PDF/ocr_dataset_106.jsonl

{"citation_id": "19930094538", "source_url": "https://ntrs.nasa.gov/api/citations/19930094538/downloads/19930094538.pdf", "page_number": 42, "total_pages": 43, "image_filename": "19930094538_p42.jpg", "text": "N. A. C. A. Technical Memorandum No. 878\n\nFigure 22.- Failure of cylinder No. II in pure twist.\n(inside view)\n\nFigure 23.- Failure of cylinder No. II in pure twist.\n(outside view)\n\nFigure 24.- Failure of cylinder No. III in pure twist (inside view).\n\nFigure 25.- Failure of cylinder No. IV in pure twist (inside view).\n\nFile. 22,23,24,25", "timestamp": "2026-07-19T18:37:08.510626+00:00"}
{"citation_id": "19930091719", "source_url": "https://ntrs.nasa.gov/api/citations/19930091719/downloads/19930091719.pdf", "page_number": 3, "total_pages": 33, "image_filename": "19930091719_p3.jpg", "text": "REPORT No. 642\n\nTESTS OF FIVE FULL-SCALE PROPELLERS  \nIN THE PRESENCE OF A RADIAL AND A LIQUID-  \nCOOLED ENGINE NACELLE, INCLUDING TESTS  \nOF TWO SPINNERS\n\nBy DAVID BIERMANN and EDWIN P. HARTMAN  \nLangley Memorial Aeronautical Laboratory\n\n80799—38——1", "timestamp": "2026-07-19T18:37:11.512816+00:00"}
{"citation_id": "19930094535", "source_url": "https://ntrs.nasa.gov/api/citations/19930094535/downloads/19930094535.pdf", "page_number": 15, "total_pages": 17, "image_filename": "19930094535_p15.jpg", "text": "N.A.C.A. Technical Memorandum No. 881\n\nFigure 3.- Formation of bubbles in three-layer safety glass using cellulose acetate.\n\nFigure 6.- Sheet of polyacrylic acid ester after 100 flying hours.\n\nAl. strips\nAl. rivet\nFabric\nCover plate screwed on.\nFabric\n\nFigure 7.- Methods of attachment of transparent plastics.\n\nFigure 10.- Multilayer safety glass after the ball-dropping test at -21°.\n\nFigs. 3,6,7,10", "timestamp": "2026-07-19T18:37:12.732862+00:00"}
{"citation_id": "19930094559", "source_url": "https://ntrs.nasa.gov/api/citations/19930094559/downloads/19930094559.pdf", "page_number": 14, "total_pages": 16, "image_filename": "19930094559_p14.jpg", "text": "Technical Memorandum No. 857\nFigs.7,8\n\n[Figure: A sequence of images showing the effect of fuel quantity on combustion.]\n\nFigure 7.- Effect of fuel quantity on combustion, B=100 mg as compared with 50 mg of fig.5. Remaining conditions unchanged.\n\n[Figure: A sequence of images showing injection and combustion in still air.]\n\nFigure 8.- Injection and combustion in still air,w=0, p=21 atm. abs, t=550° C. Remaining conditions unchanged.", "timestamp": "2026-07-19T18:37:14.251842+00:00"}
{"citation_id": "19930094549", "source_url": "https://ntrs.nasa.gov/api/citations/19930094549/downloads/19930094549.pdf", "page_number": 21, "total_pages": 76, "image_filename": "19930094549_p21.jpg", "text": "N.A.C.A. Technical Memorandum No. 367 19\n\nASCENDING GUST\n\nLet an ascending gust, characterized by $u'_2 = 0$  \n$w'_2 = 8$ m/s act suddenly on an airplane in horizontal flight. The state $R_1$ in the atmosphere $A_1$ assumed to be undisturbed, is characterized by a system of given values $u_1$, $w_1$, $\\theta_1$ (fig. 7). The state $R_1$ in the atmosphere $A_2$ is a transitory state and not one of equilibrium. The angle of attack is very large (fig. 8) and the flight path of the airplane will be modified. The airplane will try to adjust itself to a new state of equilibrium $R_2$ in the atmosphere $A_2$. It is easy to see that this state will be characterized by the same values of the relative velocity, angle of attack, and inclination to the horizontal as state $R_1$, but the absolute velocity $U_2$ will be different. The latter becomes ascending, the airplane being subject to the motion of the surrounding medium (fig. 9).\n\nWe may study this transitory period by seeking the effect of an initial disturbance $\\delta w = -8$ m/s (fig. 10). The airplane will be lifted suddenly and if it is statically stable, it will start a diving motion. These two motions have the effect of tending to close the angle included between the directions OX and V of the figure, each of these axes approaching the other.\n\nIf the airplane has great static stability, it will oscillate more rapidly and the angle of attack of the state will be regained by means of a displacement $\\delta \\theta$ larger than would be the case if the airplane had been less stable. The static stability increases the disturbance in the attitude $\\delta \\theta$ due to the rapid oscillation produced by a vertical gust. The axis of the airplane should, however, in its final state, regain its initial attitude $\\theta$. A large static stability therefore results in momentarily removing the airplane from its final attitude and increasing the nonequilibrium of the forces acting on the moving center of gravity. This explains why the amplitude of the long-period oscillations will be greater on a very stable airplane than on a less stable one. A neutral airplane will be lifted without its axis OX undergoing any change in attitude $\\delta \\theta$. It is the $U_1$ axis which will be effective in destroying the disturbance of the angle of attack $\\delta i$. The case of an unstable airplane is sufficiently explained by the figure and requires no remarks.", "timestamp": "2026-07-19T18:37:20.631858+00:00"}
{"citation_id": "19930094542", "source_url": "https://ntrs.nasa.gov/api/citations/19930094542/downloads/19930094542.pdf", "page_number": 83, "total_pages": 102, "image_filename": "19930094542_p83.jpg", "text": "N.A.C.A. Technical Memorandum No. 874\nFigs.69,70\n\n$$\\frac{dc_m}{dc_a}$$\n\n[Figure: Graph plotting $\\frac{dc_m}{dc_a}$ against $\\lambda$. The y-axis ranges from .19 to .27. The x-axis ranges from 0 to .8. A curve labeled \"Wing alone ($\\lambda=\\infty$)\" is plotted with circular data points.]\n\nFigure 69.- Stability coefficient $\\frac{dc_m}{dc_a}$ of wing in presence of propeller.\n\n$$\\frac{dc_m}{d\\lambda}$$\n\n[Figure: Graph plotting $\\frac{dc_m}{d\\lambda}$ against $\\lambda$. The y-axis ranges from 0 to 1.2. The x-axis ranges from 0 to .6. Four curves are plotted with different symbols corresponding to a legend:\n- $\\circ$ $\\kappa = 9^\\circ$\n- $\\times$ $\\kappa = 4^\\circ$\n- $+$ $\\kappa = -1^\\circ$\n- $\\triangle$ $\\kappa = -6^\\circ$]\n\nFigure 70.- Variation of $\\frac{dc_m}{d\\lambda}$ with $\\lambda$ and $\\kappa$.", "timestamp": "2026-07-19T18:37:22.085067+00:00"}
{"citation_id": "19930094568", "source_url": "https://ntrs.nasa.gov/api/citations/19930094568/downloads/19930094568.pdf", "page_number": 3, "total_pages": 28, "image_filename": "19930094568_p3.jpg", "text": "2\nN.A.C.A. Technical Memorandum No. 848\n\n| Planing surface | Airfoil | |\n| :--- | :--- | :--- |\n| | $l_p$ | distance of lift resultant from trailing edge of plate. |\n| $t = l \\beta$ | | depth of immersion. |\n| $\\beta$ | | angle of attack of plate. |\n| $\\beta_w$ | | \"effective\" angle of attack = $\\beta$ minus angle of downwash $\\beta_1$. |\n| $\\Delta \\beta_w$ | | supplementary angle of attack due to boundary-layer friction at planing surface. |\n| $\\beta_1 = \\frac{4R}{\\pi \\rho V^2 b^2}$ | $\\beta_1' = \\frac{2R'}{\\pi \\rho V^2 b}$ | angle of downwash due to infinite width of surface in motion free from gravity. |\n| $F = \\frac{V}{\\sqrt{bg}}$ | | Froude number. |\n| $R = \\frac{V l}{\\nu}$ | | Reynolds Number. |\n| $\\frac{R}{\\frac{\\rho}{2} V^2 b^2}$ | | load factor. |\n| $G_V$ | | forward plate loading. |\n| $G_H$ | | rear plate loading. |\n| $B$ | $B'$ | momentum of downward moving fluid mass. |\n| $m_0 = \\frac{\\rho}{2} \\frac{\\pi b^2}{4}$ | $m_0' = \\rho \\frac{\\pi b^2}{4}$ | mass reduced to $V_1$ moving downward per length \"one\" at infinitely small angle of attack. |\n| $m$ | $m'$ | mass reduced to $V_1$ moving downward per length \"one\" at finite angle of attack. |", "timestamp": "2026-07-19T18:37:24.670461+00:00"}
{"citation_id": "19930094544", "source_url": "https://ntrs.nasa.gov/api/citations/19930094544/downloads/19930094544.pdf", "page_number": 31, "total_pages": 43, "image_filename": "19930094544_p31.jpg", "text": "N.A.C.A. Technical Memorandum No. 872 29\n\nTABLE I\n\nStrength Data for the Newer Aluminum Alloys\n\n| Alloy and hardness | Yield point $\\sigma_{0.2}$ kg/mm² | Tensile strength $\\sigma_B$ kg/mm² | Elongation $\\delta$ (percent) | Remarks |\n| :--- | :--- | :--- | :--- | :--- |\n| 681 B, untreated hardness 1 | 26-28<br>32-34 | 38-42<br>45-48 | 18-15<br>12-10 | According to data of the Dürener Metallwerke A.-G., Düren |\n| 681 ZB, untreated hardness 1 | 28-30<br>36-38 | 42-44<br>46-48 | 18-15<br>12-10 | |\n| DM 31, untreated hardness 1 | 30-34<br>40-42 | 46-48<br>50-52 | 15-12<br>12-10 | |\n| 17SRT, average | 32 | 43 | 10-15* | According to American data |\n\n*Measured over 2 inches.", "timestamp": "2026-07-19T18:37:53.960055+00:00"}
{"citation_id": "19930094535", "source_url": "https://ntrs.nasa.gov/api/citations/19930094535/downloads/19930094535.pdf", "page_number": 16, "total_pages": 17, "image_filename": "19930094535_p16.jpg", "text": "Starting materials\ncotton, cellulose\n\nNitric +\nsulphuric acid\n\nAcetic acid anhydride\n+ catalizer\n\nPrimary cellulose\nacetate.\n\nSaponification\n\nSecondary cellulose\nacetate.\n\nCellulose\nnitrate\n\nFigure 4.- Production of cellulose acetate\nand cellulose nitrate.\n\nWater\n\nCarbon\n\nChalk\n\nCarbide\n\nAcetylene\n\nVinyl-\nchloride\n\nPolyvinyl-\nchloride\n\nFigure 5.- Production of polyvinylchloride.\n\nkg/cm²\n600\n500\n400\nσ\n300\n200\n100\n0\n-40 -20 0 20 40 °C 60\nTemperature\n\nFigure 8.- Tensile strength of cellulose\nacetates as a function of the\ntest temperature.\n\nW.A.C.A. Technical Memorandum No. 831\n\nFigs. 4,5,8", "timestamp": "2026-07-19T18:37:54.162402+00:00"}
{"citation_id": "19930094542", "source_url": "https://ntrs.nasa.gov/api/citations/19930094542/downloads/19930094542.pdf", "page_number": 84, "total_pages": 102, "image_filename": "19930094542_p84.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-19T18:37:58.264737+00:00"}
{"citation_id": "19930094538", "source_url": "https://ntrs.nasa.gov/api/citations/19930094538/downloads/19930094538.pdf", "page_number": 43, "total_pages": 43, "image_filename": "19930094538_p43.jpg", "text": "N. A. C. A. Technical Memorandum No. 878\n\nFigs. 26,27,28\n\n[Figure: Diagram showing a loading system with labeled parts A, B, P, b, V, l, and section A-B]\n\nFigure 26.- Loading system and test arrangement for buckling-bending test with steel panels.\n\n[Figure: Graph plotting Experimental failing load $P_s$ vs. Load ratio $\\lambda - \\lambda_0/\\lambda$]\n\nClosed section (of cylinder III)\n\nOpen sections (shell panels) and individual sections (of cylinder IV)\n\nPanels with cuts (of cylinder III)\n\nFigure 27.- Stringer failing loads $P_s$ against load ratio $\\lambda$.\n\n[Figure: Graph plotting Axial load on a section $P_0$ vs. Load ratio $\\lambda$]\n\nCylinder IV\n\nCylinder III\n\nCylinder IV Test\n\nCylinder III Test\n\nTheory\n\nTheory\n\nFigure 28.- Load ratio $\\lambda$, axial load $P_0$ and stringer capacity $P_{LS}$ plotted against $\\psi = T/T_0$.", "timestamp": "2026-07-19T18:38:02.480769+00:00"}
{"citation_id": "19930094549", "source_url": "https://ntrs.nasa.gov/api/citations/19930094549/downloads/19930094549.pdf", "page_number": 22, "total_pages": 76, "image_filename": "19930094549_p22.jpg", "text": "20\nN.A.C.A. Technical Memorandum No. 867\n\nHORIZONTAL GUST\n\nBy the same procedure we have calculated the effect\nof a horizontal gust of 10 m/s (22 m.p.h.) acting in the\nopposite direction to the motion of the airplane. The\ninitial disturbance is an increase $\\delta u$. After the effect\nof this disturbance has disappeared the airplane regains\nits horizontal flight at the original angle of attack but\nat a different altitude. The airplane undergoes at first\na vertical acceleration. Its angle of attack therefore\ndecreases, and if it is stable it noses up and then starts\nthe motions characteristic of the long-period oscillation.\nFigure 11 shows the action on the four stable airplanes\nconsidered, on a neutral airplane, and on an unstable air-\nplane. The disturbance of the horizontal velocity gives\nrise to disturbances in the attitude associated with the\nslow oscillation and their effects in certain cases be-\ncome far from negligible.\n\nDISTURBANCE OF THE FLIGHT PATH\n\nWe shall consider the case of a disturbance $\\delta \\theta$ act-\ning alone - the velocity V, angle of attack i, and an-\ngular velocity q not being disturbed. Such a disturb-\nance occurs under special circumstances. It would result,\nfor example, from the same angular displacement of the air-\nplane and its flight path. The case where an initial state\nand final state are characterized by sensibly the same an-\ngle of attack and same velocity V, but where the paths\nand the orientation of the OX axis differ by the same\nquantity, is one that may occur in practice, namely, that\nof the sudden failure of the engine. Denote the propel-\nler advance diameter ratio V/nD by $\\gamma$. On some air-\nplanes, when\n\n$$s \\frac{dT}{dV} + \\frac{\\partial M}{\\partial V} = 0$$\n\nthe sudden failure of the engine does not change the mo-\nment about the center of gravity. When this is true the\nairplane, for the same deflection of the elevator, will be\nin equilibrium with its engine cut off and on a descend-\ning flight path over which it will travel with the same\nvelocity and at the same angle of attack as on the hori-\nzontal path with the engine running.", "timestamp": "2026-07-19T18:38:09.381807+00:00"}
{"citation_id": "19930094559", "source_url": "https://ntrs.nasa.gov/api/citations/19930094559/downloads/19930094559.pdf", "page_number": 15, "total_pages": 16, "image_filename": "19930094559_p15.jpg", "text": "N.A.C.A. Technical Memorandum No. 857\nFigs.9,10\n\n<!-- Image (388, 159, 835, 468) -->\n\nStart of injection referred\nto compressor crank angle.\nFigure 9.- Ignition lag time $z_v$, time of combustion\n$z_{Br}$, and total time $z_{tot}$ as functions of\nstart of injection referred to compressor crank angle.\nFuel: Derop gasoil. Bosch pump PE 1B 100/100, n=400\nr.p.m. Bosch nozzle DN 12 SD 12, valve opening\npressure 85 atm.\n\n<!-- Image (390, 612, 859, 828) -->\n\nFigure 10.- Time of combustion as a function of fuel\nquantity B for two different Bosch nozzles.\nStart of injection at 350° crank angle of compressor.\nw= 67 m/s. (220 ft/sec.), p = 17.5 atm. abs, t = 780° C.", "timestamp": "2026-07-19T18:38:10.129764+00:00"}
{"citation_id": "19930093641", "source_url": "https://ntrs.nasa.gov/api/citations/19930093641/downloads/19930093641.pdf", "page_number": 25, "total_pages": 47, "image_filename": "19930093641_p25.jpg", "text": "21\n\nFigure 24.- Tuft surveys for the wing alone without fuselage or nacelles. $\\delta_f$, $0^\\circ$; approximate test air speed, 50 m.p.h.\n\nFigure 25.- Tuft surveys for the conventional wing-nacelle model. $\\delta_a$, $0^\\circ$; $\\delta_f$, $0^\\circ$; approximate test air speed, 50 m.p.h.\n\nFigure 26.- Comparison of the propulsive efficiencies of five test arrangements at a lift coefficient corresponding to high speed, $C_L = 0.25$. $\\beta$, $18\\frac{1}{2}^\\circ$.\n\nFigure 27.- Comparison of the propulsive efficiencies of five test arrangements at a lift coefficient corresponding to best climb, $C_L = 0.70$. $\\beta$, $18\\frac{1}{2}^\\circ$.\n\nFigure 28.- Propulsive efficiencies of wing-nacelle arrangement for four different blade angles.\n\nFigure 29.- Variation of propulsive efficiency with blade angle for propellers in tractor position 2.\n\nFigure 30.- Variation of propulsive efficiency with blade angle for propellers in tractor position 3.\n\nFigure 31.- Effect of power on lift coefficient for the pusher model. $\\delta_e$, $0^\\circ$; $\\delta_f$, $0^\\circ$; approximate test air speed, 30 m.p.h.\n\nFigure 32.- Effect of power on lift coefficient for tractor position 1. $\\delta_e$, $0^\\circ$; $\\delta_f$, $0^\\circ$; approximate test air speed, 30 m.p.h.\n\nFigure 33.- Effect of power on the maximum lift coefficient and on the lift-curve slope for the pusher model and for tractor position 1. $\\delta_e$, $0^\\circ$; $\\delta_f$, $0^\\circ$; approximate test air speed, 30 m.p.h.\n\nFigure 34.- Effect of power on the pitching-moment coefficient for the model with wing nacelles and external radiators. $\\delta_e$, $0^\\circ$; $\\delta_f$, $0^\\circ$.\n\nFigure 35.- Effect of power on the pitching-moment coefficient for the pusher model. $\\delta_e$, $0^\\circ$; $\\delta_f$, $0^\\circ$.\n\nFigure 36.- Effect of power on the pitching-moment coefficient for tractor position 1. $\\delta_e$, $0^\\circ$; $\\delta_f$, $0^\\circ$.", "timestamp": "2026-07-19T18:38:10.564192+00:00"}
{"citation_id": "19930091701", "source_url": "https://ntrs.nasa.gov/api/citations/19930091701/downloads/19930091701.pdf", "page_number": 9, "total_pages": 18, "image_filename": "19930091701_p9.jpg", "text": "THE TRANSITION PHASE IN THE TAKE-OFF OF AN AIRPLANE 5\n\nThe computations covered three loading conditions: gross weights of 2,060 pounds, 2,378 pounds, and 2,800 pounds. For each load the calculations were carried through for three normal take-offs at different speeds ranging from an assumed minimum allowable speed to 20 percent in excess of this value. Similarly, two zoom take-offs were calculated for take-off speeds 10 percent and 20 percent greater than the minimum allowable speed at which the final steady climb was assumed to be made in both cases. The minimum allowable speed was arbitrarily taken as 4 percent in excess of the speed corresponding to the maximum lift coefficient, 1.3. For all the foregoing conditions there was assumed to be no wind.\n\nThe effects of wind were determined for two cases: one with the heaviest loading and the other with the lightest loading. For these cases there was introduced into the calculations a wind velocity of 5-miles-per-hour magnitude at the ground, increasing with height according to the relationship given by reference 1 as representing an average wind gradient, which is\n\n$$\n\\frac{V_{w}}{V_{w_{0}}}=\\left(\\frac{H_{e}}{H_{0}}\\right)^{1 / 7}\n$$\n\nwhere $V_{w_{0}}$, which was assigned a value of 5 miles per hour, is the wind speed corresponding to $H_{0}$, the effective height of the airplane while in contact with the ground, assumed to be 5 feet; and $V_{w}$ is the wind speed at any other effective height $H_{e}$, i. e., the height of the wheels above the ground plus 5 feet.\n\nFor the same two loading conditions, the effect of ground proximity on the air-borne phase of take-off was investigated by using the excess-thrust data obtained above the 10-foot level, hence sensibly outside the influence of ground effect, throughout the integrations and comparing the results with those obtained for similar cases in which the ground effect was included.\n\nThe ground-run phase of the take-off was considered only insofar as was necessary to show the effects of variations in take-off speeds and methods on the complete take-off. In all cases only the distance required to accelerate from a common speed of 75 feet per second up to the take-off speed was calculated. In the determination of these distances, the rolling-friction coefficient was assumed to be 0.05, corresponding to an average turf surface. The air forces were taken from the data obtained within the region of ground effect.\n\nRESULTS\n\nA summary of the results obtained from the calculations is given in table II. Figures 5 through 7 show the calculated flight paths of the airplane during the transition and steady climb for all the conditions investigated. In figure 8 the distance on the ground required to accelerate from a speed of 75 feet per second to the take-off speed is plotted against take-off speed for the three loading conditions. Figures 9 through 11 show the variation due to take-off speed in the air-borne distances required to clear heights of 50 and 100 feet for both normal and zoom take-offs. These figures also show the effect of take-off speed on the over-all take-off distance, i. e., including the ground run after a velocity of 75 feet per second is attained.\n\nFigure 12 shows the percentage difference for various take-off speeds between the air-borne distance as calculated by the methods previously described, where due consideration was given to the transition, and the distance that would be obtained were the transition to be neglected. For the normal take-offs, the value for the distance with the transition neglected was taken as\n\n$$\nD=\\frac{H}{\\tan \\gamma}\n$$\n\nwhere $H$ is the obstacle height to be cleared and $\\gamma$ is the flight-path angle corresponding to a given speed. This relation was based on the assumption that steady-climbing conditions obtained from the instant of leaving the ground. For the zoom take-offs the most obvious approximate relation for the air-borne distance appeared to be\n\n$$\nD=\\frac{H-\\frac{V_{1}^{2}-V_{2}^{2}}{2 g}}{\\tan \\gamma}\n$$\n\nwhere $V_{1}$ and $V_{2}$ are the initial and final flight-path velocities, respectively, and $\\gamma$ is the flight-path angle corresponding to $V_{2}$. This equation was based on the assumptions that constant excess power was available throughout the climb and that steady conditions were realized before the height $H$ was attained. Figure 12 is intended to indicate the extent to which the take-off is affected by the transition and the magnitude of the error that might be introduced by the neglect of the transition in the calculation of take-off distances.\n\nThe effect on the air-borne distance of an average wind gradient corresponding to a surface wind velocity of 5 miles per hour is shown in figure 13 for normal take-offs with the heaviest and lightest loads. In figure 14 the influence of ground effect is shown for the same loading conditions.\n\nDISCUSSION\n\nThe nature of the flight path during the transition phase of the take-off is shown in figures 5, 6, and 7. The initially increasing slope of the path followed later by a decrease is apparently characteristic, at least for the airplane and conditions considered herein. In the case of normal take-off, the reason for this reversal of curvature lies in the fact that the airplane continues to accelerate immediately after leaving the ground, and, in being slowed to its original speed, assumes a climbing angle too steep to be maintained. The flight-path angle", "timestamp": "2026-07-19T18:38:11.256414+00:00"}
{"citation_id": "19930093640", "source_url": "https://ntrs.nasa.gov/api/citations/19930093640/downloads/19930093640.pdf", "page_number": 1, "total_pages": 28, "image_filename": "19930093640_p1.jpg", "text": "L-471\n\nACR July 1939\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nWARTIME REPORT\n\nORIGINALLY ISSUED\nJuly 1939 as\nAdvance Confidential Report\n\nAERODYNAMIC CHARACTERISTICS OF A 4-ENGINE MONOPLANE\nSHOWING COMPARISON OF AIR-COOLED AND\nLIQUID-COOLED ENGINE INSTALLATIONS\n\nBy Abe Silverstein and Herbert A. Wilson, Jr.\n\nLangley Memorial Aeronautical Laboratory\nLangley Field, Va.\n\nTECHNICAL LIBRARY\nAIRESEARCH MANUFACTURING CO.\n9851-9851 SEPULVEDA BLVD.\nINGLEWOOD,\nCALIFORNIA\n\nNACA\n\nWASHINGTON\n\nNACA WARTIME REPORTS are reprints of papers originally issued to provide rapid distribution of\nadvance research results to an authorized group requiring them for the war effort. They were pre-\nviously held under a security status but are now unclassified. Some of these reports were not tech-\nnically edited. All have been reproduced without change in order to expedite general distribution.\n\nL - 471", "timestamp": "2026-07-19T18:38:14.063610+00:00"}
{"citation_id": "19930093281", "source_url": "https://ntrs.nasa.gov/api/citations/19930093281/downloads/19930093281.pdf", "page_number": 3, "total_pages": 22, "image_filename": "19930093281_p3.jpg", "text": "L-279\n\nTHE EFFECT OF STREAMLINING THE AFTERBODY OF\nAN N.A.C.A. COWLING\nBy George W. Sticklo, John L. Crigler, and Irven Naiman\n\nSUMMARY\n\nThe drag and the power cost associated with the\nchanging of the nose of a nacelle from a streamline shape\nto a conventional N.A.C.A. cowling shape was investigated\nin the N.A.C.A. 20-foot tunnel. Full-scale propellers\nand nacelles were used. The increment of drag associated\nwith the change of nose shapes was found to be critically\ndependent on the afterbody of the nacelle. Two streamline\nafterbodies were tested. The results of the tests with\nthe more streamlined afterbody showed that the drag ap-\nproached that of an airship form and that the added drag\ndue to the open-nose cowling was only one-fourth of the\ndrag increase obtained with the other afterbody. The re-\nsults of this research indicate that the power cost, in\nexcess of that with a streamline nose, of using an N.A.C.A.\ncowling in front of a well-designed afterbody to enclose\na 1,500-horsepower engine in an airplane with a speed of\n300 miles per hour amounts to 1.5 percent of the engine\npower. If the open-nose cowling is credited with 1 per-\ncent because it cools the front of the cylinders, the non-\nuseful power cost amounts to only 0.5 percent of the en-\ngine power.\n\nINTRODUCTION\n\nThe two primary functions of an engine cowling are:\n\n(1) To provide an engine enclosure of minimum drag.\n\n(2) To pump the cooling air through the engine or\nradiator.\n\nReference 1 points out that these functions may be treated\nseparately because the definite amount of work required to\nbe done on the cooling air is distinctly different from\nthe ordinary aerodynamic drag of the cowling itself.\n\nIt is further shown in reference 1 that the drag", "timestamp": "2026-07-19T18:38:21.154215+00:00"}
{"citation_id": "19930091748", "source_url": "https://ntrs.nasa.gov/api/citations/19930091748/downloads/19930091748.pdf", "page_number": 1, "total_pages": 12, "image_filename": "19930091748_p1.jpg", "text": "To Sam with compliments.\nAlz\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\n\nREPORT No. 673\n\nEXPERIMENTAL VERIFICATION OF THE THEORY\nOF OSCILLATING AIRFOILS\n\nBy ABE SILVERSTEIN and UPSHUR T. JOYNER\n\n[Figure: Seal of the National Advisory Committee for Aeronautics]\n\n1939\n\nFor sale by the Superintendent of Documents, Washington, D. C. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 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. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .", "timestamp": "2026-07-19T18:38:26.478748+00:00"}
{"citation_id": "19930091697", "source_url": "https://ntrs.nasa.gov/api/citations/19930091697/downloads/19930091697.pdf", "page_number": 11, "total_pages": 28, "image_filename": "19930091697_p11.jpg", "text": "A PHOTOGRAPHIC STUDY OF COMBUSTION AND KNOCK IN A SPARK-IGNITION ENGINE 7\n\nThe high-speed motion pictures in figure 7 show the flames for fuels of different octane ratings with one spark plug at E. The outlines of the flames, for the first several frames, have been marked by a series of white dots for purposes of reproduction. Details of the photographs are more easily seen in the enlargements of figures 8 and 9. The photographs show an even illumination and a uniform rate of flame propagation for the nonknocking burning with the 100-octane fuel. For the knocking explosions, there was a sudden and definite increase in the intensity of illumination at the time in the cycle that knock occurred. This sudden increase in brightness at the time of occurrence of knock is characteristic of all the knocking explosions and increases in intensity with increasing violence of knock. It is not to be confused with the bright light given off by the 100-octane fuel, which has a very high actinic value, the inflamed area being very bright and uniform throughout the entire explosion. The first appearance of the bright illumination is indicated by the frame marked A. It will later be shown that the time of appearance of the bright light coincides exactly with the appearance of pressure waves, hence may be used to indicate the first appearance of knock.\n\nThe records for octane ratings of 18 and 30 show a sudden inflammation of the end gas just prior to the appearance of the bright illumination. The photographs for the 40- and 50-octane fuels also indicate that a sudden inflammation of the end gas took place but, in this case, the sudden increase in flame travel and the sudden increase in illumination are shown in the same frame. The record for 65-octane fuel shows that the flame proceeded at a uniform rate across the chamber; then, after the chamber was filled with flame, the characteristic bright light associated with knock appeared. As nearly as can be determined, the violent vibrations that appear on indicator cards of knocking explosions correspond in point of time to the appearance of this bright light. The vibrations shown on the indicator cards for the 65-octane fuel were more violent than those for explosions of some of the fuels of lower octane ratings.\n\nFigure 10 is a composite of indicator cards and streak schlieren photographs, taken with fuels of different octane ratings. The spark plug (at E, fig. 1) was located at the bottom of the strip as it is shown in the figure. For 100-octane fuel, the rate of combustion-front travel was somewhat slower than the rate for the other fuels. This result may or may not be significant. There is no indication of any vibration in the gas. The maximum pressure shown by the indicator card for the 100-octane fuel is about 800 pounds per square inch.\n\nEach of the other three records shows knocking explosions with the characteristic gas vibrations, the frequency of which corresponds approximately to the frequency recorded on the indicator cards. The slit through which the photographs were taken did not\n\n| Octane rating |\n|---------------|\n| 100           |\n| 65            |\n| 50            |\n| 40            |\n| 30            |\n| 18            |\n\n[Figure: High-speed motion pictures showing effect of fuels of different octane ratings on combustion knock. Air-fuel ratio, 14:1. First evidence of knock; engine speed, 500 r. p. m.; one spark plug.]\n\n20° B.T.C.  \nT.C.  \n20° A.T.C.\n\n49934—38——2", "timestamp": "2026-07-19T18:38:35.217766+00:00"}
{"citation_id": "19930094564", "source_url": "https://ntrs.nasa.gov/api/citations/19930094564/downloads/19930094564.pdf", "page_number": 16, "total_pages": 16, "image_filename": "19930094564_p16.jpg", "text": "N.A.C.A. Technical Memorandum No. 852\n\nFigs. 10,11,12,13\n\n.012\nTunnel Source\n• DVL 5x7m (6)\n○ VDT 1.5m (2)\nFigs. 2 and 8\nCwp(Cd=0)\n.010\n.008\n.006\n0 .08 .16 .24\nd/t\nΔcw = .00125\nΔcw = .000045\n\n.008\n△ N.A.C.A. 00\n● \" 24\n○ \" 230\nCwp(Cd = 0.1)\n.007\n.006\n.005\n0 .08 .16 .24\nd/t\nFigure 11.- Profile drag of N.A.C.A.\nairfoil series 00, 24\nand 230 against airfoil thickness\nfor R = 20 x 10⁶\n\nFigure 10.- Comparison of profile drag\nof V.D.T. and D.V.L. tunnel\nfor N.A.C.A. airfoil series 24 on the\nbasis of blunt wing tips and R = 8.2 x\n10⁶\n\n2.6\n2.4\n2.2\n2.0\n1.8\n1.6\n1.4\n1.2\n1.0\n0 .04 .08 .12 .16 .20\nd/t\n△ N.A.C.A. 00\n● \" 24\n○ \" 230\nWith split flap\nWithout split flap\nCαmax\n\n380\n340\n300\n260\n220\n180\n140\n100\n0 .08 .16 .24\nd/t\nWith\nsplit flap\n230\n24\n00\nWithout split flap.\nCαmax / Cwp(Cd = 0.1)\n\nFigure 12.- Cαmax of N.A.C.A. airfoil\nseries 00, 24, and 230\nwith and without split flap against\nairfoil thickness at R = 4 x 10⁶\n\nFigure 13.- Rating factor Cαmax /\nCwp(Cd = 0.1) for N.A.C.A.\nairfoil series 00, 24, 230 with and\nwithout split flap against profile\nthickness for average Reynolds\nNumber in free flight.", "timestamp": "2026-07-19T18:38:37.250917+00:00"}
{"citation_id": "19930091719", "source_url": "https://ntrs.nasa.gov/api/citations/19930091719/downloads/19930091719.pdf", "page_number": 4, "total_pages": 33, "image_filename": "19930091719_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\nDAVID W. TAYLOR, D. Eng., Vice Chairman,  \nWashington, D. C.\n\nWILLIS RAY GREGG, Sc. D., Chairman, Executive Committee,  \nChief, United States Weather Bureau.\n\nWILLIAM P. MACCRACKEN, J. D., Vice Chairman, Executive Committee,  \nWashington, D. C.\n\nCHARLES G. ABBOT, Sc. D.,  \nSecretary, Smithsonian Institution.\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\nHARRY F. GUGGENHEIM, M. A.,  \nPort Washington, Long Island, N. Y.\n\nSYDNEY M. KRAUS, Captain, United States Navy,  \nBureau of Aeronautics, Navy Department.\n\nCHARLES A. LINDBERGH, LL. D.,  \nNew York City.\n\nDENIS MULLIGAN, J. S. D.,  \nDirector of Air Commerce, Department of Commerce.\n\nAUGUSTINE W. ROBINS, Brigadier General, United States Army,  \nChief Matériel Division, Air Corps, Wright Field,  \nDayton, Ohio.\n\nEDWARD P. WARNER, Sc. D.,  \nGreenwich, Conn.\n\nOSCAR WESTOVER, Major General, United States Army,  \nChief of Air Corps, War Department.\n\nORVILLE WRIGHT, Sc. D.,  \nDayton, Ohio.\n\nGEORGE W. LEWIS, Director of Aeronautical Research  \nJOHN F. VICTORY, Secretary  \nHENRY J. E. REID, Engineer-in-Charge, Langley Memorial Aeronautical Laboratory, Langley Field, Va.  \nJOHN J. IDE, Technical Assistant in Europe, Paris, France\n\n## TECHNICAL COMMITTEES\n\n| AERODYNAMICS                  | AIRCRAFT STRUCTURES         |\n|-------------------------------|-----------------------------|\n| POWER PLANTS FOR AIRCRAFT     | AIRCRAFT ACCIDENTS          |\n| AIRCRAFT MATERIALS            | INVENTIONS 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  \nscientific research on the fundamental  \nproblems of flight.\n\n## OFFICE OF AERONAUTICAL INTELLIGENCE  \nWASHINGTON, D. C.\n\nCollection, classification, compilation,  \nand dissemination of scientific and technical  \ninformation on aeronautics.", "timestamp": "2026-07-19T18:39:02.258805+00:00"}
{"citation_id": "19930094559", "source_url": "https://ntrs.nasa.gov/api/citations/19930094559/downloads/19930094559.pdf", "page_number": 16, "total_pages": 16, "image_filename": "19930094559_p16.jpg", "text": "N.A.C.A. Technical Memorandum No. 857\nFigs.13,14,15,16\n\n[Figure: A series of images showing combustion in a chamber, labeled Figure 13.]\nFigure 13.\n\n[Figure: A series of images showing combustion in a chamber, labeled Figure 14.]\nFigure 14.\n\n[Figure: A series of images showing combustion in a chamber, labeled Figure 15.]\nFigure 15.\n\n[Figure: A series of images showing combustion in a chamber, labeled Figure 16.]\nFigure 16.\n\nFigures 13,14,15,16.- Effect of fuel quantity and air velocity on fuel\ndistribution and combustion in air storage type\ncombustion chamber.\nFuel: Derop gasoil. Bosch pump PE 1 B 100/100, n=400 r.p.m. Bosch\nnozzle DN 4 S 9, valve opening pressure=150 atm. seconds per\npicture=.0024 s.", "timestamp": "2026-07-19T18:39:02.823159+00:00"}
{"citation_id": "19930094493", "source_url": "https://ntrs.nasa.gov/api/citations/19930094493/downloads/19930094493.pdf", "page_number": 1, "total_pages": 31, "image_filename": "19930094493_p1.jpg", "text": "[FILE COPY\nNO. 5]\n\nTECHNICAL MEMORANDUMS\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nNo. 923\n\nMEASUREMENT OF THE AIR-FLOW VELOCITY IN THE CYLINDER\nOF AN AIRPLANE ENGINE\nBy Hermann Wenger\n\nLuftfahrtforschung\nVol. 16, No. 2, February 20, 1939\nVerlag von R. Oldenbourg, München und Berlin\n\nTHIS DOCUMENT ON LOAN FROM THE FILES OF\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\nLANGLEY AERONAUTICAL LABORATORY\nLANGLEY FIELD, HAMPTON, VIRGINIA\n\nRETURN TO THE ABOVE ADDRESS.\n\nREQUESTS FOR PUBLICATIONS SHOULD BE ADDRESSED\nAS FOLLOWS:\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n1724 I STREET, N.W.,\nWASHINGTON 25, D.C.\n\nWashington\nDecember 1939\n\n[FILE COPY\nTo be returned to\nthe files of the National\nAdvisory Committee\nfor Aeronautics\nWashington, D.C.]", "timestamp": "2026-07-19T18:39:02.965055+00:00"}
{"citation_id": "19930093641", "source_url": "https://ntrs.nasa.gov/api/citations/19930093641/downloads/19930093641.pdf", "page_number": 26, "total_pages": 47, "image_filename": "19930093641_p26.jpg", "text": "22\n\nFigure 37.- Comparison of high-speed computations for the wing-nacelle and the pusher models showing the effect of enclosing engines and radiators within the wings. Wing loading, 25.7 pounds per square foot; power loading, 17.7 pounds per horsepower, $\\beta$, $18\\frac{1}{2}^\\circ$; standard sea-level density; $\\delta_o$, $0^\\circ$; $\\delta_f$, $0^\\circ$.", "timestamp": "2026-07-19T18:39:03.168183+00:00"}
{"citation_id": "19930094568", "source_url": "https://ntrs.nasa.gov/api/citations/19930094568/downloads/19930094568.pdf", "page_number": 4, "total_pages": 28, "image_filename": "19930094568_p4.jpg", "text": "N.A.C.A. Technical Memorandum No. 848\n3\n\n| Planing surface | Airfoil | |\n| :--- | :--- | :--- |\n| $R_o$ | $R_o'$ | normal force of plate at infinitely small angle of attack. |\n| $R$ | $R'$ | normal force of plate at finite angle of attack (fig. 1). |\n| $A_o = R_o \\cos \\beta$ | $A_o' = R_o' \\cos \\beta$ | lift at infinitely small angle of attack. |\n| $A = R \\cos \\beta$ | $A' = R' \\cos \\beta$ | lift at finite angle of attack. |\n| $\\mu = \\frac{m}{m_o} = \\frac{A}{A_o}$ | $\\mu' = \\frac{m'}{m_o'} = \\frac{A'}{A_o'}$ | conversion factor. |\n| $c_a = \\frac{A}{\\frac{\\rho}{2} V^2 b l}$ | $c_a' = \\frac{A'}{\\frac{\\rho}{2} V^2 b l}$ | lift coefficient. |\n| $W_Z$ | | horizontal push-rod force. |\n| $W_R$ | | frictional drag in boundary-layer drag coefficient. |\n| $c_f = \\frac{W_R}{\\frac{\\rho}{2} V^2 b l}$ | | drag coefficient. |\n| $\\kappa$ | | auxiliary quantity (taken from fig. 18 of reference 2). |\n\nPART I\n\n1. Previous Studies\n\nIn his method of explaining the case of accelerated planing, H. Wagner (references 1 and 2) disregarded both the gravity and the fluid viscosity. He observed that, in accelerated planing at very (infinitely) small angles of attack, the lift of a planing surface is exactly half as great as that of an identical infinitely thin airfoil whose plan corresponds to the wetted surface (pressure surface) (airfoil comparison).", "timestamp": "2026-07-19T18:39:05.630986+00:00"}
{"citation_id": "19930094544", "source_url": "https://ntrs.nasa.gov/api/citations/19930094544/downloads/19930094544.pdf", "page_number": 32, "total_pages": 43, "image_filename": "19930094544_p32.jpg", "text": "N.A.C.A. Technical Memorandum No. 872\nFigs. 1,2,3,4,5,6\n\n[Figure: A long, cigar-shaped airship flying low over a body of water with a treeline in the background.]\nFigure 1.-First Zeppelin airship.\n\n[Figure: A large, white airship with a dark stripe along its side, flying at an angle above a building with a curved roof.]\nFigure 2.- Z-ship SCHWABEN.\n\n[Figure: A long, dark airship with multiple gondolas underneath, flying in the sky.]\nFigure 3.- Second Schuette-Lanz airship, SL2.\n\n[Figure: A smaller, dark, blunt-nosed airship with a single gondola and tail fins.]\nFigure 4.- First Parseval pressure ship.\n\n[Figure: A dark, cigar-shaped airship flying low over a flat landscape with some structures visible on the ground.]\nFigure 5.- Siemens-Schuckert pressure ship.\n\n[Figure: A very large, sleek airship flying low over a field with trees in the distance.]\nFigure 6.- GRAF ZEPPELIN (LZ-127).", "timestamp": "2026-07-19T18:39:08.957782+00:00"}
{"citation_id": "19930091701", "source_url": "https://ntrs.nasa.gov/api/citations/19930091701/downloads/19930091701.pdf", "page_number": 10, "total_pages": 18, "image_filename": "19930091701_p10.jpg", "text": "6\nREPORT NO. 626—NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\n<!-- Image (179, 144, 442, 390) -->\nFIGURE 5.—Weight, 2,000 pounds.\n\n<!-- Image (515, 144, 872, 390) -->\nFIGURE 6.—Weight, 2,578 pounds.\n\n<!-- Image (187, 496, 859, 760) -->\nFIGURE 7.—Weight, 2,800 pounds.\n\nFIGURES 5 TO 7.—Flight paths followed in transition and steady climb for normal and zoom take-offs for the Verville AT airplane at various speeds.", "timestamp": "2026-07-19T18:39:09.749490+00:00"}
{"citation_id": "19930093640", "source_url": "https://ntrs.nasa.gov/api/citations/19930093640/downloads/19930093640.pdf", "page_number": 2, "total_pages": 28, "image_filename": "19930093640_p2.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-19T18:39:10.683251+00:00"}
{"citation_id": "19930093281", "source_url": "https://ntrs.nasa.gov/api/citations/19930093281/downloads/19930093281.pdf", "page_number": 4, "total_pages": 22, "image_filename": "19930093281_p4.jpg", "text": "2\n\nchargeable to pumping the air through the cowling is equal to the internal work (that is, the volume multiplied by the pressure drop) divided by the free-air velocity and the pumping efficiency. The pumping efficiency is shown to be nearly 100 percent for the high-speed condition. Tests with cooling air, the results of which are to be included in another report, show that this pumping efficiency can be obtained on set-up 2, which is the subject of this report. It is shown in reference 2 that a 550-horsepower engine operating with a temperature difference of $300^\\circ$ F. required approximately 1-1/2 percent of the engine power for internal cooling work. This internal work is utilized in cooling the rear of the engine cylinders. More modern engines with improved finning and baffling have reduced this value to about 1 percent of the engine power.\n\nThe problem of providing an engine enclosure that would have minimum drag was investigated in reference 1. The best design of the nose contour was determined as well as the best method of exhausting the cooling air. It was stated that the drag of the basic blunt-nose cowling shape of an air-cooled engine has a drag somewhat in excess of that of a more properly streamlined shape, such as an airship form. In order to ascertain the reason for this increase in drag, several cowling noses varying in contour and dimensions were investigated to determine the variable of the nose shape that made the drag of the open-nose cowling larger. At the beginning of this research, an afterbody similar to that of reference 1 was used but, when the design was copied, the expansion angle of the finished nacelle was slightly larger than that of the nacelle used in reference 1. This small change in the expansion angle gave a critical flow over the after part of the nacelle and the drag coefficient changed radically with the Reynolds Number. This undesirable condition focused attention on the shape of the afterbody and work was begun to design an afterbody that would not give a critical flow condition.\n\nThe problem of reducing the form drag of the afterbody of the nacelle is similar to the problem of designing the expansion side of a venturi tube. The air must be slowed down with the least loss of energy. If the expansion angle is too large, loss in energy occurs because the kinetic energy is not transformed into potential energy. If the expansion angle is too small, skin friction over the body will make the drag too high. Further study of", "timestamp": "2026-07-19T18:39:13.337161+00:00"}
{"citation_id": "19930094498", "source_url": "https://ntrs.nasa.gov/api/citations/19930094498/downloads/19930094498.pdf", "page_number": 1, "total_pages": 37, "image_filename": "19930094498_p1.jpg", "text": "FILE COPY\nNO. 4\n\nTECHNICAL MEMORANDUMS\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nTHIS DOCUMENT ON LOAN FROM THE FILES OF\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\nLANGLEY AERONAUTICAL LABORATORY,\nLANGLEY FIELD, HAMPTON, VIRGINIA.\n\nRETURN TO THE ABOVE ADDRESS.\n\nREQUESTS FOR PUBLICATIONS SHOULD BE ADDRESSED\nAS FOLLOWS:\n\nNo. 918\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n1512 H STREET, N. W.\nWASHINGTON 25, D. C.\n\nTHE ENLARGED N.A.C.A. TANK, AND SOME OF ITS WORK\n\nBy Starr Truscott\n\nSupplementary Volume to\nJahrbuch 1938 der deutschen Luftfahrtforschung\n(Containing Papers Presented at Lilienthal Gesellschaft\nfür Luftfahrtforschung October 1938)\n\nWashington\nNovember 1939", "timestamp": "2026-07-19T18:39:31.962922+00:00"}
{"citation_id": "19930091748", "source_url": "https://ntrs.nasa.gov/api/citations/19930091748/downloads/19930091748.pdf", "page_number": 2, "total_pages": 12, "image_filename": "19930091748_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/s² 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-s² at 15° C. and 760 mm; or 0.002378 lb.-ft.⁻⁴ sec.²\nSpecific weight of \"standard\" air, 1.2255 kg/m³ 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$\\frac{Vl}{\\mu}$, 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° 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$C_{p_1}$, 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$\\alpha_s$, Angle of attack, absolute (measured from zero-lift position)\n$\\gamma$, Flight-path angle", "timestamp": "2026-07-19T18:39:35.735007+00:00"}
{"citation_id": "19930094499", "source_url": "https://ntrs.nasa.gov/api/citations/19930094499/downloads/19930094499.pdf", "page_number": 1, "total_pages": 13, "image_filename": "19930094499_p1.jpg", "text": "TECHNICAL MEMORANDUMS\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nNo. 917\n\nUniversity of Maryland\nGlenn L. Martin College\nof Engineering and Aero-\nnautical Sciences\nLibrary\n\nTHE EFFECT OF COMPRESSIBILITY ON THE PRESSURE READING\nOF A PRANDTL PITOT TUBE AT SUBSONIC FLOW VELOCITY\nBy O. Walchner\n\nDeutsche Luftfahrtforschung\nJahrbuch 1938\nVerlag von R. Oldenbourg, München und Berlin\n\nWashington\nNovember 1939", "timestamp": "2026-07-19T18:39:45.687083+00:00"}
{"citation_id": "19930091724", "source_url": "https://ntrs.nasa.gov/api/citations/19930091724/downloads/19930091724.pdf", "page_number": 2, "total_pages": 20, "image_filename": "19930091724_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 . . . . . . | $l$ | meter . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 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. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 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. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .", "timestamp": "2026-07-19T18:39:48.625988+00:00"}
{"citation_id": "19930091701", "source_url": "https://ntrs.nasa.gov/api/citations/19930091701/downloads/19930091701.pdf", "page_number": 11, "total_pages": 18, "image_filename": "19930091701_p11.jpg", "text": "THE TRANSITION PHASE IN THE TAKE-OFF OF AN AIRPLANE 7\n\nmust therefore be reduced to a value at which the airplane can climb steadily. The flight paths for the zoom take-offs have a generally similar shape, but variations in the slope are more pronounced owing to the greater changes in speed.\n\nInasmuch as most airplanes probably have, in part by virtue of the ground effect, an excess of thrust in the initial stage of the transition and hence will accelerate, it is likely that the form of the transition curve shown is representative of the form that would generally be experienced.\n\nThe procedure that would be required in controlling an airplane along a path such as that described is indicated in figure 4. The control column would first be pulled back to put the airplane in an attitude of high lift and held until the angle of climb was sufficient to cause a deceleration. It would then be pushed forward to reduce the angle of attack to a value considerably below that corresponding to the steady climb in time to prevent the speed from dropping below that prescribed for the climb. Finally, it would again be pulled back as the angle of climb decreased so that the correct flight-path angle and the angle of attack for steady climbing might be simultaneously realized. In practice, it would probably not be possible to synchronize, exactly, the attainment of the proper flight-path angle and angle of attack; consequently, an oscillatory rather than a steady flight path would result. If sufficient effort were made to maintain constant speed, however, the amplitude of the oscillation would not be great and the mean flight path would probably correspond closely to the one that would be obtained under steady conditions.\n\nFor normal take-offs, it is apparent from the curves of figures 5, 6, and 7 that, insofar as the transition alone is concerned, the optimum take-off speed, in the range considered, is the lowest value shown. Higher speeds provide an initially greater excess of lift and, consequently, a higher vertical acceleration, so that the transition is completed more quickly and with less variation in forward velocity. At the slower speed, however, there is a greater excess thrust available which, although partly converted to kinetic energy at first, eventually goes toward increasing the height or potential energy of the airplane; thus, when the transition is completed, the height attained is greater in proportion to the horizontal distance covered than that for the higher-speed take-offs.\n\nThe maximum angle of climb occurs at approximately the intermediate speed shown so that, in the range of speeds between the minimum and that for best angle of climb, the effects of variations in take-off speed on the transition and on the steady climb are opposed. For an obstacle height of 50 feet a considerable portion of the air-borne distance is occupied by the transition so that the opposing effects are nearly balanced. Hence there is little change in the air-borne distance with increasing take-off speed up to the speed for best angle of climb (figs. 9, 10, and 11); beyond this speed the distance, of course, increases. Obviously then, since the ground-run distance (fig. 8) increases with the take-off speed, the shortest overall take-off distance required to gain a height of 50 feet, in a normal take-off, would be realized with the lowest possible take-off speed.\n\nWith an obstacle height of 100 feet the transition is a relatively small part of the air-borne distance. The effect of take-off speed on the steady climb is therefore predominant and consequently the shortest air-borne distance occurs at or near the speed for best angle of climb. The reduction in air-borne distance, however, is more than offset by the increased ground run so that, in this case also, the lowest take-off speed gives the shortest overall distance.\n\nIn zoom take-offs the airplane is held in contact with the ground until the speed reaches a value considerably\n\n<!-- Image (523, 382, 842, 563) -->\n\nFIGURE 8.—Ground distance required for the Verville AT airplane to accelerate from 75 feet per second to take-off speed for all loading conditions.\n\nabove the minimum flying speed. It is then pulled off abruptly into a steep climb during which the speed is reduced. It may be shown that an airplane running along the ground at its most efficient attitude, i. e., the attitude corresponding to the minimum value of $C_D - \\mu C_L$, will ordinarily have, in the range of speeds between the minimum flying speed and a speed considerably in excess of the minimum, appreciably less resistance, hence greater excess thrust, than if it were completely air-borne at similar speeds. The excess kinetic energy gained in running a given distance along the ground would be greater, therefore, than the potential energy that might be gained in flight in the same distance. Thus, if the excess kinetic energy could be converted to potential energy without too great loss, it should be possible to attain a greater height in a given distance from a zoom take-off than from the shortest normal take-off. This argument is borne out in figures 9, 10, and 11 for an obstacle height of 100 feet where, with the lightest load, the total horizontal distance required to gain this height from a ground speed of 75 feet per second is about 5 percent less for the shortest zoom\n\n61908—38—2", "timestamp": "2026-07-19T18:40:02.001137+00:00"}
{"citation_id": "19930094542", "source_url": "https://ntrs.nasa.gov/api/citations/19930094542/downloads/19930094542.pdf", "page_number": 86, "total_pages": 102, "image_filename": "19930094542_p86.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-19T18:40:03.690987+00:00"}
{"citation_id": "19930094504", "source_url": "https://ntrs.nasa.gov/api/citations/19930094504/downloads/19930094504.pdf", "page_number": 1, "total_pages": 16, "image_filename": "19930094504_p1.jpg", "text": "National Advisory Committee\nfor Aeronautics\nMAILHD\nOCT 5 1939\nTo Mr Ryder\n\nFILE COPY\nNO 2\n\nTECHNICAL MEMORANDUMS\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nTHIS DOCUMENT ON LOAN FROM THE FILES OF\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\nLANGLEY AERONAUTICAL LABORATORY\nLANGLEY FIELD, HAMPTON, VIRGINIA\n\nRETURN TO THE ABOVE ADDRESS\n\nREQUESTS FOR PUBLICATIONS SHOULD BE ADDRESSED\nAS FOLLOWS:\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n1512 H STREET, N. W.\nWASHINGTON 25, D. C.\n\nNo. 912\n\nINCREASE OF THE SPECIFIC LOAD UNDER TENSION, COMPRESSION, AND\nBUCKLING OF WELDED STEEL TUBES IN AIRPLANE CONSTRUCTION BY\nSUITABLE TREATMENT OF STRUCTURAL STEEL AND BY PROPER DESIGN\n\nBy J. Müller\n\nLuftfahrtforschung\nVol. 16, No. 1, Jan. 10, 1939\nVerlag von R. Oldenbourg, München und Berlin\n\nWashington\nOctober 1939", "timestamp": "2026-07-19T18:40:04.357463+00:00"}
{"citation_id": "19930093281", "source_url": "https://ntrs.nasa.gov/api/citations/19930093281/downloads/19930093281.pdf", "page_number": 5, "total_pages": 22, "image_filename": "19930093281_p5.jpg", "text": "6Lb-y\n3\n\nthe effect of the shape of the afterbody on the drag is\nplanned. If the cowling is placed in front of a wing that\nhas a thickness equal to or larger than the nacelle diam-\neter, or in front of a fuselage with a diameter equal to\nor larger than the nacelle diameter, the slowing down of\nthe air is taken care of by the wing or fuselage contour.\nIf the air that flows over the wing or fuselage is at no\nplace expanded too rapidly, this source of cowling drag\ndisappears.\n\nThe results of tests using the more streamlined after-\nbody are the subject of the present report. This report\nshows that, if the correct power chargeable to the drag of\nthe nose opening is used, no reason exists, from considera-\ntions of aerodynamic efficiency only, ever to abandon the\nopen-nose cowling for any other type of engine installa-\ntion. This statement takes on added significance when it\nis realized, as is shown in references 1 and 3, that the\nopen-nose cowling provides, at no measurable internal pow-\ner loss, cooling for the front of the cylinder equivalent\nto approximately 70 percent of the cooling obtainable in\nthe free air stream.\n\nSYMBOLS\n\nV, velocity of the free air stream.\n$\\rho$, air density.\nq, dynamic pressure of the air stream, $1/2 \\rho V^2$.\nD, drag of the cowling-nacelle unit.\n$D_o$, drag of streamline shape.\nF, frontal area of the cowling, 14.75 square feet.\n$C_D$, drag coefficient, D/qF.\n$C_{D_o}$, streamline shape drag coefficient.\nR, net thrust of the propeller-nacelle unit.\nP, power input to propeller.\n$\\eta_n$, net efficient of the propeller-nacelle unit, RV/P.", "timestamp": "2026-07-19T18:40:10.978316+00:00"}
{"citation_id": "19930091748", "source_url": "https://ntrs.nasa.gov/api/citations/19930091748/downloads/19930091748.pdf", "page_number": 3, "total_pages": 12, "image_filename": "19930091748_p3.jpg", "text": "REPORT No. 673\n\nEXPERIMENTAL VERIFICATION OF THE THEORY\nOF OSCILLATING AIRFOILS\n\nBy ABE SILVERSTEIN and UPSHUR T. JOYNER\nLangley Memorial Aeronautical Laboratory\n\n16206-39", "timestamp": "2026-07-19T18:40:16.215695+00:00"}
{"citation_id": "19930094498", "source_url": "https://ntrs.nasa.gov/api/citations/19930094498/downloads/19930094498.pdf", "page_number": 2, "total_pages": 37, "image_filename": "19930094498_p2.jpg", "text": "NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nTECHNICAL MEMORANDUM NO. 918\n\nTHE ENLARGED N.A.C.A. TANK, AND SOME OF ITS WORK*\n\nBy Starr Truscott\n\nEARLY TOWING BASINS\n\nWhen work on the original N.A.C.A. tank was begun in 1929, there were few precedents that could be used in its design. The construction, equipment, and methods of testing of most of the towing basins in existence had been developed to suit the study of models of hulls of displacement craft. The resistance of such a model was determined with the expectation that it would be divided into \"frictional\" resistance and \"wave-making,\" or \"residuary,\" resistance according to the Froude method and that the resistance of the full-size craft would be estimated by computing the frictional resistance independently and adding the wave-making resistance obtained by stepping up that of the model according to Froude's law.\n\nThe principal interest was in resistance at uniform speed, corresponding to the interest of the ship operator in the ability of a ship to maintain a certain operating speed indefinitely with a minimum of power. What would happen at extremely high speeds was of little interest provided that the resistance at the designed speed was low.\n\nThe towing carriages of most of the model basins were made of structural-steel sections and plate riveted together and usually had four wheels with hardened and ground steel tires. The wheels ran on steel rails that resembled railroad rails and that were machined to provide a straight, smooth, and level course for the wheels.\n\nThe maximum speed of most of the carriages was less than 15 miles per hour.\n\nTHE ORIGINAL N.A.C.A. TANK\n\nWith such precedents, we began the design of a towing basin intended to test models of seaplane floats and hulls.\n\n*\"Der vergrösserte NACA-Schleppkanal und einiges über seine Arbeitsweise.\" Jahrbuch 1938 der deutschen Luftfahrtforschung (supplementary volume), pp. 374-95.", "timestamp": "2026-07-19T18:40:16.456294+00:00"}
{"citation_id": "19930094544", "source_url": "https://ntrs.nasa.gov/api/citations/19930094544/downloads/19930094544.pdf", "page_number": 33, "total_pages": 43, "image_filename": "19930094544_p33.jpg", "text": "N.A.C.A. Technical Memorandum No. 872\nFigs. 7,8,9,10,11,12\n\n[Figure: The English rigid airship R-100.]\nFigure 7.- The English rigid airship R-100.\n\n[Figure: The English rigid airship R-101.]\nFigure 8.- The English rigid airship R-101.\n\n[Figure: The American rigid airship AKRON.]\nFigure 9.- The American rigid airship AKRON.\n\n[Figure: Profiles of more recent rigid airships.]\nFigure 10.- Profiles of more recent rigid airships.\nLZ 113. L:D= 8,84.\n3.42 000 m³\n(1918)\n\"Graf Zeppelin\" L:D= 7,74.\n3.105 000 m³\n(1927/28)\nLZ 129. L:D= 6,0.\n3.190 000 m³\n(1932)\n\"Akron\" L:D= 5,8.\n3.141 000 m³\n(1929/31)\nR 101. L:D= 5,5.\n3.141 000 m³\n(1926/29)\n\n[Figure: Semi-rigid pressure airship PN30.]\nFigure 11.- Semi-rigid pressure airship PN30.\n\n[Figure: PN30, gangway truss.]\nFigure 12.- PN30, gangway truss.", "timestamp": "2026-07-19T18:40:18.434268+00:00"}
{"citation_id": "19930094493", "source_url": "https://ntrs.nasa.gov/api/citations/19930094493/downloads/19930094493.pdf", "page_number": 2, "total_pages": 31, "image_filename": "19930094493_p2.jpg", "text": "NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nTECHNICAL MEMORANDUM NO. 923\n\nMEASUREMENT OF THE AIR-FLOW VELOCITY IN THE CYLINDER  \nOF AN AIRPLANE ENGINE*\n\nBy Hermann Wenger\n\nThe investigation of the air flow in the cylinders of reciprocating engines plays a very important part in the design of internal combustion engines. The design of the combustion chamber and arrangement of the intake and exhaust valves have an effect on the turbulence of the air-fuel mixture. Air flows were first investigated by introducing filaments and observing the motion through glass cylinders. The magnitude of the air velocities in externally driven Diesel engines have been measured by Hintz (reference 1) and Geiger (reference 2), who determined the rotational component of the rotating air mass about the vertical cylinder axis. Hintz found the local velocity to fluctuate during the cycle between 0 and 55 meters per second for an engine speed of 200 revolutions per minute. Geiger, in his measurements carried out on a running Diesel engine but mostly without fuel injection, found velocities of from 0 to about 25 meters per second. The velocity was found to increase with increasing engine speed. Geiger found furthermore that the air velocity at the instant of ignition has a decided effect on the quality of the combustion. J. Ulsamer (reference 3) similarly measured the air velocities in his tests on an air compressor. For this purpose he made use of a hot-wire anemometer, which apparatus will also be employed in the present tests. Three speeds were investigated; namely, 63, 128, and 172 revolutions per minute. He found velocities up to 12 meters per second, which occurred during the intake stroke. The mean velocity was likewise found to increase with increasing engine speed.\n\nThe object of the present investigation is to determine the velocity in the BMW-VI cylinder of an externally driven single-cylinder test engine at high engine speeds using the hot-wire method of Ulsamer.\n\n*Messung der Strömungsgeschwindigkeit im Zylinder eines fremdangetriebenen BMW-VI Flugmotors. Luftfahrtforschung, vol. 16, no. 2, Feb. 20, 1939, pp. 62-73.", "timestamp": "2026-07-19T18:40:28.968161+00:00"}
{"citation_id": "19930094568", "source_url": "https://ntrs.nasa.gov/api/citations/19930094568/downloads/19930094568.pdf", "page_number": 5, "total_pages": 28, "image_filename": "19930094568_p5.jpg", "text": "4 N.A.C.A. Technical Memorandum No. 848\n\nThe attempt at reconciling Wagner's theory (references 1 and 2) with Sottorf's test data on flat planing surfaces (references 3, 4, and 5) was quite satisfactory (curve C, figs. 14 and 16) for short plates (i.e., for small l/b) up to about l/b = 1 at both small and large angles of attack within the entire range of Froude numbers comprising Sottorf's program. For long plates, on the other hand (large l/b), the agreement was far from satisfactory (curve D in figs. 14 to 16). This might be due in part to the increasing gravity effect with great plate length; then, too, it should be observed that Wagner's theory for the case of long planing surfaces is applicable only to very (infinitely) small angles of attack. And so the discrepancies between theory and experiment could be traced to the fact that the assumption of infinitely small angle of attack did not constitute a sufficiently close approximation for Sottorf's test range.\n\n2. Experiments\n\nThe purpose of the experiments was to elucidate these discrepancies between theory and experiment; that is, to separate as far as possible the effect of gravity and that of the finite angle of attack. With this in mind, I made a series of tests with the high-speed carriage in the Prussian Experimental Laboratory for Hydraulics and Ship Design, Berlin, on flat planing surfaces at largest possible Froude numbers; that is, at the highest possible test speed and with the smallest possible plates. (See table I at end of report.) In distinction to Sottorf's tests (reference 3) the peak speed V was raised from 9.5 to 16 meters per second; and the plate width b reduced from 30 to 15 centimeters, thus raising the highest obtained Froude number $F = V/\\sqrt{bg}$ to 2.37 times its value. The experiments with long plates and high load rating, restricted to 6 meters per second speed in Sottorf's test, were considerably extended.\n\nThe experimental arrangement, patterned largely after Sottorf set-up, is described later on.\n\nAs in Sottorf's experiments, the plates were left free to trim and loaded with weights (figs. 1 and 22). At the chosen load ratings\n$$ \\frac{R}{\\frac{\\rho}{2} V^2 b^2} = 0.218, 0.109, \\text{ and} $$", "timestamp": "2026-07-19T18:40:29.939261+00:00"}
{"citation_id": "19930091719", "source_url": "https://ntrs.nasa.gov/api/citations/19930091719/downloads/19930091719.pdf", "page_number": 5, "total_pages": 33, "image_filename": "19930091719_p5.jpg", "text": "REPORT No. 642\n\nTESTS OF FIVE FULL-SCALE PROPELLERS IN THE PRESENCE OF A RADIAL AND A LIQUID-COOLED ENGINE NACELLE, INCLUDING TESTS OF TWO SPINNERS\n\nBy DAVID BIERMANN and EDWIN P. HARTMAN\n\nSUMMARY\n\nWind-tunnel tests are reported of five 3-blade 10-foot propellers operating in front of a radial and a liquid-cooled engine nacelle. The range of blade angles investigated extended from $15^\\circ$ to $45^\\circ$. Two spinners were tested in conjunction with the liquid-cooled engine nacelle. Comparisons are made between propellers having different blade-shank shapes, blades of different thickness, and different airfoil sections.\n\nThe results show that propellers operating in front of the liquid-cooled engine nacelle had higher take-off efficiencies than when operating in front of the radial engine nacelle; the peak efficiency was higher only when spinners were employed. One spinner increased the propulsive efficiency of the liquid-cooled unit 6 percent for the highest blade-angle setting investigated and less for lower blade angles.\n\nThe propeller having airfoil sections extending into the hub was superior to one having round blade shanks.\n\nThe thick propeller having a Clark Y section had a higher take-off efficiency than the thinner one, but its maximum efficiency was possibly lower. Of the three blade sections tested, Clark Y, R. A. F. 6, and N. A. C. A. 2J00-34, the Clark Y was superior for the high-speed condition, but the R. A. F. 6 excelled for the take-off condition.\n\nAPPARATUS AND METHODS\n\nThe propeller-research tunnel has been modified since the description of reference 5 was written to the extent of installing an electric motor to drive the tunnel propeller and of replacing the balance with a more modern one capable of simultaneously recording all the forces.\n\nA 600-horsepower Curtiss Conqueror engine (GIV-1570) was used to drive the test propellers. The engine\n\n[Figure: Radial engine nacelle diagram with dimensions: 116\", 33½\", 42\", 52¼\", 51⅛\", 70½\", 39½\", 44\", 25\", 10\"]\n\n[Figure: Liquid-cooled engine nacelle diagram with Spinner 1 and Spinner 2, dimensions: 125¾\", 29¾\", 69¾\", 109¾\", 32¼\", 24¼\", 17\", 18½\", Sections A-A, B-B, C-C with diameters 37¾\", 42¾\", 36\", 12½\", 5½\"R, 11\"]\n\nFIGURE 1.—Drawings of engine nacelles.\n\nINTRODUCTION\n\nA series of tests of full-scale propellers was made in the propeller-research tunnel during the first part of 1937. Published reports of the series cover separate subjects as: compressibility effects (reference 1), solidity (reference 2), negative thrust and torque (reference 3), and blade section (reference 4). The results of tests of five propellers are published in the present report, the purpose of which is twofold: first, to present design data from tests of four 3-blade propellers made in the presence of two popular body types; and, second, from the test data for all five propellers, to make incidental comparisons regarding the effect of: body shape and size, spinners, blade-shank shape, blade thickness, and blade section. The concrete data should be of value in design work because two of the propellers are in fairly wide use and the body types are representative of those in common use. The comparisons may be of value in the determination of some of the elements of the basic design of airplanes and propellers.\n\nwas mounted in a cradle dynamometer free to rotate about an axis parallel to the propeller axis and located at one side of the engine. The torque reaction was transmitted from the other side of the engine to recording scales located on the floor of the test chamber. The propeller speed was measured by a calibrated electric tachometer.\n\nA scale drawing of each nacelle is given in figure 1.\n\n1", "timestamp": "2026-07-19T18:40:52.917417+00:00"}
{"citation_id": "19930091655", "source_url": "https://ntrs.nasa.gov/api/citations/19930091655/downloads/19930091655.pdf", "page_number": 12, "total_pages": 22, "image_filename": "19930091655_p12.jpg", "text": "8\nREPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nPRECISION OF RESULTS\n\nThe reproducibility of the experimental results depended upon the nonvariation of the fuel quantity and the indicator calibration. The maximum deviations in the observed data for apparently identical conditions amounted to roughly $\\pm 5$ percent of the average values. The variation in fuel weights, for apparently identical injection pressures, was approximately the same for all weights and amounted to about $\\pm 3.5$ percent of the actual weight for the lower injection pressures or to $\\pm 1$ percent for the higher.\n\nweight, prevented the use of the larger fuel weights. The objectionable feature of the shift arose from the fact that a given deflection before and after a particular test did not correspond to equivalent pressures. A possible error of perhaps $\\pm 5$ percent may arise in this way; at $350^\\circ$ C. the error is undoubtedly greater.\n\nWhen the data from the records were evaluated, some personal error was introduced, particularly for the A-B portion of the curve wherein the interval is more or less arbitrary and the distances on the record are often too small to be accurately measured. The magnitude of this uncertainty is shown in figure 9, for which the\n\n<!-- Image (133, 263, 886, 636) -->\n\nFIGURE 5.—Effect of nozzle design on pressure drop. Diesel fuel; fuel weight, 0.284 gram; gas-fuel ratio, 30; gas density, 14.19 grams per liter; gas temperature, 250° C.\n\nAt temperatures of $250^\\circ$ C. and above, the indicator showed a decided tendency to change its zero point as a result of creeping of the diaphragm, particularly during calibration when the deflection period was relatively great. The extent of this shift increased with the amount of deflection, the time of deflection, and the temperature. The deflection interval was diminished as much as possible during calibrations by a quick application and release of the gas pressure. The zero point immediately after deflection was taken as the proper basis for calibration in spite of its tendency in many cases to drift back toward its original position. At $350^\\circ$ C. this restoration was less evident and the shift assumed serious proportions. This fact, together with the increased deflection per unit fuel\n\ndata were taken by two observers from the same records. The individual deviations are rather great, but the mean curves seem to fit either set of data equally well. At $150^\\circ$ C. the records were so flat in the neighborhood of the minimum point that C was taken as the center of the flat portion of the curve. For the larger deflections the trace near the minimum point contained a wave of relatively low frequency. An average of the amplitudes of the first cycle was applied as a negative correction to compensate for this wave.\n\nOne other point of incidental interest is the change in fuel temperature as a result of the injection process. The passage of the fuel through the nozzle would ordinarily result in a small decrease in temperature on the basis of the Joule-Thomson effect (reference 24), assum-", "timestamp": "2026-07-19T18:40:53.234879+00:00"}
{"citation_id": "19930094499", "source_url": "https://ntrs.nasa.gov/api/citations/19930094499/downloads/19930094499.pdf", "page_number": 2, "total_pages": 13, "image_filename": "19930094499_p2.jpg", "text": "NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nTECHNICAL MEMORANDUM NO. 917\n\nTHE EFFECT OF COMPRESSIBILITY ON THE PRESSURE READING\nOF A PRANDTL PITOT TUBE AT SUBSONIC FLOW VELOCITY*\n\nBy O. Walchner\n\nSUMMARY\n\nThe effect of compressibility of a flow on the pressure reading of a normal Prandtl pitot tube (fig. 2) was investigated while the air speed approaches the velocity of sound. Errors arising from yawed flow were also determined up to 20° angle of attack. In axial flow, the Prandtl pitot tube begins at w/a = 0.6 to give an incorrect static pressure reading, while it records the tank pressure correctly, as anticipated, up to sonic velocity. Figures 5 to 7 illustrate the recorded pressure errors for different angles of attack. If errors up to within ±1 percent are permissible in the speed prediction, the pressure difference $p_1 - p_2$ recorded by the Prandtl pitot tube can be evaluated through the equation\n\n$$\np_1 - p_2 = \\frac{\\rho}{2} w^2 \\left[ 1 + \\frac{1}{4} \\left( \\frac{w}{a} \\right)^2 \\right]\n$$\n\nwhereby the range of validity extends to w/a ≤ 0.95 and yawed flow up to 10°. The equation is plotted for standard atmosphere in figure 9.\n\nOwing to the compressibility of the air, the Prandtl pitot tube manifests compression shocks when the air speed approaches velocity of sound. This affects the pressure reading of the instrument. Because of the increasing importance of high speed in aviation, this compressibility effect is investigated in detail. (The results of similar investigation by Panetti (reference 1) are not directly comparable with the present findings because his instrument was not of normal dimensions.)\n\n*Über den Einfluss der Kompressibilität auf die Druckanzeige eines Prandtl-Rohres bei Strömungen mit Unterschallgeschwindigkeit.\" Jahrbuch 1938 der deutschen Luftfahrtforschung.", "timestamp": "2026-07-19T18:40:53.838759+00:00"}
{"citation_id": "19930094544", "source_url": "https://ntrs.nasa.gov/api/citations/19930094544/downloads/19930094544.pdf", "page_number": 34, "total_pages": 43, "image_filename": "19930094544_p34.jpg", "text": "N.A.C.A. Technical Memorandum No. 872\nFigs. 13,14,15,16,17,18\n\n[Figure: A photograph of a large, cigar-shaped airship with the text \"US NAVY ZMC-2\" on its side. It is tethered to the ground with people visible below.]\nFigure 13.- Metalclad pressure airship ZMC-2.\n\n[Figure: A photograph looking down into the interior of a large, circular, metal structure with concentric rings and radial supports.]\nFigure 14.- ZMC-2, inside view.\n\n[Figure: A photograph of a large, cigar-shaped airship with the text \"GOOD YEAR\" on its side.]\nFigure 15.- Goodyear pressure ship PURITAN.\n\n[Figure: A diagram showing a side view and a cross-section of an airship's internal frame. The side view is labeled with \"HR\", \"ZR\", and \"L\".]\nFigure 16.- Usual system of airship framing. HR = main ring. ZR = intermediate ring. L = Longitudinal girder.\n\n[Figure: A wireframe diagram showing the skeletal structure of an airship.]\nFigure 17.- UNGER system.\n\n[Figure: Five circular diagrams showing different cross-sectional designs of airship rings. They are labeled \"Great Zeppelin\", \"LZ 128\", \"R 100\", \"Akron\", and \"R 101\".]\nFigure 18.- Assembly of ring types.", "timestamp": "2026-07-19T18:41:01.529767+00:00"}
{"citation_id": "19930091748", "source_url": "https://ntrs.nasa.gov/api/citations/19930091748/downloads/19930091748.pdf", "page_number": 4, "total_pages": 12, "image_filename": "19930091748_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": "2026-07-19T18:41:13.903341+00:00"}
{"citation_id": "19930094493", "source_url": "https://ntrs.nasa.gov/api/citations/19930094493/downloads/19930094493.pdf", "page_number": 3, "total_pages": 31, "image_filename": "19930094493_p3.jpg", "text": "2\nN.A.C.A. Technical Memorandum No. 923\n\nI. THE TEST SET-UP\n\na) Method of Measurement with the Hot Wire\n\nA thin metal wire was employed for measuring the air\nvelocity in the cylinder. The method is based on the fact\nthat the cooling of an electrically heated wire by the air\nincreases with increasing velocity of the latter. To each\nair velocity there will correspond a definite wire temper-\nature. Since the dimensions of the wire can, within certain\nlimits, be kept very small, the measurements will be prac-\ntically free from inertia lag. With different arrangement\nand length of the wire mean values may also be determined\nfor various measuring cross sections.\n\n1. Principles of Velocity Measurement with the Hot Wire\n\nIn tests carried out at the heat engine laboratory at\nthe Munich Technical High School, J. Ulsamer measured the\nair-flow velocities in the cylinder of an air compressor.\nThe general principles of his method will be described\nbriefly here.\n\nAn electrically heated metal wire is situated in an\nair stream. In the condition of equilibrium, the electri-\ncal energy supplied to the wire must be equal to that\ntransferred to the air from the surface of the wire. Any\nperiodical variation in the air stream must naturally be\nfollowed by a variation in the heat transferred from the\nwire surface to the air and hence also in the electrical\nenergy supplied to the wire, provided that the wire dimen-\nsions are sufficiently small for it to follow the periodic\nchanges with practically no inertia lag. Denoting the heat\nenergy in heat units supplied to the wire in unit time by\nU and the heat yielded in the same time interval to the\nair stream by Q, then in the equilibrium state the fol-\nlowing equation must be satisfied:\n\n$$U = Q \\quad (1)$$\n\nThe supplied energy U is given by\n\n$$U = 0.86 \\ i^2 \\ r \\quad (2)$$\n\nwhere i is the current in amperes through the measuring\nwire\n\nand r, the resistance in ohms.", "timestamp": "2026-07-19T18:41:18.214372+00:00"}
{"citation_id": "19930091701", "source_url": "https://ntrs.nasa.gov/api/citations/19930091701/downloads/19930091701.pdf", "page_number": 12, "total_pages": 18, "image_filename": "19930091701_p12.jpg", "text": "8\nREPORT NO. 626—NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\n<!-- Image (149, 90, 871, 883) -->\n\nFIGURE 9.—Weight, 2,000 pounds.\n\nFIGURE 10.—Weight, 2,378 pounds.\n\nFIGURE 11.—Weight, 2,800 pounds.\n\nFIGURES 9 TO 11.—Variation of take-off distances with take-off speed for the Verville AT airplane.", "timestamp": "2026-07-19T18:41:21.666754+00:00"}
{"citation_id": "19930094498", "source_url": "https://ntrs.nasa.gov/api/citations/19930094498/downloads/19930094498.pdf", "page_number": 3, "total_pages": 37, "image_filename": "19930094498_p3.jpg", "text": "```markdown\n2\nN.A.C.A. Technical Memorandum No. 918\n\nThese models were of craft that were not true water craft\nat all but used the water only as a means of temporary sup-\nport when at rest, when landing, or when taking off into\ntheir proper element - the air. When the craft was in mo-\ntion, the load on the water varied; in a take-off it changed\ncontinually from a maximum to zero as the machine passed\nfrom one element to the other. The water resistance varied\nfrom zero to a maximum and back to zero as the speed in-\ncreased, and could not be divided into frictional resistance\nand wave-making resistance in any simple manner because the\nwetted surface varied continually through the take-off run.\nFortunately, the full-size aircraft were not very large -\ncompared with ships - and models of moderate size would rep-\nresent them to relatively large scale, and thus would tend\nto reduce any difficulties from scale effects.\n\nIt was clear that the new tank must have a towing car-\nriage capable of high speed and that, in consequence, the\ntank must be much longer than the usual ship tank and much\ngreater power must be provided to propel the carriage at\nthe greater speed. More powerful devices for stopping the\ncarriage at the end of the run must be installed and in or-\nder to avoid sliding wheels and damage to tires and rails,\nadditional length must be provided for starting and stopping.\n\nThe most influential factor of all, however, was the\nrequirement that the cost must be rigidly restricted because\nthe funds available were very limited. This restriction\nmeant that attention - and money - must be concentrated on\nthe absolutely essential features of basin and carriage and\nthat everything else must be reduced to the barest minimum.\n\nA workable solution of the problem just outlined was\nobtained by devising methods of construction, types of\nequipment, and means of operation that had never before been\nused in towing basins. Among the novel features of the orig-\ninal N.A.C.A. tank, which is described quite fully in refer-\nence 1, were:\n\n1. A basin 1980 feet long, intended to give the towing\ncarriage a sufficient length of run to take usable\nreadings at its maximum speed.\n\n2. Running rails made of structural H beams which, although\nnot machined, nevertheless gave a sufficiently smooth\nsurface for the tires of the towing carriage.\n\n3. A towing carriage that had a maximum speed of 60 miles\nper hour (88 ft. per sec.) and hence could tow large\nmodels of seaplane floats at speeds corresponding to\nhigh get-away speeds.\n```", "timestamp": "2026-07-19T18:41:30.005844+00:00"}