PDF/ocr_dataset_103.jsonl

{"citation_id": "19930094552", "source_url": "https://ntrs.nasa.gov/api/citations/19930094552/downloads/19930094552.pdf", "page_number": 26, "total_pages": 32, "image_filename": "19930094552_p26.jpg", "text": "N.A.C.A. Technical Memorandum No. 864\nFigs. 9,10,11\n\nStringer forces/applied force\n+1.0\n+0.8\n+0.6\n+0.4\n+0.2\n0\n-0.2\n-0.4\n-0.6\n-0.8\n-1.0\n\nTension side\nLongitudinal stiffener 12\n\nPressure side\nLongitudinal stiffener 16\n\nII B f f I/8 a e II/8 d II/8 c I/8 b I/8 a\n\n——— Prebuckling range\n- - - - - Buckling\n\nFigure 9.- Comparison of the main spar forces in the prebuckling and the buckling range for the bending loading condition.\n\nMean bulkhead stresses\n[Figure: Polar plot of mean bulkhead stresses with concentric circles and radial lines, labeled with values such as 0, 20, 40, 60, 80, 100, 120, 140, 160, 180, 200, 220, 240, 260, 280, 300, 320, 340, 360, and points labeled 1-1, 2-2, 3-3, 4-4, 5-5, 6-6, 7-7, 8-8, 9-9, 10-10, 11-11, 12-12, 13-13, 14-14, 15-15, 16-16, 17-17, 18-18]\n\nBulkhead\n———\n- - - - -\n· · · · ·\n\nFigure 10.- Bulkhead stress under bending load.\n\n(a) Cylinder cap over main spars under bending load\n(b) Half cylinder under arching forces P\n\nFigure 11.- Equilibrium of portions cut from the cylinder.", "timestamp": "2026-07-19T18:24:00.915436+00:00"}
{"citation_id": "19930091715", "source_url": "https://ntrs.nasa.gov/api/citations/19930091715/downloads/19930091715.pdf", "page_number": 30, "total_pages": 30, "image_filename": "19930091715_p30.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-19T18:24:07.667908+00:00"}
{"citation_id": "19930094533", "source_url": "https://ntrs.nasa.gov/api/citations/19930094533/downloads/19930094533.pdf", "page_number": 41, "total_pages": 51, "image_filename": "19930094533_p41.jpg", "text": "N.A.C.A. Technical Memorandum No. 883\nFigs. 17,18\n\n[Figure: Diagram of an airfoil with lines radiating from it, labeled with circles. A scale bar is shown labeled \"Scale\" and \"0.2°\".]\n\nFigure 17.- Temperature record at .7° incidence(model)\n\n[Figure: Diagram of an airfoil with lines radiating from it, labeled with circles. A scale bar is shown labeled \"Scale\" and \"0.2°\".]\n\nFigure 18.- Temperature record at 2.20° incidence(model).", "timestamp": "2026-07-19T18:24:08.174307+00:00"}
{"citation_id": "19930094559", "source_url": "https://ntrs.nasa.gov/api/citations/19930094559/downloads/19930094559.pdf", "page_number": 8, "total_pages": 16, "image_filename": "19930094559_p8.jpg", "text": "6 N.A.C.A. Technical Memorandum No. 857\n\nthrottle. If the injection, however, occurred as in figures 15 and 14, when there was a strong air movement from the main combustion space to the air-storage chamber, almost the entire fuel was scavenged out from the main combustion space into the air cells. The pictures show clearly that in this case the main combustion space itself takes a very small part in the combustion. The slight exhaust schlieren in the main combustion space indicate that no considerable fuel quantities are carried from the second cell back into the main combustion space, so that incomplete combustion results. With a fuel quantity corresponding approximately to full load (fig. 14), soot formation due to local air insufficiency takes place in the second air cell. In the case of the pictures obtained with still greater fuel quantities, this soot formation is so strong that the windows of the second air cell are completely darkened. The fifth picture of figure 14 shows the small window of the first air cell lit up by the flame passing through. The picture frequency was not sufficient, however, to indicate definitely whether ignition first occurred in the first or in the second air cell.\n\nWith later start of injection - that is, with smaller air velocities, there is an increase in the part taken by the main combustion chamber in the combustion process. This is clearly shown by figures 15 and 16, for which the fuel quantities are the same as those of the preceding pictures. A part of the fuel apparently accumulates ahead of the mouth of the first throttle. After ignition the escaping flame drags this fuel along and throws it against the nozzle from which position it spreads laterally and fills the entire main combustion chamber. Soot formation was not observed except for fuel quantities corresponding to an overload of the engine.\n\nTable I gives the numerical results for the two extreme values of the injection times. With later injection the ignition lag decreases somewhat corresponding to the higher air density. The observed ignition lags are greater than those in the engine on account of the lower general temperature level. In general, however - also in the case of the comparison engine - the end of injection occurs before the start of ignition up to about normal load, so that with the injection arrangement here used the jet escaping from the air cell cannot affect the nozzle jet.\n\nWhereas the combustion times in the main combustion space in the case of injection at high air velocity are", "timestamp": "2026-07-19T18:24:14.240921+00:00"}
{"citation_id": "19930091692", "source_url": "https://ntrs.nasa.gov/api/citations/19930091692/downloads/19930091692.pdf", "page_number": 13, "total_pages": 20, "image_filename": "19930091692_p13.jpg", "text": "AUTO-IGNITION AND COMBUSTION OF DIESEL FUEL IN A CONSTANT-VOLUME BOMB 9\n\n[Figure: Three pressure vs. time graphs labeled 168, 165, and 159. Vertical axis: PRESSURE, LB. PER SQ. IN. Horizontal axis: TIME, SEC.]\n\n| Record | Air-fuel ratio | Density (lb./cu. ft.) | Ignition lag (sec.) | Maximum explosion pressure (lb./sq. in.) |\n|--------|----------------|------------------------|---------------------|------------------------------------------|\n| 168    | 20             | 0.59 (air)             | 0.0018              | 800                                      |\n| 165    | 13.3           | 0.59 (air)             |                     |                                          |\n|        |                | 0.29 (nitrogen)        | .0017               | 1,220                                    |\n| 159 F  | 45             | 0.89 (air)             | .0015               | 880                                      |\n| 159    | 13.5           | 0.59 (air)             |                     |                                          |\n|        |                | 0.32 (products)        | .0020               | 1,100                                    |\n\nFIGURE 11.—Effect of added inert gases on permissible air-fuel ratio. Air temperature, 1,155° F.; air density, 0.59 pound per cubic foot.\n\nthe compression ratio are varied. The manner and extent in which these trends are altered as the ignition quality of the fuel is changed must await further tests.\n\nFACTORS AFFECTING THE IGNITION LAG\n\nThe ignition lag, as shown in figures 4 to 7, decreased as either the density or temperature of the gas was increased. The fact that the lags were considerably shorter for supposedly equivalent conditions than those obtained by Michailova and Neumann (reference 11) with a fuel of superior ignition quality (cetene) indicates that the range of temperatures in their bomb must have been rather wide. The air-fuel ratio appeared to have little or no influence upon the ignition lag, presumably because the optimum conditions for ignition always exist somewhere in the spray envelope. A consideration of the data in table 1 or a comparison of record 154 (fig. 6) with the first explosion on record 155 (fig. 10) or 159 (fig. 11) and of record 167 (fig. 4) with 168 (fig. 6) shows that the spread of the points in figure 5 must have been due to slight variations in temperature; certainly there is no correlation with air-fuel ratio. It follows from the results shown in figure 5 that the decrease in ignition lag, accompanying an increase in the compression ratio of a compression-ignition engine, is due partly to the increase in temperature and partly to the increase in air density. This conclusion confirms the results of previous engine tests, in which the intake-air temperature and pressure were independently varied (references 2, 22, and 23). The fact that some methods of rating fuels, which certainly do not simulate actual engine conditions (references 12, 13, and 16), correlate other rating methods and engine requirements reasonably well indicates that the curves in these figures would be merely shifted with very little change in shape if the ignition quality were varied. This contention is partly substantiated by the curves shown in reference 24.\n\nThe results of the present tests verify, in principal, the conclusion drawn by Michailova and Neumann (reference 11) that the ignition lag shows little tendency to decrease further at the higher temperatures and densities. (See figs. 5 and 7.) Engine tests indicated the same tendency (reference 24) but the necessary conditions varied somewhat with the fuel used. A determination in this laboratory of the air consumption and compression pressure of a motored engine having a known clearance volume indicated a gas temperature of 1,290° F. and a density of 0.87 pound per cubic foot at top center for a compression ratio of 14.6. Under operating conditions the temperature should exceed this value, owing to the heating of the air charge and to the presence of residual combustion products. Figure 7 indicates that this compression temperature might", "timestamp": "2026-07-19T18:24:21.404903+00:00"}
{"citation_id": "19930091716", "source_url": "https://ntrs.nasa.gov/api/citations/19930091716/downloads/19930091716.pdf", "page_number": 18, "total_pages": 24, "image_filename": "19930091716_p18.jpg", "text": "14 REPORT NO. 641—NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nscale, project horizontally along a line of constant $Q_e$ until an intersection is made with the previously drawn radial line of constant $Q_e$ (such as line A–B intersecting line O–C at B). Now the locus of the desired points is assumed to be line A–B–C rather than the radial line O–C used in the less-refined method.\n\nThe value of the static friction torque will probably lie somewhere between 20 percent and 60 percent of\n\n35° or 40°, which is not far above the normal operating range. In case the engine fails in such a manner that it locks or if it is desired to stop the rotation to prevent damage to the airplane or the engine, it will be necessary to feather the propeller to about 85° to 90° where it will stop turning and at the same time have a very low drag.\n\nUSE OF PROPELLER BRAKING EFFECT IN REDUCING DIVING SPEEDS\n\nThe rapid development in the aerodynamic cleanness of modern airplanes has resulted in a large increase in their terminal diving speeds; in fact, it is questionable whether some of them could resist the destructive forces to which such a dive would subject them. Most airplanes are not called upon to make such dives but in certain military maneuvers, such as dive bombing, the vertical or nearly vertical dive is a routine requirement.\n\nThe accuracy of dive bombing is adversely affected by the high diving speeds, and various methods of slowing up the dive have been considered. Some of these methods depend upon a split structural surface, such as a strut or wing flap, which opens up to produce an effective air brake.\n\nThe airplane is already equipped with a convenient and very rugged mechanism for producing a large positive or negative thrust. The controllable propeller, if set at low blade angles, will provide a very effective air\n\n<!-- Image (139, 198, 468, 517) -->\n\nFIGURE 15.—Calculated power required to overcome the drag of a dead engine, idling and locked propeller, at 135 m. p. h. on a transport airplane. Propeller 5868-16, 3 blades, 11-foot diameter.\n\nthe friction torque at rated engine speed depending upon the engine temperatures.\n\nIn this example a value of 34 percent is used which, at 135 miles per hour, gives a value of $Q_e = -0.000877$ and the projected line on the chart in figure 7 is the line A–B.\n\nThe rest of the calculations may conveniently be put in tabular form:\n\n| $\\theta$ (deg.) | $-Q_e$ | $nD/V$ | $-T_e$ | $-T_p$ (lb.) | t. hp. ($-T_p V_s/375$) |\n| :--- | :--- | :--- | :--- | :--- | :--- |\n| 10 | 0.00205 | 1.08 | 0.0772 | 748 | 270.0 |\n| 15 | .00187 | .987 | .0414 | 402 | 145.0 |\n| 20 | .00157 | .823 | .0250 | 242 | 87.0 |\n| 25 | .00132 | .692 | .0152 | 150 | 54.0 |\n| 30 | .00112 | .583 | .0107 | 104 | 37.5 |\n| 40 | .000877 | .420 | .0055 | 53 | 19.1 |\n| 50 | .000877 | .292 | .0037 | 36 | 13.0 |\n| 60 | .000877 | .192 | .0046 | 45 | 16.2 |\n| 70 | .000877 | .109 | .0036 | 35 | 12.6 |\n\nThe results are plotted in figure 15 along with those for the same propeller when locked $\\left(\\frac{nD}{V}=0\\right)$. For this particular example, it is seen that, through the greater part of the blade-angle range, the power required to overcome the drag of the locked propeller is considerably greater than for the windmilling condition. It is also seen that most of the benefit gained from increasing the blade angle on the windmilling propeller is obtained at\n\n<!-- Image (535, 496, 864, 762) -->\n\nFIGURE 16.—Calculated values of terminal velocity and engine speed for a modern pursuit airplane. Vertical descent; 8,000-foot altitude; 3-blade propeller.\n\nbrake, as is shown by the curves in figure 16. The curves were obtained from the test data by a method that will shortly be explained. They represent the terminal velocity and engine speed for the modern pursuit airplane, shown in figure 17, for various blade-angle settings of its 3-blade controllable propeller. The effects of compressibility on both airplane and propeller have been neglected. In the important part of the curves (low blade angles), these effects are small.", "timestamp": "2026-07-19T18:24:32.882339+00:00"}
{"citation_id": "19930094549", "source_url": "https://ntrs.nasa.gov/api/citations/19930094549/downloads/19930094549.pdf", "page_number": 11, "total_pages": 76, "image_filename": "19930094549_p11.jpg", "text": "N.A.C.A. Technical Memorandum No. 367 9\n\nThe damping of the pitching oscillations will be characterized by a coefficient which we shall write in the form $\\frac{dC'_z}{di'}$ as if the entire damping were produced by the action of the horizontal tail surface. In the above expression $C'_z$ would be the lift coefficient of the tail, $i'$ the real angle of attack of the tail.\n\nAs a matter of fact, we shall take a numerical value somewhat larger than that corresponding to the derivative of the lift of the tail in order to take account of the damping introduced by the parts of the airplane other than the horizontal tail surface. We shall take:\n\n$$\n\\frac{dC'_z}{di'} = 0.065\n$$\n\nFinally, we shall take\n\n$$\n\\frac{dC_M}{di} = \\mu\n$$\n\nthe coefficient of static stability, as the variable element in the example.\n\nAs the variations of the static stability may be obtained by a simple displacement of the center of gravity without varying the tail surface, we may assume it to be possible to vary $\\mu$ without varying the other characteristics.\n\n3. Element characterizing the engine-propeller unit.\n\nWe write:\n\n$$\n\\frac{dT}{dV} = -h \\frac{T}{V}\n$$\n\ngiving $h$ the value 0.5.\n\n4. Element characterizing the weight of the airplane.\n\nThe weight per square meter will be taken as equal to 40 kilograms. From these data the result follows immediately that the speed of the airplane under the conditions considered is 40 meters per second (90 miles per hour).", "timestamp": "2026-07-19T18:24:47.580128+00:00"}
{"citation_id": "19930094542", "source_url": "https://ntrs.nasa.gov/api/citations/19930094542/downloads/19930094542.pdf", "page_number": 70, "total_pages": 102, "image_filename": "19930094542_p70.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-19T18:24:47.976991+00:00"}
{"citation_id": "19930094543", "source_url": "https://ntrs.nasa.gov/api/citations/19930094543/downloads/19930094543.pdf", "page_number": 45, "total_pages": 50, "image_filename": "19930094543_p45.jpg", "text": "N.A.C.A. Technical Memorandum No. 873\nFigs. 13,14\n\n<!-- Image (269, 143, 739, 377) -->\n\nFigure 13. - Effect of induced air humidity\n(anti-turbulent head) N = 1,250 r.p.m.\n\n<!-- Image (269, 436, 695, 869) -->\n\nFigure 14. - Effect of richness of mixture\n(anti-turbulent head) N = 1,250 r.p.m.", "timestamp": "2026-07-19T18:24:54.247549+00:00"}
{"citation_id": "19930094552", "source_url": "https://ntrs.nasa.gov/api/citations/19930094552/downloads/19930094552.pdf", "page_number": 27, "total_pages": 32, "image_filename": "19930094552_p27.jpg", "text": "N.A.C.A. Technical Memorandum No. 864\nFigs. 13,14,15\n\n[Figure: Graph showing Stringer forces per cm² of cross section vs. cylinder length. Y-axis ranges from -600 to +600 kg/cm². X-axis labels: II B, f, I B, α c, II' B, d, III B, c, II β, b, I D, a. Curves labeled: Tension side, Main spar, Pressure side, Main spar. Legend for Longitudinal stiffeners: 10 u. 15, 11 u. 17, 12 u. 16, 13 u. 15.]\n\nFigure 13.- Variation of stringer forces over the cylinder length under the arching loading condition.\n\n[Figure: Diagrams showing Normal stress distribution over the cross section height. Left: Circular cross section with stress profiles. Center: Rectangular cross section with labels I B, II B, III B, IV B = IV f, D B. Right: Stress distribution graph with Y-axis labels 10, 11, 12, 13, 14, 15, 16, 17, 18 and X-axis from -800 to +400 kg/cm². Legends: I B, II B, III B; V α, V β, V f.]\n\nFigure 14.- Normal stress distribution over the cross section height under the arching loading condition.\n\n[Figure: Two polar plots labeled \"Mean bulkhead stresses\" and \"Bulkhead\". Left plot: \"With ring at free end\". Right plot: \"Without ring at free end\". Both show stress contours in kg/cm² with radial lines labeled (e.g., 1-2, 2-3, ..., 18-1). Center label: \"+ Test points for bulkhead a\".]\n\nFigure 15.- Bulkhead ring stresses under the arching loading condition.", "timestamp": "2026-07-19T18:24:54.443669+00:00"}
{"citation_id": "19930091692", "source_url": "https://ntrs.nasa.gov/api/citations/19930091692/downloads/19930091692.pdf", "page_number": 14, "total_pages": 20, "image_filename": "19930091692_p14.jpg", "text": "```markdown\n10 REPORT NO. 617—NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nbe considerably reduced by decreasing the intake-air temperature, provided that the same air density is maintained, without exceeding an ignition lag of 0.0015 second. This indication is in agreement with engine results (reference 22). Such a lag and its associated rate of pressure rise have been found in this laboratory to be permissible in test engines.\n\nINFLUENCE OF IGNITION LAG ON COMBUSTION\n\nThe solid curves in figure 8 show qualitatively, for two air densities, the influence of the ignition lag on the effectiveness of the combustion within a reasonable period after ignition. Rothrock and Waldron (reference 1) have observed a corresponding decrease in engine efficiency as the ignition lag was decreased below the value giving the greatest permissible rate of pressure rise. The same trend may be seen in figures 4 and 6; the shorter the lag, the greater the ensuing combustion period or the time necessary to attain maximum pressure. This tendency accounts for the approach of the upper solid curve to the dashed curve in figure 8 at long ignition lags since the heat losses were a minimum for this condition. The lower solid curve corresponds to a density approximating that in a normally aspirated compression-ignition engine during injection.\n\nThe trends in this combustion effectiveness should be indicative of similar trends in engine mean effective pressure, provided that all but a negligible portion of the combustion occurs soon after top center. Under these conditions the mean effective pressure is a function of, but not directly proportional to, the ratio of explosion to compression pressure. The fact that this ratio becomes a maximum in an engine only when appreciable combustion occurs before top center necessarily tends to nullify the advantage, in terms of mean effective pressure, accruing from a high value of this ratio; hence an intermediate ratio gives the best results in actual practice. Moreover, the expansion of the gases in an engine prevents the attainment of as high a value of maximum explosion pressure to maximum compression pressure as would occur if the same degree of combustion could be realized at top center or in a bomb. Actual engine ratios appear to be in the neighborhood of 1.5 to 1.8 (references 23 and 25) when the maximum cylinder pressures are limited to moderate values. The ratio can be increased to some extent, of course, by permitting higher cylinder pressures. The present values of the pressure ratio after 0.004 second are of this order even for the shortest ignition lags but are not strictly comparable with the engine ratios because of different conditions and air-fuel ratios. Within limits, however, these trends should be apparent in either an engine or a bomb.\n\nTo what extent the curves in figure 8 are representative of an engine possessing considerable turbulence is unknown. A qualitative comparison can be made, however, with available combustion efficiency data for a quiescent combustion-chamber engine. Thus, the ratio of a point on the upper solid curve to a corresponding point (same ignition lag) on the dashed curve is approximately proportional to the ratio of the energy derived from burned fuel in the 0.004-second interval to the total available energy, that is, to the combustion efficiency for the particular conditions. If attention is confined to lags of 0.001 and 0.0015 second, the respective ratios for the bomb are 0.55 and 0.72, which compare favorably with efficiencies ranging from 59 to 69 percent for the total combustion in a quiescent combustion-chamber engine (reference 26). The heat losses are necessarily indicated as unburned fuel in both instances. This agreement indicates that the combustion in the bomb, even for the short lags, was comparable with that in this particular engine.\n\nThe solid curves of figure 8 show that the 0.004-second pressure ratio increased as the density was increased, particularly for ignition lags acceptable in an engine. Since this ratio should be independent of density for a given air-fuel ratio, negligible heat losses, and the same percentage of fuel burned, it follows that a greater percentage of fuel burned in the designated period at the higher density. Compression-ignition engines in this laboratory have not shown a similar trend as evidenced by a constant indicated specific fuel consumption for a given air-fuel ratio and all boost pressures. (See fig. 7, reference 22.) It is possible that, in the bomb, the combination of the higher air density and larger fuel quantity merely resulted in better mixing without any chemical effect; whereas, in the engine, this effect would be minimized by air movement. On the other hand, the data presented in reference 27 for hydrogen, carbon monoxide, and methane-air mixtures indicate that the gas density does affect the burning of these homogeneous mixtures but what the effect should be for the higher hydrocarbons is unknown. Furthermore, some reduction in indicated specific fuel consumption with increasing boost has been reported for spark-ignition engines of low compression ratio (reference 28). Whether all of this reduction can be attributed to a decrease in the percentage of residuals is not known. At higher compression ratios, for which the necessary range of ignition advance angle was greater, there was first a reduction in fuel consumption and then a continuous increase with increasing intake-air density. This fact indicates that, for the high compression ratios, other factors more than offset the improvement in burning that might have been expected on the basis of the low-compression-ratio results.\n\nEFFECT OF COMBUSTION PRODUCTS ON COMBUSTION\n\nIt has been customary in most discussions of combustion in compression-ignition engines to attribute the slow burning in the latter part of the power stroke to poor mixing of the fuel and air. This assumption is\n```", "timestamp": "2026-07-19T18:25:01.578808+00:00"}
{"citation_id": "19930091707", "source_url": "https://ntrs.nasa.gov/api/citations/19930091707/downloads/19930091707.pdf", "page_number": 18, "total_pages": 20, "image_filename": "19930091707_p18.jpg", "text": "```markdown\n14\nREPORT NO. 632—NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nIf the stresses from equations (12) and (13) are equated to the highest average column stresses (ratio of slenderness equal to zero), the highest values of $d/t$ may be found for which the column curves for the two materials apply over their entire range without cutting off at the top.¹ For chromium-molybdenum-steel tubes, the limiting column stress was found in reference 4 to be 79,400 lb. per sq. in., and, if one equates this quantity to the right-hand side of equation (12),\n\n$$79\\ 400 = \\frac{19\\ 800\\ 000}{\\frac{d}{t} + 200}$$\n\nand solves for $d/t$, one obtains $d/t=50$. Similarly, for duralumin tubes the limiting column stress was found in reference 4 to be 42,700 lb. per sq. in., and if one equates this quantity to the right-hand side of equation (13),\n\n$$42\\ 700 = 56\\ 400 \\left[ 1 - \\frac{\\frac{d}{t} - 1}{150 + \\frac{4}{3} \\left( \\frac{d}{t} - 1 \\right)} \\right]$$\n\nand solves for $d/t$, one obtains $d/t=55$. The column formulas given in reference 4 may therefore be used over their entire respective ranges for tubing for which $d/t$ does not exceed 50 in the case of chromium-molybdenum steel and $d/t$ does not exceed 55 in the case of duralumin.\n\nThe question may arise as to whether clamps used in practice for transferring transverse loads to tubing are sufficiently effective in holding the tube round to prevent failure at the clamp. Preliminary bending tests made with several types of clamps, both with clamps furnished by manufacturers and more flexible clamps made for the purpose, indicated that, so long as the transverse load was applied through a tension member, the clamp would not weaken the tube. The type of connection most likely to weaken a tube locally is a\n\nweld. A relatively small compression member transferring its load to the tube in question through a weld might easily promote a dent and cause local failure at lower stresses than those given by equation (14).\n\nIt may not be out of place here to call attention to possible failure by transverse shear. The transverse shearing strength of tubing has not been studied in the present investigation. All that can be said is that no evidence of failure due to transverse shear was observed in any specimen. Hansen (reference 2) found, for much thinner tubes, that the modulus of rupture was unaffected by shear when the cantilevered end of the tube was as short as three diameters.\n\nNATIONAL BUREAU OF STANDARDS,\nWASHINGTON, D. C., March 28, 1938.\n\nTABLE I\nNOMINAL CROSS-SECTIONAL PROPERTIES OF TUBES\n\n| Diameter $d$ (in.) | Thickness $t$ (in.) | Ratio $d/t$ | Area $A$ (sq. in.) | Section modulus $I/c$ (in.³) |\n| :--- | :--- | :--- | :--- | :--- |\n| **CHROMIUM-MOLYBDENUM STEEL** | | | | |\n| 1 | 0.0125 | 80.0 | 0.0388 | 0.00887 |\n| 1 | .028 | 35.7 | .0855 | .02021 |\n| 1 | .065 | 15.4 | .1900 | .04193 |\n| 1½ | .0120 | 100.0 | .0700 | .02572 |\n| 1½ | .022 | 68.2 | .1022 | .03720 |\n| 1½ | .028 | 53.6 | .1295 | .04678 |\n| 1½ | .035 | 42.9 | .1611 | .05765 |\n| 1½ | .049 | 25.9 | .2028 | .09121 |\n| 1⅞ | .035 | 46.4 | .1748 | .09803 |\n| 2 | .0222 | 90.0 | .1379 | .08746 |\n| 2 | .035 | 57.1 | .2161 | .10432 |\n| **DURALUMIN** | | | | |\n| 1 | 0.028 | 35.7 | 0.0855 | 0.02021 |\n| 1 | .049 | 20.4 | .1464 | .03319 |\n| 1 | .065 | 15.4 | .1900 | .04193 |\n| 1½ | .025 | 60.0 | .1158 | .04202 |\n| 1½ | .032 | 46.9 | .1476 | .05303 |\n| 1½ | .058 | 25.9 | .2528 | .09121 |\n| 1½ | .109 | 13.8 | .4785 | .1651 |\n| 2 | .020 | 100.0 | .1244 | .06097 |\n| 2 | .025 | 80.0 | .1555 | .07594 |\n| 2 | .032 | 62.5 | .1978 | .09581 |\n| 2 | .035 | 57.1 | .2161 | .10432 |\n| 2 | .042 | 47.6 | .2584 | .12386 |\n\n¹ The use of equations (12) and (13) assumes that the column curves are based on the same compressive yield strength as the crinkling curves. If the compressive yield strength of the column material is higher, as it was for the duralumin tubes of reference 4, the values of $d/t$ obtained will be on the conservative side.\n\n* No bending tests were made of this size.\n```", "timestamp": "2026-07-19T18:25:02.871780+00:00"}
{"citation_id": "19930094535", "source_url": "https://ntrs.nasa.gov/api/citations/19930094535/downloads/19930094535.pdf", "page_number": 11, "total_pages": 17, "image_filename": "19930094535_p11.jpg", "text": "10 N.A.C.A. Technical Memorandum No. 881\n\ngrams. The glasses of 4.5- to 5.5-millimeter thickness are tested by the DVL from a falling height of 1.5 meters, and those of over 5.5- to 6.5-millimeter thickness, from a height of 2 meters. The falling ball is magnetically released and must be allowed to fall perpendicularly over the center of the free surface. The ball may also be released by hand, in which case a suitable guide tube must be used. Before the test proper, the specimens are accurately weighed to 0.01 gram, brought to the test temperature, and set horizontally in the mounting. The room temperature during the test is $20^\\circ \\pm 2^\\circ$. The test temperatures selected are $-21^\\circ$, $-11^\\circ$, $0^\\circ$, $+20^\\circ$, and $+40^\\circ$. For this purpose the specimens are placed for at least 15 minutes in the following baths:\n\n- $-21^\\circ$ mixture of ice and cooking salt.\n- $-11^\\circ$ brine.\n- $0^\\circ$ ice-water mixture.\n- $+20^\\circ$ air bath.\n- $+40^\\circ$ water bath.\n\nAfter the specimens have been removed from the baths they are dried somewhat and placed in the test apparatus. The falling body is released after a definite time, depending in each case on the glass thickness and the test temperature. If the specimen is not penetrated the pieces, which in the case of the multilayer safety glasses are thrown off in the direction of the impact, are weighed as is also the pane freed from all the loose pieces, and the percent weight lost in the fragments is determined. At the same time the shapes of the pieces are noted. In case the specimen is penetrated, it is noted in what manner the bending layer has been ruptured and how many fragments result. The appearance of multiple-layer glass after the ball-falling test at $-21^\\circ$, $-11^\\circ$, and $+20^\\circ$ may be seen from figure 9. Figure 10 shows how a multiple-layer safety glass is fractured in the test at $-21^\\circ$. Similarly, large fragments occur in the fracture of artificial glasses in the ball-falling test, but in this case the edges have no cutting action in contrast to the very sharp cutting edges of silicate glass.\n\nFor the weathering-resistance tests, the specimens are placed outdoors, inclined at $45^\\circ$ to the horizontal", "timestamp": "2026-07-19T18:25:03.075605+00:00"}
{"citation_id": "19930094544", "source_url": "https://ntrs.nasa.gov/api/citations/19930094544/downloads/19930094544.pdf", "page_number": 22, "total_pages": 43, "image_filename": "19930094544_p22.jpg", "text": "20 N.A.C.A. Technical Memorandum No. 872\n\nIt is the task of the constructor to strive from the beginning for balanced distribution of the weights and lifting forces through suitable weight distribution plans. This, however, is possible only to a limited degree, so that especially after a rather large fuel consumption and in the very rare case of the deflation of a cell, the static loads can cause rather large shear forces and bending moments.\n\nThe forces of the second kind are the aerodynamic or air forces. They represent the most important group of external forces. Their determination is accomplished through pressure measurements in the wind tunnel (references 27 and 28) as well as through tests on the airship in flight (reference 29). Their theoretical determination is possible through the procedures worked out by Fuhrmann, Von Karman, and Munk (references 30, 31, and 32, respectively), the results of which in general show good agreement with the test results. The aerodynamic forces occur chiefly in trimmed flight, i.e., when the heavy or light airship flies with an upward or downward directed longitudinal axis for equalization of the static forces. Similar forces occur in curved flight. Further, the forces acting on the stern of the airship with rudder movement belong to the aerodynamic forces, and finally also the forces exerted by gusts.\n\nAs a result of the accelerations occasioned by the air forces, the third kind of forces occurs: the so-called inertia forces. They are equated to the external air forces and moments in accordance with the d'Alembert principle and depend upon the mass and the moment of inertia of the airship.\n\nIn German airship construction it is customary to select a limited number of conditions of loading. Principally, there are the case of the airship flying in the vertical plane at a fixed limiting altitude, that flying in the horizontal plane with the smallest turning circle, as well as the case of the rudder hard-over at a fixed rudder angle. More recently there has been added the consideration of the stressing due to gusts, which attack the forward part of the airship with a velocity of more than 10 m/s, as well as the forces on the airship lying at the mooring mast. The loading conditions mentioned are investigated individually and in certain combinations together with the constant static loads. In the calculation of the \"Akron\" all aerodynamic loading conditions are combined in a single loading condition, the effect of which is assumed in all longitudinal", "timestamp": "2026-07-19T18:25:08.298212+00:00"}
{"citation_id": "19930094551", "source_url": "https://ntrs.nasa.gov/api/citations/19930094551/downloads/19930094551.pdf", "page_number": 9, "total_pages": 18, "image_filename": "19930094551_p9.jpg", "text": "8 N.A.C.A. Technical Memorandum No. 855\n\n$\\alpha_{\\mathrm{s}}$, angle of incidence of a suspension-hose element to the flow (deg.)\n\n$A_{\\mathrm{s}}$, lift of a suspension-hose element (kg)\n\n$W_{\\mathrm{s}}$, drag of a suspension-hose element (kg)\n\n$G$, weight of air-speed head (kg)\n\n$W$, drag of air-speed head (kg)\n\nFigures 6 and 7 depict the lift and drag for metal-wound suspension hose of 7 and 8 mm diameter, respectively. The angle of incidence of the suspension hose ranges between 15 and 75 under practical conditions. A good approximation is:\n\n$$\nc_{\\mathrm{w}} = C \\sin^{2} \\alpha_{\\mathrm{s}}\n$$\n\n(4)\n\n$$\nc_{\\mathrm{a}} = C_{1} \\sin^{2} \\alpha_{\\mathrm{s}} \\cos \\alpha_{\\mathrm{s}}\n$$\n\n(5)\n\nIt involves empirically defined functions whose validity for different tube diameters was confirmed in wind-tunnel tests of the DVL. On these premises, the average error remains considerably below 10 percent in contrast to the 30 percent by Glauert's method.\n\nWith (4) and (5), the lift and drag of a chord element are:\n\n$$\n\\left.\n\\begin{array}{l}\nW_{\\mathrm{s}} = q \\Delta F C \\sin^{2} \\alpha_{\\mathrm{s}} \\\\\nW_{\\mathrm{s}} = C_{2} \\sin^{2} \\alpha_{\\mathrm{s}}\n\\end{array}\n\\right\\}\n$$\n\n(6)\n\n$$\n\\left.\n\\begin{array}{l}\nA_{\\mathrm{s}} = q \\Delta F C_{1} \\sin^{2} \\alpha_{\\mathrm{s}} \\cos \\alpha \\\\\nA_{\\mathrm{s}} = C_{3} \\sin^{2} \\alpha_{\\mathrm{s}} \\cos \\alpha_{\\mathrm{s}}\n\\end{array}\n\\right\\}\n$$\n\n(7)\n\nwhence figure 8, in conjunction with the equilibrium conditions for a hose element, gives:\n\n$$\n\\Sigma V = 0 : A_{n} + C_{3} \\sin^{2} \\alpha_{\\mathrm{s}} \\cos \\alpha_{\\mathrm{s}} - G_{\\mathrm{s}} - A_{n-1} = 0\n$$\n\n(8)\n\n$$\n\\Sigma H = 0 : B_{n} - B_{n-1} - C_{2} \\sin^{2} \\alpha_{\\mathrm{s}} = 0\n$$\n\n(9)", "timestamp": "2026-07-19T18:25:17.592187+00:00"}
{"citation_id": "19930094542", "source_url": "https://ntrs.nasa.gov/api/citations/19930094542/downloads/19930094542.pdf", "page_number": 71, "total_pages": 102, "image_filename": "19930094542_p71.jpg", "text": "N.A.C.A. Technical Memorandum No. 874\n\nFigs.49,50,51,52,53\n\nFigure 49.-Dimensioning of propeller  \n(American notations $S_1 F_2$  \n$A_3 P_1 H/D = 1.0$).\n\nFigure 50.-Characteristics of freely rotating propeller.\n\nFigure 51. $\\kappa = 9^\\circ$.\n\nFigure 52. $\\kappa = 4^\\circ$.\n\nFigure 53. $\\kappa = -1^\\circ$.\n\nFigure 51,52,53.-Effect of propeller slipstream on lift coefficient.", "timestamp": "2026-07-19T18:25:23.325581+00:00"}
{"citation_id": "19930091655", "source_url": "https://ntrs.nasa.gov/api/citations/19930091655/downloads/19930091655.pdf", "page_number": 6, "total_pages": 22, "image_filename": "19930091655_p6.jpg", "text": "2\nREPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nvastly different, however, for the spray considered as a unit, particularly for injection into an engine having induced air flow. Radiation to the surfaces of the droplets takes place to some extent and, whereas the actual magnitude of this exchange is uncertain, the maximum rate for energy transferred in this manner can be estimated for comparative purposes.\n\nFor conductive heat transfer there are two controlling resistances: The first is within the droplet itself and the second must be associated with the equivalent of a film surrounding the droplet. The first can be analytically treated but the second presents difficulties. Because of the net transfer of molecules from the droplet surface, the character of this film is not independent of time, as is usually assumed in theoretical treatments. (See references 9, 13, 14, and 15.) In fuel sprays the vapor films about the liquid droplets probably interpenetrate, thus necessitating some consideration of the spray as a whole. Moreover, the fuel is not uniformly distributed within the spray. In view of these difficulties no attempt will be made to establish any mathematical relations for the heat transfer through the film encompassing the droplet.\n\n**Heat transfer within a droplet.**—Ingersoll and Zobel have published equations pertaining to the internal heating of a rigid sphere suddenly inserted into a fluid possessing a higher temperature (reference 16). These equations may be modified to give, respectively, the instantaneous center temperature $t_c$ and the average temperature $t_a$ of the droplet:\n\n$$t_c = t_i + (t_s - t_i) \\left[ 1 - 2 \\left( e^{-\\frac{\\pi^2 h^2 t}{R^2}} - e^{-\\frac{4\\pi^2 h^2 t}{R^2}} + e^{-\\frac{9\\pi^2 h^2 t}{R^2}} - \\dots \\right) \\right]$$\n\n$$t_a = t_i + (t_s - t_i) \\left[ 1 - \\frac{6}{\\pi^2} \\left( e^{-\\frac{\\pi^2 h^2 t}{R^2}} + \\frac{1}{4} e^{-\\frac{4\\pi^2 h^2 t}{R^2}} + \\frac{1}{9} e^{-\\frac{9\\pi^2 h^2 t}{R^2}} + \\dots \\right) \\right]$$\n\nwhere $t_s$ is the temperature of the shell, $^\\circ$ C.\n$t_i$, the initial droplet temperature, $^\\circ$ C.\n$R$, the radius of the shell, centimeters.\n$t$, the immersion time, seconds.\n$h^2$, the thermal diffusivity of the liquid in the droplet.\n\nAs applied to liquid droplets these relations do not represent the effects of possible internal convection currents. These currents, if present, would increase the rate of temperature rise as determined by these relations.\n\nIt follows from these relations that when $h^2$ is considered constant, the increase in both the center and average temperature above the initial droplet temperature is a definite fraction of the difference $t_s - t_i$; thus $t_c - t_i = \\alpha (t_s - t_i)$ and $t_a - t_i = \\beta (t_s - t_i)$. The assumption of a constant value of $h^2$, independent of temperature in the range employed, appears justified for the purpose of qualitative comparisons in view of the uncertainty involved in its estimation. An average value for the range 49$^\\circ$ C. to 350$^\\circ$ C. can be estimated on the basis of the average values of the thermal conductivity (0.00027 calorie per second per centimeter per degree C., reference 17), the density (0.713 gram per cubic centimeter, reference 18), and the specific heat (0.662 calorie per gram per degree C., reference 19). These values result in $h^2 = 0.000572$ square centimeter per second.\n\nValues of $\\alpha$ and $\\beta$ are given in the following table for several immersion intervals and droplet radii, the largest radius corresponding to the initial average size (reference 20). No actual values of droplet temperatures are given inasmuch as there is no adequate basis on which to estimate their surface temperature. Incidentally, the foregoing relations tacitly assume that the surface temperature is instantaneously attained and thereafter remains constant. This assumption does not greatly invalidate the fact that the increase in the temperature of the droplet, particularly the average temperature as shown by the $\\beta$ values, attains a large fraction of the possible increase in a remarkably short time. Moreover, the smaller the radius the more quickly this fraction approaches unity. As a result of evaporation, the surface temperature does not attain so high a value as it would if all the heat reaching the drop served to heat it. Even so, such evaporation presumably does not alter the establishment of thermal equilibrium within the droplet and hence the $\\alpha$ and $\\beta$ quantities still have significance.\n\n**HEATING RATES FOR IMMERSED SPHERES**\n$[t_c - t_i = \\alpha(t_s - t_i); t_a - t_i = \\beta(t_s - t_i)]$\n\n| | Droplet (cm. radius) in... | 0.0015 .00059 | 0.0020 .00079 | 0.0025 .00098 |\n| :--- | :--- | :--- | :--- | :--- |\n| **Immersion time** | | $\\alpha$ $\\beta$ | $\\alpha$ $\\beta$ | $\\alpha$ $\\beta$ |\n| Second | | | | |\n| 0.0001 | | 0.002 0.464 | 0.000 0.362 | 0.000 0.296 |\n| .0005 | | .154 .706 | .018 .573 | .004 .479 |\n| .0005 | | .444 .826 | .128 .690 | .024 .587 |\n| .0007 | | .658 .865 | .256 .771 | .092 .666 |\n| .0010 | | .838 .903 | .420 .855 | .184 .749 |\n| .0015 | | .874 .932 | .710 .927 | .494 .842 |\n| .0020 | | .986 .956 | .882 .968 | .674 .900 |\n\n**Radiation from bomb wall.**—Some insight into the possible contribution of radiation from the bomb wall to the total heat transfer follows from a consideration of the maximum rate of radiation. If the very questionable assumption is made that the droplets are true black bodies suspended in a space filled with black-body radiation, the net energy transferred in calories per second is given by:\n\n$$\\Delta H = 1.37 \\times 10^{-8} S \\left[ \\left( \\frac{T_1}{100} \\right)^4 - \\left( \\frac{T_2}{100} \\right)^4 \\right]$$\n\n(reference 21) where\n$S$ is the surface area of the drop.\n$T_1$, the bomb-wall temperature, degrees $K$.\n$T_2$, the droplet temperature, degrees $K$.", "timestamp": "2026-07-19T18:25:24.997133+00:00"}
{"citation_id": "19930091716", "source_url": "https://ntrs.nasa.gov/api/citations/19930091716/downloads/19930091716.pdf", "page_number": 19, "total_pages": 24, "image_filename": "19930091716_p19.jpg", "text": "NEGATIVE THRUST AND TORQUE OF SEVERAL FULL-SCALE PROPELLERS 15\n\nThis airplane with its propeller set at $35^\\circ$ has a terminal velocity of 565 miles per hour; whereas at $2^\\circ$ its velocity has dropped to 277 miles per hour, or to 49 percent of its value at $35^\\circ$. It will be noticed that the engine speed rises to excessive values at blade angles around $15^\\circ$. These destructive engine speeds can be avoided by setting the blade angle somewhere between $5^\\circ$ and $0^\\circ$ before the dive is started, indicating that a quick-acting pitch-control mechanism would be advantageous.\n\nIf still more braking effect is required, the propeller may be set to negative angles and engine power applied.\n\n**Method of calculating $V_t$.**—Although the method of calculating terminal velocity still remains a cut-and-try process, it is made considerably easier by the use of the relation $\\frac{Q_t}{N_t} = K$ and the coefficient $Q_a$. The basic formula for a vertical dive is as follows:\n\n$$V_t = \\sqrt{\\frac{W}{\\frac{\\rho}{2} A + K T_t}}$$\n\nwhere  \n$V_t$, terminal velocity, f. p. s.  \n$W$, weight of airplane, lb.  \n$K = \\rho D^2$  \n$A = \\left( \\frac{\\text{parasite drag}}{\\frac{1}{2} \\rho V^2} \\right)$ equivalent parasite area.\n\nFrom this formula, $V_t$ may be calculated as follows:  \n1. Knowing $Q_t/N_t$ (from fig. 14), estimate a value of $V_t$ and calculate $Q_a$.  \n2. From suitable charts (figs. 3 to 9) obtain a value of $T_t$ for the desired blade angle.  \n3. Substitute $T_t$ in terminal-velocity formula and obtain $V_t$ calculated.  \n4. If $V_t$ estimated does not equal $V_t$ calculated, make a new estimate and repeat.\n\nWith a little experience two trials should be sufficient. The value of engine speed is obtained in the usual way from the value of $nD/V$ (such as point D', fig. 7).\n\n**Stability in dive.**—The destabilizing effect of a braking propeller is frequently brought up as an argument against the use of the propeller as a brake in reducing the terminal velocity of airplanes. The question will be briefly considered here in relation to the directional stability of the airplane. The stability in pitch presents a similar problem that may become critical if the center of gravity is displaced far from the thrust axis.\n\nIn a vertical dive with braking propeller, the negative thrust of the propeller will normally act upward in the vertical plane through the center of gravity. If a small displacement in yaw occurs, as shown in figure 17, the thrust and gravity forces will produce an upsetting couple that must be balanced by a lateral aerodynamic force on fin and fuselage. The situation is aggravated by the loss of energy in the slipstream, which usually passes over the tail surfaces.\n\nAn examination of forces and moments acting on the airplane, the diving characteristics of which are shown in figure 16, will be given as an example.\n\nThe upsetting-moment slope for a small displacement is given by the relation\n\n$$\\left( \\frac{dN}{d\\psi} \\right)_u = \\frac{T_t r_1}{57.3}$$\n\nwhere  \n$N$, yawing moment, ft.-lb.  \n$\\psi$, angle of yaw, deg.  \n$r_1$, distance from center of gravity to propeller disk, 6.5 ft.\n\nFor the airplane operating with a blade angle of $5^\\circ$,\n\n[Figure: Airplane in vertical dive showing unstable effect caused by propeller when used as a brake. Span, 35 feet; weight, 4,500 pounds; propeller diameter, 10 feet; 3 blades.]\n\n$T_t = 2 T_c D^2 q_0$, where $T_c$ is $-0.106$, $D$ is 10 feet, and $q_0$ is the dynamic pressure.\n\nTherefore\n\n$$\\left( \\frac{dN}{d\\psi} \\right)_u = \\frac{2 \\times -0.106 \\times 10^2 \\times 6.5}{57.3} q_0 = -2.4 q_0$$\n\nThe normal stabilizing yawing moment for this airplane may be taken from Diehl (reference 7) who gives as a reasonable value for directional stability:\n\n$$\\left( \\frac{dN}{d\\psi} \\right)_r = 0.00005 \\times W b q_1$$\n\nwhere  \n$W$ is the weight.  \n$b$, the span.  \nand $q_1$, the dynamic pressure.\n\nFor this example\n\n$$\\left( \\frac{dN}{d\\psi} \\right)_r = 7.9 q_1$$", "timestamp": "2026-07-19T18:25:32.002371+00:00"}
{"citation_id": "19930094542", "source_url": "https://ntrs.nasa.gov/api/citations/19930094542/downloads/19930094542.pdf", "page_number": 72, "total_pages": 102, "image_filename": "19930094542_p72.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-19T18:25:49.647622+00:00"}
{"citation_id": "19930094533", "source_url": "https://ntrs.nasa.gov/api/citations/19930094533/downloads/19930094533.pdf", "page_number": 42, "total_pages": 51, "image_filename": "19930094533_p42.jpg", "text": "N.A.C.A. Technical Memorandum No. 863\nFigs. 19,20\n\n[Figure: Diagram of an airfoil model with temperature recording points at 5° incidence. A scale bar labeled \"Scale 0,2°\" is shown.]\n\nFigure 19.- Temperature record at 5° incidence. (model)\n\n[Figure: Diagram of an airfoil model with temperature recording points at 7,80° incidence. A scale bar labeled \"Scale 0,2°\" is shown.]\n\nFigure 20.- Temperature record at 7,80° incidence. (model)", "timestamp": "2026-07-19T18:25:50.569372+00:00"}
{"citation_id": "19930094559", "source_url": "https://ntrs.nasa.gov/api/citations/19930094559/downloads/19930094559.pdf", "page_number": 9, "total_pages": 16, "image_filename": "19930094559_p9.jpg", "text": "N.A.C.A. Technical Memorandum No. 857 7\n\nextremely small, and complete combustion is attained in the second air cell only at part-load fuel quantities, the relations are reversed in the case of injections at low air velocities. The reason for this is to be sought in the distribution of the fuel quantities under the effect of the different air velocities before the start of ignition, as already described. Between these two extreme values of the start of injection, a favorable range will be found which depends not only on the type and quantity of the fuel but also on the injection arrangement chosen.\n\nSUMMARY\n\nThe flame photographs obtained with combustion-chamber models of engines operating respectively, with turbulence chamber and air-storage chambers or cells, provide an insight into the air and fuel movements that take place before and during combustion in the combustion chamber. The relation between air velocity, start of injection, and time of combustion was determined for the combustion process employing a turbulence chamber. For the chamber with air cells, various forms of combustion were observed in the main combustion chamber and in the air cells depending on the air velocity during injection. Due to the lower temperatures and pressures in the model as compared with the engine, greater ignition lags and lower combustion speeds were obtained in the model tests. The proper interpretation of these model tests leads, however, to good agreement with the measurements on the comparison engines. By the direct observation these tests afford, they give a clear picture of the processes that occur in the engine and confirm the significance of engine test results.\n\nTranslation by S. Reiss,\nNational Advisory Committee\nfor Aeronautics\n\nREFERENCE\n\n1. Holfelder, Otto: Ignition and Flame Development in the Case of Diesel Fuel Injection. Supplement to Forschung auf den Gebiete des Ingenieurwesens, vol. 6, September-October 1935.", "timestamp": "2026-07-19T18:25:59.075388+00:00"}
{"citation_id": "19930094538", "source_url": "https://ntrs.nasa.gov/api/citations/19930094538/downloads/19930094538.pdf", "page_number": 32, "total_pages": 43, "image_filename": "19930094538_p32.jpg", "text": "N. A. C. A. Technical Memorandum No. 878\n30\n\nTable II.\nFormulas for computing\nthe twist of thin-\nwalled, stiffened,\ncircular cylinders.\n\n| Stress / Strain | Complete | Based on Tension field | Incomplete | Calculated according to Ebner and Heck |\n| :--- | :--- | :--- | :--- | :--- |\n| Shearing stress | | $$\\tau = \\frac{T}{2\\pi r_a^2 s} = \\phi \\cdot \\tau_0$$ | | |\n| Donnell's shear stress in buckling | | $$\\tau_0 = 0.815 E \\left(\\frac{s}{r_a}\\right)^{1/2} \\left(\\frac{r_a}{t}\\right)^{1/2}$$ | | |\n| Bulkhead stress $\\sigma_y$ | | $$\\sigma_y = -\\frac{s \\cdot t}{F_y} \\tau_0 (\\theta \\operatorname{tg} \\alpha - 1)$$ | | $$\\sigma_y = -\\frac{s \\cdot t}{F_y} \\tau_0 \\operatorname{tg} \\alpha (\\theta - 1)$$ |\n| Stringer load $p$ | | $$p = s \\varphi \\tau_0 (\\theta \\operatorname{tg} \\alpha - 1)$$ | | $$p = s \\varphi \\tau_0 \\operatorname{tg} \\alpha (\\theta - 1)$$ |\n| Bending stress of stringers $\\sigma_b(x)$ | | $$\\sigma_b(x) = \\frac{s \\varphi t^2}{2 W_L} \\tau_0 (\\theta \\operatorname{tg} \\alpha - 1) \\left[ \\frac{1}{6} - \\frac{x}{t} + \\frac{x^2}{t^2} \\right]$$ | | $$\\sigma_b(x) = \\frac{s \\varphi t^2}{2 W_L} \\tau_0 \\operatorname{tg} \\alpha (\\theta - 1) \\left[ \\frac{1}{6} - \\frac{x}{t} + \\frac{x^2}{t^2} \\right]$$ |\n| Deflection of stringers (average) | | $$\\bar{f} = \\frac{s \\varphi t^4}{720 E J_L} \\tau_0 (\\theta \\operatorname{tg} \\alpha - 1)$$ | | $$\\bar{f} = \\frac{s \\varphi t^4}{720 E J_L} \\tau_0 \\operatorname{tg} \\alpha (\\theta - 1)$$ |\n| Normal stress $\\sigma_n$ | $$\\sigma_n = \\tau_0 (\\theta \\operatorname{ctg} \\alpha - 1)$$ | $$\\sigma_n = \\tau_0 (\\theta \\operatorname{ctg} \\alpha - 1) \\left[ 1 - \\sin^{1/n} \\left( \\frac{\\pi y}{b_1} \\right) (1 + \\delta R) \\right]$$ | | $$\\sigma_n = \\tau_0 \\operatorname{ctg} \\alpha (\\theta - 1)$$ |\n| | | $$- \\sigma_b(x) \\frac{\\delta}{1 + \\delta} (1 - R) \\sin^{1/n} \\frac{\\pi y}{b_1}$$ | | |\n| Principal stress $\\sigma_1$ | $$\\sigma_1 = \\tau_0 \\left( \\frac{\\theta}{\\sin \\alpha \\cos \\alpha} - 1 \\right)$$ | $$\\sigma_1 = \\sigma_n + \\tau \\operatorname{tg} \\alpha$$ | | $$\\sigma_1 = \\tau_0 \\frac{(\\theta - 1)}{\\sin \\alpha \\cos \\alpha}$$ |\n| Principal stress $\\sigma_2$ | $$\\sigma_2 = -\\tau_0$$ | $$\\sigma_2 = \\sigma_n - \\tau \\operatorname{ctg} \\alpha$$ | | In superposed tension field : $\\sigma_2 = 0$ |\n| Compressive stress in stringers $\\sigma_z$ | $$\\sigma_z = -\\delta \\tau_0 (\\theta \\operatorname{ctg} \\alpha - 1)$$ | $$\\sigma_z = -\\delta R \\tau_0 (\\theta \\operatorname{ctg} \\alpha - 1)$$ | | $$\\sigma_z = -\\delta \\tau_0 \\operatorname{ctg} \\alpha (\\theta - 1)$$ |\n| | | $$\\frac{\\delta}{1 + \\delta} (1 - R) \\sigma_b(x)$$ | | |\n| Load ratio $\\chi = P_y/Q$ | $$\\lambda = -\\frac{b_1}{\\varphi \\cdot t} \\frac{\\theta \\operatorname{ctg} \\alpha - 1}{\\theta \\operatorname{tg} \\alpha - 1}$$ | $$\\lambda = -\\frac{b_1}{\\varphi \\cdot t} \\frac{R}{\\theta \\operatorname{tg} \\alpha - 1}$$ | | $$\\lambda = -\\frac{b_1}{\\varphi \\cdot t} \\operatorname{ctg}^2 \\alpha$$ |\n| Angle $\\alpha$ of principal axes $\\operatorname{tg}^2 \\alpha = Z/N$ | $$Z = (1 + \\delta) (\\theta \\operatorname{ctg} \\alpha - 1) + \\theta \\operatorname{tg} \\alpha$$ | $$Z = R (1 + \\delta) (\\theta \\operatorname{ctg} \\alpha - 1) + \\theta \\operatorname{tg} \\alpha$$ | | $$Z = (\\theta - 1) (\\operatorname{tg} \\alpha + \\operatorname{ctg} \\alpha (1 + \\delta))$$ |\n| | $$N = \\theta (\\operatorname{ctg} \\alpha + \\operatorname{tg} \\alpha) - 1$$ | $$N = R (\\theta \\operatorname{ctg} \\alpha - 1) + \\theta \\operatorname{tg} \\alpha$$ | | $$N = (\\theta - 1) (\\operatorname{ctg} \\alpha + \\operatorname{tg} \\alpha (1 + \\gamma + \\delta))$$ |\n| | $$+ (\\gamma + \\delta) (\\theta \\operatorname{tg} \\alpha - 1)$$ | $$+ (\\gamma + \\delta) (\\theta \\operatorname{tg} \\alpha - 1)$$ | | $$+ E/r_a (\\varphi \\gamma s/24 + \\Delta r_a/r_a)$$ |\n| | $$+ \\frac{E}{r_a} \\left( \\frac{\\varphi s}{24} + \\frac{\\Delta r_a}{r_a} \\right)$$ | $$- E/r_a (\\zeta_y - \\Delta r_a/r_a)$$ | | |\n| Angle of twist $\\psi$ | | $$\\psi = 2 l/r_a \\cdot \\tau_0/E \\sqrt{Z \\cdot N}$$ | | |\n\nAbbreviations: $\\chi = \\frac{\\pi y}{b_1}$, $J(a) = \\frac{2}{\\pi} \\int_0^{\\frac{\\pi}{2}} \\sin^{1/n} \\chi \\, d\\chi$; $R = \\frac{1 - J(a)}{1 + \\delta J(a)}$; $\\delta = \\frac{b_1 \\cdot s}{F_z}$; $\\gamma = \\frac{s \\cdot t}{F_y}$; $\\kappa = \\frac{s \\varphi t^4}{720 r_a \\cdot J_L}$", "timestamp": "2026-07-19T18:26:04.145777+00:00"}
{"citation_id": "19930091693", "source_url": "https://ntrs.nasa.gov/api/citations/19930091693/downloads/19930091693.pdf", "page_number": 1, "total_pages": 13, "image_filename": "19930091693_p1.jpg", "text": "NATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\n\nREPORT No. 618\n\nCOMPARATIVE FLIGHT\nAND FULL-SCALE WIND-TUNNEL MEASUREMENTS\nOF THE MAXIMUM LIFT OF AN AIRPLANE\n\nBy ABE SILVERSTEIN, S. KATZOFF, and JAMES A. HOOTMAN\n\n[Figure: Seal of the National Advisory Committee for Aeronautics]\n\n1938\n\nFor sale by the Superintendent of Documents, Washington, D. C. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 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. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .", "timestamp": "2026-07-19T18:26:06.591363+00:00"}
{"citation_id": "19930094549", "source_url": "https://ntrs.nasa.gov/api/citations/19930094549/downloads/19930094549.pdf", "page_number": 12, "total_pages": 76, "image_filename": "19930094549_p12.jpg", "text": "10 N.A.C.A. Technical Memorandum No. 867\n\nThe above elements permit us to determine the characteristic biquadratic equation in $\\lambda^4$.\n\nCALCULATED OSCILLATIONS\n\nThe equation was solved for eight values of $\\mu$. The roots are given in the following table, the coefficient of static stability being evaluated by taking degrees for the unit angle.\n\n| Values of $\\mu$ | Values of $\\lambda$ | |\n| :--- | :--- | :--- |\n| | Short-period oscillation | Long-period oscillation |\n| $\\mu = 0.008$ | $\\lambda = -3.75 \\pm 3.02i$ | $\\lambda = -0.0408 \\pm 0.238i$ |\n| $\\mu = 0.006$ | $\\lambda = -3.75 \\pm 2.5i$ | $\\lambda = -0.0412 \\pm 0.213i$ |\n| $\\mu = 0.004$ | $\\lambda = -3.74 \\pm 1.86i$ | $\\lambda = -0.0437 \\pm 0.19i$ |\n| $\\mu = 0.002$ | $\\lambda = -3.73 \\pm 0.928i$ | $\\lambda = -0.0508 \\pm 0.141i$ |\n| $\\mu = 0.001$ | $\\lambda = -4.58 \\quad \\lambda = -2.86$ | $\\lambda = -0.056 \\pm 0.092i$ |\n| $\\mu = 0.0$ | $\\lambda = -5.18 \\quad \\lambda = -2.255$ | $\\lambda = -0.1275 \\quad \\lambda = 0$ |\n| $\\mu = -0.001$ | $\\lambda = -5.6 \\quad \\lambda = -1.8$ | $\\lambda = -0.2307 \\quad \\lambda = +0.075$ |\n| $\\mu = -0.002$ | $\\lambda = -5.97 \\quad \\lambda = -1.455$ | $\\lambda = -0.2895 \\quad \\lambda = +0.127$ |\n\nThe motion studied is stable when the roots are negative or when, if imaginary, their real parts are negative. The factor $e^{at}$ or $e^{\\lambda t}$ approaches, in this case, zero as $t$ increases.\n\nThe time after which the factor $e^{at}$ attains the value $\\frac{1}{2}$ is given by:\n\n$$T_{\\frac{1}{2}} = \\frac{\\ln 2}{|a|} = \\frac{0.692}{|a|}$$\n\nWhen the coefficient of static stability has a sufficiently large value the characteristic equation defines", "timestamp": "2026-07-19T18:26:07.887136+00:00"}
{"citation_id": "19930091697", "source_url": "https://ntrs.nasa.gov/api/citations/19930091697/downloads/19930091697.pdf", "page_number": 1, "total_pages": 28, "image_filename": "19930091697_p1.jpg", "text": "Library Mass. Inst. of Tech. AERO. & ASTRO. LIBRARY\n[Stamp: MASS. INST. TECH. AERO. ENG. LIBRARY]\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\n\nREPORT No. 622\nCopy # 23\n\nA PHOTOGRAPHIC STUDY OF COMBUSTION AND\nKNOCK IN A SPARK-IGNITION ENGINE\n\nBy A. M. ROTHROCK and R. C. SPENCER\n\n[Figure: Seal of the United States]\n\n1938\n\nFor sale by the Superintendent of Documents, Washington, D. C.\nSubscription price, $3 per year\nPrice 10 cents", "timestamp": "2026-07-19T18:26:08.553309+00:00"}
{"citation_id": "19930094564", "source_url": "https://ntrs.nasa.gov/api/citations/19930094564/downloads/19930094564.pdf", "page_number": 8, "total_pages": 16, "image_filename": "19930094564_p8.jpg", "text": "N.A.C.A. Technical Memorandum No. 852 7\n\nclearly defined that a minimum Reynolds Number effect is expected on the flap itself. Besides, its action between 60 and 80 degrees flap angle discloses a flat optimum which minimizes the importance of flap setting as a source of error at large flap angles.\n\nFor these reasons the split flap was chosen despite the fact that the data obtained with it can be no more than approximately valid for the common split flap or the Fowler flap. It was a full-span flap with 20-percent chord, hinged at 80-percent profile chord and 50-degree setting with respect to the lower wing surface (fig. 6). On the basis of subsequent special studies it would have been better to use a 70-degree setting because it strikes the average value of the optimum angles for thick and thin airfoils more accurately, or else use an adjustable flap altogether and refer the angle of attack in proper form to the airfoil median line instead of to the pressure side. The arrangement as in figure 6 discloses - although only very little - a drawback of the thick airfoils, the thickness slightly reduces the aerodynamic angle of attack of the flap.\n\nThe maximum lift obtained with the flap arrangement of figure 6 is shown in figures 7 to 9 as a function of the effective Reynolds Number. Unfortunately the completion to $R_{\\text{effective}} = 8 \\times 10^6$ is lacking because of the absence of corresponding data from the N.A.C.A. VDT or similar tunnels. We made a temporary extrapolation on the assumption that the lift increase achieved by the split flap is not affected by the Reynolds Number, that is, we joined for the Reynolds Number range of $R_{\\text{effective}} = 4$ to $8 \\times 10^6$ the experimental curve without split flap to the test data with split flap through parallel shifting in direction of higher lift coefficients. The correctness of this extrapolation is confirmed in numerous individual split-flap tests (references 5, 9), which consistently prove that the increase in lift of the split flap is not, or only very little, influenced by the Reynolds Number.\n\nWith a few exceptions (airfoil 2409, 0012, 0015, and 23012) which again manifest a slight $C_{a_{\\text{max}}}$ drop at maximum speed in the DVL tunnel, the extrapolation joins on to the test data very well. Nevertheless the extrapolation will be checked experimentally to the extent that can be achieved by addition of a turbulence screen in the DVL tunnel.", "timestamp": "2026-07-19T18:26:28.329838+00:00"}
{"citation_id": "19930091707", "source_url": "https://ntrs.nasa.gov/api/citations/19930091707/downloads/19930091707.pdf", "page_number": 19, "total_pages": 20, "image_filename": "19930091707_p19.jpg", "text": "CRINKLING STRENGTH AND BENDING STRENGTH OF ROUND AIRCRAFT TUBING 15\n\nTABLE II\nRESULTS OF CRINKLING TESTS\nCHROMIUM-MOLYBDENUM STEEL\n\n| Specimen | $d/t$ | Crinkling strength $f_{cr}$ (lb./sq. in.) | $\\frac{1}{k_2} \\frac{E}{Y} \\frac{t}{d_m}$ | $\\sigma_{cr} \\frac{f_c}{S}$ | Specimen | $d/t$ | Crinkling strength $f_{cr}$ (lb./sq. in.) | $\\frac{1}{k_2} \\frac{E}{Y} \\frac{t}{d_m}$ | $\\sigma_{cr} \\frac{f_c}{S}$ |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| 1CL-Cr | 15.5 | 112000 | 22.58 | 1.233 | 1CO-C | 54.1 | 94000 | 6.19 | 1.018 |\n| 1CE-9 | 15.6 | 112000 | 22.55 | 1.237 | 2CO-C | 58.7 | 88200 | 5.83 | .943 |\n| 2CE-9 | 24.0 | 109000 | 13.51 | 1.135 | 2CO-C2 | 56.9 | 99500 | 5.84 | .959 |\n| 2CE-C | 24.0 | 108000 | 13.51 | 1.117 | 1CM-C | 61.8 | 91500 | 5.49 | .908 |\n| 1CE-1a | 24.6 | 101500 | 13.87 | 1.108 | 2CM-C | 61.8 | 87200 | 5.22 | .917 |\n| 1CR-9 | 34.0 | 103500 | 9.68 | 1.115 | 1CS-3 | 78.4 | 82900 | 3.80 | 1.239 |\n| 1CR-C | 34.0 | 104800 | 9.68 | 1.110 | 1CS-C | 79.1 | 82900 | 3.74 | 1.234 |\n| 1CT-C | 42.7 | 96900 | 7.86 | 1.032 | 1CU-C | 89.8 | 81200 | 4.47 | 1.056 |\n| 1CP-3C2 | 45.2 | 100000 | 7.12 | 1.076 | 1CU-3 | 89.8 | 81100 | 4.47 | 1.054 |\n| 1CP-C | 45.2 | 100000 | 7.12 | 1.069 | 1CC-3 | 99.3 | 83600 | 4.16 | 1.132 |\n| 2CN-C | 52.7 | 87900 | 6.64 | 1.019 | 1CC-C | 99.4 | 83700 | 4.16 | 1.132 |\n\nDURALUMIN\n\n| Specimen | $d/t$ | Crinkling strength $f_{cr}$ (lb./sq. in.) | $\\frac{1}{k_2} \\frac{E}{Y} \\frac{t}{d_m}$ | $\\sigma_{cr} \\frac{f_c}{S}$ | Specimen | $d/t$ | Crinkling strength $f_{cr}$ (lb./sq. in.) | $\\frac{1}{k_2} \\frac{E}{Y} \\frac{t}{d_m}$ | $\\sigma_{cr} \\frac{f_c}{S}$ |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| 1DR-1 | 13.5 | 72300 | 19.74 | 1.756 | F-C | 47.4 | 49900 | 5.59 | 1.211 |\n| 1DR-C | 13.5 | 68700 | 19.74 | 1.663 | E-4 | 55.8 | 48800 | 4.73 | 1.178 |\n| 1-6C | 15.7 | 65400 | 16.43 | 1.494 | 1V-C | 55.8 | 48400 | 4.73 | 1.173 |\n| 1-C | 20.1 | 61000 | 12.77 | 1.316 | 1V-C | 61.6 | 46000 | 4.29 | 1.185 |\n| 1-6 | 20.2 | 61000 | 12.74 | 1.423 | V-C | 61.5 | 46200 | 4.57 | 1.195 |\n| 1-C | 25.6 | 59100 | 10.24 | 1.324 | V-7 | 61.6 | 46000 | 4.57 | 1.185 |\n| 1B-4C | 25.8 | 55100 | 10.73 | 1.431 | 2V-C | 62.2 | 46300 | 4.44 | 1.172 |\n| 1B-7 | 25.9 | 56100 | 10.25 | 1.339 | DH-4 | 65.7 | 47400 | 4.14 | 1.145 |\n| 1B-C | 25.9 | 54400 | 10.71 | 1.413 | DH-C | 65.7 | 47000 | 4.14 | 1.141 |\n| 1B-4C | 35.3 | 50500 | 8.14 | 1.328 | 1B-4 | 78.1 | 45000 | 3.26 | 1.038 |\n| 1B-9 | 35.3 | 50500 | 8.13 | 1.328 | 1B-5 | 78.4 | 45600 | 3.24 | 1.081 |\n| X-4 | 45.9 | 46700 | 5.98 | 1.254 | 1B-4C | 79.2 | 45300 | 3.57 | 1.109 |\n| X-C | 45.9 | 49600 | 5.98 | 1.253 | 1DQ-1 | 97.9 | 41900 | 2.62 | .994 |\n| 1F-3C | 47.3 | 46700 | 5.61 | 1.198 | 1DQ-C | 98.4 | 41800 | 2.60 | .965 |\n\nTABLE III\nRESULTS OF BENDING TESTS\nCHROMIUM-MOLYBDENUM STEEL\n\n| Specimen | $d/t$ | Modulus of Rupture $f_r$ (lb./sq. in.) | $\\frac{1}{k_2} \\frac{E}{Y} \\frac{t}{d_m}$ | $\\sigma_{cr} \\frac{f_c}{S}$ | Specimen | $d/t$ | Modulus of Rupture $f_r$ (lb./sq. in.) | $\\frac{1}{k_2} \\frac{E}{Y} \\frac{t}{d_m}$ | $\\sigma_{cr} \\frac{f_c}{S}$ |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| 1CL-1a | 15.5 | 139500 | 22.60 | 1.553 | 1CN-3 | 33.1 | 110000 | 6.31 | 1.215 |\n| 1CE-3 | 15.6 | 140000 | 22.52 | 1.559 | 1CO-3 | 54.4 | 116100 | 6.16 | 1.258 |\n| 2CE-3 | 24.0 | 128000 | 13.51 | 1.345 | 2CM-2 | 61.9 | 113300 | 5.22 | 1.191 |\n| 1CE-3a | 24.5 | 128500 | 13.62 | 1.400 | 1CM-3 | 62.1 | 118200 | 5.19 | 1.221 |\n| 1CR-7 | 34.3 | 128000 | 9.61 | 1.363 | 1CS-2 | 78.4 | 88800 | 3.80 | 1.327 |\n| 1CR-3a | 34.3 | 128000 | 9.61 | 1.354 | 1CS-1 | 79.0 | 88500 | 3.75 | 1.322 |\n| 1CT-2 | 42.8 | 120100 | 7.84 | 1.304 | 1CU-1 | 88.6 | 99600 | 4.54 | 1.288 |\n| 1CP-1 | 44.9 | 121400 | 7.16 | 1.290 | 1CU-2 | 89.8 | 95800 | 4.47 | 1.239 |\n| 1CP-2 | 44.9 | 121400 | 7.16 | 1.294 | 1CC-2 | 99.3 | 91800 | 4.16 | 1.243 |\n| 1CP-1 | 44.9 | 116800 | 7.16 | 1.246 | 1CC-1 | 100.0 | 89000 | 4.13 | 1.204 |\n\nDURALUMIN\n\n| Specimen | $d/t$ | Modulus of Rupture $f_r$ (lb./sq. in.) | $\\frac{1}{k_2} \\frac{E}{Y} \\frac{t}{d_m}$ | $\\sigma_{cr} \\frac{f_c}{S}$ | Specimen | $d/t$ | Modulus of Rupture $f_r$ (lb./sq. in.) | $\\frac{1}{k_2} \\frac{E}{Y} \\frac{t}{d_m}$ | $\\sigma_{cr} \\frac{f_c}{S}$ |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| v-3 | 15.7 | 73100 | 16.49 | 1.669 | E-3 | 55.8 | 57200 | 4.73 | 1.389 |\n| 1-4 | 20.2 | 70300 | 12.73 | 1.632 | E-4 | 61.6 | 54400 | 4.56 | 1.408 |\n| 1-5 | 20.2 | 69000 | 12.73 | 1.600 | DH-2 | 65.7 | 50100 | 4.15 | 1.391 |\n| b-5 | 25.9 | 65000 | 10.24 | 1.561 | DH-3 | 65.7 | 52600 | 4.14 | 1.350 |\n| 1b-5 | 25.9 | 65200 | 10.71 | 1.642 | DH-1 | 65.9 | 53300 | 4.13 | 1.342 |\n| 1p-6 | 35.3 | 59100 | 8.14 | 1.556 | 1B-4 | 78.4 | 52900 | 3.24 | 1.259 |\n| 1p-8 | 35.3 | 59900 | 8.13 | 1.576 | 1B-5 | 79.6 | 52100 | 3.35 | 1.277 |\n| X-3 | 45.9 | 59400 | 5.98 | 1.501 | 1DQ-2 | 98.4 | 49200 | 2.60 | 1.169 |\n| 1F-4 | 47.4 | 60900 | 5.59 | 1.467 | 1DQ-4 | 98.4 | 46800 | 2.60 | 1.115 |\n\nREFERENCES\n\n1. Timoshenko, S.: Theory of Elastic Stability. McGraw-Hill Book Co., Inc., 1936.\n2. Hansen, Knud E.: Bending Strength of Thin-Walled Cylindrical Tubes. Bygningsstatiske Meddelelser, vol. 9, no. 1 (Copenhagen), 1937.\n3. Tuckerman, L. B.: Discussion of paper \"The Determination and Significance of the Proportional Limit in the Testing of Metals,\" by R. L. Templin. A. S. T. M. Proc., vol. 29, pt. II, 1929, pp. 538-546.\n4. Osgood, William R.: Column Strength of Tubes Elastically Restrained against Rotation at the Ends. T. R. No. 615, N. A. C. A., 1938.\n5. Geckeler, J. W.: Plastisches Knicken der Wandung von Hohlzylindern und einige andere Faltungserscheinungen an Schalen und Blechen. Z. t. a. M. M., Bd. 8, Heft 5, Oktober 1928, S. 341-352.\n6. Flügge, W.: Die Stabilität der Kreiszylinderschale. Ingenieur-Archiv, Bd. III, Heft 5, Dez. 1932, S. 463-506.\n\nU. S. GOVERNMENT PRINTING OFFICE: 1939", "timestamp": "2026-07-19T18:26:31.917843+00:00"}
{"citation_id": "19930094533", "source_url": "https://ntrs.nasa.gov/api/citations/19930094533/downloads/19930094533.pdf", "page_number": 43, "total_pages": 51, "image_filename": "19930094533_p43.jpg", "text": "N.A.C.A. Technical Memorandum No. 883\nFigs. 21,22\n\n[Figure: Diagram of an airfoil with lines radiating from the surface to an outer boundary, with small circles at the endpoints. A scale bar labeled \"Scale 0.2°\" is shown to the right.]\n\nFigure 21.- Temperature record at 10.7° incidence(model)\n\n[Figure: Diagram of an airfoil with lines radiating from the surface to an outer boundary, with small circles at the endpoints. A scale bar labeled \"Scale 0.2°\" is shown to the right.]\n\nFigure 22.- Temperature record at 14.7° incidence(model).", "timestamp": "2026-07-19T18:26:40.798598+00:00"}
{"citation_id": "19930091655", "source_url": "https://ntrs.nasa.gov/api/citations/19930091655/downloads/19930091655.pdf", "page_number": 7, "total_pages": 22, "image_filename": "19930091655_p7.jpg", "text": "The extent to which the droplets are not black bodies introduces a factor that reduces this rate. The existence of black-body radiation within the bomb is actually the case prior to injection, but thereafter the radiation within the spray envelope undoubtedly corresponds to a lower temperature than that of the bomb walls. This fact again entails a diminution in the rate indicated by the equation. The area $S$ may be taken as equivalent to that of the number of droplets of average size (reference 20) required for a spray of given weight.\n\nFor comparison with the observed rate of heat transfer, expressed in this case in terms of rate of pressure drop, the calculated rate of radiant transfer must be expressed in identical units. Although not indicative of the actual mechanism, this rate can be put in terms of the units in which the experimental data are expressed by considering all the radiant energy as being derived from the ambient gas.\n\nBasic considerations of experimental method.—The observed decrease in pressure accompanied the decrease in temperature of the ambient gas caused by the flow of heat from it to the injected liquid. This process was essentially adiabatic in view of the small rate of heat transfer from the bomb wall. Cragoe's empirical relations (reference 17) for the specific and vaporization heats of oils permit the calculation of the pressure decrease that should accompany the complete vaporization of a given amount of fuel when all the heat is abstracted from the gas phase. When the total heat absorbed by the fuel in vaporizing is equated to that lost by the ambient gas, these relations lead to an expression that can be solved by trial and error to give the final equilibrium temperature:\n\n$$\nt_f = \\frac{NC_v t_i - 49.5w}{NC_v + 0.333w + 0.000444wt_i}\n$$\n\nwhere  \n$N$ is the moles of ambient gas,  \n$C_v$, the molal specific heat of this gas, taken as constant between $t_f$ and $t_i$,  \n$t_i$, the initial gas temperature, °C.  \n$w$, the weight of injected fuel, grams.\n\nIt follows at once that the temperature drop ($t_i - t_f$) should remain constant for a given initial temperature and gas-fuel ratio, i. e., essentially $N/w$. The corresponding diminutions in the partial pressure of the gas are calculable from the expression\n\n$$\nP_i - P_f = \\frac{P_i(T_i - T_f)}{T_i}\n$$\n\nin which $P_i$ is the initial pressure, atmospheres.  \n$P_f$, the final pressure, atmospheres.  \n$T_i$, the initial absolute gas temperature, degrees K.  \n$T_f$, the final equilibrium gas temperature, degrees K.\n\nThis expression is strictly applicable only to the initial stage of the heat-transfer process when little vapor exists. Later in the process, however, these diminutions should be greater than the experimentally derived maximum values to the extent of the partial pressures of the vapor, these latter being directly proportional to the initial pressure for a given initial temperature and gas-fuel ratio. Thus the calculated actual drop to be expected under these stipulated conditions is in accordance with\n\n$$\n(P_i - P_f)_{\\text{actual}} = P_i \\left[ \\frac{(T_i - T_f)}{T_i} - C \\right]\n$$\n\nwhere $C$ is a correction factor necessitated by the presence of the fuel vapor. This factor is equal to the partial pressure of the vapor of the injected liquid divided by the initial gas pressure. The partial pressure was obtained from the perfect gas law and the known bomb volume, gas temperature, fuel weight for the particular gas-fuel ratio, and an estimated average molecular weight of 200 (reference 22). It follows from this expression that the pressure drop should be directly proportional to the initial pressure if the fuel derived all its heat from the gas phase under the assumed conditions of constant initial temperature and air-fuel ratio.\n\nSome conception of the rate of heat transfer can also be obtained from the experimental results, particularly for the early part of the process in which the number of moles of gas is essentially invariant. It follows from the perfect gas law that the rate of pressure change is related to the rate of temperature change by\n\n$$\n\\frac{dP}{dt} = \\frac{NR}{V} \\frac{dT}{dt}\n$$\n\nwherein $R$ is the gas constant.  \n$V$, the volume of the bomb.\n\nAlso, the rate of change in the energy content of the gas phase must equal the rate of heat transfer, thus:\n\n$$\n\\frac{dQ}{dt} = C_v N \\frac{dT}{dt}\n$$\n\nIf minor variations in $C_v$ and $N$ are neglected, it follows that the rate of heat transfer is proportional to the rate of pressure decrease. For the practical purpose of showing the trends in the present data it is sufficient to use these rates interchangeably as though they were synonymous.\n\nThe expression $\\frac{dP}{dt} = \\frac{NR}{V} \\frac{dT}{dt}$ can also be used to compare the relative rates of temperature drop in different gases when the corresponding rates of pressure change are known for a given initial gas temperature, density, and gas-fuel ratio. The initial pressure under such conditions is very nearly proportional to $N$, hence the ratio of the initial rate of pressure drop to the initial pressure is proportional to the initial rate of temperature drop irrespective of the nature of the gas.", "timestamp": "2026-07-19T18:26:49.190764+00:00"}
{"citation_id": "19930094538", "source_url": "https://ntrs.nasa.gov/api/citations/19930094538/downloads/19930094538.pdf", "page_number": 33, "total_pages": 43, "image_filename": "19930094538_p33.jpg", "text": "N. A. C. A. Technical Memorandum No. 878\nFigs.1,5,6,7\n\n[Figure: Sketch of a stiffened circular cylinder showing cross-section with stringers and longitudinal section with spacing $t$]\n\nFigure 1.- Sketch of a stiffened circular cylinder.\n\nOriginal generating line\nSection line of the wrinkled skin\nMean skin radius\n$r_0$\nLongitudinal section of shell\n(radial plane Z-Z)\nRadius through center\nOriginal circumference\nTrace (base line) of substitute surface\nChord\nShell section\nMean skin radius\n$\\varphi$\n$\\varphi_1$\nZ\n\nFigure 5.- Determination of wrinkling $\\zeta\\gamma$.\n\nBulkhead center\nPrincipal strain direction\nStringer center\nStringer center\n$t$\n$\\sigma_r$\n$\\tau$\n$\\sigma_\\theta$\n$\\sigma_1$\n$\\sigma_2$\n$x$\n$y$\n\nFigure 7.- Equilibrium at skin element\n\nFigure 6.- Coordinate system and strain components.", "timestamp": "2026-07-19T18:26:51.058894+00:00"}
{"citation_id": "19930091692", "source_url": "https://ntrs.nasa.gov/api/citations/19930091692/downloads/19930091692.pdf", "page_number": 15, "total_pages": 20, "image_filename": "19930091692_p15.jpg", "text": "AUTO-IGNITION AND COMBUSTION OF DIESEL FUEL IN A CONSTANT-VOLUME BOMB 11\n\nundoubtedly true to some extent; otherwise increasing the air turbulence would not result in an improvement in engine performance such as has been obtained (reference 29). On the other hand, there is abundant evidence that the rates of some reactions are altered by their products, either by some specific action or by altering some physical factor such as the flame temperature. Slow burning in a compression-ignition engine could conceivably be attributed to the fact that portions of the unburned fuel are encompassed by mixtures of combustion products and air. It was to investigate this point that tests were conducted in which nitrogen or combustion products (of which a large fraction was also nitrogen) were mixed with air of a fixed concentration before the fuel in question was injected. The most pronounced effect of increasing the percentage of combustion products was an increase in the ignition lag together with some decrease in the maximum rate of pressure rise (fig. 9). Bird (reference 7) has reported similar results for repeated injections into the same air charge, but the comparisons were between different air concentrations for each injection. The effect of the combustion products on the ratio of the explosion to initial pressure after 0.004 second was less real than is apparent from a casual inspection of the records. Thus, even if the fuel burned in this period remained constant, a decrease in the pressure ratio was to be expected because of the necessity of heating a respectively greater mass of gas for records 113 and 117 than for record 103. For example, the ratio of the absolute pressure after 0.004 second to the absolute initial pressure for record 103 multiplied by the ratio of the initial absolute pressure for record 103 to that for record 113 gives an approximate value of the pressure ratio that should be observed for record 113. This calculated ratio happened to be identical with the observed ratio for this case, thus proving that the extent of the combustion within this interval was about the same for records 103 and 113. The observed pressure ratio for record 117 was slightly greater than the calculated ratio, which might have been due to the better mixing permitted by the much longer ignition lag. Figure 10 shows that the addition of nitrogen or of combustion products to an air charge of fixed concentration definitely decreased the maximum rate of pressure rise. For the lower effective air density, at least, the addition of nitrogen or combustion products had less influence on the ignition lag at the higher temperature corresponding to figure 10 than is evident from figure 9. This difference may be seen by comparing records 168 and 156 with records 103 and 113, the initial concentration of combustion products being the same for records 156 and 113. At a higher effective air density (0.89 pound per cubic foot), however, a definite change in ignition lag for the air-combustion products mixture is evident. Incidentally, the presence of water vapor in Wentzel’s tests (reference 10) probably accounted in part for the long lags he observed but, as the ignition quality of his fuel is unknown, no direct comparisons are possible. MacGregor (reference 30) has shown that variations in humidity affect the knocking characteristics of a fuel in spark-ignition engines.\n\nFigure 11 illustrates the fact that, in spite of the tendency of diluent gases to reduce the maximum rate of pressure rise, very high rates could be obtained by sufficiently decreasing the air-fuel ratio. The maximum rates of pressure rise for records 168, 165, and 159 were by no means equal; nevertheless, the permissible decrease in the air-fuel ratio made possible by the addition of inert gases to air of the same effective concentration was very definite.\n\nCONCLUSIONS\n\n1. For fuel injection into a constant-volume bomb containing stagnant air at a temperature and a pressure approximating those in a compression-ignition engine, the ignition lag was essentially independent of the injected fuel quantity and was of the same magnitude as in the engine.\n\n2. For the fuel used, the possible decrease in the ignition lag for a given increase in air temperature or density became quite small at temperatures and densities in excess of those generally occurring in compression-ignition engines.\n\n3. The combustion efficiency improved as the ignition lag was lengthened; hence it should be worth while to use those fuels in an engine whose ignition lags correspond to the higher permissible rates of pressure rise. The useless “afterburning” decreased as the ignition lag was lengthened.\n\n4. The ignition lag tended to increase and the maximum rate of pressure rise definitely decreased upon the addition of inert gases to an air charge of fixed concentration.\n\nLANGLEY MEMORIAL AERONAUTICAL LABORATORY,  \nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS,  \nLANGLEY FIELD, VA., October 5, 1937.\n\nREFERENCES\n\n1. 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\n2. Boerlage, G. D., and Broeze, J. J.: Ignition Quality of Diesel Fuels as Expressed in Cetene Numbers. S. A. E. Jour., vol. 31, no. 1, July 1932, pp. 283–293.\n\n3. Good, R. F.: Cetane Numbers—Life Size. S. A. E. Jour., vol. 40, no. 6, June 1937, pp. 232–241.\n\n4. Tausz, J., and Schultz, F.: Ignition Points and Combustion Reactions in Diesel Engines. Part I, T. M. No. 483, and Part II, T. M. No. 484, N. A. C. A., 1928.", "timestamp": "2026-07-19T18:26:58.320809+00:00"}
{"citation_id": "19930094542", "source_url": "https://ntrs.nasa.gov/api/citations/19930094542/downloads/19930094542.pdf", "page_number": 73, "total_pages": 102, "image_filename": "19930094542_p73.jpg", "text": "N.A.C.A. Technical Memorandum No. 874\n\nFigs.54,55,62,63,64\n\n[Figure: Graph showing $C_L$ vs $\\alpha$ for various $\\lambda$ values and wing alone, with $\\kappa = -6^\\circ$. Legend includes $\\lambda=.025$, $\\lambda=.035$, $\\lambda=.050$, $\\lambda=.070$, $\\lambda=.080$, $\\lambda=.013$, and wing alone.]\n\nFigure 54. $\\kappa = -6^\\circ$.\n\n[Figure: Graph showing $C_L$ vs $\\alpha$ for $\\lambda = 0.13$, with curves for $\\kappa = 0^\\circ$, $\\kappa = -4^\\circ$, $\\kappa = -7^\\circ$, $\\kappa = -9^\\circ$.]\n\nFigure 55. $\\lambda = 0.13$.\n\nFigures 54,55.—Effect of propeller slipstream on lift coefficient.\n\n[Figure: Polar plots of $C_L$ vs $C_D$ for different $\\lambda$ values ($\\lambda=.020$, $\\lambda=.028$, $\\lambda=.036$) and wing alone, with arrows indicating direction of increasing $\\alpha$.]\n\nFigure 62.—Polars of wing in presence of propeller.\n\n[Figure: Graph showing $C_T$ vs $\\lambda$, with curves labeled “propeller alone”, “propeller on wing”, and “wing alone”.]\n\nFigure 63.—Thrust coefficient as function of $\\lambda$.\n\n[Figure: Graph showing $C_M$ vs $C_L$ for various $\\lambda$ values ($\\lambda=.025$, $\\lambda=.035$, $\\lambda=.050$, $\\lambda=.070$, $\\lambda=.080$, $\\lambda=.013$) and wing alone, with $\\kappa = -9^\\circ$.]\n\nFigure 64.—Moment curves of wing in presence of propeller.", "timestamp": "2026-07-19T18:27:01.249848+00:00"}
{"citation_id": "19930094534", "source_url": "https://ntrs.nasa.gov/api/citations/19930094534/downloads/19930094534.pdf", "page_number": 17, "total_pages": 21, "image_filename": "19930094534_p17.jpg", "text": "N.A.C.A. Technical Memorandum No. 882\nFigs.1,2,3,4,6,7,8\n\nPunch\nTension flange with\nRotary locking.\nSleeve\nMandrel\nDrawplate\n\nPipe\nVise\nTable\n\nFigure 1.- Outside drawing\n(or expanding) tool.\n\nFigure 2.- Contracted thin-walled tube.\n\nFigure 4.- Contracted thin-wall tube.\n\nContracting pressure $\\frac{\\sin \\alpha + \\mu}{\\cos \\alpha \\times \\sin \\alpha}$\nContraction angle $\\alpha$\n\nFigure 3.- Effect of contraction angle on the contracting pressure for different coefficients of friction.\n\nFigure 6.- Push rod with adjustable eye.\n\nBuckling test\nSection A Tensile test Break 1\n\nFigures 7,8.- Loading arrangement for buckling and tensile tests.", "timestamp": "2026-07-19T18:27:02.483277+00:00"}
{"citation_id": "19930094535", "source_url": "https://ntrs.nasa.gov/api/citations/19930094535/downloads/19930094535.pdf", "page_number": 12, "total_pages": 17, "image_filename": "19930094535_p12.jpg", "text": "N.A.C.A. Technical Memorandum No. 881 11\n\nand facing the south. The specimens lie unprotected, so that they are freely exposed to the action of heat and cold, sunlight and moisture. The data in table I, giving the mean ratios of effective sun exposure per day and month were used as a basis for the measurement of the duration of the exposure, since the principal action is that of the sun's rays. After a weathering of 300, 450, and 600 working hours, the specimens are tested at 20° by the falling-ball test.\n\nFor the bending-stress test, the specimens are placed on two supports and loaded in the center. Steel cylindrical rollers 50 millimeters in diameter and 300 millimeters long, ground as accurately as possible, serve to transmit the load. The distance between the supports is 200 millimeters. The bending strength is then computed from the formula\n\n$$\n\\sigma = \\frac{P}{d^2} \\text{ (kg/cm}^2\\text{)}\n$$\n\nwhere P is the load, and d the thickness of the specimen.\n\nThis test procedure, with some further details and explanations, is described in the DIN standard test-procedure sheet DVM 2302 \"Safety glass for Vehicles, test Procedure.\"\n\nAs may be seen from the foregoing brief discussion, the greatest efforts are being made to remove the source of danger to passengers, arising from the fracturing of silicate glass. The problem, of course, has not yet been completely solved with the artificial types of glass at present available, but the danger of serious injuries from broken glass has been considerably lessened by the application of the safety glasses. It is to the interest of the entire transportation industry - both airplane and automobile - to insure maximum safety to the lives of passengers entrusted to its care, and this idea has been gaining ground to an ever-increasing extent.\n\nTranslation by S. Reiss,\nNational Advisory Committee\nfor Aeronautics", "timestamp": "2026-07-19T18:27:02.909927+00:00"}
{"citation_id": "19930094549", "source_url": "https://ntrs.nasa.gov/api/citations/19930094549/downloads/19930094549.pdf", "page_number": 13, "total_pages": 76, "image_filename": "19930094549_p13.jpg", "text": "N.A.C.A. Technical Memorandum No. 867 11\n\nin general two very different types of oscillation. One oscillation is of short period and is very damped; the second has a long period and is less damped. As the static stability decreases the motion ceases to be oscillatory. In the example given, the rapid oscillation loses its character somewhat earlier than the slow oscillation. The periods and damping are given in the table below.\n\n| μ | Short-period oscillation (seconds) | | Long-period oscillation (seconds) | |\n|---|---|---|---|---|\n| | T | τ½ | T | τ½ |\n| 0.008 | 2.08 | 0.185 | 26.4 | 17 |\n| .006 | 2.51 | .185 | 29.5 | 16.8 |\n| .004 | 3.38 | .185 | 33 | 15.9 |\n| .002 | 6.7 | .186 | 44.5 | 13.6 |\n| .001 | aperiodic | | 68 | 12.4 |\n| .00 | aperiodic | | aperiodic | |\n\nWhen μ = 0, the two kinds of oscillations vanish, the roots becoming real and one of them zero.\n\nIn passing to the condition of static instability the total motion is the sum of four periodic motions, three of them corresponding to disturbances decreasing with time, the fourth corresponding to a disturbance increasing with time and indicating dynamic instability. The table of roots shows clearly that the unstable motion corresponds to one of the components of the slow oscillation. The aperiodic motions which replace the rapid oscillation do not cease to be stable within the limits of static instability considered.\n\nWhen the motions are aperiodic the equations defining the motion cannot be put in the sinusoidal form VIII and must remain in the exponential form IV.", "timestamp": "2026-07-19T18:27:03.064335+00:00"}
{"citation_id": "19930094544", "source_url": "https://ntrs.nasa.gov/api/citations/19930094544/downloads/19930094544.pdf", "page_number": 23, "total_pages": 43, "image_filename": "19930094544_p23.jpg", "text": "N.A.C.A. Technical Memorandum No. 372 21\n\nplanes of the airship (references 17 and 33). In the construction of the English airships, on the other hand, combinations of loading conditions are considered in much greater number than was previously customary (reference 34).\n\nAlso in the matter of safety, distinct progress has been made in the newer airship structures. In German airship construction a uniform factor of safety (against breaking) of 2 for tension and compression is taken as a basis. With this the factor of safety for tension is applied to the tensile strength of the material and that for compression is applied to the experimentally established compression strength of the member concerned. In the American construction, on the other hand, the factor of safety 2 applies against exceeding the so-called \"yield point,\" which in the alloy used, 17SRT, lies approximately around $30 \\text{ kg/mm}^2$ (reference 33). Since this limit agrees approximately with the compressive stress attained in the compression members, this gives, even more severely than in airplane construction, a distinct security against the breaking of tension and compression members. A still more extensive graduation of factors of safety is followed out in the English constructions. The required factors of safety (against breaking) lie, depending on the kind of stress, between 2 and 4 (reference 35).\n\nWith the high degree of static indeterminateness, the exact calculation of an airship framework as a statically indeterminate space framework practically can not be accomplished. On this account one is compelled to adopt approximate methods (references 36 and 37). The simplest and, under certain hypotheses, also the most suitable approximate method consists in considering the entire airship frame to be a homogeneous beam, and to calculate according to the usual bending theory. In the determination of the moment of inertia of such a beam one must, however, consider not only the circular cross sections, but also the diagonal reinforcement of the tension zone by the outer panel and the inner net stressing, and under certain circumstances also that by the outer covering. In what magnitudes the individual portions are to be taken depends on the transverse force acting at the section considered.\n\nAnother approximate method consists in calculating the forces in the diagonals of the outer surface under the hypothesis that the transverse rings are rigid in and perpendicular to their planes and that only a parallel dis-", "timestamp": "2026-07-19T18:27:06.946124+00:00"}
{"citation_id": "19930091697", "source_url": "https://ntrs.nasa.gov/api/citations/19930091697/downloads/19930091697.pdf", "page_number": 2, "total_pages": 28, "image_filename": "19930091697_p2.jpg", "text": "# AERONAUTIC SYMBOLS\n\n## 1. FUNDAMENTAL AND DERIVED UNITS\n\n| | Symbol | Metric | | English | |\n| :--- | :---: | :--- | :--- | :--- | :--- |\n| | | **Unit** | **Abbreviation** | **Unit** | **Abbreviation** |\n| Length<br>Time<br>Force | $l$<br>$t$<br>$F$ | meter<br>second<br>weight of 1 kilogram | m<br>s<br>kg | foot (or mile)<br>second (or hour)<br>weight of 1 pound | ft. (or mi.)<br>sec. (or hr.)<br>lb. |\n| Power<br>Speed | $P$<br>$V$ | horsepower (metric)<br>kilometers per hour<br>meters per second | k.p.h.<br>m.p.s. | horsepower<br>miles per hour<br>feet per second | hp.<br>m.p.h.<br>f.p.s. |\n\n## 2. GENERAL SYMBOLS\n\n$W$, Weight=$mg$\n$g$, Standard acceleration of gravity=9.80665 m/s$^2$ or 32.1740 ft./sec.$^2$\n$m$, Mass=$\\frac{W}{g}$\n$I$, Moment of inertia=$mk^2$. (Indicate axis of radius of gyration $k$ by proper subscript.)\n$\\mu$, Coefficient of viscosity\n$\\nu$, Kinematic viscosity\n$\\rho$, Density (mass per unit volume)\nStandard density of dry air, 0.12497 kg-m$^{-4}$-s$^2$ at 15° C. and 760 mm; or 0.002378 lb.-ft.$^{-4}$ sec.$^2$\nSpecific weight of \"standard\" air, 1.2255 kg/m$^3$ or 0.07651 lb./cu. ft.\n\n## 3. AERODYNAMIC SYMBOLS\n\n$S$, Area\n$S_w$, Area of wing\n$G$, Gap\n$b$, Span\n$c$, Chord\n$b^2$, Aspect ratio\n$\\frac{b}{S}$, Aspect ratio\n$V$, True air speed\n$q$, Dynamic pressure=$\\frac{1}{2}\\rho V^2$\n$L$, Lift, absolute coefficient $C_L=\\frac{L}{qS}$\n$D$, Drag, absolute coefficient $C_D=\\frac{D}{qS}$\n$D_0$, Profile drag, absolute coefficient $C_{D_0}=\\frac{D_0}{qS}$\n$D_i$, Induced drag, absolute coefficient $C_{D_i}=\\frac{D_i}{qS}$\n$D_p$, Parasite drag, absolute coefficient $C_{D_p}=\\frac{D_p}{qS}$\n$C$, Cross-wind force, absolute coefficient $C_C=\\frac{C}{qS}$\n$R$, Resultant force\n$i_w$, Angle of setting of wings (relative to thrust line)\n$i_t$, Angle of stabilizer setting (relative to thrust line)\n$Q$, Resultant moment\n$\\Omega$, Resultant angular velocity\n$\\frac{Vl}{\\mu}$, Reynolds Number, where $l$ is a linear dimension (e.g., for a model airfoil 3 in. chord, 100 m.p.h. normal pressure at 15° C., the corresponding number is 234,000; or for a model of 10 cm chord, 40 m.p.s., the corresponding number is 274,000)\n$C_p$, Center-of-pressure coefficient (ratio of distance of c. p. from leading edge to chord length)\n$\\alpha$, Angle of attack\n$\\epsilon$, Angle of downwash\n$\\alpha_0$, Angle of attack, infinite aspect ratio\n$\\alpha_i$, Angle of attack, induced\n$\\alpha_a$, Angle of attack, absolute (measured from zero-lift position)\n$\\gamma$, Flight-path angle", "timestamp": "2026-07-19T18:27:07.138025+00:00"}
{"citation_id": "19930091693", "source_url": "https://ntrs.nasa.gov/api/citations/19930091693/downloads/19930091693.pdf", "page_number": 2, "total_pages": 13, "image_filename": "19930091693_p2.jpg", "text": "# AERONAUTIC SYMBOLS\n\n## 1. FUNDAMENTAL AND DERIVED UNITS\n\n| Symbol | Metric | | English | |\n| :--- | :--- | :--- | :--- | :--- |\n| | **Unit** | **Abbreviation** | **Unit** | **Abbreviation** |\n| Length... | $l$ | meter | m | foot (or mile) | ft. (or mi.) |\n| Time... | $t$ | second | s | second (or hour) | sec. (or hr.) |\n| Force... | $F$ | weight of 1 kilogram | kg | weight of 1 pound | lb. |\n| Power... | $P$ | horsepower (metric) | | horsepower | hp. |\n| | | kilometers per hour | k.p.h. | miles per hour | m.p.h. |\n| Speed... | $V$ | meters per second | m.p.s. | feet per second | f.p.s. |\n\n## 2. GENERAL SYMBOLS\n\n$W$, Weight=$mg$\n$g$, Standard acceleration of gravity=9.80665 m/s$^2$ or 32.1740 ft./sec.$^2$\n$m$, Mass=$\\frac{W}{g}$\n$I$, Moment of inertia=$mk^2$. (Indicate axis of radius of gyration $k$ by proper subscript.)\n$\\mu$, Coefficient of viscosity\n$\\nu$, Kinematic viscosity\n$\\rho$, Density (mass per unit volume)\nStandard density of dry air, 0.12497 kg-m$^{-4}$s$^2$ at 15$^\\circ$ C. and 760 mm; or 0.002378 lb.-ft.$^{-4}$ sec.$^2$\nSpecific weight of \"standard\" air, 1.2255 kg/m$^3$ or 0.07651 lb./cu. ft.\n\n## 3. AERODYNAMIC SYMBOLS\n\n$S$, Area\n$S_w$, Area of wing\n$G$, Gap\n$b$, Span\n$c$, Chord\n$b^2$, Aspect ratio\n$S'$\n$V$, True air speed\n$q$, Dynamic pressure=$\\frac{1}{2}\\rho V^2$\n$L$, Lift, absolute coefficient $C_L=\\frac{L}{qS}$\n$D$, Drag, absolute coefficient $C_D=\\frac{D}{qS}$\n$D_0$, Profile drag, absolute coefficient $C_{D_0}=\\frac{D_0}{qS}$\n$D_i$, Induced drag, absolute coefficient $C_{D_i}=\\frac{D_i}{qS}$\n$D_p$, Parasite drag, absolute coefficient $C_{D_p}=\\frac{D_p}{qS}$\n$C$, Cross-wind force, absolute coefficient $C_C=\\frac{C}{qS}$\n$R$, Resultant force\n$i_w$, Angle of setting of wings (relative to thrust line)\n$i_p$, Angle of stabilizer setting (relative to thrust line)\n$Q$, Resultant moment\n$\\Omega$, Resultant angular velocity\n$\\rho \\frac{Vl}{\\mu}$, Reynolds Number, where $l$ is a linear dimension (e.g., for a model airfoil 3 in. chord, 100 m.p.h. normal pressure at 15$^\\circ$ C., the corresponding number is 234,000; or for a model of 10 cm chord, 40 m.p.s., the corresponding number is 274,000)\n$C_p$, Center-of-pressure coefficient (ratio of distance of c.p. from leading edge to chord length)\n$\\alpha$, Angle of attack\n$\\epsilon$, Angle of downwash\n$\\alpha_0$, Angle of attack, infinite aspect ratio\n$\\alpha_i$, Angle of attack, induced\n$\\alpha_a$, Angle of attack, absolute (measured from zero-lift position)\n$\\gamma$, Flight-path angle", "timestamp": "2026-07-19T18:27:08.169648+00:00"}
{"citation_id": "19930094564", "source_url": "https://ntrs.nasa.gov/api/citations/19930094564/downloads/19930094564.pdf", "page_number": 9, "total_pages": 16, "image_filename": "19930094564_p9.jpg", "text": "8 N.A.C.A. Technical Memorandum No. 852\n\nV. RESULTS OF PROFILE DRAG MEASUREMENTS AT LOW LIFT\n\nThe method of profile drag measurements developed by the DVL has been described in detail in reference 6. A résumé therefore suffices.\n\nThe profile drag measurements on the airfoil series 00, 24, and 230 were fundamentally made in two ways: first, by employing the usual method of measuring the forces on the balance, then by measuring the loss of momentum according to Betz. In these measurements the normal airfoils of 4-meter span and 0.8-meter chord were used. The fact that both methods gave the same result after rounding off the wing tips and subtracting the drag corresponding to the area of rounding, is proof that the profile drag of the plane problem had been reached very closely.\n\nAs regards the effect of the jet turbulence on profile drag, there was not and is not even today any clear perception. From comparing the DVL data with those of the N.A.C.A. VDT on the basis of the same Reynolds Number ($R_{\\text{effective}} \\cong 8.2 \\times 10^6$) and the same tip shape (blunt tips) it may be assumed that the turbulence effect is influenced by the thickness (fig. 10). For thick airfoils the rise in profile drag due to turbulence is substantially greater than for thin airfoils. Elsewhere (reference 5) it had been attempted to convert the profile drag of an airfoil to effective Reynolds Number by subtracting the drag difference between $R_{\\text{test}}$ and $R_{\\text{effective}}$ of the fully turbulent friction curve of the flat plate. The corresponding drag difference ($\\Delta c_w$) has been subtracted from the two drag curves of figure 10. It is seen that, while for very small profile thickness the correction effects an approximate agreement, it is unsatisfactory for the practical range of thicknesses. So long as this difference remains to be cleared up, it is a mistake even at present to make profile drag tests in low-turbulence tunnels, because they conform much better to free flight conditions.\n\nWith low turbulence the effective Reynolds Number reached on normal airfoils in the DVL tunnel ($R_{\\text{effective}} \\cong 3.5 \\times 10^6$) is very low compared with actual values obtained in high-speed flight ($R_{\\text{effective}} \\cong 10$ to $30 \\times 10^6$).", "timestamp": "2026-07-19T18:27:14.032850+00:00"}
{"citation_id": "19930091707", "source_url": "https://ntrs.nasa.gov/api/citations/19930091707/downloads/19930091707.pdf", "page_number": 20, "total_pages": 20, "image_filename": "19930091707_p20.jpg", "text": "Positive directions of axes and angles (forces and moments) are shown by arrows\n\n| Axis |  | Force (parallel to axis) symbol | Moment about axis |  |  | Angle |  | Velocities |  |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| Designation | Symbol |  | Designation | Symbol | Positive direction | Designation | Symbol | Linear (component along axis) | Angular |\n| Longitudinal<br>Lateral<br>Normal | $X$<br>$Y$<br>$Z$ | $X$<br>$Y$<br>$Z$ | Rolling<br>Pitching<br>Yawing | $L$<br>$M$<br>$N$ | $Y \\longrightarrow Z$<br>$Z \\longrightarrow X$<br>$X \\longrightarrow Y$ | Roll<br>Pitch<br>Yaw | $\\phi$<br>$\\theta$<br>$\\psi$ | $u$<br>$v$<br>$w$ | $p$<br>$q$<br>$r$ |\n\nAbsolute coefficients of moment\n$C_l = \\frac{L}{qbS}$ (rolling)\n$C_m = \\frac{M}{qcS}$ (pitching)\n$C_n = \\frac{N}{qbS}$ (yawing)\n\nAngle of set of control surface (relative to neutral position), $\\delta$. (Indicate surface by proper subscript.)\n\n**4. PROPELLER SYMBOLS**\n\n$D$, Diameter\n$p$, Geometric pitch\n$p/D$, Pitch ratio\n$V$, Inflow velocity\n$V_s$, Slipstream velocity\n$T$, Thrust, absolute coefficient $C_T = \\frac{T}{\\rho n^2 D^4}$\n$Q$, Torque, absolute coefficient $C_Q = \\frac{Q}{\\rho n^2 D^5}$\n\n$P$, Power, absolute coefficient $C_P = \\frac{P}{\\rho n^3 D^5}$\n$C_s$, Speed-power coefficient $= \\sqrt[5]{\\frac{\\rho V^5}{P n^2}}$\n$\\eta$, Efficiency\n$n$, Revolutions per second, r.p.s.\n$\\Phi$, Effective helix angle $= \\tan^{-1} \\left( \\frac{V}{2\\pi r n} \\right)$\n\n**5. NUMERICAL RELATIONS**\n\n1 hp. = 76.04 kg-m/s = 550 ft-lb./sec.\n1 metric horsepower = 1.0132 hp.\n1 m.p.h. = 0.4470 m.p.s.\n1 m.p.s. = 2.2369 m.p.h.\n\n1 lb. = 0.4536 kg.\n1 kg = 2.2046 lb.\n1 mi. = 1,609.35 m = 5,280 ft.\n1 m = 3.2808 ft.\n\n-760-\n\nU. S. GOVERNMENT PRINTING OFFICE: 1938", "timestamp": "2026-07-19T18:27:17.282310+00:00"}
{"citation_id": "19930091714", "source_url": "https://ntrs.nasa.gov/api/citations/19930091714/downloads/19930091714.pdf", "page_number": 33, "total_pages": 36, "image_filename": "19930091714_p33.jpg", "text": "EFFECT OF COMPRESSIBILITY ON PROPELLERS IN TAKE-OFF AND CLIMBING RANGE 29\n\nTABLE II\nEXAMPLE 4, FIXED-PITCH PROPELLER, DESIGN B\n\n| 1 | 2 | 3 | 4 | 5 | 6 | 7 | 8 | 9 | 10 | 11 | 12 | 13 | 14 | 15 |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| $\\frac{V}{nD}$ | $C_{P_2}$ | $C_{T_2}$ | $\\frac{N}{N_{max}}$ for $C_{P_2}$ | $\\frac{V}{V_s}$ | $\\frac{C_{T_2}}{C_{T_2(at\\ stall)}}$ | $C_P$ | $C_T$ | $C_{P_2}$ | $C_{T_2}$ | $\\frac{N}{N_{max}}$ for $C_{P_2}$ | $\\frac{V}{V_{max}}$ | Air speed (m.p.h.) | Thrust (lb.) | Remarks |\n| | | | | | | $C_P \\left(\\frac{V}{V_s}=0.5\\right)$ | $C_T \\left(\\frac{V}{V_s}=0.5\\right)$ | | | | | | | |\n| 0 | 0.1129 | 0.1420 | 0.780 | 0.080 | 1.01 | 1.030 | : .020 | 0.1178 | 0.1448 | 0.796 | 0 | 0 | 1,214 | Blades stalled. |\n| .1 | .1105 | .1420 | .780 | .080 | 1.01 | 1.030 | 1.020 | .1160 | .1448 | .802 | .092 | 17.1 | 1,330 | |\n| .2 | .1080 | .1410 | .780 | .090 | 1.01 | 1.035 | 1.020 | .1130 | .1438 | .806 | .185 | 34.5 | 1,333 | |\n| .3 | .1065 | .1400 | .800 | .098 | 1.00 | 1.060 | 1.010 | .1130 | .1427 | .813 | .277 | 51.2 | 1,330 | |\n| .4 | .1040 | .1330 | .809 | .705 | .96 | 1.060 | 1.025 | .1100 | .1383 | .825 | .370 | 70.1 | 1,343 | |\n| .5 | .1000 | .1280 | .824 | .719 | .91 | 1.060 | 1.030 | .1060 | .1318 | .839 | .482 | 90.5 | 1,325 | Blades not stalled. |\n| .6 | .0940 | .1130 | .850 | .742 | .81 | 1.052 | 1.040 | .0990 | .1175 | .869 | .599 | 111.0 | 1,270 | |\n| .7 | .0840 | .0970 | .890 | .777 | .69 | 1.065 | 1.050 | .0918 | .1030 | .902 | .725 | 134.0 | 1,185 | |\n| .8 | .0740 | .0780 | .945 | .825 | .56 | 1.085 | 1.080 | .0825 | .0842 | .952 | .875 | 162.0 | 1,050 | |\n| .87 | .0680 | .0660 | 1.000 | .873 | .48 | 1.100 | 1.100 | .0747 | .0726 | 1.000 | 1.000 | 185.0 | 1,038 | |\n\nU. S. GOVERNMENT PRINTING OFFICE: 1939", "timestamp": "2026-07-19T18:27:38.828016+00:00"}
{"citation_id": "19930094549", "source_url": "https://ntrs.nasa.gov/api/citations/19930094549/downloads/19930094549.pdf", "page_number": 14, "total_pages": 76, "image_filename": "19930094549_p14.jpg", "text": "12 N.A.C.A. Technical Memorandum No. 867\n\nCHARACTER OF THE OSCILLATIONS\n\nThe character of the two types of oscillations is well known. The rapid oscillation is primarily that about the center of gravity of the airplane, while the slow oscillation, on the contrary, is one of the path described by the center of gravity in the vertical plane. The latter type of oscillation originates in the irregularities produced by any disturbance of the equilibrium of the forces applied to the airplane. It is possible by approximate methods to study each of these types of oscillations independently.\n\na) Rapid oscillation.- In studying the rapid oscillation by the method utilized by Munk - taking account of the loss in altitude of the airplane - it will be found that the restoring moment is not proportional to $dC_M/di$, but to\n\n$$\n\\frac{dC_M}{di} + \\frac{dC_Z}{di} \\frac{dC'_Z}{di'} \\frac{S'l'}{Sl} \\frac{gl'}{V^2} \\frac{1}{C_Z}\n$$\n\nand the damping is not proportional to\n\n$$\n\\frac{dC'_Z}{di'} \\frac{S'l'^2}{Sr^2}\n$$\n\nbut to\n\n$$\n\\frac{dC'_Z}{di'} \\frac{S'l'^2}{Sr^2} + \\frac{dC_Z}{di}\n$$\n\nThis explains why, so long as the maximum lift is not attained, the rapid oscillation could be stable even if the coefficient of static stability is negative.\n\nThe rapid oscillation, at small angles of attack, is strongly damped and the function $e^{at}p \\sin (bt + \\phi)$, as a result of the values of a and b, respectively, practically ceases to have an oscillatory character. If, for example:\n\n$$\na = -3.75 \\quad b = 1.56\n$$\n\nthen", "timestamp": "2026-07-19T18:27:41.949777+00:00"}
{"citation_id": "19930091655", "source_url": "https://ntrs.nasa.gov/api/citations/19930091655/downloads/19930091655.pdf", "page_number": 8, "total_pages": 22, "image_filename": "19930091655_p8.jpg", "text": "```markdown\n4\nREPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\n# APPARATUS AND METHOD\n\nThe experimental method employed in this investigation consisted in photographically recording with a suitable indicator the decrease in pressure following the injection of a definite quantity of liquid into a spherical bomb containing a gas at a known temperature and pressure. With the exception of a few minor modifications this apparatus was essentially as described in reference 23. The present arrangement is shown diagrammatically in figure 1. The essential parts were a bomb, a constant-temperature bath, a fuel-injection system, and an optical-type differential-pressure indicator. The stainless-steel bomb has a volume of 600 cubic centimeters and is provided with openings for the injection valve, the gas inlet and exhaust fittings, and the indicator.\n\nThe liquids used in the constant-temperature bath were S. A. E. 30 lubricating oil for the low temperatures and an approximately 1:1 mixture of sodium and potassium nitrates for the high temperatures. The bath temperature was kept within $\\pm 2^\\circ$ C. of the desired value by an automatic control.\n\nThe injection system delivered a single fuel charge of the desired weight upon the release of a trip mechanism. The injection valve was so constructed that fuel could be continuously circulated through it, thereby maintaining a constant fuel temperature of $49^\\circ \\pm 1.5^\\circ$ C. Three nozzles were used, all having equivalent orifice areas: A 13-orifice, a 2-impinging-jets, and a single-orifice nozzle with an orifice diameter of 0.050 inch (fig. 2). The fuel weights were varied by changing the injection pressures; the latter varied from about 194 to 600 atmospheres (2,850 to 9,000 pounds per square inch) for each nozzle. The requisite injection pressure for a given fuel quantity was determined just prior to a series of tests at each temperature. The injection period ranged between 0.002 and 0.006 second, depending upon the injection pressure used.\n\nThe high-pressure indicator employed in earlier work (references 6, 11, and 23) was altered to record small pressure differences by substituting a thin corrugated phosphor-bronze diaphragm for the heavy steel diaphragm and by providing a gas connection between the sealed chamber above the diaphragm and the bomb proper. The same initial pressure was applied to both sides of the diaphragm but, just before injection, a valve inserted in this connection was closed. This procedure permitted the subsequent pressure difference to actuate the indicator and thus to generate a trace of the pressure-difference variation with time on the film. This valve was opened again immediately after injection in order to minimize the interval within which the diaphragm remained deflected. A spark, recorded as a vertical line on certain records, marked the start of injection. This spark and injection start were synchronized by observing the spray with a neon-tube stroboscope actuated by the switching device on the injection system that ordinarily controlled the spark. The film drum was driven by a synchronous motor to provide the time scale.\n\n<!-- Image (136, 283, 874, 647) -->\n\nA, air gage.\nB, bimetallic strip.\nC, bomb.\nD, cam.\nE, check valve.\nF, clamping rings.\nG, clutch.\nH, condenser, 8 microfarad.\nI, contact points.\nJ, cooling coil.\nK, exhaust.\nL, film drum.\nM, from compressed-gas bottle.\nN, from high-pressure pump.\nO, fuel-circulating pressure gage.\nP, fuel high-pressure gage.\nQ, fuel reservoir.\nR, gear pump.\nS, heating coil.\nT, high-pressure reservoir.\nU, holder for bomb.\nV, indicator diaphragm.\nW, injection tube.\nX, injection valve.\nY, lamp.\nZ, lens.\nA', motor.\nB', oil bath.\nC', orifice, 0.020 inch.\nD', phase-changing gears.\nE', pivoted mirror.\nF', poppet valve.\nG', relay.\nH', resistance lamps.\nI', spark coil.\nJ', spark gap.\nK', quick-acting valves.\nL', spark-timing switch.\nM', stirrer.\nN', synchronous motor.\nO', thermometer.\nP', voltage 220 a. c.\nQ', voltage 230 d. c.\n\nFIGURE 1.—Diagrammatic sketch of the apparatus.\n```", "timestamp": "2026-07-19T18:27:50.169381+00:00"}
{"citation_id": "19930094552", "source_url": "https://ntrs.nasa.gov/api/citations/19930094552/downloads/19930094552.pdf", "page_number": 29, "total_pages": 32, "image_filename": "19930094552_p29.jpg", "text": "N.A.C.A. Technical Memorandum No. 864\n\nStrip$_{A}$\nA\nB\nC\nD\nShear x thickness T x S\n+8\n+6\n+4\n+2\n0\n-2\n-4\n-6\n-8\nkg/cm\n10\n20\n30\n40\n50\n60\n70\n80\n90\n100\nx\nSection at VI$\\beta$ Section at V$\\beta$\n\nStrip$_{A}$\nA\nB\nC\nD\nE\nShear x thickness T x S\n+8\n+6\n+4\n+2\n0\n-2\n-4\n-6\n-8\nkg/cm\n10\n20\n30\n40\n50\n60\n70\n80\n90\n100\nx\nSection at VI$\\beta$ Section at V$\\beta$\n\nStrip$_{A}$\nA\nB\nC\nD\nE\nShear x thickness T x S\n+8\n+6\n+4\n+2\n0\n-2\n-4\n-6\n-8\nkg/cm\n10\n20\n30\n40\n50\n60\n70\n80\n90\n100\nx\nSection at VI$\\beta$ Section at V$\\beta$\n\nLongitudinal stiffeners\nBending load\nShear times\nthickness\nat sections\n\nLongitudinal stiffeners.\nArching load with bulkheads\ne and f riveted.\n\nLongitudinal stiffeners.\nArching load with bulkheads\ne and f unattached.\n\nCylinder\nperimeter\nCylinder\nperimeter\nCylinder\nperimeter\n\nFigure 18.- Plot of the shear-times-thickness over the cylinder length and cylinder\nperimeter for P = 1,000 kg\n\nFig. 18", "timestamp": "2026-07-19T18:27:51.291039+00:00"}
{"citation_id": "19930091701", "source_url": "https://ntrs.nasa.gov/api/citations/19930091701/downloads/19930091701.pdf", "page_number": 1, "total_pages": 18, "image_filename": "19930091701_p1.jpg", "text": "CASE FILE\nCOPY\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\n\nREPORT No. 626\n\nTHE TRANSITION PHASE IN THE TAKE-OFF\nOF AN AIRPLANE\n\nBy J. W. WETMORE\n\n[Figure: Seal of the National Advisory Committee for Aeronautics]\n\nTHIS DOCUMENT ON LOAN FROM THE FILES OF\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\nLANGLEY AERONAUTICAL LABORATORY\nLANGLEY FIELD, HAMPTON, VIRGINIA\n\nRETURN TO THE ABOVE ADDRESS.\n\nREQUESTS FOR PUBLICATIONS SHOULD BE ADDRESSED\nAS FOLLOWS:\n\n1938\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n1724 - STREET, N. W.,\nWASHINGTON 25, D. C.\n\nFor sale by the Superintendent of Documents, Washington, D. C.\nSubscription price, $3.00 per year\nPrice 10 cents\n\nFILE COPY\nNO. 4", "timestamp": "2026-07-19T18:27:54.330280+00:00"}
{"citation_id": "19930091697", "source_url": "https://ntrs.nasa.gov/api/citations/19930091697/downloads/19930091697.pdf", "page_number": 3, "total_pages": 28, "image_filename": "19930091697_p3.jpg", "text": "REPORT No. 622\n\nA PHOTOGRAPHIC STUDY OF COMBUSTION AND KNOCK IN A SPARK-IGNITION ENGINE\n\nBy A. M. ROTHROCK and R. C. SPENCER\n\nLangley Memorial Aeronautical Laboratory\n\n49934—38—1", "timestamp": "2026-07-19T18:27:56.733321+00:00"}
{"citation_id": "19930094544", "source_url": "https://ntrs.nasa.gov/api/citations/19930094544/downloads/19930094544.pdf", "page_number": 24, "total_pages": 43, "image_filename": "19930094544_p24.jpg", "text": "22 N.A.C.A. Technical Memorandum No. 872\n\nplacement of these rings with respect to each other takes place. The circumferential forces are then determined from the components of the diagonal forces so determined. In contrast to the previously mentioned bending theory, this method is designated as the shear theory (references 4, 38, and 39).\n\nStress and bending measurements on the framing with definite conditions of loading can give an indication concerning the accuracy of the approximate methods discussed. A loading test of that kind was undertaken early in 1929 by the DVL with the framing of the LZ 127 in the hangar. The measurements were made on the weighed-off airship and the various loading conditions were obtained by shifting of the weights provided. The measurement of the stretch of longitudinal girders was mostly by the electro-acoustic method with Maihak strain gauges, tensions in wires were determined with the tensiometers developed by Luftschiffbau Zeppelin.\n\nFrom the great number of measurements taken, there are selected in figure 50 the stress measurements in the longitudinal girders over an airship's cross section approximately amidships for two significant conditions of loading. In the first case a large bending moment acts in conjunction with a small transverse force; in the second case a small bending moment in conjunction with a large transverse force. The curves a show the variation of the stresses measured in the longitudinal girders under these conditions of loading. Superimposed on these are three calculated curves b, c, d, which were obtained in accordance with the above-mentioned beam theory b under the hypothesis that only the longitudinals alone, c, that the longitudinals and all diagonals, and d, that the longitudinals and only the diagonals lying in the tension zone contribute to the moment of inertia. In the case of the diagonals a cooperation of the net stressing and outer cover is considered. The course of the curves shows that the stress distribution measured lies in general between the two lines b and c, and, indeed, agrees well with b in the compression zone and well with c in the tension zone. The line d is in good agreement with whole course.\n\nA somewhat expensive procedure for checking the stresses is the carrying out of static tests on models, which in their elastic properties duplicate the full size. Such model tests are in preparation at the DVL.", "timestamp": "2026-07-19T18:28:03.931572+00:00"}
{"citation_id": "19930094542", "source_url": "https://ntrs.nasa.gov/api/citations/19930094542/downloads/19930094542.pdf", "page_number": 74, "total_pages": 102, "image_filename": "19930094542_p74.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-19T18:28:06.016538+00:00"}
{"citation_id": "19930091693", "source_url": "https://ntrs.nasa.gov/api/citations/19930091693/downloads/19930091693.pdf", "page_number": 3, "total_pages": 13, "image_filename": "19930091693_p3.jpg", "text": "REPORT No. 618\n\nCOMPARATIVE FLIGHT\nAND FULL-SCALE WIND-TUNNEL MEASUREMENTS\nOF THE MAXIMUM LIFT OF AN AIRPLANE\n\nBy ABE SILVERSTEIN, S. KATZOFF, and JAMES A. HOOTMAN\n\nLangley Memorial Aeronautical Laboratory\n\n38572—38", "timestamp": "2026-07-19T18:28:07.673619+00:00"}
{"citation_id": "19930094543", "source_url": "https://ntrs.nasa.gov/api/citations/19930094543/downloads/19930094543.pdf", "page_number": 47, "total_pages": 50, "image_filename": "19930094543_p47.jpg", "text": "N.A.C.A. Technical Memorandum No. 873\nFigs. 17,18,19\n\n1,000\nHot-spot temperature, °C\n950\n900\n0\n50\n100\nCharge in percent\nFigure 17. - Effect of engine charge\n(semi-turbulent head)\nN = 1,250 r.p.m.\n\n1,100\nHot-spot temperature, °C\n1,050\n1,000\n950\n0\n50\n100\nCharge in percent\nFigure 18. - Effect of engine charge\n(turbulent head)\nN = 1,250 r.p.m.\n\n1,000\nHot-spot temperature, °C\n950\n900\nSpecific consumption, g/hp.-h\n300\n250\n0\n500\n1,000\n1,500\nr.p.m.\n10.0\n7.5\n5.0\n2.5\nHorsepower, hp.\nFigure 19. - Effect of r.p.m.\n(anti-turbulent head)", "timestamp": "2026-07-19T18:28:09.521583+00:00"}