PDF/ocr_dataset_1058.jsonl

{"citation_id": "19930082613", "source_url": "https://ntrs.nasa.gov/api/citations/19930082613/downloads/19930082613.pdf", "page_number": 16, "total_pages": 46, "image_filename": "19930082613_p16.jpg", "text": "```markdown\nNACA TN 1938\n15\n\nTABLE II - VARIATION OF MICROSTRUCTURE AND HARDNESS OF INCONEL SHEET WITH HEAT TREATMENT\n\n| Sample | Heat treatment | A.S.T.M. grain size | Etching time in 10-percent sodium cyanide (electrolytic) (min) | Range of Rockwell B90T superficial hardness | Range of Rockwell B hardness (commercial values) | Description of microstructure |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| 1 | Dead-soft stock | 3-7 | 10 | 61-62 | 66.5-68 | Grains at edges are worked to approximate depths of 0.001 to 0.002 inch. Boundaries contain fine solid precipitates and some globules. Interior grains are equiaxed (fig. 18 (a)). |\n| 2 | Annealed at 900° F for 2 hours, air-cooled | 5-7 | 15 | 58-61 | 62-66.5 | Same as sample 1 (fig. 16 (b)). |\n| 3 | Annealed at 1600° F for 2 hours, air-cooled | 6-8 | 5 | 61-62 | 66.5-68 | Grain boundaries at edges of punched hole are channeled and show mechanical working. Slight evidence of grain growth exists in interior. Boundaries of some grains contain a large number of solid and globular precipitates (fig. 16 (c)). |\n| 4 | Annealed at 1600° F for 2 hours, water-quenched | 4-8 | 45 | 59.5-62.5 | 64-69 | Same as sample 3. |\n| 5 | Annealed at 2200° F for 2 hours, air-cooled | 1-4 | 30 | 48-51 | 47.5-52 | Grains are not larger at punched edge as compared with interior. Boundaries are very difficult to etch because there are almost no visible precipitates (fig. 16 (d)). Tight scale was formed on edges. |\n| 6 | Annealed at 2200° F for 2 hours, water-quenched | 3-5 | 45 | 52-53.5 | 53.5-55.5 | Same as sample 5. |\n\n[Figure: NACA]\n\n$^a$Small hole drilled through the specimens caused an irregular current density and an abnormal etching time.\n```", "timestamp": "2026-07-22T06:11:11.994508+00:00"}
{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 7, "total_pages": 72, "image_filename": "19930085491_p7.jpg", "text": "6\nCONFIDENTIAL\nNACA RM No. A8J04\n\nX streamwise distance from midchord of mean aerodynamic chord to center of pressure measured positive when center of pressure is ahead of midchord, inches\n\nY lateral coordinate, inches\n\n$\\Lambda_{L.E.}$ sweep angle of leading edge, degrees\n\n[Figure: Diagram showing wing chord, velocity vector $V_o$, lift coefficient change $\\Delta C_L$, drag coefficient change $\\Delta C_D$, resultant force vector change, angle of attack $\\alpha$, change in angle of attack $\\Delta \\alpha$, minimum drag angle $\\alpha_{D=min}$, and rearward inclination $\\alpha_{\\Delta L}$]\n\n$\\alpha$ angle of attack, radians (unless otherwise specified)\n\n$\\Delta \\alpha$ change in angle of attack from value for minimum drag, radians (unless otherwise specified)\n\n$\\alpha_{\\Delta L}$ rearward inclination of the change in resultant force corresponding to the change in lift coefficient $\\Delta C_L$, radians (unless otherwise specified)\n\nSubscripts\n\nL=0 value at zero lift\n\nD=min value at minimum drag\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:11:14.893380+00:00"}
{"citation_id": "19930082566", "source_url": "https://ntrs.nasa.gov/api/citations/19930082566/downloads/19930082566.pdf", "page_number": 20, "total_pages": 44, "image_filename": "19930082566_p20.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:11:15.771767+00:00"}
{"citation_id": "19930082646", "source_url": "https://ntrs.nasa.gov/api/citations/19930082646/downloads/19930082646.pdf", "page_number": 8, "total_pages": 37, "image_filename": "19930082646_p8.jpg", "text": "NACA TN 1980\n7\n\nand 39 percent less than those of the extended afterbody alone. The maximum negative angular accelerations remained relatively unchanged.\n\nThe maximum and minimum values of trim and rise at the greatest cycle of oscillation during each landing are plotted against wave length in figure 18. The maximum amplitudes of oscillation in trim of the modified hull were approximately 25 percent lower than those of the basic hull and of the warped-forebody configuration, and were substantially the same as those of the extended-afterbody configuration. The modified hull reduced the maximum amplitudes of oscillation in rise of the basic hull and of the warped-forebody configuration by approximately 20 percent, and increased those of the extended-afterbody configuration by approximately the same percentage.\n\nSummary Chart\n\nThe hydrodynamic qualities in smooth water of the flying boat with a high-length-beam-ratio hull having a warped forebody and an extended afterbody are summarized in figure 19. This chart gives an over-all picture of the hydrodynamic characteristics in terms of full-scale operational parameters. It is therefore useful for comparisons with similar data regarding other seaplanes for which operating experience is available.\n\nCONCLUSIONS\n\nThe effects of combining a warped forebody and an extended afterbody on a hull having a high length-beam ratio are as follows:\n\n1. The stable range of trims available for take-off was substantially increased throughout the entire speed range to take-off.\n\n2. The center-of-gravity limits of stability were improved, chiefly by the elimination of a practical after limit.\n\n3. The smooth-water landing stability characteristics were approximately the same as those for the basic hull.\n\n4. Spray characteristics were considerably improved; there was no propeller or flap spray at design gross load. Tail spray encountered during landings was considerably reduced.", "timestamp": "2026-07-22T06:11:16.355768+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 6, "total_pages": 149, "image_filename": "19930083192_p6.jpg", "text": "2\nNACA TN 1976\n\nwithin a logical framework is needed and, for this purpose, the present paper has been prepared. The scope includes all NACA material available up to about October 1947 arranged according to the three principal phases previously mentioned.\n\nBecause airplanes are used to obtain gust data and the various phases of research are interdependent, the first section of the paper is a presentation and discussion of the basic equations of the reactions of an airplane as used in gust-load calculations. This first section is followed by sections covering gust characteristics, airplane reactions, and operating statistics and by a short concluding section on the coordination of information and the calculation of applied loads. The material contained in this paper is based mainly on past work of the NACA, much of which has been accomplished through the cooperation of the agencies and airlines listed in appendix A.\n\nBACKGROUND AND BASIC EQUATIONS OF\nGUST-LOADS RESEARCH\n\nIn the earliest stages of research on gust loads, an immediate practical answer to the question of appropriate design loads was needed. The problem was approached by recourse to the obvious and necessary expedient of measuring the accelerations under actual operating conditions so that some data could be made available to designers pending a more complete understanding of gust-load phenomena.\n\nAccelerometers were used in this work although it was evident that accelerations measured on one type of airplane would not be applicable to other types, or even to the same type operated at a different speed. The main effects of airplane characteristics and airspeed were taken into consideration in overcoming this difficulty by utilizing a formula based on the most elementary concepts of the nature and action of gusts. This formula, called the \"sharp-edge-gust formula,\" was applied in such a way that the accelerometer data were reduced to \"effective\" gust velocities, the term effective being employed to indicate the fictitious nature of these gust velocities.\n\nSince that time, several advances have been made so that at present three concepts of evaluating airplane reactions are being utilized: (1) the very simple sharp-edge-gust formula, (2) extended equations taking into account the gust shape and additional airplane characteristics, and, finally, (3) the sharp-edge-gust formula modified by an alleviation factor. These relations and concepts are presented subsequently and form the basis of NACA research on gust loads.", "timestamp": "2026-07-22T06:11:17.571055+00:00"}
{"citation_id": "19930093769", "source_url": "https://ntrs.nasa.gov/api/citations/19930093769/downloads/19930093769.pdf", "page_number": 35, "total_pages": 39, "image_filename": "19930093769_p35.jpg", "text": "34\nCONFIDENTIAL\nNACA RM No. E8L10a\n\n1070\n\nO Satisfactory start\n$\\Delta$ Burner ignited but\nacceleration impossible\n$\\diamond$ No ignition\n\nAltitude, ft\n50,000\n40,000\n30,000\n20,000\n10,000\n0\n\nAN-F-58\nfuel\n\nRegion of unsatisfactory\nstarts\n\nNACA\n\n0 .2 .4 .6 .8 1.0 1.2\nFlight Mach number\n\n(c) Fuel, gasoline; standard spark plug; dashed line from\nfigure 9(b).\nFigure 9. - Concluded. Windmilling-starting characteristics.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:11:17.850510+00:00"}
{"citation_id": "19930082918", "source_url": "https://ntrs.nasa.gov/api/citations/19930082918/downloads/19930082918.pdf", "page_number": 9, "total_pages": 62, "image_filename": "19930082918_p9.jpg", "text": "8\nNACA TN 1940\n\naddition, chromium radiation was used to eliminate, in part,\nfluorescence of the sample. Attention was centered on the highest-\norder line of the austenite matrix obtainable with the chromium\nradiation, the (220) line (at a diffraction angle, $\\theta = 65^\\circ$) resulting\nfrom the K$\\alpha_1\\alpha_2$ wave length. In order to record this line photo-\ngraphically, the samples were mounted at the center of a 20-centimeter\ncircular camera with the irradiated area being oscillated $\\pm 10^\\circ$ about\nan axis perpendicular to both the incoming X-ray beam and a diameter\nof the camera. Moll microphotometer recordings of the film showing\nthe (220) $\\alpha_1\\alpha_2$ doublet were then made and the widths of the $\\alpha_1$ line,\ndetermined at half peak intensity, by graphical means. To correct\nfor the presence of the $\\alpha_2$ line, the widths were actually measured as\nhalf widths on the side of the $\\alpha_1$ line away from the $\\alpha_2$ line, the\ndivision being a perpendicular through the apparent $\\alpha_1$ peak. While\nmore rigorous methods are available to correct for the presence of\nthe $\\alpha_2$ line, it is felt that the differences between them and the\nmethod actually used are of second order and that the accuracy of\nthe rest of the technique did not warrant such corrections. Care\nwas taken to make certain that the density range of the $\\alpha_1$-line\nrecording lay within the linear-density and log-intensity range of\nthe film. The microphotometer data were also corrected to read\nintensity against diffraction angle before broadening measurements\nwere made.\n\nPrincipally for use in future research, the line-broadening data\nwere broken down to give the mean lattice deviation $\\frac{\\Delta d}{d}$, where d\nrepresents the interplanar spacings of any given set of planes,\naccording to the method proposed by Haworth (see reference 3). In\nthis connection, it should be mentioned that correction for broadening\ndue to all sources other than lattice strain was done by the method\noriginally put forth by Warren (see reference 4) using unaged solution-\ntreated low-carbon N-155 as a standard. It should be noted then that\nthe line-broadening results are relative to the unaged material.\n\nLattice parameters.- Lattice parameters were measured with a\nSachs and Weerts type camera using copper K$\\alpha_1\\alpha_2$ radiation. The\nsamples were mounted with the prepared plane surface perpendicular\nto the incoming X-ray beam and rotated about an axis parallel with\nand slightly to one side of the axis of the X-ray beam. On the\nirradiated surface a light film of a mixture of mineral oil and\nchemically precipitated silver powder was placed. With the\ncopper K$\\alpha_1\\alpha_2$ radiation used, two lines, among others, were photo-\ngraphically recorded, the (420) lines of the austenite matrix of\nlow-carbon N-155 (at $\\theta = 75^\\circ$) and the (333) lines of the silver\n(at $\\theta = 78^\\circ$). From the measured spacing of the silver (333)\nlines and a lattice parameter of 4.0778Å for the silver, the camera-\nto-film distance was calculated for each exposure. With this", "timestamp": "2026-07-22T06:11:18.130947+00:00"}
{"citation_id": "19930082592", "source_url": "https://ntrs.nasa.gov/api/citations/19930082592/downloads/19930082592.pdf", "page_number": 19, "total_pages": 50, "image_filename": "19930082592_p19.jpg", "text": "```markdown\n18\nNACA TN 1914\n\nMolybdenum oxidation-rate constant, $\\Delta p/\\Delta t = K_{Mo}$\n\n<!-- Image (116, 126, 936, 806) -->\n\nFigure 3. - Effect of molybdenum content on oxidation-rate constant at 2000° F for titanium carbide - molybdenum ceramals.\n```", "timestamp": "2026-07-22T06:11:19.119999+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 31, "total_pages": 78, "image_filename": "19930082483_p31.jpg", "text": "```markdown\nNACA TN No. 1807\n29\n\nA numerical illustration of these calculations is presented in appendix B.\n\nLoss-presentation method. - The power losses corrected to sea level are plotted against corrected turbine rotor speed in figure 9 at a total-pressure ratio of 2.0 with the degree of admission as parameter where applicable. Disk-friction losses are so small that they are not shown in figure 9.\n\nThe ideal power, net power, and applicable losses per pound of driving fluid, all corrected to sea-level conditions, at a pressure ratio of 2.0, are plotted against corrected rotor speed for $360^\\circ$ (full), $180^\\circ$, and $120^\\circ$ admissions in figure 10.\n\nPerformance-Estimation Method\n\nThe turbine power outputs at any two amounts of admission are related by the change in the loss quantities. It is therefore possible to obtain estimates of power output at any degree of admission by making use of the full-admission carpet plot and the loss curves. From equation (30) it can be seen that\n\n(net power estimated)$_F$ = F (power observed)$_{360^\\circ}$ +\n(shaft losses)$_{360^\\circ}$ - (shaft losses)$_F$\n- (driving-fluid losses)$_F$\n(53)\n\nIn order to estimate power at any degree of admission, it remains to evaluate these losses at some degree of admission and to know the manner of variation of the losses with admission. As previously described, disk-friction and bearing losses are invariant with admission, and pumping loss and driving-fluid losses are proportional to the inactive nozzle arc.\n\nWhen the power output has been estimated at any degree of admission with equation (53) the efficiency for that degree of admission can be easily estimated from equation (33). A numerical example is provided in appendix B.\n\nThis simplified method quickly yields good approximations. Upon correction of the turbine output and of correctable losses to sea level, the bearing loss, which has no correction applied, forms a disproportionately large part of the corrected power. This disproportion in turn affects the accuracy of the efficiency estimations based\n\n```", "timestamp": "2026-07-22T06:11:21.412003+00:00"}
{"citation_id": "19930085859", "source_url": "https://ntrs.nasa.gov/api/citations/19930085859/downloads/19930085859.pdf", "page_number": 1, "total_pages": 31, "image_filename": "19930085859_p1.jpg", "text": "NACA RM No. L9B25\nRM No. L9B25\n\n[Figure: NACA logo]\n\nRESEARCH MEMORANDUM\n\nAERODYNAMIC CHARACTERISTICS OF A WING WITH QUARTER-CHORD LINE\nSWEPT BACK $35^\\circ$, ASPECT RATIO 4, TAPER RATIO 0.6,\nAND NACA 65A006 AIRFOIL SECTION\n\nTRANSONIC-BUMP METHOD\n\nBy\n\nWilliam C. Sleeman, Jr. and Robert E. Becht\n\nLangley Aeronautical Laboratory\nLangley Air Force Base, Va.\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\nWASHINGTON\n\nApril 21, 1949\nDeclassified August 18, 1954", "timestamp": "2026-07-22T06:11:29.012712+00:00"}
{"citation_id": "19930082703", "source_url": "https://ntrs.nasa.gov/api/citations/19930082703/downloads/19930082703.pdf", "page_number": 13, "total_pages": 28, "image_filename": "19930082703_p13.jpg", "text": "NACA TN 1983\n\nreducing the helicopter pitching inertia while keeping the same gross weight. Another means would be to alter the manner in which the control action is transmitted to the rotor blades so that, when the stick is deflected, a part of the corresponding blade cyclic-pitch change is delayed somewhat.\n\nInasmuch as the important pull-up characteristics of helicopters A and B could be theoretically predicted, it is concluded that the important longitudinal characteristics of any reasonably similar helicopter for which the necessary parameters are available can also be theoretically predicted.\n\nPractical Requirements\n\nExamination of the time histories presented suggests that a requirement intended to preclude dangerous stick-fixed divergent tendencies, as regards longitudinal stability and control in forward flight, might be worded as follows:\n\nWhen the longitudinal control stick is suddenly displaced rearward 1 inch from trim (while in level flight at the maximum placard speed) and held fixed at this displacement, the time history of normal acceleration shall become concave downward within 2 seconds following the start of the maneuver.\n\nFurther consideration of the present results suggests that a requirement aimed at reducing the difficulty of anticipating the results of a control deflection and, hence, reducing pilot fatigue and thereby further increasing the safety of operation might be worded as follows:\n\nWhen the longitudinal control stick is suddenly displaced rearward 1 inch from trim (while in level flight at the maximum placard speed) and held fixed at this displacement, the time history of normal acceleration should preferably be concave downward throughout the period between the start of the maneuver and the attainment of maximum acceleration, and, in any event, the slope of the normal-acceleration curve must remain positive from the start of the maneuver until the maximum acceleration is approached.\n\nThe demonstration of fulfillment of these requirements involves instrumentation which is not always available and, in any case, involves judgment in the fairing of record lines which have a \"hash\" due to rotor and engine vibrations. As a supplementary requirement, therefore, a requirement based on time histories of oscillations or attempted", "timestamp": "2026-07-22T06:11:30.086209+00:00"}
{"citation_id": "19930083221", "source_url": "https://ntrs.nasa.gov/api/citations/19930083221/downloads/19930083221.pdf", "page_number": 8, "total_pages": 47, "image_filename": "19930083221_p8.jpg", "text": "6\nNACA TN No. 1824\n\nrequired expression is\n\n$$\n\\frac{(\\gamma+1)}{a^*} \\Phi_x \\Phi_{xx} - \\Phi_{yy} - \\Phi_{zz} + \\frac{2}{a^*} \\Phi_z \\Phi_{xz} + \\frac{2}{a^*} \\Phi_y \\Phi_{xy} = 0 \\quad (11)\n$$\n\nTwo- and three-dimensional linear equations, $M_o = 1$. -\n\nEquation (10) has been used by von Kármán (reference 1) to establish similarity rules for two-dimensional transonic flow and is the basis for work continuing at the present time. (See also reference 5.) If at $M_o = 1$ the assumptions made in the linearization process still hold, it follows from either equation (10) or (8) that the differential equation reduces to the form\n\n$$\n\\Phi_{zz} = 0 \\quad (12)\n$$\n\nIt is possible, however, to predict independently from this relation that linearized methods cannot be applied to the calculation of arbitrary airfoil pressure distributions. The range of applicability of such an equation is thus almost nonexistent. On the other hand, the linearized form of equation (11) or (7) at $M_o = 1$ is\n\n$$\n\\Phi_{yy} + \\Phi_{zz} = 0 \\quad (13)\n$$\n\nand from this equation a class of nontrivial solutions can be obtained for particular boundary conditions. Both equations are of parabolic form in the number of dimensions for which they are defined. In the present report, formal solutions satisfying the imposed conditions will be obtained in three dimensions for flat lifting surfaces with swept-back leading edges and for an infinitely long, symmetrical, swept-back wing.\n\nUnsteady State\n\nThe derivation of the steady-state equations for the velocity potential was developed in some detail because of the various results to be obtained. Similar methods can be used when unsteady conditions are to be considered, the differential equation for the velocity potential being now in the form\n\n$$\n\\begin{aligned}\n& - \\frac{1}{a^2} \\left( \\Phi_{tt} + 2\\Phi_x \\Phi_{xt} + 2\\Phi_y \\Phi_{yt} + 2\\Phi_z \\Phi_{zt} \\right) \\\\\n& + \\Phi_{xx} \\left( 1 - \\frac{\\Phi_x^2}{a^2} \\right) + \\Phi_{yy} \\left( 1 - \\frac{\\Phi_y^2}{a^2} \\right) + \\Phi_{zz} \\left( 1 - \\frac{\\Phi_z^2}{a^2} \\right) \\\\\n& - \\frac{2}{a^2} \\Phi_{yz} \\Phi_y \\Phi_z - \\frac{2}{a^2} \\Phi_{zx} \\Phi_z \\Phi_x - \\frac{2}{a^2} \\Phi_{xy} \\Phi_x \\Phi_y = 0 \\quad (14)\n\\end{aligned}\n$$", "timestamp": "2026-07-22T06:11:33.726560+00:00"}
{"citation_id": "19930082546", "source_url": "https://ntrs.nasa.gov/api/citations/19930082546/downloads/19930082546.pdf", "page_number": 22, "total_pages": 65, "image_filename": "19930082546_p22.jpg", "text": "NACA TN No. 1870\n\nradiation resistance is given by equation 7(b) of the appendix. This equation gives the vibration amplitude if the structural damping, mass, and natural frequency of the panel are known. Calculations for a resonant condition using equation 7(b) have been made for comparison with experimental results and these values are shown in figure 22(a). For these calculations, $f_{o} = 130$ cycles per second, $\\frac{C}{C_{c}} = 0.02$ (estimated from shape of resonance peak), and the weight of the panel was 7 pounds per square foot. Equation 7(b) shows that for lower values of the mass and frequency the acoustical radiation resistance becomes of greater importance. A conventional fuselage will therefore have greater damping and the resonances will not be so sharply peaked as in figures 22(a) and 22(b).\n\nEffect of fuselage parameters on fuselage vibration.- The appendix shows that the panel vibration amplitude of the fuselage is a function of oscillating pressure and frequency as well as the mass, rigidity, and damping of the structure. Rigidity is effective in reducing low-frequency vibrations, mass is the most effective in reducing high-frequency vibrations, and wall damping is the most effective in reducing the amplitude of the resonant peaks.\n\nThe present tests showed that the panel vibrated predominantly at the fundamental or lowest excitation frequency of the propeller. This fact has also been found to be the case for an airplane fuselage. Since rigidity is the most effective at the low frequencies, wall vibration may be reduced by increasing wall rigidity, provided, of course, that the resonant condition is far enough removed from the range in which the propeller operates. This increase in wall rigidity was accomplished for the test panels by means of reinforcements which raised the panel resonance frequency to a value higher than the fundamental excitation frequency. This procedure necessarily increases the possibility that the panel may be in resonance with the higher harmonics of the propeller. An inspection of figure 22(c) shows that when the reinforced wooden panel was excited by the four-blade propeller several small resonances occurred at higher frequencies; however, these small resonances seemed to be of little importance.\n\nSince the propeller has numerous exciting harmonics and the walls have numerous modes of vibration, eliminating all resonant conditions is impractical. It is therefore desirable to apply a damping material to the walls to reduce the amplitude of the resonant peaks.\n\nThe first section of the present paper shows that as the tip Mach number is increased, more of the pressure energy goes into the higher harmonics. As indicated in the appendix, the mass of the wall becomes most effective in reducing wall vibration at the higher frequencies. The wall must therefore have sufficient mass to prevent excessive vibration at the high frequencies which predominate at high tip speeds.", "timestamp": "2026-07-22T06:11:36.771850+00:00"}
{"citation_id": "19930082613", "source_url": "https://ntrs.nasa.gov/api/citations/19930082613/downloads/19930082613.pdf", "page_number": 17, "total_pages": 46, "image_filename": "19930082613_p17.jpg", "text": "16\nNACA TN 1938\n\nTABLE III - EFFECTS THAT REAMING, SANDING, AND VAPOR\nBLASTING OF PUNCHED EDGES HAVE UPON CRACKING\n\n| Time in accelerated life runs | | Average number of cracks in seven as-fabricated liners | Average number of cracks in seven reamed, sanded, and vapor-blasted liners |\n| :--- | :--- | :--- | :--- |\n| (hr) | (min) | | |\n| 8 | 20 | 8.43 | 1.86 |\n| 16 | 40 | 20.29 | 9 |\n\n[Figure: NACA logo]", "timestamp": "2026-07-22T06:11:37.440073+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 83, "total_pages": 96, "image_filename": "19930085880_p83.jpg", "text": "NACA RM No. L9C03\n81\n\n[Figure: A line graph plotting Resistance against Wetted area for various speeds. The graph includes a small inset diagram of a triangle above a horizontal line in the upper left corner.]\n\nResistance, lb\n8\n7\n6\n5\n4\n3\n2\n1\n0\n\nSpeed (fps)\n30\n25\n20\n15\n10\n\n0 .05 .10 .15 .20 .25 .30 .35\nWetted area, sq ft\n(b) $\\tau = 8^\\circ$.\nFigure 23.- Continued.\nNACA", "timestamp": "2026-07-22T06:11:37.851879+00:00"}
{"citation_id": "19930082566", "source_url": "https://ntrs.nasa.gov/api/citations/19930082566/downloads/19930082566.pdf", "page_number": 21, "total_pages": 44, "image_filename": "19930082566_p21.jpg", "text": "NACA TN No. 1889\n19\n\n[Figure: Biaxial fatigue testing machine with labeled components: M, P, S, Y, F, G, H, L, C, A, R, T, I, U, B, D, N, K]\n\nFigure 6.- Biaxial fatigue testing machine.", "timestamp": "2026-07-22T06:11:39.224970+00:00"}
{"citation_id": "19930085485", "source_url": "https://ntrs.nasa.gov/api/citations/19930085485/downloads/19930085485.pdf", "page_number": 13, "total_pages": 26, "image_filename": "19930085485_p13.jpg", "text": "CONFIDENTIAL\n\nNACA RM No. L58K02\n\n[Figure: Three-quarter front view of a model mounted on a bump in a wind tunnel, with a small fin-like structure protruding from the top surface.]\n\nNACA\nL-55528\n\nFigure 3.- Three-quarter front view of the model as mounted on the bump\nin the Langley high-speed 7- by 10-foot tunnel.\n\nCONFIDENTIAL\n\n11", "timestamp": "2026-07-22T06:11:40.332199+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 42, "total_pages": 114, "image_filename": "19930086061_p42.jpg", "text": "```markdown\n38\n\n$\\alpha=4.1^\\circ$\n$C_L=0.14$\n\n$\\alpha=8.1^\\circ$\n$C_L=0.32$\n\n$\\alpha=14.1^\\circ$\n$C_L=0.54$\n\n$\\alpha=24.1^\\circ$\n$C_L=0.85$\n\nP\n-4\n-3\n-2\n-1\n0\n1\n\nStation 1\n$\\frac{y}{b/2}, 0$\n\n[Figure: Triangle symbol]\n\nP\n-4\n-3\n-2\n-1\n0\n1\n\n--- Upper\n--- Lower\nTwo dimensional\n(calculated at\nequal $C_L$)\n\nStation 2\n$\\frac{y}{b/2}, 0.167$\n\nP\n-3\n-2\n-1\n0\n1\n\nStation 3\n$\\frac{y}{b/2}, 0.333$\n\n0 2 4 6 8 10\nx/c\n\n0 2 4 6 8 10\nx/c\n\n0 2 4 6 8 10\nx/c\n\n0 2 4 6 8 10\nx/c\n\n(a) Stations: 1, 2, 3.\n\nFigure 8.- Chordwise pressure distribution about wing 2 at angles of attack of $4.1^\\circ$, $8.1^\\circ$, $14.1^\\circ$, and $24.1^\\circ$.\n\nNACA RM L9J07\n```", "timestamp": "2026-07-22T06:11:41.282896+00:00"}
{"citation_id": "19930085572", "source_url": "https://ntrs.nasa.gov/api/citations/19930085572/downloads/19930085572.pdf", "page_number": 6, "total_pages": 17, "image_filename": "19930085572_p6.jpg", "text": "4\nNACA RM No. E8L02\n\nengine throttle was slowly closed. The actual engine speed at which the combustors blew out varied from 2900 to 3450 rpm for the AN-F-58 fuel and from 2600 to 3250 rpm for the AN-F-32 fuel over the range of altitudes. Although the data show the blow-out limit for the AN-F-58 fuel to be somewhat inferior by a value of 200 to 250 rpm, it is believed that this difference in the two fuels is within the normal scatter of data obtained for this type of investigation.\n\nData obtained from windmilling starts and engine accelerations at altitudes from 5,000 to 30,000 feet with the fuels are presented in table II. Successful starts were made at all the altitudes with both fuels. A longer period of time was required, however, to accelerate the engine with AN-F-58 fuel. Different engine operators conducted the experiments with the two fuels and the difference in acceleration time could be attributed to differences in the starting technique. Because reproduction of an acceleration maintaining the same tail-pipe temperature is almost impossible, such a difference in acceleration time would normally be experienced in practice.\n\nVisual observations of the jet exhaust showed that no objectionable smoke trail was produced by either fuel. After $2\\frac{1}{2}$ and $7\\frac{1}{2}$ hours of engine operation with the AN-F-58 fuel, several burners and fuel injectors were removed and examined for carbon deposits. Traces of carbon were found both times on the burner liner, the burner dome, and the fuel injector. These deposits, however, were no more severe than the deposits found in the engine when operated with AN-F-32 fuel for approximately the same periods.\n\nSUMMARY OF RESULTS\n\nFrom a flight investigation conducted to compare the performances of AN-F-58 and AN-F-32 fuels in a 4000-pound-thrust turbojet engine, the following results were obtained:\n\n1. The performance of AN-F-58 fuel was equivalent to that of AN-F-32 fuel for the range of conditions investigated.\n\n2. The investigation of AN-F-58 fuel, compared with that of AN-F-32 fuel, indicated a 3-percent-higher thrust and fuel consumption (same specific fuel consumption) at the high engine speed; a slightly inferior blow-out limit (250 rpm higher); equally successful starts at altitudes between 5,000 and 30,000 feet, but somewhat longer acceleration time; and similarly small carbon deposits", "timestamp": "2026-07-22T06:11:45.446788+00:00"}
{"citation_id": "19930085847", "source_url": "https://ntrs.nasa.gov/api/citations/19930085847/downloads/19930085847.pdf", "page_number": 4, "total_pages": 32, "image_filename": "19930085847_p4.jpg", "text": "2\nCONFIDENTIAL\nNACA RM A9D04\n\nas a means of reducing airplane drag at high Mach numbers, especially\nthose Mach numbers at which separation occurs behind the shock. As far\nas is known, research to date on boundary-layer control at high Mach\nnumbers has been confined to tests on very small models. References 1\nand 2 present the results of tests where suction control was applied\nbehind the shock on 2-inch-chord airfoils yielding a drag reduction of\nthe order of 50 percent.\n\nThe purpose of the present investigation was to study at large\nscale the effect of boundary-layer suction on airfoil drag at super-\ncritical Mach numbers and to determine whether the drag increase due\nto separation could be reduced. Accordingly, an airplane was fitted\nwith a suction slot on the upper surface of the left wing at about 70-\npercent chord. Measurements were made of profile drag (by the wake\nsurvey method) and of pressure distribution (to evaluate chordwise\nforce) at Mach numbers beyond that of drag divergence to determine the\neffect at these speeds of boundary-layer removal on drag.\n\nTEST EQUIPMENT\n\nThe tests of boundary-layer control reported herein were carried\nout on a portion of the left wing of a jet-propelled fighter airplane.\nA picture of the airplane as instrumented for the tests is shown in\nfigure 1.\n\nThe slot configuration used in the tests is shown in figures 2\nand 3. As can be seen from the figures, the slot was located at about\n70-percent chord. The slot was 10.3 percent (24 in.) of the wing semi-\nspan and 2.52 percent (2 in.) of the local wing chord and was located\nat about 42 percent (8 ft, 1-1/4 in.) of the wing semispan from the\nfuselage center line. Air flow was induced through the slot by the\nlow pressure existing at the duct exit. (See figs. 2 and 4.) A wing\nroot bump was installed to increase the pressure difference between\nthe slot entrance and the exit. For a limited series of tests the\nslot length was reduced by one-half in an effort to increase the flow\ncoefficient obtainable.\n\nStandard NACA recording instruments synchronized by a standard\nNACA timer were used to record the following variables: indicated\nairspeed, pressure altitude, normal acceleration, wake-survey total-\nhead decrement and static pressures, boundary-layer total head and\nstatic pressures, and the amount of suction air flow. All recording\ninstruments were installed in the nose compartment except the acceler-\nometer which was installed in the pilot's cockpit. The air temperature\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:11:45.716690+00:00"}
{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 8, "total_pages": 72, "image_filename": "19930085491_p8.jpg", "text": "NACA RM No. A8J04 CONFIDENTIAL 7\n\nexp experimental value\n\ntheo theoretical value\n\nopt value at optimum lift coefficient\n\nl value for lower (parabolic) range of drag curve\n\nEXPERIMENTAL CONSIDERATIONS\n\nWind Tunnel and Balance\n\nThe investigation was performed in the Ames 1- by 3-foot supersonic wind tunnel No. 1 which was fitted temporarily with a fixed nozzle designed for a Mach number of 1.5 providing a 1- by 2-1/2-foot test section. The wind tunnel, electric strain-gage balance, and instrumentation are described in detail in references 5 and 6. A cutaway drawing of the strain-gage balance is shown in figure 1.\n\nModels and Supports\n\nPhotographs of the wings and fuselage used in the investigation are shown in figures 2 and 3 and the design dimensions of the basic configuration are shown in figure 4. The fuselage and wings were constructed of steel and no attempt was made to fair in their junctures. In addition to the design setting of $63^\\circ$ the model was constructed so that leading-edge settings of $57.0^\\circ$, $60.4^\\circ$, $67.0^\\circ$, and $69.9^\\circ$ could also be tested. The setting of $60.4^\\circ$ locates the trailing edge at the Mach angle corresponding to a Mach number of 1.53 and those of $57.0^\\circ$ and $69.9^\\circ$ were the limits of the range attainable. For purposes of brevity, the wing-fuselage configurations will be designated by the letters WF followed by a two-digit number giving the leading-edge sweep angle to the nearest whole degree. Thus the basic wing-fuselage configuration is designated WF-63.\n\nThe streamwise airfoil section of the wing of WF-63 is an NACA 64A006. A section having a rounded leading edge was employed in an attempt to realize the leading-edge suction force predicted by theory when the wing leading edge is swept within the Mach cone. Because it was desirable to have the wing thickness-chord ratio as small as possible to minimize the pressure drag and still obtain a wing that was structurally practical, the limitations of present-day\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:11:47.084372+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 44, "total_pages": 98, "image_filename": "19930086073_p44.jpg", "text": "42\nNACA RM A9H04\n\nLift coefficient, $C_L$\nPitching-moment coefficient, $C_m$\n\n| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | |", "timestamp": "2026-07-22T06:11:48.302283+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 7, "total_pages": 149, "image_filename": "19930083192_p7.jpg", "text": "NACA TN 1976\n3\n\nSHARP-EDGE-GUST FORMULA\n\nThe simple sharp-edge-gust formula was derived on the basis of numerous simplifying assumptions which include the following:\n\n1. The gust is sharp-edged in the direction of flight and represents an instantaneous change in wind direction or speed.\n\n2. The gust velocity is uniform across the span of the airplane at any instant of time or position of the airplane in space.\n\n3. The gust direction is normal to the lateral axis of the airplane.\n\n4. The airplane is in steady level flight prior to entry into the gust, and the airplane flight path, attitude, and ground speed are not affected by the action of the gust on the airplane. (This assumption might be visualized by imagining the airplane to be driven along a fixed track through the gust so that the airplane has no resulting motions due to the action of the gust.)\n\n5. The primary effect of encountering a gust is to change the lift on the airplane.\n\n6. The lift increment of the horizontal tail due to the action of the gust is negligible as compared to the wing lift increment.\n\n7. The lift coefficient of a wing is a unique function of angle of attack and independent of time.\n\nFor the assumptions noted, the expression for lift or load factor for a gust of any angle is:\n\n$$\n\\frac{L}{W} = n = 1 \\pm \\Delta n = \\left[ 1 + \\left( \\frac{U}{V} \\right)^2 + \\frac{2U}{V} \\cos \\beta \\right] \\left( 1 + \\frac{dC_L}{d\\alpha} \\frac{S}{W} \\frac{\\rho V^2}{2} \\tan^{-1} \\frac{\\frac{U}{V}}{1 + \\frac{U}{V} \\cos \\beta} \\right)\n$$\n\n(See appendix B for definitions of symbols.) The term in brackets represents the effect of the increased resultant speed and the term in parentheses, the lift in steady level flight plus the lift increment", "timestamp": "2026-07-22T06:11:50.886593+00:00"}
{"citation_id": "19930093769", "source_url": "https://ntrs.nasa.gov/api/citations/19930093769/downloads/19930093769.pdf", "page_number": 36, "total_pages": 39, "image_filename": "19930093769_p36.jpg", "text": "```markdown\nCONFIDENTIAL\n\nNACA RM No. E8L10a\n\nFuel\n- $\\Delta$ AN-F-58\n- --$\\square$-- Gasoline\n\nTurbine-outlet temperature, $^\\circ$F\n\nBlade tip\n\nBlade root\n\nMaximum\nAverage\nMinimum\n\nMaximum\nAverage\nMinimum\n\nDistance across turbine annulus, in.\n\n(a) Altitude, 20,000 feet; engine speed, approximately 12,025 rpm.\nFigure 10. - Radial temperature distribution at turbine outlet for AN-F-58 fuel and gasoline.\nFlight Mach number, 0.85.\n\nNACA\n\nCONFIDENTIAL\n\n35\n```", "timestamp": "2026-07-22T06:11:51.440253+00:00"}
{"citation_id": "19930085859", "source_url": "https://ntrs.nasa.gov/api/citations/19930085859/downloads/19930085859.pdf", "page_number": 2, "total_pages": 31, "image_filename": "19930085859_p2.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:11:53.165076+00:00"}
{"citation_id": "19930082646", "source_url": "https://ntrs.nasa.gov/api/citations/19930082646/downloads/19930082646.pdf", "page_number": 9, "total_pages": 37, "image_filename": "19930082646_p9.jpg", "text": "8\nNACA TN 1980\n\n5. Rough-water landing behavior was improved; reductions in the maximum impact acceleration of greater than 50 percent and in the maximum angular acceleration of approximately 60 percent were obtained as compared with those for the basic model.\n\n6. During landings in waves the maximum amplitudes of oscillation in trim and rise attained during the high-speed portion of the run-out were reduced.\n\n7. In general, the effect of the combination of the hull modifications (warped forebody and extended afterbody) on hydrodynamic characteristics was in the direction expected on the basis of the cumulative effect of the separate modifications.\n\nLangley Aeronautical Laboratory\nNational Advisory Committee for Aeronautics\nLangley Air Force Base, Va., March 9, 1949\n\nREFERENCES\n\n1. Carter, Arthur W., and Weinstein, Irving: Effect of Forebody Warp on the Hydrodynamic Qualities of a Hypothetical Flying Boat Having a Hull Length-Beam Ratio of 15. NACA TN 1828, 1949.\n\n2. Kapryan, Walter J., and Clement, Eugene P.: Effect of Increase in Afterbody Length on the Hydrodynamic Qualities of a Flying-Boat Hull of High Length-Beam Ratio. NACA TN 1893, 1949.\n\n3. Carter, Arthur W., and Haar, Marvin I.: Hydrodynamic Qualities of a Hypothetical Flying Boat with a Low-Drag Hull Having a Length-Beam Ratio of 15. NACA TN 1570, 1948.\n\n4. Carter, Arthur W.: Effect of Hull Length-Beam Ratio on the Hydrodynamic Characteristics of Flying Boats in Waves. NACA TN 1782, 1949.\n\n5. Yates, Campbell C., and Riebe, John M.: Effect of Length-Beam Ratio on the Aerodynamic Characteristics of Flying-Boat Hulls. NACA TN 1305, 1947.\n\n6. Truscott, Starr: The Enlarged N.A.C.A. Tank, and Some of Its Work. NACA TM 918, 1939.\n\n7. Benson, James M., and Bidwell, Jerold M.: Bibliography and Review of Information Relating to the Hydrodynamics of Seaplanes. NACA ACR L5G28, 1945.", "timestamp": "2026-07-22T06:11:57.515580+00:00"}
{"citation_id": "19930082592", "source_url": "https://ntrs.nasa.gov/api/citations/19930082592/downloads/19930082592.pdf", "page_number": 20, "total_pages": 50, "image_filename": "19930082592_p20.jpg", "text": "NACA TN 1914\n19\n\n<!-- Image (185, 258, 688, 676) -->\n\n(a) Temperature, 1625° F.\nFigure 4. - Effect of time, temperature, and tungsten content on oxidation penetration of titanium carbide - tungsten cermamals.", "timestamp": "2026-07-22T06:11:57.653551+00:00"}
{"citation_id": "19930082918", "source_url": "https://ntrs.nasa.gov/api/citations/19930082918/downloads/19930082918.pdf", "page_number": 10, "total_pages": 62, "image_filename": "19930082918_p10.jpg", "text": "NACA TN 1940\n\ncalculated distance (which averaged approximately 10 cm) and measurement of the low-carbon N-155 (420)-line spacing, the lattice parameter for each of the various aged samples of solution-treated low-carbon N-155 was calculated. The absolute accuracy of these calculated parameters was not established, but the relative error was estimated to be ±0.0005Å.\n\nHardness Surveys\n\nHardness surveys were made on the various samples with a Brinell machine using a 10-millimeter ball and a 3000-kilogram load. Two impressions were made on each sample, on planes which were originally transverse planes of the bar stock. Two perpendicular diameters of each impression were measured, and the resulting four readings were averaged and converted to the Brinell hardness number.\n\nCreep Measurements\n\nFor the measurement of creep, 0.250-inch-diameter specimens were prepared with the axis of each specimen corresponding to the original axis of the bar stock. For the majority of tests the gage section was $1\\frac{1}{2}$ inches long terminated at each end with a 1/8-inch radius to the shoulder which in turn was 1/2 inch in diameter. Through the shoulders, 0.100-inch-diameter holes were diametrically drilled in which Chromel pins were driven. A modified Martens extensometer was attached to these pins for measurement of the extension under load. The least reading of this extensometer and associated telescope and scale was $10^{-5}$ inch.\n\nFor an additional check against the accuracy of extension measurements obtained in this manner, duplicate tests were run at the 60,000-psi stress level on samples aged at 1400° F and a different extensometer system was used. Here the samples had a gage length of $1\\frac{1}{4}$ inches terminated at each end and with a 1/4-inch radius to a 3/8-inch-diameter shoulder threaded its entire length. Collars were then threaded onto these shoulders, down to the 1/4-inch radii, and locked in place with setscrews. The extensometer system, in turn, was suspended from these collars with the use of pins mounted in the collars. Temperatures were controlled to ±5° F throughout the tests and the temperature differences along the gage length held to ±3° F. The furnaces used were electrical-resistance type split along a transverse section at the center for ease of control of the uniformity of the gage-length temperature. Load was applied to the specimen through a beam system with a mechanical advantage of 23.", "timestamp": "2026-07-22T06:11:59.736528+00:00"}
{"citation_id": "19930082613", "source_url": "https://ntrs.nasa.gov/api/citations/19930082613/downloads/19930082613.pdf", "page_number": 18, "total_pages": 46, "image_filename": "19930082613_p18.jpg", "text": "NACA TN 1938\n17\n\n[Figure: Two diagrams of cylindrical combustion-chamber liners, labeled \"Type-A liner\" and \"Type-B liner\", showing intake and exhaust ends, with rows of holes and numbered sections.]\n\nIntake\nRow\n1\n2\n3\nExhaust\nType-A liner\n\nIntake\nRow\n1\n2\n3\n4\nExhaust\nType-B liner\n\nNACA\n\nFigure 1. - Combustion-chamber liner investigated.", "timestamp": "2026-07-22T06:12:03.387648+00:00"}
{"citation_id": "19930083221", "source_url": "https://ntrs.nasa.gov/api/citations/19930083221/downloads/19930083221.pdf", "page_number": 9, "total_pages": 47, "image_filename": "19930083221_p9.jpg", "text": "NACA TN No. 1824\n\nwhere $t'$ represents time. The details of the derivation can, however, be avoided by referring directly to the equation satisfied by the velocity potential for the propagation of sound waves of small amplitude. (See reference 6, p. 492.) In this form of the equation the Cartesian coordinate system $x$, $y$, $z$ is assumed fixed in the medium so that free-stream velocity is zero, while the wing, which moves in the direction of the negative $x$ axis with velocity $V_0$, generates small pressure disturbances. As a consequence, the velocity potential of the field satisfies the well-known wave equation in three space dimensions:\n\n$$\n\\frac{1}{a^2} \\varphi_{t't'} - \\varphi_{xx} - \\varphi_{yy} - \\varphi_{zz} = 0 \\tag{15}\n$$\n\nEquation (15) is reducible to canonical form by means of the relation\n\n$$\nt = a_0 t'\n$$\n\nand the three-dimensional form of the equation is therefore\n\n$$\n\\varphi_{tt} - \\varphi_{xx} - \\varphi_{yy} - \\varphi_{zz} = 0 \\tag{16}\n$$\n\nwhile in the two-dimensional case independence with respect to $y$ yields\n\n$$\n\\varphi_{tt} - \\varphi_{xx} - \\varphi_{zz} = 0 \\tag{17}\n$$\n\nPART II. TWO-DIMENSIONAL LINEAR PROBLEMS FOR $M_0$ NEAR ONE\n\nUnsteady State, $M_0 \\geq 1$\n\nIt was pointed out in the derivation of equation (12) that the linear equation for the velocity potential is not applicable to airfoil problems in either the subsonic or supersonic regimes for $M_0$ near one. The possibility still remains, however, of analyzing unsteady flows during the period in which the perturbation velocities remain small. As an example of such a problem, consider the case of a flat lifting surface at a small angle of attack $\\alpha$ starting from rest at a velocity $V_0$ near the speed of sound. The perturbation potential for such a motion is equivalent to the change in potential brought about by an abrupt change $\\alpha$ in angle of attack of an airfoil flying in a steady-state condition at velocity equal to $V_0$.", "timestamp": "2026-07-22T06:12:05.860239+00:00"}
{"citation_id": "19930082703", "source_url": "https://ntrs.nasa.gov/api/citations/19930082703/downloads/19930082703.pdf", "page_number": 14, "total_pages": 28, "image_filename": "19930082703_p14.jpg", "text": "12\nNACA TN 1983\n\noscillations such as that of figure 6 and reflecting the results of both\nreference 1 and of the present paper might be worded as follows:\n\nWhen a disturbance is produced by displacing the longi-\ntudinal control stick rearward 1/2 inch from trim for\n1/2 second and then returning to trim and holding the trim\nsetting, the following qualities shall be demonstrated:\n(1) The value of normal acceleration $g$ shall not increase by\nmore than 1/4g (total, $1\\frac{1}{4}g$) within 10 seconds from the start\nof the disturbance; and (2) during the subsequent nose-down\nmotion (with controls still fixed at trim), the value of\nacceleration shall not fall below 3/4g within 10 seconds,\nthe 10 seconds being measured from the time of initial\nreturn to 1 g.\n\nThis supplementary requirement primarily tends to insure that an\noscillation rather than a sudden divergence will occur and, as a rule,\ncomparatively simple instrumentation should suffice. Likewise, on the\nbasis of existing experience, the exact time or amount that the stick\nis held displaced should seldom be critical.\n\nWith any of these checks, a mechanical device providing adjustable\nstops for limiting the stick travel would be desirable to aid the pilot\nin obtaining rectangular control-displacement time histories. For\nreasons of safety, however, such a device must be designed so that the\npilot can instantly remove the stops, in event of difficulty, yet will\nnot unintentionally over-ride them.\n\nCONCLUSIONS\n\nThe indications of brief studies of forward-flight longitudinal\nflying qualities of several single-rotor helicopters may be summarized\nas follows:\n\n1. In relation to the pilot's satisfaction with the flying qualities,\nthe most important consideration is the prevention of prolonged stick-\nfixed divergence.\n\n2. When prolonged stick-fixed divergence is eliminated, additional\nimprovement is concluded to relate to the continuous development of\nnormal acceleration in contrast to a pause in the development of\nacceleration during the first second following abrupt control deflection.", "timestamp": "2026-07-22T06:12:05.890434+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 84, "total_pages": 96, "image_filename": "19930085880_p84.jpg", "text": "82\nNACA RM No. L9C03\n\nResistance, lb\nSpeed\n(fps)\n30\n25\n20\n15\n10\nWetted area, sq ft\n(c) $\\tau = 120^\\circ$.\nFigure 23.-- Continued.\nNACA", "timestamp": "2026-07-22T06:12:07.532166+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 32, "total_pages": 78, "image_filename": "19930082483_p32.jpg", "text": "30\nNACA TN No. 1807\n\non the use of corrected data. Some additional accuracy is therefore gained by use of operational data uncorrected to sea level in equation (53).\n\nThis modified procedure has been followed in the construction of figure 11 in which is plotted over-all turbine efficiency against percentage of full-power output with degree of admission as parameter. Included in the range of turbine operation at full admission is the point close to the design operating condition, which is a corrected rotor speed of 8650 rpm, a total-pressure ratio of 2.08, and an inlet total pressure of 45 inches of mercury absolute at the test operating temperature of 800° R (based on constant design Reynolds number index $p_1^{\\prime}/(T_1^{\\prime})^{1.1}$, reference 6). In order to obtain an indication of the turbine performance at other values of admission, the conditions of operation were arbitrarily selected to be maintained constant at this equivalent design point while estimates of performance were made at several reduced admissions. Because the maximum corrected power output in these experiments occurs at a pressure ratio of 2.4 rather than 2.0, the power at the equivalent design point for full admission represents approximately 83 percent of this value.\n\nComparison of Partial Admission with Other Methods\nof Power Reduction\n\nTurbine-inlet pressure, pressure ratio, and active nozzle arc are evaluated as power-control parameters in figure 12. Turbine over-all efficiency is plotted against the percentage of full-power output for a constant inlet total temperature of 800° R and a corrected rotor speed of 8650 rpm. Lines of constant inlet total pressure and total-pressure ratio enclose areas at both full and 120° admission within which turbine performance was investigated. In order to isolate the turbine performance from any interaction with a compressor, the driving fluid was assumed to be delivered to the turbine from an external source at a constant pressure of 45 inches of mercury absolute. Reduction of the inlet pressure to lower values is accomplished by throttling, the accompanying loss being charged to the turbine. This procedure accounts for the largest part of the severe loss in efficiency at the lower pressures. The projected distance of these areas on the abscissa gives an indication of the flexibility of power control.\n\nSuperimposed on the plot is the curve representing power regulation by means of active nozzle-arc control at a constant pressure ratio of 2.0, which is shown in figure 11. The partial-admission", "timestamp": "2026-07-22T06:12:08.159867+00:00"}
{"citation_id": "19930082546", "source_url": "https://ntrs.nasa.gov/api/citations/19930082546/downloads/19930082546.pdf", "page_number": 23, "total_pages": 65, "image_filename": "19930082546_p23.jpg", "text": "22\nNACA TN No. 1870\n\nSound Levels in Fuselage\n\nThe difference in pressure level of sound as it passes into an enclosure such as a fuselage is given by reference 3 as\n\n$$\n\\text{Attenuation in decibels} = 10 \\log_{10} \\left( 1 + \\frac{A_C}{T_C} \\right)\n$$\n\nwhere $A_C$ is the absorption coefficient in the enclosure and $T_C$ is the transmission coefficient of sound through the walls. The transmission is given by the square of the ratio of wall vibration amplitude to the amplitude of the external sound wave. (See appendix.) The lower the wall vibration for a given external excitation, the lower the transmission, and, hence, the greater the sound reduction. Such reduction is possible only if $A_C$ is greater than zero; that is, only if sound-absorbing material is present in the fuselage can the sound intensity inside be less than the intensity outside. It may also be noted from the equation for attenuation that even though $A_C$ be unity (its maximum value), the sound reduction will not be appreciable unless $T_C$ is quite small. In the interest of crew comfort, a nominal value of absorption and a low value of transmission are therefore necessary.\n\nThe designer may reduce sound pressures in the fuselage: (1) by moving the engines outboard to increase tip clearance, (2) by increasing the number of blades, (3) by choosing the optimum fuselage shape, (4) by increasing fuselage rigidity, mass, and damping, and (5) by applying sound-absorbing material. Each of these variables is most effective over a certain range of conditions.\n\nCONCLUSIONS\n\nFree-space oscillating-pressure measurements for static conditions near the propeller tips (tip Mach number range 0.45 to 1.00) for five different propellers indicate the following conclusions:\n\n1. Pressures measured on a line parallel to the propeller axis are increased as tip clearance is decreased; however, only the pressures in a region one-half radius ahead of the plane of rotation to one-half radius behind it are greatly increased.\n\n2. At a constant power the pressure amplitudes of the lower harmonics tend to decrease and the higher harmonics tend to increase with an increase in tip Mach number. The fundamental frequency of pressure produced by a four-blade propeller is essentially independent of tip Mach number in the useful tip Mach number range.", "timestamp": "2026-07-22T06:12:14.017475+00:00"}
{"citation_id": "19930082566", "source_url": "https://ntrs.nasa.gov/api/citations/19930082566/downloads/19930082566.pdf", "page_number": 22, "total_pages": 44, "image_filename": "19930082566_p22.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:12:24.416244+00:00"}
{"citation_id": "19930085572", "source_url": "https://ntrs.nasa.gov/api/citations/19930085572/downloads/19930085572.pdf", "page_number": 7, "total_pages": 17, "image_filename": "19930085572_p7.jpg", "text": "NACA RM No. E8L02\n5\n\nafter $7\\frac{1}{2}$ hours of operation. These small differences, however, are attributed to the normal reproducibility of test conditions and scatter of data for this type of investigation.\n\nLewis Flight Propulsion Laboratory,\nNational Advisory Committee for Aeronautics,\nCleveland, Ohio.\n\nREFERENCES\n\n1. Sanders, Newell D.: Performance Parameters for Jet-Propulsion Engines. NACA TN No. 1106, 1946.\n\n2. Gooding, Richard M., and Hopkins, Ralph L.: The Determination of Aromatics in Petroleum Distillates. Paper presented before Div. Petroleum Chem., Am. Chem. Soc. (Chicago, Ill.) Sept. 9-13, 1946, pp. 131-141.", "timestamp": "2026-07-22T06:12:25.410669+00:00"}
{"citation_id": "19930082592", "source_url": "https://ntrs.nasa.gov/api/citations/19930082592/downloads/19930082592.pdf", "page_number": 21, "total_pages": 50, "image_filename": "19930082592_p21.jpg", "text": "20\nNACA TN 1914\n\n<!-- Image (223, 246, 804, 635) -->\n\n(b) Temperature, 1785° F.\nFigure 4. - Continued. Effect of time, temperature, and tungsten content on oxidation penetration of titanium carbide - tungsten cermamals.", "timestamp": "2026-07-22T06:12:28.475000+00:00"}
{"citation_id": "19930085485", "source_url": "https://ntrs.nasa.gov/api/citations/19930085485/downloads/19930085485.pdf", "page_number": 14, "total_pages": 26, "image_filename": "19930085485_p14.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:12:33.558643+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 43, "total_pages": 114, "image_filename": "19930086061_p43.jpg", "text": "```markdown\nNACA RM L9J07\n\n$\\alpha=4.1^\\circ$\n$C_L=0.14$\n\n$\\alpha=8.1^\\circ$\n$C_L=0.32$\n\n$\\alpha=14.1^\\circ$\n$C_L=0.54$\n\n$\\alpha=24.1^\\circ$\n$C_L=0.85$\n\nP\n-2\n-1\n0\n1\n\nStation 4\n$\\frac{y}{b/2}, 0.500$\n\nStation 5\n$\\frac{y}{b/2}, 0.667$\n\nStation 6\n$\\frac{y}{b/2}, 0.833$\n\nStation 7\n$\\frac{y}{b/2}, 0.916$\n\nUpper\nLower\nTwo dimensional\n(calculated at\nequal $C_L$)\n\nx/c\n0 2 4 6 8 10\n\nx/c\n0 2 4 6 8 10\n\nx/c\n0 2 4 6 8 10\n\nx/c\n0 2 4 6 8 10\n\nNACA\n\n(b) Stations: 4,5,6,7.\n\nFigure 8.- Concluded.\n\n39\n```", "timestamp": "2026-07-22T06:12:38.813795+00:00"}
{"citation_id": "19930085847", "source_url": "https://ntrs.nasa.gov/api/citations/19930085847/downloads/19930085847.pdf", "page_number": 5, "total_pages": 32, "image_filename": "19930085847_p5.jpg", "text": "```markdown\nNACA RM A9D04 CONFIDENTIAL 3\n\nused in the Reynolds number calculations was obtained from radiosonde data. In addition, a 16-millimeter gunsight aiming-point camera was installed in the canopy to photograph tuft action on the test panel.\n\nFor the recording airspeed system a freely swiveling airspeed head was mounted on the end of an airspeed boom attached to the left wing tip and extending two chord lengths ahead of the wing leading edge, as shown in figures 1 and 2. The airspeed calibration error for the installation was almost negligible, the maximum correction to the measured Mach number throughout the test range being 0.01.\n\nThe profile-drag rake (shown in fig. 4) was mounted in line with the center line of the test section on the end of a cantilevered strut extending outward from the fuselage as shown in that figure. The rake contained 54 total head tubes and 6 static tubes with the tube openings located 14.2-percent chord (11-1/4 in.) aft of the wing trailing edge.\n\nTESTS AND RESULTS\n\nThe pressure-distribution and wake-survey measurements were obtained in steady dives of substantially linear flight path. The wake surveys were conducted at a pressure altitude of approximately 15,000 feet over a Mach number range from 0.70 to 0.78 (airplane lift coefficient varied from 0.12 to 0.08). (Maximum speed was limited by rake vibration and expansion of the wake, which exceeded the limited extent of the rake at Mach numbers above 0.78.) The pressure-distribution tests were conducted at an approximate pressure altitude of 30,000 feet over a Mach number range from 0.70 to 0.83 (airplane lift coefficient varied from 0.21 to 0.12). The Reynolds number-Mach number relation for both sets of tests is shown in figure 5. As can be seen from the figure, the Reynolds number range for the higher altitude was from $13 \\times 10^6$ to $16 \\times 10^6$ and for the lower altitude was from $20 \\times 10^6$ to $23 \\times 10^6$ based on the test section mean chord.\n\nA study of the flow conditions on the test panel upper surface was made with the following results:\n\n1. From the boundary-layer surveys, transition was found to occur at about 20-percent chord due to surface waviness.\n\n2. Inspection of the pressure distributions indicated that the shock location varied from about 52- to 63-percent chord. Hence the shock was always forward of the slot by at least 7-percent chord.\n\n3. The tuft study indicated that a pronounced cross flow over the\n\nCONFIDENTIAL\n```", "timestamp": "2026-07-22T06:12:39.961728+00:00"}
{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 9, "total_pages": 72, "image_filename": "19930085491_p9.jpg", "text": "8 CONFIDENTIAL NACA RM No. A8J04\n\nconstruction were considered. Since at the present time the minimum depth at the root is approximately one-fifteenth the spar distance from the root to the centroid of panel area, the ratio of 1:13.6 obtained with the wing of WF-63 indicates a slightly greater wing thickness was employed than was required by this structural criterion.\n\nTo obtain the four additional leading-edge sweep settings, the half-wings were rotated about the midpoint of the root chord. Thus, increasing the sweep angle resulted in a decrease in streamwise thickness-chord ratio, a decrease in aspect ratio, and a rearward movement of the position of maximum streamwise section thickness. (The structural criterion was not violated in any case.) Table I shows the variation of the geometric parameters affected by rotating the wing panels.\n\nThe fuselage shape used has been determined by Haack, reference 7, to have the minimum pressure drag for a given length and volume assuming closure at the tail as is shown by the broken lines in figure 4. The model fuselage shape, however, had a base to permit installation on the balance sting, the area of the base being large enough to shield the sting shroud. In order to obtain a variation of the incidence angle of the fuselage on the sting, the model fuselage was constructed in two parts as is shown in figure 2. The fuselage used for obtaining force data had $4^\\circ$ incidence to the sting axis so that with the balance beam travel of $\\pm 5^\\circ$ indicated in figure 1, the total angle-of-attack range was from $-1^\\circ$ to $9^\\circ$. A photograph of the model mounted in the tunnel prior to a force test is shown in figure 3. Plan-form schlieren and liquid-film photographs were obtained during special tests with the model rotated $90^\\circ$ from the position shown in figure 3. The balance beam was set at zero angle of attack for these tests and the desired lift coefficients were obtained by selection of the afterbody with the required incidence angle.\n\nTo obtain a fuselage for the fuselage-alone force tests, the fuselage wing slots were filled and the metal formed in a manner that gave circular sections normal to the longitudinal axis.\n\nTest Methods\n\nThe methods used for determining the aerodynamic forces on the model were the same as those of reference 5. Measurements were made of lift, drag, and pitching moment.\n\nThe liquid-film technique employed in reference 5 was used to determine the nature of the boundary-layer flow on the model surfaces\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:12:40.224617+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 45, "total_pages": 98, "image_filename": "19930086073_p45.jpg", "text": "NACA RM A9H04\n\n1.4\n1.2\n1.0\n.8\n.6\n.4\n.2\n0\n-.2\n-.06 -.04 -.02 0 .02\nRolling-moment\ncoefficient, $C_l$\n\n1.4\n1.2\n1.0\n.8\n.6\n.4\n.2\n0\n-.2\n-.08 -.06 -.04 -.02 0 .02\nYawing-moment\ncoefficient, $C_n$\n\n1.4\n1.2\n1.0\n.8\n.6\n.4\n.2\n0\n-.2\n-.20 -.16 -.12 -.08 -.04 0 .04\nSide-force coefficient,\n$C_Y$\n\n$\\beta$, deg\n$\\circ$ 0.0\n$\\square$ 6.0\n$\\diamond$ 12.0\n$\\triangle$ 15.9\n\n(d) $C_L$ vs $C_l$, $C_n$ and $C_Y$.\n\nFigure 8.— Concluded.\n\nNACA\n\n43", "timestamp": "2026-07-22T06:12:46.656046+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 8, "total_pages": 149, "image_filename": "19930083192_p8.jpg", "text": "4\nNACA TN 1976\n\ndue to the change in angle of attack. If the particular gust angles of $90^\\circ$, $0^\\circ$, and $180^\\circ$ are considered, then the expression reduces to\n\n$$\n\\frac{L}{W} = 1 + \\Delta n = 1 + \\frac{dC_L}{d\\alpha} \\frac{S}{W} \\frac{\\rho}{2} UV\n$$\n\nfor $\\beta = 90^\\circ$ and\n\n$$\n\\frac{L}{W} = 1 \\pm \\Delta n = 1 \\pm \\frac{2U}{V}\n$$\n\nfor $\\beta = 0^\\circ$ and $180^\\circ$.\n\nThe so-called sharp-edge-gust formula is the equation for $\\beta = 90^\\circ$ for sea-level density and is generally written\n\n$$\n\\Delta n = \\frac{\\rho_0 U_e V_e \\frac{dC_L}{d\\alpha}}{\\frac{W}{2S}}\n$$\n\nThe effect of gust direction on the acceleration increment is shown in figure 1 for airplanes with wing loadings from 15 to 60 pounds per square foot flying at 200 miles per hour through a gust velocity of 15 feet per second. The ratio of the acceleration increment for given values of wing loading to the maximum value for each wing loading is plotted against the gust angle $\\beta$. Figure 1 indicates, as do references 2 and 3, that the acceleration is a maximum for angles close to $90^\\circ$; figure 1 also indicates that the vertical or near-vertical gust is 4 to 15 times as effective in producing acceleration as the horizontal gust. Therefore, in reducing acceleration data, the assumption is usually made that the significant accelerations caused by gusts result from vertical gusts, that is, $\\beta = 90^\\circ$.\n\nThe sharp-edge-gust formula, sometimes used with a correction factor, requires fairly rigid definition of the quantities to be substituted. The equation is used for general research studies where masses of acceleration data are to be reduced and compared for evaluation of past airplane gust-load experience and for the prediction of load experience. For research studies, the actual weight of the airplane is used to determine the wing loading since the data from different sources", "timestamp": "2026-07-22T06:12:48.707053+00:00"}
{"citation_id": "19930085548", "source_url": "https://ntrs.nasa.gov/api/citations/19930085548/downloads/19930085548.pdf", "page_number": 7, "total_pages": 46, "image_filename": "19930085548_p7.jpg", "text": "6\nNACA RM No. E8L30\n\nthe charge at a constant pressure. (See appendix for method of\ncalculation of M.) Although the flow processes do not actually\noccur in this manner, this assumption permits a convenient method\nor index for comparing the capacities of various valve systems,\ninasmuch as it takes into account the valve areas, the rates of\nopening, the flow coefficients, the total opening periods, and the\ncylinder dimensions. With this method of comparison, the lower the\nvalue of M, the greater the capacity of the valve under the oper-\nating conditions considered. The comparative values of M for the\nported cylinder and the cylinder of a conventional aircraft engine\nat piston speeds of 1350 and 2400 feet per minute and arbitrarily\nchosen inlet and exhaust sonic velocities of 1100 and 2500 feet per\nsecond, respectively, are:\n\n| | Ported cylinder | Conventional aircraft-engine cylinder |\n| :--- | :--- | :--- |\n| Engine speed at piston speed of 1350 feet per minute | 1800 rpm | 1350 rpm |\n| Inlet M . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 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. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 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. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .", "timestamp": "2026-07-22T06:12:52.019484+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 85, "total_pages": 96, "image_filename": "19930085880_p85.jpg", "text": "NACA RM No. L9C03\n83\n\nResistance, lb\nSpeed (fps)\nWetted area, sq ft\n(d) $\\tau = 16^\\circ$.\nFigure 23.- Continued.\nNACA", "timestamp": "2026-07-22T06:12:55.834188+00:00"}
{"citation_id": "19930082703", "source_url": "https://ntrs.nasa.gov/api/citations/19930082703/downloads/19930082703.pdf", "page_number": 15, "total_pages": 28, "image_filename": "19930082703_p15.jpg", "text": "NACA TN 1983\n\n3. The normal-acceleration characteristics appreciated by the pilot can be theoretically predicted.\n\nLangley Aeronautical Laboratory\nNational Advisory Committee for Aeronautics\nLangley Air Force Base, Va. September 8, 1949\n\nREFERENCES\n\n1. Reeder, John P., and Gustafson, F. B.: On the Flying Qualities of Helicopters. NACA TN 1799, 1949.\n\n2. Phillips, William H.: Appreciation and Prediction of Flying Qualities. NACA TN 1670, 1948.\n\n3. Gilruth, R. R.: Requirements for Satisfactory Flying Qualities of Airplanes. NACA Rep. 755, 1943.\n\n4. Hunter, Paul A., and Reeder, John P.: Flight Measurements to Determine Effect of a Spring-Loaded Tab on Longitudinal Stability of an Airplane. NACA ARR L5I20, 1946.\n\n5. Wheatley, John B.: An Aerodynamic Analysis of the Autogiro Rotor with a Comparison between Calculated and Experimental Results. NACA Rep. 487, 1934.\n\n6. Bailey, F. J., Jr.: A Simplified Theoretical Method of Determining the Characteristics of a Lifting Rotor in Forward Flight. NACA Rep. 716, 1941.\n\n7. Gessow, Alfred, and Amer, Kenneth B.: An Introduction to the Physical Aspects of Helicopter Stability. NACA TN 1982, 1949.", "timestamp": "2026-07-22T06:12:56.202118+00:00"}
{"citation_id": "19930082592", "source_url": "https://ntrs.nasa.gov/api/citations/19930082592/downloads/19930082592.pdf", "page_number": 22, "total_pages": 50, "image_filename": "19930082592_p22.jpg", "text": "```markdown\nNACA TN 1914\n21\n\n<!-- Image (182, 199, 746, 733) -->\n\n(c) Temperature, 2000° F.\nFigure 4. - Continued. Effect of time, temperature, and\ntungsten content on oxidation penetration of titanium\ncarbide - tungsten cermamals.\n```", "timestamp": "2026-07-22T06:12:59.259890+00:00"}
{"citation_id": "19930085485", "source_url": "https://ntrs.nasa.gov/api/citations/19930085485/downloads/19930085485.pdf", "page_number": 15, "total_pages": 26, "image_filename": "19930085485_p15.jpg", "text": "Balance\ncenter line\n42.7°\nCONFIDENTIAL\nAll dimensions\nin inches\n14\n40.0\n3.7°\n2 1/4\nTunnel floor\nNACA\nFigure 4.- Schematic sketch of relative position of model, balance, and\ntransonic bump as mounted in the Langley high-speed 7- by 10-foot\ntunnel.\nCONFIDENTIAL\nNACA RM No. L8E02\n13", "timestamp": "2026-07-22T06:13:04.644834+00:00"}
{"citation_id": "19930082613", "source_url": "https://ntrs.nasa.gov/api/citations/19930082613/downloads/19930082613.pdf", "page_number": 19, "total_pages": 46, "image_filename": "19930082613_p19.jpg", "text": "18\nNACA TN 1938\n\nFront\nRear\n\nCombustion-chamber\nliner\n\nCombustion-chamber numbers\nviewed from rear of engine\n\n[Figure: Diagram of a turbojet engine showing a typical type-B combustion-chamber liner installation. The diagram includes a cross-section of the engine and a circular arrangement of numbered combustion chambers.]\n\nFigure 2. - Turbojet engine showing typical type-B combustion-chamber liner installation.\n\nNACA", "timestamp": "2026-07-22T06:13:09.848242+00:00"}
{"citation_id": "19930082918", "source_url": "https://ntrs.nasa.gov/api/citations/19930082918/downloads/19930082918.pdf", "page_number": 11, "total_pages": 62, "image_filename": "19930082918_p11.jpg", "text": "10\nNACA TN 1940\n\nIn all tests the furnace was brought to the proper operating temperature before placing the specimen in it. The specimen was allowed to come to thermal equilibrium for an average period of $1\\frac{1}{2}$ hours during which time minor controller and temperature-uniformity adjustments were also made. At the expiration of the average period of $1\\frac{1}{2}$ hours, the specimen was loaded. It was felt that in this way modification of the known initial structure of the specimens in the creep testing equipment prior to loading was held to a minimum. Thus the short-time creep characteristics found, it is felt, truly represent the creep characteristics of the known initial structures without appreciable modification by time at the test temperature.\n\nTwo stresses were used in creep testing, 30,000 and 60,000 psi, and one temperature, 1200° F. The 30,000-psi stress was approximately the highest possible without excessive plastic deformation upon loading. The higher stress was used to determine how the stress level affected the conclusions concerning the effects of aging on creep resistance at 30,000 psi.\n\nThe creep tests were run for an average of 50 hours provided fracture had not occurred. These tests were restricted to 50 hours in order to obtain creep properties as characteristic as possible of the known initial structures and not the properties of the known initial structure plus modifications induced by time at the test temperature. At the end of 50 hours all the tests covered herein had reached the so-called second stage of creep, with a reasonably steady creep rate, or had fractured. The creep rates reported are either these second-stage rates at 50 hours or the minimum rates occurring before fractures. It is obvious then that complete evaluation of decreasing secondary rates was not carried out.\n\nRupture Testing\n\nRupture testing was carried out in three units. The tests under stresses above 60,000 psi were run in a hydraulic tensile machine equipped with a transversely split electric-resistance furnace. The load was held constant during the test to within $\\pm 1$ percent with the rate of initial loading of the specimens approximately 50,000 psi per minute and comparable with the rate of loading of the more conventional rupture tests. Temperature control and uniformity over the gage length were the same as for the creep tests. Specimens for these tests were obtained by longitudinal quartering of the original 7/8-inch-square bar stock and using only those corners of the original bar which were uniformly fine grained. Gage lengths $1\\frac{3}{8}$ inches long by 0.250 inch", "timestamp": "2026-07-22T06:13:10.303269+00:00"}