PDF/ocr_dataset_1057.jsonl

{"citation_id": "19930082546", "source_url": "https://ntrs.nasa.gov/api/citations/19930082546/downloads/19930082546.pdf", "page_number": 20, "total_pages": 65, "image_filename": "19930082546_p20.jpg", "text": "NACA TN No. 1870\n19\n\nFUSELAGE RESPONSE TO OSCILLATING PRESSURES\n\nVibration\n\nTheory and experiments have been discussed which make possible the prediction of the oscillating pressures acting on the fuselage. The present discussion deals with the fuselage response to these pressures and indicates some of the factors to be considered in solving the problem of fuselage vibration and noise. Since references 2 and 3 consider in detail the acoustical treatment for aircraft fuselages, no experiments were made on soundproofing. Some amplitude and frequency measurements, however, were made on vibration of these panels which were subjected to pressure impulses from propellers.\n\nExperimental data.- The test panels were designed primarily as reflectors and were not intended for use in vibration studies. Thus, heavy construction was used in order to minimize the effect of panel vibration on the pressure measurements. The panel weights were approximately 8 pounds per square foot for the flat wall and approximately 5.5 pounds per square foot for the circular wall. This is appreciably greater than the normal fuselage weight of about 1 pound per square foot. Despite these weight differences the vibration data taken during the course of these tests are of interest in that they indicate the way in which the vibration amplitudes are affected by panel resonances.\n\nFigure 22(a) gives the vibration response of the flat wooden panel at position of greatest vibration amplitude both before and after reinforcing. As a result of excitation by a two-blade propeller a resonance peak occurred at 130 cycles per second. Reinforcing the panel removed the resonant condition from the operating range. The response curve for the circular steel panel figure 22(b) shows a narrow resonance peak at 107 cycles per second. The steel shell has a more narrow frequency response than the wooden panel and thus indicates less damping. The peak amplitude of the circular wall is less than that for the flat wall even though the flat wall had more damping. Thus it is indicated that pressures on the circular wall are less than those on the flat wall. This is further indicated by the curves for the reinforced walls, because the flat wall has about twice the amplitude of the circular shell. Figures 22(a) and 22(b) indicate the necessity of removing any large wall resonances from the operating range. They also indicate that a curved wall has less vibration amplitude than a flat wall for comparable tip clearance and operating conditions.\n\nResponse of the reinforced flat wooden panel to excitation by a four-blade propeller, which absorbs slightly less power than the two-blade propeller of figures 22(a) and 22(b), is shown in figure 22(c). A number of small resonance peaks appear in this figure; however, the over-all value of the amplitude is considerably less than for the two-blade", "timestamp": "2026-07-22T06:08:38.546691+00:00"}
{"citation_id": "19930086076", "source_url": "https://ntrs.nasa.gov/api/citations/19930086076/downloads/19930086076.pdf", "page_number": 48, "total_pages": 50, "image_filename": "19930086076_p48.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:08:40.045136+00:00"}
{"citation_id": "19930093769", "source_url": "https://ntrs.nasa.gov/api/citations/19930093769/downloads/19930093769.pdf", "page_number": 32, "total_pages": 39, "image_filename": "19930093769_p32.jpg", "text": "NACA RM No. E8L10a\nCONFIDENTIAL\n31\n\n[Figure: A line graph plotting Altitude (ft) against Engine speed (rpm). The Y-axis ranges from 10,000 to 60,000 ft. The X-axis ranges from 0 to 10,000 rpm. The graph contains four data series with markers and lines representing different flight conditions.]\n\n| Flight Mach number | Fuel |\n| :--- | :--- |\n| 0.25 | AN-F-58 |\n| .25 | Gasoline |\n| .60 | AN-F-58 |\n| .60 | Gasoline |\n\nFigure 8. - Low engine-speed blow-out limits for AN-F-58 fuel and gasoline.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:08:45.036799+00:00"}
{"citation_id": "19930083221", "source_url": "https://ntrs.nasa.gov/api/citations/19930083221/downloads/19930083221.pdf", "page_number": 4, "total_pages": 47, "image_filename": "19930083221_p4.jpg", "text": "2\nNACA TN No. 1824\n\nbecome arbitrarily large. In this case, linearized theory therefore\npredicts its inability to treat such problems. On the other hand,\nif linear methods are applied to nonstationary two-dimensional\nairfoil and particular steady-state, three-dimensional, lifting-\nsurface problems at sonic speeds, a consistent theory results since\nsolutions are found which yield perturbation velocities of the same\norder of magnitude as those calculated for free-stream Mach numbers\nof, say, 0.6 or 1.5.\n\nUnfortunately, arbitrary thickness distributions at sonic speeds\ncannot be studied by linear theory in the steady state since, in\ngeneral, the theory predicts infinite pressure differences between\nthe wing surface and infinity. In the particular case of a yawed,\nsymmetrical wing of infinite aspect ratio, the results are, however,\nagain consistent with the theory and yield pressure distributions\nwhich are the same as those determined by using only the component\nof free-stream velocity normal to the leading edge. The derivation\nof this latter result for a free-stream Mach number of one will be\ngiven.\n\nThe difficulty of not being able to include thickness effects\nin general, together with the uncertainty of the magnitude of the\nviscous effects, leaves the question as to the limitations of such\na linear theory in application to practical wing shapes. Such a\nquestion can certainly not be resolved by mathematical reasoning\nalone. The extent to which the fluid medium can be idealized at\nthese speeds is left, for the time being, unsettled and it remains\nfor experiment to determine whether the consistent mathematical\nresults which are obtained from the linearized equations provide\nreasonably exact predictions. In this connection, it should be\nmentioned that the few experimental results available for the total\nlift on thin triangular wings at Mach numbers near one tend to\nconfirm the theory. But even if more detailed experimental results\nindicate that further refinements are necessary, there is still\nlittle doubt but that the linear potential solutions will provide a\nvaluable basis for more exact extensions of theory.\n\nThe present report is divided into three parts. In the first\npart, the linearization of the partial differential equation for\nthe velocity potential is carried out in some detail for steady-\nstate conditions. A by-product of this derivation is the nonlinear\nform of the equation for two-dimensional flow which was used by\nvon Kármán (reference 1) to determine his similarity rules for tran-\nsonic flow. The equation for unsteady two-dimensional flow based on\nthe same assumptions is also given. The second part of the report", "timestamp": "2026-07-22T06:08:45.639530+00:00"}
{"citation_id": "19930086097", "source_url": "https://ntrs.nasa.gov/api/citations/19930086097/downloads/19930086097.pdf", "page_number": 36, "total_pages": 36, "image_filename": "19930086097_p36.jpg", "text": "UNCLASSIFIED\n\nThis document contains information affecting the National Defense of the United States within the meaning of the Espionage Laws, Title 18, U.S.C., Sections 793 and 794; the transmission or revelation of its contents in any manner to an unauthorized person is prohibited by law.\n\nUNCLASSIFIED", "timestamp": "2026-07-22T06:08:48.247725+00:00"}
{"citation_id": "19930082592", "source_url": "https://ntrs.nasa.gov/api/citations/19930082592/downloads/19930082592.pdf", "page_number": 16, "total_pages": 50, "image_filename": "19930082592_p16.jpg", "text": "NACA TN 1914\n15\n\n<!-- Image (43, 109, 884, 999) -->\n\nFigure 1. - Effect of time, temperature, and molybdenum content on oxidation penetration of titanium-carbide - molybdenum cermets.", "timestamp": "2026-07-22T06:08:49.561172+00:00"}
{"citation_id": "19930082613", "source_url": "https://ntrs.nasa.gov/api/citations/19930082613/downloads/19930082613.pdf", "page_number": 14, "total_pages": 46, "image_filename": "19930082613_p14.jpg", "text": "NACA TN 1938\n13\n\n2. Buckling, which is produced at or near most of the cracks by thermal stresses, is the result of the production of over-all temperature gradients in the liner and of large temperature gradients formed at individual louvers and air-intake holes by the entrance of secondary combustion air.\n\n3. Cracks that form in the buckle are believed to be caused principally by thermal fatigue of the buckle.\n\n4. Cracking may be retarded and liner life prolonged by removing stress raisers produced during punching operations by reaming, sanding, and vapor blasting the edges of the punched holes. Some cracks probably originated from small fissures produced by the punching operation.\n\n5. The surface and subsurface scales, which formed, are believed to lengthen and widen cracks, act as stress raisers thereby lowering resistance to fatigue, and, in some cases, so weaken or stress grain boundaries that cracks originate in these boundaries.\n\nLewis Flight Propulsion Laboratory,\nNational Advisory Committee for Aeronautics,\nCleveland, Ohio, October 11, 1948.\n\nREFERENCES\n\n1. Anon.: Engineering Properties of Inconel. Bull. T-7, Development and Res. Div., The International Nickel Co., Inc., March 1943.\n\n2. Evans, Ulick R.: Metallic Corrosion Passivity and Protection. Longmans, Green & Co. (New York), 2d. ed., 1946, p. 116.\n[annotation: Liquid?]\n\n3. McKay, Robert J., and Worthington, Robert: Corrosion Resistance of Metals and Alloys, chs. XIV and XVII. Reinhold Pub. Corp. (New York), 1936, pp. 285-326, 350-378.", "timestamp": "2026-07-22T06:08:55.292464+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 28, "total_pages": 78, "image_filename": "19930082483_p28.jpg", "text": "26\nNACA TN No. 1807\n\nthe operating characteristics of the turbine, which is either\nindicated directly or is obtained by readily performed calcula-\ntions. As an illustration, the corrected power output, total-\npressure ratio, corrected rotor speed, and over-all efficiency\nfor a given inlet total pressure and temperature may be observed\ndirectly for two amounts of admission; whereas such quantities\nas weight flow, theoretical power input, discharge static pressure,\nratio of inlet total pressure to discharge static pressure, theo-\nretical blade-to-jet speed ratio, and turbine efficiency based on\nratio of inlet total pressure to discharge static pressure may be\ncomputed. The method by which these quantities may be computed\nfrom a performance presentation like figure 8 follows.\n\n(1) theoretical power input based on stagnation conditions\n\n$$= \\frac{\\text{power output corrected to sea level}}{\\text{efficiency based on total-pressure ratio}}$$\n\n$$\\frac{W(\\Delta_s h')}{0.707} = \\frac{P}{\\eta^t} \\quad (42)$$\n\n(2) weight flow\n\n$$= \\frac{\\text{theoretical power input based on stagnation conditions}}{\\text{ideal enthalpy drop per pound fluid based on stagnation conditions}}$$\n\n$$W = \\frac{\\frac{W(\\Delta_s h')}{0.707}}{\\frac{(\\Delta_s h')}{0.707}} \\quad (43)$$\n\nWhen the inlet total temperature and total-pressure ratio are\nknown, the ideal power per pound of driving fluid may be obtained\nby use of tables in reference 8.\n\n(3) discharge total pressure = $\\frac{\\text{inlet total pressure}}{\\text{total-pressure ratio}}$\n\n$$P_e' = \\frac{P_i'}{\\left(\\frac{P_i'}{P_e'}\\right)} \\quad (44)$$", "timestamp": "2026-07-22T06:08:55.398224+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 43, "total_pages": 98, "image_filename": "19930086073_p43.jpg", "text": "NACA RM A9E04\n\nLift coefficient, $C_L$\n\nDrag coefficient, $C_D$\n\nAngle of sideslip, $\\beta$, deg\n\n(b) $C_L$ vs $C_D$.\n\nFigure 8.—Continued.\n\n41", "timestamp": "2026-07-22T06:09:01.533118+00:00"}
{"citation_id": "19930082703", "source_url": "https://ntrs.nasa.gov/api/citations/19930082703/downloads/19930082703.pdf", "page_number": 10, "total_pages": 28, "image_filename": "19930082703_p10.jpg", "text": "8\nNACA TN 1983\n\ncriterion. This situation is analogous to airplane requirements for\nstick-fixed and stick-free stability in which requirement of stick-free\nstability has generally been sufficient for the simple reason that, in\nachieving stick-free stability, stick-fixed stability was automatically\nobtained. When exceptions occurred and stick-free stability was\nobtained without achieving stick-fixed stability, the characteristics,\nin some cases at least, were not considered satisfactory (reference 4).\n\nThe time interval between stick deflection and attainment of\nmaximum normal acceleration could be made to provide a criterion for\navoidance of prolonged divergence, but not a criterion for completely\nsatisfactory characteristics, inasmuch as the satisfactory configuration\n(helicopter C) showed a time interval almost identical to that for\nhelicopter B, which had an objectionable delay in development of\nacceleration. It seems noteworthy that, provided that the manner of\ndevelopment is logical, a time interval of $2\\frac{1}{2}$ seconds as shown for\nhelicopter C is not objectionable to the pilot. For operation in close\nquarters, as in crop dusting, the time interval might have more signifi-\ncance, at least to the extent that the collective pitch control would\nbe used when immediate acceleration is needed. The present study did\nnot include such operations.\n\nThe long-period stick-fixed oscillation characteristics for the\nthree helicopters show improvements coincident with those of the pull-up\ncharacteristics; however, to require simply that these oscillations\ndamp out would indicate that both helicopters B and C were fully satis-\nfactory, whereas helicopter B was found to cause the pilot undue\ndifficulty in anticipating the final result of a control deflection.\nAlso, although the change from helicopter A to helicopter B resulted in\na change of the long-period motion from divergent to convergent, the\npause in the development of acceleration appears, if anything, to be\nsomewhat increased rather than diminished. Furthermore, long-period\noscillations, even if moderately divergent rather than damped, have\ngenerally been found not to influence the pilot's liking for an aircraft.\nAlthough the oscillations of the helicopter involve more attitude and\nacceleration changes than do the long-period (phugoid) oscillation of\nthe airplane, nevertheless from present knowledge (including the\ntoleration of long-period helicopter oscillations in hovering; see\nreference 1) it does not appear logical to require rapid damping of\nthese oscillations. Furthermore stick-free stability may be found to\nmask adequately a tendency toward slow divergence of these long-period\nstick-fixed oscillations (but not a tendency toward rapid divergence).\nFinally, consideration of the analysis given in the following section,\ntogether with consideration of the factors known to affect the\noscillations, indicates that no unique relation exists between the\ndetails of the early part of the pull-up and the damping of the long-\nperiod oscillations. It is concluded, therefore, that a requirement", "timestamp": "2026-07-22T06:09:03.336403+00:00"}
{"citation_id": "19930082566", "source_url": "https://ntrs.nasa.gov/api/citations/19930082566/downloads/19930082566.pdf", "page_number": 17, "total_pages": 44, "image_filename": "19930082566_p17.jpg", "text": "```markdown\nNACA TN NO. 1889\n\nFlat, 1/16\n\nView A-A\n\nRadius, 1/8\n\n$2\\frac{1}{2}$ - 16 EF - 2\n(SAE)\n\nRadius, 20\n\n$59^\\circ$\n\n$2.000 \\pm 0.001$\n$2.10 \\pm 0.001$\n\n$1\\frac{1}{8}$\n$\\frac{1}{4}$\n$1\\frac{1}{2}$\n\nGround and polished\n\n8\n16\n\n$1\\frac{1}{2}$\n$\\frac{1}{4}$\n$1\\frac{1}{8}$\n\n$\\frac{3}{8}$\n$\\frac{23}{32}$\n$\\frac{16}{32}$\n\nNACA\n\nFigure 3.- Biaxial stress specimen. All dimensions are in inches.\n\n15\n```", "timestamp": "2026-07-22T06:09:04.910424+00:00"}
{"citation_id": "19930085485", "source_url": "https://ntrs.nasa.gov/api/citations/19930085485/downloads/19930085485.pdf", "page_number": 9, "total_pages": 26, "image_filename": "19930085485_p9.jpg", "text": "NACA RM No. L8K02 CONFIDENTIAL 7\n\nCONCLUSIONS\n\nComparative tests of aileron-contour modifications at subsonic and transonic speeds on a $42.7^\\circ$ sweptback circular-arc wing indicated the following conclusions:\n\n1. The aileron showed greater effectiveness with flat sides and thickened trailing edge than it did with circular-arc contour throughout the Mach number range tested. At an angle of attack of $0^\\circ$ there was no indication of reversal of control of the aileron with the trailing edge one-half as thick or as thick as the aileron at the hinge line throughout the Mach number range tested. The aileron with flat sides and trailing edge one-half as thick as the aileron at the hinge line gave the most linear variation of rolling-moment coefficient with aileron deflection at any Mach number and generally gave the highest effectiveness of any configuration investigated.\n\n2. The drag coefficient was increased by thickening the aileron trailing edge at subsonic Mach numbers, but a decrease was indicated in some cases in the transonic range.\n\n3. Thickening the aileron trailing edge shifted the aerodynamic center rearward by approximately 2 to 10 percent at subcritical Mach numbers.\n\nLangley Aeronautical Laboratory\nNational Advisory Committee for Aeronautics\nLangley Field, Va.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:09:07.540073+00:00"}
{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 4, "total_pages": 72, "image_filename": "19930085491_p4.jpg", "text": "NACA RM No. A8J04 CONFIDENTIAL 3\n\nleading edge. An increase in maximum lift-drag ratio then results from decreases in both minimum drag coefficient and the drag due to lift. The effect of sweeping the wing leading edge well within the Mach cone in reducing the minimum drag coefficient has been shown by Jones in reference 2. The reduction in drag due to lift results from the realization of a leading-edge suction force associated with the up-flow at the wing leading edge that is not obtained with wings swept ahead of the Mach cone. The flow on the sections farthest from the wing root, exclusive of those within the tip Mach cones, most closely approach ideal, two-dimensional, subsonic flow and thereby realize the greatest reduction in minimum drag coefficient and drag due to lift. Thus the use of the highest practicable aspect ratio is indicated.\n\nA general wind-tunnel investigation is being undertaken at the Ames Aeronautical Laboratory with wing-fuselage combinations having wings with leading edges swept back $63^\\circ$ to determine experimentally the characteristics of a configuration similar to the types shown by Jones in reference 1 to be theoretically efficient at supersonic flight speeds. The facilities employed permit a study at several Reynolds numbers for both subsonic and supersonic Mach numbers. Results obtained to date at subsonic speeds with this configuration are presented in references 3 and 4. The present investigation is primarily concerned with the characteristics of the $63^\\circ$ uncambered, untwisted wing and fuselage combination at a Mach number of 1.53. The leading-edge sweep angle in the present tests was variable within the range of $57.0^\\circ$ to $69.9^\\circ$ and, as a secondary phase of the study, an experimental determination of the optimum leading-edge sweep angle for maximum lift-drag ratio at a Mach number of 1.53 was undertaken. This secondary phase of the investigation also served to indicate any possible adverse effects, particularly on longitudinal stability characteristics, of a subsonic, sonic, or supersonic trailing edge.\n\nSYMBOLS\n\nBasic Symbols\n\n| Symbol | Definition |\n| :--- | :--- |\n| A | aspect ratio $\\left( \\frac{b^2}{S} \\right)$ |\n| b | wing span measured perpendicular to plane of symmetry, inches |\n| c | wing chord measured parallel to plane of symmetry, inches |\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:09:12.547623+00:00"}
{"citation_id": "19930083221", "source_url": "https://ntrs.nasa.gov/api/citations/19930083221/downloads/19930083221.pdf", "page_number": 5, "total_pages": 47, "image_filename": "19930083221_p5.jpg", "text": "```markdown\nNACA TN No. 1824\n3\n\nis restricted to two-dimensional unsteady problems for values of\nMach number near one. The principal contribution of this section is\nthe evaluation of the change with time of the pressure distribution\nover an airfoil starting suddenly from rest at a speed close to that\nof sound. Such an idealized problem involves a step function in\nvelocity in which the airfoil has zero velocity for all negative\nvalues and near sonic velocity for all positive values of time. From\nthese results the initial build-up of lift can be calculated for Mach\nnumbers near one, although the eventual value of the lift cannot be\nfound by linear methods. Further application can also be made to\nproblems in flutter and gust loads. The third part of the report\ntreats the steady-state three-dimensional problem. Both lifting\nsurfaces and symmetrical nonlifting wings are considered and it is\nseen that in the former case consistent solutions are obtained by\nparticularly simple means. These solutions represent the limiting\ncase of both subsonic and supersonic lifting-surface theory and\ngive, for example, the same value of lift-curve slope at the speed\nof sound that was obtained for the supersonic triangular wing by\nStewart (reference 2).\n\nA list of symbols is given in the appendix.\n\nPART I - THE LINEARIZED EQUATIONS OF MOTION\n\nSteady State\n\nThe nonlinear partial differential equation satisfied by the\nvelocity potential $\\phi$ of an isentropic flow field can be expressed\nin the form\n\n$$\n\\phi_{xx} \\left( 1 - \\frac{\\phi_x^2}{a^2} \\right) + \\phi_{yy} \\left( 1 - \\frac{\\phi_y^2}{a^2} \\right) + \\phi_{zz} \\left( 1 - \\frac{\\phi_z^2}{a^2} \\right)\n$$\n$$\n- \\frac{2}{a^2} \\phi_{yz} \\phi_y \\phi_z - \\frac{2}{a^2} \\phi_{zx} \\phi_z \\phi_x - \\frac{2}{a^2} \\phi_{xy} \\phi_x \\phi_y = 0 \\quad (1)\n$$\n\nwhere the subscript notation is used to indicate differentiation and\na is the local speed of sound given by the relation\n\n$$\n\\left( \\frac{a}{V_o} \\right)^2 = \\frac{1}{M_o^2} \\left\\{ 1 - \\frac{(\\gamma-1)}{2} M_o^2 \\left[ \\left( \\frac{V}{V_o} \\right)^2 - 1 \\right] \\right\\} \\quad (2)\n$$\n\nIn this latter equation $V_o$ and $M_o$ are, respectively, velocity\nand Mach number of the free stream, $\\gamma$ is the ratio of specific\n```", "timestamp": "2026-07-22T06:09:33.485509+00:00"}
{"citation_id": "19930085485", "source_url": "https://ntrs.nasa.gov/api/citations/19930085485/downloads/19930085485.pdf", "page_number": 10, "total_pages": 26, "image_filename": "19930085485_p10.jpg", "text": "8\nCONFIDENTIAL\nNACA RM No. L8K02\n\nREFERENCES\n\n1. Turner, Thomas R., Lockwood, Vernard E., and Vogler, Raymond D.:\nPreliminary Investigation of Various Ailerons on a 42° Sweptback\nWing for Lateral Control at Transonic Speeds. NACA RM No. L8D21,\n1948.\n\n2. Swanson, Robert S., and Toll, Thomas H.: Jet-Boundary Corrections\nfor Reflection-Plane Models in Rectangular Wind Tunnels. NACA\nRep. No. 770, 1943.\n\n3. Gilruth, R. R., and Wetmore, J. W.: Preliminary Tests of Several\nAirfoil Models in the Transonic Speed Range. NACA ACR No. L5B08,\n1945.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:09:37.525623+00:00"}
{"citation_id": "19930082918", "source_url": "https://ntrs.nasa.gov/api/citations/19930082918/downloads/19930082918.pdf", "page_number": 7, "total_pages": 62, "image_filename": "19930082918_p7.jpg", "text": "6\nNACA TN 1940\n\nIn order to maintain a surface that was flat during the electrolytic metal removal, a special cell was designed as shown in figure 4. It consisted essentially of a 250-cubic-centimeter cylindrical container with a copper plate on the bottom as a cathode. This plate was connected to the source of current by a lead through a glass-metal seal. About one-half way up the cell a watch glass containing a 5/8-inch hole was mounted horizontally. The metal sample acting as the anode was mounted 1/2 inch above the hole with the polished surface facing the hole. This hole acted to distribute the current evenly over the 7/8-inch-square surface at this distance, and thus a plane surface was maintained during metal removal. Water-cooling was used to prevent pitting associated with electrolyte temperatures above 100° C.\n\nThe most satisfactory electrolyte was experimentally found to be a mixture of one-third concentrated hydrochloric acid (37 percent) and two-thirds glycerine. This mixture had the best current efficiency, approximately 0.0000625 inch of metal removed per ampere-minute at 8 amperes per square inch, without excessive pitting. A quantity of 200 cubic centimeters of this solution was sufficient for 6 to 10 samples. Pitting occurred when the metal ion concentration became too high. Phosphoric acid and glycerine combinations gave good polished surfaces, but had low current efficiencies. Sulfuric acid and glycerine mixtures caused passivation; mixtures of chromic acid or hydrofluoric acid with glycerine left the surface badly pitted.\n\n(2) After the gross metal removal of step (1), the surface was given a high polish using undiluted Du Pont electropolishing solution for 5 minutes at 5 amperes per square inch. Metal removal in this step was negligible. The simple cell consisting of a beaker with a copper plate in the bottom was used for this step. No water-cooling was necessary. Of several other electrolytes tried for this step, only a mixture of 40 percent phosphoric acid and 60 percent glycerine was nearly as satisfactory as the Du Pont solution.\n\n(3) After surface preparation, diffraction patterns were taken with the specimens either rotating or oscillating in the X-ray beam. The grain size of the low-carbon N-155 specimens used was too large to present an effectively random distribution in the X-ray beam when the specimen was stationary. By moving the specimen under the beam during the time the diffraction pattern was being taken, a number of grains were presented to the beam in an effort to make the specimen approximately one of random grain orientation. In the case of line intensity measurements on a Norelco spectrometer, a specimen holder was designed and built which rotated the specimens at approximately 17 rps. Care was taken to insure that the plane of polish was perpendicular to the axis of spin. In the case of line width measurements, the specimens were oscillated $\\pm 10^\\circ$ about an axis which was normal to the incoming X-ray beam, but parallel to the", "timestamp": "2026-07-22T06:09:38.819889+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 29, "total_pages": 78, "image_filename": "19930082483_p29.jpg", "text": "(4) discharge static pressure $P_e$, from the following equation\n\n$$\n\\frac{P_e'}{P_e} = \\left[ \\frac{1}{2} + \\sqrt{ \\frac{1}{4} + \\frac{1}{2g} \\frac{\\gamma - 1}{\\gamma} \\left( RT_1' - \\frac{\\gamma - 1}{\\gamma} 550 \\frac{P}{W} \\right) \\left( \\frac{W}{70.73 P_e A_e} \\right)^2 } \\right]^{\\frac{\\gamma}{\\gamma - 1}}\n$$\n\nwhere\n\n$A_e$ discharge area, (sq ft)\n\nIt is to be noted that the power term in equation (45) should utilize blade power; however, shaft power, an easily measurable quantity, has been substituted with resulting error so slight as to be negligible.\n\n(5) ratio of inlet total pressure to discharge static pressure\n\n$$\n= \\frac{\\text{inlet total pressure}}{\\text{discharge static pressure}}\n$$\n\n$$\n= \\frac{P_1'}{P_e}\n$$\n\n(6) efficiency based on ratio of inlet total pressure to discharge static pressure\n\n$$\n= \\frac{\\text{net power output}}{\\text{ideal power input based on ratio of inlet total pressure to discharge static pressure}}\n$$\n\n$$\n\\eta = \\frac{P}{\\left[ \\frac{W (\\Delta g h)}{0.707} \\right]}\n$$\n\nNACA TN No. 1907", "timestamp": "2026-07-22T06:09:49.404743+00:00"}
{"citation_id": "19930085847", "source_url": "https://ntrs.nasa.gov/api/citations/19930085847/downloads/19930085847.pdf", "page_number": 1, "total_pages": 32, "image_filename": "19930085847_p1.jpg", "text": "```markdown\nNACA RM A9D04\n\nCopy 204\nRM A9D04\n\nCONFIDENTIAL\nCLASSIFICATION CANCELLED\n\nNACA\nCASE FILE\nCOPY\n\nRESEARCH MEMORANDUM\n\nFLIGHT INVESTIGATION OF THE EFFECT OF BOUNDARY-\nLAYER SUCTION ON PROFILE-DRAG COEFFICIENT\nAT SUPERCRITICAL MACH NUMBERS\n\nBy Richard B. Skoog\n\nAmes Aeronautical Laboratory\nMoffett Field, Calif.\n\nCLASSIFICATION CANCELLED\nAUTHORITY CHOWLEY CHANGE#1548\nDATE 6-29-53 J.H. LOVINEK\n\nCLASSIFIED DOCUMENT\nThis document contains classified information\naffecting the National Defense of the United\nStates within the meaning of the Espionage Act,\nUSC 50:31 and 32. Its transmission or the\nrevelation of its contents in any manner to an\nunauthorized person is prohibited by law.\nDistribution to civilians may be restricted\nonly to persons in the military and naval\nservices of the United States, appropriate\ncivilian officers and employees of the Federal\nGovernment who have a legitimate interest\ntherein, and to United States citizens of known\nloyalty and discretion who of necessity must be\ninformed thereof.\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\nWASHINGTON\nSeptember 20, 1949\n\nCONFIDENTIAL\nCLASSIFICATION CANCELLED\n```", "timestamp": "2026-07-22T06:09:49.786516+00:00"}
{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 5, "total_pages": 72, "image_filename": "19930085491_p5.jpg", "text": "4\nCONFIDENTIAL\nNACA RM No. A8J04\n\n$\\bar{c}$\nmean aerodynamic chord $\\left( \\frac{\\int_{0}^{b/2} c^2 dY}{\\int_{0}^{b/2} c \\ dY} \\right)$, inches\n\n$\\bar{c}_g$\nmean geometric chord $\\left( \\frac{S}{b} \\right)$, inches\n\n$c_r$\nwing root chord, inches\n\n$c_t$\nwing tip chord, inches\n\n$C_D$\ntotal drag coefficient $\\left( \\frac{\\text{drag}}{q_o S} \\right)$\n\n$C_{D_{\\text{min}}}$\nminimum total drag coefficient\n\n$\\Delta C_D$\nrise in drag coefficient above minimum $(C_D - C_{D_{\\text{min}}})$\n\n$C_L$\nlift coefficient $\\left( \\frac{\\text{lift}}{q_o S} \\right)$\n\n$C_{L_{\\text{opt}}}$\nlift coefficient for maximum lift-drag ratio\n\n$\\frac{dC_L}{d\\alpha}$\nlift-curve slope, per radian unless otherwise specified\n\n$\\Delta C_L$\nchange in lift coefficient from value for minimum drag\n$(C_L - C_{L_{D=\\text{min}}})$\n\n$\\frac{\\Delta C_D}{(\\Delta C_L)^2}$\ndrag-rise factor\n\n$\\left( \\frac{L}{D} \\right)_{\\text{max}}$\nmaximum lift-drag ratio\n\n$C_{m_{\\frac{\\bar{c}}{2}}}$\npitching-moment coefficient about 50 percent mean aerodynamic\nchord $\\left( \\frac{\\text{pitching moment about 50 percent mean aerodynamic chord}}{q_o S \\bar{c}} \\right)$\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:09:50.668583+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 40, "total_pages": 114, "image_filename": "19930086061_p40.jpg", "text": "```markdown\n36\n\n| $\\alpha=34.1^\\circ$ | $\\alpha=39.1^\\circ$ | $\\alpha=44.1^\\circ$ | $\\alpha=48.1^\\circ$ |\n| :--- | :--- | :--- | :--- |\n| $C_L=0.98$ | $C_L=0.91$ | $C_L=0.67$ | $C_L=0.59$ |\n| [Graph: y-axis P from -4 to 1, x-axis x/c from 0 to 10. Curve shows pressure distribution.] | [Graph: Station 1, $y/b/2, 0$. Curve shows pressure distribution.] | [Graph: Curve shows pressure distribution.] | [Graph: Curve shows pressure distribution.] |\n| [Graph: y-axis P from -3 to 1. Legend: Upper (solid line), Lower (dashed line). Curve shows pressure distribution.] | [Graph: Station 2, $y/b/2, 0.167$. Curve shows pressure distribution.] | [Graph: Curve shows pressure distribution.] | [Graph: Curve shows pressure distribution.] |\n| [Graph: y-axis P from -3 to 1. Curve shows pressure distribution.] | [Graph: Station 3, $y/b/2, 0.333$. Curve shows pressure distribution.] | [Graph: Curve shows pressure distribution.] | [Graph: Curve shows pressure distribution.] |\n\n(a) Stations 1, 2, 3.\n\nFigure 7.- Chordwise pressure distribution about wing 1 at angles of attack of $34.1^\\circ$, $39.1^\\circ$, $44.1^\\circ$, and $48.1^\\circ$.\n\nNACA RM L9J07\n```", "timestamp": "2026-07-22T06:09:51.095412+00:00"}
{"citation_id": "19930082646", "source_url": "https://ntrs.nasa.gov/api/citations/19930082646/downloads/19930082646.pdf", "page_number": 7, "total_pages": 37, "image_filename": "19930082646_p7.jpg", "text": "6\nNACA TN 1980\n\nthe smooth-water landing characteristics of the modified hull were considered satisfactory.\n\nSpray Characteristics\n\nA plot of gross load against the speed range over which spray entered the propellers and struck the flaps is presented in figure 11. The modified hull encountered no propeller spray at the design gross load of 75,000 pounds. (See fig. 12.) Observations indicate that propeller spray comparable to that of the basic hull at design gross load was approximated at a gross load of 85,000 pounds (fig. 13) which represents an overload of approximately 11 percent. No flap spray was encountered at design gross load and no heavy flap spray was encountered by the modified hull at any load investigated. This improvement in spray characteristics was attributed to the effectiveness of the warped forebody in reducing the height of the bow blister even though the trims were lower. Photographs showing the maximum flap spray at design gross load for both models are presented in figure 14.\n\nPhotographs of heaviest spray striking the horizontal tail surfaces during landings are shown in figure 15. Spray on the horizontal tail was considerably lessened with the extended afterbody.\n\nThe spray diagram obtained during taxying tests in waves 2 feet high and 110 feet long is presented in figure 16. This spray was fairly heavy but occurred over a smaller range of speed and load than for the basic hull. The net effect of combining the warped forebody with the extended afterbody was a definite improvement in all spray characteristics.\n\nLandings in Waves\n\nThe rough-water landings were made in oncoming waves 4 feet high varying in length from 130 to 360 feet. Pertinent data obtained from records of these landings are presented in table I.\n\nThe maximum vertical and angular accelerations are plotted against wave length in figure 17. The maximum vertical acceleration of 4g encountered by the modified hull was approximately 55 percent lower than the maximum encountered by the basic hull and approximately 30 percent lower than the maximum obtained during tests of configurations incorporating either the warped forebody or extended afterbody modifications alone (references 1 and 2). The maximum positive angular accelerations for the modified hull were 59 percent less than those of the basic hull, 19 percent less than those of the warped forebody alone,", "timestamp": "2026-07-22T06:09:52.865016+00:00"}
{"citation_id": "19930086076", "source_url": "https://ntrs.nasa.gov/api/citations/19930086076/downloads/19930086076.pdf", "page_number": 49, "total_pages": 50, "image_filename": "19930086076_p49.jpg", "text": "NACA RM E9F09\n47\n\n<!-- Image (113, 206, 887, 366) -->\n\nTime, sec\n(a) No combustion\n\n<!-- Image (113, 495, 887, 665) -->\n\nTime, sec\n(b) Combustion\n\nNACA\nC-23570\n6-10-49\n\nFigure 19. - Pressure variation with time at 4- by 8-inch combustor inlet with flame holder 13. Inlet-air velocity, 200 feet per second; inlet-air temperature, 200° F.\n\nNACA - Langley Field, Va.", "timestamp": "2026-07-22T06:09:54.395161+00:00"}
{"citation_id": "19930082566", "source_url": "https://ntrs.nasa.gov/api/citations/19930082566/downloads/19930082566.pdf", "page_number": 18, "total_pages": 44, "image_filename": "19930082566_p18.jpg", "text": "16\n\nInjection\npump I\n\nGear drive\n\nDynamometer N\n\nPressure gages\n\nSpecimen S\n\nLoading lever K\n\nNACA\n\nFigure 4.- Axial loading arrangement.\n\nNACA TN No. 1889", "timestamp": "2026-07-22T06:09:54.961854+00:00"}
{"citation_id": "19930093769", "source_url": "https://ntrs.nasa.gov/api/citations/19930093769/downloads/19930093769.pdf", "page_number": 33, "total_pages": 39, "image_filename": "19930093769_p33.jpg", "text": "32\nCONFIDENTIAL\nNACA RM No. E8L10a\n\n[Figure: A graph plotting Altitude (ft) against Flight Mach number. The y-axis ranges from 0 to 50,000 ft. The x-axis ranges from 0 to 1.0. The graph contains a legend, a shaded region, and several data points.]\n\nAltitude, ft\n50,000\n40,000\n30,000\n20,000\n10,000\n0\n\n0 .2 .4 .6 .8 1.0\nFlight Mach number\n\no Satisfactory start\n□ Burner ignited but\nacceleration impossible\n◇ No ignition\n\nRegion of unsatisfactory\nstarts\n\nNACA\n\n(a) Fuel, AN-F-58; standard spark plug; plug cleaned\nfor every run.\nFigure 9. - Windmilling-starting characteristics.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:09:58.460817+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 81, "total_pages": 96, "image_filename": "19930085880_p81.jpg", "text": "NACA RM No. L9C03\n79\n\n[Figure: A graph plotting Load (lb) against Wetted area (sq ft). The graph contains five curves representing different speeds (10, 15, 20, 25, 30 fps). The curves are marked with different symbols: circles for 10 fps, squares for 15 fps, diamonds for 20 fps, triangles for 25 fps, and inverted triangles for 30 fps. There is also a small inset diagram showing a triangle with a line underneath it.]\n\nLoad, lb\n32\n28\n24\n20\n16\n12\n8\n4\n0\n\nSpeed\n(fps)\n30\n25\n20\n15\n10\n\n0 .05 .10 .15 .20 .25 .30 .35\nWetted area, sq ft\n\n(e) $\\tau = 20^\\circ$.\n\nFigure 22.- Concluded.\n\nNACA", "timestamp": "2026-07-22T06:10:01.171783+00:00"}
{"citation_id": "19930085572", "source_url": "https://ntrs.nasa.gov/api/citations/19930085572/downloads/19930085572.pdf", "page_number": 4, "total_pages": 17, "image_filename": "19930085572_p4.jpg", "text": "2\nNACA RM No. E8L02\n\nAPPARATUS\n\nTwo turbojet engines were used for the investigation, one a J35-C5 with a J35 fuel-control system and the other a J35-C-3A with a J33 fuel-control system. The engine change was made in compliance with an U.S. Air Force technical order, which was issued during the investigation. Both engines, however, have the same manufacturer's rating of 4000 pounds of static thrust at an engine speed of 7700 rpm. The principal components of the engines are the same and include an 11-stage axial-flow compressor, a single-stage turbine, and eight individual combustion chambers.\n\nFor the investigation, the engine was mounted on a carriage, which was lowered in flight from the forward bomb bay of a medium-bomber-type airplane (fig. 1). Pressure and temperature instrumentation was provided at the compressor inlet and the tail-pipe outlet for calculations of net thrust. A positive-displacement-type volumetric flowmeter was provided for measuring fuel flow to the engine.\n\nThe specifications and analysis of the properties for the two fuels used in this investigation are given in table I. Both fuels are within the specification limits.\n\nPROCEDURE\n\nNormal performance data using each fuel were obtained on the J35-C5 engine at pressure altitudes of 5,000, 10,000, 20,000, and 30,000 feet. At each altitude, the engine was operated over a speed range from 4000 to approximately 7000 rpm at a Mach number of 0.37 (ram pressure ratio, 1.1). This program was individually conducted on each fuel because the airplane could not simultaneously carry an ample supply of both fuels.\n\nEngine starting and blow-out data, also using each fuel separately, were obtained on the J35-C3A engine at pressure altitudes of 5,000, 10,000, 20,000, and 30,000 feet at a Mach number of approximately 0.37. With the engine windmilling, engine starts were effected by opening the throttle until the small-slot fuel pressure reached a value of 30 to 40 pounds per square inch and then turning on the ignition. After the starts were made, the engine was accelerated as rapidly as possible to a speed of approximately 7000 rpm without exceeding the tail-pipe-temperature limit. The engine blow-out speed was obtained at a given altitude by slowly reducing the engine speed with the throttle until a sudden drop in tail-pipe temperature occurred.", "timestamp": "2026-07-22T06:10:01.356709+00:00"}
{"citation_id": "19930082703", "source_url": "https://ntrs.nasa.gov/api/citations/19930082703/downloads/19930082703.pdf", "page_number": 11, "total_pages": 28, "image_filename": "19930082703_p11.jpg", "text": "NACA TN 1983\n\nbased on long-period oscillations could not be used as a complete substitute for the pull-and-hold requirements.\n\nTheoretical Analysis of Pull-Up Characteristics\n\nA theoretical analysis of helicopters A and B in pull-ups has been made in order to determine whether the pull-up characteristics previously discussed can be theoretically predicted and in what way the tail surface causes the measured change in these characteristics. (No theoretical analysis of helicopter C could be made because some of the necessary parameters were not available.)\n\nBecause of the complexity of the phenomenon, several simplifying assumptions were made. The most important of these assumptions are:\n\n(1) Constant rotor speed and collective pitch (the collective pitch on these helicopters varies with lag angle and coning angle)\n\n(2) Small displacements\n\n(3) Representation of the dynamic motion by variations in steady state conditions, for example, the lag in the changes of induced velocity is neglected\n\nFor the present purpose at least, these simplifying assumptions should not qualitatively alter the theoretical results and conclusions.\n\nIn the analysis, flight-path axes were used and four variables were considered: forward speed, pitching velocity, angle of climb (the time derivative of which is proportional to normal-acceleration increment), and rotor angle of attack (which differs from the fuselage angle of attack by an amount equal to the longitudinal cyclic control). An instantaneous rearward motion of the longitudinal control resulting in a change of $1^\\circ$ in cyclic pitch was assumed, the resulting control position being maintained indefinitely. Four differential equations were set up: three of them expressing the equilibrium of pitching moments and of forces along and perpendicular to the flight path and the fourth expressing the rotor angle of attack as a function of pitching velocity, angle of climb, and control displacement. Most of the rotor terms in these equations were based on the theory of references 5 and 6. The rotor terms which depended on pitching velocity were the only ones not so derived. The values of these latter terms were based on the commonly used parameter $\\frac{16}{\\gamma \\Omega}$, where $\\gamma$ is the blade mass factor and $\\Omega$ is the rotor speed. This parameter, which is discussed in reference 7, expresses the longitudinal tilt of the thrust vector per unit pitching velocity of the rotor shaft and, while not necessarily precise, is", "timestamp": "2026-07-22T06:10:02.812102+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 4, "total_pages": 149, "image_filename": "19930083192_p4.jpg", "text": "```markdown\nPage\n\nOPERATING STATISTICS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 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. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .", "timestamp": "2026-07-22T06:10:03.713099+00:00"}
{"citation_id": "19930085485", "source_url": "https://ntrs.nasa.gov/api/citations/19930085485/downloads/19930085485.pdf", "page_number": 11, "total_pages": 26, "image_filename": "19930085485_p11.jpg", "text": "CONFIDENTIAL\nTypical section\nA\nCopper insert\n1/32-gap, filled with wax\nA\nNACA RM No. L8E02\nBalance\ncenter line\n2.0\n4.0\nA\n42.7°\nMGC\n3.11\n3.0\n6.0\nA\n2.67\n.75\n1.5\n4.6\n4.0\n13.4\nNACA\nFigure 1.- Drawing of the 42.7° sweptback wing and fuselage combination.\nAll dimensions in inches.\nCONFIDENTIAL\n9", "timestamp": "2026-07-22T06:10:05.079728+00:00"}
{"citation_id": "19930082592", "source_url": "https://ntrs.nasa.gov/api/citations/19930082592/downloads/19930082592.pdf", "page_number": 17, "total_pages": 50, "image_filename": "19930082592_p17.jpg", "text": "16\nNACA TN 1914\n\n<!-- Image (116, 153, 829, 871) -->\n\n(b) Magnification of area outlined in figure 1(a).\nEffect of time, temperature, and molybdenum content on oxidation penetration of titanium carbide - molybdenum cermets.\n\nFigure 1. - Concluded.", "timestamp": "2026-07-22T06:10:05.266086+00:00"}
{"citation_id": "19930083221", "source_url": "https://ntrs.nasa.gov/api/citations/19930083221/downloads/19930083221.pdf", "page_number": 6, "total_pages": 47, "image_filename": "19930083221_p6.jpg", "text": "4\nNACA TN No. 1824\n\nheats (for air, $\\gamma=1.4$), and $V$ is local velocity.\n\nIntroducing the perturbation velocity potential $\\Phi$, where\n$$ \\Phi = -V_0 x + \\phi \\quad (3) $$\nit is possible to express equation (1) in terms of the derivatives of $\\phi$ and the parameters $M_0$ and $V_0$. To begin the linearization of the resulting equation, the coefficients of the second ordered derivatives of $\\Phi$ are expanded in Maclaurin series with ascending powers of $\\frac{u}{V_0}, \\frac{v}{V_0}, \\frac{w}{V_0}$. The convergence is assured provided\n$$ \\left| \\frac{(\\gamma-1)}{2} M_0^2 \\left( \\frac{2u}{V_0} + \\frac{u^2+v^2+w^2}{V_0^2} \\right) \\right| < 1 \\quad (4) $$\nor, in a slightly modified form, provided\n$$ |V^2 - V_0^2| < \\frac{2a_0^2}{\\gamma-1} = 5a_0^2 \\quad (5) $$\nIf the assumption is now made that $\\frac{u}{V_0}, \\frac{v}{V_0}, \\frac{w}{V_0} \\ll 1$ so that second and higher powers in the perturbation velocities can be neglected in comparison with one, the partial differential equation can be simplified to the form\n$$ \\Phi_{xx} \\left\\{ 1 - M_0^2 \\left[ 1 + \\frac{2u}{V_0} + (\\gamma-1) M_0^2 \\frac{u}{V_0} \\right] \\right\\} + \\Phi_{yy} + \\Phi_{zz} $$\n$$ - 2 \\Phi_{xz} \\frac{w}{V_0} M_0^2 - 2 \\Phi_{xy} \\frac{v}{V_0} M_0^2 = 0 \\quad (6) $$\nFrom this equation all the succeeding expressions will be derived.\n\nTwo- and three-dimensional linear equations, $M_0 \\neq 1$. - Since equation (6) is obviously nonlinear, additional assumptions must be made to reduce it to a linear form. Clearly, these assumptions must involve the relative magnitudes of all the terms in order to determine which ones may be neglected. Perhaps one of the least restrictive set of conditions is that:", "timestamp": "2026-07-22T06:10:05.352076+00:00"}
{"citation_id": "19930082546", "source_url": "https://ntrs.nasa.gov/api/citations/19930082546/downloads/19930082546.pdf", "page_number": 21, "total_pages": 65, "image_filename": "19930082546_p21.jpg", "text": "propeller. Even though the pressures associated with the four-blade propeller at high tip Mach numbers will be nearly equal in amplitude to those for a two-blade propeller, the corresponding wall vibration amplitudes may be much smaller. This reduction is attributable to the greater wall inertia at the higher frequencies produced by the four-blade propeller.\n\nComparison of experimental data with theory.- A body such as a fuselage has an infinite number of vibration modes. The determination of the response to a forced vibration load such as a sound wave would require the vector summation of all the responses to the particular sound wave. Such a procedure is difficult, if not impossible. It has been found experimentally that at a particular exciting frequency the response of a body is predominately determined by the vibration mode which is near the exciting frequencies. If the excitation is far from a resonant condition the amplitude of vibration may be estimated by considering only the inertia or mass of the panel. (See p. 219, reference 4.) As a first approximation, the natural frequency of the panel may be assumed to be zero and the material damping and radiation resistance may be neglected. Under such assumptions, the response of a panel to an oscillating force may be simply calculated as (p. 62, reference 4)\n\n$$\n\\xi_{02} = \\frac{P_s}{M \\omega_1^2}\n$$\n\nwhere $\\xi_{02}$ is the displacement each side of the neutral position, $P_s$ is the pressure measured at the panel surface, $M$ is the mass of panel per unit area, and $\\omega_1$ is the angular frequency of sound in radians per second. Calculations of the vibration amplitudes of the test panels for the fundamental propeller frequencies have been made by equation (3) and are plotted in figure 22. The maximum pressures measured for the first harmonic near the plane of rotation and corrected for wall reflection were used in these calculations. Wall pressures used were 2 times free-space values for the flat surface and 1.5 times free-space values for the curved surface, as indicated by results given in figure 10. Total amplitude is $2\\xi_{02}$. The calculated values are seen to be in good agreement with the vibration amplitudes measured for the reinforced panel except where resonant peaks occur (fig. 22). Since the calculations were made for an assumed natural frequency of zero, the calculated curve does not indicate the response at resonance. A simple calculation such as this may be useful for predicting vibration amplitudes for heavy walls far from resonance.\n\nFor conventional fuselage walls, which weigh much less than those tested, the acoustical radiation resistance and damping cannot be neglected. A more refined method for calculating the response of an idealized panel and which gives the effect of rigidity, panel damping, and acoustical", "timestamp": "2026-07-22T06:10:08.818198+00:00"}
{"citation_id": "19930085847", "source_url": "https://ntrs.nasa.gov/api/citations/19930085847/downloads/19930085847.pdf", "page_number": 2, "total_pages": 32, "image_filename": "19930085847_p2.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:10:10.721266+00:00"}
{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 6, "total_pages": 72, "image_filename": "19930085491_p6.jpg", "text": "NACA RM No. A8J04 CONFIDENTIAL 5\n\n$C_{m_{\\frac{1}{4}}}$ pitching-moment coefficient about 25 percent mean aerodynamic chord $\\left(\\frac{\\text{pitching moment about 25 percent mean aerodynamic chord}}{q_o S \\bar{c}}\\right)$\n\n$\\frac{dC_m}{dC_L}$ moment-curve slope\n\nh location of maximum airfoil thickness, measured from leading edge in streamwise direction, inches\n\nh/c chordwise location of maximum thickness t/c\n\n$k_a$ angle ratio, $\\frac{\\alpha_{AL}}{\\Delta\\alpha}$\n\nm ratio of the cotangent of sweep angle of the leading edge to the cotangent of the sweep angle of the Mach line\n\n$M_n$ Mach number corresponding to velocity component perpendicular to wing leading edge\n\n$M_o$ free-stream Mach number\n\nP pressure coefficient $\\left(\\frac{p-p_o}{q_o}\\right)$\n\np local static pressure, pounds per square inch\n\n$p_o$ free-stream static pressure, pounds per square inch\n\n$q_o$ free-stream dynamic pressure, pounds per square inch\n\nR Reynolds number based on mean geometric chord of wing\n\nS wing plan-form area including that blanketed by the fuselage, square inches\n\n$S_T$ wing area of triangular wing having the same leading-edge length and sweep angle as the given swept wing, square inches\n\nt/c maximum thickness of streamwise wing section\n\n$V_o$ free-stream velocity, feet per second\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:10:16.558257+00:00"}
{"citation_id": "19930082918", "source_url": "https://ntrs.nasa.gov/api/citations/19930082918/downloads/19930082918.pdf", "page_number": 8, "total_pages": 62, "image_filename": "19930082918_p8.jpg", "text": "```markdown\nNACA TN 1940\n7\n\nfilm. Finally, in making lattice-parameter measurements, the specimens were rotated about 2 rps, with the axis of rotation parallel with, but offset from, the axis of the X-ray beam.\n\nDiffraction-line peak-intensity studies.- The studies of diffraction-line peak intensity were confined to the (111) line of the austenite matrix, this line being the strongest line and within the range of the Norelco spectrometer. Copper $K\\alpha_1\\alpha_2$ radiation was used. Plots of the (111) line obtained with this spectrometer and an automatic recorder were measured graphically for peak height. Because the basic measurement needed was the line height of a given aged sample relative to unaged material, an unaged sample was run alternately with each aged sample.\n\nChecks of the reproducibility of the (111)-line measurements were made by the following procedures:\n\n(1) The unaged solution-treated sample, used as a standard of comparison, was taken twice through the surface preparation step with, however, removal of only an additional 0.0025 inch of metal the second time. After taking the (111)-line height from the first surface, the X-ray tube and counter circuits were left on during the repolishing. The (111)-line measurements were immediately taken from the second surface and found to check within the accuracy of the spectrometer. This surface was carefully preserved and used as a standard for subsequent measurements on aged samples.\n\n(2) The samples aged at $1200^\\circ$ and $1600^\\circ$ F were alternately measured for (111)-line peak intensity against the standard and repolished for a minimum of two and a maximum of six times. Between successive measurements 0.0025 inch of metal was removed. This was sufficient depth to bring up a new set of grains, since the average grain size of the samples used was approximately 0.001 inch.\n\n(3) Despite the precautions, some scatter was still found in the measurements. In an effort to reduce this scatter all the duplicate measurements on the samples aged at $1400^\\circ$ F were carried out on the same surface of each sample. These measurements were made on this surface with a 1-millimeter lateral shift in the holder between each run so that a new spot on the surface was covered by the X-ray beam each time. Two and usually three such measurements were made on each sample.\n\nDiffraction-line width studies.- Line intensity studies on the (111) line of the low-carbon N-155 matrix with the Norelco spectrometer revealed no evidences of line broadening at any stage of the aging process. As a further search for broadening effects, a photographic back-reflection technique was chosen in order to take advantage of the increased resolving power in the back-reflection region. In\n```", "timestamp": "2026-07-22T06:10:18.838742+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 82, "total_pages": 96, "image_filename": "19930085880_p82.jpg", "text": "80\nNACA RM No. L9C03\n\nResistance, lb\n\nSpeed\n(fps)\n30\n25\n20\n15\n10\n\nWetted area, sq ft\n(a) $\\tau = 40$.\n\nFigure 23.- Variation of resistance with wetted area. Model 250D.", "timestamp": "2026-07-22T06:10:20.684990+00:00"}
{"citation_id": "19930082613", "source_url": "https://ntrs.nasa.gov/api/citations/19930082613/downloads/19930082613.pdf", "page_number": 15, "total_pages": 46, "image_filename": "19930082613_p15.jpg", "text": "```markdown\n14\nNACA TN 1938\n\nTABLE I - EXAMINATION AND COMPARISON OF CRACKED AND PUNCHED EDGES\n\n<!-- Table (114, 188, 691, 785) -->\n\\begin{tabular}{|c|c|c|c|c|c|c|c|c|c|c|c|}\n\\hline\n\\multicolumn{1}{|c|}{\\multirow{3}{*}{\\textbf{Specimen}}} & \\multicolumn{1}{c|}{\\multirow{3}{*}{\\textbf{Type of}}} & \\multicolumn{1}{c|}{\\multirow{3}{*}{\\textbf{Location}}} & \\multicolumn{1}{c|}{\\multirow{3}{*}{\\textbf{Description}}} & \\multicolumn{1}{c|}{\\multirow{3}{*}{\\textbf{Depth of}}} & \\multicolumn{1}{c|}{\\multirow{3}{*}{\\textbf{Depth of}}} & \\multicolumn{1}{c|}{\\multirow{3}{*}{\\textbf{Distance}}} & \\multicolumn{1}{c|}{\\multirow{3}{*}{\\textbf{Length}}} & \\multicolumn{1}{c|}{\\multirow{3}{*}{\\textbf{Edges of Crack}}} & \\multicolumn{1}{c|}{\\multirow{3}{*}{\\textbf{Location}}} & \\multicolumn{1}{c|}{\\multirow{3}{*}{\\textbf{Location}}} & \\multicolumn{1}{c|}{\\multirow{3}{*}{\\textbf{Location}}} \\\\\n\\multicolumn{1}{|c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} \\\\\n\\multicolumn{1}{|c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} \\\\\n\\cline{1-12}\n\\multicolumn{1}{|c|}{\\textbf{number}} & \\multicolumn{1}{c|}{\\textbf{section}} & \\multicolumn{1}{c|}{\\textbf{of crack}} & \\multicolumn{1}{c|}{\\textbf{of crack}} & \\multicolumn{1}{c|}{\\textbf{crack}} & \\multicolumn{1}{c|}{\\textbf{crack}} & \\multicolumn{1}{c|}{\\textbf{of crack}} & \\multicolumn{1}{c|}{\\textbf{of crack}} & \\multicolumn{1}{c|}{\\textbf{visible}} & \\multicolumn{1}{c|}{\\textbf{of}} & \\multicolumn{1}{c|}{\\textbf{of}} & \\multicolumn{1}{c|}{\\textbf{of}} \\\\\n\\multicolumn{1}{|c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{\\textbf{edge}} & \\multicolumn{1}{c|}{\\textbf{edge}} & \\multicolumn{1}{c|}{\\textbf{(in.)}} & \\multicolumn{1}{c|}{\\textbf{(in.)}} & \\multicolumn{1}{c|}{\\textbf{from}} & \\multicolumn{1}{c|}{\\textbf{visible}} & \\multicolumn{1}{c|}{\\textbf{under}} & \\multicolumn{1}{c|}{\\textbf{surface}} & \\multicolumn{1}{c|}{\\textbf{surface}} & \\multicolumn{1}{c|}{\\textbf{intergranular}} \\\\\n\\multicolumn{1}{|c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{\\textbf{edge}} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{\\textbf{microscope}} & \\multicolumn{1}{c|}{\\textbf{crack}} & \\multicolumn{1}{c|}{\\textbf{crack}} & \\multicolumn{1}{c|}{\\textbf{or}} \\\\\n\\multicolumn{1}{|c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{\\textbf{(in.)}} & \\multicolumn{1}{c|}{\\textbf{(in.)}} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{\\textbf{transgranular}} \\\\\n\\multicolumn{1}{|c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{\\textbf{crack}} \\\\\n\\hline\n\\multicolumn{12}{|c|}{Specimens obtained from as-fabricated type-A liner} \\\\\n\\hline\n1 & Long & Lower & Rough & 0.0019 & 0.0017 & ----- & ----- & ----- & ----- & ----- & ----- \\\\\n1 & Long & Lower & Rough & ----- & 0.0017 & ----- & ----- & ----- & ----- & ----- & ----- \\\\\n1 & Long & Lower & Rough & ----- & 0.0017 & ----- & ----- & ----- & ----- & ----- & ----- \\\\\n1 & Long & Lower & Smooth & 0.0009 & 0.0011 & ----- & ----- & ----- & ----- & ----- & ----- \\\\\n\\hline\n\\multicolumn{12}{|c|}{Specimens obtained from type-A liner, which failed by cracking during service} \\\\\n\\hline\n2 & Long & Lower & Rough & 0.0033 & ----- & 0.0015 & 0.0008 & .157 & Both & ----- & Both \\\\\n2 & Long & Lower & Rough & ----- & ----- & 0.0009 & 0.0008 & .157 & Both & ----- & Both \\\\\n2 & Long & Lower & Rough & ----- & ----- & ----- & ----- & .25 & Both & ----- & ----- \\\\\n\\hline\n\\multicolumn{12}{|c|}{Specimens obtained from type-B liner, which failed by cracking during accelerated runs of 66 hours and 57 minutes} \\\\\n\\hline\n3 & Long & Near lower & Rough & 0.0008 & ----- & 0.0016 & 0.0028 & 0.0013 & Both & ----- & Both \\\\\n3 & Long & Air intake & Rough & ----- & ----- & 0.0009 & 0.0009 & .23 & Both & ----- & Both \\\\\n3 & Long & Air intake & Rough & ----- & ----- & 0.0009 & 0.0009 & .23 & Both & ----- & Both \\\\\n3 & Long & Air intake & Rough & ----- & ----- & ----- & ----- & ----- & ----- & ----- & ----- \\\\\n\\hline\n4 & Long & Near lower & Smooth & ----- & 0.00032 & ----- & 0.0008 & ----- & ----- & ----- & Chiefly transcrystalline \\\\\n4 & Long & Near lower & Smooth & ----- & 0.00032 & ----- & 0.0003 & 0.0003 & ----- & ----- & Chiefly transcrystalline \\\\\n4 & Long & Near lower & Smooth & ----- & 0.00032 & ----- & 0.0003 & 0.0003 & ----- & ----- & Chiefly transcrystalline \\\\\n4 & Long & Counter & Smooth & ----- & 0.00022 & ----- & 0.0003 & 0.0002 & ----- & ----- & Chiefly transcrystalline \\\\\n4 & Long & Counter & Smooth & ----- & 0.00022 & ----- & 0.0003 & 0.0002 & ----- & ----- & Chiefly transcrystalline \\\\\n4 & Long & Counter & Smooth & ----- & 0.00022 & ----- & 0.0003 & 0.0002 & ----- & ----- & Chiefly transcrystalline \\\\\n4 & Long & Counter & Smooth & ----- & 0.00011 & ----- & 0.0003 & 0.00012 & .25 & Both & Chiefly transcrystalline \\\\\n\\hline\n5 & Long & Counter & Smooth & ----- & ----- & ----- & ----- & ----- & ----- & ----- & ----- \\\\\n5 & Long & Air intake & Rough & 0.0117 & ----- & 0.00015 & 0.00025 & .50 & Both & ----- & Both \\\\\n5 & Long & Counter & Rough & 0.0110 & ----- & 0.00008 & 0.00025 & .50 & Both & ----- & Both \\\\\n\\hline\n6 & Long & Counter & Smooth & 0.00032 & ----- & 0.00012 & 0.00036 & 0.18 & Both & ----- & Chiefly transcrystalline \\\\\n6 & Long & Counter & Smooth & 0.00032 & ----- & 0.00012 & 0.00036 & 0.18 & Both & ----- & Chiefly transcrystalline \\\\\n6 & Long & Counter & Smooth & 0.00030 & ----- & 0.00013 & 0.00017 & .065 & Both & ----- & Crack extends to des- \\\\\n6 & Long & Counter & Smooth & ----- & ----- & ----- & ----- & ----- & ----- & ----- & ignated point. Chiefly \\\\\n6 & Long & Counter & Smooth & ----- & ----- & ----- & ----- & ----- & ----- & ----- & granular or transcrystalline \\\\\n6 & Long & Lower & Smooth & ----- & ----- & 0.0012 & 0.0003 & .50 & Both & ----- & Both \\\\\n6 & Long & Lower & Smooth & ----- & ----- & 0.0021 & 0.00048 & .50 & Both & ----- & Both \\\\\n\\hline\n\\end{tabular}\n\nNotes:\n1.  Absence of transverse sections perpendicular to rolling direction.\n2.  Long, longitudinal section parallel to rolling direction.\n3.  Rough edge, edge which appears jagged or irregular at a magnification of X200 and was caused by punching or by corrosion. An edge is considered rough if cracks, pits, or tears are present. If one or more large cracks are present they are not considered.\n4.  Smooth edge, edge which appears smooth at a magnification of X200.\n5.  Depth of deepest fissure, applies only to machined tears or jaggedness at punched edges and does not apply to main cracks.\n6.  Depth of working groove at punched edges are elongated or worked. This elongation is noticeable at high magnifications because normal grains of metal are elongated. Distortion of grain structure was not evident with a research metallograph and film micrographs.\n7.  Distance of crack from edge was measured on one that was least distorted by polishing. Average thicknesses rather than maximum or minimum values are recorded. The purpose was to obtain approximate measurements when both surface and subsurface cracks were present, the combined thicknesses were measured.\n8.  Fawed, not punched.\n9.  Spot punched, stepped surfaces (side edges).\n\n[Figure: NACA logo]\n```", "timestamp": "2026-07-22T06:10:30.145360+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 5, "total_pages": 149, "image_filename": "19930083192_p5.jpg", "text": "NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nTECHNICAL NOTE 1976\n\nSUMMARY OF INFORMATION RELATING TO GUST\n\nLOADS ON AIRPLANES\n\nBy Philip Donely\n\nS U M M A R Y\n\nAvailable information on gust structure, airplane reactions, and pertinent operating statistics has been examined. This paper attempts to coordinate this information with reference to the prediction of gust loads on airplanes. The material covered represents research up to October 1947.\n\nI N T R O D U C T I O N\n\nThe fact that all airplanes fly in rough air at some time poses a number of problems relative to safe flight. One of the most important of these problems is that of designing the airplane structure to withstand the loads imposed by gusts. The three principal phases of the gust-load problem are: (1) the determination of the gust structure (that is, the size, shape, intensity, and frequency of occurrence), (2) the reaction of any airplane to gusts of known structure, and (3) the determination of the operating statistics.\n\nNo order of importance can be given to the three phases of the problem since the final loads are a function of the gust, the airplane, and how the airplane is flown. The characteristics of gusts, where they occur, and the variation in their characteristics are of fundamental importance because the gust is the source of the problem. Analytical and experimental work on what happens to an airplane when it strikes a known gust is of importance since a knowledge of airplane reactions permits the load calculation for any airplane. Finally, operating statistics are of importance in setting the level of loading for operating airplanes since they define the gusts encountered under actual operating conditions and the speeds at which the gusts are encountered.\n\nAlthough research on gust loads has been carried on for many years and many of the results have been incorporated in design rules (reference 1), these results have either been issued piecemeal or not published at all. Under these conditions, a compilation of available information", "timestamp": "2026-07-22T06:10:33.800553+00:00"}
{"citation_id": "19930085485", "source_url": "https://ntrs.nasa.gov/api/citations/19930085485/downloads/19930085485.pdf", "page_number": 12, "total_pages": 26, "image_filename": "19930085485_p12.jpg", "text": "10\nNACA RM No. L8K02\n\nCONFIDENTIAL\n\nHinge line\n\nFlat-sided\nt = 1.00\n\nFlat-sided\nt = 0.50\n\nFlat-sided\nt = 0.37\n\nCircular-arc\nt = 0\n\nNACA\n\nFigure 2.- Section profiles of the 0.20c aileron tested.\nCONFIDENTIAL", "timestamp": "2026-07-22T06:10:34.706557+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 30, "total_pages": 78, "image_filename": "19930082483_p30.jpg", "text": "28\nNACA TN No. 1807\n\nWhen the inlet total temperature and ratio of inlet total pressure to discharge static pressure are known, the ideal power input per pound of driving fluid based on ratio of inlet total pressure to discharge static pressure may be obtained by use of tables in reference 8.\n\n(7) blade pitch-line velocity $u_m$ from the following equation\n\n$$u_m = \\pi D \\left( \\frac{N}{60} \\right) \\tag{48}$$\n\n(8) theoretical jet velocity $V_j'$ based on ratio of inlet stagnation pressure to discharge stagnation pressure, from the following equation\n\n$$V_j' = \\sqrt{2gJc_p T_1'} \\left[ 1 - \\frac{1}{\\left( \\frac{p_1'}{p_e'} \\right)^{\\frac{\\gamma-1}{\\gamma}}} \\right] \\tag{49}$$\n\nwhere J is the mechanical equivalent of heat.\n\n(9) theoretical jet velocity $V_j$ based on ratio of inlet total pressure to discharge static pressure, from the following equation\n\n$$V_j = \\sqrt{2gJc_p T_1'} \\left[ 1 - \\frac{1}{\\left( \\frac{p_1'}{p_e} \\right)^{\\frac{\\gamma-1}{\\gamma}}} \\right] \\tag{50}$$\n\n(10) velocity ratio $v'$ based on ratio of inlet total pressure to discharge total pressure, from the following equation:\n\n$$v' = \\frac{u_m}{V_j'} \\tag{51}$$\n\n(11) velocity ratio $v$ based on inlet total pressure to discharge static pressure, from the following equation:\n\n$$v = \\frac{u_m}{V_j} \\tag{52}$$", "timestamp": "2026-07-22T06:10:40.738142+00:00"}
{"citation_id": "19930082566", "source_url": "https://ntrs.nasa.gov/api/citations/19930082566/downloads/19930082566.pdf", "page_number": 19, "total_pages": 44, "image_filename": "19930082566_p19.jpg", "text": "NACA TN No. 1889\n\nAir-escape\nplug E\n\nSpecimen S\n\nLow-pressure\ngage L\n\nHigh-pressure\ngage H\n\nElectrical leads from\nmicroswitch\n\nMicroswitch M\nPressure valve P\nActuating microswitch\n\nReservoir R\n\nValve F\n\nValve block G\n\nCheck valve C\n\nAccumulator A\n\nValve D\n\nPump B\n\nNACA\n\nDiesel injection\npump I\n\nFigure 5.- Internal-pressure loading arrangement.\n\n17", "timestamp": "2026-07-22T06:10:42.031005+00:00"}
{"citation_id": "19930093769", "source_url": "https://ntrs.nasa.gov/api/citations/19930093769/downloads/19930093769.pdf", "page_number": 34, "total_pages": 39, "image_filename": "19930093769_p34.jpg", "text": "NACA RM No. E8L10a\nCONFIDENTIAL\n33\n\n1070\n\nAltitude, ft\n50,000\n40,000\n30,000\n20,000\n10,000\n0\n0 .2 .4 .6\nFlight Mach number\n\no Satisfactory start\n□ Burner ignited but\nacceleration impossible\n◇ No ignition\n\nRegion of unsatisfactory\nstarts\n\nNACA\n\n(b) Fuel, AN-F-58; extended-electrode\nspark plug; plug not cleaned\nbetween runs.\n\nFigure 9. - Continued. Windmilling-starting characteristics.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:10:43.049674+00:00"}
{"citation_id": "19930082592", "source_url": "https://ntrs.nasa.gov/api/citations/19930082592/downloads/19930082592.pdf", "page_number": 18, "total_pages": 50, "image_filename": "19930082592_p18.jpg", "text": "NACA TN 1914\n17\n\nMolybdenum oxidation-rate constant, $\\Delta p/\\Delta t = K_{Mo}$\n\nMolybdenum\n(percent)\n$\\circ$ 5\n$\\square$ 10\n$\\diamond$ 20\n$\\triangle$ 30\n\nReciprocal of absolute temperature in $^\\circ R$, $1/T$\nTemperature, $^\\circ F$\n\n[Figure: Graph showing the effect of temperature and molybdenum content on oxidation-rate constant of titanium carbide - molybdenum cermamels.]\n\nFigure 2. - Effect of temperature and molybdenum content\non oxidation-rate constant of titanium carbide - molybdenum\nceramels.", "timestamp": "2026-07-22T06:10:43.436372+00:00"}
{"citation_id": "19930083221", "source_url": "https://ntrs.nasa.gov/api/citations/19930083221/downloads/19930083221.pdf", "page_number": 7, "total_pages": 47, "image_filename": "19930083221_p7.jpg", "text": "NACA TN No. 1824\n\n(a) The ratios of the perturbation velocities to the free-stream velocity are small enough to be neglected when compared to one.\n\n(b) The velocity gradients at a given point of the flow field are all of similar magnitude.\n\nWith the aid of these assumptions, it follows that, to the order of the approximations made, the perturbation velocity potential $\\varphi$ satisfies the well-known linear equation\n\n$$\n(1 - M_0^2)\\ \\varphi_{xx} + \\varphi_{yy} + \\varphi_{zz} = 0 \\tag{7}\n$$\n\nIn the case of two-dimensional flow, the equation is independent of $y$ and thus may be written in the form\n\n$$\n(1 - M_0^2)\\ \\varphi_{xx} + \\varphi_{zz} = 0 \\tag{8}\n$$\n\nTwo- and three-dimensional nonlinear equations, $M_0 = 1$. —\n\nThe study of equation (8) in both subsonic and supersonic flow has shown that for arbitrary lifting surfaces or symmetrical nonlifting airfoils the value of the induced velocity $u$ on the surface of a fixed geometric configuration is proportional to $(|1 - M_0^2|)^{-1/2}$. In all airfoil problems, the value of $u$ becomes infinitely large as $M_0$ approaches one, either from above or below, and the basic assumptions are thus violated. Such a difficulty led Oswatitsch and Wieghardt (reference 3) and Sauer (reference 4) to abandon the restriction of linearity and to seek a more exact equation at $M_0 = 1$. Retaining the assumptions underlying equation (6) and setting $V_0 = a^*$ where $a^*$ is the critical speed of sound, it follows that at $M_0 = 1$ the perturbation velocity potential satisfies the equation\n\n$$\n\\frac{(\\gamma+1)}{a^*}\\ \\varphi_x\\ \\varphi_{xx} - \\varphi_{zz} + \\frac{2}{a^*}\\ \\varphi_z\\ \\varphi_{xz} = 0 \\tag{9}\n$$\n\nSince $\\varphi_x$ is much larger than $\\varphi_z$ as the Mach number approaches one, equation (9) may be further simplified to\n\n$$\n\\frac{(\\gamma+1)}{a^*}\\ \\varphi_x\\ \\varphi_{xx} - \\varphi_{zz} = 0 \\tag{10}\n$$\n\nIf, in three dimensions, the perturbation velocities do not remain small, equation (6) again supplies the necessary form of the differential equation at $M_0 = 1$. From the relation $V_0 = a^*$, the", "timestamp": "2026-07-22T06:10:47.871590+00:00"}
{"citation_id": "19930085548", "source_url": "https://ntrs.nasa.gov/api/citations/19930085548/downloads/19930085548.pdf", "page_number": 6, "total_pages": 46, "image_filename": "19930085548_p6.jpg", "text": "NACA RM No. E8L30\n\nData were recorded at a minimum of 15-minute intervals to permit stabilization of operating temperatures. Friction losses were determined by motoring the engine after each point at which data were taken.\n\nThe following engine operating conditions were used:\n\nPort timing, deg A.T.C.\n\nExhaust opens . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 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. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 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. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .", "timestamp": "2026-07-22T06:10:48.283476+00:00"}
{"citation_id": "19930082703", "source_url": "https://ntrs.nasa.gov/api/citations/19930082703/downloads/19930082703.pdf", "page_number": 12, "total_pages": 28, "image_filename": "19930082703_p12.jpg", "text": "```markdown\n10\nNACA TN 1983\n\nadequate for present purposes. The equation for the time history of\nthe normal-acceleration increment above 1 g was obtained by solving the\nfour differential equations simultaneously by the method of the Laplace\ntransformation. The conventional method could also have been used but\nwith considerable sacrifice in ease of solution. The following two\nequations were obtained:\n\nHelicopter A:\n\n$$ \\Delta N = 0.10e^{-2.06t} - 0.088e^{-0.28t} + 0.48e^{0.38t} \\sin(14.5t + 5.31)^\\circ $$\n\nHelicopter B:\n\n$$ \\Delta N = 0.34e^{-0.028t} \\sin(23.15t + 58.1)^\\circ - 0.45e^{-0.865t} \\sin(47.0t + 30.8)^\\circ $$\n\nwhere $\\Delta N$ is the acceleration increment caused by the control displace-\nment which was made at time $t = 0$.\n\nThese two equations are plotted in figure 9, which shows that the\ndivergent characteristics of helicopter A and the nondivergent charac-\nteristics of helicopter B can be theoretically predicted. Helicopter B,\nbecause of its tail surface, has different values from helicopter A\nfor three stability parameters. Moments due to pitching velocity are\nincreased 20 to 30 percent, moments due to speed change are increased\n80 to 90 percent, and moments due to angle-of-attack change are changed\nfrom unstable to stable with about one-half the magnitude. Further\ncalculations show, however, that the change in pull-up characteristics\nfrom helicopter A to helicopter B is primarily due to the change in the\nvalue of the moment increment per unit angle-of-attack increment.\n\nFigure 9 also shows that for both helicopters the jump in accelera-\ntion at $t = 0$ and the flat spot at the start of the acceleration time\nhistory, as well as the general shape of the curve can be theoretically\npredicted. The cause of this flat spot can be explained as follows:\nThe abrupt rearward control displacement causes an abrupt increase in\nrotor angle of attack and thus an increase in rotor thrust and normal\nacceleration. The resulting curvature of the flight path results in a\nclimb, which tends to reduce the rotor angle of attack and normal\nacceleration from their abruptly increased values back to their trim\nvalues. In the meantime, however, the nose-up moment produced by the\ncontrol displacement tends to cause a nose-up pitching velocity and\nthus an increase in rotor angle of attack and normal acceleration.\nThese two opposing tendencies result in a flat spot in the normal-\nacceleration time history. Various means exist whereby the relative\nproportions of these two tendencies can be altered, for example, by\n```", "timestamp": "2026-07-22T06:10:50.593012+00:00"}
{"citation_id": "19930086076", "source_url": "https://ntrs.nasa.gov/api/citations/19930086076/downloads/19930086076.pdf", "page_number": 50, "total_pages": 50, "image_filename": "19930086076_p50.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:10:52.288901+00:00"}
{"citation_id": "19930085847", "source_url": "https://ntrs.nasa.gov/api/citations/19930085847/downloads/19930085847.pdf", "page_number": 3, "total_pages": 32, "image_filename": "19930085847_p3.jpg", "text": "NACA RM A9D04\nCLASSIFICATION CANCELLED\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nRESEARCH MEMORANDUM\n\nFLIGHT INVESTIGATION OF THE EFFECT OF BOUNDARY-\nLAYER SUCTION ON PROFILE-DRAG COEFFICIENT\nAT SUPERCRITICAL MACH NUMBERS\n\nBy Richard B. Skoog\n\nSUMMARY\n\nFlight tests were conducted with a fighter airplane to study the\neffect of boundary-layer suction aft of the shock wave on airfoil drag\nat supercritical Mach numbers and high Reynolds numbers. A suction\nslot was placed at about 70-percent chord, approximately 7-percent chord\naft of the shock location at the highest test Mach number. Airfoil\nchord force was determined from pressure-distribution measurements\nobtained at Mach numbers of 0.70 to 0.83 in steady dives. Wake survey\nmeasurements were also made but over the lesser Mach number range from\n0.70 to 0.78. The approximate Mach number for drag divergence was 0.73.\n\nResults of the tests showed no measurable effect of suction for\nthe suction coefficient available. Even under conditions where flow\nseparation was present the drag increase with Mach number was due\nprimarily to pressure changes associated with supersonic flow on upper\nand lower surfaces which resulted in increased pressure drag.\n\nINTRODUCTION\n\nAs has been known for some time, an abrupt rise in the drag\ncoefficient of an airfoil occurs at high subsonic Mach numbers due\nto flow changes associated with the occurrence of local regions of\nsupersonic flow near the airfoil. One of these flow changes is the\nboundary-layer growth or separation accompanying the shock formation\nwhich terminates a supersonic region. This boundary-layer behavior\ndevelops because of the steep adverse pressure gradient at the base\nof the shock. In view of this action of the boundary layer, there has\nbeen renewed interest in the possibilities of boundary-layer control\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:11:03.682714+00:00"}
{"citation_id": "19930085572", "source_url": "https://ntrs.nasa.gov/api/citations/19930085572/downloads/19930085572.pdf", "page_number": 5, "total_pages": 17, "image_filename": "19930085572_p5.jpg", "text": "NACA RM No. E5L02\n3\n\nRESULTS AND DISCUSSION\n\nA comparative performance of AN-F-58 and AN-F-32 fuels is presented by means of a direct comparison of the jet-engine performance parameters obtained when operating with each fuel.\n\nThe comparison of corrected net thrust at a Mach number of 0.37 with both fuels is shown in figure 2. Above a corrected engine speed of 4000 rpm, the thrust obtained with AN-F-58 fuel was approximately 150 to 200 pounds greater than that obtained with AN-F-32 fuel. The difference is approximately 3 percent at a corrected engine speed of 8000 rpm. This small difference in the fuels is believed to be a result of the error in reproducibility of test conditions in flight.\n\nA comparison of corrected jet-fuel consumption with the two fuels is presented in figure 3. Below a corrected engine speed of 5500 rpm, no difference occurred. At higher engine speeds, however, AN-F-58 fuel consumption increased gradually to a value of approximately 160 pounds per hour higher than the AN-F-32 fuel at the maximum engine speed. The value of 160 pounds per hour is approximately a 3-percent difference at a corrected engine speed of 8000 rpm. The net thrust, however, was also 3 percent higher for the AN-F-58 fuel than for the AN-F-32 fuel (fig. 2). Consequently, the corrected specific fuel consumption was the same at the higher speed as shown in figure 4. At engine speeds below 6750 rpm, the AN-F-58 fuel results show a somewhat lower specific fuel consumption than the AN-F-32 fuel. The close agreement of the specific-fuel-consumption data indicates that the combustion efficiency with the two fuels was equal at the high engine speeds.\n\nThe variation of corrected tail-pipe temperature with corrected engine speed for the two fuels is shown in figure 5. These data show no difference in tail-pipe temperature at the maximum engine speed but show an approximately 500° R higher temperature at the low engine speeds with AN-F-58 fuel. This temperature difference is consistent with the differences in net thrust and jet-fuel consumption (figs. 2 and 3) and is additional evidence that the performance discrepancy is due to the reproducibility limits of engine operation or test conditions rather than to a difference in fuel performance.\n\nCombustor blow-out speeds at pressure altitudes from 5,000 to 30,000 feet at a Mach number of 0.37 are presented in figure 6. These blow-out speeds are the engine speeds at which a sudden drop in tail-pipe temperature to approximately 100° F occurred as the", "timestamp": "2026-07-22T06:11:04.975660+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 41, "total_pages": 114, "image_filename": "19930086061_p41.jpg", "text": "```markdown\nNACA RM L59J07\n\n$\\alpha=34.1^\\circ$\n$C_L=.038$\n\n$\\alpha=39.1^\\circ$\n$C_L=.031$\n\n$\\alpha=44.1^\\circ$\n$C_L=.067$\n\n$\\alpha=48.1^\\circ$\n$C_L=.059$\n\n-3\n-2\nP-1\n0\n1\n\nStation 4\n$\\frac{y}{b/2}, .0500$\n\n-3\n-2\nP-1\n0\n1\n\n--- Upper\n--- Lower\n\nStation 5\n$\\frac{y}{b/2}, .0667$\n\n-2\n-1\nP\n0\n1\n\nStation 6\n$\\frac{y}{b/2}, .0833$\n\n-2\n-1\nP\n0\n1\n\nStation 7\n$\\frac{y}{b/2}, .0916$\n\n0 2 4 6 8 10\nx/c\n\n0 2 4 6 8 10\nx/c\n\n0 2 4 6 8 10\nx/c\n\n0 2 4 6 8 10\nx/c\n\n(b) Stations: 4,5,6,7.\n\nFigure 7.- Concluded.\n\n37\n```", "timestamp": "2026-07-22T06:11:09.959801+00:00"}