\nTip Mach number, $M_t$\nEfficiency, $\\eta$\n\nPower coefficient, $C_P$\nThrust coefficient, $C_T$\n\nAdvance ratio, J\n$\\beta_{0.7R}=65^\\circ$\n(in) M=0.925.\n\nFigure 5 - Concluded.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T04:44:16.470887+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 11, "total_pages": 99, "image_filename": "19930082511_p11.jpg", "text": "NACA TN No. 1826\n\nwhereas air flows up a velocity-potential gradient); in order to remove this difficulty, the sign convention for electrical potential is reversed in the following discussion.\n\nFor greater clarity of exposition, the two-dimensional analogies are treated in detail, the three-dimensional analogies appearing as reasonably obvious extensions or modifications. It will be remembered, however, that any application will be found in three-dimensional problems, inasmuch as most of the two-dimensional problems can be solved analytically.\n\nA two-dimensional vortex may be represented by two long metal plates separated by a thin insulator. (See fig. 6(a).) A flow corresponding to a vortex located at the edge of the plates is set up by applying a difference of potential across the plates. If the perturbation flow that results from the presence of the vortex in the tunnel has a horizontal velocity component, this representation is no longer adequate because it requires the potential to be uniform along each plate. Rigor in this case would require that the plates be composed of a number of separate sections, with each pair separately activated. (See fig. 6(b).) In this way it is possible to provide a potential difference between upper and lower surfaces that is everywhere equal to the desired circulation, without requiring that the potential be uniform along the entire upper surface or lower surface. A horizontal velocity component normally occurs only when the lifting vortex is asymmetrically located in the tunnel. For simplification, only the simpler representation of figure 6(a) is used in the remaining sketches.\n\nThe element of lift in three-dimensional flow is the horseshoe vortex of zero span, which is the same as a semi-infinite line of doublets. It may be represented by a pair of long narrow metal strips separated by an insulator. (See fig. 6(c).) As in the two-dimensional analogy, if the lifting element is asymmetrically located in the field the strips must be made up of short pieces, with each pair separately activated. The horseshoe vortex of finite span is represented as in figure 6(d), provided there are no appreciable perturbation velocities in its plane.\n\nEvaluation of interference velocities.- The vertical velocity component in the tunnel corresponds to the vertical voltage gradient in the electrolyte, which can be determined by measuring the voltage difference between a pair of short wire electrodes mounted one above the other a fixed distance apart. The tunnel interference at any point is found by measuring this voltage difference (relative to that across the two plates representing the vortex) first in the simulated tunnel and then in a large tank for which the boundary interference is either negligible or so small that it can be adequately computed by simple methods. Since the theoretical flow field for the second case is known, the ratio of these two gradients, together with the distance from the pair of wires to the lifting vortex, should suffice to evaluate the boundary interference. The distance between the pair of wires need not be measured because only the ratio of the gradients is required. Similarly, the exact design and dimensions of the simulated lifting vortex are of no", "timestamp": "2026-07-22T04:44:16.555426+00:00"}
{"citation_id": "19930082585", "source_url": "https://ntrs.nasa.gov/api/citations/19930082585/downloads/19930082585.pdf", "page_number": 2, "total_pages": 30, "image_filename": "19930082585_p2.jpg", "text": "NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nTECHNICAL NOTE 1907\n\nAN ANALYSIS OF THE TRANSITION OF A HELICOPTER FROM\nHOVERING TO STEADY AUTOROTATIVE VERTICAL DESCENT\n\nBy A. A. Nikolsky and Edward Seckel\n\nSUMMARY\n\nAn analytical study is presented of the transition from hovering\nflight (power on) to steady vertical descent (in autorotation) following\npower failure while hovering. The effects of hinging the blades, of blade\nmoment of inertia, and of rapidity of pitch reduction after power failure\nare investigated.\n\nThe results indicate that the effect of blade flapping is negligible\nas far as the establishment of steady autorotation is concerned. From\nthe standpoint of avoiding excessive blade stalling during the transition,\nit is desirable that the blade pitch be reduced as rapidly as possible\nafter power failure, and that the blade moment of inertia be large.\n\nSample calculations and graphs of computed transitions for a typical\nhelicopter are presented.\n\nINTRODUCTION\n\nThis report is the second phase of a broad program of study of the\ntransient motions of helicopters in autorotative flight. The first\nphase (reference 1) dealt with the steady-state condition of autorotative\nvertical descent. This report is concerned with the transition from the\nsteady condition of hovering to the steady autorotative descent.\n\nIt has long been realized that the ability of a helicopter to\nestablish steady autorotation following power failure would be an\nimportant safety feature. Although some helicopters have successfully\ndemonstrated such an ability, there has been no analytical method avail-\nable for assessing the influence of the design variables on the ease and\nsafety of performing such a maneuver.\n\nSuch a method, based on numerical integration of the equations of\nmotion, is presented in this report.", "timestamp": "2026-07-22T04:44:18.955784+00:00"}
{"citation_id": "19930085838", "source_url": "https://ntrs.nasa.gov/api/citations/19930085838/downloads/19930085838.pdf", "page_number": 104, "total_pages": 118, "image_filename": "19930085838_p104.jpg", "text": "102\nNACA RM No. L9B23\n\n[Figure: A graph plotting \"Flap section hinge-moment coefficient, $c_{h_f}$\" on the y-axis against \"Section angle of attack, $\\alpha_o$, deg\" on the x-axis. The y-axis ranges from +.20 to -.32. The x-axis ranges from -16 to 20. The graph contains multiple curves with different symbols. A legend is present in the upper right quadrant. A NACA logo is in the bottom right corner.]\n\nFlap section hinge-moment coefficient, $c_{h_f}$\n\n| $\\delta_a$ (deg) | $\\delta_1$ (deg) |\n| :--- | :--- |\n| $\\circ$ -10 | 0 |\n| $\\square$ -10 | -15 |\n| $\\diamond$ -10 | +20 |\n| $\\triangle$ -10 | +25 |\n| $\\nabla$ -15 | 0 |\n| $\\blacktriangleright$ -15 | -25 |\n\nSection angle of attack, $\\alpha_o$, deg\n\n(j) $\\delta_f = 40^\\circ$\nFigure 12.- Concluded.", "timestamp": "2026-07-22T04:44:19.828873+00:00"}
{"citation_id": "19930082542", "source_url": "https://ntrs.nasa.gov/api/citations/19930082542/downloads/19930082542.pdf", "page_number": 3, "total_pages": 53, "image_filename": "19930082542_p3.jpg", "text": "2\nNACA TN No. 1867\n\nAging treatments are especially beneficial to ductility in the rupture test and to alleviation of the stress-concentration brittleness.\n\nA rough correlation exists between Brinell hardness and yield strengths and rupture strengths. The correlations, together with the specific trends shown by the detailed data, permit quite close predictions of the properties of large forgings. Large forgings have properties which are slightly lower than those of the bar stock studied.\n\nWide ranges in properties of most of the better alloys developed for gas-turbine service have been due to the influence of heat treatment and processing conditions on properties. The treatments used have been of more influence than wide ranges in chemical composition. The properties of alloys in the hot-worked condition are quite variable because hot-working simultaneously involves solution treatments, aging, and hot-cold-work; and it is very difficult to control these under normal hot-working conditions.\n\nIt is expected that the trends shown in this investigation for the various treatments will hold for other alloys. It is not expected, however, that the optimum treatments will be the same for all alloys. A certain amount of test work on any other alloy will have to be done to establish the best conditions of treatment. It is also quite probable that the influence of the treatments on rupture strength may not be the same as their influence on the strengths for limited deformations.\n\nInsofar as could be determined, both precipitation reactions and strain hardening control the properties. High strength at room temperature and for short time periods at $1200^\\circ$ F probably depends on the presence of strain hardening. High long-time strength at $1200^\\circ$ F is dependent on aging during testing. Hot-cold-work at temperatures below $1400^\\circ$ F provides the best preparation for aging to high strength during testing at $1200^\\circ$ F. Aging treatments to be effective in a reasonable time period require a higher temperature treatment than $1200^\\circ$ F. The resulting precipitation has very little effect on room-temperature strength, improves the rupture strength at shorter time periods, but results in a very low strength at prolonged time periods.\n\nINTRODUCTION\n\nA program of research on heat-resistance alloys has been in progress at the University of Michigan for the National Advisory Committee for Aeronautics. At the outset of this program the major emphasis was placed on a search for new alloys with outstanding properties for gas turbines on the basis of chemical composition. From this work several alloys appeared to have outstanding properties. Further experience with these alloys indicated that their properties were quite dependent on the schedule used in processing and heat-treating. (See references 1, and 3.)", "timestamp": "2026-07-22T04:44:21.277207+00:00"}
{"citation_id": "19930082476", "source_url": "https://ntrs.nasa.gov/api/citations/19930082476/downloads/19930082476.pdf", "page_number": 18, "total_pages": 41, "image_filename": "19930082476_p18.jpg", "text": "16\nNACA TN No. 1801\n\nCHART 1.- SPIN CHARACTERISTICS OF MODEL FOR NORMAL LOADING (LINKED RUDDER AND AILERON CONTROLS)\n$$ \\left[ \\frac{I_x - I_y}{mb^2} = -3 \\times 10^{-3}; \\mu = 5.04 \\text{ (loading 1 in table II and point 1 in fig. 4); right erect spins} \\right] $$\n\n| | Wheel setting | | | | |\n| :--- | :---: | :---: | :---: | :---: | :---: |\n| | **Left** | **0** | **$\\frac{1}{2}$** | **$\\frac{1}{2}$** | **$\\frac{1}{2}$** | **Full** |\n| | | | | **a** | **c** | **a** |\n| **Up** | **20** | No spin | | | 21 5D <br> 31 16D <br> 145 | |\n| | **13** | No spin | No spin | **b** | 21 8U <br> 33 11D <br> 167 0.63 | **b** |\n| **Elevator setting, degrees** | **8** | | | | 28 1D <br> 175 0.77 | |\n| | **6** | | | **b** | | |\n| | **5** | | | | No spin | |\n| | **0** | No spin | | | No spin | No spin |\n| **Down** | **12** | No spin | | | No spin | No spin |\n\n$^a$Steep spin, vertical velocity too high to permit obtaining test data.\n$^b$Steep spiral.\n$^c$Oscillatory spin, range of values or average value given.\n\nNACA\n\nModel values converted to corresponding full-scale values.\nU Inner wing up\nD Inner wing down\n\n| $\\alpha$ (deg) | $\\beta$ (deg) |\n| :---: | :---: |\n| V (fps) | $\\Omega$ (rps) |", "timestamp": "2026-07-22T04:44:27.953475+00:00"}
{"citation_id": "19930086015", "source_url": "https://ntrs.nasa.gov/api/citations/19930086015/downloads/19930086015.pdf", "page_number": 52, "total_pages": 54, "image_filename": "19930086015_p52.jpg", "text": "CONFIDENTIAL\n\nTheory —\nExperiment : ○ $\\Delta \\alpha = 5.35^\\circ$ (Model Vertical)\n△ $\\Delta \\alpha = 3.70^\\circ$ (Model Horizontal)\n□ $\\Delta \\alpha = 5.74^\\circ$ (Model Horizontal)\n\nLoading coefficient\nper unit angle of attack, $\\beta_0$, per deg\n\n.18\n.16\n.14\n.12\n.10\n.08\n.06\n.04\n.02\n0\n\nPercent of local chord\n0    20    40    60    80    100\n\n(d) $M = 1.60$.\nFigure 14— continued.\n\nNACA\n\nNACA RM A9E24\nCONFIDENTIAL\n51", "timestamp": "2026-07-22T04:44:29.413502+00:00"}
{"citation_id": "19930082496", "source_url": "https://ntrs.nasa.gov/api/citations/19930082496/downloads/19930082496.pdf", "page_number": 9, "total_pages": 50, "image_filename": "19930082496_p9.jpg", "text": "8\nNACA TN No. 1836\n\nabout the wheel periphery to serve as a control for the three new\nceramal blades. Remaining in the wheel were 127 of the original\nmetal blades. All unnecessary turbine-case components that were in\nthe plane of wheel rotation or in the path of flying fragments were\nremoved at this time.\n\nThe turbine was operated in the following manner for both phases.\nCombustion air was supplied, and the turbine was motored at approxi-\nmately 6000 rpm for 5 minutes. The motoring was done in order to\nmeet a safety requirement for a 5-minute preoperation scavenge (motor\ntime). Combustion was then initiated and test conditions attained in\napproximately 3 minutes (power time). The wheel was then operated at\nrequired conditions (condition time) until a blade failure occurred.\nSpeed was maintained within $\\pm$200 rpm and temperature within $\\pm$15° F.\nFailure was indicated by a change in pitch of sound from the unit.\nUpon failure, combustion was discontinued and air flow was reduced to\na value at which the turbine motored at approximately 6000 rpm. This\nflow was maintained for approximately 10 minutes in order to cool the\nassembly and the turbine was then removed for overhaul.\n\nAll fractured blades were removed and replaced with new metal\nblades. All severely cracked metal blades were considered to be\nfailures and were replaced; this replacement was made because failure\nin this instance was imminent and the blades were, for practical\npurposes, fractured; additional complete failures would result in\nmore shutdowns and also increase the risk of a flying fragment\ninjuring sound blades. All metal replacement blades were installed\nin accordance with applicable U. S. Air Force technical orders.\n\nRecords were kept of operating conditions and of blade condition\nat each overhaul.\n\nX-Ray-Diffraction Study\n\nComparison of the structure of the material as received and\nafter operation was made by X-ray-diffraction studies. Debye-Scherrer\npowder cameras, 114.6 and 143.2 millimeters in diameter, were used\nwith filtered cobalt and copper radiations in obtaining diffraction\npatterns of the ceramal in the \"as-received\" condition and of ceramal\nblades after failure in the quasi-service investigation. Identifi-\ncation of components was made by comparing spacings of the pattern\nobtained with the interplanar spacing of compounds of the various\nelements listed in the literature and in the A.S.T.M. card index.", "timestamp": "2026-07-22T04:44:33.473882+00:00"}
{"citation_id": "19930086081", "source_url": "https://ntrs.nasa.gov/api/citations/19930086081/downloads/19930086081.pdf", "page_number": 44, "total_pages": 44, "image_filename": "19930086081_p44.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T04:44:34.646290+00:00"}
{"citation_id": "19930085842", "source_url": "https://ntrs.nasa.gov/api/citations/19930085842/downloads/19930085842.pdf", "page_number": 92, "total_pages": 104, "image_filename": "19930085842_p92.jpg", "text": "88\nNACA RM L9C29\n\n<!-- Image (100, 109, 837, 856) -->\n\n(b) $\\alpha_u = 11^\\circ$.\nFigure 44.- Continued.", "timestamp": "2026-07-22T04:44:35.377596+00:00"}
{"citation_id": "19930082487", "source_url": "https://ntrs.nasa.gov/api/citations/19930082487/downloads/19930082487.pdf", "page_number": 17, "total_pages": 33, "image_filename": "19930082487_p17.jpg", "text": "NACA TN No. 1813\n15\n\n[Figure: Schlieren photograph of a wing model in a wind tunnel, showing shock waves and flow patterns.]\n\n[Figure: Graph plotting Pressure coefficient, P (y-axis, range -2.4 to 1.6) against Chordwise station, x/c (x-axis, range 0 to 1.0). The graph shows two curves: one solid line and one dashed line labeled M, 1.0. An arrow points to a peak on the solid curve labeled Crest. The caption below reads: (c) $M_0$, 0.70]\n\n[Figure: Graph plotting Pressure coefficient, P (y-axis) against Chordwise station, x/c (x-axis, range 0 to 1.0). The graph shows two curves: one solid line and one dashed line labeled M, 1.0. An arrow points to a peak on the solid curve labeled Crest. The caption below reads: (d) $M_0$, 0.73]\n\nFigure 2. - Concluded.\n\nNACA", "timestamp": "2026-07-22T04:44:35.677462+00:00"}
{"citation_id": "19930082498", "source_url": "https://ntrs.nasa.gov/api/citations/19930082498/downloads/19930082498.pdf", "page_number": 8, "total_pages": 49, "image_filename": "19930082498_p8.jpg", "text": "NACA TN No. 1838\n\nlevel of 83.5 decibels at 2000 rpm, with high back pressure. Another, and supposedly identical, muffler (muffler 14, table II) had a slightly lower back pressure and a higher sound level (88.5 db at 2000 rpm) than the first. Mufflers 13 and 14 and also muffler 12, discussed in the preceding section, have excessive back pressure even though they are straight-through types. The back pressure can be reduced by increasing the diameter of the central tube. The equations presented in reference 1 show, however, that if the central-tube diameter is increased to lower the back pressure, the outside diameter of the muffler must be increased in approximately the same proportion to avoid loss in attenuation. Consequently, the lower the required back pressure, the larger the muffler must be.\n\nBurgess-Farnborough Silencer, Type K\n\nA Burgess-Farnborough silencer, type K (muffler 16 of table II and fig. 1) was designed to silence one of two banks of cylinders on a 450-horsepower engine (reference 3). When installed on the engine used in this investigation it reduced the sound-pressure level to 83 decibels at 2000 rpm. This muffler is the most effective one so far discussed, but it is about 8 feet long. Because of the generous proportions of this muffler the internal velocities are low. The highest internal dynamic pressure is about one-fourth that in the $2\\frac{3}{4}$-inch exhaust pipe. The flow losses in the muffler, consequently, are small and the back pressure is quite low in spite of the baffle at the rear of the second chamber. When the rear half of this muffler was used alone (muffler 17, table II) the sound level increased from 83 to 92 decibels at 2000 rpm. The large chamber at the front of the muffler therefore apparently produces most of the silencing at the lower frequencies. At the lowest engine speed, 1650 rpm, where the exhaust frequencies are low, more noise passes through the muffler than at the higher engine speeds.\n\nNACA-Designed Mufflers\n\nInasmuch as none of the mufflers discussed so far seemed wholly satisfactory for the test engine, an attempt was made to design a more suitable muffler. In the course of this investigation a large number of configurations were tried in order to study the basic characteristics of resonant-chamber and expansion-chamber mufflers.\n\nResonant-chamber types.- The development of muffler 19 (table II and fig. 1) proved especially informative because of the significant muffler properties demonstrated. This muffler was originally expected to work largely through the absorption of sound-pressure waves by the steel-wool packing. The results show that it reduces the sound-pressure level to 89 decibels at 2000 rpm and that its effectiveness is least at the lower frequencies. Closing the exit from the outer steel-wool chamber (muffler 20, table II) sharply increases the low-frequency", "timestamp": "2026-07-22T04:44:36.984690+00:00"}
{"citation_id": "19930093773", "source_url": "https://ntrs.nasa.gov/api/citations/19930093773/downloads/19930093773.pdf", "page_number": 21, "total_pages": 47, "image_filename": "19930093773_p21.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T04:44:40.241305+00:00"}
{"citation_id": "19930085965", "source_url": "https://ntrs.nasa.gov/api/citations/19930085965/downloads/19930085965.pdf", "page_number": 58, "total_pages": 67, "image_filename": "19930085965_p58.jpg", "text": "1125\n\nNACA RM E9E06\n\nMagnetic-flux density, gausses\n\n15,000\nF\nD\nB\nC\nE\nG\n\n10,000\n\n5,000\n\n0\nA\n0\n10\n20\n30\n40\n50\n\nMagnetic-field intensity, oersteds\n\nFigure 14. - Magnetization curve for Hipernik.\n(Data from reference 8.)\n\nNACA\n\n57", "timestamp": "2026-07-22T04:44:43.582686+00:00"}
{"citation_id": "19930090382", "source_url": "https://ntrs.nasa.gov/api/citations/19930090382/downloads/19930090382.pdf", "page_number": 35, "total_pages": 37, "image_filename": "19930090382_p35.jpg", "text": "NACA RM L9I07\n37\n\nCONFIDENTIAL\n\nMaximum efficiency, $\\eta_{max}$\nForward Mach number, M\n\n$\\beta_{0.7}R = 20^\\circ$\n25°\n30°\n35°\n40°\n45°\n50°\n55°\n60°\n65°\n70°\n65°\n70°\n60°\n65°\n70°\n\nCONFIDENTIAL\n\nFigure 6.- Effect of forward Mach number on maximum efficiency.\n\nNACA", "timestamp": "2026-07-22T04:44:46.538427+00:00"}
{"citation_id": "19930082245", "source_url": "https://ntrs.nasa.gov/api/citations/19930082245/downloads/19930082245.pdf", "page_number": 25, "total_pages": 66, "image_filename": "19930082245_p25.jpg", "text": "```markdown\n24\n\nPressure coefficient, $p$\n\n| | | |\n| :--- | :--- | :--- |\n| [Graph: M=0.250] | [Graph: M=0.548] | [Graph: M=0.673] |\n| | | |\n| [Graph: M=0.697] | [Graph: M=0.720] | [Graph: M=0.762] |\n\n--- Upper surface\n--- Lower surface\n\n$P_{cr}$\n\nNACA\n\n$x/c$\n\nFigure 5.- Pressure distribution about an NACA 66,1-115 airfoil section equipped with an unsealed 0.20c plain aileron of true-airfoil-contour profile. $\\delta_a=0^\\circ$; $\\alpha=1^\\circ$.\n\nNACA TN NO. 1596\n```", "timestamp": "2026-07-22T04:44:48.260821+00:00"}
{"citation_id": "19930082485", "source_url": "https://ntrs.nasa.gov/api/citations/19930082485/downloads/19930082485.pdf", "page_number": 18, "total_pages": 62, "image_filename": "19930082485_p18.jpg", "text": "NACA TN No. 1810\n17\n\nThen,\n\n$$\nJ = \\frac{\\exp\\left(\\frac{n_o C_m^2}{2\\Delta C}\\right)}{\\sqrt{-\\frac{n_o \\Delta C}{2}}} \\int_{t_2}^{t_1} \\exp\\left(-\\frac{n_o C^2}{2\\Delta C}\\right) dt\n$$\n\n$$\n= \\frac{\\exp\\left[-\\frac{t_m^2}{2}\\right]}{\\sqrt{-\\frac{\\Delta C n_o}{2}}} \\int_{t_2}^{t_1} \\exp(t^2) dt \\tag{26}\n$$\n\nand\n\n$$\nK = \\frac{\\exp\\left[-\\frac{3}{2} t_m^2\\right]}{\\sqrt{-\\frac{3}{2} \\Delta C n_o}} \\int_{\\sqrt{3} t_2}^{\\sqrt{3} t_1} \\exp(3t^2) d\\sqrt{3}t \\tag{27}\n$$\n\nThe terms $t_1$ and $t_2$ are known from the channel dimensions and a table of $\\int_0^x \\exp(t^2)dt$ can be found in reference 7.\n\nEvaluation of velocity at channel center. - The factors $f$ and $g$ are functions only of $\\sqrt{Z_m}$ and $\\gamma$ (equations (15) and (16)). Thus, for given values of $\\sqrt{Z_m}$ and $\\gamma$, curves of $f$ plotted against $\\sqrt{Z_m}$ and $g$ can be made, as shown in figure 14. In order to evaluate $\\sqrt{Z_m}$, a line is drawn at $\\tan^{-1} \\frac{K}{J}$ intersecting the ordinate of the plot of $f$ against $g$ at $\\frac{1}{J}$ (fig. 14). The line will, in general, intersect the curve at two points, one intersection corresponding to the subsonic solution and the other, the supersonic solution. If the line is tangent to the curve, the channel is choked. If the line lies above the curve, no solution exists inasmuch as $\\mu$ is larger than physically possible. The value of $\\sqrt{Z_m}$ is then found by drawing a horizontal line from the subsonic or supersonic intersection point, whichever is desired, to the curve of $f$ against $\\sqrt{Z_m}$; then $\\sqrt{Z_m}$ is read on the abscissa as shown in figure 14.", "timestamp": "2026-07-22T04:44:50.491767+00:00"}
{"citation_id": "19930085838", "source_url": "https://ntrs.nasa.gov/api/citations/19930085838/downloads/19930085838.pdf", "page_number": 105, "total_pages": 118, "image_filename": "19930085838_p105.jpg", "text": "NACA RM No. L9B23\n103\n\n[Figure: Graph plotting Section drag coefficient, $c_d$ (y-axis) against Section lift coefficient, $c_l$ (x-axis). The graph contains three curves labeled with $\\delta_f$ (deg) values of 40, 25, and 0. The NACA logo is present in the bottom right corner of the plot area.]\n\nFigure 13.- Drag characteristics of the approximately 17.7 percent-chord thick NACA 7-series-type airfoil with double slotted flap, straight-sided Frise aileron, and flip. $R = 6 \\times 10^6$ (approx.); $\\delta_a = 0^\\circ$, $\\delta_i = 0^\\circ$.", "timestamp": "2026-07-22T04:44:52.231417+00:00"}
{"citation_id": "19930082585", "source_url": "https://ntrs.nasa.gov/api/citations/19930082585/downloads/19930082585.pdf", "page_number": 3, "total_pages": 30, "image_filename": "19930082585_p3.jpg", "text": "2\nNACA TN 1907\n\nValidity of the analysis presented here is limited to cases where blade stalling can be neglected throughout the maneuver, and therefore where the rotor is in no danger of stopping. Nevertheless the investigation is sufficient to indicate the effects of certain important variables.\n\nThis work was conducted at Princeton University under the sponsorship and with the financial assistance of the National Advisory Committee for Aeronautics.\n\nSYMBOLS\n\nPhysical Quantities\n\n| Symbol | Definition |\n| :--- | :--- |\n| W | gross weight of helicopter, pounds |\n| b | number of blades per rotor |\n| R | blade radius, feet |\n| r | radial distance to blade element, feet |\n| c | blade-section chord, feet |\n| $c_e$ | equivalent blade chord, feet $\\left( c_e = \\frac{\\int_0^R cr^2 \\, dr}{\\int_0^R r^2 \\, dr} \\right)$ |\n| $\\sigma$ | rotor solidity ratio $\\left( \\frac{bc_e}{\\pi R} \\right)$ |\n| $\\theta$ | blade-section pitch angle from zero lift, radians unless otherwise stated |\n| $I_1$ | mass moment of inertia of blade about flapping hinge, slug-feet$^2$ |\n| $\\rho$ | mass density of air, slugs per cubic foot |\n| t | time, seconds |\n| g | acceleration of gravity (32.2 ft/sec$^2$) |", "timestamp": "2026-07-22T04:44:53.670768+00:00"}