Buckets:
| {"citation_id": "19930094557", "source_url": "https://ntrs.nasa.gov/api/citations/19930094557/downloads/19930094557.pdf", "page_number": 10, "total_pages": 20, "image_filename": "19930094557_p10.jpg", "text": "8 N.A.C.A. Technical Memorandum No. 859\n\ntransmitted to the wing, the casing of the motor was pivoted and connected by a wire with a scale. An interruptor disk on the motor axle recorded the revolutions by electric timer and stop watch.\n\nThe peripheral speed of the jet was measured with a light wind vane (fig. 4) carried along by the jet without slip (fig. 10). Every rotation closed a small electric contact so that the revolutions could be recorded by electric timer and stop watch (as on the motor). The axial velocity was recorded with Prandtl tube and micromanometer.\n\nThe wing was either measured in the spinning range through the air forces or in the unstalled flow range driven by motor and the respective autorotation and damping moment measured on the balance. This moment was occasionally held constant for a test series and the wing rotation progressively replaced by jet rotation. It was found that the rotations of the wing decreased exactly by the amount of jet rotations, until finally the wing came to rest, when the jet rotations reached the initial rotations of the wing for static jet. Figures 5 to 8 illustrate the results, the abscissas denoting the jet rotations and the ordinates the wing rotations. It is readily seen how for different angles of attack and autorotation or damping moments the rotation of the model can be replaced by the corresponding rotation of the jet. Disregarding minor discrepancies probably due to imperfections of the first attempts, the practicability of the method has been proved by the tests.\n\nOn conclusion of the experiments we measured the rolling and damping moments on an M 5 airfoil section at 10, 20, and 30 degrees angle of attack for different rates of roll. First, came the measurements on the rotating wing with the aid of the above described electrical rotation device. Two sets of measurements were taken; one in the calm stream with the screen necessary for the jet rotation, the other without screens. Then the same measurements were repeated in the rotating jet; the model being suspended from the six-component balance (figs. 9 and 10). Figure 11 shows the recorded moments about the jet axis against the ratio $U/v$ at the wing tip for all three arrangements. The curves for the rotating jet reach only as far as $\\frac{U}{v} = 0.3$, since technical defects of the original version prohibited higher tip speeds of the", "timestamp": "2026-07-19T18:10:07.078690+00:00"} | |
| {"citation_id": "19930091717", "source_url": "https://ntrs.nasa.gov/api/citations/19930091717/downloads/19930091717.pdf", "page_number": 14, "total_pages": 14, "image_filename": "19930091717_p14.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-19T18:10:08.759077+00:00"} | |
| {"citation_id": "19930094554", "source_url": "https://ntrs.nasa.gov/api/citations/19930094554/downloads/19930094554.pdf", "page_number": 33, "total_pages": 37, "image_filename": "19930094554_p33.jpg", "text": "N.A.C.A. Technical Memorandum No. 862\nFigs. 11,12\n\nd/D\n0 .25 .5 .75 1\n20\n15\nFailing load, t\n10\n5\n0\n5 10 15 20\nNotch diameter, d\nD = 20$\\phi$\n25\n90°\nFigure 11.\n\n200\n150\nFailing stress $\\sigma_B$, kg/mm$^2$\n100\n50\n0\n1 2 3 4\nD/d\nFigure 12.", "timestamp": "2026-07-19T18:10:11.308289+00:00"} | |
| {"citation_id": "19930091692", "source_url": "https://ntrs.nasa.gov/api/citations/19930091692/downloads/19930091692.pdf", "page_number": 6, "total_pages": 20, "image_filename": "19930091692_p6.jpg", "text": "2\nREPORT NO. 617—NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\n(reference 12) has employed a similarly defined lag in his investigation of methods of rating fuels in a modified C. F. R. engine.\n\nAPPARATUS AND PROCEDURE\n\nThe apparatus consisted essentially of an electrically heated stainless-steel bomb, provided with an injection system capable of delivering a single charge of fuel and with an optical-type indicator for recording pressures photographically. Figure 1 is a diagrammatic sketch with guides to fix the position of the bomb with respect to the optical system. The furnace (C—C, fig. 1) is divided into two parts with the top half hinged to the lower rigid section. The inlet valve J and the exhaust valve E' are so designed that they can be quickly opened or closed. The thermocouple P', inserted through a lug in the side of the bomb to within ¼ inch of the inner wall, served to indicate the bomb temperature. The thermocouple F in the top, or hottest part, of the furnace and the pyrometer U' served to control\n\n[Figure: Diagrammatic sketch of apparatus]\n\nA, compressed-air line.\nB, needle valve.\nC, asbestos-board lagging.\nD, fuel-return line.\nE, air-pressure gage.\nF, furnace-control thermocouple.\nG, fuel-suction line.\nH, high-pressure gage.\nI, fuel-circulating pump.\nJ, quick-closing valve.\nK, furnace shell.\nL, valve to protect gage H.\nM, fuel-pressure gage.\nN, valve to adjust pressure across C'.\nO, coolant-fuel line.\nP, injection valve.\nQ, bomb.\nR, check valve.\nS, high-pressure reservoir.\nT, poppet-valve assembly.\nU, thermometer.\nV, combustion space.\nW, indicator diaphragm.\nX, platinized quartz mirror and staff.\nY, quartz lens.\nZ, light source.\nA', film drum.\nB', hydraulic-pump discharge.\nC', 0.030-inch orifice.\nD', injection tube.\nE', exhaust valve.\nF', monel-metal gasket.\nG', asbestos board.\nH', spark.\nI', light box.\nJ', motor.\nK', phase-timing gears.\nL', power line to furnace.\nM', clutch.\nN', cam.\nO', timing switch.\nP', thermocouple for bomb temperature.\nQ', leads to thermocouple F.\nR', resistance lamps in 230-volt d.-c. line.\nS', condenser, 4 microfarad.\nT', automobile spark coil.\nU', pyrometer controller.\n\nFIGURE 1.—Diagrammatic sketch of apparatus.\n\nof the assembled apparatus. A manually operated hydraulic pump (not shown) was employed to force the fuel through tube B' into reservoir S. The fuel tank contained heating and cooling units that maintained the desired temperature of the circulating fuel (130° F.) at thermometer U for all save a few tests at the highest bomb temperature. The increase in temperature as the fuel passed through the injection valve was approximately 15° F.\n\nThe bomb has a maximum inside diameter of 3 inches, a length from nozzle to indicator diaphragm of about 3½ inches, and a measured volume of 21.7 cubic inches (356 cm³). The bomb support was arranged the furnace temperature and, indirectly, the bomb temperature. The large thermal lag of the bomb relative to that of the heating elements necessitated this arrangement to avoid destruction of these elements.\n\nA sketch of the injection valve and of an enlarged section of the nozzle is shown in figure 2. The heat flow in the neighborhood of the nozzle was minimized by placing the narrow seal (K₁, fig. 2) between the valve body and the bomb some distance back from the nozzle. Thermocouples spot-welded to the interior surface of the bomb wall showed that the cooled region was confined to the curved surface at the valve end of the bomb and that the total temperature difference", "timestamp": "2026-07-19T18:10:13.202470+00:00"} | |
| {"citation_id": "19930091716", "source_url": "https://ntrs.nasa.gov/api/citations/19930091716/downloads/19930091716.pdf", "page_number": 7, "total_pages": 24, "image_filename": "19930091716_p7.jpg", "text": "NEGATIVE THRUST AND TORQUE OF SEVERAL FULL-SCALE PROPELLERS 3\n\n$$nD/V$$\n\n| $T_c$ | 0 | .1 | .2 | .3 | .4 | .5 | .6 | .7 | .8 | .9 | 1.0 | 1.1 | 1.2 | 1.3 |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| 0 | | | | | | | | | | | | | | |\n| -.01 | | | | | | | | | | | | | | |\n| -.02 | | | | | | | | | | | | | | |\n| -.03 | | | | | | | | | | | | | | |\n| -.04 | | | | | | | | | | | | | | |\n| -.05 | | | | | | | | | | | | | | |\n| -.06 | | | | | | | | | | | | | | |\n| -.07 | | | | | | | | | | | | | | |\n| -.08 | | | | | | | | | | | | | | |\n\n[Figure: Graph showing curves for blade angles 70°, 65°, 60°, 55°, 50°, 45°, 40°, 35°, 30°, 25°, 20°, and 15° Blade angle at 0.75R]\n\n| $Q_c$ | 0 | | | | | | | | | | | | | |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| -.001 | | | | | | | | | | | | | | |\n| -.002 | | | | | | | | | | | | | | |\n| -.003 | | | | | | | | | | | | | | |\n| -.004 | | | | | | | | | | | | | | |\n| -.005 | | | | | | | | | | | | | | |\n| -.006 | | | | | | | | | | | | | | |\n\n[Figure: Graph showing curves for blade angles 70°, 65°, 60°, 55°, 50°, 45°, 40°, 35°, 30°, 25°, 20°, and 15° Blade angle at 0.75R. Includes dashed lines for $Q_h$ director lines and locus of points having $Q_h = -0.0050$. Axes labeled $Q_h$ with values -.010, -.0075, -.0050, -.0025, -.0010]\n\nFIGURE 3.—Negative thrust and torque coefficients for propeller 5868-9, Clark Y section, 2 blades.", "timestamp": "2026-07-19T18:10:14.527527+00:00"} | |
| {"citation_id": "19930094534", "source_url": "https://ntrs.nasa.gov/api/citations/19930094534/downloads/19930094534.pdf", "page_number": 3, "total_pages": 21, "image_filename": "19930094534_p3.jpg", "text": "2 N.A.C.A. Technical Memorandum No. 882\n\nthe manufacturing tolerances of tubes in their absolute value are smaller than the permissible departures (up or down), tubular shafts require only outside drawing (contraction) and tube holes only inside drawing (expansion). If greater accuracy (grade 8) is specified, a change of shape is effected in such a manner that, by outside drawing the expansion is effected first and, by inside drawing, the contraction first. Both processes are effected in one operation through suitable arrangement of the tools.\n\nThe dimensions of the tools must be so determined that the size of the subsequently drawn tubing remains within the desired limits. Owing to the elastic spring-back of the material, the tool dimensions do not agree with the nominal dimensions of the piece of work. For higher requirements on accuracy (grade 8) they are determined empirically, whereby material as well as diameter and wall thickness must be considered separately.\n\nThe design of the drawing tools is conceivably simple although the tools do make severe demands on the quality of workmanship as regards surface, accuracy, and wear resistance. (Here the pioneer work was performed by the C. H. Bernhardt Special Tool factory, Dresden, and the success of the method is in a large measure due to their efforts.) The operation of the tools is shown in figure 1.\n\nFor inside drawing (or expanding) to $D_i$ diameter H 11, only a mandrel is used; for outside drawing (contracting) to $D_o$ diameter h 11, a drawplate is sufficient. For more accurate (inside expanding drawing), say, to $D_i$ diameter H 8, the mandrel is followed by a drawplate, which first reduces the diameter from the outside so that on return of the tool the mandrel gives the desired inside dimension. For outside drawing (contracting), as illustrated in figure 1, the process is, of course, reversed. The subsequent drawing of the tube ends is appropriately made on a vertical hydraulic press, which offers the simplest solution for smooth operation and satisfactory lubrication of the tools. A drilling-oil emulsion (50 percent) is used.\n\nThe working speed should not exceed 1 m/min, to allow the material sufficient time for form changing. For the starting and return movements, the higher speed easily obtainable with hydraulic presses is useful and therefore recommended for economical reasons.", "timestamp": "2026-07-19T18:10:20.904301+00:00"} | |
| {"citation_id": "19930093641", "source_url": "https://ntrs.nasa.gov/api/citations/19930093641/downloads/19930093641.pdf", "page_number": 5, "total_pages": 47, "image_filename": "19930093641_p5.jpg", "text": "3\n\n1. Propellers located about 0.39c ahead of the leading edge of the wing (tractor position 1). Extension-shaft housings 4 inches in diameter (fig. 4).\n\n2. Propellers located about 0.26c ahead of the leading edge of the wing (tractor position 2).\n\n a. Extension-shaft housings 4 inches in diameter.\n\n b. Extension-shaft housings 8 inches in diameter to represent air-cooled engine cowlings on the same airplane scaled to 100 tons gross weight (fig. 5).\n\n3. Propellers located about 0.13c ahead of the leading edge of the wing (tractor position 3).\n\n a. Extension-shaft housings 4 inches in diameter (fig. 6).\n\n b. Extension-shaft housings 8 inches in diameter.\n\nE. Wing alone without fuselage or nacelles.\n\nFor all the arrangements with motors enclosed in the wing, there were no radiators on the model. For convenience of reference, arrangements with enclosed motors and extension shafts to propellers have been designated by the propeller position, e.g., pusher, tractor position 1, etc.\n\nSYMBOLS\n\n$\\alpha_T$, angle of attack of the fuselage reference axis relative to the wind axis, deg.\n\n$q$, dynamic pressure, lb. per sq. ft.\n\n$S$, Wing area, sq. ft.\n\n$\\bar{c}$, mean chord of the wing, area/span, ft.", "timestamp": "2026-07-19T18:10:25.446455+00:00"} | |
| {"citation_id": "19930094533", "source_url": "https://ntrs.nasa.gov/api/citations/19930094533/downloads/19930094533.pdf", "page_number": 30, "total_pages": 51, "image_filename": "19930094533_p30.jpg", "text": "28 N.A.C.A. Technical Memorandum No. 883\n\nIn the face of such discrepancies, which are much greater than in the model test, the interest in the wing measurement drops considerably. To compare the theoretical (table VI) with the experimental results (table IV), we limit ourselves to plotting the temperature readings.\n\nVI. Temperature Readings\n\nThe procedure is the same as for the model; the curves present the theoretical $\\theta - t_0$, and the points, the corresponding experimental values. The eight readings (figs. 35 to 42) relate to the eight explored incidences. The graph for $6.1^\\circ$ was plotted by means of the values observed on November 17.\n\nVII. COMPARISON\n\nThe tests made on November 17 must be separated from those made on November 18.\n\nAs to the five test series of November 17, it may be said that the experimental check with the theoretical results. The discrepancies do not exceed $10^\\circ$ except for two points over 65 and they correspond to two consecutive measurements (test orifices 4 and 5, incidence $3.1^\\circ$).\n\nContrariwise, the three test series of November 18, disclosed large and systematic discrepancies at orifices 2, 3, 4, and 5 which, at great positive incidences, correspond to the depression zone over the top surface. One plausible explanation for this strange behavior is that the imputed observations are precisely those made at the beginning of the tests. The tests may have been made rather fast on November 18 — the steady regime not quite reached — so that the junctions had not as yet attained their steady temperature at the time the measurement was taken. In fact, as already stated in V (p. 27), the November 18 measurements do not seem to confirm those of November 17, and this holds for both the temperature and the pressures. I do not think that much importance attaches to the discrepancies of November 18.\n\nIn conclusion, and bearing in mind — above all — the fidelity of the results observed on the model, it may be said that the simple theory explained in section A, is", "timestamp": "2026-07-19T18:10:26.916209+00:00"} | |
| {"citation_id": "19930091712", "source_url": "https://ntrs.nasa.gov/api/citations/19930091712/downloads/19930091712.pdf", "page_number": 22, "total_pages": 22, "image_filename": "19930091712_p22.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-19T18:10:50.374111+00:00"} | |
| {"citation_id": "19930094558", "source_url": "https://ntrs.nasa.gov/api/citations/19930094558/downloads/19930094558.pdf", "page_number": 24, "total_pages": 24, "image_filename": "19930094558_p24.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-19T18:10:51.446647+00:00"} | |
| {"citation_id": "19930094542", "source_url": "https://ntrs.nasa.gov/api/citations/19930094542/downloads/19930094542.pdf", "page_number": 58, "total_pages": 102, "image_filename": "19930094542_p58.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-19T18:10:57.773361+00:00"} | |
| {"citation_id": "19930094543", "source_url": "https://ntrs.nasa.gov/api/citations/19930094543/downloads/19930094543.pdf", "page_number": 37, "total_pages": 50, "image_filename": "19930094543_p37.jpg", "text": "N.A.C.A. Technical Memorandum No. 873 35\n\nIn this case, auto-ignition probably induces a combustion substantially the same as that by normal ignition.\n\nDiagram (4) at the top of figure 24 corresponds to critical temperature $T_2 = 1,045^\\circ$ C. The action of the hot spot produces auto-ignition with an advance superior to $0^\\circ$. At times the auto-ignition is accompanied by detonation; the combustion is very irregular.\n\nLastly, diagram (5) of figure 25 corresponds to a hot-spot temperature distinctly superior to the critical values, reaching, in fact, $1,190^\\circ$.\n\nThe auto-ignition is advanced so that combustion terminates under a very low pressure. Then follows the compression of the burned gases. (The power output of the engine is very low or zero, and electricity must be re-sorted to, to maintain the speed.)\n\nThe interesting fact here is that detonation disappears in the case of materially advanced auto-ignition as already pointed out by the author several years ago.\n\nSUMMARY\n\nThe working up of the different diagrams discloses the following:\n\n1. At minimum temperature on appearance of auto-ignition the combustion produced by hot spot proceeds along a regime substantially the same as with electric ignition at zero advance.\n\n2. At the temperature of regularization of auto-ignition the combustion released by it is more advanced than the normal combustion, which may induce the detonation and tend to further increase the temperature of the hot spot.\n\n3. An increase of less than $200^\\circ$ in hot-spot temperature ushers in a regime for which auto-ignition is practically indistinguishable from a regime of zero horsepower.", "timestamp": "2026-07-19T18:11:00.115052+00:00"} | |
| {"citation_id": "19930094554", "source_url": "https://ntrs.nasa.gov/api/citations/19930094554/downloads/19930094554.pdf", "page_number": 34, "total_pages": 37, "image_filename": "19930094554_p34.jpg", "text": "N.A.C.A. Technical Memorandum No. 862\nFigs. 15,18,22\n\n[Figure: Graph with x-axis labeled $\\frac{a}{b}$, % ranging from 0 to 50. Left y-axis labeled $\\frac{\\sigma_{BL}}{\\sigma_B}$, % ranging from 80 to 110. Right y-axis labeled $\\sigma_{BL}$ kg/mm$^2$ ranging from 26 to 34. The graph contains multiple curves with data points marked by circles, triangles, and crosses. Annotations include \"Hole diameter\" with values 0.625, 1.25, 2.5, 3 mm, 5, 10, 20.]\n\nFigure 15.\n\n[Figure: Graph with x-axis labeled \"Hole diameter, mm\" ranging from 0 to 12. Left y-axis labeled $\\sigma_{BL}$ kg/mm$^2$ ranging from 22 to 32. Right y-axis labeled $\\frac{\\sigma_{BL}}{\\sigma_B}$ ranging from 80 to 110. Two curves are plotted with data points marked by circles. Annotations indicate $\\frac{a}{b} = 30\\%$ and $\\frac{a}{b} = 12.5\\%$.]\n\nFigure 18.\n\n[Figure: Graph with x-axis labeled $\\frac{\\sigma_{BL}}{\\sigma_B}$, % ranging from 90 to 100. Y-axis labeled \"Frequency, %\" ranging from 0 to 30. A bell-shaped curve is plotted with data points marked by circles. A dashed vertical line indicates the peak at 94.4.]\n\nFigure 22.", "timestamp": "2026-07-19T18:11:02.858118+00:00"} | |
| {"citation_id": "19930094534", "source_url": "https://ntrs.nasa.gov/api/citations/19930094534/downloads/19930094534.pdf", "page_number": 4, "total_pages": 21, "image_filename": "19930094534_p4.jpg", "text": "N.A.C.A. Technical Memorandum No. 882 3\n\nThe pressures necessary for the drawing of thin-walled tubing commonly used in airplane design range up to about 2 tons.\n\nEmployed for shaping the process is the same as for drawing, though the necessary pressures may, depending on tube diameter, wall thickness, and degree of contraction, assume values approaching the stability of the original tube. By dividing the contraction process into several stages interspersed with heat treatment, very elaborate form changes can be effected.\n\nIn new constructions it is frequently necessary to find the pressure required for a certain contraction in advance, so that a decision may be made as to a particular machine and the number of necessary stages. To follow this process mathematically is very difficult and the attack of this problem by some competent party is earnestly desired. For the present Heinkel follows a formula which, in the explored cases varies <15 percent from the experimental values:\n\n$$\nP = \\frac{\\pi k}{4} \\frac{\\sin \\alpha + \\mu}{\\sin \\alpha \\cos \\alpha} \\frac{(\\epsilon^4 - 1)}{\\epsilon^2} D s \\sigma_{0.2}\n$$\n\nwhere\n\nP is contracting pressure (kg)\n\nk, material constant\n\n$\\epsilon = \\frac{D}{d}$, degree of contraction\n\n$\\mu$, coefficient of friction\n\n$\\sigma_{0.2}$ yield point (kg/mm²)\n\nD, d, and s in mm (cf. fig. 2).\n\nThe material constant k ranges around 1.3 for the tools used by Heinkel. The design of the tools is governed by the angle of contraction $\\alpha$ and the probable coefficient of friction. With predetermined degree of contraction, the contracting pressure is conditional on the factor $\\frac{\\sin \\alpha + \\mu}{\\sin \\alpha \\cos \\alpha}$. Thus figure 3 shows the influence", "timestamp": "2026-07-19T18:11:03.913863+00:00"} | |
| {"citation_id": "19930091716", "source_url": "https://ntrs.nasa.gov/api/citations/19930091716/downloads/19930091716.pdf", "page_number": 8, "total_pages": 24, "image_filename": "19930091716_p8.jpg", "text": "4\nREPORT NO. 641—NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\n<!-- Image (160, 120, 879, 860) -->\n\nFIGURE 4.—Negative thrust and torque coefficients for propeller 5868-9, Clark Y section, 3 blades.", "timestamp": "2026-07-19T18:11:05.530143+00:00"} | |
| {"citation_id": "19930091692", "source_url": "https://ntrs.nasa.gov/api/citations/19930091692/downloads/19930091692.pdf", "page_number": 7, "total_pages": 20, "image_filename": "19930091692_p7.jpg", "text": "never exceeded 270° F. The valve was cooled by two fuel streams (see figs. 1 and 2): The main stream passed through tube O, then passage H₁, jacket L₁, and the four passages (two of which are indicated as I₁) to the return line B₁; the smaller stream passed through orifice C', injection tube D', passage G₁, and the two small holes in the valve stem to be mixed with the main stream in the return line B₁ (or D). The area of these stem holes and of orifice C' is so much smaller than that of the nozzle orifices that a relatively small portion of the hydraulic injection impulse was dissipated at these points. The pressure in the line between pump 1 and orifice C' was adjusted to 150 pounds per square inch\n\nThe indicator was calibrated by recording photographically the deflections corresponding to several static gas pressures over the range of interest with the indicator bolted to the hot bomb as for an explosion test. The calibration pressures were determined by a Bourdon gage, which had been checked against a dead-weight gage tester. The error involved in measuring the records is believed to be larger than any involved in the calibration procedure and, since most deflections were relatively small, the derived data may involve appreciable errors. These errors are not particularly important, however, as trends, rather than exact magnitudes, are of primary interest. Except, perhaps, for extremely high rates of pressure rise, it is believed that this indicator, in view of its high natural frequency, satisfactorily recorded the instantaneous explosion pressures. Very great rates of pressure rise, such as accompanied the larger fuel weights under conditions giving long ignition lags, invariably led to severe vibration, which loosened the mirror staff in its bearings.\n\nThe fuel employed in these tests was found by the U. S. Naval Engineering Experiment Station, Annapolis, Md., to have the following characteristics:\n\n| Diesel index | 70.1 |\n| --- | --- |\n| Aniline point | 183.7 |\n| Specific gravity, 60/60° F | 0.834 |\n| A. P. I. gravity | 38.2 |\n| Flash point, closed cup | ° F | 242 |\n| Cloud point | ° F | 28 |\n| Pour point | ° F | 25 |\n| Color N. P. A | 1.0 |\n| Saybolt Universal viscosity: | |\n| At 32° F | seconds | 87 |\n| At 100° F | do | 43 |\n| Carbon residue | percent | 0.01 |\n| Sulphur | do | 0.04 |\n| Heat value, calorimeter gross | B. t. u. per pound | 1,9996 |\n\nDistillation characteristics:\n\n| First drop | ° F | 526 |\n| --- | --- | --- |\n| 5 cm³ | | 531 |\n| 10 cm³ | | 532 |\n| 20 cm³ | | 538 |\n| 30 cm³ | | 546 |\n| 40 cm³ | | 553 |\n| 50 cm³ | | 562 |\n| 60 cm³ | | 572 |\n| 70 cm³ | | 584 |\n| 80 cm³ | | 599 |\n| 90 cm³ | | 627 |\n| End point | | 681 |\n| Recovered | percent | 98.3 |\n\nThe cetane number was found to be 64 on the Diesel conversion of the C. F. R. engine in conjunction with the modified magnetic pick-up method recommended in reference 16. This ignition quality compares favorably with the better commercial fuels (reference 17).\n\nDuring tests the pyrometer controller was set to give a furnace temperature corresponding to the desired bomb temperature, the circulating pump 1 (see fig. 1)\n\n[Figure: Twelve-orifice nozzle. Plane DD' is identical with BB'. Plane CC' is identical with AA'. A₁, thermometer for fuel temperature. B₁, connection between valve and fuel return line. C₁, upper lock nut for stem stop. D₁, lock-nut and stem-top support. E₁, nut for adjusting spring load. F₁, lock washers. G₁, injection passage. H₁, passage for main coolant stream. I₁, return passages from jacket K. J₁, oil seal around valve body. K₁, sealing surface. L₁, coolant jacket around end of valve. Injection-valve assembly.]\n\nby means of the bypass valve N. Small readjustments of this valve were necessary for each bomb temperature. Seal J₁, which prevents fuel leakage into the furnace, consists of a number of turns of soft electric fuse wire forced tightly against the threads on the valve body by the clamping nut shown. The fact that this wire is satisfactory in spite of its low melting point is a good indication of the effectiveness of the cooling system.\n\nThe optical indicator is an adaptation of the one described in reference 15. All parts were constructed of a high-tungsten steel for which the manufacturer claims a yield point of 120,000 pounds per square inch at 1,100° F. The platinized quartz mirror proved fairly satisfactory although it gradually lost its mirror finish and reflectivity at the higher temperatures. It was necessary, therefore, to retouch certain of the prints of the original records in order to get satisfactory half-tone reproduction.", "timestamp": "2026-07-19T18:11:10.413330+00:00"} | |
| {"citation_id": "19930094533", "source_url": "https://ntrs.nasa.gov/api/citations/19930094533/downloads/19930094533.pdf", "page_number": 31, "total_pages": 51, "image_filename": "19930094533_p31.jpg", "text": "N.A.C.A. Technical Memorandum No. 883 29\n\npractical enough for evaluating the temperatures at different points of a single wing and deducing the distribution of the pressures and temperatures. In particular, in the case of the wing investigated here, we take as results to be examined, those which express the theoretical curves plotted above (figs. 35 to 42).\n\nD. APPLICATION OF TEMPERATURE MEASUREMENTS ON A WING\n\nI. Wing Warmer Than Air\n\nThe plotted temperature readings prove that in steady regime the different zones of a wing are all warmer than the air in which the wing moves.\n\n1. At points of the wing where positive pressure prevails, the heat effects due to friction and those due to compression, all contribute separately to the heating of the wing. The positive pressure, incidentally, is maximum in the dead point, for which we have the equality\n\n$p = p_0 + \\rho_0 v_0^2/2$; for this point $(p - p_0)/(\\rho_0 v_0^2/2)$ is maximum and equal to 1. Consequently, according to equation (13), the heating is also maximum and equal to\n\n$\\theta_m - t_0 = \\frac{v_0^2}{2c_p}$\n\nwhich, with $c_p = 10^7$ ergs/gram deg., gives:\n\n$\\theta_m - t_0 = 5 \\times 10^{-8} v_0^2$\n\nHence, at points of the wing where it is maximum, the heating reaches $5^\\circ$ for a speed of 360 k.p.h. (223 m.p.h.).\n\n2. At the points on the wing where depression prevails, the thermic effects due to friction and those due to depression are restrained, but the usual result is a heating because the thermic effect due to friction exceeds, as a rule, that due to depression. What is necessary, in fact, in order that a point of the wing shall have the temperature of the undisturbed flow? It is necessary that the value of $(p - p_0)/(\\rho_0 v_0^2/2)$ is such that the difference $\\theta - t_0$ becomes zero:", "timestamp": "2026-07-19T18:11:13.075014+00:00"} | |
| {"citation_id": "19930091705", "source_url": "https://ntrs.nasa.gov/api/citations/19930091705/downloads/19930091705.pdf", "page_number": 5, "total_pages": 10, "image_filename": "19930091705_p5.jpg", "text": "REPORT No. 630\n\nA FLIGHT COMPARISON OF CONVENTIONAL AILERONS ON A RECTANGULAR WING AND OF CONVENTIONAL AND FLOATING WING-TIP AILERONS ON A TAPERED WING\n\nBy H. A. SOULÉ and W. GRACY\n\nSUMMARY\n\nFlight tests comparing the relative effectiveness of conventional ailerons of the same size on wings of rectangular and tapered plan forms were made with a Fairchild 22 airplane. Information is included comparing conventional and floating wing-tip ailerons on a tapered wing. The results showed that the conventional ailerons were somewhat more effective on the tapered than on the rectangular wing. The difference, however, was so small as to be imperceptible to the pilots. The floating wing-tip ailerons were only half as effective as the conventional ailerons and, for this reason, were considered unsatisfactory.\n\nINTRODUCTION\n\nAt the request of the Matériel Division of the Army Air Corps, the N. A. C. A. has conducted a series of flight tests to compare the relative effectiveness of conventional ailerons of a given size on wings having rectangular and tapered plan forms. Earlier wind-tunnel tests are reported in references 1, 2, and 3. The flight tests were made with two Fairchild 22 airplanes. The two wings used in the investigation were of the same area and span. One had a rectangular plan form with semicircular tips and the other a taper ratio of 2:1. The conventional ailerons with which these wings were fitted had the same plan-form dimensions and were arranged during the flight tests to have approximately the same deflections.\n\nThe tests consisted of the determination of the effectiveness of the ailerons (1) for different degrees of deflection at two air speeds, and (2) for full deflection at various air speeds throughout the speed range of the airplane. The comparisons are based on the maximum measured rolling accelerations and velocities, the observed yawing action, and the computed rolling-moment coefficients.\n\nIn addition to being fitted with the conventional ailerons, the tapered wing was equipped with detachable wing tips that could be replaced by floating wing-tip ailerons. The floating wing-tip ailerons were also tested during the investigation and were compared with the conventional ailerons on the same wing.\n\nAIRPLANES AND WINGS\n\nThe Fairchild 22 airplanes used in the investigation are shown in figures 1 and 2. The rectangular wing, which had the same plan form as the standard wing for the Fairchild 22 airplanes, had a span of 32 feet 10 inches, a chord of 5 feet 6 inches, an area of 171 square feet, and an N. A. C. A. 2R₁₂ airfoil section. The conventional ailerons with which this wing was fitted had a span of 13 feet 3⁵⁄₁₆ inches (81 percent b/2) and a chord of 12 inches (18 percent c). They were\n\n[Figure: Fairchild 22 airplane used for tests of conventional ailerons on a rectangular wing.]\n\n[Figure: Fairchild 22 airplane used for tests of conventional ailerons on a tapered wing.]\n\noperated differentially, having a maximum upward deflection of 17° and a downward deflection of 9°.\n\nThe tapered wing (figs. 2 to 5) had the same span and area as the rectangular wing. It had a 2:1 taper ratio with a straight trailing edge. The trailing edge was made straight so that the aerodynamic centers of the tapered and rectangular wings could be located at the same point relative to the fuselage while still permitting access to the rear cockpit. In external dimensions the tapered wing was comparable with an in-", "timestamp": "2026-07-19T18:11:18.036098+00:00"} | |
| {"citation_id": "19930094549", "source_url": "https://ntrs.nasa.gov/api/citations/19930094549/downloads/19930094549.pdf", "page_number": 1, "total_pages": 76, "image_filename": "19930094549_p1.jpg", "text": "FILE COPY\nNO. 2\nFILE COPY\nNO. I-W\nN 62 57867\n\nTECHNICAL MEMORANDUMS\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nNo. 967\n\nTHEORETICAL STUDY OF VARIOUS AIRPLANE MOTIONS\nAFTER INITIAL DISTURBANCE\nBy Fr. Haus\n\nBulletin du Service Technique de L'Aéronautique\nNo. 17, June 1937\n\nWashington\nJune 1938\n\nFILE COPY\nTo be returned to\nthe files of the National\nAdvisory Committee\nfor Aeronautics\nWashington, D.C.", "timestamp": "2026-07-19T18:11:25.781053+00:00"} | |
| {"citation_id": "19930094552", "source_url": "https://ntrs.nasa.gov/api/citations/19930094552/downloads/19930094552.pdf", "page_number": 14, "total_pages": 32, "image_filename": "19930094552_p14.jpg", "text": "12 N.A.C.A. Technical Memorandum No. 864\n\nstarting from an arbitrary zero value and the final zero line is so determined that the sum of the areas of the $\\tau(y)$ curves over the semiperimeter (for example, between the intersections with the vertical plane) vanishes (the semiperimeter is sufficient on account of the antisymmetry of the stress distribution).\n\nThe shear stresses thus determined are plotted against the longitudinal coordinate $x$ and the derivatives $\\frac{\\partial \\tau}{\\partial x}$ obtained graphically. For determining the transverse stresses, the equation\n\n$$\n\\frac{\\partial \\sigma_y}{\\partial y} = - \\frac{\\partial \\tau}{\\partial x}\n\\tag{1b}\n$$\n\nis integrated with respect to $y$, it being necessary to determine initial values for the transverse stresses.\n\nIn the bonding case the halves of the cylinder on each side of the horizontal plane are antisymmetrically loaded. The setting up of a peripheral stress in this intersecting plane would be contrary to the mutual interaction principle and therefore in the horizontal plane $\\sigma_y = 0$. In the arching case both the vertical and horizontal planes are antisymmetric planes and hence free from transverse stresses. Maxima (over $y$) occur in the bonding case in the section of the vertical plane ($\\frac{\\partial \\tau}{\\partial x} = 0$ for all values of $x$) and in the arching case at the main spars.\n\nThe determination of the shear stresses may be checked by cutting the cylinder along the length and comparing the shearing forces in the intersecting planes with the applied forces ($P$ at each spar, fig. 11) and the normal forces in the radial cuts. If in the bending case the cylinder is cut in the horizontal plane, then\n\n$$\n\\int_0^x \\tau s \\, dx = 2P - \\int \\sigma_x \\, dF\n$$", "timestamp": "2026-07-19T18:11:31.767526+00:00"} | |
| {"citation_id": "19930091715", "source_url": "https://ntrs.nasa.gov/api/citations/19930091715/downloads/19930091715.pdf", "page_number": 22, "total_pages": 30, "image_filename": "19930091715_p22.jpg", "text": "```markdown\n18\nREPORT NO. 640—NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nAt peak efficiency the value of the ratio $C_T/J^2$ decreases with blade angle, so that the loss in peak efficiency caused by an increase of solidity may be expected to be less at the higher blade angles.\n\nThe effect on propeller characteristics of changing the number of blades is illustrated by the $C_T$, $C_P$, and $\\eta$ coefficient curves slope less steeply. The curves come together at $C_T=0$ because here the slipstream velocity becomes zero.\n\nThe efficiency curves show a loss for the 3-blade and 4-blade propellers less than would be calculated from the simple momentum theory; in fact, in the portion of the curves where the blades begin to stall, the 3-blade and 4-blade propellers have a somewhat higher efficiency than the 2-blade propeller. The higher efficiency of the 3-blade and 4-blade propellers in the stalled range may be accounted for by the fact that their higher inflow velocities delay and reduce the severity of their stalling.\n\nThe effect of a change in solidity has been commonly thought to be the same whether the change results from variation in blade width or number of blades. Modern\n\n<!-- Image (66, 176, 492, 432) -->\nFIGURE 33.—Comparison of thrust coefficients for propellers having 2, 3, and 4 blades.\n\n<!-- Image (526, 329, 932, 576) -->\nFIGURE 35.—Comparison of efficiency curves for propellers having 2, 3, and 4 blades.\n\n<!-- Image (66, 462, 492, 787) -->\nFIGURE 34.—Comparison of power coefficients for propellers having 2, 3, and 4 blades.\n\ncurves of figures 33, 34, and 35. Values are given for blade angles of $25^\\circ$ and $35^\\circ$ for all three propellers of Clark Y section.\n\nOwing to the increase in inflow velocity, which theoretically is equal to $(V_s+V)/2$, with increasing number of blades, the thrust-coefficient and power- theory and experimental evidence show this belief to be untrue.\n\nThe modern vortex theory of propellers pictures each propeller blade, as it describes its helical path through the air, leaving a continuous sheet of vortices behind it which, if no rotational interference velocity is assumed, moves straight backward with slipstream velocity in a manner somewhat similar to the way a screw conveyor appears to move. The strength and backward velocity of the vortex sheets depend on the strength of circulation around the blade itself which, of course, varies with the thrust and therefore with the blade width.\n\nThe air trapped between the sheets moves backward with them except for the part that slips forward around the edges of the sheets and produces a tip or edge vortex. The edge vortex destroys some of the circulation of the blade and produces what is known as a \"tip loss.\"\n\nPrandtl has shown (reference 4) that the edge flow, and therefore the tip loss, is reduced if the normal distance between two consecutive vortex sheets is reduced. The distance between vortex sheets is reduced\n```", "timestamp": "2026-07-19T18:11:45.495574+00:00"} | |
| {"citation_id": "19930094554", "source_url": "https://ntrs.nasa.gov/api/citations/19930094554/downloads/19930094554.pdf", "page_number": 35, "total_pages": 37, "image_filename": "19930094554_p35.jpg", "text": "N.A.C.A. Technical Memorandum No. 862\nFigs.16,17,19,20,21,23,24\n\n<!-- Image (117, 110, 907, 999) -->\n\nFigure 16.\n\nFigure 20.\n\nFigure 17.\n\nFigure 21.\n\nFigure 19.\n\nFigure 23.\n\nFigure 24.", "timestamp": "2026-07-19T18:12:04.130484+00:00"} | |
| {"citation_id": "19930094543", "source_url": "https://ntrs.nasa.gov/api/citations/19930094543/downloads/19930094543.pdf", "page_number": 38, "total_pages": 50, "image_filename": "19930094543_p38.jpg", "text": "36 N.A.C.A. Technical Memorandum No. 873\n\nInformation Supplied by the Diagrams Regarding \nthe Phenomena of Dissociation of Carbureted Mixture \nat High Temperatures\n\nAn analysis of diagrams (1) and (4) has revealed, \namong other things, some particularly interesting information. On transforming these diagrams into p v axes \n(figs. 26 and 27), we find that the mean polytropic coefficient of expansion (between 0.05 and 0.75 of the stroke) \nis:\n\n1.25 if no auto-ignition exists, and\n\n1.26 for very advanced auto-ignition (diagram 4).\n\nThe mean polytropic coefficient of compression is, in \nthe latter case, only 0.96. This seems to point to release of heat during expansion, even for very advanced \nauto-ignition.\n\nOwing to the abnormally small value of mean polytropic coefficient of compression, it cannot be admitted that \nthis release of heat is due to a simple phenomenon of late burning when there is auto-ignition. The specific heat \nchanges of the gases with temperature are no longer of \nsufficient order of magnitude to allow for the result obtained. It seems very likely that the products of the advanced combustion due to the auto-ignition are subject \nduring compression to a very material dissociation followed by an equally very intense recombination during expansion.\n\nIt is difficult to directly verify the proper base of \nthis conclusion with the few diagrams, by reason of the \nrelatively great importance of the wall losses in the test \nengine and, in any case, the question is clearly beyond \nthe scope of the present study. Even so, these facts appeared striking enough to merit particular mentioning.\n\nCONCLUSION\n\nFrom the results outlined in Parts II and III, the \nfollowing conclusions can be reached.\n\nFrom the theoretical point of view:", "timestamp": "2026-07-19T18:12:05.937383+00:00"} | |
| {"citation_id": "19930094533", "source_url": "https://ntrs.nasa.gov/api/citations/19930094533/downloads/19930094533.pdf", "page_number": 32, "total_pages": 51, "image_filename": "19930094533_p32.jpg", "text": "30 N.A.C.A. Technical Memorandum No. 883\n\n$$\n\\theta - t_0 = v_0^2 \\left[ A + \\frac{p - p_0}{\\rho_0 \\, v_0^2 / 2} \\left( \\frac{1}{2c_p} - A \\right) \\right] = 0\n$$\n\nwhich gives:\n\n$$\n\\frac{p - p_0}{\\rho_0 \\, v_0^2 / 2} = \\frac{A}{A - \\frac{1}{2} \\, c_p}\n$$\n\nWith $ A = 4.2 \\times 10^{-8} \\text{ deg.}/(\\text{cm/s})^2 $\n\n$$\n\\frac{p - p_0}{\\rho_0 \\, v_0^2 / 2} = - \\frac{4.2}{0.8} \\sim 5\n$$\n\nThis value is not encountered in ordinary flying conditions.\n\nII. Change of Temperature with the Speed\n\nThe temperature at the various points increases proportionally to the square of the speed if the law of pressure distribution — that is, to say, the local values of $(p - p_0)/(\\rho_0 \\, v_0^2 / 2)$ — does not vary with the Reynolds Number (equation 13).\n\nIn passing from the model to the wing tests, the dimensions are altered in the ratio of 5 to 1.1, and the speed in the ratio of 38 to 45, so that the Reynolds Number is approximately multiplied by 4. Are the values of $(p - p_0)/(\\rho_0 \\, v_0^2 / 2)$ noticeably changed? The test taps, numbers 3 and 12 on the wing (one on top, the other on bottom surface) are plainly coincident with those on the model. Plotting for taps 3 and 12, the curves giving $(p - p_0)/(\\rho_0 \\, v_0^2 / 2)$ against the incidence (figs. 43 and 44) it is observed that, even when disregarding that which occurs at high incidence where the discrepancies are very pronounced, the differences between the curves for the wings and those of the model are largely above 10 percent. These differences can even become greater if one passes from the case of the wing at 38 m/s to the practical case of a wing at 76 m/s (270 k.p.h.). Hence the statement that this local heating is proportional to the square of the speed, is an approximation.", "timestamp": "2026-07-19T18:12:09.246521+00:00"} | |
| {"citation_id": "19930091692", "source_url": "https://ntrs.nasa.gov/api/citations/19930091692/downloads/19930091692.pdf", "page_number": 8, "total_pages": 20, "image_filename": "19930091692_p8.jpg", "text": "was started, and the flow of cooling water through the coils in the fuel tank was adjusted to give the requisite fuel temperature. The bomb was filled with air to the desired pressure, needle valve B being used for close control of the air flow. Fuel was then forced into reservoir S to a predetermined pressure such that the desired weight of fuel could be injected. Finally, the motor J' was started, valve J closed to protect gage E, and the injection made by means of a “trip-hammer” mechanism\n\never mass-action effects there may have been in the ignition and combustion of the fuel. The air-fuel ratios were based upon the weight of air in the bomb prior to injection without considering the small amount of air compressed, as a result of combustion, into the small inlet and exhaust passages and into the space about the end of the injection valve. In any case, these ratios are not indicative of the wide range of actual air-fuel ratios from point to point in the fuel spray. The desired fuel quantities were obtained by varying the pressure in the reservoir S between 4,600 and 7,600 pounds per square inch.\n\nRESULTS\n\nTypical records for an air-fuel ratio of 30, reproduced in figure 4, show the effect of air temperature on the ignition lag at a density of 0.59 pound per cubic foot. Ignition-lag data taken from these and similar records for air-fuel ratios ranging from 20 to 80 are plotted in figure 5, together with data obtained at twice this density, 1.18 pounds per cubic foot.\n\nA similar set of records (fig. 6) for an air-fuel ratio of 20 shows the effect of air density on the ignition\n\nthat permitted the engagement of clutch M' for a single revolution of the driving shaft. The motion of the trip-mechanism handle automatically closed the circuit for lamp Z for a period beginning before injection and extending beyond the combustion period. The resulting film trace shows both the initial pressure and the pressure changes resulting from combustion. This film, together with the reference trace (zero gage pressure) taken before admitting air to the bomb, constitutes the pressure record.\n\nThe engagement of the clutch M' lifted a poppet valve T by means of cam N', thus admitting the full pressure from reservoir S to the injection line D'. This operation also closed switch O', producing a spark at gap H' and a corresponding trace on the film at right angles to the constant-pressure traces. This spark served to denote the start of injection on the film record, the two having been properly phased by means of the gears K' as described in reference 18, before the tests were begun.\n\nThe air pressures used were arbitrarily selected to give densities corresponding to 5, 10, 15, 20, 25, and 30 atmospheres absolute at 212° F. For convenience in the examination of the experimental results, the relations of temperature, density, and gage pressure are shown graphically in figure 3. Air density rather than air pressure was used as one of the primary variables because of the better correlation with spray development (reference 19) and of the better control of what-\n\nlag at the highest temperature (1,155° F.) for which the indicator was suitable for continuous service. Ignition-lag data obtained with various air-fuel ratios at this and several other gas temperatures are shown in figure 7. A summary of the data for figures 5 and 7 is given in table I.\n\nThe effectiveness of combustion (insofar as it is defined by the ratio of the pressure 0.004 second after ignition to the initial pressure) is shown as a function of the ignition lag by the solid curves of figure 8. These curves correspond to data obtained at two air densities and at several air temperatures (870°, 1,060°, 1,155°, and 1,255° F.). The 0.004-second period was arbitrarily taken as the longest in which combustion of the fuel would efficiently produce power in a moderately high-speed engine. For this reason the highest pressure indicated on the records reproduced in the", "timestamp": "2026-07-19T18:12:11.142777+00:00"} | |
| {"citation_id": "19930094534", "source_url": "https://ntrs.nasa.gov/api/citations/19930094534/downloads/19930094534.pdf", "page_number": 5, "total_pages": 21, "image_filename": "19930094534_p5.jpg", "text": "4 N.A.C.A. Technical Memorandum No. 882\n\nof the contraction angle on the contracting pressure for different $\\mu$.\n\nThe high surface pressures rule out fluid friction ($\\mu \\approx 0.2$). The best contracting angle, even with regard to form rigidity was found to be $\\alpha = 15^\\circ$.\n\nThe limit value of the degree of contraction follows from the condition\n\n$$\n\\left(1 + \\frac{\\mu}{\\sin \\alpha}\\right) \\frac{1}{4 \\cos \\alpha} \\frac{\\epsilon^4 - 1}{\\epsilon^3} k = 1\n$$\n\nwhereat the pressure of contraction would become equal to the crushing limit of the tube. In practice the diameter is contracted in stages of 10 mm; for tube diameters of less than 30 mm, the stages are 5 mm. To avoid waste of aluminum tubing, the cold hardening caused by contracting is removed after each stage by heat treatment. Even tubes contracted in one stage only are given a final heat treatment. Only after-drawn tubes, that is, those whose nominal diameter has not been changed, are excluded from heat treatment. During the contraction process the original cross-sectional area of the tube is not altogether preserved, since part of the material flows in the direction of the tube axis and so lengthens the contracted tube. The reduction in area approximately follows the empirical formula:\n\n$$\n\\frac{F_1 - F_2}{F_1} \\approx 0.4 \\frac{D_1 - D_2}{D_1}\n$$\n\nSubscripts 1 and 2 refer to the areas and diameters before and after form change (fig. 4).\n\nTubes of any origin, provided the material and dimensions are approved for airplane construction, are suitable for after-drawing. According to the investigations in the Heinkel shops on coarse recrystallization in the cold-shaping of light alloys, however, duralumin tubes are not suitable for extended form-changing unless their structure is fine-grained.\n\nThe expanding process is similar to the contracting", "timestamp": "2026-07-19T18:12:15.651905+00:00"} | |
| {"citation_id": "19930094542", "source_url": "https://ntrs.nasa.gov/api/citations/19930094542/downloads/19930094542.pdf", "page_number": 59, "total_pages": 102, "image_filename": "19930094542_p59.jpg", "text": "N.A.C.A. Technical Memorandum No. 874\nFig.34\n\nO — Without propeller\nx — With \"\n\n[Figure: Four graphs arranged in a 2x2 grid. Each graph plots two curves (marked with O and x) against a horizontal axis labeled 0, 3, 6, 9, 12, 15, α and a vertical axis labeled 0, 2, 4, 6, 8, δ°. The top-left graph is labeled a/t = 0. The top-right graph is labeled a/t = 0.351. The bottom-left graph is labeled a/t = 0.715. The bottom-right graph is labeled a/t = 1.125.]\n\nFigure 34.- Variation of downwash angle.", "timestamp": "2026-07-19T18:12:32.068760+00:00"} | |
| {"citation_id": "19930094556", "source_url": "https://ntrs.nasa.gov/api/citations/19930094556/downloads/19930094556.pdf", "page_number": 28, "total_pages": 29, "image_filename": "19930094556_p28.jpg", "text": "N.A.C.A. Technical Memorandum No. 860\nFigs. 20, 21 22, 23\n\n<!-- Image (261, 117, 775, 412) -->\n\nFigure 20.- Optimum resistance for all models at the same load.\n\n<!-- Image (199, 492, 492, 631) -->\n<!-- Image (575, 483, 888, 631) -->\n\nFigure 21.- Check on beam at step. Figure 22.- Check on beam at step.\n\n<!-- Image (379, 703, 637, 946) -->\n\nFigure 23.-\nDetermination\nof angle\nbetween\nwing\nand\nplaning\nbottom.", "timestamp": "2026-07-19T18:12:34.614563+00:00"} | |
| {"citation_id": "19930094549", "source_url": "https://ntrs.nasa.gov/api/citations/19930094549/downloads/19930094549.pdf", "page_number": 2, "total_pages": 76, "image_filename": "19930094549_p2.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-19T18:12:35.092808+00:00"} | |
| {"citation_id": "19930091705", "source_url": "https://ntrs.nasa.gov/api/citations/19930091705/downloads/19930091705.pdf", "page_number": 6, "total_pages": 10, "image_filename": "19930091705_p6.jpg", "text": "2\nREPORT NO. 630—NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nternally braced wing although it was supported externally for the tests. The airfoil section varied from an N. A. C. A. 2218 section at the root to an N. A. C. A. 2209 section at 15 feet from the axis of symmetry. The\n\naileron on the rectangular wing. They were operated differentially and, for the tests, were limited so that the maximum upward deflection was 18° and the downward deflection 9°. Owing to differences in the aileron-operating mechanism, the maximum aileron deflections on the tapered wing were obtained with a stick deflection of 14°; whereas, with the rectangular wing, the maximum deflections were obtained with a stick deflection of 20°. The plan view in figure 3, on which the rectangular wing has been drawn in outline, gives a direct comparison of the wings and the conventional-aileron installation.\n\n<!-- Image (100, 148, 508, 503) -->\n\nFIGURE 3.—Three-view drawing showing the installation of the tapered wing on a Fairchild 22 airplane.\n\n<!-- Image (541, 282, 917, 452) -->\n\n(a) Installation of fixed wing tip.\n\n<!-- Image (541, 472, 917, 653) -->\n\n(b) Installation of floating wing-tip aileron.\nFIGURE 5.—View of right wing.\n\n<!-- Image (100, 539, 508, 798) -->\n\nFIGURE 4.—Section through tapered wing at outboard end of conventional aileron.\n\nwing tips were rounded. The chord varied from 7 feet 4 inches at the root to 3 feet 8 inches at the 15-foot station.\n\nThe conventional ailerons on this wing had the same span and chord and were located in the same position relative to the wing span as were the conventional\n\nFor the installation of the floating wing-tip ailerons, the fixed tips of the tapered wing outboard of the 15-foot station were removed and the conventional ailerons were locked in their neutral position. The floating wing-tip ailerons had a symmetrical N. A. C. A. 0009 airfoil section at the root. Each aileron had an area of 7.9 square feet and a span of 35 inches; the wing area and the span with these ailerons were 177 square feet and 35 feet 10 inches, respectively. These ailerons were statically balanced about a hinge axis 17 percent back of their leading edges and were permitted to float freely between limiting positions of 40° up and 30° down. The ailerons could be deflected relative to one another to obtain a maximum angular difference of 30° with a stick movement of 24°.", "timestamp": "2026-07-19T18:12:40.197693+00:00"} | |
| {"citation_id": "19930093641", "source_url": "https://ntrs.nasa.gov/api/citations/19930093641/downloads/19930093641.pdf", "page_number": 6, "total_pages": 47, "image_filename": "19930093641_p6.jpg", "text": "4\n\nV, air speed, f.p.s.\n\nL, lift, or force normal to the relative wind, lb.\n\nD, drag, or force parallel to the relative wind, lb.\n\nDc, power-off drag of combination, lb.\n\nR, resultant drag, force of a propeller-body combination, lb.\n\nT, thrust of propellers operating in front of a body (tension in propeller shafts), lb.\n\nM, pitching moment, lb.-ft.\n\nΔD, increase in drag of the body behind the propellers due to the action of the propellers.\n\nT - ΔD, effective thrust of the propeller-body combination.\n\n$C_L = L/qS$\n\n$C_D = D/qS$ (Subscript w refers to power-off drag of the model with bare wing; c, to power-off drag of the model with engine-propeller arrangement; h.s., to drag at high speed; min, to minimum drag.)\n\n$C_m = M/qS\\bar{c}$\n\nP, total power input to propellers.\n\n$\\eta = \\frac{(T - \\Delta D) V}{P} =$ propulsive efficiency.\n\n$\\eta_t = \\eta \\left( \\frac{C_{D_w}}{C_{D_c}} \\right) =$ over-all efficiency.\n\n$T_{c_o}' = \\frac{P \\eta_o}{\\frac{1}{2} \\rho V^3 S} =$ index thrust coefficient.\n\n$\\eta_o = \\eta$ at $C_L = 0.25$\n\nn, propeller revolution speed, r.p.s.", "timestamp": "2026-07-19T18:12:42.351513+00:00"} | |
| {"citation_id": "19930094543", "source_url": "https://ntrs.nasa.gov/api/citations/19930094543/downloads/19930094543.pdf", "page_number": 39, "total_pages": 50, "image_filename": "19930094543_p39.jpg", "text": "N.A.C.A. Technical Memorandum No. 873 37\n\n1. The critical temperatures of auto-ignition by hot spot is increased as:\n\na) the pressure of carburetion is decreased;\n\nb) the period of contact between this mixture and the hot spot is increased;\n\nc) the carburetion is farther away from a richness corresponding to 2 percent of CO at exhaust;\n\nd) the octane number of the fuel is increased.\n\n2. For an extremely low speed and sufficiently weak turbulence, the hot-spot temperature capable of causing auto-ignition in the engine approaches the spontaneous ignition temperature of the employed fuel.\n\n3. Taken as a whole, the variations observed for the critical temperatures of auto-ignition, are the same as those assumed by the author regarding the spontaneous ignition temperature of fuels in his detonation theory.\n\nHowever, the agreement between these variations is more qualitative than quantitative, as is easily proved either by comparing the ignition-temperature variations with the pressure, as may be deduced from figures 10, 15, or 17 with the experimental relation previously determined in the case of nuclear ignition (No. 103 of this series), or by comparing the effect of ethyl fluid with the critical temperatures of auto-ignition and with the critical temperature corresponding to the appearance of detonation.\n\nIn any case, the critical temperatures of ignition by hot spot remain, for the same ignition lag, distinctly above the critical temperatures which we have had to consider regarding the detonation, and it remains doubtful whether the difference can be attributed to a temperature difference between hot spot and gas which becomes heated on contact.\n\nThe laws governing the ignition by hot spot of carbureted mixtures, though similar in entirety to the laws governing the ignition in mass of the same mixtures, do not seem to harmonize with the latter. This, it is said, conforms to the result of applying the reaction theory \"by chains\" to the ignition of combustible mixtures.", "timestamp": "2026-07-19T18:12:49.436384+00:00"} | |
| {"citation_id": "19930094542", "source_url": "https://ntrs.nasa.gov/api/citations/19930094542/downloads/19930094542.pdf", "page_number": 60, "total_pages": 102, "image_filename": "19930094542_p60.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-19T18:12:54.055851+00:00"} | |
| {"citation_id": "19930094534", "source_url": "https://ntrs.nasa.gov/api/citations/19930094534/downloads/19930094534.pdf", "page_number": 6, "total_pages": 21, "image_filename": "19930094534_p6.jpg", "text": "N.A.C.A. Technical Memorandum No. 832 5\n\nprocess. Duralumin tubing can also be expanded in original condition (aircraft material 3115,5) up to 8 percent of its diameter. Contracted or expanded tubes can be subsequently drawn with tolerances up to grade 8 in the same way as tubes in the original condition.\n\nThe proper preparation within the specified tolerances of the tube ends is almost exclusively carried out in the cited manner by cold-working, whether the tubes are of light metal or steel (VCM, Cr-Mo). It was natural to apply the same method to the calibration of tubes of greater lengths; first, in order to adapt tubular shafts for receiving levers, bearings, etc.; second, to obtain cylinders of thinner walls than are obtainable by machining. Undoubtedly, the tubing purchased from supply houses is accurate enough, but from the point of view of stock-keeping it is desirable to get along with the common commercial products and to do the necessary redrawing in one's own shop. The tools required for this are of the same type as for the redrawing of tube ends. All that is necessary is the proper control so that the fluctuations in wall thickness occurring in the commercial tubes do not lead to rejections. As it pertains to long work, the tools are clamped in a horizontal hydraulic machine of special design. The lubrication is assured by pressure pump. The chief requisite for an unobjectionable finish of the piece of work and careful treatment of the tools is the scrupulous removal of dirt and rust from the tubes before redrawing.\n\nDrawn cylinders are spot-welded onto the cylinder bottom. A cylinder of this kind of 60 mm diameter and 1 mm gage Cr-Mo steel tubing withstood an inside pressure of 280 atm. - equivalent to the strength of the material - and even then the weld remained undamaged. Figure 5 illustrates the experimental cylinder after machining off the collar on the bottom of the cylinder for a study of the weld. Cylinders of this type offer considerable advantages both as regards weight as well as economy.\n\nOwing to the temperature drop at high altitude the operation of airplane controls is preferably accomplished with rods of material of the same heat expansion as used for the fuselage and the wings. For this purpose especially, the contracting process offers great advantages because it makes it possible to give the push-pull control rods, which are primarily stressed in buckling, a structurally beneficial shape, so that their ends require less", "timestamp": "2026-07-19T18:12:55.622728+00:00"} | |
| {"citation_id": "19930094554", "source_url": "https://ntrs.nasa.gov/api/citations/19930094554/downloads/19930094554.pdf", "page_number": 36, "total_pages": 37, "image_filename": "19930094554_p36.jpg", "text": "N.A.C.A. Technical Memorandum No. 862\nFigs.25,26,30,32,33,34\n\n200\n$\\sigma_{B2c}$\n$kg/mm^2$\n150\n100\n50\n0\n$\\circ$ $d=1 mm$\n$+$ $d=0.5 mm$\n50\n100\n$\\sigma_B$ $kg/mm^2$ 200\nFigure 25.\n\n200\n$kg/mm^2$\n150\n100\n50\n0\n$\\sigma_B$ $\\sigma_{BL}$\n$\\circ$ $\\bullet$ 10 mm\n$\\square$ $\\blacksquare$ 0.5 \"\n$\\triangle$ $\\blacktriangle$ 0.6 \"\n$\\circledcirc$ $\\bullet$ 0.10 \"\nSpecimen\nthickness\n\"\n\"\n\"\n\"\n0\n2\n4\n6\n8 Max./mm 10\n12\nCorrosion period/sheet thickness\nFigure 32.\n\n200\n$\\sigma_B$\n$kg/mm^2$\n160\n120\n$\\sigma_{BL}$\n$kg/mm^2$\n80\n40\n0\nStrength\n200\n400\n600\n$^\\circ C$ 800\nAnnealing temperature.\nFigure 26.\n\n120\n$\\sigma_{BL}/\\sigma_B$\n%\n110\n100\n90\n80\n0\n10\n20\n30 $d/b$ 40\n50 %\nHalf round rivet\n$\\sigma_{BL}$\n$kg/mm^2$\n34\n32\n30\n28\n26\n24\nFlathead rivet\nWithout\nrivet\nFigure 33.\n\n200\n$kg/mm^2$\n150\n100\n50\n0\nStrength\n2\n4\n6 Months 8\nCorrosion period\n$\\sigma_B$\n$\\sigma_{B2c}$\nSteel 1\n$\\sigma_{BL}$\n$\\sigma_{B2c}$\nSteel 2\nFigure 30.\n\n120\n$\\sigma_{B}/kg$\n110\n100\n90\n80\n0\n10\n20\n30 $d/b$ 40\n50 %\nNumber of\nrivets\nFigure 34.", "timestamp": "2026-07-19T18:12:58.815275+00:00"} | |
| {"citation_id": "19930094549", "source_url": "https://ntrs.nasa.gov/api/citations/19930094549/downloads/19930094549.pdf", "page_number": 3, "total_pages": 76, "image_filename": "19930094549_p3.jpg", "text": "NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nTECHNICAL MEMORANDUM NO. 867\n\nTHEORETICAL STUDY OF VARIOUS AIRPLANE MOTIONS\nAFTER INITIAL DISTURBANCE*\n\nBy Fr. Haus\n\nI. INTRODUCTION\n\nThe study of the dynamic stability of airplanes by\nthe method of small motions involves the determination of\noscillatory and aperiodic motions. There are a great\nnumber of well-established methods that permit us to cal-\nculate the periods and the damping factors of these mo-\ntions but which do not, in general, give any indication of\nthe amplitudes of the motions set up. We shall in the\npresent paper, with the aid of a number of numerical ex-\namples, attempt to clarify the phenomena arising after a\nseries of typical initial disturbances. It will be found\nthat the results of these calculations contribute to an\nunderstanding of the mechanism of the motions which are\nset up, and it is our belief that the publication of these\nresults, although limited to particular cases, will facil-\nitate the study of the motions and flight paths. In what\nfollows, we shall strictly separate the study of the lon-\ngitudinal from that of the lateral motions and shall not\nconsider the complete theory.\n\nNOTATION\n\nThe OX, OY, and OZ axes will be assumed fixed on the\nairplane (fig. 1). The projections of the velocity on the\naxes will be denoted by u, v, and w, respectively. They\ndefine the angles of attack i, and of yaw j, these two\nangles fixing the position of the airplane on its path.\nWhen the components v and w are small with respect to\nu, we may write:\n\n$$i = - \\frac{w}{u} \\quad j = + \\frac{v}{u}$$\n\n*\"Etude théorique de quelques trajectoires d'avions apres\nperturbation initiale.\" Bulletin du Service Tech-\nnique de L'Aéronautique, No. 17, June 1937.", "timestamp": "2026-07-19T18:13:09.832677+00:00"} | |
| {"citation_id": "19930094535", "source_url": "https://ntrs.nasa.gov/api/citations/19930094535/downloads/19930094535.pdf", "page_number": 2, "total_pages": 17, "image_filename": "19930094535_p2.jpg", "text": "NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nTECHNICAL MEMORANDUM NO. 881\n\nAPPLICATION AND TESTING OF TRANSPARENT PLASTICS USED\nIN AIRPLANE CONSTRUCTION*\n\nBy K. Riechers and J. Olms\n\nTransparent plastics are used in airplane construction principally to replace silicate glass, having recently become indispensable in the construction of the gunner's cockpit for increasing the field of view. The advantages of the artificial material are its low specific gravity, and the absence of the great brittleness characteristic of ordinary glass. The fracturing of silicate glass is always accompanied by the formation of dagger-shaped, very sharp-edged splinters (fig. 1) that may cause severe or even fatal injuries. The disk shown in figure 1 was tested for its resistance to a high pressure on one side by being subjected to a bursting test through the application on one side of high pressure air. Similar disturbance phenomena arise whenever impulsive stresses are set up as, for example, occur in difficult landings.\n\nIn general, two types of dangerous conditions may arise:\n\n1. If the disk is fractured as a result of the impact, the splinters may be hurled against the occupants of the airplane and may lead to cuts and eye injuries. Eloquent testimony of this are the automobile accident lists, for example, in the newspapers.\n\n2. The occupants may be hurled against or even through the windshield, in which case - while the glass breaks away at the point of collision - there generally remains a ring of sharp-edged splinters in the frame, with the result that serious and even fatal arm and throat injuries may hardly be avoided.\n\n*\"Verwendung und Prüfung durchsichtiger Werkstoffe im Flugzeugbau.\" Luftwissen, vol. V, no. 6, June 1938, pp. 197-202.", "timestamp": "2026-07-19T18:13:13.680078+00:00"} | |
| {"citation_id": "19930094557", "source_url": "https://ntrs.nasa.gov/api/citations/19930094557/downloads/19930094557.pdf", "page_number": 11, "total_pages": 20, "image_filename": "19930094557_p11.jpg", "text": "N.A.C.A. Technical Memorandum No. 859\n\nscreens. Figure 11 discloses the following: for $\\alpha = 10^\\circ$, that is, for unstalled flow over the whole wing in the low U/v range, the measurements on the rotating wing with screen coincide with those for the rotating jet and fixed wing. The same holds true for the measurements with completely separated flow at $\\alpha = 30^\\circ$, against minor discrepancies for the ranges of partial separation of flow ($\\alpha = 20^\\circ$ and $\\alpha = 10^\\circ$, respectively). Inasmuch as it is common knowledge that, in the range of incipient separation, the reproduction of the test values is accompanied by scattering even if the test method is not changed, the agreement of both test methods must be pronounced good.\n\nThe jet rotation method makes investigations possible which were heretofore very difficult; first among these is the measurement of the yawing and pitching moments, which with rotating jet can be read on the normal 6-component balance along with the other quantities, whereas, even on the very latest rotating spinning balances, these components must be measured separately and with the most careful balancing of all parts.\n\nTranslation by J. Vanier, \nNational Advisory Committee \nfor Aeronautics.\n\n---\n\n**REFERENCES**\n\n1. Francis, R. H.: Senkrechter Windkanal für Trudelversuche. Übersetzung in Z.V.D.I., vol. 79, p. 65.\n\n2. Jourawtschenko, A. N.: Eine Methode zur Lösung des Trudelproblems und des Problems der Stabilität und Steuerbarkeit von Flugzeugen bei Fahrtverlust. Arbeiten des Zentralen Aero-Hydrodynamischen Instituts Moskau 1932. Ausgabe 167.\n\n3. Von Doepp, Ph.: Luftschraubenberechnung nach dem Verfahren der gleichwertigen Tragflügel-Polaren. Luftf.-Forsch. vol. 13, no. 2, Feb. 20, 1936.", "timestamp": "2026-07-19T18:13:14.074458+00:00"} | |
| {"citation_id": "19930091707", "source_url": "https://ntrs.nasa.gov/api/citations/19930091707/downloads/19930091707.pdf", "page_number": 14, "total_pages": 20, "image_filename": "19930091707_p14.jpg", "text": "10\nREPORT NO. 632—NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\n<!-- Image (96, 88, 916, 736) -->\n\nFIGURE 12.—Diagram of $f_c$, $d/t$ for chromium-molybdenum steel (transverse loading).\n\nnormalized, it is not unreasonable to expect similar stress-strain curves in compression. If this is the case, then for a given yield strength $S$ for the two types of material and at a given stress $f>S$ at failure, the value of $\\tau$, which depends on the tangent modulus, will be greater for the material with the blunt-knee stress-strain curve than for the material with the sharp-knee stress-strain curve. Now, unquestionably, the strength increases with $\\tau$ and decreases with $d/t$ (whether equation (1) is right or not) and, consequently, for a given strength, represented by $f$, a high value of $\\tau$ will be associated with a high value of $d/t$ and a low value of $\\tau$ with a low value of $d/t$. It is to be expected, then, that the three sets of points representing the tests on the thin tubes in figures 7 and 11 would be shifted to the left. They are \"off the beaten track\" of the other\n\n$^2$ The statement that follows is true not only for $f>S$ but for $f$ greater than the stress at which the slopes of the two stress-strain curves become equal. The values of $E$ are assumed to be equal.", "timestamp": "2026-07-19T18:13:14.773718+00:00"} | |
| {"citation_id": "19930091705", "source_url": "https://ntrs.nasa.gov/api/citations/19930091705/downloads/19930091705.pdf", "page_number": 7, "total_pages": 10, "image_filename": "19930091705_p7.jpg", "text": "COMPARISON OF AILERONS ON A RECTANGULAR AND A TAPERED WING\n3\n\n[Figure: Graphs showing Maximum rolling acceleration and Maximum rolling velocity vs. Aileron deflection for rectangular wing]\nAir speed, f.p.s.\no--- 90\nΔ--- 138\n\n[Figure: Graphs showing Maximum rolling acceleration and Maximum rolling velocity vs. Aileron deflection for tapered wing]\nAir speed, f.p.s.\no--- 84\nΔ--- 138\n\nFIGURE 6.—Variation of the maximum rolling velocities and accelerations with deflection of conventional ailerons on the rectangular wing.\nFIGURE 7.—Variation of the maximum rolling velocities and accelerations with deflection of conventional ailerons on the tapered wing.\n\n[Figure: Graphs showing Maximum rolling acceleration and Maximum rolling velocity vs. Air speed for rectangular and tapered wings]\nConventional ailerons on tapered wing\nΔ--- rectangular wing\no--- Standard Fairchild 22 ailerons and wing (from reference 4)\n\n[Figure: Graphs showing Maximum rolling acceleration and Maximum rolling velocity vs. Aileron deflection for tapered wing with floating wing-tip ailerons]\nAir speed, f.p.s.\no--- 83\nΔ--- 131\n\nFIGURE 8.—Comparison of the maximum rolling velocities and accelerations with full deflection of conventional ailerons on the rectangular and the tapered wings.\nFIGURE 9.—Variation of the maximum rolling velocities and accelerations with deflection of floating wing-tip ailerons on the tapered wing.", "timestamp": "2026-07-19T18:13:17.661120+00:00"} | |
| {"citation_id": "19930094533", "source_url": "https://ntrs.nasa.gov/api/citations/19930094533/downloads/19930094533.pdf", "page_number": 33, "total_pages": 51, "image_filename": "19930094533_p33.jpg", "text": "N.A.C.A. Technical Memorandum No. 863 31\n\nIII. Temperature Distribution\n\n1. Toward the tail the temperature always approaches the value\n\n$\\theta = t_0 + 4.2 \\times 10^{-8} v^2$\n\nIt is little affected by the incidence of the wing.\n\n2. But, near the leading edge the results vary with the incidence. For negative incidence — that is, toward the top camber — the temperature is higher, while on the bottom camber a depression zone defines the cooler region. At positive incidence the depression passes to the top camber; it is there that elevation of the temperature will be least. The bottom camber, on the contrary, is the hot zone.\n\nSome figures are given for an airplane at 300 k.p.h. (obtained on the premise of local heating proportional to the square of the speed). At $-6^\\circ$ incidence, the excess $\\theta - t_0$ of local temperature $\\theta$ over the air temperature $t_0$ is in proximity of the leading edge — $1.5^\\circ$ toward the bottom camber, while reaching $3.5^\\circ$ toward the top camber. At positive incidence the excess $\\theta - t_0$ is $3.5^\\circ$ at the leading edge but may drop over the top camber as much as $1.5^\\circ$ for $15^\\circ$ incidence.\n\nIV. Effect of the Nature of the Wing\n\nIt is to be remembered that the experiments were made on a single wing and that no allowance was made for the thermal conductivity in the wing mass. Incidentally, the heat exchanges of a wing with the atmosphere (by forced convection) are of greater importance than those which may be observed in the material (by conduction); the thickness of the metal of which the wing is made is always small. The substitution of a metal wing for an isolated wing evidently acts in the sense of making the temperatures uniform, tied to the existence of a heat flow from the warmer toward the cooler points, but the results will nevertheless not be substantially different.\n\nTranslation by J. Vanier, \nNational Advisory Committee \nfor Aeronautics.", "timestamp": "2026-07-19T18:13:17.978520+00:00"} | |
| {"citation_id": "19930093641", "source_url": "https://ntrs.nasa.gov/api/citations/19930093641/downloads/19930093641.pdf", "page_number": 7, "total_pages": 47, "image_filename": "19930093641_p7.jpg", "text": "```markdown\nD, propeller diameter, ft.\n\n$\\beta$, propeller blade angle at 0.75 R, deg.\n\n$\\delta_e$, angle of the elevator to the stabilizer (positive when trailing edge of elevator is down), deg.\n\n$\\delta_f$, flap deflection from closed position, deg.\n\n$i_w$, $i_s$, angle of wing and stabilizer setting, respectively, to the reference axis, deg.\n\n$a$, slope of lift curve, $dC_L/d\\alpha$.\n\nAIRPLANE 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 1.\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, exclusive of the tail, with dimensions of the various nacelle-propeller arrangements tested, is shown in figure 7.\n\nEach propeller was driven by a 25-horsepower squirrel-cage induction motor. The speed of the motors was regulated by varying the impressed frequency and was measured by a Weston electrical tachometer. In order that the motor torques might be computed, the motors were calibrated on a dynamometer to determine the power output from the measured electrical input for various combinations of impressed voltage and frequency.\n\nThe motors for the wing-nacelle arrangement were supported in the nacelles ahead of the leading edge of the wing; for the enclosed-engine arrangements, the motors were mounted within the wing between the front and rear spars (fig. 7). The propeller axes for the wing-nacelle\n\n```", "timestamp": "2026-07-19T18:13:23.165131+00:00"} | |
| {"citation_id": "19930094556", "source_url": "https://ntrs.nasa.gov/api/citations/19930094556/downloads/19930094556.pdf", "page_number": 29, "total_pages": 29, "image_filename": "19930094556_p29.jpg", "text": "N.A.C.A. Technical Memorandum No. 860\nFigs. 24,25\n\n<!-- Image (398, 117, 711, 293) -->\n\nFigure 24.- Equilibrium of moments.\n\n<!-- Image (299, 329, 801, 878) -->\n\nFigure 25.- Comparison of DVL standard floats\nwith the N.A.C.A. model No. 35.", "timestamp": "2026-07-19T18:13:28.066527+00:00"} | |
| {"citation_id": "19930094551", "source_url": "https://ntrs.nasa.gov/api/citations/19930094551/downloads/19930094551.pdf", "page_number": 1, "total_pages": 18, "image_filename": "19930094551_p1.jpg", "text": "Y3.N21/5:8/865\nm865\nHartford Public Library\n\nTECHNICAL MEMORANDUMS\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nNo. 955\n\nBEHAVIOR OF STATIC PRESSURE HEADS AT HIGH SPEEDS\nBy Helmut Danielzig\n\nLuftfahrtforschung\nVol. 14, No. 6, June 20, 1937\nVerlag von R. Oldenbourg, München und Berlin\n\nWashington\nJune 1938", "timestamp": "2026-07-19T18:13:38.968685+00:00"} | |
| {"citation_id": "19930091692", "source_url": "https://ntrs.nasa.gov/api/citations/19930091692/downloads/19930091692.pdf", "page_number": 9, "total_pages": 20, "image_filename": "19930091692_p9.jpg", "text": "AUTO-IGNITION AND COMBUSTION OF DIESEL FUEL IN A CONSTANT-VOLUME BOMB 5\n\n[Figure: Pressure vs. Time graphs for records 186, 102, 167, 227]\n\n| Record | Temperature (°F.) | Ignition lag (sec.) | Maximum explosion pressure (lb./sq. in.) |\n| :--- | :--- | :--- | :--- |\n| 186 | 870 | 0.0070 | 750 |\n| 102 | 1,040 | .0021 | 680 |\n| 167 | 1,135 | .0018 | 690 |\n| 227 | 1,255 | .0015 | 640 |\n\nFIGURE 4.—Effect of temperature on ignition lag. Air density, 0.59 pound per cubic foot; air-fuel ratio, 20.\n\n[Figure: Pressure vs. Time graphs for records 57, 168, 154]\n\n| Record | Density (lb./cu. ft.) | Ignition lag (sec.) | Maximum explosion pressure (lb./sq. in.) |\n| :--- | :--- | :--- | :--- |\n| 57 | 0.29 | 0.0036 | 380 |\n| 168 | .59 | .0018 | 800 |\n| 154 | .89 | .0013 | 1,260 |\n\nFIGURE 6.—Effect of air density on ignition lag. Air temperature, 1,165° F.; air-fuel ratio, 20.", "timestamp": "2026-07-19T18:13:57.147733+00:00"} | |
| {"citation_id": "19930094534", "source_url": "https://ntrs.nasa.gov/api/citations/19930094534/downloads/19930094534.pdf", "page_number": 7, "total_pages": 21, "image_filename": "19930094534_p7.jpg", "text": "5 N.A.C.A. Technical Memorandum No. 882\n\nspace for openings, etc. The number of structural parts, while preserving standardization, can be reduced by rolling in the thread in the contracted end. This possibility is the result of the thickening of the wall on contracting, according to the previously cited formula.\n\nA particularly expedient solution is afforded by the use of contracted push rods with rolled thread in combination with the Elf eyes and adjustable eyes with self-aligning bearings (German patent), shown in figure 6. In these the outer ring of the bearings is combined into one with the threaded bolt and made of carburized steel so that only the bearing surface has the necessary hardness while the thread connection itself remains soft and can be drilled with the tube for \"safetying\" with a pin. The dust protection of the bearing is oil and acidproof rubber or similar material in form of a sliding cover.\n\nHeinkel has now accepted this method for duralumin push-pull rods as standard practice after exhaustive tests had proved their practicability. Six different tubes of 1 mm gage with diameter ranging from 25 to 50 mm were contracted and fitted with M 14 x 1.5 rolled thread. Then eight different tubes of 1 mm gage with diameters ranging from 15 to 50 mm were contracted and fitted with rolled thread M 12 x 1.5. The tubes were tested in tension, compression, buckling, and alternating fatigue load, and checked for cross-sectional area reduction. In addition, the tubing material was tested in original and in finished condition.\n\nThe loading arrangement is shown in figures 7 and 8. All tubes invariably failed at the point marked 1, except on the 30 x 1.0 tube with thread M 14 x 1.5, where the break occurred in section A.\n\nTable 1 contains a compilation of the strength test data; the result of the subsequent check is shown in figure 9 as mean value against the corresponding tubes. Figures 10 and 11 show the experimental tubes after buckling.\n\nThe tubes intended for fatigue-testing were loaded together with the eye and adjustable eye terminals, according to figure 6. Three 20 x 1 tubes with M 12 x 1.5 thread, twice with and once without check nut were subjected to a load of $P = \\pm 200$ kg at $n = 82$ stress reversals per minute. No strain, cracks, etc., were noticed after 100,000 stress reversals.", "timestamp": "2026-07-19T18:14:03.790161+00:00"} | |
| {"citation_id": "19930094542", "source_url": "https://ntrs.nasa.gov/api/citations/19930094542/downloads/19930094542.pdf", "page_number": 61, "total_pages": 102, "image_filename": "19930094542_p61.jpg", "text": "N.A.C.A. Technical Memorandum No. 874\nFigs. 37, 38\n\n<!-- Image (173, 95, 847, 936) -->\n\nFigure 37. $\\alpha = 0^\\circ$.\n$$\n\\begin{array}{l}\n\\text{---} d' \\\\\n\\text{---} \\frac{q-q_0}{q}\n\\end{array}\n$$\n$$\n\\alpha = 0^\\circ \\quad \\rightarrow \\frac{y}{R}\n$$\n\nFigure 38. $\\alpha = 4^\\circ$.\n$$\n\\begin{array}{l}\n\\text{---} d' \\\\\n\\text{---} \\frac{q-q_0}{q}\n\\end{array}\n$$\n$$\n\\alpha = 4^\\circ \\quad \\rightarrow \\frac{y}{R}\n$$\n\nDownwash angle and dynamic pressure distribution.", "timestamp": "2026-07-19T18:14:07.327629+00:00"} | |
| {"citation_id": "19930094544", "source_url": "https://ntrs.nasa.gov/api/citations/19930094544/downloads/19930094544.pdf", "page_number": 13, "total_pages": 43, "image_filename": "19930094544_p13.jpg", "text": "N.A.C.A. Technical Memorandum No. 872 11\n\nof two plane trusses in the vertical and horizontal longitudinal planes, intersecting in the airship's axis. The rings are built around those plane trusses and attach to the two plane trusses at the edges of the latter. The obvious disadvantage of this construction is the practically unobtainable lateral stability of the deep plane trusses and, in addition, their deficient torsional stiffness. An advantage of this construction is, perhaps, that a natural attachment of the stabilizing surfaces results and that the vertical plane trusses can be used for supporting weights and the nose for mast mooring without anything additional.\n\nIn all of the more recent airships, however, the previously described basket-work framing has been used. In this construction the transverse rings are designated as main and intermediate rings, depending on whether or not they are stiffened in their own planes. The stiff main rings serve a double purpose. Firstly, they take care of a proportionate share of the external forces on the outer cover which affect the framing; secondly, they divide the total gas space into the individual compartments which serve for the accommodation of the gas cells. In the design of the framing the case of a deflated gas cell is considered. Then the adjacent cells which are still inflated are subjected to large side gas forces, for which either the main rings themselves must be carefully designed, or some other structural provision must be made.\n\nIn the main rings of the more recent rigid airships one may distinguish two different arrangements. In figure 18 they are shown in contrast, above and below. The \"Graf Zeppelin,\" as well as the new airship LZ 129, now under construction, have wire-braced rings. The wire bracing is attached to alternate ring corners; the intermediate sides are constructed as trusses. Also, in the one English airship R 100 no departure from wire-braced rings has been made; the wire forces are here led to each ring corner. On the other hand, the \"Akron\" and the English airship R 101 have so-called inherently stiff rings. These are built up in such a manner that two external ring members lying in the outer surface of the airship are joined with an inner ring member by means of wall struts to form a stable triangular girder. The question, which of the two main ring types is the better for the present size and form defined by the framing, can not be definitely decided. This is due to the two opposing functions of the main ring, on", "timestamp": "2026-07-19T18:14:11.285197+00:00"} | |
| {"citation_id": "19930091707", "source_url": "https://ntrs.nasa.gov/api/citations/19930091707/downloads/19930091707.pdf", "page_number": 15, "total_pages": 20, "image_filename": "19930091707_p15.jpg", "text": "CRINKLING STRENGTH AND BENDING STRENGTH OF ROUND AIRCRAFT TUBING 11\n\n$\\sigma_{cy}$\n\n$\\sigma_{cy} = \\frac{1.525}{\\frac{D}{t} + 1.4}$\n\n$\\Delta$ $1\" \\times .0125\"$ \n$\\circ$ $1 \\times .020$ \n$\\square$ $1 \\times .065$ \n$\\triangle$ $1\\frac{1}{2} \\times .0150$ \n$\\bigcirc$ $1\\frac{1}{2} \\times .022$ \n$\\nabla$ $1\\frac{1}{2} \\times .020$ \n$\\triangledown$ $1\\frac{1}{2} \\times .035$ \n$\\blacktriangle$ $1\\frac{1}{2} \\times .050$ \n$\\times$ $1\\frac{3}{8} \\times .025$ \n$\\blacktriangledown$ $2 \\times .0222$ \n$\\blacktriangleright$ $2 \\times .035$\n\n$\\frac{D}{t}$\n\nFIGURE 13.—Diagram of $\\sigma_{cy}$, $D/t$ for chromium-molybdenum steel (transverse loading).\n\n$f_c$, 1000 lb./sq. in.\n\n$f_c = 61,500 - 131 \\frac{D}{t}$\n\n$\\circ$ $1\" \\times .028\"$ \n$\\triangle$ $1 \\times .045$ \n$\\square$ $1 \\times .065$ \n$\\blacktriangle$ $1\\frac{1}{2} \\times .025$ \n$\\nabla$ $1\\frac{1}{2} \\times .032$ \n$\\blacktriangledown$ $1\\frac{1}{2} \\times .050$ \n$\\times$ $2 \\times .050$ \n$\\blacktriangleleft$ $2 \\times .025$ \n$\\blacktriangleright$ $2 \\times .032$ \n$\\blacktriangledown$ $2 \\times .035$ \n$\\blacktriangle$ $2 \\times .042$\n\n$\\frac{D}{t}$\n\nFIGURE 14.—Diagram of $f_c$, $D/t$ for duralumin (transverse loading).", "timestamp": "2026-07-19T18:14:21.664379+00:00"} | |
| {"citation_id": "19930094552", "source_url": "https://ntrs.nasa.gov/api/citations/19930094552/downloads/19930094552.pdf", "page_number": 16, "total_pages": 32, "image_filename": "19930094552_p16.jpg", "text": "14 N.A.C.A. Technical Memorandum No. 864\n\n(See fig. 19.) The resultant force increases in magnitude because the moment applied remains constant. Such an increase, however, is only possible if there exists at the intersection with the horizontal plane a shearing action which is oppositely directed to that along the main spar.\n\nIn the bending load case the shear along the cylinder varies continuously; the distribution in the two inner strips (A and B) in bay V is related to the variation of longitudinal stresses in the stiffeners between the main spars and the longitudinal neutral axis. With the arching loading condition and with bulkheads e and f riveted, sharp discontinuities of shear occur in the latter ring. In the strips along the extreme fibers, the shear at first increases somewhat from bulkhead e toward bulkhead a and then again diminishes slowly to zero. In contrast to the bending loading condition, the changes in the shear between two main spars are not considerable and the change due to the longitudinal stresses in the strips negligibly small. If bulkheads e and f are unattached, the shears rise steeply in bay V to a maximum value and then again decrease steadily toward bulkhead a.\n\nUnder the arching loading condition with attached bulkheads, buckling occurs first in bay III, strip A¹ (figs. 1 and 18) for a spar load P = 2630 kg. The corresponding shear stress is obtained as τ = 128 kg/cm². From the formula of H. Ebner (reference 3) the buckling strength of the sheet metal strip is obtained as 87.4 kg/cm² if the shear buckling strength is computed according to Donnell and 172 kg/cm² if it is determined by the formula of H. Wagner. The value computed from the test therefore lies in between. The fact that buckling first occurs in bay III and not in the somewhat more highly stressed bay IV may be explained by the effect of pre-buckling.\n\nThe check tests mentioned under IV, 3 were found to be in satisfactory agreement with the computed values of the shears, the deviations remaining within 10 percent of the applied spar forces.\n\nThe shears in the full skin portion in the peripheral direction generally vary little within a strip. An exception is formed by the strips which, in the bending case, lie in the neighborhood of the extreme fibers. Along the longitudinal direction, however, a continuous variation", "timestamp": "2026-07-19T18:14:24.493559+00:00"} | |
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