PDF/ocr_dataset_107.jsonl

{"citation_id": "19930091724", "source_url": "https://ntrs.nasa.gov/api/citations/19930091724/downloads/19930091724.pdf", "page_number": 3, "total_pages": 20, "image_filename": "19930091724_p3.jpg", "text": "REPORT No. 649\n\nTHE “PACK” METHOD FOR COMPRESSIVE TESTS  \nOF THIN SPECIMENS OF MATERIALS USED  \nIN THIN-WALL STRUCTURES\n\nBy C. S. AITCHISON and L. B. TUCKERMAN  \nNational Bureau of Standards\n\n114539—39", "timestamp": "2026-07-19T18:41:33.204862+00:00"}
{"citation_id": "19930093641", "source_url": "https://ntrs.nasa.gov/api/citations/19930093641/downloads/19930093641.pdf", "page_number": 28, "total_pages": 47, "image_filename": "19930093641_p28.jpg", "text": "N.A.C.A.\nFigs. 2, 2(a)\n\n294-7\n\n[Figure: A model of a 4-engine aircraft mounted on supports inside a large wind tunnel structure. The aircraft has conventional nacelles and external radiators.]\n\nFigure 2.- Installation of the 4-engine model in the full-scale wind tunnel: Conventional nacelles and external radiators for liquid-cooled engines.\n\n[Figure: Bottom view of the same 4-engine aircraft model installed in the wind tunnel, showing the underside of the wings and fuselage with engines and radiators visible.]\n\nFigure 2(a) Bottom view.- Installation of the 4-engine model in the full-scale wind tunnel: Conventional nacelles and external radiators for liquid-cooled engines.", "timestamp": "2026-07-19T18:41:34.458650+00:00"}
{"citation_id": "19930094504", "source_url": "https://ntrs.nasa.gov/api/citations/19930094504/downloads/19930094504.pdf", "page_number": 2, "total_pages": 16, "image_filename": "19930094504_p2.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-19T18:41:36.821121+00:00"}
{"citation_id": "19930091697", "source_url": "https://ntrs.nasa.gov/api/citations/19930091697/downloads/19930091697.pdf", "page_number": 13, "total_pages": 28, "image_filename": "19930091697_p13.jpg", "text": "152\n40\nA\nT.C.\n5°\nA.T.C.\n148\n30\nOctane rating\nA\nB.T.C.\n5°\nT.C.\n145\n18\nA\nB.T.C.\n5°\nT.C.\n5°\nA.T.C.\n\nFIGURE 9.—Enlargements of high-speed motion pictures showing effect of different octane ratings on combustion knock. Air-fuel ratio, 14; A, first evidence of knock; engine speed, 500 r. p. m.; one spark plug. Octane ratings, 40, 30, and 18.\n\nA PHOTOGRAPHIC STUDY OF COMBUSTION AND KNOCK IN A SPARK-IGNITION ENGINE\n9", "timestamp": "2026-07-19T18:41:39.376147+00:00"}
{"citation_id": "19930091655", "source_url": "https://ntrs.nasa.gov/api/citations/19930091655/downloads/19930091655.pdf", "page_number": 13, "total_pages": 22, "image_filename": "19930091655_p13.jpg", "text": "HEAT TRANSFER TO FUEL SPRAYS INJECTED INTO HEATED GASES 9\n\ning the coefficient to be positive as for certain other hydrocarbons (reference 25). Qualitative experience indicates, however, that the net effect is a temperature increase due to friction in the orifice and the conversion of the kinetic energy of the spray into heat. The and vaporization. Lee has shown (reference 20) that the degree of subdivision attainable with a hydraulic injection system under operating conditions approaches a practical limit. For the practical range of gas densities and injection pressures, however, it is impossible\n\n[Figure: Graph showing pressure drop vs. time for different injection numbers (1, 3, 5, 11, 23), with labeled curves A, B, C and values 380, 382, 384, 386, 388. Y-axis labeled “PRESSURE DROP, ATMOSPHERE”, X-axis labeled “TIME, SECOND”. Handwritten annotations include “0.88”, “228”, “448” near respective curves.]\n\nFIGURE 6.—Influence of fuel-vapor concentration, prior to injection, upon pressure drop. Diesel fuel; fuel weight per injection, 0.508 gram; gas density, 10.35 grams per liter; gas temperature, 250° C.\n\nchange is believed to be too small to be of any interest in the interpretation of the present results and will therefore be ignored.\n\nDISCUSSION\n\nOn the basis of diffusion and heat-transfer concepts the size of a droplet must influence its rate of heating to vary the distribution of droplet sizes without at the same time varying the rate of spray penetration. This concomitant variation prevents the isolation of any effect that can be associated solely with the distribution of droplet sizes. In the subsequent discussion it is well to bear in mind that the same condition should be true of certain other quantities that may represent an aggregation of variables.", "timestamp": "2026-07-19T18:41:43.086220+00:00"}
{"citation_id": "19930094568", "source_url": "https://ntrs.nasa.gov/api/citations/19930094568/downloads/19930094568.pdf", "page_number": 6, "total_pages": 28, "image_filename": "19930094568_p6.jpg", "text": "N.A.C.A. Technical Memorandum No. 848 5\n\n0.0545* (see table I), this assured three groups of tests (figs. 2 to 4, 14 to 16, and 17 to 19) which, in turn, were subdivided according to Froude number (speeds) into test series with constant weight loading but variable center-of-gravity position. Table I also indicates the test series, run with the 15- and 30-centimeter plates. To assure greater accuracy and at the same time afford a check on Sottorf's measurements, the tests were made - as far as the experimental set-up allowed - with the 30-centimeter plate. Because it was occasionally expedient to analyze the test data in relation to the speed rather than the Froude number (figs. 2-4, 14-16, and 17-19), and also because the speeds at equal Froude number are proportional to the roots of the plate widths, the speed was referred to one plate width. In the following it was referred and converted to the 30-centimeter plate (figs. 2-4, 14-16, and 17-19), the same as Sottorf employed in his tests.\n\nAs in Sottorf's case (figs. 1 and 22), our experiments covered:\n\n1. The produced wetted length $l$ of the pressure surface;\n2. The ensuing angle of attack $\\beta$ of the planing surface; and\n3. The drag $W_Z$ of the planing surface (push-rod force).\n\nThe position of the lift resultants (fig. 1 and table I, columns 6 and 15) and their components normal to the planing surface (= R) and in the plane of the planing surface (= $W_R$ = friction) were mathematically established from Gy and Gx, the plate weight and $W_Z$; from $W_R$ followed the coefficient of friction $c_f = \\frac{W_R}{\\frac{\\rho}{2} V^2 b l}$ (table I, columns 8 and 15).\n\n3. Results of Experiments (Airfoil Comparison)\n\nThe results are shown in figures 2 to 4, with $1/F^2$, the reciprocal value of the squared Froude number $F$ as\n\n*Corresponding to Sottorf's load ratings $C_B = 0.218$, 0.109, and 0.0545 (reference 3).", "timestamp": "2026-07-19T18:41:44.623654+00:00"}
{"citation_id": "19930091716", "source_url": "https://ntrs.nasa.gov/api/citations/19930091716/downloads/19930091716.pdf", "page_number": 23, "total_pages": 24, "image_filename": "19930091716_p23.jpg", "text": "16.14 Reserves\n\n[Figure: Diagram showing an airplane with coordinate axes X, Y, Z and angles φ, θ, ψ indicated by arrows. Caption: Positive directions of axes and angles (forces and moments) are shown by arrows]\n\n| Axis | | Force (parallel to axis) symbol | Moment about axis | | | Angle | | Velocities | |\n|---|---|---|---|---|---|---|---|---|---|\n| Designation | Symbol | | Designation | Symbol | Positive direction | Designation | Symbol | Linear (component along axis) | Angular |\n| Longitudinal…… | $X$ | $X$ | Rolling…… | $L$ | $Y \\longrightarrow Z$ | Roll…… | $\\phi$ | $p$ | |\n| Lateral……… | $Y$ | $Y$ | Pitching…… | $M$ | $Z \\longrightarrow X$ | Pitch…… | $\\theta$ | $q$ | |\n| Normal……… | $Z$ | $Z$ | Yawing…… | $N$ | $X \\longrightarrow Y$ | Yaw…… | $\\psi$ | $r$ | |\n\nAbsolute coefficients of moment  \n$C_l = \\frac{L}{q b S}$ (rolling)  \n$C_m = \\frac{M}{q c S}$ (pitching)  \n$C_n = \\frac{N}{q b S}$ (yawing)\n\nAngle of set of control surface relative to neutral position), $\\delta$. (Indicate surface by proper subscript.)\n\n---\n\n**4. PROPELLER SYMBOLS**\n\n$D$, Diameter  \n$p$, Geometric pitch  \n$p/D$, Pitch ratio  \n$V'$, Inflow velocity  \n$V_s$, Slipstream velocity  \n$T$, Thrust, absolute coefficient $C_T = \\frac{T}{\\rho n^2 D^4}$  \n$Q$, Torque, absolute coefficient $C_Q = \\frac{Q}{\\rho n^2 D^5}$  \n\n$P$, Power, absolute coefficient $C_P = \\frac{P}{\\rho n^3 D^5}$  \n$C_s$, Speed-power coefficient $\\frac{\\sqrt[5]{P}}{n^2}$  \n$\\eta$, Efficiency  \n$n$, Revolutions per second,  \n$\\Phi$, Effective helix angle $= \\tan^{-1} \\left( \\frac{V}{2 \\pi r n} \\right)$\n\n---\n\n**5. NUMERICAL RELATIONS**\n\n1 hp = 76.04 kg-m/s = 550 ft-lb./sec.  \n1 metric horsepower = 1.0132 hp.  \n1 m.p.h. = 0.4470 m.p.s.  \n1 m.p.s. = 2.2369 m.p.h.  \n\n1 lb. = 0.4536 kg.  \n1 kg = 2.2046 lb.  \n1 mi. = 1,609.35 m = 5,280 ft.  \n1 m = 3.2808 ft.", "timestamp": "2026-07-19T18:41:47.517472+00:00"}
{"citation_id": "19930094542", "source_url": "https://ntrs.nasa.gov/api/citations/19930094542/downloads/19930094542.pdf", "page_number": 87, "total_pages": 102, "image_filename": "19930094542_p87.jpg", "text": "N.A.C.A. Technical Memorandum No. 874\nFigs. 73, 74\n\n[Figure: Graph showing spanwise lift distribution for $\\kappa = -1^\\circ$. The graph plots $c_l$ against $y/b$. It contains multiple curves for different $\\lambda$ values and angles of attack $\\alpha$. A small diagram of an airfoil is shown. Labels include \"End plate\" on the left and right sides of the x-axis.]\n\nFigure 73. $\\kappa = -1^\\circ$.\n\n[Figure: Graph showing spanwise lift distribution for $\\kappa = -6^\\circ$. The graph plots $c_l$ against $y/b$. It contains multiple curves for different $\\lambda$ values and angles of attack $\\alpha$. A small diagram of an airfoil is shown. Labels include \"End plate\" on the left and right sides of the x-axis.]\n\nFigure 74. $\\kappa = -6^\\circ$.\nSpanwise lift distribution.", "timestamp": "2026-07-19T18:41:58.209870+00:00"}
{"citation_id": "19930093281", "source_url": "https://ntrs.nasa.gov/api/citations/19930093281/downloads/19930093281.pdf", "page_number": 6, "total_pages": 22, "image_filename": "19930093281_p6.jpg", "text": "4\n\n$\\eta_{o}$, net efficiency of the propeller-nacelle unit on the basic nose shape.\n\nS, propeller disk area.\n\n$P_{c}$, disk-loading coefficient or unit disk loading, $P/qSV$.\n\n$\\Delta C_{D}$, effective change in drag coefficient caused by the nose shape, $(\\eta_{o} - \\eta_{n}) P_{c} \\frac{S}{F}$.\n\n$1/\\sqrt[3]{P_{c}}$, propeller disk-loading coefficient, $\\sqrt[3]{\\frac{\\rho S}{2P}}$.\n\nAPPARATUS AND METHODS\n\nThe investigation was conducted in the N.A.C.A. 20-foot tunnel, which with its standard equipment is described in reference 4.\n\nFigure 1 presents a line drawing of the arrangements tested, with the designations of the noses and the nacelles used in each arrangement. Set-up 1 was used in reference 1; set-up 2 was used in the present investigation. The nose shapes that were used in reference 1 are shown in figures 2 to 4. The results presented in this paper were obtained with a pointed tail as shown in figure 1 and not with the tail pump shown in figure 3. Figures 5 to 7 show the nose shapes used in the tests for this report. The results in this report were obtained with all slots closed and faired.\n\nBecause the engine-nacelle installation for a tractor propeller is located in the slipstream of the propeller, it is necessary to study the nacelle with the propeller operating to obtain the possible secondary effects of the propeller. In order to include as many details as possible with a reasonable number of tests, three selected 10-foot-diameter propellers were tested over a range of blade angles from $20^{\\circ}$ to $55^{\\circ}$ at the 75-percent radius. Propeller B is Navy plan form 4893 with airfoil sections near the propeller hub; propeller C is Navy plan form 5368-9 with the conventional round blade shanks near the hub. Both propellers B and C have a constant pitch distribution when set at a blade angle of $15^{\\circ}$ at the 75-percent radius. Propeller $C_{x}$ is the same as propeller C except that it has a constant pitch distribution from the 50-percent radius to the tip when set at a blade angle of $35^{\\circ}$ at the 75-percent radius. Figure 8 shows one blade of each of the three 10-foot-diameter 3-blade propellers.", "timestamp": "2026-07-19T18:42:01.572753+00:00"}
{"citation_id": "19930094544", "source_url": "https://ntrs.nasa.gov/api/citations/19930094544/downloads/19930094544.pdf", "page_number": 35, "total_pages": 43, "image_filename": "19930094544_p35.jpg", "text": "N.A.C.A. Technical Memorandum No. 872\n\nFigs. 19,20,21\n\n[Figure: Figure 19.-LZ127-Framework during assembly, showing a view of rings. The rings are suspended from the roof trusses during assembly.]\n\n[Figure: Figure 20.- LZ127-Main ring on the floor. The rings are completely finished on the floor and are erected by the aid of stiff assembly frames.]\n\n[Figure: Figure 21.- LZ127 - Partial view of the framework showing the wire-braced main rings with the truss work, and the two unbraced intermediate auxiliary rings.]", "timestamp": "2026-07-19T18:42:18.988277+00:00"}
{"citation_id": "19930091748", "source_url": "https://ntrs.nasa.gov/api/citations/19930091748/downloads/19930091748.pdf", "page_number": 5, "total_pages": 12, "image_filename": "19930091748_p5.jpg", "text": "REPORT No. 673\n\nEXPERIMENTAL VERIFICATION OF THE THEORY OF OSCILLATING AIRFOILS\n\nBy ABE SILVERSTEIN AND UPSHUR T. JOYNER\n\nSUMMARY\n\nMeasurements have been made of the lift on an airfoil in pitching oscillation with a continuous-recording, instantaneous-force balance. The experimental values for the phase difference between the angle of attack and the lift are shown to be in close agreement with the theory.\n\nINTRODUCTION\n\nThe theory for the lift of infinite-span airfoils oscillating in pitch with small amplitude in a uniform stream of perfect fluid has been exhaustively studied and now provides a basis for the aerodynamic analysis of the flutter problem. Wagner’s theory (reference 1) for calculating the lift on an airfoil in nonuniform motion has been followed by those of Küssner (reference 2), Glauert (reference 3), and Theodorsen (reference 4); Garrick has indicated (reference 5) that the several theories are in agreement. Jones has given certain approximations (reference 6) to account for the effect of finite span.\n\nAt the suggestion of Theodorsen, tests have been made to obtain experimental data for a direct comparison of the measured lift on an oscillating airfoil with that predicted by the theory. The accuracy of the theory has been essentially substantiated in a less direct manner by the agreement of experimental and theoretically predicted flutter phenomena.\n\nIn a comparison of an oscillating airfoil with one in uniform motion, the theory indicates that the principal effect of the oscillation is to change the angle of attack at which a given lift occurs; for example, zero lift on an oscillating symmetrical airfoil does not occur at zero angle of attack. The phase difference between the lift and the angle of attack depends on the location of the axis along the airfoil chord and on a nondimensional parameter describing the wave length of the oscillating vortex sheet in the airfoil wake. For an infinite frequency of oscillation and a forward location of the axis, the lift would lead the angular displacement by $180^\\circ$. At finite frequencies, the countervorticity of the oscillating vortex sheet produces, in general, a lag that opposes the inertia effect causing the leading force. At low frequencies, the lag predominates; and, at zero frequency (steady motion), the lag again disappears.\n\nIt was planned to verify the theory by measurements of the phase difference between the lift and the angle of attack for an airfoil in rotational oscillation at various frequencies and air speeds. For these measurements, an instantaneous-force balance was designed with which the lift and the angle of attack of an oscillating airfoil could be continuously recorded. The measurements were made in a 2- by 3-foot tunnel on a symmetrical airfoil of about 5-inch chord and 18 percent thickness. The axis of rotation was located at the quarter-chord point of the airfoil. Measurements were taken for values of frequencies and air speeds that covered the useful flutter range.\n\nSYMBOLS\n\n$\\alpha$, angle of attack ($\\alpha = \\alpha_0 \\sin pt$).\n\n$\\alpha_0$, amplitude of oscillation.\n\n$b$, half chord of airfoil.\n\n$v$, air speed at infinity.\n\n$p$, $2\\pi$ times the frequency of oscillations.\n\n$k$, reduced frequency ($pb/v$); wave length in vortex sheet is $2\\pi b/k$.\n\n$a$, coordinate of axis of oscillation. (See reference 4.)\n\n$L$, lift force on airfoil.\n\n$t$, time.\n\n$F$ and $G$, circulation functions. (See reference 4.)\n\n$\\delta$, phase difference between angle of attack and lift for oscillating airfoil. Positive values indicate a leading force.\n\n$\\theta$, phase difference due to natural frequency of recording instrument.\n\n$n$, damping constant of airfoil and balance.\n\n$r$, $2\\pi$ times natural frequency of vibration of airfoil and balance.\n\n$A$, aspect ratio.", "timestamp": "2026-07-19T18:42:20.394495+00:00"}
{"citation_id": "19930094549", "source_url": "https://ntrs.nasa.gov/api/citations/19930094549/downloads/19930094549.pdf", "page_number": 25, "total_pages": 76, "image_filename": "19930094549_p25.jpg", "text": "N.A.C.A. Technical Memorandum No. 867 23\n\nwill not diminish or complicate the action. Numerous works based on more probable assumptions have been published. In a remarkable paper published many years ago, E. B. Wilson assumed as the law for the set-up of the gust, the following expression:\n\n$$\nw = J \\left(1 - e^{-rt}\\right)\n$$\n\nOther authors have assumed a sinusoidal law which is generally arbitrary. The latter assumption leads immediately to resonance phenomena if the period of the airplane is the same as that of the gust. Fisher, Bryan, and Jones have studied the effect of sudden and gradual gusts by methods which have become more and more accurate. Kussner, in a very well-known work of his, takes into effect the elasticity of the wings. Since our object is to compare the effect of different disturbances, we have not gone into the subject in more detail.\n\nEFFECT OF THE ACTION OF THE CONTROLS\n\nThe method outlined above enables us in certain particular cases to determine how an airplane responds to a manipulation of the controls. Let us examine figure 13, which gives $C_M$ as a function of $i$ and $\\beta$. Assume the airplane to be flying level at the angle of attack $i_1$, with the control deflected at angle $\\beta_1$, this condition corresponding to a velocity $V_1$. If the pilot gives the control a deflection $\\beta_2$, the point A passes to B. At B the moment M is no longer zero and the point will be displaced from B to C along the curve $\\beta_2$. The airplane will be in rotational equilibrium about the center of gravity only at the angle of attack corresponding to point C. Let $i_2$ be this angle of attack, to which will correspond a speed $V_2$ different from $V_1$. We shall assume that the maneuver is made without altering the throttle and that the velocity $V_1$ is greater than the velocity corresponding to minimum power. Under these conditions the useful power sufficient for making the airplane fly at the velocity $V_1$ will not be sufficient to make the airplane fly at velocity $V_2$. The path of the airplane will not be able to remain horizontal but will incline by the angle $\\xi$. The axis of the airplane will be lowered by $\\xi + (i_1 - i_2)$. At the instant when the pilot in deflect-", "timestamp": "2026-07-19T18:42:24.652834+00:00"}
{"citation_id": "19930094504", "source_url": "https://ntrs.nasa.gov/api/citations/19930094504/downloads/19930094504.pdf", "page_number": 3, "total_pages": 16, "image_filename": "19930094504_p3.jpg", "text": "NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nTECHNICAL MEMORANDUM NO. 912\n\nINCREASE OF THE SPECIFIC LOAD UNDER TENSION, COMPRESSION, AND BUCKLING OF WELDED STEEL TUBES IN AIRPLANE CONSTRUCTION BY SUITABLE TREATMENT OF STRUCTURAL STEEL AND BY PROPER DESIGN*\n\nBy J. Müller\n\nAlthough recently the light-metal stressed-skin construction has largely replaced the welded-steel-tube framework construction for airplane structures, it will nevertheless not be possible entirely to dispense with the steel tube as a tension and compression element. In the landing gear, engine mounting and also as surface supporting strut the steel tube, on account of the simplicity of the welded joints it permits and its outstanding strength characteristics for airplanes of all types, will continue as before to maintain its place.\n\nIn this report some considerations and test results are presented that may lead to higher tension, compression, and buckling stresses than is possible with the welded-steel-tube struts of the usual familiar construction. The new construction method indicated, which makes possible a considerably better material utilization and hence a saving in weight, has been tested in a number of Focke-Wulf types also in series production and has fully justified itself.\n\nIn any structural design problem, the lengths of the struts and the loads to be taken by them are generally given. An airplane framework structure must, on account of the various flight and landing conditions encountered, often be able to take up a certain tensile and also a definite pressure loading. Often, too, only one type of loading occurs.\n\nAs is known, the properties of the material that determine the cross-sectional area and hence the weight of the\n\n*\"Erhöhung der spezifischen Belastbarkeit bei Zug, Druck und Knickung von eingeschweißten Stahlrohr-Fachwerkstreben im Flugzeugbau durch Massnahmen werkstofftechnischer und konstruktiver Art.\" Luftfahrtforschung, vol. 16, no. 1, Jan. 10, 1939, pp. 14-17.", "timestamp": "2026-07-19T18:42:27.267703+00:00"}
{"citation_id": "19930094498", "source_url": "https://ntrs.nasa.gov/api/citations/19930094498/downloads/19930094498.pdf", "page_number": 4, "total_pages": 37, "image_filename": "19930094498_p4.jpg", "text": "N.A.C.A. Technical Memorandum No. 918\n\n4. Pneumatic rubber tires on the running wheels of the towing carriage, upon which the carriage ran smoothly in spite of the slight roughness of the surface of the rails and which gave greater adhesion than steel tires, thus making it possible to accelerate and decelerate at very high rates.\n\nIn general, each of the novel features incorporated into the original design has worked well and, to compensate for a few difficulties, advantages have appeared that were not foreseen when the original ideas were proposed. The use of the H boom rails and the rubber tires was proposed by the writer as a method of making a considerable saving in the cost of the tank. It was not until the detail design was begun that the potential effect on the starting and stopping of the carriage, because of the greater coefficient of friction on the rails, was perceived and advantage taken of it.\n\nThe high maximum speed of the towing carriage in combination with the rubber tires provided the anticipated ability to attain speed quickly and thus to use the length of the tank to the full in normal running. The length of the tank was found to be sufficient for a test run of about 10 seconds at 60 miles per hour but, at speeds under 30 miles per hour, from 2 to 6 test points could be obtained during a single run of the length of the tank, the larger number naturally corresponding to the lowest speeds used. This method of operation considerably increased the amount of work that could be done in a given time.\n\nGood as were the results obtained with the original N.A.C.A. tank, the rapid development of seaplanes and the increasing work required of the tank soon made it plain that even better performance would be required, and in 1936 serious consideration was given to proposals to enlarge it.\n\nThe primary reasons for enlarging the tank were to increase the amount of work that could be done in a given time and to be able to tow larger models at higher speeds. As has been stated, the length of the tank was found to be sufficient to provide a run of about 10 seconds at 88 feet per second. It was also found, however, that generally only one point could be gotten in a run if the speed exceeded 40 feet per second. Inasmuch as a good many tests with large models required test runs at speeds up to 50 feet per second and, occasionally, to 60 feet per second, many of the runs gave but one point. It was clear that, if the tank were 500 feet longer, at least one additional point per run could be obtained at the higher speeds, and more at the lower ones, and the time lost in accelerating and braking would become a smaller part of the total time of the run.", "timestamp": "2026-07-19T18:42:30.293590+00:00"}
{"citation_id": "19930094493", "source_url": "https://ntrs.nasa.gov/api/citations/19930094493/downloads/19930094493.pdf", "page_number": 4, "total_pages": 31, "image_filename": "19930094493_p4.jpg", "text": "N.A.C.A. Technical Memorandum No. 923 3\n\nNow the resistance of the wire depends on its temperature, according to the following relation:\n\n$$\nr = \\varphi(T_W)\n\\tag{3}\n$$\n\nwhere $T_W$ is the absolute temperature. Equation (2) therefore becomes\n\n$$\nU = 0.86 \\, \\psi(i^2, T_W)\n\\tag{4}\n$$\n\nThe wire is heated by the supplied energy, $U$. The variable velocity results in a change in the wire temperature and hence in the wire resistance. The latter is measured by comparison with a precision resistance and the current by means of an oscillograph loop.\n\nThe heat transmitted to the air stream (reference 4), neglecting small flow velocities, is given by the following equation:\n\n$$\n\\frac{\\alpha \\, d}{\\lambda_m} = f \\left[ \\frac{d \\, \\rho_m \\, w}{\\eta_m} \\right]\n\\tag{5}\n$$\n\nIn order to evaluate the above relation, there is required a knowledge of the temperature of the surrounding air. The latter is also measured with the hot wire, which is now employed as a resistance thermometer.\n\nIn equation (5):\n\n- $d$, is the diameter of the wire (m);\n- $\\alpha$, the mean heat-transfer coefficient for the entire wire surface (kcal/m² h °C.);\n- $\\lambda_m$, the heat conductivity of the medium (kcal/mh °C.);\n- $\\eta_m$, the viscosity coefficient of the medium (kgs/m²);\n- $\\rho_m$, the density of the medium (kgs²/m⁴);\n- $w$, the velocity of the medium at some distance from the wire surface (m/s).\n\nIf $T_W$ is the absolute temperature of the wire surface and $T_0$ the absolute temperature of the surrounding medium at some distance from the wire, the mean values in equation (5) are defined as follows:", "timestamp": "2026-07-19T18:42:31.175769+00:00"}
{"citation_id": "19930091724", "source_url": "https://ntrs.nasa.gov/api/citations/19930091724/downloads/19930091724.pdf", "page_number": 4, "total_pages": 20, "image_filename": "19930091724_p4.jpg", "text": "# NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nHEADQUARTERS, NAVY BUILDING, WASHINGTON, D. C.\nLABORATORIES, LANGLEY FIELD, VA.\n\nCreated by act of Congress approved March 3, 1915, for the supervision and direction of the scientific study of the problems of flight (U. S. Code, Title 50, Sec. 151). Its membership was increased to 15 by act approved March 2, 1929. The members are appointed by the President, and serve as such without compensation.\n\nJOSEPH S. AMES, Ph. D., *Chairman*,\nBaltimore, Md.\n\nVANNEVAR BUSH, Sc. D., *Vice Chairman*.\nWashington, D. C.\n\nCHARLES G. ABBOT, Sc. D.,\nSecretary, Smithsonian Institution.\n\nHENRY H. ARNOLD, Major General, United States Army,\nChief of Air Corps, War Department.\n\nGEORGE H. BRETT, Brigadier General, United States Army,\nChief Matériel Division, Air Corps, Wright Field, Dayton,\nOhio.\n\nLYMAN J. BRIGGS, Ph. D.,\nDirector, National Bureau of Standards.\n\nARTHUR B. COOK, Rear Admiral, United States Navy,\nChief, Bureau of Aeronautics, Navy Department.\n\nCLINTON M. HESTER, A. B., LL. B.,\nAdministrator, Civil Aeronautics Authority.\n\nJEROME C. HUNSAKER, Sc. D.,\nCambridge, Massachusetts.\n\nSYDNEY M. KRAUS, Captain, United States Navy,\nBureau of Aeronautics, Navy Department.\n\nCHARLES A. LINDBERGH, LL. D.,\nNew York City.\n\nEDWARD J. NOBLE, A. B.,\nChairman, Civil Aeronautics Authority.\n\nFRANCIS W. REICHELDERFER, A. B.,\nChief, United States Weather Bureau.\n\nEDWARD P. WARNER, Sc. D.,\nGreenwich, Conn.\n\nORVILLE WRIGHT, Sc. D.,\nDayton, Ohio.\n\nGEORGE W. LEWIS, *Director of Aeronautical Research*\n\nJOHN F. VICTORY, *Secretary*\n\nHENRY J. E. REID, *Engineer-in-Charge, Langley Memorial Aeronautical Laboratory, Langley Field, Va.*\n\nJOHN J. IDE, *Technical Assistant in Europe, Paris, France*\n\n## TECHNICAL COMMITTEES\n\n| | |\n| :--- | :--- |\n| AERODYNAMICS | AIRCRAFT STRUCTURES |\n| POWER PLANTS FOR AIRCRAFT | AIRCRAFT ACCIDENTS |\n| AIRCRAFT MATERIALS | INVENTIONS AND DESIGNS |\n\n*Coordination of Research Needs of Military and Civil Aviation*\n\n*Preparation of Research Programs*\n\n*Allocation of Problems*\n\n*Prevention of Duplication*\n\n*Consideration of Inventions*\n\n## LANGLEY MEMORIAL AERONAUTICAL LABORATORY\nLANGLEY FIELD, VA.\n\nUnified conduct, for all agencies, of scientific research on the fundamental problems of flight.\n\n## OFFICE OF AERONAUTICAL INTELLIGENCE\nWASHINGTON, D. C.\n\nCollection, classification, compilation, and dissemination of scientific and technical information on aeronautics.\n\n3-15-39", "timestamp": "2026-07-19T18:42:31.693158+00:00"}
{"citation_id": "19930091731", "source_url": "https://ntrs.nasa.gov/api/citations/19930091731/downloads/19930091731.pdf", "page_number": 2, "total_pages": 32, "image_filename": "19930091731_p2.jpg", "text": "# AERONAUTIC SYMBOLS\n\n## 1. FUNDAMENTAL AND DERIVED UNITS\n\n| | Symbol | Metric | | English | |\n| :--- | :---: | :--- | :--- | :--- | :--- |\n| | | Unit | Abbreviation | Unit | Abbreviation |\n| Length<br>Time<br>Force | $l$<br>$t$<br>$F$ | meter<br>second<br>weight of 1 kilogram | m<br>s<br>kg | foot (or mile)<br>second (or hour)<br>weight of 1 pound | ft. (or mi.)<br>sec. (or hr.)<br>lb. |\n| Power<br>Speed | $P$<br>$V$ | horsepower (metric)<br>kilometers per hour<br>meters per second | k.p.h.<br>m.p.s. | horsepower<br>miles per hour<br>feet per second | hp.<br>m.p.h.<br>f.p.s. |\n\n## 2. GENERAL SYMBOLS\n\n$W$, Weight=$mg$\n$g$, Standard acceleration of gravity=9.80665 m/s$^2$ or 32.1740 ft./sec.$^2$\n$m$, Mass=$\\frac{W}{g}$\n$I$, Moment of inertia=$mk^2$. (Indicate axis of radius of gyration $k$ by proper subscript.)\n$\\mu$, Coefficient of viscosity\n$\\nu$, Kinematic viscosity\n$\\rho$, Density (mass per unit volume)\nStandard density of dry air, 0.12497 kg-m$^{-4}$-s$^2$ at 15$^\\circ$ C. and 760 mm; or 0.002378 lb.-ft.$^{-4}$ sec.$^2$\nSpecific weight of \"standard\" air, 1.2255 kg/m$^3$ or 0.07651 lb./cu. ft.\n\n## 3. AERODYNAMIC SYMBOLS\n\n$S$, Area\n$S_w$, Area of wing\n$G$, Gap\n$b$, Span\n$c$, Chord\n$b^2$, Aspect ratio\n$S'$\n$V$, True air speed\n$q$, Dynamic pressure=$\\frac{1}{2}\\rho V^2$\n$L$, Lift, absolute coefficient $C_L=\\frac{L}{qS}$\n$D$, Drag, absolute coefficient $C_D=\\frac{D}{qS}$\n$D_0$, Profile drag, absolute coefficient $C_{D_0}=\\frac{D_0}{qS}$\n$D_i$, Induced drag, absolute coefficient $C_{D_i}=\\frac{D_i}{qS}$\n$D_p$, Parasite drag, absolute coefficient $C_{D_p}=\\frac{D_p}{qS}$\n$C$, Cross-wind force, absolute coefficient $C_C=\\frac{C}{qS}$\n$R$, Resultant force\n$i_w$, Angle of setting of wings (relative to thrust line)\n$i_t$, Angle of stabilizer setting (relative to thrust line)\n$Q$, Resultant moment\n$\\Omega$, Resultant angular velocity\n$\\rho \\frac{Vl}{\\mu}$, Reynolds Number, where $l$ is a linear dimension (e.g., for a model airfoil 3 in. chord, 100 m.p.h. normal pressure at 15$^\\circ$ C., the corresponding number is 234,000; or for a model of 10 cm chord, 40 m.p.s., the corresponding number is 274,000)\n$C_p$, Center-of-pressure coefficient (ratio of distance of c.p. from leading edge to chord length)\n$\\alpha$, Angle of attack\n$\\epsilon$, Angle of downwash\n$\\alpha_0$, Angle of attack, infinite aspect ratio\n$\\alpha_i$, Angle of attack, induced\n$\\alpha_a$, Angle of attack, absolute (measured from zero-lift position)\n$\\gamma$, Flight-path angle", "timestamp": "2026-07-19T18:42:35.171555+00:00"}
{"citation_id": "19930091697", "source_url": "https://ntrs.nasa.gov/api/citations/19930091697/downloads/19930091697.pdf", "page_number": 14, "total_pages": 28, "image_filename": "19930091697_p14.jpg", "text": "10\nREPORT NO. 622—NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nnecessarily include the knocking zone. With the two fuels of lowest octane number, there is a marked reverse movement in the combustion front before the appearance of the gas vibrations. The indicator cards for these two fuels show a sudden slight rise in pressure just prior to the violent vibrations in the pressure record. Evidently there is a sudden release of energy during this latter part of the burning, and it appears probable that the slight rise in pressure prior to knock bands is the same as the frequency of the waves in the schlieren record, visible in the right-hand portion of the record; they begin simultaneously with the pressure waves and are believed to be caused by successive reillumination of the charge as the pressure waves passed through it. Therefore either a sudden appearance of a brighter light, as at A in figure 7, or the appearance of pressure waves may be used to indicate the start and occurrence of knocking.\n\n[Figure: Indicator cards and schlieren photographs showing effect of different octane ratings on combustion knock. Air-fuel ratio, 14; engine speed, 500 r. p. m.; one spark plug.]\n\nand the reverse movement in the combustion gases are caused by the auto-ignition of the end gas. It is important to note that only the records for explosions in which auto-ignition is indicated show this slight rise. For the record presented in figure 11, adjustments were made to the optical system to permit more of the intensely brilliant light from the knocking combustion to record on the film. In this case, some of the light from the combustion prior to knock also recorded in the ½-inch space nearest the spark plug, where the mirror did not cover the piston. The series of vertical bright bands across the record are evidently successive images of the ⅛-inch slit. The “frequency” of these Motion pictures of the flame and spark schlieren motion pictures taken simultaneously but from slightly different angles are reproduced in figure 12 for a single spark plug at position G. When the refractive index of the charge is uniform throughout, the spark schlieren field appears uniformly illuminated except for spots, lines, and shaded areas caused, respectively, by dirty spots, strain lines, and irregularities in the mirror or window surfaces. When combustion appears in the field, it is visible because of the temperature change and, with the type of schlieren arrangement used, the combustion front appears darker than the field. The region back of the combustion front also appears dark", "timestamp": "2026-07-19T18:42:50.901507+00:00"}
{"citation_id": "19930093641", "source_url": "https://ntrs.nasa.gov/api/citations/19930093641/downloads/19930093641.pdf", "page_number": 29, "total_pages": 47, "image_filename": "19930093641_p29.jpg", "text": "L-452\nN.A.C.A.\nFigs. 3,4\n\n[Figure: A large model of a four-engine aircraft is shown installed in a full-scale wind tunnel. The aircraft is mounted on a support structure, and the image shows the pusher-propeller arrangement.]\n\nFigure 3.- Installation of the 4-engine model in full-scale wind tunnel.\nFour-inch diameter extension shaft housings and pusher-propeller\narrangement.\n\n[Figure: A large model of a four-engine aircraft is shown installed in a full-scale wind tunnel. The aircraft is mounted on a support structure, and the image shows the tractor propellers.]\n\nFigure 4.- Installation of the 4-engine model in the full-scale wind tunnel:\nFour-inch-diameter extension-shaft housings and tractor propellers\n0.39c ahead of wing.", "timestamp": "2026-07-19T18:42:52.055315+00:00"}
{"citation_id": "19930091716", "source_url": "https://ntrs.nasa.gov/api/citations/19930091716/downloads/19930091716.pdf", "page_number": 24, "total_pages": 24, "image_filename": "19930091716_p24.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-19T18:43:00.317869+00:00"}
{"citation_id": "19930091701", "source_url": "https://ntrs.nasa.gov/api/citations/19930091701/downloads/19930091701.pdf", "page_number": 13, "total_pages": 18, "image_filename": "19930091701_p13.jpg", "text": "THE TRANSITION PHASE IN THE TAKE-OFF OF AN AIRPLANE 9\n\ntake-off than for the shortest normal take-off; with the heaviest load there is a larger difference, about 17 percent, owing to the fact that, with other conditions remaining equal, the difference between the excess thrust on the ground and that in flight increases with increasing weight.\n\nIn figures 5, 6, and 7 it will be noted that in some cases, particularly with the lighter loads and higher take-off speeds, the height attained before the conversion of energy is completed is greater than 50 feet. In these cases, at an obstacle height of 50 feet, there is still an excess of kinetic energy remaining, which is equivalent to a loss. Therefore the zoom take-off provides little or no advantage over the normal take-off, as may be seen in figures 9, 10, and 11.\n\nAn indication of the extent of the error that might be introduced into the calculation of take-off performance by the neglect of the transition is provided in figure 12. This figure shows the percentage difference between\n\nthe 50-foot obstacle height, the error increases rapidly with take-off speed to very large values, particularly with the lighter loads. The large errors are due to the fact that the conversion from kinetic to potential energy is not completed until after the 50-foot height has been reached, in which case the assumption of an instantaneous change from ground-run to steady-climb conditions is not justified.\n\nThe scope of this investigation is not sufficiently wide for a definite determination of the relationship that might exist between this error and the airplane characteristics, but it is believed that this relationship could be established with the aid of similar data for other types of airplanes. It would then be possible to obtain a measure of the inherent take-off capabilities of a given airplane, exclusive of the troublesome factor of piloting procedure, by means of a rather simple method. The relation between ground-run distance and speed would be determined in one series of tests;\n\n<!-- Image (93, 369, 867, 575) -->\n\nFIGURE 12.—Error in air-borne distance due to neglect of transition for the Verville AT airplane.\n\nFIGURE 13.—Effect of wind and wind gradient on the flight path of the Verville AT airplane during transition and steady climb. Surface wind velocity, 5 miles per hour.\n\nthe air-borne distance as calculated by the rigorous method and the distance resulting from the assumption that the change from the conditions of the ground run to those of the steady climb occurs instantaneously and without effective loss of energy. For normal take-offs over a 100-foot obstacle the error ranges from a maximum positive value of about 4 percent with the heaviest load, i. e., the approximate distance is too great, to a maximum negative value of the same magnitude. The fact of a positive error is undoubtedly attributable to the influence of ground effect. With a 50-foot obstacle height the error is about twice as great in the same sense for corresponding conditions, since the error in actual distance is about the same.\n\nFor the zoom take-offs over a 100-foot obstacle, the error is comparable at the lower take-off speeds with that for the normal take-offs but becomes increasingly negative as the take-off speed departs more from the minimum value. The largest error in this case, in the range of conditions covered, occurs with the lightest load and has a negative value of about 10 percent. For\n\nother tests, made at some safe altitude providing steady-air conditions, would establish the relationship between angle of climb and speed. These quantities, which should be largely independent of piloting effects, could then be combined, with a suitable correction for a standard type of transition, to give the total distance required to take off over obstacles of any desired height.\n\nThe effects on the air-borne portion of the take-off of a wind increasing in velocity with height are: a reduction in the speed of the airplane relative to the ground, consequently a reduction in the horizontal distance covered in a given time; and an increased vertical velocity due to the velocity gradient. These effects in combination and the effect of the wind gradient alone are shown in figure 13 for normal take-offs with the heaviest and the lightest loads. For the heavy-load condition, the over-all reduction in the distance required to clear a 50-foot obstacle is 25 percent; the reduction due to the wind gradient alone is 16 percent. For an obstacle height of 100 feet, the reductions are 21 percent and 11 percent, respectively. With the light", "timestamp": "2026-07-19T18:43:02.510447+00:00"}
{"citation_id": "19930094498", "source_url": "https://ntrs.nasa.gov/api/citations/19930094498/downloads/19930094498.pdf", "page_number": 5, "total_pages": 37, "image_filename": "19930094498_p5.jpg", "text": "4 N.A.C.A. Technical Memorandum No. 918\n\nFurthermore, the development of even larger and faster seaplanes was being discussed as a serious matter; and, inasmuch as tests of models of such craft would require higher speeds of the towing carriage, the proportion of single-joint runs would be increased, with a consequent reduction in the amount of work accomplished. These facts indicated that an increase in the length of the basin and increases in the speed and the power of the towing carriage would be desirable.\n\nA survey of the available space showed that the maximum amount by which the length of the basin could be increased was 900 feet and the extension of the tank by that amount was authorized. The enlarged tank was opened for work in October 1937. The relative lengths of the original and the enlarged tank can be seen in figure 1.\n\nPRINCIPAL CHARACTERISTICS OF THE ENLARGED N.A.C.A. TANK\n\nThe most conspicuous of the features of the enlarged N.A.C.A. tank are derived directly from those of the original tank and owe their present form not only to the reasons for their first use but also to the experience obtained with them. As in the original tank, there are:\n\n1. A basin of great length (now 2,880 ft.).\n\n2. Rails made of structural H beams, without machining.\n\n3. A towing carriage of very high speed (now 80 m.p.h., maximum).\n\n4. Rubber tires on all the wheels, pneumatic on the running wheels and solid on the guide wheels.\n\nThese features, together with some related matters, will now be discussed in more detail, in order that their effects on the methods of testing models and the methods of recording data may be more clearly seen.\n\nBasin.- The reinforced concrete basin of the enlarged N.A.C.A. tank has the following dimensions:\n\n| | Feet |\n| :--- | :--- |\n| Length on water, extreme | 2,920 |\n| Normal width of water surface | 24 |", "timestamp": "2026-07-19T18:43:10.958136+00:00"}
{"citation_id": "19930094542", "source_url": "https://ntrs.nasa.gov/api/citations/19930094542/downloads/19930094542.pdf", "page_number": 88, "total_pages": 102, "image_filename": "19930094542_p88.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-19T18:43:12.467153+00:00"}
{"citation_id": "19930094549", "source_url": "https://ntrs.nasa.gov/api/citations/19930094549/downloads/19930094549.pdf", "page_number": 26, "total_pages": 76, "image_filename": "19930094549_p26.jpg", "text": "24 N.A.C.A. Technical Memorandum No. 867\n\ning the control to $\\beta_2$, passes from point A to point B he may be considered to have departed from his final state C by amounts: $\\delta i = i_1 - i_2$, $\\delta V = V_1 - V_2$, $\\delta \\xi = (i_1 - i_2) + \\zeta$, where $\\zeta$ is computed from the polar of the airplane and the variation of the thrust T with velocity.\n\nWe have applied this computation to a numerical example, choosing the same airplane as above. At about the velocity of 40 m/s (90 m.p.h.), it was found that states differing from each other by a difference in velocity $\\delta V = \\pm 1$ m/s are characterized by a difference in angle of attack of $\\mp 0.308^\\circ$.\n\nSimilarly, basing our computations on an assumed polar, we found that a variation in velocity of $\\pm 1$ m/s requires that the slope of the flight path vary by $\\mp 0.0070$, corresponding to $d\\zeta = -0.400^\\circ$. These data were applied by us to the numerical example by multiplying them by the factor necessary to obtain amplitudes that may easily be studied.\n\nConsider, for example, the passage from the level flight condition at velocity $V = 27$ m/s to a state where $V = 40$ m/s by means of a deflection $\\Delta \\beta$ required to produce a $\\Delta i = -4^\\circ$. If this deflection $\\Delta \\beta$ is suddenly applied, the airplane is deviated from its final condition by:\n\n$$\n\\delta V = -13; \\quad \\delta i = +4^\\circ; \\quad \\delta \\xi = -5.2^\\circ\n$$\n\nNow $\\delta \\theta = \\delta \\xi - \\delta i$, so that $\\delta \\theta = -9.2^\\circ$, one portion of which represents the change in angle of attack, and the other the change in the flight path. It suffices to reduce the amplitudes of the disturbances of figures 10, 11, and 12 in the desired ratio and add them. Figure 14 (continuous line) shows the result of this computation for the airplane of coefficient $\\mu = 0.004$. The diagram of the angles of attack shows that the final angle of attack is not established rapidly. We may trace separately the effect of each of the elements of the initial disturbance on the variable $\\delta i$ (fig. 15). The effect of the initial disturbance in the angle of attack decreases very rapidly. The initial disturbances of velocity and attitude, however, also have an effect on the variable $i$ — an insufficient velocity tending to increase the angle of attack, and similarly for a not sufficiently descending path. These two initial disturbances act on the long-period oscillation. It may be seen, in the example chosen, that in spite of the rapidity", "timestamp": "2026-07-19T18:43:17.219470+00:00"}
{"citation_id": "19930094493", "source_url": "https://ntrs.nasa.gov/api/citations/19930094493/downloads/19930094493.pdf", "page_number": 5, "total_pages": 31, "image_filename": "19930094493_p5.jpg", "text": "4 N.A.C.A. Technical Memorandum No. 923\n\n$$\n\\lambda_{\\mathrm{m}} = \\frac{1}{T_{\\mathrm{w}} - T_{\\mathrm{o}}} \\int_{T_{\\mathrm{o}}}^{T_{\\mathrm{w}}} \\lambda \\, \\mathrm{d}t\n\\tag{6}\n$$\n\n$$\n\\eta_{\\mathrm{m}} = \\frac{1}{T_{\\mathrm{w}} - T_{\\mathrm{o}}} \\int_{T_{\\mathrm{o}}}^{T_{\\mathrm{w}}} \\eta \\, \\mathrm{d}t\n\\tag{7}\n$$\n\n$$\n\\begin{aligned}\n\\rho_{\\mathrm{m}} &= \\frac{1}{\\xi} \\frac{1}{T_{\\mathrm{w}} - T_{\\mathrm{o}}} \\int_{T_{\\mathrm{o}}}^{T_{\\mathrm{w}}} \\gamma \\, \\mathrm{d}t \\\\\n&= \\frac{1}{\\xi} \\gamma_{\\mathrm{o}} \\frac{T_{\\mathrm{o}}}{T_{\\mathrm{m}}}\n\\end{aligned}\n\\tag{8}\n$$\n\nwhere $\\gamma_{\\mathrm{o}}$ (kg/m³) is the specific weight of the medium at temperature $T_{\\mathrm{o}}$,\n\n$$\nT_{\\mathrm{m}} = \\frac{T_{\\mathrm{w}} - T_{\\mathrm{o}}}{\\ln \\frac{T_{\\mathrm{w}}}{T_{\\mathrm{o}}}}\n\\tag{9}\n$$\n\nand $\\xi$ (m/s²) is the acceleration of gravity.\n\nThe function $f$ must be determined experimentally. There are a number of tests available on the heat transfer of thin wires. On the basis of the results of these tests, the function $f$ is found to be:\n\n$$\n\\frac{\\alpha \\cdot d}{\\lambda_{\\mathrm{m}}} = m \\left[ \\frac{d \\cdot \\rho_{\\mathrm{m}} \\cdot w}{\\eta_{\\mathrm{m}}} \\right]^n = m (\\mathrm{Re})^n\n\\tag{10}\n$$\n\nThe values of the coefficient $m$ and the exponents $n$ are given according to J. Ulsamer in the table below:\n\n| For | Re | n | m |\n| :--- | :--- | :--- | :--- |\n| | $0.1 < \\mathrm{Re} < 4$ | 0.305 | 0.875 |\n| | $4 < \\mathrm{Re} < 50$ | .41 | .764 |\n| | $50 < \\mathrm{Re}$ to 1,000 | .5 | .537 |", "timestamp": "2026-07-19T18:43:26.367015+00:00"}
{"citation_id": "19930093281", "source_url": "https://ntrs.nasa.gov/api/citations/19930093281/downloads/19930093281.pdf", "page_number": 7, "total_pages": 22, "image_filename": "19930093281_p7.jpg", "text": "```markdown\n5\n\nAll the tests were made with zero air flow through the nacelle to eliminate the effect of cowling pumping efficiency on the results. The struts were shielded from the air stream as shown in figures 2 to 7. Because the tare drag remained constant for each set-up, the results are not corrected for this effect. The results are corrected, however, for the effect of horizontal buoyancy because this effect varies with the body shape and the location of the test arrangement in the tunnel. The magnitude of the effect of horizontal buoyancy can be seen in table I. Set-up 1 was located 2 feet farther forward in the tunnel than set-up 2. Inasmuch as the static pressure in the air stream increases toward the entrance cone, the buoyancy corrections for set-up 1 were larger than for set-up 2.\n\nDISCUSSION OF FIGURES\n\nThe condensed results of the drag tests are given in table I. This table shows that the more streamlined afterbody reduced the drag increment chargeable to the open-nose N.A.C.A. cowling, $C_D - C_{D_0}$, from 0.0350 to 0.0061.\n\nThe net efficiency was computed from the net force on the tunnel balance. The net efficiency for each test was plotted against $1/\\sqrt[3]{P_C} \\left( = \\sqrt[3]{\\frac{6S}{2P}} \\right)$. Envelopes were drawn from the composite of all the propeller tests for each arrangement. The net efficiency envelopes are shown in figures 9, 10, and 11. A comparison of the envelopes for any propeller at constant values of $1/\\sqrt[3]{P_C}$ shows the power cost of the front opening of the cowling when in the presence of that propeller.\n\nThese propeller results strictly apply only to the ratio of F/S used in the test arrangement. If the value of F/S were larger, the effect of the nose opening would be somewhat greater than noted and, if smaller, the reverse would be true. Since the present trend is to put more engine power into the same engine diameter, this ratio has been decreasing because the greater engine power requires larger propeller diameters. The test arrangement is near the upper end of the range of F/S used and, consequently, the effect of the nose opening discussed in this report is larger than will be experienced in most modern installations.\n```", "timestamp": "2026-07-19T18:43:32.907156+00:00"}
{"citation_id": "19930091724", "source_url": "https://ntrs.nasa.gov/api/citations/19930091724/downloads/19930091724.pdf", "page_number": 5, "total_pages": 20, "image_filename": "19930091724_p5.jpg", "text": "REPORT No. 649\n\nTHE “PACK” METHOD FOR COMPRESSIVE TESTS OF THIN SPECIMENS OF MATERIALS USED IN THIN-WALL STRUCTURES\n\nBy C. S. AITCHISON and L. B. TUCKERMAN\n\nSUMMARY\n\nThe strength of modern lightweight thin-wall structures is generally limited by the strength of the compression members. An adequate design of these members requires a knowledge of the compressive stress-strain graph of the thin-wall material. The “pack” method was developed at the National Bureau of Standards with the support of the National Advisory Committee for Aeronautics to make possible a determination of compressive stress-strain graphs for such material.\n\nIn the “pack” test an odd number of specimens are assembled into a relatively stable pack, like a “pack of cards.” Additional lateral stability is obtained from lateral supports between the external sheet faces of the pack and outside reactions. Studies have been made of the reproducibility of the test results by testing packs taken from sheets of aluminum alloy 17ST and steel. The largest spread in yield strength was about 2 percent. Tests were also made to determine whether the results from packs were like those obtained from compact solid specimens. The results indicated that the method of transverse support had no appreciable effect on the yield strength. The largest difference between a pack and a solid specimen was 1.60 percent. Experience gathered in developing the test emphasized the fact that, while the method seemed to furnish results within the same order of accuracy as was usually obtained from other mechanical tests, such as the tensile test, it must be simplified before it can be used economically for inspection testing. The test seems adequate, however, for many problems in structural research.\n\nINTRODUCTION\n\nDuring recent years a remarkable expansion has taken place in the use of thin sheet and thin-wall material in lightweight structures such as airplane wings, and airplane fuselages. The strength of these structures is generally limited by the strength of certain members carrying compressive loads. These members have frequently been designed on the basis of the tensile properties of the material. This is convenient as the tensile test is relatively simple and is widely used. However, it may lead to an unsafe structure, on the one hand, or an uneconomical structure, on the other hand, if the compressive properties of the material differ from the tensile properties. There is an urgent need for a method which makes possible a direct determination of compressive stress-strain graphs for thin-wall material. In recognition of this need an investigation has been undertaken by the National Bureau of Standards with the financial support of the National Advisory Committee for Aeronautics.\n\nSpecimens of thin sheet usually fail through instability before the yield strength is reached. Some methods have been reported for overcoming this difficulty by assembling the material under consideration into a compact unit similar to a compact solid. By these methods failures through instability occur at higher compressive loads.\n\nE. B. Wolff and L. J. G. Van Ewijk (reference 1) made compressive tests on carefully selected wood and compared the results from “massive” bars with those from bars built up by gluing together lamellae taken from the same wood. They reported that the elastic properties for both kinds of specimens were the same.\n\nA. Robertson (reference 2), in his investigation of “The Strength of Tubular Struts,” gives compressive results on various tubes which were made from strips of wood, about 0.025 inch thick, “* * * by wrapping the necessary number of strips round a mandril having first spread a fine coating of glue on all the faces that were to come together.” He adds that “* * * the collapsing stress is uniform and practically that of the solid specimen for all values of * * * ratios of thickness of the wall to radius of the tube greater than 0.08. In his report Robertson suggests, also, the possibility of combining sheet metal into compact units. He made some experiments on high tensile steel strip, about 0.015 inch thick. He states that “It is very difficult to get a good compression test of the material when in the form of such thin strips. An attempt was made to make a test piece by soldering together a large number of pieces and then machining the resulting block to a square section. The result, however, was not satisfactory.”\n\n1", "timestamp": "2026-07-19T18:43:35.698165+00:00"}
{"citation_id": "19930091748", "source_url": "https://ntrs.nasa.gov/api/citations/19930091748/downloads/19930091748.pdf", "page_number": 6, "total_pages": 12, "image_filename": "19930091748_p6.jpg", "text": "```markdown\n2\nREPORT NO. 673—NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nPHASE RELATIONS FOR OSCILLATING AIRFOILS\n\nFollowing the method of Theodorsen (reference 4), the lift on an airfoil oscillating in pitch, in a uniform stream of perfect fluid, with one degree of freedom about an axis parallel to the span is given (reference 5) as\n\n$$\n\\begin{aligned}\nL &= \\rho b^2 (\\pi v \\alpha_0 p \\cos pt + \\pi b a \\alpha_0 p^2 \\sin pt) \\\\\n&+ 2\\pi \\rho b V \\left[ v \\alpha_0 \\sin pt + b \\left( \\frac{1}{2} - a \\right) \\alpha_0 p \\cos pt \\right] \\\\\n&+ 2\\pi \\rho b G \\left[ v \\alpha_0 \\cos pt - b \\left( \\frac{1}{2} - a \\right) \\alpha_0 p \\sin pt \\right]\n\\end{aligned}\n\\tag{1}\n$$\n\nRegrouping terms and substituting $k = pb/v$ and $a = -\\frac{1}{2}$ (axis at quarter-chord line), the expression becomes\n\n$$\nL = \\pi \\rho b \\alpha_0 v^2 (k + 2kF + 2G) \\cos pt + \\pi \\rho b \\alpha_0 v^2 \\left( -\\frac{1}{2}k^2 + 2F - 2kG \\right) \\sin pt\n\\tag{2}\n$$\n\nor\n\n$$\nL = A_0 \\sin pt + B_0 \\cos pt\n\\tag{3}\n$$\n\nin which\n\n$$\n\\begin{aligned}\nA_0 &= \\pi \\rho b \\alpha_0 v^2 \\left( -\\frac{1}{2}k^2 + 2F - 2kG \\right) \\\\\nB_0 &= \\pi \\rho b \\alpha_0 v^2 (k + 2kF + 2G)\n\\end{aligned}\n\\tag{4}\n$$\n\nExpression (3) may be readily rewritten as\n\n$$\nL = C_0 \\sin (pt + \\delta)\n\\tag{5}\n$$\n\nin which\n\n$$\n\\delta = \\tan^{-1} (B_0 / A_0)\n\\tag{6}\n$$\n\nand\n\n$$\nC_0 = \\sqrt{A_0^2 + B_0^2}\n\\tag{7}\n$$\n\nThe angle $\\delta$ is the phase difference due to the oscillation, and $C_0$ is a measure of the slope of the lift curve for the oscillating wing. It will be noted from equations (4) and (6) that the value of the phase angle $\\delta$ is a function of $k$, $F$, and $G$. The term $k$ is a fundamental parameter that links together the frequency of the wing oscillation and the air speed. It may be noted that $2\\pi/k$ is the distance between successive waves in the vortex sheet in terms of the half-chord length $b$ as a reference length. The functions $F$ and $G$ determine the circulation so as to satisfy the Kutta condition for the oscillating airfoil, and their values are given as functions of $1/k$ for the case of an infinite-span airfoil in reference 4. For finite aspect ratio, reference 6 gives certain approximate corrections employing \"effective\" values of $F$ and $G$.\n\nAPPARATUS AND TEST METHOD\n\nThe tests were conducted in the $\\frac{1}{15}$-scale model of the N. A. C. A. full-scale wind tunnel, which is described in reference 7. The test section was modified to a 2- by 3-foot rectangle with sides but without top or bottom. The side walls served as end plates for the airfoil that spanned the 3-foot width of the jet. Surveys across the tunnel air stream showed variations of $\\pm$ 2 percent in the dynamic pressure and $\\pm$ 0.6° in the air-stream direction, the effects of which were not considered important enough to warrant correction.\n\nA diagram and a photograph of the apparatus used are shown in figures 1 and 2, respectively. The airfoil F (fig. 1) was an 18-percent-thick symmetrical section with a chord of $5\\frac{3}{16}$ inches and a span of $36\\frac{1}{2}$ inches. It was of hollow construction with 0.016-inch sheet-aluminum covering and weighed only 0.66 pound. The counterweight I, the end plates C, the mounting shafts, and the self-aligning ball bearings B brought the total weight of the oscillating assembly up to 1.52 pounds. The airfoil projected through the sides of the test section L, and separate end plates C were attached that rotated with the airfoil to prevent air leakage through the walls due to the local pressures. This arrangement was used to achieve an effective aspect ratio approaching that for an infinite-span airfoil. The axis of oscillation, i. e., the axis of the mounting shafts, was located at the quarter-chord point of the airfoil.\n\nThe instantaneous force balance was designed to measure the force on one end of the airfoil by measuring the deflection of a stiff spring P to which the airfoil was attached. The attachment was made by means of a self-aligning ball bearing, which provided freedom for the wing to rotate about its axis and for the spring to deflect without restraint. In order to record the deflection of the spring, two styluses were arranged to rotate the mirrors J and N and thereby displace light beams that were focused on a sensitized film. The spring deflections in both the lift and the drag directions were recorded by mirrors J and N. (This paper is confined to a discussion of the lift forces only.) In order to obtain a continuous record, the recording film was attached to a circular drum rotated by a synchronous motor H at a constant film speed of 39 inches per second. The balance was calibrated by means of loads applied at the center of the airfoil.\n\nThe design of the spring P was dictated by the consideration that accurate measurements of the phase difference by means of a spring require the natural frequency of the vibrating system to be considerably higher than the impressed frequency. The angle of lag $\\theta$ of the recorded force behind the impressed force in terms of the natural frequency of the vibrating system $r$ is given (reference 8) as\n\n$$\n\\theta = \\tan^{-1} \\frac{2pn}{r^2 - p^2}\n$$\n\nin which $n$, the damping coefficient, is obtained from a measurement of the decrease in amplitude of successive vibrations. That is,\n\n$$\nA_2 = A_1 e^{-n\\tau}\n$$\n\nin which $A_1$ and $A_2$ are the amplitudes of successive vibrations occurring in the period $\\tau$. The values of the constants $n$ and $r/2\\pi$ for the wing and the balance are\n```", "timestamp": "2026-07-19T18:43:43.151331+00:00"}
{"citation_id": "19930091719", "source_url": "https://ntrs.nasa.gov/api/citations/19930091719/downloads/19930091719.pdf", "page_number": 7, "total_pages": 33, "image_filename": "19930091719_p7.jpg", "text": "FIVE FULL-SCALE PROPELLERS IN THE PRESENCE OF A RADIAL AND A LIQUID-COOLED NACELLE 3\n\nlining them with thin sheet-metal cuffs. These cuffs, shown in figure 5, extended along the blade shanks for a distance of about 4 inches beyond the spinner and were secured to the spinner. The blades were thus enclosed for a distance of about 24 percent of the radius.\n\nFive 3-blade propellers (fig. 6), all having diameters of 10 feet, were tested. Blade-form curves are given in figures 7 and 8. The propeller dimensions are given by the following notation: $D$, diameter; $R$, radius to the tip; $r$, station radius; $h$, section thickness; $b$, station chord; $p$, geometric pitch. Figure 9 shows the section\n\nThe principal propeller dimensions are given in the following table:\n\n| Propeller drawing number | Diameter (feet) | Section | $b/D$ at $0.75R$ | $h/b$ at $0.75R$ | Shank shape |\n| :--- | :--- | :--- | :--- | :--- | :--- |\n| Bureau Aeronautics 5808-9. | 10 | Clark Y | 0.061 | 0.09 | Round. |\n| Hamilton Standard 1C1-6 | 10 | ...do... | .059 | .07 | Airfoil. |\n| Hamilton Standard 6101 | 10 | ...do... | .059 | .07 | Round. |\n| Hamilton Standard 6129 | 10 | R. A. F. 6 | .059 | .07 | Do. |\n| Hamilton Standard 6131 | 10 | N. A. C. A. 2400-34. | .059 | .07 | Do. |\n\n$^1$ Controllable.\n\n[Figure: Three propeller blades mounted on hubs. From left to right, they are labeled 6101, 6129, 6131; 1C1-6; and 5808-9.]\n\nFIGURE 6.—The propeller blades tested.\n\noutline and gives the ordinates for the three blade sections incorporated in the different propellers. It may be noted that the N. A. C. A. 2400-34 airfoil section is modified for propeller design by changing the thickness with respect to the mean camber line. The camber therefore remains constant for the whole blade, whereas the camber increases with blade section thickness for propellers having the Clark Y and R. A. F. 6 sections.\n\nIt may be noted from the table that the essential difference between propellers 5808-9 and 6101 is the blade thickness although propeller 6101 has a slightly larger shank diameter and a different hub, which should not appreciably affect the results. These two propellers probably represent the upper and lower limits in thickness ratios for present-day aluminum-alloy propellers.", "timestamp": "2026-07-19T18:43:45.559578+00:00"}
{"citation_id": "19930091731", "source_url": "https://ntrs.nasa.gov/api/citations/19930091731/downloads/19930091731.pdf", "page_number": 3, "total_pages": 32, "image_filename": "19930091731_p3.jpg", "text": "---\n\nREPORT No. 656\n\nTHE COLUMN STRENGTH OF TWO EXTRUDED  \nALUMINUM-ALLOY H-SECTIONS\n\nBy WILLIAM R. OSGOOD and MARSHALL HOLT\n\nNational Bureau of Standards and Aluminum Company of America\n\nI", "timestamp": "2026-07-19T18:43:54.630610+00:00"}
{"citation_id": "19930094542", "source_url": "https://ntrs.nasa.gov/api/citations/19930094542/downloads/19930094542.pdf", "page_number": 89, "total_pages": 102, "image_filename": "19930094542_p89.jpg", "text": "N.A.C.A. Technical Memorandum No. 874\n\n[Figure: Graph showing spanwise lift distribution with multiple curves for different angles of attack (α = 2°, 4°, 6°, 8°) and wing alone, plotted against y/b/2 from -35 to 35; vertical axis labeled C_L from 0.2 to 2.2; annotations include λ=0.16, α=8°, α=-2°, and “End plate” at right edge]\n\nFigure 75.-Spanwise lift distribution.λ = 0.16\n\n[Figure: Series of 10 rows, each containing four panels: leftmost shows C_L vs x/c curve; second shows top-down view of wing with shaded separation region; third shows cross-section with flow separation pattern; rightmost shows C_L vs x/c curve again — all illustrating progression of flow separation over the wing as angle of attack increases]\n\nFigure 76.-Spread of flow separation over the wing.\n\nFigs. 75, 76", "timestamp": "2026-07-19T18:44:12.475577+00:00"}
{"citation_id": "19930094498", "source_url": "https://ntrs.nasa.gov/api/citations/19930094498/downloads/19930094498.pdf", "page_number": 6, "total_pages": 37, "image_filename": "19930094498_p6.jpg", "text": "N.A.C.A. Technical Memorandum No. 918 5\n\nFeet\nNormal depth of water 12\nLength of 12-foot depth 2,860\n\nIn the old part of the tank the side walls are coved\nin above the water in order to bring the rails closer to-\ngether and also to help reduce the waves. In the exten-\nsion the side walls extend vertically above the water\nlevel for 15 inches and then meet the horizontal lower\nsurface of the overhangs that correspond to the original\ncoves. The purpose of this change in section is to make\nit possible for waves to run freely in the extension when\nthe wave suppressors are removed, although they will begin\nto dissipate when they strike the coves in the old section.\n\nThe two sections are compared in figure 2.\n\nThe canopy of the extension is practically the same\nin structure and arrangement as that of the original tank.\n\nThe rails upon which the towing carriage runs are\nstructural H beams set with the web vertical, as in the\noriginal tank, and are supported on chairs of the same\ntype as in the original tank.\n\nTowing carriage.- The original towing carriage of the\nN.A.C.A. tank had four wheels, each fitted with a large\npneumatic tire of the type used on high-speed buses. These\ntires were not a standard type of tire but were specially\nmade with smooth treads. The loads on each tire was about\n5,000 pounds and the wheels and tires were large and heavy.\n\nChanging wheels and tires was a difficult and labori-\nous process and it was concluded that operation would be\nbetter if a standard-size tire could be used and the load\nper tire be thereby reduced. These objectives were accom-\nplished by doubling the number of wheels and reducing their\nsize to suit a tire that is regularly made with a smooth\ntread. Doubling the number of wheels made it possible to\ndouble the number of propelling motors and thus to increase\nthe acceleration and the maximum speed of the towing carri-\nage.\n\nThe towing carriage used on the enlarged N.A.C.A.\ntank, as shown in the diagram (fig. 3), was made by remov-\ning from the four corners of the old carriage the structure", "timestamp": "2026-07-19T18:44:12.830364+00:00"}
{"citation_id": "19930094506", "source_url": "https://ntrs.nasa.gov/api/citations/19930094506/downloads/19930094506.pdf", "page_number": 1, "total_pages": 24, "image_filename": "19930094506_p1.jpg", "text": "FILE COPY\nNO 4\n\nTECHNICAL MEMORANDUMS\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nNo. 910\n\nMEASUREMENTS ON A LOW-WING MODEL IN THE ROTATING JET\nAND COMPARISON WITH FLIGHT MEASUREMENTS\n\nBy W. Eader\n\nLuftfahrtforschung\nVol. 16, No. 2, February 20, 1939\nVerlag von R. Oldenbourg, München und Berlin\n\n[annotation: THIS DOCUMENT ON LOAN FROM THE FILES OF\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\nLANGLEY AERONAUTICAL LABORATORY\nLANGLEY FIELD, HAMPTON, VIRGINIA\nRETURN TO THE ABOVE ADDRESS.\nREQUESTS FOR PUBLICATIONS SHOULD BE ADDRESSED\nAS FOLLOWS:\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n1512 H STREET, N. W.\nWASHINGTON 25, D. C.]\n\nWashington\nSeptember 1939", "timestamp": "2026-07-19T18:44:17.274722+00:00"}
{"citation_id": "19930094549", "source_url": "https://ntrs.nasa.gov/api/citations/19930094549/downloads/19930094549.pdf", "page_number": 27, "total_pages": 76, "image_filename": "19930094549_p27.jpg", "text": "N.A.C.A. Technical Memorandum No. 867 25\n\nwith which the short-period oscillation tends to bring the airplane to its final angle of attack, the airplane may in certain cases remain considerably deviated from its final position by the effect of the components of $\\delta i$ that depend on the long-period oscillation.\n\nRemarks\n\n1. If we assume that the deflection $\\Delta \\beta$ has not been instantaneously applied, but in a progressive manner (for example, one-third at time zero, the second third at time 0.5 second, and the remainder at time $t = 1$ second), the curve of variation of angle of attack loses the abrupt character which it has during the first second and becomes regular with its appearance not appreciably modified thereafter. This condition corresponds to the dash-dot curve of figure 14.\n\n2. The airplane will more quickly attain its final angle of attack if the pilot, to start the maneuver, gives the control a greater deflection than the amount $\\Delta \\beta$, which he must maintain at the end of the phenomenon.\n\n3. The computation is made on the assumption that the pilot does not vary the engine throttle. If the pilot increases the power the final path will no longer be descending. The effect of such a maneuver would have to be separately studied.\n\nMORE COMPLICATED MANIPULATIONS OF THE CONTROLS\n\nThe same procedure may, of course, be applied to analyze the effect of more complicated maneuvers. Let an initial state be characterized by a given velocity attitude, angle of attack, and deflection $\\beta$. For example, in the case of the airplane considered above, $V = 27$ m/s; $i = +4^\\circ$; $\\theta = -4^\\circ$.\n\nThe pilot deflects his control by $\\Delta \\beta$, and the airplane starts the motions which tend to bring it to a final state B. He applies this deflection $\\beta + \\Delta \\beta$, however, only within a limited time interval, and re-establishes the deflection $\\beta$ after $n$ seconds. At this instant the airplane has not yet attained its state B but is in an intermediate state. We see on the figure that the momentary characteristics after 7 seconds are: $V = 34.7$ m/s; $i = 1^\\circ$; $\\theta = +13.4^\\circ$; $q = +1^\\circ$ per second.", "timestamp": "2026-07-19T18:44:21.316674+00:00"}
{"citation_id": "19930094499", "source_url": "https://ntrs.nasa.gov/api/citations/19930094499/downloads/19930094499.pdf", "page_number": 4, "total_pages": 13, "image_filename": "19930094499_p4.jpg", "text": "N.A.C.A. Technical Memorandum No. 917 3\n\nor less severe disturbance in the flow at the static orifice of the Prandtl pitot, with the result that the pressure of undisturbed flow can no longer be measured at that point.\n\nStrictly speaking, the speed prediction with the Prandtl pitot therefore is contingent upon the knowledge of the errors of the pressure readings from calibration of the instrument and their correction. The experimental determination of this correction is related in the following.\n\nNotation\n\nw, velocity in undisturbed flow;\n\na, velocity of sound in undisturbed flow;\n\nρ, density in undisturbed flow;\n\np, pressure in undisturbed flow;\n\np₀, tank pressure;\n\np₁, pressure at forward orifice of Prandtl pitot;\n\np₂, pressure at lateral orifice of Prandtl pitot;\n\nκ, ratio of specific heat at constant pressure and constant volume, respectively.\n\nDESCRIPTION OF TEST EQUIPMENT\n\nThe high-speed tunnel (Prandtl type) employed is shown in figure 1. A container B is pumped empty so that the atmospheric air (pressure p₀, density ρ₀, and speed w₀ = 0) flows, after opening of cock H, as a free jet through the test chamber M into the container. On leaving the entrance cone E the air stream attains the speed w, the pressure expands to p and the density to ρ; a indicates the velocity of sound in the free jet. The wind velocity w is regulated with the nozzle V whose narrowest section is adjustable. So long as the pressure in B stays low enough to produce velocity of sound in the narrowest section of V, the jet remains stationary and the flow volume and jet velocity, respectively, change only", "timestamp": "2026-07-19T18:44:32.748884+00:00"}
{"citation_id": "19930093281", "source_url": "https://ntrs.nasa.gov/api/citations/19930093281/downloads/19930093281.pdf", "page_number": 8, "total_pages": 22, "image_filename": "19930093281_p8.jpg", "text": "6\n\nThe change in the net efficiency may be defined as\n\n$$\n\\eta_0 - \\eta_n = \\frac{\\Delta C_D}{P_c} \\frac{F}{S}\n$$\n\nwhere $\\Delta C_D$ is the effective change in the drag coefficient caused by the nose shape. The curves used as the basis of comparison are designated $\\eta_0$ curves. The $\\eta_0$ curves for propellers C and $C_x$ were obtained with nose 4, and the $\\eta_0$ curve for propeller B was obtained with nose 5 and spinner 1 because the spinner for nose 4 would not fit propeller B. The change in drag coefficient $\\Delta C_D$ is a combination of the increment of drag of the body and the change in the propeller efficiency caused by body interference and by the drag of the exposed propeller hub and blade shanks.\n\nIn figure 12, the effective $\\Delta C_D$ caused by the nose shapes is plotted against $1/\\sqrt[3]{P_c}$. For small values of $1/\\sqrt[3]{P_c}$, the main effect is the change in body drag produced by high velocities over the nacelle; for larger values of $1/\\sqrt[3]{P_c}$, the main effect is the change in propeller efficiency.\n\nThe preceding fact is illustrated in figure 12(b) in the $\\Delta C_D$ curve for nose 5 without spinner. This arrangement has the smallest value of $\\Delta C_D$ for any nose tested with this propeller for values of $1/\\sqrt[3]{P_c}$ below 2.0 and the highest value of $\\Delta C_D$ for values of $1/\\sqrt[3]{P_c}$ of 3.4 or more. The fact that the values of $\\Delta C_D$ up to $1/\\sqrt[3]{P_c} = 2.0$ are low shows that the slipstream-drag effect is small. The fact that the values of $\\Delta C_D$ at $1/\\sqrt[3]{P_c} = 3.4$ or more are high shows that the power absorbed by the propeller hub and the blade shanks in the relatively high-velocity air stream of nose 5 without spinner is large.\n\nThe addition of spinner 1 to nose 5 decreases the power absorbed by the inner part of the propeller and makes the arrangement of nose 5 with spinner as good as any tested in the high-speed range, except nose 4.", "timestamp": "2026-07-19T18:44:33.926440+00:00"}
{"citation_id": "19930091724", "source_url": "https://ntrs.nasa.gov/api/citations/19930091724/downloads/19930091724.pdf", "page_number": 6, "total_pages": 20, "image_filename": "19930091724_p6.jpg", "text": "2\nREPORT NO. 649 NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\n\"PACK\" TEST\n\nThe successful results for tests where pieces of wood are combined into compact units suggest that the compressive properties can be obtained when there is sufficient lateral stability so that the yield strength is reached before the unit buckles.\n\nWith this approach a number of methods were tried at this Bureau to develop an adequate technique for compressive tests of thin-wall material. A compressive test (reference 3), which has become known as the \"pack\" test, has resulted from this preliminary work. The \"pack\" test is described in detail in the following pages. The details are given very fully because minor deviations from these details have, in some cases, produced unsatisfactory results and the necessary time has not been available to investigate just which of these are essential and which are unessential to the success of the test.\n\nThe method was developed at this Bureau in 1933 and has given satisfactory results in all those cases in which the detail procedure, given below, was closely followed.\n\nThe \"pack\" test involves the use of external support supplied by a number of transverse members between outside reactions and the external sheet faces of a \"pack\" of specimens. The test was intended to simulate a block compressive test on a compact solid specimen of the kind described (reference 4), in a tentative specification of the American Society for Testing Materials as a \"medium-length\" specimen to determine the \"general compressive strength properties of metallic materials.\" It was not intended, however, to determine the modulus of elasticity for which the \"long\" specimen described in this specification would be preferable.\n\nThe \"pack\" was composed of an odd number of rectangular specimens taken from the same material. These were assembled with sheet faces in contact to form a compact unit. The strains were measured on the middle specimen of the pack which, therefore, acted as a compression specimen supported on both faces by the remaining specimens of the pack. The specimens were machined using procedures similar to those normally employed for tensile specimens. This avoids other operations, such as forming, riveting, or welding, which are frequently used to stiffen structures and which might change the properties of the specimens.\n\nThe lateral supports of the pack were designed to give adequate support against buckling combined with a minimum resistance to displacements parallel to the load. Emphasis was placed on this requirement in order to assure that the method of support would not alter the stress distribution in the compression specimen.\n\nTHE \"PACK\"\n\"MIDDLE\" SPECIMEN\n\nThe middle specimen M of a pack composed of 9 specimens taken from a piece of steel tubing is shown in fig. 1. The middle specimen is also shown in fig. 2 in a pack of 13 specimens taken from aluminum alloy sheet.\n\nThe compressive load $P$ was applied parallel to the length of the pack and was distributed over the ends E of the pack. The lateral edge faces were nominally parallel to the load axis. These faces were left clear so that gages to measure the strain could be attached to the specimen. Stability in the direction of the width was obtained by making the specimen sufficiently wide in comparison to the length. The width $h$ was about $\\frac{25}{32}$ inch. The length $L$ was twice the width plus one inch or about $2\\frac{7}{16}$ inches.\n\n[Figure 1: \"Pack\" taken from tubing.]\n\n[Figure 2: \"Pack\" taken from sheet.]\n\n\"SUPPORTING\" SPECIMENS\n\nAll of the specimens in the pack were subjected to the axial load. For this reason the supporting specimens S, fig. 2, were made the same length as the middle specimen. In order that they would not interfere", "timestamp": "2026-07-19T18:44:36.695661+00:00"}
{"citation_id": "19930093640", "source_url": "https://ntrs.nasa.gov/api/citations/19930093640/downloads/19930093640.pdf", "page_number": 5, "total_pages": 28, "image_filename": "19930093640_p5.jpg", "text": "```markdown\n3\n\nV, air speed, f.p.s.\n\nL, lift, or force normal to the relative wind, lb.\n\nD, drag, or force parallel to the relative wind, lb.\n\n$D_o$, power-off drag of combination, lb.\n\nM, pitching moment, lb.-ft.\n\n$C_L = L/qS$\n\n$C_D = D/qS$ (Subscript w refers to power-off drag of the model with bare wing; c, to power-off drag of the model with engine-nacelle installation.)\n\n$C_m = M/qS\\bar{c}$\n\nR, resultant drag force of a propeller-body combination, lb.\n\nT, thrust of propellers operating in front of a body (tension in propeller shafts), lb.\n\n$\\Delta D$, increase in drag of the body behind the propellers due to the action of the propellers.\n\n$T - \\Delta D$, effective thrust of the propeller-body combination.\n\n$T_o$, index thrust.\n\nP, power input per propeller.\n\n$P_{tot}$, total power input to propellers.\n\n$C_T = \\frac{T - \\Delta D}{\\rho n^2 D^4}$\n\n$C_P = \\frac{P}{\\rho n^3 D^5}$\n\n$\\eta = \\frac{(T - \\Delta D) V}{P} = \\text{propulsive efficiency.}$\n\n$\\eta_t = \\eta \\left( \\frac{C_{D_w}}{C_{D_c}} \\right) = \\text{over-all efficiency.}$\n```", "timestamp": "2026-07-19T18:44:40.517428+00:00"}
{"citation_id": "19930094504", "source_url": "https://ntrs.nasa.gov/api/citations/19930094504/downloads/19930094504.pdf", "page_number": 4, "total_pages": 16, "image_filename": "19930094504_p4.jpg", "text": "2 N.A.C.A. Technical Memorandum No. 912\n\ntube welded at each end are: in the case of tensile load, the strength and the elastic limit of the material in the region annealed by the weld; in the case of compressive load, assuming sufficiently high slenderness ratio of the strut, the elasticity modulus corresponding to the buckling relation of Euler, this modulus being independent of the condition of the material whether welded or unwelded; in the case of lower slenderness ratio, the stretching or buckling limit of the welded tube.\n\nIt may be seen that with the exception of the case where only buckling stress in the Euler region occurs, the zone affected by the weld is always of decided significance for the dimensions and weight of the strut. The manner in which the steel tubes applied in airplane construction are affected as regards metallography and strength by the weld is known (references 1 and 2). An essential point brought out is that the tensile and compressive strength of butt-welded tubes corresponds to the strength of the unwelded, annealed tube, so that it assumes a minimum value which is characteristic of each type of steel. Cold or hot joining of the strut to be welded therefore, also outside the Euler range, is of no significance and the heat treatment of a tube cannot be utilized either for tensile or compressive stress unless the welded framework is treated as a whole, which treatment, however, on account of the size and deformation feared in hardening is generally impossible.\n\nThere was no disadvantage in this as long as for airplane construction only unalloyed steel tube was employed whose strength properties cannot be essentially increased over those of the welded state. This was also the result reached by A. Rochlich in his comprehensive investigations of 1930 (reference 1, p. 410), namely: \"that the weld up to the very small slenderness ratios is without effect also in the non-elastic range.\" The buckling values of welded unalloyed tubes for slenderness ratios greater than 15 lie as high as for those of the unwelded tubes.\n\nThe main results of the compression and buckling tests of Rochlich with two kinds of tube of various carbon content (0.1 to 0.15 percent and 0.3 to 0.5 percent) are shown in figure 8.\n\nSince the introduction of chrome-molybdenum steel tubing in the Albatros works at Berlin-Johannisthal, the object at Albatros and later at Focke-Wulf was consistently followed of utilizing the heat treatment possibilities of", "timestamp": "2026-07-19T18:44:50.668353+00:00"}
{"citation_id": "19930091655", "source_url": "https://ntrs.nasa.gov/api/citations/19930091655/downloads/19930091655.pdf", "page_number": 15, "total_pages": 22, "image_filename": "19930091655_p15.jpg", "text": "HEAT TRANSFER TO FUEL SPRAYS INJECTED INTO HEATED GASES 11\n\nTHE A-B INTERVAL\n\nThe A-B intervals correspond to the early part of the spray development for which the rate of heat transfer is essentially constant for a particular record. It is possible that this constancy is in some way associated with the approximately constant initial rate of spray-tip penetration (references 1 and 26). The magnitude of this interval is comparable with the ignition lag in compression-ignition engines. For this reason any conclusions based upon this interval are also applicable to such engines, provided that proper allowances are made for differences in chamber size and air temperature. The photographs shown in figure 8, together with more extensive penetration data (reference 1), show that this interval is essentially equivalent to the time (0.002 to 0.003 second) required by the sprays to traverse a distance of 4 inches, the approximate diameter of the bomb\n\nThis association of the moment of impingement with point B is supported by the fact that the interval decreases as the gas density decreases, i. e., as the penetration increases. (See table 1, column 7.) On the contrary, the interval is not appreciably shorter for the single-orifice nozzle in spite of the greater penetration to be expected with it. The period is about the same for carbon dioxide as for nitrogen in contradistinction to the longer A-C interval with carbon dioxide. Increasing the fuel quantity increases the injection period by a maximum factor of 3, yet the interval remains essentially the same. The interval also proved to be independent of the fuel used.\n\nInitial rate of heat transfer.—The magnitude of the initial rate of pressure drop, as shown by the particular pressure-time curve, is representative of the total rate of heat exchange between the gas and the fuel for the early part of the spray. The manner in which this initial rate is influenced by the temperature gradient between the gas and the fuel at two gas densities and several fuel weights is illustrated in figure 9. These initial slopes become more negative, i. e., the initial rate of heat transfer increases, as either the temperature difference or the fuel weight increases. Increasing the gas density decreases the numerical magnitude of the slope for a given fuel quantity but does not greatly alter the temperature dependence of the initial rate of pressure drop of the pressure-time curve: corresponding lines in figure 9 have roughly the same slope.\n\nThe increased density evidently decreases the effective transfer area in the early part of the spray as might be expected from the slower rate of spray development shown by the photographs reproduced in reference 1. The decrease cannot be attributed to a lower rate of heat transfer per unit area because the coefficient of heat conductivity should be nearly independent of density and the coefficient of heat transfer might be expected to increase with gas density (reference 27). Carbon dioxide gave rise to a greater rate of temperature drop than did nitrogen, even though its rate of pressure drop was smaller. This fact may be demonstrated by dividing the values of the initial slope in column 9 of tables I and III by their respective initial pressures, as outlined earlier in this paper. As the specific heats of nitrogen and carbon dioxide do not differ greatly on a weight basis, carbon dioxide must have given a greater initial rate of heat transfer. Since carbon dioxide has a lower coefficient of heat conductivity, it must give a greater effective heat-transfer area. The slopes for benzene (table III) are slightly greater than for Diesel fuel (table I) owing perhaps to a combination of the differences in the properties (molecular weight, specific heat, heat of vaporization, etc.) of the two fuels.\n\n[Figure: FIGURE 9.—Effect of gas temperature on initial rate of pressure drop at different gas-fuel ratios.]\n\nInitial temperature difference between gas and fuel, °C.\n| | 151 | 201 | 251 | 301 | 101 | 151 | 201 | 251 | 301 | 351 |\n|---|---|---|---|---|---|---|---|---|---|---|\n| Gas density, 4.73 grams per liter | | | | | Gas density, 14.13 grams per liter | | | | | |\n| + ○ △ □ Observer A | | | | | + × △ Observer A | | | | | |\n| × △ □ \" B | | | | | × \" B | | | | | |\n| Fuel weight, gram | | | | | Fuel weight, gram | | | | | |\n| .142 | | | | | .142 | | | | | |\n| .189 | | | | | .284 | | | | | |\n| .284 | | | | | | | | | | |\n| .568 | | | | | | | | | | |\n\nInitial rate of pressure drop, atmospheres per second\n-40\n-80\n-120\n-160\n-200", "timestamp": "2026-07-19T18:44:52.969859+00:00"}
{"citation_id": "19930094544", "source_url": "https://ntrs.nasa.gov/api/citations/19930094544/downloads/19930094544.pdf", "page_number": 36, "total_pages": 43, "image_filename": "19930094544_p36.jpg", "text": "N.A.C.A. Technical Memorandum No. 872\nFigs. 22,23,24\n\nFigure 22.- LZ127 - Inside view\nof framework. The axial\ngirder may be seen between the\nupper lift gas cells and the lower\nfuel gas cells. Below is seen the\ngangway girder.\n\nFigure 23.- AKRON-\nFraming\nwith tip of stern\nsuspended beside\nit. The inherently\nstiff, three boom,\nmain rings with\ntheir zig-zag strut\nbracing are easily\nvisible. The framing\nof the AKRON was\nassembled on \"framing\ntowers\". Two of these\nare placed under each\nmain ring.\n\nFigure 24.-AKRON-\nMain\nring lying down\nwith resilient\nbulkhead netting.\nThe casings\nattached to the\ncorners of the\ninner ring mem-\nber in the upper\npart of the ring\ncontain the\nresiliency devices.\nAt the left is\nseen the junction\nof the side\ncorridor with\nthe main ring.", "timestamp": "2026-07-19T18:45:01.204025+00:00"}
{"citation_id": "19930091719", "source_url": "https://ntrs.nasa.gov/api/citations/19930091719/downloads/19930091719.pdf", "page_number": 8, "total_pages": 33, "image_filename": "19930091719_p8.jpg", "text": "4 REPORT NO. 642—NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nPropeller 1C1-0 was included in the series because it differed from 6101 only in the shank shape and, incidentally, in the hub design.\n\nPropellers 6101, 6129, and 6131 constitute a series differing only in blade section. These propellers were whirl-tested (reference 6) and flight-tested at Wright Field previous to the present investigation.\n\nThe method of testing in the propeller-research tunnel consists in maintaining the propeller speed constant and increasing the tunnel speed in steps up to the maximum value of 115 miles per hour. Higher values engine power; the following schedule was therefore adhered to:\n\nPropeller speeds for tunnel speeds below 115 miles per hour\n\n| Blade angle, degrees: | Initial propeller speed, r. p. m. |\n| :--- | :--- |\n| 15. | 1,000 |\n| 20. | 1,000 |\n| 25. | 800 |\n| 30. | 800 |\n| 35. | 800 |\n| 40. | 700 |\n| 45. | 700 |\n\n[Figure: Graph showing Blade-form curves for propeller 3868-9. Axes: $b/D$, $h/b$, $p/D$, $r/R$. Curves labeled: Blade angle 45° at 0.75R, 35°, 25°, 15°.]\n\nFIGURE 7.—Blade-form curves for propeller 3868-9.\n\n[Figure: Graph showing Blade-form curves for propellers 6101, 6129, 6131, and 1C1-0. Axes: $b/D$, $h/b$, $p/D$, $r/R$. Curves labeled: Blade angle 45° at 0.75R, 35°, 25°, 15°, Propeller 1C1-0.]\n\nFIGURE 8.—Blade-form curves for propellers 6101, 6129, 6131, and 1C1-0.\n\nof $V/nD$ are obtained by reducing the engine speed until zero thrust is reached.\n\nThe tests reported in reference 1 showed that losses in efficiency occurred at tip speeds above 600 to 800 feet per second, depending principally on the blade angle and the $V/nD$ range. At slightly lower tip speeds the values of the thrust and the power coefficients, but not the efficiencies, were affected by compressibility. The present tests were therefore run at tip speeds of 525 feet per second and less to avoid complications arising from compressibility. The standard initial testing propeller speed of 1,000 r. p. m. could not be maintained for the higher blade-angle settings owing to the limitation of The approximate test propeller speed may be computed from the relation r. p. m.=$\\frac{K}{V/nD}$, for $V/nD$ values higher than can be obtained from the foregoing schedule, where $K$=1,000 for $V$=115 miles per hour and $D$=10 feet. The tests reported in reference 1 were confined to tip speeds above about 600 feet per second, so the use of the data in this reference for correcting coefficients for normal-flight operating speeds would necessitate neglecting any effects occurring at lower speeds. Unreported data obtained during these tests indicate that this procedure would entail little, if any, error.", "timestamp": "2026-07-19T18:45:08.225477+00:00"}
{"citation_id": "19930091693", "source_url": "https://ntrs.nasa.gov/api/citations/19930091693/downloads/19930091693.pdf", "page_number": 13, "total_pages": 13, "image_filename": "19930091693_p13.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-19T18:45:15.066950+00:00"}
{"citation_id": "19930094506", "source_url": "https://ntrs.nasa.gov/api/citations/19930094506/downloads/19930094506.pdf", "page_number": 2, "total_pages": 24, "image_filename": "19930094506_p2.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-19T18:45:18.461846+00:00"}
{"citation_id": "19930094499", "source_url": "https://ntrs.nasa.gov/api/citations/19930094499/downloads/19930094499.pdf", "page_number": 5, "total_pages": 13, "image_filename": "19930094499_p5.jpg", "text": "4 N.A.C.A. Technical Memorandum No. 917\n\nwhen the narrowest section of V is changed, with the provision that the section of the entrance cone is greater than the narrowest section of nozzle V.\n\nThe speed w and the Mach number $\\frac{w}{a}$, respectively, in the free jet follow, according to Bernoulli, from the pressure ratio $\\frac{p_0}{p}$. The pressure p is tapped through a wall orifice shortly before the exit from the entrance cone, and the tank pressure $p_0$, which, of course, must not be confused with the pressure in container B, corresponds to the atmospheric air pressure. Then Bernoulli's equation for stationary compressible flow without substantial local height differences reads:\n\n$$\n\\frac{w^2}{2} + \\int \\frac{d\\,p}{\\rho} = \\text{const} \\tag{1}\n$$\n\nSince adiabatic change of state may be assumed ($\\kappa = 1.405$ for air) in the flow through the entrance cone, hence\n\n$$\n\\frac{1}{\\rho} = \\frac{1}{\\rho_0} \\left( \\frac{p_0}{p} \\right)^{1/\\kappa} \\tag{2}\n$$\n\nthe integral in equation (1) becomes\n\n$$\n\\int \\frac{d\\,p}{\\rho} = \\frac{\\kappa}{\\kappa - 1} \\frac{p_0}{\\rho_0} p^{1/\\kappa} p^{\\frac{\\kappa - 1}{\\kappa}} = \\frac{\\kappa}{\\kappa - 1} \\frac{p}{\\rho} \\tag{3}\n$$\n\nFor the state of rest p is replaced by $p_0$, $\\rho$ by $\\rho_0$, and w by $w_0 = 0$, whence the constant of the Bernoulli equation\n\n$$\n\\text{const} = \\frac{\\kappa}{\\kappa - 1} \\frac{p_0}{\\rho_0} = \\frac{\\kappa}{\\kappa - 1} \\frac{p}{\\rho} \\left( \\frac{p_0}{p} \\right)^{\\frac{\\kappa - 1}{\\kappa}} \\tag{4}\n$$\n\nEntering equations (3) and (4) in equation (1) gives the flow velocity in the free jet at\n\n$$\nw = \\sqrt{ \\frac{2\\,\\kappa}{\\kappa - 1} \\frac{p}{\\rho} \\left[ \\left( \\frac{p_0}{p} \\right)^{\\frac{\\kappa - 1}{\\kappa}} - 1 \\right] } \\tag{5}\n$$", "timestamp": "2026-07-19T18:45:19.234349+00:00"}
{"citation_id": "19930094549", "source_url": "https://ntrs.nasa.gov/api/citations/19930094549/downloads/19930094549.pdf", "page_number": 28, "total_pages": 76, "image_filename": "19930094549_p28.jpg", "text": "26 N.A.C.A. Technical Memorandum No. 867\n\nThe final state of equilibrium, however, defined by the position of the control, has become state A, and the airplane deviates from it by $\\delta V = + 7.7$; $\\delta i = - 3^\\circ$; $\\delta \\theta = 17.4^\\circ$; $\\delta q = + 1^\\circ$ per second. Taking these values as an initial disturbance with respect to state A, the changes in the variables may be easily determined. The results of the computation are indicated as discontinuous lines on figure 14. The computations have been made neglecting the effect of the disturbance $\\delta q$, which is less important than the three preceding ones.\n\nRemarks\n\n1. The curve of the angle of attack drawn as a thin dot-dash line corresponds to the assumption of a more gradual maneuvering of the elevator.\n\n2. We have calculated the curve of $J_z$, the apparent weight or acceleration, for the maneuver described - consisting of nosing down the airplane for 7 seconds, then nosing up. The result of the computation is given on figure 16.\n\nTests\n\nIt was possible for us to obtain records of the variables defining the longitudinal motion of an airplane. The test was carried out on a Fairey \"Fox\" with the Bouny equipment, which will be described in detail in a succeeding Bulletin. Records of \"phugoid\" oscillations are sufficiently numerous. Nevertheless, we believe it useful to publish here the results of the measurements we have made.\n\nAlthough this test has been conducted on an airplane differing from the one employed in our calculations, the curves obtained show a qualitative correspondence with the calculated curves that appears to justify comparison. The polar and the curve of moments were known from a tunnel test conducted under the usual conditions on a model not provided with a propeller. The airplane itself had a loading of 58.5 kg/m². At the speed of 45.2 m/s, it flew at a lift coefficient of 0.458, corresponding to an angle of attack of 3.6°. The coefficient of mean static stability measured in the interval $i = + 6$ to $i = - 4$ may be determined from figure 17,\n\n$$\n\\frac{dC_M}{di} = 0.00417\n$$", "timestamp": "2026-07-19T18:45:19.647244+00:00"}
{"citation_id": "19930091731", "source_url": "https://ntrs.nasa.gov/api/citations/19930091731/downloads/19930091731.pdf", "page_number": 4, "total_pages": 32, "image_filename": "19930091731_p4.jpg", "text": "# NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nHEADQUARTERS, NAVY BUILDING, WASHINGTON, D. C.  \nLABORATORIES, LANGLEY FIELD, VA.\n\nCreated by act of Congress approved March 3, 1915, for the supervision and direction of the scientific study of the problems of flight (U. S. Code, Title 50, Sec. 151). Its membership was increased to 15 by act approved March 2, 1929. The members are appointed by the President, and serve as such without compensation.\n\nJOSEPH S. AMES, Ph. D., Chairman,  \nBaltimore, Md.\n\nVANNEVAR BUSH, Sc. D., Vice Chairman,  \nWashington, D. C.\n\nCHARLES G. ABBOT, Sc. D.,  \nSecretary, Smithsonian Institution.\n\nHENRY H. ARNOLD, Major General, United States Army,  \nChief of Air Corps, War Department.\n\nGEORGE H. BRETT, Brigadier General, United States Army,  \nChief Matériel Division, Air Corps, Wright Field, Dayton, Ohio.\n\nLYMAN J. BRIGGS, Ph. D.,  \nDirector, National Bureau of Standards.\n\nCLINTON M. HESTER, A. B., LL. B.,  \nAdministrator, Civil Aeronautics Authority.\n\nROBERT H. HINCKLEY, A. B.,  \nChairman, Civil Aeronautics Authority.\n\nJEROME C. HUNSAKER, Sc. D.,  \nCambridge, Mass.\n\nSYDNEY M. KRAUS, Captain, United States Navy,  \nBureau of Aeronautics, Navy Department.\n\nCHARLES A. LINDBERGH, LL. D.,  \nNew York City.\n\nFRANCIS W. REICHELDERFER, A. B.,  \nChief, United States Weather Bureau.\n\nJOHN H. TOWERS, Rear Admiral, United States Navy,  \nChief, Bureau of Aeronautics, Navy Department.\n\nEDWARD WARNER, Sc. D.,  \nGreenwich, Conn.\n\nORVILLE WRIGHT, Sc. D.,  \nDayton, Ohio.\n\nGEORGE W. LEWIS, Director of Aeronautical Research\n\nJOHN F. VICTORY, Secretary\n\nHENRY J. E. REID, Engineer-in-Charge, Langley Memorial Aeronautical Laboratory, Langley Field, Va.\n\nJOHN J. IDE, Technical Assistant in Europe, Paris, France\n\n## TECHNICAL COMMITTEES\n\nAERODYNAMICS  \nPOWER PLANTS FOR AIRCRAFT  \nAIRCRAFT MATERIALS  \n\nAIRCRAFT STRUCTURES  \nAIRCRAFT ACCIDENTS  \nINVENTIONS AND DESIGNS  \n\nCoordination of Research Needs of Military and Civil Aviation  \nPreparation of Research Programs  \nAllocation of Problems  \nPrevention of Duplication  \nConsideration of Inventions  \n\n## LANGLEY MEMORIAL AERONAUTICAL LABORATORY  \nLANGLEY FIELD, VA.\n\nUnified conduct, for all agencies, of scientific research on the fundamental problems of flight.\n\n## OFFICE OF AERONAUTICAL INTELLIGENCE  \nWASHINGTON, D. C.\n\nCollection, classification, compilation, and dissemination of scientific and technical information on aeronautics.", "timestamp": "2026-07-19T18:45:21.103316+00:00"}
{"citation_id": "19930091701", "source_url": "https://ntrs.nasa.gov/api/citations/19930091701/downloads/19930091701.pdf", "page_number": 14, "total_pages": 18, "image_filename": "19930091701_p14.jpg", "text": "10 REPORT NO. 626—NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nload the distance to clear the 50-foot obstacle is reduced 19 percent by the over-all effect of wind and 9 percent by reason of the wind gradient alone. For a 100-foot obstacle the reductions are 21 percent and 10 percent, respectively.\n\nThe method used in this report, i. e., step-by-step integration, would be too laborious for general use in evaluating the corrections for wind; but it has been found that these corrections can be determined with sufficient accuracy through the aid of rather simple relations: Still regarding the effects of wind velocity and velocity gradient as separate, the correction to the air-borne distance for the effect of wind velocity is\n\n$$\n\\Delta D_1 = \\int_0^T V_w dt\n$$\n\nwhere $V_w$ is the wind velocity at any time $t$ and $T$ is the time required, from the instant of leaving the ground, to attain the height $H$.\n\nFor the average wind gradient, previously defined, the correction becomes, for $H=50$ feet,\n\n$$\n\\Delta D_1 = 1.27 V_{w_0} T\n$$\n\nwhere $V_{w_0}$ is the wind velocity at the ground. For $H=100$ feet\n\n$$\n\\Delta D_1 = 1.38 V_{w_0} T\n$$\n\nThe effect of a wind-velocity gradient on the height attained in a given time $T$ is found from the energy relations to be\n\n$$\n\\Delta H = \\frac{V \\cos \\gamma \\Delta V_w}{g}\n$$\n\nwhere $\\Delta H$ is the difference between the heights attained with and without the benefit of a wind gradient, in the same period of time, which is very nearly equivalent\n\nheight $H_e$. The correction to the observed air-borne distance for the wind gradient is then\n\n$$\n\\Delta D_2 = \\frac{\\Delta H}{\\tan \\gamma_1}\n$$\n\nwhere $\\gamma_1$ is the angle of climb that would obtain were there no wind gradient; it is given closely enough by\n\n$$\n\\gamma_1 = \\gamma - \\tan^{-1} \\frac{dV_w}{dH} \\times \\frac{V \\sin \\gamma}{g}\n$$\n\nFor the average wind gradient\n\n$$\n\\Delta V_w = 0.41 V_{w_0} \\text{ at a height of 50 feet,}\n$$\n\nand\n\n$$\n\\Delta V_w = 0.55 V_{w_0} \\text{ at a height of 100 feet.}\n$$\n\nThe angle of climb for no wind is given, for an obstacle height of 50 feet, by\n\n$$\n\\gamma_1 = \\gamma - \\tan^{-1} 0.0037 V_{w_0} \\times \\frac{V \\sin \\gamma}{g}\n$$\n\nand, for an obstacle height of 100 feet, by\n\n$$\n\\gamma_1 = \\gamma - \\tan^{-1} 0.0021 V_{w_0} \\times \\frac{V \\sin \\gamma}{g}\n$$\n\nThe over-all correction to no-wind conditions is then, for an obstacle height of 50 feet,\n\n$$\n\\Delta D = 1.27 V_{w_0} T + \\frac{V \\cos \\gamma \\times 0.41 V_{w_0}}{g \\tan \\left( \\gamma - \\tan^{-1} 0.0037 \\frac{V_{w_0} V \\sin \\gamma}{g} \\right)}\n$$\n\nand, for an obstacle height of 100 feet,\n\n$$\n\\Delta D = 1.38 V_{w_0} T + \\frac{V \\cos \\gamma \\times 0.55 V_{w_0}}{g \\tan \\left( \\gamma - \\tan^{-1} 0.0021 \\frac{V_{w_0} V \\sin \\gamma}{g} \\right)}\n$$\n\nThe corrections as computed from the foregoing equations agreed closely with those determined by the step-by-step integrations, the difference being less than 2 percent of the air-borne distance in all the cases considered. In the absence of specific data on the variation of the wind velocity with altitude, it is believed that the assumption of an average wind gradient will provide a good approximation.\n\nThe effect of proximity of the ground on the distance required for the air-borne stages of the take-off is shown in figure 14. The ground effect reduces the distance required to attain an altitude of 50 feet by 10 percent with the lightest load and by 16 percent with the heaviest load. For an obstacle height of 100 feet, the percentage reductions are about one-half of those for the 50-foot obstacle. The greater difference for the heavier load is probably due to the fact that the airplane climbs more slowly than with the light load; hence it is in the region of strongest ground effect for a longer period.\n\n[Figure: Graph showing Vertical distance, ft. vs. Horizontal distance, ft. with curves for Weight = 2,060 lb. and Weight = 2,600 lb., and lines for No ground effect and With ground effect]\n\nFIGURE 14.—Effect of ground proximity on the flight path of the Verville AT airplane during transition and steady climb.\n\nto the same horizontal distance; $V$ is the air speed at the height $H$; $\\gamma$ is the flight-path angle relative to the air at the height $H$; $\\Delta V_w$ is the difference between the wind speed at 5 feet from the ground and at the effective", "timestamp": "2026-07-19T18:45:24.374687+00:00"}
{"citation_id": "19930091724", "source_url": "https://ntrs.nasa.gov/api/citations/19930091724/downloads/19930091724.pdf", "page_number": 7, "total_pages": 20, "image_filename": "19930091724_p7.jpg", "text": "```markdown\n\"PACK\" COMPRESSIVE TEST\n3\n\nwith the seating of the strain gages, their width $h'$ was made 0.02 to 0.05 inch less than the middle specimen.\n\nIn practice the specimens were usually slightly warped, bowed, or irregular on the surface. The effect of these deviations from a plane surface was minimized by assembling the supporting specimens, whenever possible, so that they bowed towards the middle specimen. The number of supporting specimens was kept as small as possible consistent with obtaining sufficient stability with the transverse support employed. This was done to limit the sample from the piece, so that specimens would be taken from like material and to obtain packs where only a small amount of material was available. This, also, reduced the cost of machining.\n\nThe specimens were sawed in a milling machine about $\\frac{1}{32}$ inch longer than the finished length. The ends E were then finished in a surface grinder using a Norton alundum grinding wheel, number 1936:G and kerosene as a lubricant. No attempt was made, however, to make them more than nominally parallel and perpendicular to the axis of the pack. The edges were neither rounded nor marred, appreciably, when the burrs were removed.\n\n### MACHINING PROCEDURE\n\nThe specimens were finished to width using a series of light cuts in order that the underlying material\n\n### TRANSVERSE SUPPORT\n\nTransverse support was supplied as shown in fig. 4 by thirty steel pins A on each side of the pack. The pins were in three columns and ten rows. They were $\\frac{3}{8}$ inch in diameter and about two inches long. One end G was hardened and ground to a conical point. The other end H was machined to a hemisphere. The pointed end rested the external sheet face of the pack. The hemispherical end rested in a conical seat in the end of a size 8 machine screw Y, one inch long. The screws were threaded through the webs of two pieces of three inch structural steel channel R and were spaced on $\\frac{1}{2}$ inch centers.\n\nThe channels were bolted at the bottom to the rectangular steel block B and were prevented from spreading at the top by a heavy yoke clamp, not shown in the figure.\n\n### TEST PROCEDURE\n\n#### TESTING MACHINE\n\nThe packs were tested in a vertical, fluid-support, Bourdon-tube hydraulic type of testing machine of 100 kips capacity, using the 10 kip dial and the 50 kip dial to indicate the load. The testing machine is shown in fig. 5.\n\n#### BEARING BLOCKS\n\nThe surfaces of bearing blocks which transfer the load from the heads of the testing machine to the pack were flat. They were inspected frequently for dirt or mars. A paper shim D, shown in fig. 4, was used between the $5\\frac{1}{8}$ by 3 by $1\\frac{3}{4}$ inch block E and the surface of the lower head of the testing machine. The bearing block F was a disk of hardened steel $1\\frac{15}{16}$ inches in diameter and $1\\frac{3}{16}$ inch thick with top and bottom surfaces smooth-ground.\n\nThe bearing block U was attached to the upper head of the testing machine through the $8\\frac{1}{8}$ by $5\\frac{7}{8}$ by $1\\frac{3}{16}$ inch plate I. The upper contact surface of the bearing block was $\\frac{3}{4}$ inches in diameter and the lower contact surface was $1\\frac{15}{16}$ inches in diameter.\n\nSlight deviations from parallelism of the bearing blocks, the heads of the testing machine, and the ends of the pack which are within the limits of good machine shop practice, may appreciably affect the results of compressive tests. To eliminate these effects and to equalize the load on the specimen a cap-block V and a plaster of paris shim N were used, as shown in fig. 4.\n\n<!-- Image (112, 369, 442, 648) -->\nFIGURE 3.—Machining jig.\n\nwould be disturbed as little as possible. The lateral edges were finished smooth and the burrs were removed.\n\nThe specimens were finished to length after they were assembled in the pack. The machining jig, shown in fig. 3, was used to hold the specimens while they were being machined. This jig was a small vise. The contact surfaces of the jaws were $1\\frac{13}{16}$ inches long and $\\frac{1}{2}$ inch wide. They were plane and smooth. The movable jaw J pivoted at the end of the screw Q. The surfaces of the body of the vise were planes parallel and perpendicular to the stationary jaw K so that the pack could be readily aligned with the machine tool. A small clamp C was attached at each end of the pack before the ends were machined to hold the specimens in position after the pack was removed from the jig.\n```", "timestamp": "2026-07-19T18:45:33.563539+00:00"}
{"citation_id": "19930094493", "source_url": "https://ntrs.nasa.gov/api/citations/19930094493/downloads/19930094493.pdf", "page_number": 6, "total_pages": 31, "image_filename": "19930094493_p6.jpg", "text": "N.A.C.A. Technical Memorandum No. 923 5\n\n2. The Effect of the Humidity and the Direction of the Air Flow on the Heat Transfer from Thin Wires\n\nIt was shown by J. Ulsamer (reference 3) that the effect of the air humidity on the heat transfer from thin wires may be neglected, the error that arises from the neglect amounting to about ±2 percent, which lies within the limits of accuracy required of the function f.\n\nIn setting up the function f, the case was assumed where the wire is situated at right angles to the flow direction. It was found by J. Ulsamer (reference 3) that the heat transfer is lowered considerably with reduction in the angle α. For a ratio of the length to diameter of the wire equal to 400, the heat transfer with the wire parallel to the flow direction is three-fourths of the value for the wire at right angles. The transverse position is thus characterized by the maximum power absorption. In the present investigation, a still more favorable length to diameter ratio equal to 1,300 was employed.\n\n3. Objects of the Measurements\n\nAccording to Newton's law, the heat transmitted to the air stream is given by\n\n$$Q = \\alpha F (T_w - T_o) \\text{(kcal/h)} \\tag{11}$$\n\nwhere $F = \\pi d l \\text{ (m}^2\\text{)}$ and\n\n$$\\alpha = \\frac{\\lambda_m}{d} m \\text{ Re}^n$$\n\nTransforming equation (11) by means of equations (5), (6), (7), and (8), there is obtained\n\n$$Q = \\Phi (T_w, T_o, P_o, w) \\tag{12}$$\n\nWe have furthermore equation (4):\n\n$$U = 0.86 \\psi (i^2, T_w) \\tag{4}$$\n\nFor a condition of equilibrium, the electrical energy supplied U in each time interval must be equal to that conducted away Q:", "timestamp": "2026-07-19T18:45:40.184431+00:00"}