PDF/ocr_dataset_108.jsonl

{"citation_id": "19930091655", "source_url": "https://ntrs.nasa.gov/api/citations/19930091655/downloads/19930091655.pdf", "page_number": 16, "total_pages": 22, "image_filename": "19930091655_p16.jpg", "text": "```markdown\n12\nREPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nFor the lowest density the impinging-jets nozzle gave a heat-transfer rate similar to the 13-orifice nozzle but at the highest density its rate was substantially less. The single-orifice nozzle gave a smaller rate at all densities. The maximum cylinder pressures obtained with similar nozzles and the N. A. C. A. combustion apparatus show the same trends, indicating that better initial mixing of fuel and air, together with the resulting improvement in heat transfer, occurs with the high-dispersion nozzles (reference 3). No information relevant to the effect of vapor concentration on the ignition lag can be obtained from such an engine study, presumably because all nozzles giving at least moderate fuel dispersion permit the optimum air-vapor mixture somewhere within the spray and thus give approximately the same ignition lag.\n\nThe contribution of radiation to the total heat-transfer rate can be shown to be negligible on the basis of the treatment given in an earlier section. If record 418 is considered to be typical of the others, the rate of pressure drop equivalent to the maximum rate of radiation that could occur is only 1 percent of the observed rate. It appears that radiation contributes little toward heating the fuel injected into an engine except for the possibility of unvaporized fuel becoming surrounded by a cloud of radiating combustion products. Even in this case, the conductive heat exchange can be shown to predominate if its rate per degree temperature difference remained constant and independent of the gas temperature to the extent indicated by the data in column 10 of table I.\n\n**Effect of temperature on initial heat transfer.**—Straight lines seem to agree with the data plotted in figure 9 within the limits of the uncertainty involved and, moreover, such lines are in agreement with a rate of heat transfer directly proportional to the temperature difference. The fact that the lines are straight indicates that the gas temperature has little influence on spray development within the range employed (reference 28), measured in this case by the effective area available for heat transfer. This area appears to be constant for a given density and fuel weight; otherwise a compensating change in the heat-transfer coefficient must be assumed. There is no indication that the slopes of the lines of figure 9, and hence the corresponding heat-transfer coefficients, will assume different values at the higher temperatures attained in an engine. The extrapolation, however, is too great to be of more than qualitative interest. The mass flow of gas inherent in an engine (reference 4) would lead to greater effective transfer areas and thus increase the apparent rate of pressure decrease indicated in this figure.\n\nThe ratios of the initial slope values given in column 9 of tables I and III to the respective products of fuel weight and initial fuel-gas temperature difference give a fundamental basis for comparing the relative efficacy of the heat transfer in all cases for a given ambient gas. It follows from such ratios that the rate of heat transfer varies directly with the initial temperature difference, as stated earlier in connection with figure 9. Increasing the weight of Diesel fuel leads to considerable decrease in these values but with benzene the tendency is not so evident. This difference indicates that the effective heat-transfer area is more nearly proportional to the fuel weight for benzene than for Diesel fuel. Again, as with the initial slopes, these ratios are somewhat greater for benzene, but it is not known whether this situation arises from a greater heat requirement or from better spatial distribution of the spray. The latter seems most probable in view of the effect of fuel viscosity on the distribution of fuel within the spray (references 2 and 29).\n\n**Fuel vaporization.**—The records reproduced in figure 6 show that some evaporation of the fuel occurs during the A-B interval. If all the heat transferred served merely to heat the liquid fuel, it is evident that the initial rate of heat transfer should not decrease as it does in these records. As more and more fuel is injected into the same gas charge, thermal equilibrium being reestablished before each injection, such a condition is approximated as the partial pressure of the vapor and the saturation pressure of the liquid approach one another. Certainly the relatively small molecular concentrations of vapor that produce the diminutions in initial heat-transfer rate evident even after a single injection can only be effective in the observed manner by retarding the evaporation of the fuel. These records show that the heat transferred to the vapor or to the fuel in effecting vaporization represents an appreciable part of the total heat transferred to an ordinary spray during the A-B interval. Rothrock and Waldron (reference 8) have presented conclusive evidence that considerable vaporization does occur in a high-speed engine but the rate, of course, is indeterminate as in the present case. The speed of the engine proved to be influential, presumably for two reasons: differences in mechanical mixing of the spray with the air, and certain changes in the thermal boundary conditions of the spray. Photographs in reference 30 of sprays injected into cold and heated air show a distinct decrease in the spray penetration with the hot air. It is quite probable that vaporization of the fuel within the spray envelope contributed to this decrease in addition to the changes in fuel temperature and air viscosity, which were cited in explanation of this phenomenon.\n\nTHE A-C INTERVAL\n\n**Effectiveness of heat transfer.**—Even though the B-C portion of the A-C interval has no particular connection with engine operation, it does present some information of interest on the effectiveness of the heat transfer. This effectiveness is shown most readily by comparing with the actual pressure drop the calculated pressure drop that should take place if all the fuel had vaporized. The nearer the experimental value approaches the calculated value the greater the effective-\n```", "timestamp": "2026-07-19T18:45:53.868418+00:00"}
{"citation_id": "19930094506", "source_url": "https://ntrs.nasa.gov/api/citations/19930094506/downloads/19930094506.pdf", "page_number": 3, "total_pages": 24, "image_filename": "19930094506_p3.jpg", "text": "NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nTECHNICAL MEMORANDUM NO. 910\n\nMEASUREMENTS ON A LOW-WING MODEL IN THE ROTATING JET\nAND COMPARISON WITH FLIGHT MEASUREMENTS*\n\nBy W. Bader\n\nSUMMARY\n\nThe present report deals with six-component measurements in the small tunnel of the DVL on a model of the BFW-M 27b₁, which were made to determine the effect of rolling and yawing on the air forces and moments. The experiments were carried out in a rotating air stream. The wind was given a spiral motion by means of a rotating screen, the model being suspended in the conventional manner.\n\nFrom the findings, the following points are of special interest:\n\n1) With markedly increasing angles of yaw the maximum lift shifts to higher angles of attack;\n\n2) At lower angles of attack the drag is reduced during rolling;\n\n3) In the stalling range the drag is increased during rolling;\n\n4) At high angles of attack the lateral force shows a reversal of sign; the effect of rolling on the lateral force is quite considerable;\n\n5) The pitching moment appears to be relatively independent of rolling;\n\n6) The yawing-rolling moment is very high in the region well beyond stalling.\n\n*\"Messungen an einem Tiefdeckermodell in rotierenden Strahl und ihr Vergleich mit Flugmessungen.\" Luftfahrtforschung, Bd. 16, Lfg. 2, pp. 104-111.", "timestamp": "2026-07-19T18:45:58.216865+00:00"}
{"citation_id": "19930093641", "source_url": "https://ntrs.nasa.gov/api/citations/19930093641/downloads/19930093641.pdf", "page_number": 31, "total_pages": 47, "image_filename": "19930093641_p31.jpg", "text": "L-4152\n\nN.A.C.A.\nFig. 7\n\n-223 1/2\"\n\nA\n-34\"\n-47 1/2\"\n\nPusher position\n\n-10 1/2\"\n-100 3/4\"\n\n3\n2\n1\nTractor positions\n\n-1 1/2\"\n-2 1/2\"\n-3 1/2\"\n\nA\n\nEnclosed-engine arrangements\n\nExtension-shaft housing, 4\" dia.\nElectric motor\nEngine cowling, 8 1/2\" dia.\n\n-2 5/8\"\n\nSection A-A\n\nB\n-37\"\n-44 1/2\"\n\nC.g. 0.45 in.\nabove T.E. of\nroot chord\n\n-60 7/8\"\n-60 1/4\"\n\n-60\"\n-27\"\n-30\"\n\nB\nConventional nacelle-type tractor\n\n-46\"\nElectric motor\n\nSection B-B\n\nFIG. 7.-DIAGRAM OF MODEL.", "timestamp": "2026-07-19T18:46:06.640187+00:00"}
{"citation_id": "19930093281", "source_url": "https://ntrs.nasa.gov/api/citations/19930093281/downloads/19930093281.pdf", "page_number": 9, "total_pages": 22, "image_filename": "19930093281_p9.jpg", "text": "```markdown\n7\n\nGENERAL DISCUSSION\n\nThe results of tests without propellers show that the increase in the drag coefficient due to replacing a streamline nose with an open-nose N.A.C.A. cowling is equal to 0.0081. The propeller tests were made with a 10-foot-diameter propeller and a 52-inch-diameter nacelle, which gives a value of $F/S = 0.188$. The maximum power that can be efficiently utilized with a 10-foot-diameter propeller at a speed of 300 miles per hour is approximately 750 horsepower. These conditions give a value of $1/\\sqrt[3]{F_C} = 2.68$. From figure 12(b), the value of $\\Delta C_D$ for nose 1 at $1/\\sqrt[3]{F_C} = 2.68$ is 0.0094. This value of $\\Delta C_D$ includes the effect of the nose opening and the change in the propeller efficiency caused by exposing the propeller hub and the round blade shanks. The change in $\\Delta C_D$ caused by shielding the hub and the blade shanks with spinner 1 on nose 5 is equal to 0.0033. A similar application of a spinner with nose 1 would result in a reduction of $\\Delta C_D$ from 0.0094 to approximately 0.008, the value obtained by the drag tests.\n\nFrom the definition of $\\Delta C_D$ in terms of propeller efficiency, a $\\Delta C_D$ of 0.008 gives a change in propeller efficiency of 2.9 percent at $1/\\sqrt[3]{F_C} = 2.68$ and $F/S = 0.188$. If the same $\\Delta C_D$ were applied at the same value of $1/\\sqrt[3]{F_C}$ to a 14-foot-diameter propeller and a 52-inch-diameter nacelle, the percentage change of propeller efficiency would be 1.5. This example would apply to a 1,470-horsepower engine and a speed of 300 miles per hour.\n\nThis same result may be calculated from the drag results in the following manner. A 52-inch-diameter cowling in an air stream of 300 miles per hour with a drag coefficient of $\\Delta C_D = 0.008$ absorbs 22 horsepower. This power amounts to 2.9 percent of the engine power for a 750-horsepower engine or to 1.5 percent for a 1,500-horsepower engine.\n\nAs stated previously, about 1 percent of the engine power is required for internal work in cooling the rear of the engine cylinders. Since the open-nose cowling provides sufficient cooling for the front of the cylinders, its aerodynamic power cost should be credited with 1 percent\n```", "timestamp": "2026-07-19T18:46:07.397424+00:00"}
{"citation_id": "19930094568", "source_url": "https://ntrs.nasa.gov/api/citations/19930094568/downloads/19930094568.pdf", "page_number": 9, "total_pages": 28, "image_filename": "19930094568_p9.jpg", "text": "8 N.A.C.A. Technical Memorandum No. 848\n\nAfter this digression, we return to our discussion of figures 2 to 4.\n\nAccording to Wagner's theoretical reasoning, the drag (angle of attack) of wide and short plates and given loading increases with increasing gravity effect (decreasing Froude number). The experiments (figs. 2-4) actually manifest this effect on short plates (at least up to $l/b = 2$) at all load stages. In fact, these short plates disclosed, even at the highest Froude numbers reached in the test, a change in angle of attack (drag) with the Froude number.\n\nAs regards very long plates ($\\frac{l}{b} = 3$) the gravity appears to have a drag decreasing rather than increasing effect ($\\beta =$ approximately constant), according to figures 2 to 4. At least, this holds true for the highest Froude numbers reached in the test, which is in line with Sottorf's results (figs. 2-4). At very low speeds the loading of the planing surfaces is finally borne free from drag by the buoyant lift. But it is precisely the long plates which, in view of the persistence of the gravity on the short plates at equal Froude numbers, seem to raise the doubt as to whether the tests actually correspond already to the gravity-free problem.\n\nBut that this is actually the case is suggested from the airfoil data plotted at $\\frac{1}{F^2} = 0$, provided these values themselves are correct and not, perhaps, afflicted with an appreciable error due to chamfering of the leading and trailing edges, or caused by the conversion of the test data (figs. 26 and 27).\n\nIn order to bring out the accord between the planing surface and airfoil data even more clearly, figure 7 shows the lift coefficients $c_a = \\frac{R}{\\frac{\\rho}{2} v^2 b l}$ against angle of attack for the three load stages, the region of maximum Froude numbers reached in the test $\\overline{F} = 13.05$ and $\\overline{F} = 3.47$ being shown as shaded area. (It corresponds in figs. 2-4 to the range between $V = 22.6$ and 6 meters per second.) The graph also includes half the lift coefficients $c_a$ of flat planing surfaces with the aid of figure 29. The good agreement extends far beyond that predicted by Wagner for infinitely small angles of attack and includes, in fact, a considerable range of Froude numbers.", "timestamp": "2026-07-19T18:46:14.140673+00:00"}
{"citation_id": "19930094544", "source_url": "https://ntrs.nasa.gov/api/citations/19930094544/downloads/19930094544.pdf", "page_number": 37, "total_pages": 43, "image_filename": "19930094544_p37.jpg", "text": "N.A.C.A. Technical Memorandum No. 872\nFigs. 25,26,27,28\n\n[Figure: Tip of the bow with mooring spindle]\nFigure 25.- AKRON - Tip of the bow with mooring spindle. The mooring spindle is at the tip of the bow and in the middle of the background a cruciform ring is seen. The mooring cone, here still lacking, hangs from the tip of the spindle.\n\n[Figure: Assembly view of the framing]\nFigure 26.- R-100 - Assembly view of the framing. The great ring and longitudinal spacings, as well as the single-panel bracing, are noteworthy.\n\n[Figure: Partial view showing cell and ring bracing]\nFigure 27.- R-100 - Partial view showing cell and ring bracing. The ring bracing is distinctly marked on the end of the cell. The axial girder seen above supports the wire netting at the center and is inclosed by the gas cell.\n\n[Figure: Inside view at the bow]\nFigure 28.- R-100 - Inside view at the bow. In the foreground the ramie cord net is visible between the ring wires.", "timestamp": "2026-07-19T18:46:14.660519+00:00"}
{"citation_id": "19930091724", "source_url": "https://ntrs.nasa.gov/api/citations/19930091724/downloads/19930091724.pdf", "page_number": 8, "total_pages": 20, "image_filename": "19930091724_p8.jpg", "text": "4 REPORT NO. 649 NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\n[Figure: Diagram of a mechanical test apparatus labeled with letters: I, U, X, N, T, V, A, R, F, B, D, W, Z, H, G, Y]\n\nFIGURE 4.—“Pack” ready for test.", "timestamp": "2026-07-19T18:46:15.445404+00:00"}
{"citation_id": "19930091748", "source_url": "https://ntrs.nasa.gov/api/citations/19930091748/downloads/19930091748.pdf", "page_number": 8, "total_pages": 12, "image_filename": "19930091748_p8.jpg", "text": "4\nREPORT NO. 673—NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\n[Figure: Typical photographic record obtained with oscillating-airfoil balance.]\n\nFIGURE 3.—Typical photographic record obtained with oscillating-airfoil balance.\n\n| v, f.p.s | $C_l$ |\n| :--- | :--- |\n| $\\square$ 23.0 | 0 |\n| $+$ 114.5 | 0 |\n| $\\triangle$ 23.0 | .74 |\n| $\\circ$ 46.4 | .74 |\n\nTheory for $A=12.5$\n$\\infty$\n\nPhase difference, $\\delta$, deg.\n100\n80\n60\n40\n20\n0\n-1 Lead\n0 4 8 12 16 20 24 28\n$1/k$\n\nFIGURE 4.—Comparison of experimental and theoretical values of the phase difference $\\delta$ for various values of the parameter $1/k$.\n\n1.0\n.8\n.6\n.4\n.2\n0\n0 4 8 12 16 20 24 28\n$1/k$\n\n$A=\\infty$\n$A=12.5$\n$F$\n$G$\n$A=\\infty$\n$A=12.5$\n\nFIGURE 6.—Theoretical values of circulation functions $F$ and $G$.\n\n1.4\n1.2\n1.0\n.8\n.6\n.4\n.2\n0\n-.2\n0 4 8 12 16\nAngle of attack, $\\alpha$, deg.\n\nLift coefficient, $C_L$\nLift at fixed angles\nAngle suddenly increased\n\nFIGURE 5.—Lift curves for the airfoil tested in steady motion at fixed angles and for a test in which the angle of attack was suddenly increased.", "timestamp": "2026-07-19T18:46:18.941469+00:00"}
{"citation_id": "19930094499", "source_url": "https://ntrs.nasa.gov/api/citations/19930094499/downloads/19930094499.pdf", "page_number": 6, "total_pages": 13, "image_filename": "19930094499_p6.jpg", "text": "N.A.C.A. Technical Memorandum No. 917 5\n\nThis equation is known as the discharge formula for the case where a flow with velocity $w$, pressure $p$, and density $\\rho$ is produced through adiabatic discharge from a tank in which the medium rests under pressure $p_0$.\n\nThen the introduction of the velocity of sound in the free jet\n\n$$\na = \\sqrt{\\frac{d}{d} \\frac{p}{\\rho}} = \\sqrt{\\kappa \\frac{p}{\\rho}}\n\\tag{6}\n$$\n\ngives the Mach number\n\n$$\n\\frac{w}{a} = \\sqrt{\\frac{2}{\\kappa - 1} \\left[ \\left( \\frac{p_0}{p} \\right)^{\\frac{\\kappa - 1}{\\kappa}} - 1 \\right]}\n\\tag{7}\n$$\n\nThus, $w$, $\\rho$, $a$, $p$, and $\\kappa$ denote the state quantities of undisturbed flow for the Prandtl pitot tube, the dimensions of which are given in figure 2. The pressure at the forward orifice is hereafter indicated with $p_1$ and the pressure at static orifice on the side with $p_2$. For determining the effect of yawed flow, the tube could be turned in the sense of the arrow in figure 1.\n\nThe measurements included the incorrect readings caused by compressibility effects and yawed flow, i.e., the pressure differences $p_0 - p_1$ and $p - p_2$ over a speed range from around 0.55 times sonic velocity up to near the velocity of sound. The pressure field in the vicinity of the forward part of the Prandtl tube was treated by the Schlieren method and photographed.\n\nRESULTS\n\nThe first appearance of a shock wave was recorded by the Schlieren method at a Mach number of $w/a \\approx 0.7$. The intensity of the shock increases very little at first by an increasing Mach number. Figure 3 is a Schlieren record at $w/a = 0.88$. The dark area denotes a compression of the medium in flow direction and the light areas an expansion. Note the compression toward the stagnation point followed by expansion at circulation about the head. Behind this expansion a compression follows again. The surprising fact now is the location of the shock wave in the middle of the", "timestamp": "2026-07-19T18:46:20.090163+00:00"}
{"citation_id": "19930093640", "source_url": "https://ntrs.nasa.gov/api/citations/19930093640/downloads/19930093640.pdf", "page_number": 6, "total_pages": 28, "image_filename": "19930093640_p6.jpg", "text": "4\n\n$$\nT_{c_o}' = \\frac{P_{tot.} \\eta_o}{\\frac{1}{2} \\rho V^3 S} = \\text{index thrust coefficient.}\n$$\n\n$\\eta_o = \\eta$ at $C_L = 0.25$.\n\nn, propeller revolution speed, r.p.s.\n\nD, propeller diameter, ft.\n\n$\\beta$, propeller blade angle at 0.75 R, deg.\n\n$\\delta_f$, flap deflection from closed position, deg.\n\na, slope of lift curve, $dC_L/da$, deg.\n\nMODEL AND TEST EQUIPMENT\n\nThe tests were conducted in the N.A.C.A. full-scale wind tunnel, a description of which is given in reference 2.\n\nThe model was a metal-covered, midwing monoplane with a span of 37.25 feet. The wing sections were symmetrical and tapered in thickness from 0.18c at the root to 0.10c at the tip. The wing had a plan form tapered 4:1, with a root chord of 7.28 feet and an area of 172 square feet. Split trailing-edge flaps with an average chord of 0.15c extended over the middle 60 percent of the span with the exception of a short gap at the fuselage. The angle of wing setting to the fuselage reference line was 4.6°. A line diagram of the model with dimensions of the various nacelle-propeller arrangements tested is shown in figure 4.\n\nFour 3-blade aluminum alloy model propellers were used throughout the tests. Blade dimensions and sections for the propellers are given in figure 5. Each propeller was driven by a 25-horsepower squirrel-cage induction motor, the speed of which was regulated by varying the frequency. The propeller speed was measured with a Weston electrical tachometer. Propeller torques were determined from an electrical calibration of the motors.\n\nPerforated metal plates were used to simulate the radiators for the liquid-cooled engine installation, and the engines for the air-cooled engine installations. The plates simulating the radiators were proportioned to have", "timestamp": "2026-07-19T18:46:28.413109+00:00"}
{"citation_id": "19930091731", "source_url": "https://ntrs.nasa.gov/api/citations/19930091731/downloads/19930091731.pdf", "page_number": 5, "total_pages": 32, "image_filename": "19930091731_p5.jpg", "text": "REPORT No. 656\n\nTHE COLUMN STRENGTH OF TWO EXTRUDED  \nALUMINUM-ALLOY H-SECTIONS  \n\nBy WILLIAM R. OSGOOD and MARSHALL HOLT\n\nSUMMARY\n\nExtruded aluminum-alloy members of various cross sections are used in aircraft as compression members either singly or as stiffeners for aluminum-alloy sheet. In order to design such members, it is necessary to know their column strength or, in the case of stiffeners, the value of the double modulus, which is best obtained for practical purposes from column tests.\n\nColumn tests made on two extruded H-sections are described, and column formulas and formulas for the ratio of the double modulus to Young’s modulus, based on the tests, are given.\n\nINTRODUCTION\n\nExtruded aluminum-alloy members of various cross sections are used in aircraft as compression members either singly or as stiffeners for aluminum-alloy sheet. In order to design such members, it is necessary to know their column strength, or in the case of stiffeners, the value of the double modulus (references 1 and 2), which is best obtained for practical purposes from column tests.\n\nThe interest of the National Advisory Committee for Aeronautics in stiffened-sheet construction as applied to monocoque design led to the allotment of funds to the National Bureau of Standards for research in this field, and a part of these funds was used to investigate the column strength of an extruded aluminum-alloy shape comparable with those used in stiffened-sheet construction. The data obtained in the tests made at the National Bureau of Standards are presented and discussed in part I of this report. The material for this investigation was supplied by the Aluminum Company of America.\n\nColumn tests were conducted at the Aluminum Research Laboratories on pieces of extruded aluminum alloy taken from the same lot of material supplied to the National Bureau of Standards. Column tests were also made at the Aluminum Research Laboratories on another extruded aluminum-alloy shape, the data on which had been requested by the National Advisory Committee for Aeronautics. The results of these tests are presented and discussed in part II of this report.\n\nA correlation of the test data from the National Bureau of Standards and those from the Aluminum Research Laboratories is made in part III of this report.\n\nMATERIAL\n\nThe material used in these investigations of column strength is designated Aleoa 24S-T by the Aluminum Company of America and complies with Navy Department Specifications 46A9a, June 1, 1938: Aluminum-alloy (aluminum-copper-magnesium (1.5 percent)-manganese): Bars, Rods, Shapes, and Wire. The material was furnished in the form of extruded H-beams. The nominal dimensions of the cross sections are shown in\n\n[Figure: Cross Section A and Cross Section B diagrams with labeled dimensions]\n\nFIGURE 1.—Dimensions of extruded 24S-T H-sections.\n\nfigure 1. The National Bureau of Standards tests were made only on cross section A and the Aluminum Company tests included both cross sections.\n\nThe mechanical tests to determine properties and the results of these tests are discussed in the following three parts of this report.\n\nI. TESTS MADE AT THE NATIONAL BUREAU OF STANDARDS\n\nPREPARATION OF SPECIMENS\n\nThe tensile specimens were three standard type-5 tensile-test specimens, as defined in Navy Department General Specifications for Inspection of Material, Appendix II (Metals). They were cut from the same length of extruded shape, one specimen from the middle of the web and the other two from diagonally opposite positions in the two flanges. The cross-sectional areas of the reduced portions of these specimens were determined by calipering them.", "timestamp": "2026-07-19T18:46:29.291438+00:00"}
{"citation_id": "19930094549", "source_url": "https://ntrs.nasa.gov/api/citations/19930094549/downloads/19930094549.pdf", "page_number": 29, "total_pages": 76, "image_filename": "19930094549_p29.jpg", "text": "N.A.C.A. Technical Memorandum No. 867 27\n\nThe coefficient of effectiveness of the elevator for negative deflections was:\n\n$$\n\\frac{dC_H}{d\\beta} = 0.01308\n$$\n\nThe conditions of equilibrium, according to the tunnel tests, show that the flight at $C_z = 0.458$ and $i = 3.6^\\circ$, should be made at an angle of elevator deflection of $\\beta = -2.55^\\circ$. Actually, equilibrium was obtained at $\\beta = -3.6^\\circ$, which is a sufficiently good agreement.\n\nThe test consisted in carrying out the following maneuver: Having attained the steady state at velocity $V = 45.2$ m/s, which here is state A, the pilot pushes on the stick and lowers the elevator, the mean deflection becoming $\\beta = -1.5^\\circ$, the pilot not touching the throttle. The airplane dives and tends toward a new position of equilibrium. The pilot, however, does not wait until the final state B is attained. After having maintained the deflection $\\beta = -1.5^\\circ$ for 7 seconds, he pulls back on the stick, fixes it in its initial position and allows the airplane, after some oscillations, to return to its initial state A. The test is carried out while recording:\n\n| Symbol | Description |\n| :--- | :--- |\n| $V$ | is the velocity on the flight path. |\n| $V_y$ | vertical component of the velocity measured with a variometer. |\n| $\\beta$ | setting of the elevator. |\n| $i$ | angle of attack. |\n| $J_z$ | component of the apparent weight in the direction of the OZ axis. |\n| $J_x$ | component of the apparent weight in the direction of the OX axis. |\n| $q$ | angular velocity about the OY axis. |\n\nWe thus have 7 test curves.\n\nThe variometer is an apparatus whose readings show considerable lag, and can only be used when corrected. The curve of corrected vertical velocities is shown on one of", "timestamp": "2026-07-19T18:46:29.556286+00:00"}
{"citation_id": "19930094498", "source_url": "https://ntrs.nasa.gov/api/citations/19930094498/downloads/19930094498.pdf", "page_number": 7, "total_pages": 37, "image_filename": "19930094498_p7.jpg", "text": "6 N.A.C.A. Technical Memorandum No. 916\n\nthat supported the wheels, the motors, and the gears, and replacing it by a new structure, within which are supported the trucks carrying the new wheels, the new motors, and the new gears. The new structure was made of steel tubing of the same type and sizes used in the old structure; it was welded in itself and to the old carriage structure.\n\nTrucks.- The towing carriage now operates on eight wheels arranged in four groups of two. Each wheel is driven by a 75-horsepower electric motor through a worm and gear and is not mechanically connected to any other wheel. Each pair of wheels supports a truck that carries the two motors and supports the carriage through a pivot pin. The truck can rock freely on ball bearings on the pin and any slight irregularities in the track cause less vertical motion of the carriage itself than they did with the four-wheel arrangement. Automotive practice has been followed throughout and wheels now can be removed - for grinding tires, or repairs - about as easily as from the axles of an automobile.\n\nThis is believed to be the first use of equalized wheels arranged in trucks on carriages for towing basins, although it is realized that it reflects primarily the very special nature of the construction, equipment, and methods of work in use in the N.A.C.A. tank.\n\nThe service brakes are automotive and can also be operated by hand or by standard automotive air-brake equipment controlled by a pedal. In order to save weight, the compressed-air reservoirs are charged through a hose from a fixed compressor and reservoirs \"ashore\" and not by a compressor on the carriage.\n\nElectrical braking - regenerative and dynamic - is provided as it was on the old carriage; a track switch is provided so that when the carriage passes a certain point the dynamic braking is automatically applied. If the carriage passes another point farther on, a second track switch operates and the air brake is applied full strength as in an emergency. The last resort in braking remains the grab brakes that were fitted on the original carriage.\n\nThe possibility that a tire may fail has been provided for by fitting under each truck, but attached to the main structure, a steel roller to receive the weight of the carriage and roll on the rail. The car need sink only half an inch to bring the roller into play.", "timestamp": "2026-07-19T18:46:34.109909+00:00"}
{"citation_id": "19930094542", "source_url": "https://ntrs.nasa.gov/api/citations/19930094542/downloads/19930094542.pdf", "page_number": 91, "total_pages": 102, "image_filename": "19930094542_p91.jpg", "text": "N.A.C.A. Technical Memorandum No. 874\nFigs.78,79\n\n[Figure: Diagram showing a Wing and a Sphere. The wing has a chord length labeled 't'. The sphere is located at a distance of '2.5 t' behind the wing. The sphere is shown at two vertical positions relative to the wing chord line: 'Position I' at 0 and 'Position II' at 0.25 t.]\n\nFigure 78.- Arrangement for downwash measurements.\n\n[Figure: Two line graphs plotting $\\delta$ versus $\\frac{y}{R}$.\nTop graph title: Position I.\nBottom graph title: Position II.\nY-axis (left): $\\delta$, range 0 to 12.0°.\nY-axis (right): $\\alpha$, range -8° to 16°.\nX-axis: $\\frac{y}{R}$, range -2.0 to 2.0.]\n\nFigure 79.- Downwash behind wing without propeller.", "timestamp": "2026-07-19T18:46:57.849165+00:00"}
{"citation_id": "19930094544", "source_url": "https://ntrs.nasa.gov/api/citations/19930094544/downloads/19930094544.pdf", "page_number": 38, "total_pages": 43, "image_filename": "19930094544_p38.jpg", "text": "N.A.C.A. Technical Memorandum No. 872\nFigs. 29,30,31\n\n[Figure: Inside view of a structure with a walkway and fabric-covered walls]\n\nFigure 29.- R-100 - Inside view.\nIn the foreground the\npromenade deck of the passenger\nspace located inside the ship.\nThe walls are fabric covered.\n\n[Figure: View of a large circular ring structure with wire bracing]\n\nFigure 31.- R-101 - View of rings.\nThe three boom ring has\nrectangular panels, which are\nwire braced. The wire netting\nsurrounding the cell and its\nattachment to the lower part of\nthe ring are easily seen.\n\n[Figure: View of the bow framing of a large structure with girders and struts]\n\nFigure 30.- R-101.- View of the bow framing. Between\nthe widely spaced wire braced\nlongitudinal girders are located the numerous strut\nbraced intermediate longitudinals. These can be used\nfor final tensioning of the outer cover in the radial\ndirection.", "timestamp": "2026-07-19T18:47:09.634405+00:00"}
{"citation_id": "19930091701", "source_url": "https://ntrs.nasa.gov/api/citations/19930091701/downloads/19930091701.pdf", "page_number": 15, "total_pages": 18, "image_filename": "19930091701_p15.jpg", "text": "THE TRANSITION PHASE IN THE TAKE-OFF OF AN AIRPLANE 11\n\nCONCLUSIONS\n\n1. For normal take-offs the horizontal distances covered in the transition in proportion to the heights attained were least at the slowest possible take-off speed. Likewise, the shortest over-all distance required in taking off over an obstacle was obtained with the slowest speed.\n\n2. For normal ground conditions, zoom take-offs required shorter over-all distances than normal take-offs, particularly with heavy loads if the obstacle to be surmounted was sufficiently high. With light loadings and low obstacle heights, the zoom take-offs provided no advantage.\n\n3. The error resulting from neglect of the transition in calculating the air-borne distance in take-off varied from 8 percent with the heaviest load considered to -8 percent with the lightest load for normal take-offs over a 50-foot obstacle. For a 100-foot obstacle the percentage error was about one-half of that for the 50-foot obstacle. For zoom take-offs the error arising from neglect of the transition was much greater.\n\n4. The effect of the average wind gradient corresponding to a 5-mile-per-hour surface wind was a reduction in the air-borne distance to clear a 50-foot obstacle of about 9 percent with the lightest load and about 16 percent with the heaviest load. For the 100-foot obstacle height the reduction was about 10 percent for both loads. The over-all reduction due to this wind was approximately twice that due to the wind gradient alone. The correction of observed take-off performance to no-wind conditions can be accomplished through the use of relatively simple expressions.\n\n5. The ground effect reduced the air-borne distance required to attain a height of 50 feet by about 10 percent with the lightest loading and by about 16 percent with the heaviest loading. For an obstacle height of 100 feet the percentage reduction was about one-half as great.\n\nLANGLEY MEMORIAL AERONAUTICAL LABORATORY,\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS,\nLANGLEY FIELD, VA., October 26, 1937.\n\nREFERENCES\n\n1. Thompson, F. L., Peck, W. C., and Beard, A. P.: Air Conditions Close to the Ground and the Effect on Airplane Landings. T. R. No. 489, N. A. C. A., 1934.\n2. Hartman, Edwin P.: Working Charts for the Determination of Propeller Thrust at Various Air Speeds. T. R. No. 481, N. A. C. A., 1934.\n3. Freeman, Hugh B.: Comparison of Full-Scale Propellers Having R. A. F-6 and Clark Y Airfoil Sections. T. R. No. 378, N. A. C. A., 1931.\n\nTABLE I.—CHARACTERISTICS OF THE VERVILLE AT AIRPLANE\n\n| Engine—Continental A-70 | 165 hp. at 2,000 r. p. m. |\n| :--- | :--- |\n| Propeller—Metal, fixed pitch: | Clark Y |\n| Blade section | 8 ft. 5 in. |\n| Diameter | 12.8° |\n| Blade-angle setting at 0.75 R | |\n| Wing dimensions—Clark Y-15: | |\n| Total area | 362.5 sq. ft. |\n| Span, upper wing | 31 ft. |\n| Span, lower wing | 31 ft. |\n| Chord, upper wing | 50 in. |\n| Chord, lower wing | 50 in. |\n| Test loadings: | |\n| Gross weight | 2,950 lb. |\n| Wing loading | 7.8 lb. per sq. ft. |\n| Power loading | 12.3 lb. per hp. |\n| Gross weight | 2,378 lb. |\n| Wing loading | 9.1 lb. per sq. ft. |\n| Power loading | 14.4 lb. per hp. |\n\nTABLE II.—TAKE-OFF DISTANCES FOR THE VERVILLE AT AIRPLANE FROM STEP-BY-STEP INTEGRATIONS\n\n| Weight (lb.) | Power loading (lb./hp.) | Wing loading (lb./sq. ft.) | Take-off speed (m.p.h.) | Take-off speed (f. p. s.) | Climb-ing speed (f. p. s.) | Ground-run distance from V=75 f. p. s. (Feet) | Height attained in transition (Feet) | Horizon-tal distance for transition (Feet) | Tan-gent of angle of climb γ | Horizon-tal distance for steady climb (Feet) | Total air-borne distance (Feet) | Total distance from V=75 f. p. s. (Feet) | Horizon-tal distance for steady climb (Feet) | Total air-borne distance (Feet) | Total distance from V=75 f. p. s. (Feet) | Remarks |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| 2,950 | 12.5 | 7.8 | 75.5 | 75.5 | 7 | 74.5 | 534 | 0.1441 | | 370 | 377 | 176 | 710 | 717 | Normal. |\n| | | | 82.0 | 82.0 | 89 | 39.2 | 318 | 1.366 | 20 | 368 | 457 | 359 | 707 | 796 | |\n| | | | 90.0 | 90.0 | 98 | 29.2 | 318 | 1.366 | 78 | 362 | 600 | 444 | 758 | 866 | |\n| | | | 73.5 | 73.5 | 208 | 82.8 | 483 | 1.141 | | 287 | 374 | 121 | 604 | 692 | Zoom. |\n| | | | 90.0 | 73.5 | 208 | 101.0 | 490 | 1.441 | | 237 | 445 | | 482 | 690 | |\n| 2,378 | 14.4 | 9.1 | 81.0 | 81.0 | 98 | 46.2 | 335 | 0.934 | 40 | 631 | 629 | 122 | 1,082 | 1,101 | Normal. |\n| | | | 88.0 | 88.0 | 217 | 36.5 | 296 | 0.969 | 139 | 535 | 752 | 654 | 1,059 | 1,267 | |\n| | | | 96.0 | 96.0 | 262 | 28.6 | 335 | 0.934 | 228 | 573 | 905 | 789 | 1,114 | 1,526 | No wind. |\n| | | | 88.0 | 81.0 | 217 | 65.0 | 519 | 0.620 | | 367 | 584 | 376 | 895 | 1,112 | Zoom. |\n| | | | 96.0 | 81.0 | 262 | 88.3 | 550 | 0.610 | | 399 | 601 | 125 | 675 | 1,067 | |\n| | | | 88.0 | 88.0 | 274 | 28.0 | 526 | 0.455 | 482 | 1,008 | 1,282 | 1,581 | 2,107 | 2,381 | Normal. |\n| 2,850 | 17.0 | 10.7 | 96.0 | 81.0 | 490 | 22.7 | 432 | 0.503 | 543 | 975 | 1,465 | 1,540 | 1,972 | 2,462 | |\n| | | | 104.0 | 104.0 | 782 | 15.0 | 312 | 0.441 | 786 | 1,098 | 1,880 | 1,917 | 2,229 | 3,011 | |\n| | | | 96.0 | 88.0 | 490 | 69.5 | 542 | 0.455 | 13 | 555 | 1,045 | 1,108 | 1,650 | 2,140 | Zoom. |\n| | | | 104.0 | 88.0 | 782 | 68.7 | 509 | 0.455 | | 302 | 1,084 | 686 | 1,195 | 1,977 | |\n| 2,950 | 12.5 | 7.8 | 75.5 | 75.5 | | | | | | 390 | | | 561 | | 5 m. p. h. wind+gradient. |\n| | | | 73.5 | 73.5 | | | | | | 357 | | | 642 | | Wind gradient only. |\n| 2,850 | 17.0 | 10.7 | 88.0 | 88.0 | | | | | | 756 | | | 1,465 | | 5 m. p. h. wind+gradient. |\n| | | | 88.0 | 88.0 | | | | | | 844 | | | 1,838 | | Wind gradient only. |\n| 2,950 | 12.5 | 7.8 | 75.5 | 75.5 | | 63.3 | 495 | | | 404 | | 254 | 749 | | No ground effect. |\n| 2,850 | 17.0 | 10.7 | 88.0 | 88.0 | | 10.6 | 308 | | 865 | 1,173 | | 1,952 | 2,270 | | No ground effect. |\n\nU. S. GOVERNMENT PRINTING OFFICE: 1938", "timestamp": "2026-07-19T18:47:10.615268+00:00"}
{"citation_id": "19930091724", "source_url": "https://ntrs.nasa.gov/api/citations/19930091724/downloads/19930091724.pdf", "page_number": 9, "total_pages": 20, "image_filename": "19930091724_p9.jpg", "text": "\"PACK\" COMPRESSIVE TEST\n5\n\n[Figure: Testing machine]\n\nFIGURE 5.—Testing machine.", "timestamp": "2026-07-19T18:47:11.363673+00:00"}
{"citation_id": "19930094516", "source_url": "https://ntrs.nasa.gov/api/citations/19930094516/downloads/19930094516.pdf", "page_number": 1, "total_pages": 32, "image_filename": "19930094516_p1.jpg", "text": "FILE COPY\nNO 7\nFILE COPY\nNO 1\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\nRETURN TO THE ABOVE ADDRESS.\nREQUESTS FOR PUBLICATIONS SHOULD BE ADDRESSED\nAS FOLLOWS:\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n3224 F STREET, N.W.,\nWASHINGTON 25, D.C.\n\nNo. 901\n\nINVESTIGATIONS AND EXPERIMENTS\nIN THE GUIDONIA SUPERSONIC WIND TUNNEL\nBy Antonio Ferri\n\nHauptversammlung der Lilienthal-Gesellschaft für\nLuftfahrtforschung, Berlin, October 12-15, 1938\n\nWashington\nJuly 1939\n\nFILE COPY\nTo be returned to\nthe files of the National\nAdvisory Committee\nfor Aeronautics\nWashington, D. C.", "timestamp": "2026-07-19T18:47:11.587429+00:00"}
{"citation_id": "19930093640", "source_url": "https://ntrs.nasa.gov/api/citations/19930093640/downloads/19930093640.pdf", "page_number": 7, "total_pages": 28, "image_filename": "19930093640_p7.jpg", "text": "5\n\nthe same resistance as a standard Army Air Corps radiator of 9-inch depth. Holes were spaced in the 13-inch-diameter plate used to simulate the air-cooled engine so as to obtain a conductivity, k, of 0.124 (see reference 3), which approximates that of a twin-row radial engine. The cowling was tested with the originally designed exit slot 1-3/16 inches and the reduced slot of 3/4 inch width (fig. 4) which have been designated as large exit slot and refaired exit slot, respectively. A pressure drop across the conductivity plates of 1.29 q was measured with the large exit slots and 0.63 q with the refaired slots.\n\nTESTS\n\nWith the propellers removed from the model, measurements of forces and pitching moments were made for all the test arrangements over an angle-of-attack range from zero lift through the stall at an air speed of about 60 miles per hour. Scale effect on the drag at low lift coefficients was also measured over a range of air speeds from 30 to 120 miles per hour.\n\nWith the propellers operating, propulsive characteristics of the nacelle-propeller arrangements were determined for an angle of attack corresponding to high-speed flight. In addition to the usual aerodynamic forces and pitching moment, the measurements included the power input to the propellers and the propeller speed. The procedure followed in the propeller tests was to hold the torque constant and increase the tunnel air speed in steps from 30 miles per hour to 100 miles per hour, after which the propeller speed was reduced until zero thrust was reached. The effect of the propeller operation upon the lift and the pitching moment was determined at a tunnel speed of approximately 50 miles per hour for several thrust conditions.\n\nPOWER-OFF CHARACTERISTICS\n\nAerodynamic characteristics of the model with the propellers removed are shown in figures 6 to 13. The data shown in figures 6 to 10 were obtained at a test speed of about 60 miles per hour corresponding to a Reynolds Number of approximately 2,500,000, based on the average wing chord of 4.62 feet. The coefficients are based on a wing area of", "timestamp": "2026-07-19T18:47:23.408968+00:00"}
{"citation_id": "19930093641", "source_url": "https://ntrs.nasa.gov/api/citations/19930093641/downloads/19930093641.pdf", "page_number": 32, "total_pages": 47, "image_filename": "19930093641_p32.jpg", "text": "L-452\n\nN.A.C.A\n\n39\" DIAM.\n\nTRAILING EDGE\n\nNote: All linear dimensions\ngiven in inches.\n\nFigure 8. - Blade dimensions for\n3-blade model propellers.\n\n| 3 | 4 1/2 | 6 | 7 1/2 | 9 | 10 1/2 | 12 | 13 1/2 | 15 | 16 1/2 | 18 | 19 1/2 |\n|---|---|---|---|---|---|---|---|---|---|---|---|\n| 2 5/8 | | | | | | | | | | | |\n\nLEADING EDGE\n\n| .920 | .670 | .500 | .380 | .300 | .245 | .208 | .172 | .141 | .108 | .076 |\n|---|---|---|---|---|---|---|---|---|---|---|\n| MAX. THICKNESS OF SECTIONS | | | | | | | | | | |\n\n| 1.365 | 1.818 | 2.238 | 2.565 | 2.685 | 2.660 |\n|---|---|---|---|---|---|\n| .920 | .670 | .500 | .380 | .300 | .245 |\n| 50.5° | 47.3° | 41.4° | 34.3° | 29.9° | 26.5° |\n\n| 2.550 | 2.336 | 2.020 | 1.602 | 1.113 |\n|---|---|---|---|---|\n| .208 | .172 | .141 | .108 | .076 |\n| 23.75° | 21.5° | 19.5° | 17.8° | 16.3° |\n\nFIG. 8", "timestamp": "2026-07-19T18:47:25.169456+00:00"}
{"citation_id": "19930094549", "source_url": "https://ntrs.nasa.gov/api/citations/19930094549/downloads/19930094549.pdf", "page_number": 30, "total_pages": 76, "image_filename": "19930094549_p30.jpg", "text": "28 N.A.C.A. Technical Memorandum No. 867\n\nthe diagrams of figure 18. A knowledge of $V_{\\gamma}$ and V gives the angle of slope of the path and the angle $\\theta$ may be calculated from $\\xi$ and i. Finally, having determined $\\theta$, it is possible to calculate q and check the readings of the indicator of the angular velocity of pitch.\n\nWe shall now consider the curves distinguishing three portions:\n\n1. The flight is assumed rectilinear and uniform at state A, the deflection being -7.6°, the part of the curves to the left of point 1.\n\n2. A period of 7 seconds duration, during which the airplane undergoes maneuvering which would lead to state B if the deflection $\\beta = -1.5^\\circ$ were maintained. This period corresponds to the portion 1-2 of the curves.\n\n3. The period following the return to the setting $\\beta = -7.6^\\circ$, during which the airplane oscillates and tends to regain its initial state A - that is, the part of the curves to the right of point 2. The relation:\n\n$$\n\\frac{J_z}{g} \\frac{P}{S} = C_z \\frac{nV^2}{rg}\n$$\n\npermits us to determine $C_z$ and hence, i.\n\nFor the angle-of-attack curve the computation gave a curve which differed slightly from the one recorded, the latter lagging behind the computed curve and at the minimum angle of attack not coming down so low. It was possible for us to show that there was a systematic error in the reading of the vane due to a play of about 1.5°, a fact which explains in part the lag of the record. The Bouny accelerometer is an instrument with which we are very familiar and whose readings are very accurate. Under those conditions we think that the computed curve corresponds to that of actual angles of attack.\n\nThe curve of angular velocities has been obtained with the aid of an apparatus designed to measure the angular velocities which are produced in spinning and which are of the order of 1 to 2 radians per second. This apparatus did not have the desired sensitivity for measuring angular", "timestamp": "2026-07-19T18:47:28.076916+00:00"}
{"citation_id": "19930091697", "source_url": "https://ntrs.nasa.gov/api/citations/19930091697/downloads/19930091697.pdf", "page_number": 17, "total_pages": 28, "image_filename": "19930091697_p17.jpg", "text": "A PHOTOGRAPHIC STUDY OF COMBUSTION AND KNOCK IN A SPARK-IGNITION ENGINE 13\n\nRecords (not shown) taken with the wire exploding after the combustion crossed the chamber always showed knock occurring at the same time the wire exploded. Apparently if the charge was about ready to knock of its own accord, the shock from the exploding wire acted as a “trigger” to set off the knock.\n\nThe indicator cards (fig. 13) were taken at conditions similar to those represented by the schlieren records spark plugs on opposite sides of the chamber (E and F). In both photographs of figure 14 the two combustion fronts apparently meet and continue through each other, covering about half of the remaining distance across the chamber before merging with the general pattern of the burning. The combustion space is about 1 inch deep at this time, so there is little possibility that the two fronts passed each other at different\n\n[Figure: Five rows of schlieren photographs labeled (a) through (e), each with an associated indicator card on the left. Labels include B, A, T.C., B.T.C. 15°, T.C., 15° A.T.C.]\n\nFIGURE 13.—Indicator cards and schlieren photographs showing effect of exploding wires on combustion knock. Air-fuel ratio, 14; A, first evidence of knock; B, wire explodes; engine speed, 500 r. p. m.; one spark plug; S-octane fuel.\n\nand show the gas vibrations caused by the exploding wire. Records 566 and 561 show no knock, whereas each of the other cards shows that knock occurred at or near the end of the combustion period.\n\nThe amplitude of the shock waves shown in figure 13 is comparable with the amplitude of the waves set up by moderate knock. There is no indication, however, that waves of this intensity will in themselves cause knock.\n\nFigures 14 and 15 are schlieren photographs showing combustion, with and without knock, started from two levels in the chamber. In the knocking explosion, the two flame fronts had met before knock occurred.\n\nFigure 15 also shows that the two combustion fronts in the knocking explosion met each other before knock occurred, so that no region of noninflamed charge remained in the visible field to auto-ignite. A region of noninflamed charge may have existed in the half of the combustion chamber not covered by the window. The photographs of the nonknocking explosion show the reaction zones meeting at the fifth frame from the left. A small, roughly triangular region still remained on", "timestamp": "2026-07-19T18:47:32.510435+00:00"}
{"citation_id": "19930093281", "source_url": "https://ntrs.nasa.gov/api/citations/19930093281/downloads/19930093281.pdf", "page_number": 10, "total_pages": 22, "image_filename": "19930093281_p10.jpg", "text": "8\n\nfor this useful work. Thus, only 1.9 percent of the engine power is chargeable to the open-nose cowling at 300 miles per hour for a 750-horsepower engine and 0.5 percent for a 1,500-horsepower engine.\n\nThe total percentage of power chargeable to the installation of a radial engine with N.A.C.A. cowling in front of a thick wing or a fuselage is equal to that power chargeable to the nose opening plus the power chargeable to cooling the cylinders. Thus, at 300 miles per hour, this installation cost is 3.9 percent of the engine power for the 750-horsepower engine and 2.5 percent for the 1,500-horsepower engine.\n\nAlthough the preceding result is extremely important as regards radial-engine installations, it is even more important in its general application to airplane design. The greatest drawback to radial-engine installations, namely, the supposedly high aerodynamic drag of the large frontal area, has been eliminated.\n\nThe fact that the power cost of the blunt nose is so markedly affected by the afterbody helps to explain why many test results of cowling installations on airplanes have shown the power cost to be of the order of 25 percent of the engine power. This high power cost means that the nacelles produced some bad flow condition. The test results in this report also explain how some modern airplane-engine installations have given speeds much higher than can be computed from existing cowling-performance data. The installations that gave the high-speed performance were free from bad flow conditions and consequently gave results comparable with those discussed in this report. More exact information on this problem in relation to modern airplanes is an important subject for further research.\n\nCONCLUSIONS\n\n1. The increase in drag of a conventional N.A.C.A. open-nose cowling over that of a streamline nose is greatly affected by the shape of the afterbody. Of the two streamline afterbodies tested, the more streamlined afterbody showed the increment of drag associated with changing", "timestamp": "2026-07-19T18:47:33.644725+00:00"}
{"citation_id": "19930094504", "source_url": "https://ntrs.nasa.gov/api/citations/19930094504/downloads/19930094504.pdf", "page_number": 6, "total_pages": 16, "image_filename": "19930094504_p6.jpg", "text": "4 N.A.C.A. Technical Memorandum No. 912\n\nhigher buckling strength in the plastic region. The grain structure and properties of the 40 x 1 tube may be attained with the steel 1452 by heat treatment.\n\nIn 1932 a series of buckling rods of average slenderness ratio 35 to 60 were improved to a strength of about 110 kg/mm². The buckling-strength curve obtained rises along the Euler curve up to about 60 kg/mm² and in the Tetmajer region still higher as shown in figure 5 (numerical values given in table 6). As a result of more pressing problems, this work was not ended until 1934 with the following two larger series of tests:\n\nFirst there was determined the effect of an annealed weld at the end of the heat-treated buckling rod, since it was supposed that such an effect on the buckling curve would be relatively small. Eighteen rods of the slenderness ratio under consideration were first improved by heat treatment from 110 to 125 kg/mm². Then at both ends 10 millimeters distant from the latter a butt joint was made by welding a saw cut that did not quite go through.\n\nFigure 6 and table IV show that in the entire region lying below the Euler curve, that is, up to the slenderness ratio 55 a minimum buckling stress of 60 kg/mm² may with certainty be attained. It is to be noted that also for this high slenderness ratio buckling failure occurred at the ends while the rod itself remained unchanged over its entire length. Compression pieces of the length of their diameter which pieces were treated and welded in the center gave the same compressive strengths (fig. 6). Those values for which failure could not be attained with the test machines available are indicated with an arrow pointing upward.\n\nConsiderably higher buckling stresses below the Euler curve may be attained if care is taken to see that the above-mentioned buckling failures are avoided, i.e.: if the struts before the treatment are so designed that in the regions which are again annealed by welding to the framework the thickness of the walls is made correspondingly greater than in the remaining region not affected by the heat of the weld*. The reinforcing of the ends is most economically\n\n*This process is legally protected by the firm of Focke-Wulf Flugzeugbau G.m.b.H., Bremen, through DRP. Inventor: Dr.-Ing. Müller, Bremen.", "timestamp": "2026-07-19T18:47:40.137898+00:00"}
{"citation_id": "19930091655", "source_url": "https://ntrs.nasa.gov/api/citations/19930091655/downloads/19930091655.pdf", "page_number": 17, "total_pages": 22, "image_filename": "19930091655_p17.jpg", "text": "ness of the transfer. Calculated and observed pressure changes are plotted against the initial nitrogen (or air) pressure in figure 10 for several temperatures and a gas-fuel ratio of 20.\n\nThe disagreement between the calculated and observed pressure drops is very striking and is much too great to be associated with heat transferred from the bomb wall to the gas phase during the A–C interval, as evidenced by the slow rate of pressure rise after point C. A probable explanation is that a good fraction of the fuel struck the wall, deriving most of its heat therefrom. This assumption is supported by earlier observations that a definite pattern of the sprays could be seen on the bomb wall after certain explosion tests (reference 23) and particularly by the photographs in figure 8. At lower gas densities or with the single-orifice nozzle, the penetration should be greater (reference 1) and the time required to traverse the bomb somewhat shorter. In any case the sprays struck the bomb wall long before minimum pressure was attained.\n\nIn view of the discrepancy between the calculated and observed pressure changes it is rather surprising that the experimental pressure drops are directly proportional to the initial pressure. There is no particular reason for believing that the vapor left the wall in temperature equilibrium with it; i. e., that this vapor could abstract little or no heat from the gas phase, unless perhaps the mass motion of the gas was too slow to effect the removal of the vapor from the immediate neighborhood of the wall in the interval examined.\n\nThe ratio of observed to calculated pressure drop is indicative of the fraction of the total heat contributed by the gas phase. It follows from figure 10 that above $250^\\circ$ C. the fraction of the total heat contributed by the walls became relatively constant at all temperatures for a given density and a gas-fuel ratio of 20, indicating that a constant fraction of the fuel charge struck the wall at temperatures above $250^\\circ$ C., the gas density being almost noninfluential.\n\nThe total pressure drop subsequent to injection increases with initial temperature, fuel quantity, and to some extent with initial density, although in the higher range this latter change is not very evident. There is also a slight decrease in this drop (table I, section 7) with a moderate increase in fuel temperature, showing that in this case less total heat is transferred to the portion of the fuel charge that normally absorbs heat from the gas phase. With carbon dioxide as the ambient gas, the drop is less than that for nitrogen, but a consideration of the relative initial pressures shows that the corresponding temperature drops are of the same magnitude. This similarity might be expected because the spray development is about the same for a given density irrespective of the nature of the gas (reference 28) and, on a weight basis, the specific heat of carbon dioxide is not greatly different from that of nitrogen in this temperature range. For a given fuel weight benzene gives a greater drop than does the\n\nDiesel fuel, presumably owing to the greater heat required for vaporization. This presumption assumes that the same fraction of the fuel (benzene or Diesel fuel) fails to strike the wall under identical circumstances. The benzene tests also indicate that the surface temperature of the drops is well below the ambient-gas temperature; although the gas temperatures employed were near to or above the critical temperature of benzene, the fact that the A–C interval was about the same for benzene as for Diesel fuel indicates a droplet temperature much below the critical point.\n\nTime to attain minimum pressure.—Small variations of the A–C interval are evident but, because of possible errors, these variations may not be real. In any case the variations cannot be associated with any primary variable. The interval is greatest for carbon dioxide, intermediate for air, and least for nitrogen; it increases with the fuel quantity for the lower but not for the\n\n[Figure: Comparison of calculated and observed pressure drops at various temperatures.]\n\nhigher weights; and there appears to be a slight increase with gas density. As the total pressure drop increases with an increase in the fuel weight and to some extent with an increase in gas density, it is conceivable that the latter trends arise from an “overshooting” of the true decrease in pressure because of the increased amplitude of the wave evident after point C. The records for benzene, however, fail to show such trends.\n\nIn view of the wide variation of the fraction of the fuel that strikes the bomb wall with varying fuel weights and given gas density, the minimum point cannot be logically associated with the moment of complete evaporation of the fuel on the wall. This contention is further substantiated by the failure of benzene to give a shorter interval; its greater volatility should enable it to evaporate more rapidly from the bomb surface. It has previously been shown that non-uniformity of the gas-vapor mixture exists for at least 0.06 second after injection (reference 23). An attempt was made to mix the charge with a 4-blade fan driven at 7,000 r. p. m. but, as the A–C interval corresponded to only two revolutions of the fan, it is not surprising", "timestamp": "2026-07-19T18:47:44.236447+00:00"}
{"citation_id": "19930091731", "source_url": "https://ntrs.nasa.gov/api/citations/19930091731/downloads/19930091731.pdf", "page_number": 6, "total_pages": 32, "image_filename": "19930091731_p6.jpg", "text": "2\nREPORT NO. 656—NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nThe lengths of all the compressive and the column specimens were measured. In order to determine the required cross-sectional properties of the compressive and the column specimens, more than half of them were weighed and, for each of these specimens, measurements were made of the thickness and the width of each flange and the depth of the section at the middle and of the thickness of the web at each end. The density of a sample of the material was determined by the Division of Weights and Measures of the National Bureau of Standards. The cross-sectional areas were computed\n\nThe most suitable machine available for making the compressive tests was considered to be a fluid-support, Bourdon-tube, hydraulic machine. Auxiliary nuts on the screws of this machine were tightened against the lower surface of the adjustable head to bring it into contact with the lower surface of the threads on the screws, so that rotation of the head relative to the platen of the machine due to clearance between the nuts of the head and the screws was obviated. The unsymmetrical position of the motor, the handwheel, and the other mechanism for raising and lowering the adjustable\n\n<!-- Image (131, 274, 902, 630) -->\n\nFIGURE 2.—Typical tensile stress-strain diagrams for 24S-T of cross section A. Strains measured on 2-inch gage length with Ewing extensometer. National Bureau of Standards.\n\nfrom the weights, the lengths, and the densities; and the least radii of gyration were computed from the measured cross-sectional dimensions and the nominal radii of the roundings and fillets.\n\nTENSILE AND COMPRESSIVE TESTS\n\nTensile tests were made in screw-driven, beam-and-poise testing machines. The specimens were held in Templin grips supported by spherical bearings. Strains were measured in 2-inch gage lengths by means of Ewing extensometers. Three typical tensile stress-strain diagrams are shown in figure 2. Specimens 5CTC and 5CTA were taken from the flanges and specimen 5CTE from the web. The tensile yield strength was obtained from the stress-strain diagram as the stress at a strain 0.002 in excess of the elastic strain corresponding to this stress.\n\nhead causes it to exert on the portion of the two screws below it a constant moment of roughly 1,000 pound-inches in a plane normal to that of the screws. Consequently the screws are slightly bent elastically and, as they tend to straighten under load, produce rotation of the head. This condition causes a slight eccentricity of loading, which is especially undesirable in compression testing; but, with the short specimens and comparatively low loads (maximum, one-third the capacity of the machine) of the present investigation, the effect was not considered serious. Another possible source of error in making compressive tests in this type of machine arises from the possibility of rotation of the platen about a horizontal axis. The platen is rigidly connected to the piston of the hydraulic jack, which is packed, and the clearance between the cylinder and the piston permits rotation of the platen under eccentric", "timestamp": "2026-07-19T18:48:02.309199+00:00"}
{"citation_id": "19930091748", "source_url": "https://ntrs.nasa.gov/api/citations/19930091748/downloads/19930091748.pdf", "page_number": 9, "total_pages": 12, "image_filename": "19930091748_p9.jpg", "text": "14.8 and 53.5, respectively. The value of $r/2\\pi$ for the spring P is 280 cycles per second. Since all the tests with the balance were made at oscillation frequencies lower than 17 cycles per second, the maximum phase error was 1.8°, which is of the same magnitude as the experimental errors.\n\nIt was necessary to balance dynamically the oscillating mass about the axis of rotation. For this purpose, a weight I was placed ahead of the axis of rotation to provide static and dynamic balance within the accuracy of the balance measurements.\n\nThe wing was oscillated in approximately sinusoidal motion about its axis by the direct-current motor D, which was controllable so that any desired frequency could be obtained. The wing was oscillated at the end opposite to the one at which the forces were measured. A time history of the angular position of the airfoil was recorded on the same film that was used to record the forces. This record was obtained by means of a rotating contact E attached to the driving motor. Once each revolution, this contact energized an electromagnet M, which deflected a mirror and moved a light beam on the film G. The lag of the timing system was measured by an oscillograph and found to be 0.0014 second.\n\nA typical record of the measurements taken photographically with the oscillating-airfoil balance is shown in figure 3. Line 1 (fig. 3) is a stationary reference line; line 2 is the record of the timing element; and line 3 is the trace of the deflection of spring P and therefore is a record of the lift on the airfoil. The oscillations occurring in the timing line 2 are of no interest.\n\nIf, for example, oscillation about the angle of zero lift is assumed, the intersection of the lift line 3 with the zero line 4 determines the point of zero lift. By static calibration, the angular position of the airfoil corresponding to this point on the record is readily determined. A small correction is applied to take into account the lag in the timing unit. The phase angle is then directly obtained as the difference between the angular position of the curves for zero lift and zero angle.\n\n### RESULTS AND DISCUSSION\n\nThe measured phase difference $\\delta$ for various values of the parameter $1/k$ are shown in figure 4. It will be noted in some cases that the same value of $1/k$ was obtained with several values of $p$ and $v$.\n\nBefore the theoretical values of $\\delta$ could be computed, it was necessary to determine the effective aspect ratio of the airfoil for the test conditions. This value was derived from the value of the lift-curve slope determined by means of static lift observations at numerous angles of attack (fig. 5). The effective aspect ratio is 12.5, which may be compared with the geometric value of 7.07. The leakage around the end plates of the airfoil prevented the realization of a higher value of the effective aspect ratio. For comparison, the lift curve is given (fig. 5) for the case in which the angle of attack of the airfoil is suddenly increased. The slope of the curve is about the same; the stalling of the moving airfoil at the higher angles of attack, however, does not occur.\n\nThe values of $F$ and $G$ corresponding to the effective aspect ratio were obtained from reference 6 and are shown in figure 6 against $1/k$. Values for an infinite-span airfoil from reference 4 are also shown. The theoretical values of $\\delta$ computed by equation (6) with values of $F$ and $G$ from figure 6 are shown with the experimental results in figure 4. Consideration was given to changing the inertia terms in equation (6) to take into account the decreased virtual volume due to the finite span of the airfoil. Calculations indicate this effect to be inappreciable.\n\nThe experimental and the theoretical values for the finite-span airfoil show phase differences of no more than about 5°. The corrections applied to the infinite-aspect-ratio theory to take the finite span into account are in the direction of improving the agreement between the theory and experiments.\n\nIt is noted that, at low frequencies of oscillation (large values of $1/k$, fig. 4), a lead that is consistently larger than expected appears. The study of the cause of this discrepancy will be left for a future investigation.\n\nLANGLEY MEMORIAL AERONAUTICAL LABORATORY,\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS,\nLANGLEY FIELD, VA., April 24, 1939.\n\n### REFERENCES\n\n1. Wagner, Herbert: Über die Entstehung des Dynamischen Auftriebes von Tragflügeln. Z. f. a. M. M., Bd. 5, Heft 1, Feb. 1925, S. 17-35.\n2. Küssner, Hans Georg: Schwingungen von Flugzeugflügeln. Jahrb. 1929, DVL S. 313-334.\n3. Glauert, H.: The Force and Moment on an Oscillating Aerofoil. R. & M. No. 1242, British A. R. C., 1929.\n4. Theodorsen, Theodore: General Theory of Aerodynamic Instability and the Mechanism of Flutter. T. R. No. 496, N. A. C. A., 1935.\n5. Garrick, I. E.: Propulsion of a Flapping and Oscillating Airfoil. T. R. No. 567, N. A. C. A., 1936.\n6. Jones, Robert T.: The Unsteady Lift of a Finite Wing. T. N. No. 682, N. A. C. A., 1939.\n7. Theodorsen, Theodore, and Silverstein, Abe: Experimental Verification of the Theory of Wind-Tunnel Boundary Interference. T. R. No. 478, N. A. C. A., 1934.\n8. Timoshenko, S.: Vibration Problems in Engineering. D. Van Nostrand Co., Inc., 1928, pp. 24-27.\n\nU. S. GOVERNMENT PRINTING OFFICE: 1939", "timestamp": "2026-07-19T18:48:02.545061+00:00"}
{"citation_id": "19930094544", "source_url": "https://ntrs.nasa.gov/api/citations/19930094544/downloads/19930094544.pdf", "page_number": 39, "total_pages": 43, "image_filename": "19930094544_p39.jpg", "text": "N.A.C.A. Technical Memorandum No. 872\n\nFigs. 32,33,34\n\n[Figure: Figure 32.- R-101 - Ring lying on the floor. The columns in the outer ring plane consist of longitudinal girder sections.]\n\n[Figure: Figure 33.- R-101- Stabilizing surface structure. The two rings in way of the surfaces are of cruciform type, extensions of which form the spars for the surfaces.]\n\n[Figure: Figure 34.- R-101. Inside view. In the foreground at the left the corridor made up of weak framing, and at the right a portion of the three boom ring, are visible.]", "timestamp": "2026-07-19T18:48:05.091347+00:00"}
{"citation_id": "19930094542", "source_url": "https://ntrs.nasa.gov/api/citations/19930094542/downloads/19930094542.pdf", "page_number": 92, "total_pages": 102, "image_filename": "19930094542_p92.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-19T18:48:05.294770+00:00"}
{"citation_id": "19930094549", "source_url": "https://ntrs.nasa.gov/api/citations/19930094549/downloads/19930094549.pdf", "page_number": 31, "total_pages": 76, "image_filename": "19930094549_p31.jpg", "text": "N.A.C.A. Technical Memorandum No. 867 29\n\nvelocities which are of the order of 0.05 radian per second. We were able, however, to calculate the angular velocities q from the angles θ. The angular velocities thus calculated are, however, obtained after a rather larger number of operations, among which is the correction for the indications of the variometer. Under these conditions, the agreement of the two curves is not to be considered as very poor.\n\nWe shall now consider the curves in detail. After the elevator has been deflected from β = -3.2° to β = -1.5° the airplane tends toward a final state B. The latter, corresponding to a displacement Δβ = 1.7°, is characterized by a final angle of attack:\n\n$$\ni_f = 3.6^\\circ + \\Delta i\n$$\n\nwhere\n\n$$\n\\Delta i = -\\frac{0.0138}{0.00477} \\Delta \\beta = -5^\\circ\n$$\n\nTherefore,\n\n$$\ni_f = 3.6^\\circ - 5^\\circ = 1.4^\\circ\n$$\n\ncorresponding to a lift coefficient:\n\n$$\nC_z = 0.458 - 0.0695 \\Delta i = 0.111\n$$\n\nand a velocity\n\n$$\nV = 92 \\text{ m/s}\n$$\n\nThe airplane used is thus extremely sensitive to the elevator controls. The deflection was applied during 7 seconds, within which time the airplane began a series of motions which would have led to state B if the deflection had been applied long enough. This period is characterized by the immediate appearance of a component of the velocity directed downward by the decrease in the angles of attack and acceleration $J_z$, and by the increase in the velocity along the flight path. We may note that the increase in the velocity is made at first rather slowly, thus confirming the computations given in figure 14.\n\nLet us now examine the effect of the return to the initial elevator deflection. At the instant when the stick is thrown back to its original position (point 2), the curve of accelerations is suddenly modified, this being the curve which indicates most exactly the instant when the maneuver has been executed. The angle of attack is instant-", "timestamp": "2026-07-19T18:48:10.506913+00:00"}
{"citation_id": "19930091697", "source_url": "https://ntrs.nasa.gov/api/citations/19930091697/downloads/19930091697.pdf", "page_number": 18, "total_pages": 28, "image_filename": "19930091697_p18.jpg", "text": "14\nREPORT NO. 622—NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\n5282-93\n30° B.T.C. 20° 10° T.C. A.T.C. 10°\nKnocking explosion, 65-octane fuel\n\n5282-101\n30° B.T.C. 20° 10° T.C. A.T.C. 10°\nNonknocking explosion, 65-octane fuel plus ethyl fluid\n\nFIGURE 14.—Schlieren photographs showing two flame fronts passing through each other. Air-fuel ratio, 14; engine speed, 900 r. p. m.; two spark plugs.", "timestamp": "2026-07-19T18:48:10.935486+00:00"}
{"citation_id": "19930093281", "source_url": "https://ntrs.nasa.gov/api/citations/19930093281/downloads/19930093281.pdf", "page_number": 11, "total_pages": 22, "image_filename": "19930093281_p11.jpg", "text": "L-379\n\nthe nose to be about one-fourth of that with the other afterbody.\n\n2. The results show that the drag measurements obtained without the use of the propeller on a neutral afterbody need not be corrected in applying them to the condition of the propeller operating.\n\n3. The results from this investigation indicate that the power cost, in excess of that with a streamline nose, of using an N.A.C.A. cowling in front of a well-designed afterbody to enclose a 1,500-horsepower engine on an airplane with a speed of 300 miles per hour amounts to 1.5 percent of the engine power. To this value must be added 1 percent for the internal work of cooling the rear of the engine cylinders, giving a total installation power cost of 2.5 percent. If the open-nose cowling is credited with 1 percent because it cools the front of the cylinders, the nonuseful power cost of the N.A.C.A. installation amounts to only 0.5 percent of the engine power.\n\nLangley Memorial Aeronautical Laboratory,\nNational Advisory Committee for Aeronautics,\nLangley Field, Va., April 28, 1939.", "timestamp": "2026-07-19T18:48:14.343677+00:00"}
{"citation_id": "19930094516", "source_url": "https://ntrs.nasa.gov/api/citations/19930094516/downloads/19930094516.pdf", "page_number": 2, "total_pages": 32, "image_filename": "19930094516_p2.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-19T18:48:15.751863+00:00"}
{"citation_id": "19930094493", "source_url": "https://ntrs.nasa.gov/api/citations/19930094493/downloads/19930094493.pdf", "page_number": 8, "total_pages": 31, "image_filename": "19930094493_p8.jpg", "text": "N.A.C.A. Technical Memorandum No. 923 7\n\nThe resistance itself is determined by comparison with a precision resistance of the firm, Hartmann & Braun.\n\nWhen used as a hot wire for measuring the velocity, it is connected to circuit II (fig. 1). Here the wire is put in series with an oscillograph loop and a precision ammeter and connected to a battery of 4 volts. In order to keep the voltage of the battery of circuit II as constant as possible, the storage battery is provided with a resistance during the entire test period. Here again the resistance is determined by comparison with a precision resistance and the current with the aid of the precision ammeter.\n\nThe oscillogram is obtained in the following manner. First, the resistance box is connected to the circuit I. By inserting the proper values of the resistance, lines of constant resistance, and therefore lines of constant temperature are obtained on the recording paper. The resistance box is then disconnected and the wire element switched in. The resistance of the wire varies with the temperature of the surrounding air in the cylinder according to the relation\n\n$$\nr = \\varphi(T_w)\n$$\n\n(3)\n\nAdditional heat is received by the wire by the measuring current. Also, investigation was made to determine how high the measuring current may be before the change in heat resistance is practically zero. This was found to be the case for a current of 10 mA. The current used was half this value; that is, a maximum of 5 mA.\n\nIn obtaining the hot-wire curves, the wire must be strongly loaded, since it must receive a higher temperature than that of the surrounding air in the cylinder. For this purpose, it was connected to circuit II. Through the changes in the air velocity in the cylinder, the rate of cooling of the wire varies. This results in a change in the wire temperature or its resistance and also the power absorbed. The calibration is the same as with circuit I. There are again obtained calibration lines of constant resistance and also of constant current, the current through the resistance being measured with the precision ammeter. The voltage can therefore also be obtained, being given by Ohm's law:\n\n$$\nE = i r \\text{ (V)}\n$$\n\n(15)", "timestamp": "2026-07-19T18:48:15.937046+00:00"}
{"citation_id": "19930093640", "source_url": "https://ntrs.nasa.gov/api/citations/19930093640/downloads/19930093640.pdf", "page_number": 8, "total_pages": 28, "image_filename": "19930093640_p8.jpg", "text": "6\n\n172 square feet and are corrected for wind-tunnel effects. Pitching-moment coefficients are computed about the assumed center-of-gravity position shown in figure 5. A comparison of the more important characteristics such as $L/D_{\\text{max}}$, $C_{L_{\\text{max}}}$, $C_D$ at $C_L = 0.25$, etc., is given in table I.\n\n**Drag.**— The scale effect on the drag coefficients of the various model arrangements at $C_L = 0.25$ (assumed high-speed lift coefficient) is shown in figures 11 and 12. The drag coefficients obtained at 100 miles per hour are used for the comparison of the arrangements in table I. The drag increments due to the nacelles, radiators, cowlings, etc., are shown in figure 13.\n\nBased on the bare-wing model drag, the tests show that the liquid-cooled engine nacelles increase the drag coefficient of the model by 0.0014, or 7.9 percent; the oil coolers increase the drag by 0.0007, or 3.9 percent; and the Prestone radiators increase the drag by 0.0024, or 13.5 percent. The total increase in drag coefficient due to the liquid-cooled engine installation is 0.0045, or 25.3 percent.\n\nThe increase in drag coefficient due to the air-cooled engine nacelles and cowlings with no cooling air is 0.0030 or 16.8 percent of the bare-wing model drag. With the cooling air flowing through the large exit slot of the cowlings the drag coefficient of the nacelles is increased to 0.0060 or 33.7 percent. Including the 3.9-percent increase due to the oil coolers, the total drag of the air-cooled engine installations with large exit slots is 0.0067 or 37.6 percent of the bare-wing model drag. By reducing the exit slot gap to 3/4 inch, eliminating the sharp corner of the nacelle at the cowling exit slot, and providing a smooth contour, the drag of the air-cooled installation was reduced to 0.0054 or 30.4 percent of the bare-wing model drag.\n\n**Maximum lift.**— Values of maximum lift for the various arrangements are shown in table I. There is little variation in the maximum lift coefficients for the air-cooled engine arrangements; however, they show a small increase over the values obtained for the bare-wing case. This increase may possibly be attributed to an increase in the effective area of the wing due to the nacelles.\n\nOf particular interest is the comparatively low value of the maximum lift coefficient for the liquid-cooled en-", "timestamp": "2026-07-19T18:48:22.743846+00:00"}
{"citation_id": "19930093641", "source_url": "https://ntrs.nasa.gov/api/citations/19930093641/downloads/19930093641.pdf", "page_number": 33, "total_pages": 47, "image_filename": "19930093641_p33.jpg", "text": "```markdown\nL-458\n\nN.A.C.A.\n\nOriginal\nO--x Repeat\n\nPitching-moment\ncoefficient, $C_m$\n\nRatio of lift to drag, L/D\n\nLift coefficient, $C_L$\n\nDrag coefficient, $C_D$\n\nAngle of attack of reference axis, $\\alpha_r$, deg.\n\nFigure 9\n\nPitching-moment\ncoefficient, $C_m$\n\nRatio of lift to drag, L/D\n\nLift coefficient, $C_L$\n\nDrag coefficient, $C_D$\n\nAngle of attack of reference axis, $\\alpha_r$, deg.\n\nLanding gear $\\delta$, deg.\nRetracted 0\nExtended 60.8\n\nFigure 10\n\nFigs. 9, 10\n```", "timestamp": "2026-07-19T18:48:23.653140+00:00"}
{"citation_id": "19930091701", "source_url": "https://ntrs.nasa.gov/api/citations/19930091701/downloads/19930091701.pdf", "page_number": 16, "total_pages": 18, "image_filename": "19930091701_p16.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-19T18:48:37.148750+00:00"}
{"citation_id": "19930091724", "source_url": "https://ntrs.nasa.gov/api/citations/19930091724/downloads/19930091724.pdf", "page_number": 10, "total_pages": 20, "image_filename": "19930091724_p10.jpg", "text": "6\nREPORT NO. 649 NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nThe cap-block was a disk of hardened steel $1\\frac{13}{16}$ inches in diameter and $\\frac{3}{4}$ inch thick. Its lower surface, which was smooth-ground, made direct contact with the upper end of the pack. The plaster of paris shim did not exceed 0.1 inch in thickness. It was cast, under an initial load of about one kip, between bearing block U and the upper surface of the cap-block V. The plaster was allowed to harden about 10 minutes before the pins were placed in position.\n\nASSEMBLING PINS\n\nThe pins providing transverse support were located by a definite procedure. They were first positioned, using the tweezers Z to guide the pins to their proper location and the small wrenches W to turn the screws. The perforated strips of tracing cloth T were used to support the pins in approximately the right position. The screws in the same locations on opposite sides of the pack were then tightened simultaneously and progressively until the points of the pins were slightly embedded in the surface of the pack.\n\nIt was considered important to locate the pins in a definite order. Satisfactory results were obtained by using the following sequence. Numbering the pins in rows from 1 to 10 beginning at the top, the pins in rows 5 and 6 of the middle column were located first. Next, those in rows 3, 4, 7, and 8, of the middle column, and then those in rows 3 to 8, inclusive, of the outside columns were located. The clamp at the upper end was then removed and the end of the strip of tracing cloth T placed under the clamp X. The pins in rows 1 and 2 were then located. The clamp at the lower end of the pack was removed and pins in rows 9 and 10 located. All screws were then systematically tried with the wrenches to insure that the ends of all the pins were bearing against the pack.\n\nSTRAIN GAGES\n\nThe strain was measured by a pair of Tuckerman 1-inch optical strain gages (reference 5). These gages were attached on each side of the pack to the edge of the middle specimen.\n\nCROSS-SECTIONAL AREA\n\nThe cross-sectional area of a pack was computed by dividing the weight of the pack by its length and the density of the material.\n\nLIMITATIONS\n\nThe limitations of this method of test have not been thoroughly explored. When preliminary results were obtained which apparently furnished satisfactory information for some of the materials generally used in aircraft, tests on a greater number of materials were desired. This has limited the time available for a thorough investigation into the capacity and accuracy of the method under various conditions.\n\nExperience from tests, however, has shown that packs taken from aluminum alloy sheet composed of 13, 7, and 5 specimens of 0.032, 0.064, and 0.081 inch material, respectively, sustained compressive stresses in excess of 60 kips/in.$^2$ before the packs failed through major instability. Packs composed of five specimens taken from heat-treated chromium-molybdenum steel sheet, 0.05 inch, were subjected to compressive stresses up to 180 kips/in.$^2$ without failure through major instability. Within these limitations the pack test appears to give the compressive properties of a material within the same order of accuracy as is usually obtained in other mechanical tests, such as the tensile test.\n\nTESTS ON BARS\n\nPURPOSE\n\nThe \"pack\" test is based on the assumption that it will give compressive results like those obtained from block compressive tests. A number of comparative tests on packs and on compact solid specimens taken from metal bars were made to see whether or not this assumption was justified.\n\nMATERIALS\n\nThe following materials were used in making these tests:\n\na. Carbon steel bar.\n   Condition, cold rolled.\n   Shape, round.\n   Size, one-inch diameter.\n\nb. Brass bar.\n   Condition, rolled.\n   Shape, square.\n   Size, one-inch on side.\n\nc. Aluminum alloy.\n   Condition, rolled.\n   Shape, round.\n   Size, one-inch diameter.\n\nSPECIMENS AND PACKS\n\nCompact solid specimens and packs were obtained from alternate locations along each bar. The compact solid specimens were cut with symmetry to the axis of the bar to a size of $\\frac{23}{32}$ by $\\frac{23}{32}$ by $2\\frac{7}{16}$ inch.\n\nThe \"pack\" specimens were obtained from the same location in the cross section of the bar as the compact solid specimens. The pack was composed of five specimens, 0.1 inch thick. These specimens were prepared by machining with light cuts so that the underlying material was deformed as little as possible. The finished surfaces were smooth and the burrs were removed from the edges.\n\nPROCEDURE\n\nThe packs were tested using the procedure for \"pack\" tests as previously outlined in the section on Test Procedure (p. 3).", "timestamp": "2026-07-19T18:48:52.043089+00:00"}
{"citation_id": "19930093641", "source_url": "https://ntrs.nasa.gov/api/citations/19930093641/downloads/19930093641.pdf", "page_number": 34, "total_pages": 47, "image_filename": "19930093641_p34.jpg", "text": "```markdown\nL-452\n\nN.A.C.A.\n\nFigs. 11, 12\n\n<!-- Image (45, 120, 900, 860) -->\n\nFigure 11\n\nFigure 12\n```", "timestamp": "2026-07-19T18:49:03.333160+00:00"}
{"citation_id": "19930093640", "source_url": "https://ntrs.nasa.gov/api/citations/19930093640/downloads/19930093640.pdf", "page_number": 9, "total_pages": 28, "image_filename": "19930093640_p9.jpg", "text": "7\n\ngine arrangement with oil coolers open. Unfortunately, the maximum lift coefficient was not determined for the air-cooled engine arrangement with oil coolers open; however, a study of tuft surveys made on the liquid-cooled engine arrangement (reference 1) indicates that the oil coolers seriously disturb the air flow over the wing at large angles of attack, thereby inducing an earlier separation and lower maximum lift coefficient.\n\nPROPULSIVE AND OVER-ALL EFFICIENCIES\n\nEngine-propeller combinations should be compared by means of an over-all efficiency including both drag and propulsive efficiency. The over-all efficiency is defined as the ratio of the power required for the bare-wing model at a given level flight speed to the power input actually required at this speed for the model with the engine-propeller installation.\n\nThe over-all efficiency of the bare-wing model is therefore 100 percent and, for an engine-propeller combination, is given by\n\n$$\n\\eta_t = \\eta \\left( \\frac{C_{D_w}}{C_{D_c}} \\right)\n$$\n\nValues of over-all efficiency given in table I are based on a lift coefficient, $ C_L = 0.25 $, and a blade angle, $ \\beta = 23\\frac{1}{2}^\\circ $ at 0.75 R, which are assumed high-speed conditions.\n\nThe effective thrust of a propeller-body combination may be computed from wind-tunnel data by means of the relation\n\n$$\nR = D_c + \\Delta D - T\n$$\n\nfrom which,\n\n$$\nT - \\Delta D = D_c - R\n$$\n\nFor tests without a lifting surface behind the propeller, $ T - \\Delta D $ may be obtained from measurements of $ D_c $ and R made at the same angle of attack and dynamic pressure.", "timestamp": "2026-07-19T18:49:05.884126+00:00"}
{"citation_id": "19930091655", "source_url": "https://ntrs.nasa.gov/api/citations/19930091655/downloads/19930091655.pdf", "page_number": 18, "total_pages": 22, "image_filename": "19930091655_p18.jpg", "text": "that the interval was unaltered. Since the rates of heat transfer from the wall and to the fuel are equal at C, it would seem that the interval should depend upon the nonuniformity of the mixture, which in turn should be dependent upon the injected fuel weight and the gas density. Actually, the interval is practically independent of both variables.\n\nCONCLUSIONS\n\n1. The injection of liquid fuel into a heated and compressed gas has furnished data on the initial rate of heat exchange between the ambient gas and the fuel. The actual rates of vaporization were indeterminate, but it is shown that vaporization began immediately after injection started. The same situation must also be true for engines.\n\n2. For given experimental conditions, the initial rate of heat transfer was essentially constant during the time required for the spray to traverse the bomb. This initial rate was found to be proportional to the initial temperature difference between the fuel and the gas. The total heat transferred in engines must be greater owing to the greater initial temperature difference.\n\n3. The initial heat-transfer period was approximately constant (0.0020 ± 0.0005 second) for the 13-orifice, 2-impinging-jets, and single-orifice nozzles tested and also for benzene and Diesel fuel, which have quite different volatilities and viscosities.\n\n4. At the temperatures investigated the transfer of heat by radiation was negligible as compared with that transferred by conduction. This situation must also exist in an engine until the start of flame combustion.\n\n5. The efficacy with which heat transfer took place decreased considerably with increasing fuel quantity at all densities and temperatures investigated.\n\n6. Under all conditions a good fraction of the total heat absorbed after the spray had traversed the bomb must have occurred at the bomb wall.\n\nLANGLEY MEMORIAL AERONAUTICAL LABORATORY,  \nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS,  \nLANGLEY FIELD, VA., August 25, 1939.\n\nREFERENCES\n\n1. Lee, Dana W.: A Comparison of Fuel Sprays from Several Types of Injection Nozzles. T. R. No. 520, N. A. C. A., 1935.\n\n2. Lee, Dana W.: Measurements of Fuel Distribution within Sprays for Fuel-Injection Engines. T. R. No. 565, N. A. C. A., 1936.\n\n3. Rothrock, A. M., and Waldron, C. D.: Effect of Nozzle Design on Fuel Spray and Flame Formation in a High-Speed Compression-Ignition Engine. T. R. No. 561, N. A. C. A., 1936.\n\n4. Rothrock, A. M., and Waldron, C. D.: Fuel Spray and Flame Formation in a Compression-Ignition Engine Employing Air Flow. T. R. No. 588, N. A. C. A., 1937.\n\n5. Boerlage, G. D., and van Dyck, W. J. D.: Causes of Detonation in Petrol and Diesel Engines. R. A. S. Jour., Dec. 1934, pp. 953–986.\n\n6. Rothrock, A. M., and Waldron, C. D.: Some Effects of Injection Advance Angle, Engine-Jacket Temperature, and Speed on Combustion in a Compression-Ignition Engine. T. R. No. 525, N. A. C. A., 1935.\n\n7. Gerrish, Harold C., and Ayer, Bruce E.: Influence of Fuel-Oil Temperature on the Combustion in a Prechamber Compression-Ignition Engine. T. N. No. 565, N. A. C. A., 1936.\n\n8. Rothrock, A. M., and Waldron, C. D.: Fuel Vaporization and Its Effect on Combustion in a High-Speed Compression-Ignition Engine. T. R. No. 435, N. A. C. A., 1932.\n\n9. Wentzel, W.: Ignition Process in Diesel Engines. T. M. No. 797, N. A. C. A., 1936.\n\n10. Ellenwood, F. O., Evans, F. C., and Chwang, C. T.: Efficiencies of Otto and Diesel Engines. A. S. M. E. Trans. OGP-50-6, Jan.–April 1928, pp. 1–22.\n\n11. Rothrock, A. M., and Cohn, Mildred: Some Factors Affecting Combustion in an Internal-Combustion Engine. T. R. No. 512, N. A. C. A., 1934.\n\n12. Rothrock, A. M., and Spencer, R. C.: Effect of Moderate Air Flow on the Distribution of Fuel Sprays after Injection Cut-Off. T. R. No. 483, N. A. C. A., 1934.\n\n13. Nusselt, Wilhelm: Wärmeübergang, Diffusion und Verdunstung. Z. f. a. M. M., vol. 10, 1930, pp. 105–121.\n\n14. Fuchs, N.: Über die Verdampfungsgeschwindigkeit kleiner Tröpfchen in einer Gasatmosphäre. Phys. Zeit. Sowjet-un., vol. 6.3, 1934, pp. 224–243.\n\n15. Sherwood, T. K., and Gilliland, E. R.: Diffusion of Vapors through Gas Films. Indus. Eng. Chem., vol. 26, 1934, pp. 1093–1096.\n\n16. Ingersoll, L. R., and Zobel, O. J.: An Introduction to the Mathematical Theory of Heat Conduction. Ginn and Co. 1913, p. 133.\n\n17. Cragoe, C. S.: Thermal Properties of Petroleum Products. Misc. Publication No. 97, Bur. Standards, 1929.\n\n18. National Bureau of Standards: National Standard Petroleum Oil Tables. Circular No. 154, Bur. Standards, 1924, pp. 95–113.\n\n19. Gauchier, L. P.: Specific Heat of Liquid Pure Hydrocarbons and Petroleum Fractions. Indus. Eng. Chem., vol. 27, 1935, pp. 57–64.\n\n20. Lee, Dana W.: The Effect of Nozzle Design and Operating Conditions on the Atomization and Distribution of Fuel Sprays. T. R. No. 425, N. A. C. A., 1932.\n\n21. Walker, William H., Lewis, Warren K., and McAdams, William H.: Principles of Chemical Engineering. McGraw-Hill Book Co., Inc., 1927, p. 162.\n\n22. Watson, K. M., and Nelson, E. F.: Improved Methods for Approximating Critical and Thermal Properties of Petroleum Fractions. Indus. Eng. Chem., vol. 25, 1933, pp. 880–887.\n\n23. Cohn, Mildred, and Spencer, Robert C.: Combustion in a Bomb with a Fuel-Injection System. T. R. No. 544, N. A. C. A., 1935.\n\n24. Lewis, Gilbert Newton, and Randall, Merle: Thermodynamics and the Free Energy of Chemical Substances. McGraw-Hill Book Co., Inc., 1923, p. 68.\n\n25. National Research Council: International Critical Tables, vol. V. McGraw-Hill Book Co., Inc., 1929, p. 146.\n\n26. Schweitzer, P. H.: The Penetration of Oil Sprays in Dense Air. Tech. Bull. No. 20, Penn. State Coll., 1934, pp. 108–124.\n\n27. McAdams, William H.: Heat Transmission. McGraw-Hill Book Co., Inc., 1933, pp. 20, 96, 216, and 246.\n\n28. Joachim, W. F., and Beardsley, Edward G.: The Effects of Fuel and Cylinder Gas Densities on the Characteristics of Fuel Sprays for Oil Engines. T. R. No. 281, N. A. C. A., 1927.\n\n29. Lee, Dana W., and Spencer, Robert C.: Photomicrographic Studies of Fuel Sprays. T. R. No. 454, N. A. C. A., 1933.\n\n30. Gelalles, A. G.: Some Effects of Air and Fuel Oil Temperatures on Spray Penetration and Dispersion. T. N. No. 338, N. A. C. A., 1930.", "timestamp": "2026-07-19T18:49:06.932996+00:00"}
{"citation_id": "19930091697", "source_url": "https://ntrs.nasa.gov/api/citations/19930091697/downloads/19930091697.pdf", "page_number": 19, "total_pages": 28, "image_filename": "19930091697_p19.jpg", "text": "A PHOTOGRAPHIC STUDY OF COMBUSTION AND KNOCK IN A SPARK-IGNITION ENGINE\n\nB.T.C. 15° 10° 5° T.C. 5° 10° A.T.C.\nKnocking explosion, 65-octane fuel\n\nB.T.C. 15° 10° 5° T.C. 5° 10° A.T.C.\nNonknocking explosion, 100-octane fuel\n\nFIGURE 15.—Spark schlieren photographs of knocking and nonknocking explosions. Air-fuel ratio, 14; A, first evidence of knock; engine speed, 500 r. p. m.; two spark plugs.\n\n15", "timestamp": "2026-07-19T18:49:07.959993+00:00"}
{"citation_id": "19930094506", "source_url": "https://ntrs.nasa.gov/api/citations/19930094506/downloads/19930094506.pdf", "page_number": 5, "total_pages": 24, "image_filename": "19930094506_p5.jpg", "text": "N.A.C.A. Technical Memorandum No. 910 3\n\nwere to explain the question deciding the usefulness of the method whether or not the effect of the different influential quantities is correctly reproduced in model testing.\n\nIn the new spinning balance the model is suspended from the conventional six-component balance in the tunnel jet which is given a spiral motion by a rotating screen. This type of experiment has, from the recording point of view, fundamental advantages over the operation on the spinning balance; but, as pointed out at the same time by Kramer and Krüger, it also has one fundamental defect: the static pressure of the free flow is not constant. As a result of the centrifugal force applied at the jet the static pressure decreases a little according to a parabolic law from the circumference to the jet center. Measurements disclosed a very close accord between theory and experiment.\n\nThe moments induced by the variable static pressure themselves may, at higher angles of attack where the tail is perceptibly away from the jet center, be ignored in the face of the elsewhere existing instrumental inaccuracies. Another question is whether, as a result of the pressure gradient, a movement of the boundary layer might occur which could effect a substantial change of profile characteristics. This change would be in the opposite sense from the flight test. Since, on the other hand, the speed of jet rotation in the tests was fairly low and furthermore, the drop in static pressure at jet center remained small, no appreciable effect on the profile characteristics through boundary-layer movement was anticipated.\n\nTEST PROCEDURE\n\nIn view of the original intention to include measurements with the introduction of a spinning radius, the model was made comparatively small. The lessened instrumental accuracy resulting therefrom was, to a certain degree, ameliorated through the use of sufficiently sensitive metering diaphragms. This was most difficult to achieve in the drag component measurements where, because of the smallness of the righting forces, it was difficult to get an exact reading of the zero reference values. Another drawback resulting from the smallness of the model was that the", "timestamp": "2026-07-19T18:49:15.423145+00:00"}
{"citation_id": "19930094516", "source_url": "https://ntrs.nasa.gov/api/citations/19930094516/downloads/19930094516.pdf", "page_number": 3, "total_pages": 32, "image_filename": "19930094516_p3.jpg", "text": "NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nTECHNICAL MEMORANDUM NO. 901\n\nINVESTIGATIONS AND EXPERIMENTS\n\nIN THE GUIDONIA SUPERSONIC WIND TUNNEL*\n\nBy Antonio Ferri\n\nINTRODUCTION\n\nThe initial period of activity at the supersonic wind tunnel at Guidonia was devoted to the problem of designing and building the experimental equipment necessary for systematic research. This equipment consisted primarily of\n\n1) A number of different subsonic and supersonic cones or ducts designed to generate the desired speed in the experiment chamber.\n\n2) An aerodynamic balance.\n\n3) Optical instruments operating on the Schlieren and interferometric principle.\n\nThis period of study was long and laborious since practically no previous experimental data were available for guidance. Progress was of necessity slow before any definite decision could be made. During this period we designed a number of speed cones, an aerodynamic balance, and an optical plant, which, although in part still under construction, may be looked upon as being the final products.\n\nIt might be of interest to point out the points of view which emerged from completed experiments with the aid of such provisory equipment. Parallel with this activity some systematic studies in the field of sound were carried out, resulting in part in altogether new results.\n\n---\n\n*\"Untersuchungen und Versuche im Überschallwindkanal zu Guidonia,\" Reprint of paper presented at meeting of Lilienthal-Gesellschaft für Luftfahrtforschung, October 12-15, 1938, Berlin.", "timestamp": "2026-07-19T18:49:24.949754+00:00"}
{"citation_id": "19930094493", "source_url": "https://ntrs.nasa.gov/api/citations/19930094493/downloads/19930094493.pdf", "page_number": 9, "total_pages": 31, "image_filename": "19930094493_p9.jpg", "text": "8\nN.A.C.A. Technical Memorandum No. 923\n\nThe power supplied to the wire is given by\n$$U = 0.86 \\ i^2 \\ r(\\text{kcal/h}) \\quad (2)$$\n\nThe voltage on the wire changes within certain limits as a result of the voltage drop in the instrument and lead resistances. With the wire placed transverse to the air flow, the cooling of the wire is strongest, the resistance thus the lowest and the current the strongest. The voltage drop will therefore also be the maximum. If care is taken, however, to see that the latter is small compared to the wire voltage, as may be done by keeping the instrument and lead resistances small, then the heat conducted to the air stream, which heat is equal to the electrical energy supplied, is a maximum when the current through the wire is a maximum. The transverse air-flow case is thus characterized not only by the maximum current in the wire, but also by the maximum power absorption of the latter. If the ratio of the wire resistance to the instrument and conducting lead resistances were equal to about 1 or less, then the power need not increase with increasing current but will even decrease. In the present case, the most unfavorable ratio was about 8.\n\nFigure 2 shows the oscillogram curves for the wire used as thermometer and hot wire. For the temperature curves about 10 cycles were photographed above one another. For the hot-wire curves, it is necessary, in order to obtain the maximum speeds for each piston position, to rotate the wire at right angles to the air flow. If the flow curves are therefore obtained for different positions of the hot wire in the cylinder, the envelope of the family of curves at each position gives the maximum velocities as a function of the crank angle. The hot-wire curves are also taken for nine spindle settings (from 20° to 20°) for about 10 cycles. This gives altogether about 90 cycles. The curves for any definite spindle setting do not cover each other but give a scattered band which practically coincides with the bands of the other eight spindle settings. The direction of the velocity vector during the individual cycles cannot be determined on account of the strong scattering of the curves at a definite spindle setting.\n\nc) The Measuring Instrument\n\n1. Construction of the Measuring Instrument\n\nThe wire element employed for the measurements is", "timestamp": "2026-07-19T18:49:25.202662+00:00"}
{"citation_id": "19930094498", "source_url": "https://ntrs.nasa.gov/api/citations/19930094498/downloads/19930094498.pdf", "page_number": 9, "total_pages": 37, "image_filename": "19930094498_p9.jpg", "text": "8\nN.A.C.A. Technical Memorandum No. 918\n\nin use in most European tanks, particularly in that no\nbatteries are used and in that the uniformity of speed of\nthe towing carriage is dependent entirely on the uniform-\nity of the voltage output of the motor generator sets.\n\nIn connection with the extension of the tank, a two-\nstory office building was added at the south end of the\ntank and the north end of the shop was extended 100 feet.\n\nOperation of tank.- The data that are obtained on the\ntowing carriage of the N.A.C.A. tank during the tests of a\nmodel are:\n\nSpeed\n\nResistance\n\nTrim\n\nTrimming moment\n\nLift developed by the hydrofoil device\n\nDraft (or rise of center of gravity)\n\nThe equipment for obtaining these quantities (the\ntowing gear) consists of the dynamometer, the towing girder,\nand the balance linkage and is arranged as shown on figure\n4.\n\nSpeed is measured by determining the distance traveled\nin a definite time. The distance is obtained from the dis-\ntance tape, a steel tape 1 inch wide that extends from one\nend of the tank to the other. It is secured at the south\nend of the basin and rests on supporting brackets that ex-\ntend below the bottom chords of the roof trusses. At the\nnorth end of the basin it passes over a sheave and is held\nunder tension by a weight. Holes 1/2 inch in diameter occur\nin the tape every 5 feet throughout its length.\n\nTwo sheaves, carried above the top of the main girder\nof the carriage, lift the tape off the brackets and guide\nit through a horizontal slot in one side of a small box.\nOn one side of the slot is a source of light and on the\nother a photoelectric tube. The light beam normally is in-\ntercepted by the tape but falls on the tube each time a\nhole passes through the slot and the energy generated in\nthe tube causes a small solenoid to tilt a tiny mirror and\ndeflect a beam of light.", "timestamp": "2026-07-19T18:49:35.596011+00:00"}
{"citation_id": "19930094568", "source_url": "https://ntrs.nasa.gov/api/citations/19930094568/downloads/19930094568.pdf", "page_number": 11, "total_pages": 28, "image_filename": "19930094568_p11.jpg", "text": "10 N.A.C.A. Technical Memorandum No. 848\n\nsmall $\\beta$ or, in other words,\n\n| On the planing surface | On airfoil |\n| :--- | :--- |\n| $m = \\mu \\ m_o$ | $m' = \\mu' \\ m_o'$ (3) |\n\nthen $\\mu$ (and $\\mu'$) can only be dependent on $t/b$.\n\nFor the lift of this long plate, the following is therefore applicable:\n\n| On the planing surface | On airfoil |\n| :--- | :--- |\n| $R = \\mu \\ P_o$ | $R' = \\mu' \\ R_o'$ |\n| $= \\mu \\ m_o \\ V^2 \\ \\beta$ | $= \\mu' \\ m_o' \\ V^2 \\ \\beta$ |\n| $= \\mu \\ \\frac{1}{2} \\ \\rho \\ \\frac{\\pi b^2}{4} \\ V^2 \\ \\beta$ | $= \\mu' \\ \\rho \\ \\frac{\\pi b^2}{4} \\ V^2 \\ \\beta$ (4) |\n\nwhereby $R_o$ and $R_o'$ denote the lift coefficients computed according to the theory for infinitely small $\\beta$ (that is, according to equation (2), fig. 9).\n\nWintor's studies afford a double check on this line of reasoning for the airfoil:\n\n1. Does $\\mu'$ actually depend on $t/b$ only?\n2. Does the result of the airfoil tests for the case of very small immersion depth (i.e., small $\\beta$) with $\\mu' = 1$ agree with equation (4)?\n\nNote 1: Figure 12 gives the lift ratio $\\mu' = \\frac{A'}{A_o'} = \\frac{m'}{m_o'}$ for various airfoil lengths (parameter) and various $\\beta$ (= $5^\\circ$, $10^\\circ$, and $15^\\circ$) against $\\frac{t}{b} = \\frac{l \\beta}{b}$. The individual values of $\\mu$ were obtained from\n\n$$ \\mu' \\ \\rho \\ \\frac{\\pi b^2}{4} \\ V^2 \\ \\beta = c_a \\ \\frac{\\rho}{2} \\ V^2 \\ b \\ l $$\n\n$$ \\mu' = \\frac{2}{\\pi} \\ \\frac{l}{b} \\ \\frac{c_a}{\\beta} $$", "timestamp": "2026-07-19T18:49:41.961194+00:00"}
{"citation_id": "19930093281", "source_url": "https://ntrs.nasa.gov/api/citations/19930093281/downloads/19930093281.pdf", "page_number": 12, "total_pages": 22, "image_filename": "19930093281_p12.jpg", "text": "10\n\nREFERENCES\n\n1. Theodorsen, Theodore, Brevoort, M. J., and Stickle, George W.: Full-Scale Tests of N.A.C.A. Cowlings. T.R. No. 592, N.A.C.A., 1937.\n\n2. Brevoort, M. J., Stickle, George W., and Ellerbrock, Herman H., Jr.: Cooling Tests of a Single-Row Radial Engine with Several N.A.C.A. Cowlings. T.R. No. 596, N.A.C.A., 1937.\n\n3. Brevoort, M. J., and Joyner, U. T.: Cooling on the Front of an Air-Cooled Engine Cylinder in a Conventional Engine Cowling. T.R. No. 674, N.A.C.A., 1939.\n\n4. Weick, Fred E., and Wood, Donald H.: The Twenty-Foot Propeller Research Tunnel of the National Advisory Committee for Aeronautics. T.R. No. 300, N.A.C.A., 1928.", "timestamp": "2026-07-19T18:49:47.441529+00:00"}
{"citation_id": "19930091719", "source_url": "https://ntrs.nasa.gov/api/citations/19930091719/downloads/19930091719.pdf", "page_number": 11, "total_pages": 33, "image_filename": "19930091719_p11.jpg", "text": "FIVE FULL-SCALE PROPELLERS IN THE PRESENCE OF A RADIAL AND A LIQUID-COOLED NACELLE 7\n\n[Figure: Thrust-coefficient curves for propeller 5868-9, 3 blades, radial engine nacelle.]\n\n[Figure: Design chart for propeller 5868-9, 3 blades, radial engine nacelle.]\n\n80709-38-2", "timestamp": "2026-07-19T18:50:03.183264+00:00"}
{"citation_id": "19930094506", "source_url": "https://ntrs.nasa.gov/api/citations/19930094506/downloads/19930094506.pdf", "page_number": 6, "total_pages": 24, "image_filename": "19930094506_p6.jpg", "text": "4 N.A.C.A. Technical Memorandum No. 910\n\nvalue of $\\lambda = \\frac{Qs}{V}$ could not increase excessively in the experiments. The best way for obtaining high speeds would have been with high angular velocities at sufficiently high tunnel speeds, in order to run the test at the largest possible Reynolds Numbers.\n\nOn account of the severe speed decrease due to the rotating screen, it was impractical to raise the dynamic pressure above 30 kg/m², the maximum wind-tunnel speeds used in spinning-balance tests, at which the strength of the model and incipient oscillations form an upper dynamic pressure limit.\n\nAt this first trial of the new arrangement, the highest dynamic pressure could not yet be utilized in spite of various improvements, because of difficulties with the rotating screen, which, for lack of time, could not be remedied. Measurements made for maximum-lift appraisal at 15, 20, and 25 kg/m² dynamic pressures manifested, in agreement with other investigations, no appreciable influence of the Reynolds Number in this range; as a result the tests were in general run at the low dynamic pressure of 15 kg/m². The choice of low tunnel speeds made it possible to obtain fairly satisfactory rotation values and hence of the effect of rotation on the loads and moments.\n\nSince, in view of the difficulties, readily removed in subsequent tests with the rotating jet, the Reynolds Number was disproportionately small, it was attempted to increase the equivalent Reynolds Number by means of a turbulence grid built up of parallel round bars. It could not be mounted downstream from the rotating screen, as it would have destroyed part of the created turbulence. A turbulence grid made of radial bars which would have to be solidly mounted on the rotating screen would obviate this difficulty.\n\nLacking a hot-wire anemometer, the turbulence measurements were made with a calibrating sphere (140 mm diameter). They gave a turbulence factor of 2.7 for the non-rotating jet and a much lower figure for jet rotation (at a jet rotation of $n = 2.5 \\text{ s}^{-1}$, it dropped to around 2). However, it is very likely that the still somewhat crude test method is unsuitable for the rotating jet. The effective Reynolds Number at 15 kg/m² dynamic pressure was $7.5 \\times 10^5$ (reference length: mean wing chord).", "timestamp": "2026-07-19T18:50:07.859500+00:00"}