PDF/ocr_dataset_101.jsonl

{"citation_id": "19930094535", "source_url": "https://ntrs.nasa.gov/api/citations/19930094535/downloads/19930094535.pdf", "page_number": 3, "total_pages": 17, "image_filename": "19930094535_p3.jpg", "text": "2\nN.A.C.A. Technical Memorandum No. 881\n\nIt was attempted to meet both of these defects of normal glass either by preventing formation of the large dagger-shaped splinters or by binding the splinters so as to prevent their scattering. These attempts led to the development of the so-called laminated safety glass of one or several layers.\n\nSINGLE-LAYER SAFETY GLASS\n\nIn the so-called hard-glass type, ordinary plate glass is subjected, by heat treatment, to large inner stresses which, when released during fracture, give rise to small crumbly fragments as shown in figure 2. These small fragments are generally not very sharp-edged, and so cannot lead to the same type of injuries as the splinters of normal glass; besides, there never remain in the frame those extremely dangerous knife-like splinters. Nevertheless, the occupants may be covered by numerous small fragments and thus possibly suffer light injuries. If the disk, in breaking, remains in the frame, the visibility through an obliquely set windshield becomes seriously impaired.\n\nMULTILAYER SAFETY GLASS\n\nBy the above type of safety glass is generally meant the type consisting of two glass sheets which are firmly attached by means of a transparent elastic layer, sandwiched in binding, being thus a combination of normal and artificial glass. The sandwiched layers are products of cellulose nitrates, cellulose acetates, and polymerization products, the oldest in use being cellulose nitrate. Disadvantages are the strong discoloring and formation of bubbles and haziness. Moreover, the binding force of the cellulose-nitrate layer weakens considerably by the dispersive precipitations in ageing, so that the protective action is diminished. Safety-glass panes with celluloid layer as binder, acquire after some time the appearance shown in figure 3. If safety glass with a cellulose acetate layer is used, there is likewise observed after weathering for some time, a slight discoloring which, however, only affects the appearance. Great progress has been made with the highly polymerized plastics which are not sensitive to weathering and moisture.", "timestamp": "2026-07-19T18:14:28.595060+00:00"}
{"citation_id": "19930094533", "source_url": "https://ntrs.nasa.gov/api/citations/19930094533/downloads/19930094533.pdf", "page_number": 34, "total_pages": 51, "image_filename": "19930094533_p34.jpg", "text": "N.A.C.A. Technical Memorandum No. 883\nFig. 1\n\n[Figure: Graph with y-axis labeled $T_2-T_a$ ranging from 0 to 50, and x-axis labeled \"Wind velocity (km/h)\" ranging from 0 to 800. The graph contains four curves labeled $W=41.5$, $W=27.2$, $W=15.9$, and $W=7.35$.]\n\nFigure 1.- Curves of constant heating giving excess temperature of a Balin antenna at altitude $0^0$.", "timestamp": "2026-07-19T18:14:34.837767+00:00"}
{"citation_id": "19930094557", "source_url": "https://ntrs.nasa.gov/api/citations/19930094557/downloads/19930094557.pdf", "page_number": 12, "total_pages": 20, "image_filename": "19930094557_p12.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-19T18:14:41.001613+00:00"}
{"citation_id": "19930094542", "source_url": "https://ntrs.nasa.gov/api/citations/19930094542/downloads/19930094542.pdf", "page_number": 62, "total_pages": 102, "image_filename": "19930094542_p62.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-19T18:14:54.973636+00:00"}
{"citation_id": "19930091715", "source_url": "https://ntrs.nasa.gov/api/citations/19930091715/downloads/19930091715.pdf", "page_number": 24, "total_pages": 30, "image_filename": "19930091715_p24.jpg", "text": "20\nREPORT NO. 640—NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nand 4-blade propellers should grow less at higher values of blade angle. The opposite appears to be true, however, when, on a more practical basis, the efficiency-curve envelope is plotted against $C_s$, as in figure 37.\n\nFigure 40 shows efficiency-curve envelopes for 2-blade and 3-blade propellers having the same solidity. The separation of the two curves is about 2 percent in this case. It should be pointed out, in connection with the results indicated in figures 36 and 40, that the blade thickness of the wide propeller was increased in proportion to its breadth (to maintain a constant thickness ratio and airfoil section) so that it is probably somewhat thicker than necessary for strength purposes. Part of the difference in the efficiency between the\n\nThe lack of data for the 3-blade propeller in the past has resulted in the use of empirical methods of making 3-blade and 4-blade propeller selections from 2-blade-propeller data. As the propeller with the greater number of blades absorbs more power, it is customary to use a certain fraction of the available power in computing the value of $C_s$ to be used with the 2-blade-propeller charts. This method is an approximation and will not give the optimum propeller diameter and blade angle for the design condition, although the difference may not be large. The convenience of this approximation has more than offset its faults and, now that data for 3-blade and 4-blade propellers are available, it is interesting to compare the ratios of the power absorbed\n\n<!-- Image (58, 297, 500, 437) -->\nFIGURE 39.—Efficiency-curve envelopes (against $V/nD$) for propellers having 2, 3, and 4 blades of Clark Y section.\n\n<!-- Image (547, 297, 927, 437) -->\nFIGURE 40.—Efficiency-curve envelopes for two propellers having the same solidity but a different number of blades.\n\n2-blade and the 3-blade propellers having the same solidity would undoubtedly be offset by thinning the 2-blade propeller, although such a procedure would, of course, change the airfoil section.\n\nA general comparison of the take-off qualities of the various propellers was not attempted as there was no basis of comparison that would have been entirely fair\n\nby the 2-blade, 3-blade, and 4-blade propellers. Such a comparison is shown in figure 41; in figure 42 is shown a similar comparison for the 2-blade and 3-blade propellers having the same solidity. The curves in figure 41 represent the mean of the curves for the Clark Y and R. A. F. 6 propellers, which were separated by a small amount.\n\n<!-- Image (58, 575, 500, 780) -->\nFIGURE 41.—Ratios of power absorbed by propellers having 2, 3, and 4 blades for the high-speed design condition.\n\n<!-- Image (547, 575, 927, 780) -->\nFIGURE 42.—Ratio of power absorbed by 2-blade and 3-blade propellers having the same solidity for the high-speed design condition.\n\nto all propellers. Any designer having a choice of two or more propellers can calculate their thrusts in the take-off range by the methods given in the appendix of this report. The designer knowing the design limitations peculiar to his particular problem will thus be able to make a satisfactory comparison.\n\nThroughout the $V/nD$ range shown in figure 41, the 2-blade propeller absorbs from 70 to 75 percent of the power absorbed by the 3-blade propeller and from 53 to 58 percent of the power absorbed by the 4-blade propeller. The 3-blade propeller 5868-R6 absorbs more power than the 2-blade propeller 37-3647, which has the same solidity. The ratio of their powers, $P_2/P_3$, varies (fig. 42) from 0.88 to 0.91.", "timestamp": "2026-07-19T18:14:57.367122+00:00"}
{"citation_id": "19930094534", "source_url": "https://ntrs.nasa.gov/api/citations/19930094534/downloads/19930094534.pdf", "page_number": 8, "total_pages": 21, "image_filename": "19930094534_p8.jpg", "text": "N.A.C.A. Technical Memorandum No. 828 7\n\nTwo 25 x 1 tubes with M 14 x 1.5 thread, once with and once without check nut disclosed no sign of failure at 100,000 stress reversals.\n\nBy the same method were investigated: two 40 x 1 tubes with M 12 x 1.5 thread under ±400 kg load, as well as two 50 x 1 tubes with M 14 x 1.5 thread under ±500 kg load. Neither case disclosed anything objectionable after 100,000 stress reversals.\n\nIn figure 12 two push rods of the same size (35 x 1) and identical attachment fitting (pin: 8 mm diameter) are compared. By foregoing all contraction the sample with inserted piece would appear even plumper than the new version.\n\nThe success with the push rods prompted the application of the method of contracting the tubes and rolling-in of the thread to other structural parts, such as adjustable struts of 55 x 2 chromium-molybdenum steel tubing. But, owing to the high stress of the tools, it is impossible to roll a thread smaller than M 23 x 1.5 in steel tubing. For rolled female thread in duralumin tubing, the lower limit is M 12 x 1.5.\n\nThe contracting process can be carried farther, so that smaller thread-cutting becomes possible. The Heinkel Company, for example, uses a turnbuckle shrunk from 15 x 1 duralumin tubing to 7.9 mm outside diameter and fitted with M 5 female thread. Its failing load was 892 kg, according to tensile tests. The strength of rolled thread is, as is known, 15 percent higher than the thread cut by conventional method.\n\nTools similar to those used for contraction take care of any necessary flanging. One such flange pressed on to a 30 x 1 duralumin tube for the purpose of holding a spring withstood a failing load of 2,350 kg. The break occurred in the unstrained tube, figure 13.\n\nSuch a flange can be continued and put on the inside, as illustrated in the handle, figure 13. The fluting also was effected by drawplate. Following the contraction and calibrating of the fitting end, the grip portion is flanged, contoured and flattened to conform to the turning motions of the hand.\n\nExpansions and contractions of every kind, symmetri-", "timestamp": "2026-07-19T18:14:58.804428+00:00"}
{"citation_id": "19930091716", "source_url": "https://ntrs.nasa.gov/api/citations/19930091716/downloads/19930091716.pdf", "page_number": 11, "total_pages": 24, "image_filename": "19930091716_p11.jpg", "text": "NEGATIVE THRUST AND TORQUE OF SEVERAL FULL-SCALE PROPELLERS\n7\n\n<!-- Image (124, 129, 842, 852) -->\n\nFIGURE 7.—Negative thrust and torque coefficients for propeller 5868-R6, R. A. F. 6 section, 3 blades.\n\n80778-38-2", "timestamp": "2026-07-19T18:15:00.321309+00:00"}
{"citation_id": "19930094559", "source_url": "https://ntrs.nasa.gov/api/citations/19930094559/downloads/19930094559.pdf", "page_number": 1, "total_pages": 16, "image_filename": "19930094559_p1.jpg", "text": "FILE COPY\nNO 4\n\nTECHNICAL MEMORANDUMS\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nTHIS DOCUMENT ON LOAN FROM THE FILES OF\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\nLANGLEY AERONAUTICAL LABORATORY\nLANGLEY FIELD, HAMPTON, VIRGINIA\n\nRETURN TO THE ABOVE ADDRESS.\nREQUESTS FOR PUBLICATIONS SHOULD BE ADDRESSED\nAS FOLLOWS:\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n1512 H STREET, N. W.\nWASHINGTON 25, D. C.\n\nNo. 857\n\nINVESTIGATION OF IGNITION AND COMBUSTION PROCESSES OF\nDIESEL ENGINES OPERATING WITH TURBULENCE AND\nAIR-STORAGE CHAMBERS\n\nBy Hans Petersen\n\nForschung auf dem Gebiete des Ingenieurwesens\nVol. 8, No. 6, November-December 1937\n\nWashington\nApril 1938", "timestamp": "2026-07-19T18:15:07.484263+00:00"}
{"citation_id": "19930094554", "source_url": "https://ntrs.nasa.gov/api/citations/19930094554/downloads/19930094554.pdf", "page_number": 37, "total_pages": 37, "image_filename": "19930094554_p37.jpg", "text": "N.A.C.A. Technical Memorandum No. 862\nFigs. 27,28,29,31\n\nFigure 27.\nAnnealing temp.\n$\\sigma_{\\text{Bl}}/\\sigma_{\\text{B}}$, %\n$\\sigma_{\\text{B}}$, kg/mm$^2$\n630°\n500°\n350°\n250°\n100°\nNot annealed\n\nFigure 28.\nStrength, kg/mm$^2$\nCorrosion period, months\n$\\sigma_{\\text{B}}$\n$\\sigma_{\\text{BLe}}$\n\nFigure 29\nStrength, kg/mm$^2$\nCorrosion period, months\n$\\sigma_{\\text{B}}$\n$\\sigma_{\\text{BLe}}$\n\nFigure 31.\nStrength, kg/mm$^2$\nCorrosion period, months\n$\\sigma_{\\text{B}}$\n$\\sigma_{\\text{BLe}}$", "timestamp": "2026-07-19T18:15:13.502041+00:00"}
{"citation_id": "19930094549", "source_url": "https://ntrs.nasa.gov/api/citations/19930094549/downloads/19930094549.pdf", "page_number": 4, "total_pages": 76, "image_filename": "19930094549_p4.jpg", "text": "2\nN.A.C.A. Technical Memorandum No. 867\n\nand have, moreover, the relation\n\n$$V = \\sqrt{u^2 + v^2 + w^2} = u$$\n\nThe projections of the instantaneous rotational velocity $\\Omega$ on the three axes are $p$, $q$, $r$, constituting the rolling, pitching, and yawing angular velocities. The rolling is positive if it tends to lower the right wing and raise the left; the pitching is positive if it tends to nose the airplane down, while the yawing is positive if it tends to turn the airplane toward the left. The controls corresponding to these rotations are, respectively, the ailerons, elevator, and rudder, their deflections $\\alpha$, $\\beta$, $\\gamma$ being positive when they tend to turn the airplane in the positive sense of the rotations.\n\nThe position of the system of axes $OXYZ$ fixed to the airplane, will be determined as a function of a system of spatially fixed axes $OX_oY_oZ_o$ by means of the three angular displacements defined in figure 2. The derivatives of these angles\n\n$$\\frac{d\\phi}{dt} = \\phi'$$\n\n$$\\frac{d\\theta}{dt} = \\theta'$$\n\n$$\\frac{d\\psi}{dt} = \\psi'$$\n\nare connected with the angular velocities $p$, $q$, $r$ by the geometric relations\n\n$$p = \\phi' \\cos \\theta - \\psi' \\cos \\phi \\sin \\theta$$\n\n$$q = \\theta' + \\psi' \\sin \\phi$$\n\n$$r = \\phi' \\sin \\theta + \\psi' \\cos \\phi \\cos \\theta$$", "timestamp": "2026-07-19T18:15:19.443573+00:00"}
{"citation_id": "19930093641", "source_url": "https://ntrs.nasa.gov/api/citations/19930093641/downloads/19930093641.pdf", "page_number": 8, "total_pages": 47, "image_filename": "19930093641_p8.jpg", "text": "6\n\narrangement were parallel to the fuselage reference axis, at an angle of -4.6° with the wing chord line, whereas the thrust axes for the enclosed-engine arrangements lie along the wing chord line. The difference was tolerated to aid in a clean design for the extension-shaft arrangements.\n\nThe extension shafts for the enclosed-engine arrangements were supported by tubular housings 4 inches in diameter which were bolted to either the front or the rear spar of the model.\n\nWood fairings 8 inches in diameter were placed concentrically over the 4-inch housings for some of the tests to simulate air-cooled engine nacelles for the case of a hypothetical 100-ton airplane.\n\nFour 3-blade aluminum-alloy propellers 39 inches in diameter were used throughout the tests; the dimensions of the blades are given in figure 8. Blade settings are given with reference to the 0.75 R station.\n\nTESTS\n\nPower-off measurements of forces and pitching moments were made for all the test arrangements over an angle-of-attack range from zero lift through the stall at an air speed of about 60 miles per hour. Scale effects on the over-all airplane drag and on the drag of the radiators, spinners, nacelles, and extension shafts were obtained in the low angle-of-attack range at air speeds from 30 to 120 miles per hour. Tests of the model with a bare wing (without nacelles, extension shafts, etc.) were made twice during the investigation to isolate the effects of suspected variations in the smoothness of the wing surface. Support tares and interferences were measured over the test range of tunnel speeds and angles of attack.\n\nThe nature and the spread of the wing stall for the cases of the wing-nacelle model and of the wing alone were observed by means of wool tufts glued to the upper wing surface.\n\nIn addition to the usual balance readings of force and moment, the power-on tests included measurements of electrical input to the motors and of propeller speed.", "timestamp": "2026-07-19T18:15:24.855893+00:00"}
{"citation_id": "19930094552", "source_url": "https://ntrs.nasa.gov/api/citations/19930094552/downloads/19930094552.pdf", "page_number": 17, "total_pages": 32, "image_filename": "19930094552_p17.jpg", "text": "N.A.C.A. Technical Memorandum No. 864 15\n\nof the shear within a bay is observed in all cases.\n\nIf the shears at the main spars determined from tests (fig. 20) are compared with those determined from the statically indeterminate computations, it is found that the test gives a somewhat larger concentration of the shear in the neighborhood of an applied force.\n\nb) Transverse stress distribution.- The distribution of the transverse or circumferential stresses over the cylinder length and cylinder circumference is shown on figures 21 and 22. It was necessary to plot the ring stresses in the arching loading condition with unattached bulkheads e and f to a smaller scale since their maximum values were twenty times as great as the maximum values under the arching load with attached rings. In the polar coordinate representation, the compressive peripheral stresses are plotted outward in order that the radial loads produced by them in the stiffeners may be shown more clearly.\n\nUnder the bending loading condition, transverse stresses of importance occur only in bay V (between the reinforced bulkheads). The sign of the stresses does not change over the cylinder length. Under the arching loading condition, the transverse stresses become very small if the rings e and f are attached to the skin, and the stresses change sign several times over the cylinder length. If bulkheads e and f are unattached, the transverse stresses in bay V assume very large values, which are positive on the loading side (near ring f) and negative on the side of ring e. In the remaining bays the peripheral stresses remain insignificantly small.\n\nThe transverse stresses are obtained through the integration,\n\n$$\n\\sigma_y = \\sigma_{yo} - \\int_0^y \\frac{\\partial \\tau}{\\partial x} \\, dy.\n$$\n\nThe derivative $\\frac{\\partial \\tau}{\\partial x}$ on which the distribution of the magnitude of the peripheral stress essentially depends is obtained from equations (1) and (2). Differentiating equation (2) with respect to x, we obtain\n\n$$\n\\left( \\frac{\\partial \\tau}{\\partial x} \\right)_b - \\left( \\frac{\\partial \\tau}{\\partial x} \\right)_a = - \\frac{1}{s} \\frac{d^2 P_x}{d x^2}\n$$", "timestamp": "2026-07-19T18:15:30.301497+00:00"}
{"citation_id": "19930094557", "source_url": "https://ntrs.nasa.gov/api/citations/19930094557/downloads/19930094557.pdf", "page_number": 13, "total_pages": 20, "image_filename": "19930094557_p13.jpg", "text": "N.A.C.A. Technical Memorandum No. 859\nFigs.1,2,3,11\n\nLeads to 6\ncomponent\nbalance\nExit cone\nModel\nEntrance\ncone\nPivoted ring\nscreen\nSupplementary\ncone\n\nFigure 1.- Spinning test setup in the\n1.2 m wind tunnel schemat-\nical plan.\n\n[Figure: Graph showing Angle between tip and axial speed vs. Distance from jet axis. Curves labeled 1, 2, 3.]\nDistance from jet axis.\nFigure 2.- Radial tip speed dis-\ntribution with a screen\nhaving 40 percent solidity for\nthree different screen speeds.\n\n[Figure: Graph showing Angle between peripheral and axial velocity vs. Distance from jet axis. Curves labeled 4 screens, 3 screens, 2 screens.]\nFigure 3.- Radial tip speed distribution for\ndifferent screen speeds.\n\n[Figure: Graph showing Moment about jet axis vs. $\\alpha$. Curves labeled Without screen, With, Rotating airfoil-fixed jet, Rotating jet-fixed airfoil, $\\alpha=0^\\circ$, $\\alpha=10^\\circ$, $\\alpha=20^\\circ$, $\\alpha=30^\\circ$, $\\alpha=40^\\circ$, $\\alpha=50^\\circ$ (Wing tip).]\nFigure 11.- Spinning moments of airfoil section M5.", "timestamp": "2026-07-19T18:15:31.534883+00:00"}
{"citation_id": "19930094538", "source_url": "https://ntrs.nasa.gov/api/citations/19930094538/downloads/19930094538.pdf", "page_number": 24, "total_pages": 43, "image_filename": "19930094538_p24.jpg", "text": "22 N.A.C.A. Technical Memorandum No. 878.\n\nFor the range from end of test range to failure, the factor of rise was figured at half as great as in the test range for the purpose of allowing for the fact that, due to the effect of the webs, the bulkheads become increasingly harder against growing local indentations. The indentations may be seen in figures 24 and 25.\n\nThe cylinders Nos. III and IV were computed on the basis of the complete and of the incomplete tension field, according to the method described in section III, 1, while the method of Ebner-Heck was checked also. For the incomplete tension field the free values $R_0$ and $\\delta_z$ were so chosen that the discrepancies of the theoretical from the experimental figures at the angles of twist and failing strengths (explained in section IV) became minimum. The figures are appended in table IV.\n\nb) Shell stresses.- All stress changes treated in the following, relate to the stages of torque cited under III, la. The changes in the principal tensile stress $\\sigma_{1m}$ obtained for the panel center between stringers, are plotted against x in figure 15; the section of the wrinkles is indicated below it. The test values on cylinder III (thin skin, many wrinkles) are substantially the same at the various points of the wrinkles, whereas at the highest and lowest points of the wrinkles on cylinder No. IV, they are maximum and disclose minimum values at the turning points. On cylinder No. III, the test values are considerably scattered around a mean value approximately constant over the length. The theoretical values, constant over x, according to all methods, lie on the average in the center of the scattered experimental values, thus confirming the aspect of a complete tension field in the center of the panel.\n\nThe bending moments $M_b$ (fig. 16), computed from the test data with the aid of equation (28), follow over x, according to a kind of bending moment line for uniformly distributed load and two-end restraint. The curves for both the complete and incomplete tension field are coincident and yield excessive values for cylinder III. For cylinder IV, the agreement is better although a few experimental values diverge in the vicinity of x = t.\n\nAs regards the compressive stresses $\\sigma_x$ in the stringers (fig. 17), the test values in the center between bulkheads are lower than at the bulkhead points x = 0 and x = t. On cylinder No. III, the curve computed for incomplete tension field runs through the scatter range of", "timestamp": "2026-07-19T18:15:42.458016+00:00"}
{"citation_id": "19930094549", "source_url": "https://ntrs.nasa.gov/api/citations/19930094549/downloads/19930094549.pdf", "page_number": 5, "total_pages": 76, "image_filename": "19930094549_p5.jpg", "text": "N.A.C.A. Technical Memorandum No. 867\n\nII. LONGITUDINAL MOTION\n\nMethod Used\n\nThe equations of motion referred to axes fixed to the airplane are, for nonuniform flight (fig. 3)*\n\n$$\n\\left.\n\\begin{aligned}\nT + X + P \\sin \\theta &= \\frac{P}{g} \\left( \\frac{du}{dt} + qv \\right) \\\\\nZ - P \\cos \\theta &= \\frac{P}{g} \\left( \\frac{dv}{dt} - qu \\right) \\\\\nM + Ts &= \\frac{Pr^2}{g} \\frac{dq}{dt}\n\\end{aligned}\n\\right\\} \\quad \\text{(I)}\n$$\n\nto which should be added the relation\n\n$$\nq = \\frac{d\\theta}{dt}\n$$\n\nThe equations may be written:\n\n$$\n\\left.\n\\begin{aligned}\n\\frac{du}{dt} &= f_1(u, w, q, \\theta) \\\\\n\\frac{dw}{dt} &= f_2(u, w, q, \\theta) \\\\\n\\frac{dq}{dt} &= f_3(u, w, q, \\theta) \\\\\n\\frac{d\\theta}{dt} &= f_4(u, w, q, \\theta)\n\\end{aligned}\n\\right\\} \\quad \\text{(II)}\n$$\n\nAt the instant $t_0$, let the values satisfying the system be $u_0, w_0, q_0, \\theta_0$, and let us give the airplane an initial disturbance, so that at the instant $t_0 + \\delta t$ the variables have the values $u_0 + \\delta u, w_0 + \\delta w$, etc. We may replace the four variables by their increments $\\delta u, \\delta w, \\delta q, \\delta \\theta$ with respect to their initial values. This\n\n---\n\n* In the section on longitudinal motion, $r$ denotes the radius of gyration with respect to OY.", "timestamp": "2026-07-19T18:15:55.551176+00:00"}
{"citation_id": "19930091716", "source_url": "https://ntrs.nasa.gov/api/citations/19930091716/downloads/19930091716.pdf", "page_number": 12, "total_pages": 24, "image_filename": "19930091716_p12.jpg", "text": "8\nREPORT NO. 641—NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\n$$nD/V$$\n0 .2 .4 .6 .8 1.0 1.2 1.4 1.6 1.8\n\n$$T_c$$\n-0.02\n-0.04\n-0.06\n-0.08\n-0.10\n-0.12\n-0.14\n\nBlade angle at 0.75 R.\n10°\n5°\n0°\n\n$$Q_c$$\n.008\n.004\n.002\n0\n-.002\n-.004\n-.006\n\n0°\n5°\n10°\n\n$$Q_c$$\n-.0010\n-.0015\n-.0020\n-.0025\n-.0030\n-.0035\n\n$$Q_c$$ director line---\n\nFIGURE 8.—Negative thrust and torque coefficients for propeller 5868-R6 at small blade-angle settings. R. A. F. 6 section, 3 blades.", "timestamp": "2026-07-19T18:16:04.070962+00:00"}
{"citation_id": "19930094564", "source_url": "https://ntrs.nasa.gov/api/citations/19930094564/downloads/19930094564.pdf", "page_number": 1, "total_pages": 16, "image_filename": "19930094564_p1.jpg", "text": "[Stamp: FILE COPY NO 5]\n\nTECHNICAL MEMORANDUMS\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nTHIS DOCUMENT ON LOAN FROM THE FILES OF\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\nLANGLEY AERONAUTICAL LABORATORY\nLANGLEY FIELD, HAMPTON, VIRGINIA\n\nNo. 852\n\nRETURN TO THE ABOVE ADDRESS.\nREQUESTS FOR PUBLICATIONS SHOULD BE ADDRESSED\nAS FOLLOWS:\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n1512 H STREET, N. W.\nWASHINGTON 25, D. C.\n\nSYSTEMATIC AIRFOIL TESTS\nIN THE LARGE WIND TUNNEL OF THE DVL\nBy H. Doetsch and M. Kramer\n\nLuftfahrtforschung\nVol. XIV, No. 10, October 12, 1937\nVerlag von R. Oldenbourg, München und Berlin\n\nWashington\nMarch 1938", "timestamp": "2026-07-19T18:16:05.113941+00:00"}
{"citation_id": "19930094542", "source_url": "https://ntrs.nasa.gov/api/citations/19930094542/downloads/19930094542.pdf", "page_number": 63, "total_pages": 102, "image_filename": "19930094542_p63.jpg", "text": "N.A.C.A. Technical Memorandum No. 874\n\nFigs. 39, 40\n\n[Figure: Graphs showing downwash angle and dynamic pressure distribution for α = 8° and α = 12°, with multiple curves plotted against x/R from -3.0 to 3.0, and y/R values labeled on right axis including 2.133, 1.600, 1.067, 0.533, -0.533, -1.067, -1.600, -2.133. Shaded regions indicate areas of interest. Axes labeled d' and (t₂-t₁)/v. Circular annotation highlights central region of plots.]\n\nFigure 39. α = 8°.\n\n[Figure: Same structure as above but for α = 12°, with similar axes and shaded regions. Circular annotation again highlights central area.]\n\nFigure 40. α = 12°.\n\nDownwash angle and dynamic pressure distribution.", "timestamp": "2026-07-19T18:16:06.125445+00:00"}
{"citation_id": "19930094534", "source_url": "https://ntrs.nasa.gov/api/citations/19930094534/downloads/19930094534.pdf", "page_number": 9, "total_pages": 21, "image_filename": "19930094534_p9.jpg", "text": "8 N.A.C.A. Technical Memorandum No. 882\n\ncal or asymmetrical with the axis of the original tube are feasible.\n\nII. SKIN RIVETING\n\nThe Customary Riveting Methods,\n\nRange of Application, and Special Tools\n\nThe demand for perfectly smooth surfaces, for aerodynamical reasons, has led to a multiplicity of riveting methods. After sufficient experience had been accumulated, the need for some kind of standardization in riveting methods and shop practice became imperative.\n\nIn the appended tabulation, figure 14, the riveting methods used by Heinkel are, so far as they have not as yet been included in the standardization, indicated by the letters x, y, and z.\n\nThe most important types of riveting for covering large surfaces are: mushroom head (F) and flat, countersunk riveting with dimpled sheets (F S x). As regards strength, both are about equal. Economy and quality of workmanship depend upon type and degree of accessibility. The following data serve for their appraisal:\n\n1. Mushroom-Head Riveting (F)\n\nThis lends itself to hand and machine operation, figure 15. To assure a satisfactory closing head, rivets of finer than standard length graduation are used; cutting-off therefore is usually inevitable. Moreover, since very short rivets are difficult to insert, the patented mushroom-head rivet-set tool (fig. 15) is employed for machine operations; it combines the rivet shank-cutting tool, the pulling-tight or dimpling tool, and the dolly; hence the loss of time due to changing of tools is a minimum. After the rivet is pulled tight, the free shank is cut to its exact length by the laterally disposed tongs.\n\nHowever, this type of machine-riveting requires adequate clearance in rivet axis direction, since any deviation of the hammer axis toward the rivet axis, say due to sloping or wrong shape of the snap, results in defective", "timestamp": "2026-07-19T18:16:08.157158+00:00"}
{"citation_id": "19930094557", "source_url": "https://ntrs.nasa.gov/api/citations/19930094557/downloads/19930094557.pdf", "page_number": 14, "total_pages": 20, "image_filename": "19930094557_p14.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-19T18:16:09.949587+00:00"}
{"citation_id": "19930094552", "source_url": "https://ntrs.nasa.gov/api/citations/19930094552/downloads/19930094552.pdf", "page_number": 18, "total_pages": 32, "image_filename": "19930094552_p18.jpg", "text": "16 N.A.C.A. Technical Memorandum No. 864\n\nDifferentiation of equation (1) for the sheet bays with respect to x gives\n\n$$\n\\frac{\\partial^2 \\tau}{\\partial x \\partial y} = - \\frac{\\partial^2 \\sigma_x}{\\partial x^2}\n$$\n\nand by integrating over the periphery, there is obtained\n\n$$\n\\frac{\\partial \\tau}{\\partial x} = \\left( \\frac{\\partial \\tau}{\\partial x} \\right)_0 - \\int_0^y \\frac{\\partial^2 \\sigma_x}{\\partial x^2} \\, dy\n$$\n\nThe derivatives $\\frac{\\partial \\tau}{\\partial x}$ therefore depend on the curvatures of the $\\sigma_x(x)$ or $F_x(x)$ curves. The greatest transverse stresses in the cross sections will thus be expected to occur in those sections in which the $F_x(x)$ or $\\sigma_x(x)$ curves are most sharply curved. If the bulkheads e and f are not attached to the skin the breaks (fig. 17) of the $F_x(x)$ curves at the bulkhead positions vanish and in their place, particularly between ring f and the center of bay V, there appear curve portions of very sharp curvature. This is the mathematical explanation for the high values of the transverse stresses. With unattached bulkheads no shears can be transferred locally. For this reason, the shear behind ring f in bay V (fig. 18) must sharply increase starting from zero. In the rear half of bay V, the shear decreases rapidly at first then more slowly. The internal equilibrium of the skin for such a stress distribution requires very high transverse stresses with change of sign in panel V.\n\nc) The loading of the stiffener system.- The distribution of the longitudinal stresses along the cylinder is shown on figures 8 and 13 and that of the bulkhead stresses on figures 10 and 15. In the latter two figures are plotted the mean values of the stresses measured at the bulkheads and no effective supporting width of skin has been taken into account. The rings could be loaded tangentially through the transfer of the shear differences at rows of rivets between the bulkheads and skin, or radially through transferred peripheral forces of the peripheral", "timestamp": "2026-07-19T18:16:21.884701+00:00"}
{"citation_id": "19930094535", "source_url": "https://ntrs.nasa.gov/api/citations/19930094535/downloads/19930094535.pdf", "page_number": 4, "total_pages": 17, "image_filename": "19930094535_p4.jpg", "text": "N.A.C.A. Technical Memorandum No. 881\n\nIn the above binding materials the silicare glass is always used as a thin cover sheet, the object being to save weight since the specific weight of silicate glass is 2.6 as compared with 1.4 of the artificial materials.\n\nTRANSPARENT PLASTIC RESINS\n\nThe transparent plastics in use are celluloid, cellulose acetate, mixtures of polymers and esters of polyacrylic acid. In the following paragraphs these materials will be considered in more detail.\n\nA. Celluloid\n\nThe starting material for celluloid is cellulose nitrate, which is developed according to the scheme of figure 4. The production of celluloid is as follows: The cellulose nitrate is dissolved in alcohol and camphor, gelatinized in kneading machines, during which process a part of the alcohol evaporates so that a doughy mass results. The latter is again strongly kneaded under heated rollers, colors in some cases being added. There is thus obtained a homogeneous mass in the form of thin, now completely gelatinized, plates. The latter are put up in layers and compressed while still warm into blocks and, to remove entrapped bubbles, are baked for several hours at high temperatures. The blocks are then cut up by suitable machines into the desired shapes (sheets, plates, rods, etc.). The alcohol still remaining (up to about 15 percent) is driven off by drying.\n\nBy further working of the rough material, as, for example, rolling and pressing, the final product is obtained. Plates which are to receive a high surface polish are pressed between highly polished, chromed layers.\n\nPrincipal Properties of Celluloid\n\nSpecific weight . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 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. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 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{"citation_id": "19930094564", "source_url": "https://ntrs.nasa.gov/api/citations/19930094564/downloads/19930094564.pdf", "page_number": 2, "total_pages": 16, "image_filename": "19930094564_p2.jpg", "text": "NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nTECHNICAL MEMORANDUM NO. 852\n\nSYSTEMATIC AIRFOIL TESTS\n\nIN THE LARGE WIND TUNNEL OF THE DVL*\n\nBy H. Doetsch and M. Kramer\n\nSUMMARY\n\nThe present report is a description of systematic tests at maximum lift on airfoils with and without split flap and of profile drag at low lift. The program included, respectively, the symmetrical and 2-percent camber N.A.C.A. airfoil sections 00, 24, and 230, with 9- to 21-percent thickness range. The maximum lift of the airfoil series without split flap was established for the entire practical flying range by comparing the DVL data with the findings from other wind tunnels. In order to obtain an opinion as to the suitability of the airfoils with flaps, the maximum-lift measurements were repeated on airfoils with split flaps.\n\nThe profile drag at low lift was arrived at by direct weighing and momentum measurements and, since the profiles were of unusual depth, extended to large Reynolds Numbers. It results in very carefully developed curves $c_{a_{max}}/c_{wp}(c_a = 0.1)$ with and without split flap, which as regards Reynolds Number correspond to actual flight conditions.\n\nI. INTRODUCTION\n\nAs the 5- by 7-meter wind tunnel of the DVL did not begin to operate until in the fall of 1935 (reference 1) only the utmost restrictions in the scope of the research program made it possible to catch up with other countries which were years ahead. For this reason only two airfoil\n\n*\"Systematische Profiluntersuchungen im grossen Windkanal der DVL,\" Luftfahrtforschung, vol. 14, no. 10, October 12, 1937, pp. 480-485.", "timestamp": "2026-07-19T18:16:45.953119+00:00"}
{"citation_id": "19930094544", "source_url": "https://ntrs.nasa.gov/api/citations/19930094544/downloads/19930094544.pdf", "page_number": 15, "total_pages": 43, "image_filename": "19930094544_p15.jpg", "text": "N.A.C.A. Technical Memorandum No. 872 13\n\nwire net inside the inherently stiff ring has been avoided. Instead, the whole cell is surrounded by a parachute-like wire net, which leads the side gas forces into the joints of the longitudinal girders (fig. 31).\n\nThe fact that the inherently stiff ring occupies some of the available gas space and thereby reduces the lift is always emphasized as an unfortunate disadvantage of such rings. To avoid this, it has been suggested that the ring be made as deep as possible and its inside filled with a special ring cell. However, this solution is accompanied by great structural difficulties and also results in an additional weight of cell material and valves, apart from the consideration that the increased surface of the whole cell installation involved in this solution causes increased gas loss.\n\nAlso, with respect to the spacing of the main rings, the newer airships differ very substantially. To minimize the ring and cell weights, it would be desirable to subdivide the gas space as little as possible. The size of the cells and therewith the main-ring spacing is, however, limited by the condition that the loss of lift in the event of the deflation of a cell, and the ensuing trim moment, may not exceed a definite maximum value. This maximum value depends upon what matter in the airship can be expended to offset the loss of lift and the trim of the ship with deflation of this cell. Besides this, a limitation of the cell size results from the requirement that the stressing of the framing with deflation of a cell may not be too unfavorable. The spacing of the main rings selected in the case of the \"Graf Zeppelin\" is 15 m. Between the main rings, two intermediate rings are placed (fig. 21). They serve to reduce the column length of the longitudinal girders to the most favorable figure of 5 m and also to provide a favorable angle of inclination for the shear wires. In the LZ 129, in spite of the large increase in the gas content, a cell length of 15.0 m, as well as the scheme of two intermediate rings, have been retained. Only amidships is the main ring spacing increased to 16.5 m. On the other hand, the wide main ring spacing in the \"Akron\" has been increased to 20 m amidships and to subdivide the the column length of the longitudinals three intermediate rings have been used (fig. 23). In the English constructions, R 100 and R 101, the intermediate rings have been entirely omitted and, instead, the main rings have been put close together (fig. 28). This resulted in a relatively", "timestamp": "2026-07-19T18:16:48.367925+00:00"}
{"citation_id": "19930091716", "source_url": "https://ntrs.nasa.gov/api/citations/19930091716/downloads/19930091716.pdf", "page_number": 13, "total_pages": 24, "image_filename": "19930091716_p13.jpg", "text": "NEGATIVE THRUST AND TORQUE OF SEVERAL FULL-SCALE PROPELLERS\n9\n\n<!-- Image (108, 78, 861, 905) -->\n\nFIGURE 9.—Negative thrust and torque coefficients for propeller 5868-R6, R. A. F. 6 section, 4 blades.", "timestamp": "2026-07-19T18:16:55.591023+00:00"}
{"citation_id": "19930094542", "source_url": "https://ntrs.nasa.gov/api/citations/19930094542/downloads/19930094542.pdf", "page_number": 64, "total_pages": 102, "image_filename": "19930094542_p64.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-19T18:16:59.335570+00:00"}
{"citation_id": "19930094549", "source_url": "https://ntrs.nasa.gov/api/citations/19930094549/downloads/19930094549.pdf", "page_number": 6, "total_pages": 76, "image_filename": "19930094549_p6.jpg", "text": "4 N.A.C.A. Technical Memorandum No. 867\n\nsubstitution, if the increments are assumed to be small, makes it possible to linearize the system of differential equations. We thus obtain:\n\n$$\n\\left.\n\\begin{aligned}\n\\frac{d(\\delta u)}{dt} &= \\frac{\\partial f_1}{\\partial u} \\delta u + \\frac{\\partial f_1}{\\partial w} \\delta w + \\frac{\\partial f_1}{\\partial q} \\delta q + \\frac{\\partial f_1}{\\partial \\theta} \\delta \\theta \\\\\n\\frac{d(\\delta w)}{dt} &= \\frac{\\partial f_2}{\\partial u} \\delta u + \\frac{\\partial f_2}{\\partial w} \\delta w + \\frac{\\partial f_2}{\\partial q} \\delta q + \\frac{\\partial f_2}{\\partial \\theta} \\delta \\theta \\\\\n\\frac{d(\\delta q)}{dt} &= \\frac{\\partial f_3}{\\partial u} \\delta u + \\frac{\\partial f_3}{\\partial w} \\delta w + \\frac{\\partial f_3}{\\partial q} \\delta q + \\frac{\\partial f_3}{\\partial \\theta} \\delta \\theta \\\\\n\\frac{d(\\delta \\theta)}{dt} &= \\frac{\\partial f_4}{\\partial u} \\delta u + \\frac{\\partial f_4}{\\partial w} \\delta w + \\frac{\\partial f_4}{\\partial q} \\delta q + \\frac{\\partial f_4}{\\partial \\theta} \\delta \\theta\n\\end{aligned}\n\\right\\} \\quad \\text{(III)}\n$$\n\nIn this new system the disturbances $\\delta u$, $\\delta w$, $\\delta q$, $\\delta \\theta$ are the variables and the partial derivatives are constants whose values are determined by the initial conditions. This system is linear and readily integrated, yielding the variables $\\delta u$, $\\delta w$, $\\delta q$, $\\delta \\theta$ as functions of the time after an initial disturbance. The integrated system is of the form:\n\n$$\n\\left.\n\\begin{aligned}\n\\delta u &= C_1 e^{\\lambda_1 t} + C_2 e^{\\lambda_2 t} + C_3 e^{\\lambda_3 t} + C_4 e^{\\lambda_4 t} \\\\\n\\delta w &= l_1 C_1 e^{\\lambda_1 t} + l_2 C_2 e^{\\lambda_2 t} + l_3 C_3 e^{\\lambda_3 t} + l_4 C_4 e^{\\lambda_4 t} \\\\\n\\delta q &= m_1 C_1 e^{\\lambda_1 t} + m_2 C_2 e^{\\lambda_2 t} + m_3 C_3 e^{\\lambda_3 t} + m_4 C_4 e^{\\lambda_4 t} \\\\\n\\delta \\theta &= n_1 C_1 e^{\\lambda_1 t} + n_2 C_2 e^{\\lambda_2 t} + n_3 C_3 e^{\\lambda_3 t} + n_4 C_4 e^{\\lambda_4 t}\n\\end{aligned}\n\\right\\} \\quad \\text{(IV)}\n$$\n\nIn the above system the four values of $\\lambda$ are the solutions of the characteristic equation of the system:\n\n$$\n\\left|\n\\begin{array}{cccc}\n\\frac{\\partial f_1}{\\partial u} - \\lambda & \\frac{\\partial f_1}{\\partial w} & \\frac{\\partial f_1}{\\partial q} & \\frac{\\partial f_1}{\\partial \\theta} \\\\\n\\frac{\\partial f_2}{\\partial u} & \\frac{\\partial f_2}{\\partial w} - \\lambda & \\frac{\\partial f_2}{\\partial q} & \\frac{\\partial f_2}{\\partial \\theta} \\\\\n\\frac{\\partial f_3}{\\partial u} & \\frac{\\partial f_3}{\\partial w} & \\frac{\\partial f_3}{\\partial q} - \\lambda & \\frac{\\partial f_3}{\\partial \\theta} \\\\\n\\frac{\\partial f_4}{\\partial u} & \\frac{\\partial f_4}{\\partial w} & \\frac{\\partial f_4}{\\partial q} & \\frac{\\partial f_4}{\\partial \\theta} - \\lambda\n\\end{array}\n\\right| = 0 \\quad \\text{(V)}\n$$", "timestamp": "2026-07-19T18:17:01.085782+00:00"}
{"citation_id": "19930094533", "source_url": "https://ntrs.nasa.gov/api/citations/19930094533/downloads/19930094533.pdf", "page_number": 36, "total_pages": 51, "image_filename": "19930094533_p36.jpg", "text": "N.A.C.A. Technical Memorandum No. 883\nFigs. 4,5,6\n\n<!-- Image (257, 102, 874, 318) -->\n\nFigure 4.- Effect of altitude on the power dissipation of the\nBadin anemometer having a temperature 25° higher\nthan the atmosphere.\n\n<!-- Image (288, 377, 814, 638) -->\n\nFigure 5.- Curve of energy required for de-icing in relation\nto initial temperature.\n\n<!-- Image (299, 666, 802, 902) -->\n\nFigure 6.- Curve of energy required for de-icing plotted against\nhorse power. (initial constant temperature)", "timestamp": "2026-07-19T18:17:03.136987+00:00"}
{"citation_id": "19930091714", "source_url": "https://ntrs.nasa.gov/api/citations/19930091714/downloads/19930091714.pdf", "page_number": 30, "total_pages": 36, "image_filename": "19930091714_p30.jpg", "text": "26 REPORT NO. 639—NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nsults in the following characteristics:\n\n$C_{t}=1.503$\n\n$\\frac{V}{nD}=0.9$\n\n$D=7.62 \\text{ ft.}$\n\n$\\frac{V}{V_{c}}=0.847$\n\n$\\frac{C_{P}}{C_{P}\\left(\\frac{V}{V_{c}}^{0.5}\\right)}=1.085$\n\nIn view of the change in the power correction factor incurred by the reduced diameter and tip speed, this propeller will only absorb $538 \\times 1.085 = 585 \\text{ hp}$. A third approximation using the average of the first and second values of $\\frac{C_{P}}{C_{P}\\left(\\frac{V}{V_{c}}^{0.5}\\right)}$ should result in approximately the correct answer; $\\frac{1.115 + 1.085}{2} = 1.100$. The third approximation results in the following characteristics designated “design B”:\n\n$\\text{hp.} = \\frac{600}{1.1} = 545 \\text{ (for design purpose).}$\n\n$C_{t}=1.497$\n\n$\\frac{V}{nD}=0.87$\n\n$\\beta=22.7^\\circ$\n\n$D=7.88 \\text{ ft.}$\n\nTip speed $= 978 \\text{ f. p. s.}$\n\n$\\frac{V}{V_{c}}=0.873$\n\n$\\frac{C_{P}}{C_{P}\\left(\\frac{V}{V_{c}}^{0.5}\\right)}=1.1 \\text{ (check).}$\n\nIn table II the thrust is computed for design B according to the following procedure:\n\n1. In column 1, values of $V/nD$ are assumed and, in addition, the design value for high speed is included.\n2. From figure 10 (reference 10), the low-tip-speed power coefficients $C_{P_2}$ are read following the line for a blade angle of $22.7^\\circ$.\n3. The corresponding thrust coefficient, $C_{T_2}$, is also read from figure 10 of reference 10.\n4. The ratio $N/N_{max}$ is computed from the relation $\\frac{N}{N_{max}} = \\sqrt{\\frac{C_{P_2(at\\ high\\ speed)}}{C_{P_2}}}$, assuming that the torque remains constant for small changes in rotational speed. This condition is substantially true for unsupercharged engines.\n5. The ratio $V/V_c$ is equal to $\\frac{N}{N_{max}} \\times 0.873$.\n6. The ratio $\\frac{C_{T_2}}{C_{T(at\\ static)}}$ is computed using $C_{T(at\\ static)} = 0.140$.\n7. $\\frac{C_P}{C_P\\left(\\frac{V}{V_c}^{0.5}\\right)}$ is read from figure 58 for different values of $V/V_c$ and $\\frac{C_{T_2}}{C_{T(at\\ static)}}$.\n8. $\\frac{C_P}{C_P\\left(\\frac{V}{V_c}^{0.5}\\right)}$ is also read from figure 58.\n9. Corrected values of power coefficient $C_{P_2}$ are computed, $C_{P_2} \\times \\frac{C_P}{C_P\\left(\\frac{V}{V_c}^{0.5}\\right)} = C_{P_2}$.\n10. Corrected values of thrust coefficient $C_{T_2}$ are computed in a similar manner.\n11. Corrected values of $N/N_{max}$ are computed using $C_{P_2}$.\n12. Values of $V/V_{max}$ are computed from the relation $\\frac{V}{V_{max}} = \\frac{\\left(\\frac{V}{nD}\\right) N}{\\left(\\frac{V}{nD}\\right)_{max} N_{max}}$.\n13. The air speed is computed from the relation $V/V_{max}$, knowing $V_{max}$.\n14. The thrust is computed from the relation $T = C_T \\frac{C_{P_2(at\\ high\\ speed)}}{C_{P_2}} \\times K$, where $K = \\rho n^2 D^4$.\n\nIf the method of reducing the blade angle is followed, to offset the increase in power coefficients from low to high tip speed, design A is used directly but it is necessary to determine the blade-angle reduction. The value of $C_{P_1}$ is divided by $\\frac{C_P}{C_P\\left(\\frac{V}{V_c}^{0.5}\\right)}$ to determine the $C_{P_2}$ corresponding to the low-tip-speed data: $\\frac{0.0653}{1.115} = 0.0585$. Unfortunately, this value is only the first approximation because the low-tip-speed thrust coefficient is likewise reduced, changing the value of $\\frac{C_{T_1}}{C_{T(at\\ static)}}$ to $\\frac{0.059}{0.14}$ or $0.42$. The value of $\\frac{C_P}{C_P\\left(\\frac{V}{V_c}^{0.5}\\right)}$ then becomes $1.105$. The second approximate value of power coefficient becomes $\\frac{0.653}{1.105} = 0.0592$. This $C_{P_2}$ defines the blade angle so the $C_{P_2}$ and $C_{T_2}$ can be read from figure 10 (reference 10) for different values of $V/nD$. The corrected thrust is then computed in the manner outlined in table II. No table is included for design A", "timestamp": "2026-07-19T18:17:05.798233+00:00"}
{"citation_id": "19930094534", "source_url": "https://ntrs.nasa.gov/api/citations/19930094534/downloads/19930094534.pdf", "page_number": 10, "total_pages": 21, "image_filename": "19930094534_p10.jpg", "text": "N.A.C.A. Technical Memorandum No. 832 9\n\nwork. On the other hand, in fuselage riveting, the continuously changing position of this axis is subordinate because the snap is correctly guided by the rivet head. Another advantage of this riveting method is that by forming the closing head as a flat, flush head it becomes practically impossible to spoil the rivet.\n\nFor driving mushroom rivets by hand the snap holding the rivet head can be the wrong shape, since it merely represents a shoulder of the dolly. For this reason, this method is applicable even in places not readily accessible; but it is expensive and should be used only in special cases. For instance, it takes less time to drive single mushroom rivets by hand than to use a rivet hammer for heading individual F S x.\n\nIn such cases a dolly with elastic mass is employed which positively prevents the snap from jumping off the rivet head (fig. 17).\n\n2. Flat, Countersunk Riveting with Dimpled Sheets (F S x)\n\nOnly machine riveting.- The process for this is illustrated in figure 18. Given ready accessibility on approximately upright surfaces, this method is about as fast as mushroom riveting. But for overhead or slanting use of the dolly, as in fuselage riveting, for example, the time involved and the danger of inferior workmanship are greater, since the forming of the closing head (F) cannot be watched as closely. In case of restricted accessibility, the F S x machine method is definitely superior; the requisite dollies are adaptable to almost any form and comparatively easy to fabricate. Figures 19 to 21 show various dollies for riveting in places not readily accessible.\n\nThe rivets are inserted from the outside, thus eliminating the irksome task of the \"threading.\" It is not advisable to use the lower limit of accessibility given in figure 14 any more than is absolutely necessary, because working with the necessarily thin and light dollies imposes added exertions.", "timestamp": "2026-07-19T18:17:09.532800+00:00"}
{"citation_id": "19930094551", "source_url": "https://ntrs.nasa.gov/api/citations/19930094551/downloads/19930094551.pdf", "page_number": 3, "total_pages": 18, "image_filename": "19930094551_p3.jpg", "text": "2 N.A.C.A. Technical Memorandum No. 865\n\nwidespread favor over the previously customary quadrangular flights for relative speed determinations and exact dynamic pressure calibrations. The pressure-head method is, as is known, based upon Bernoulli's fundamental equation:\n\n$$\nP_{\\text{total}} = P_{\\text{static}} + P_{\\text{dynamic}}\n$$\n\nwhere the total pressure $P_{\\text{total}}$ is obtained with a DVL total-head meter and the static pressure $P_{\\text{static}}$, by means of a static pressure head trailed several meters below the airplane. The two pressures yield as difference the true flight dynamic pressure. The conventional static pressure head heretofore used by the DVL weighs 2.4 kg while the tubing connecting with the airplane is a 20 m length of rubber hose of 8 mm outside and 3 mm inside diameter housed in wire netting. This instrument is not practical at speeds above 250 km/h, as the dynamic stability of the suspension cable becomes inferior at such speeds. The vibrations set up by gusts or irregularities in propeller slipstream no longer die out, but are propagated in direction of the hose end where the static pressure head itself is thrown into erratic oscillations of such amplitude that not only reading is falsified, but the airplane itself is endangered.\n\nAnother reason for not exceeding the cited speed of 250 km/h is that at higher speeds the head itself may be raised so high that it eventually finds itself in the wake of the airplane.\n\nII. TEST PROCEDURE\n\nIn order to arrive at a satisfactory solution of the problem, the DVL made systematic flight tests with a Heinkel He 70. The floor plates were removed during these tests to provide an uninterrupted view of the pressure head and of the tubing. Above the two openings we mounted two bearing loops which, after sighting the pressure head on a sector scale, permitted the reading of the angle of the pressure head below the airplane with respect to a body-fixed reference axis (fig. 1). The distance of the bearing-loop axes amounted to 1.57 m. The measuring accuracy of the bearing device was close enough to remain", "timestamp": "2026-07-19T18:17:12.733054+00:00"}
{"citation_id": "19930094535", "source_url": "https://ntrs.nasa.gov/api/citations/19930094535/downloads/19930094535.pdf", "page_number": 5, "total_pages": 17, "image_filename": "19930094535_p5.jpg", "text": "4\nN.A.C.A. Technical Memorandum No. 881\n\nImpact bending strength . . . 100-200 cm kg/cm²\nHeat resistance according\nto Martens, about . . . . . . . 40° C.\nExpansion coefficient . . 100 x 10⁻⁶\n\nB. Cellulose Acetates\n\nIn the year 1907 the gelatinizing of cellulose acetate\nwas successful, and it thereby became possible to work this\nmaterial in exactly the same manner and with the same appa-\nratus used in the manufacture of celluloid. Instead of al-\ncohol, however, a mixture of benzol and alcohol is used.\nThese cellulose nitrates, while not inflammable to the same\ndegree, have properties similar to those of celluloid. The\nproduction of cellulose acetate is shown schematically in\nfigure 4.\n\nPrincipal Properties of Cellulose Acetate\n\n| Property | Value |\n| :--- | :--- |\n| Specific weight | 1.3 |\n| Tensile strength, about | 500 kg/cm² |\n| Bending strength, about | 350 kg/cm² |\n| Elasticity modulus | 30,000-60,000 kg/cm² |\n| Impact bending strength | 100-200 cm kg/cm² |\n| Heat resistance according to Martens, about | 35° C. |\n| Expansion coefficient | 110 x 10⁻⁶ |\n\nThe artificial materials to be described next, namely,\nthe polymers, were first developed in recent years. In\nspite of a different chemical basis, they are similar to\ncelluloid and acetate sheets in their properties without,\nhowever, possessing their disadvantageous properties\n\nC. Mixture of Polymers\n\nThe polymerized plastics are thermoplastic materials\nwhich do not consist of cellulose nitrate or acetate but\nare built up of water, carbon, and chalk (fig. 5). By mix-", "timestamp": "2026-07-19T18:17:17.482532+00:00"}
{"citation_id": "19930091705", "source_url": "https://ntrs.nasa.gov/api/citations/19930091705/downloads/19930091705.pdf", "page_number": 10, "total_pages": 10, "image_filename": "19930091705_p10.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 | $X$ | $X$ | Rolling | $L$ | $Y \\longrightarrow Z$ | Roll | $\\phi$ | $u$ | $p$ |\n| Lateral | $Y$ | $Y$ | Pitching | $M$ | $Z \\longrightarrow X$ | Pitch | $\\theta$ | $v$ | $q$ |\n| Normal | $Z$ | $Z$ | Yawing | $N$ | $X \\longrightarrow Y$ | Yaw | $\\psi$ | $w$ | $r$ |\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\n4. 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\n5. NUMERICAL RELATIONS\n\n1 hp. = 76.04 kg-m/s = 550 ft-lb./sec.\n1 metric horsepower = 1.0132 hp.\n1 m.p.h. = 0.4470 m.p.s.\n1 m.p.s. = 2.2369 m.p.h.\n\n1 lb. = 0.4536 kg.\n1 kg. = 2.2046 lb.\n1 mi. = 1,609.35 m = 5,280 ft.\n1 m. = 3.2808 ft.", "timestamp": "2026-07-19T18:17:24.629631+00:00"}
{"citation_id": "19930094557", "source_url": "https://ntrs.nasa.gov/api/citations/19930094557/downloads/19930094557.pdf", "page_number": 15, "total_pages": 20, "image_filename": "19930094557_p15.jpg", "text": "N.A.C.A. Technical Memorandum No. 859\nFigs.4,9,10\n\n[Figure: Model airfoil mounted on rotation device.]\n\nFigure 4.- Model airfoil mounted on rotation device.\n\n[Figure: Side view of airfoil.]\n[Figure: Front view of airfoil.]\n\nFig.9- Side view.\nFig.10- Front view.\n\nFigure 9,10.- Side and front view of airfoil and six component balance.", "timestamp": "2026-07-19T18:17:39.886316+00:00"}
{"citation_id": "19930094533", "source_url": "https://ntrs.nasa.gov/api/citations/19930094533/downloads/19930094533.pdf", "page_number": 37, "total_pages": 51, "image_filename": "19930094533_p37.jpg", "text": "N.A.C.A. Technical Memorandum No. 883\n\nFigs. 7, 10, 14\n\n[Figure: Close-up of model surface with small circular taps and a protruding tab.]\n\nFigure 7.— Pressure and temperature taps on model.\n\n[Figure: Model mounted inside wind tunnel structure, showing support struts and tunnel walls.]\n\nFigure 10.— Model mounted in tunnel.\n\n[Figure: Wing assembly mounted vertically on test rig within large wind tunnel facility.]\n\nFigure 14.— Wing mounted for test.", "timestamp": "2026-07-19T18:17:48.244777+00:00"}
{"citation_id": "19930091716", "source_url": "https://ntrs.nasa.gov/api/citations/19930091716/downloads/19930091716.pdf", "page_number": 14, "total_pages": 24, "image_filename": "19930091716_p14.jpg", "text": "10\nREPORT NO. 641—NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nAlthough the original plotted curves of thrust and torque coefficient were, in general, fairly smooth, the few irregularities in some of the curves and their spacing made it seem desirable to cross-fair them. An illustration of the appearance of one of the typical original\n\nComparison of propellers having 2, 3, and 4 blades.—A comparison of propellers having 2, 3, and 4 blades is shown in figure 12. The coefficients of the three propellers were divided by the number of blades and then multiplied by 2 to permit comparison on the basis of\n\n<!-- Image (124, 165, 488, 437) -->\n\nFIGURE 10.—Typical pair of negative thrust and torque curves showing test points. Propeller 5868-9, 3 blades, set 30° at 0.75 R.\n\n<!-- Image (500, 165, 888, 437) -->\n\nFIGURE 11.—Comparison of thrust and torque coefficients for propellers having Clark Y and R. A. F. 6 sections; 3-blade, 10-foot propellers set 25° at 0.75 R.\n\nplots is given in figure 10 to show the extent of the dispersion of the test points.\n\nComparison of Clark Y and R. A. F. 6 propeller characteristics.—It will be noted that, in general, the values of thrust and torque coefficients are greater for the Clark Y propellers than for the R. A. F. 6 propellers,\n\ntwo blades. The curves in figure 12 show no consistent variation, probably owing to the process of cross-fairing. They do show that, compared on this basis, there is no great difference between the characteristics of propellers with 2, 3, and 4 blades.\n\nCoefficients for locked propellers.—Figure 13 shows\n\n<!-- Image (124, 535, 488, 752) -->\n\nFIGURE 12.—Comparison of thrust and torque coefficients for propellers having 2, 3, and 4 blades of Clark Y section; set 25° at 0.75 R.\n\n<!-- Image (500, 535, 888, 752) -->\n\nFIGURE 13.—Thrust and torque coefficients with propeller locked ($nD/V=0$); blades.\n\nFor an easier comparison, figure 11 was prepared to show the thrust-coefficient and torque-coefficient curves for the 3-blade Clark Y and R. A. F. 6 propellers set at a blade angle of 25°. At zero $nD/V$, the thrust coefficients of the two propellers are nearly the same; however, the difference in the shapes of the two sections (see fig. 2) causes a considerable difference in both thrust and torque throughout most of the $nD/V$ range.\n\nthe thrust and torque coefficients, at $\\frac{nD}{V}=0$ (propeller locked) and through a 90° blade-angle range, for both the R. A. F. 6 and Clark Y 3-blade propellers.\n\nThe difference in the shape of the curves, which is negligible at 0° and quite large at high angles, may be attributed to the difference in the shape of the leading edges of the two airfoil sections. The static thrust has apparently not reached its peak at 0°.", "timestamp": "2026-07-19T18:17:56.226796+00:00"}
{"citation_id": "19930094559", "source_url": "https://ntrs.nasa.gov/api/citations/19930094559/downloads/19930094559.pdf", "page_number": 3, "total_pages": 16, "image_filename": "19930094559_p3.jpg", "text": "NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nTECHNICAL MEMORANDUM NO. 857\n\nINVESTIGATION OF IGNITION AND COMBUSTION PROCESSES OF\nDIESEL ENGINES OPERATING WITH TURBULENCE AND\nAIR-STORAGE CHAMBERS*\n\nBy Hans Petersen\n\nModel tests were conducted at the machine laboratory\nof the Technical High School of Dresden (1935-36) with\nthe object of investigating the processes in Diesel en-\ngines before and during combustion. The tests are a con-\ntinuation of the work of Holfelder (reference 1). In the\npresent paper some of the results are presented that have\nbeen obtained in the study of the effect of air movement\non flame development and combustion in engines operating\nwith turbulence chamber and air-storage chamber, respec-\ntively.\n\n1. TEST SET-UP\n\nA reciprocating compressor delivers compressed air\ninto a combustion bomb mounted on the cylinder head. The\nair then flows through a throttle valve over an electric\nheater back to the suction valve of the compressor, the\nair circulation being continued until the desired temper-\nature condition of the bomb is attained. At the time the\ntest photographs are taken the throttle valve is closed\nand the bomb charged, according to the test conditions de-\nsired. The injection and combustion process can be ob-\nserved through two windows in the bomb and photographically\nrecorded. For the purposes of the present investigation,\nthe following modifications were introduced in\nthe test set-up of Holfelder. The fuel pump, instead of\nbeing separately driven, was gear-coupled to the compres-\nsor so that the start of injection of the pump could be\nadjusted to any compressor crank angle. In what follows,\nthe injection starting time will be referred to the com-\n\n*\"Untersuchung des Zünd- und Verbrennungsvorganges der\nnach dem Wirbelkammer- und Luftspeicher- Verfahren ar-\nbeitenden Dieselmotoren.\" Forschung auf dem Gebiete\ndes Ingenieurwesens, vol. 8, no. 6, November-December\n1937, pp. 279-284.", "timestamp": "2026-07-19T18:17:57.980596+00:00"}
{"citation_id": "19930094544", "source_url": "https://ntrs.nasa.gov/api/citations/19930094544/downloads/19930094544.pdf", "page_number": 16, "total_pages": 43, "image_filename": "19930094544_p16.jpg", "text": "14 N.A.C.A. Technical Memorandum No. 872\n\nlarge number of main rings and the rather large column lengths of about 11 to 13 m in the longitudinals. The close subdivision of the gas space may well have contributed to the fact that the structural weight in the two English airships has turned out to be relatively high.\n\nThe spacing of longitudinals is limited by the condition that a certain figure should not be exceeded for the free span width of the outer cover, which is laced to the outer booms of the longitudinals. In the German constructions LZ 127 and LZ 129, as well as in the \"Akron,\" it amounts to around 3.50 m. Also with respect to these figures, those previously customary have been exceeded in the English airships. In order to reduce the distortion and fluttering of the outer cover resulting from the great span width, a special supporting structure has been provided in the R 100, which pulls the cover inward. On the other hand, in the R 101 portable intermediate longitudinals are placed between adjacent main longitudinals (figure 30), which serve to tension the cover radially. However, since these intermediate longitudinals are not adapted to taking tension, they represent a useless excess weight; a further reason for the high structural weight in the R 101.\n\nAll previous German rigid airships have a frame-stiffening keel girder, which serves to transfer to the main rings the weights located in the lower part of the airship (fig. 19). In contrast to this, in the R 101 such a keel girder has been entirely avoided, since for the greater part it was possible to place the weights in the spacious main rings. The corridors provided are made up of relatively weak framing (fig. 34). In the \"Akron\" three corridors in all are provided, one at the top and one on each side in the lower part of the airship at 45° to the longitudinal plane. In the forward part of the airship a corridor runs from the control car to the extreme bow. The engines are inside the airship in properly fitted rooms at the intersections of the side corridors with four midship main rings.\n\nFor the attachment of the stabilizing surfaces it has been heretofore customary to construct a stiff cruciform frame in one or more of the main rings in the longitudinal location of the surfaces, to which the surfaces can then be attached without bracing (fig. 33). In the German and English airships, this manner of construction has been retained. In the \"Akron,\" on the other hand, the surfaces", "timestamp": "2026-07-19T18:18:01.836430+00:00"}
{"citation_id": "19930094542", "source_url": "https://ntrs.nasa.gov/api/citations/19930094542/downloads/19930094542.pdf", "page_number": 65, "total_pages": 102, "image_filename": "19930094542_p65.jpg", "text": "N.A.C.A. Technical Memorandum No. 874\nFigs. 41, 42\n\n[Figure: Graph showing multiple curves plotted against a grid. The horizontal axis is labeled $y/R$ with values from -3.0 to 3.0. The vertical axes on the left and right have various scales including $\\epsilon - \\epsilon_0$, $d'$, $q$, and $z/R$. The legend indicates curves for $d'$ and $\\frac{q-p}{p}$. An annotation $\\alpha = 14.9^\\circ$ is present.]\n\nFigure 41.- Downwash angle and dynamic pressure distribution.\n\n[Figure: Diagram of a mechanical assembly. Labels point to \"Fairing\", \"Motor\", and \"Exchangeable bearing piece\".]\n\nFigure 42.- Model wing with propeller.", "timestamp": "2026-07-19T18:18:02.036903+00:00"}
{"citation_id": "19930094534", "source_url": "https://ntrs.nasa.gov/api/citations/19930094534/downloads/19930094534.pdf", "page_number": 11, "total_pages": 21, "image_filename": "19930094534_p11.jpg", "text": "10 N.A.C.A. Technical Memorandum No. 882\n\n3. Flat, Countersunk Riveting with\nCountersunk and Dimpled Skin\n(F S y/F and R/F S y)\n\nThis method is used when the maximum sheet thickness\nor the maximum gripping lengths for P and F S x are ex-\nceeded. The different operating stages are illustrated\nin figures 22 to 25 for flush rivets and for mushroom-\nhead rivets, once with countersinking beginning at the\nthickest sheet and then with countersinking beginning at\nthe third sheet. F S y/F in conjunction with F S x is\nprimarily a machine method, while R/F S y in conjunc-\ntion with P lends itself to either machine or hand oper-\nation, depending upon accessibility.\n\n4. Flush-Riveting Dimpled Holes and Driven-In Sheets\n(R/F S z)\n\nHand riveting only: The process is illustrated in\nfigure 26. Standard half-round rivet heads are headed\nwith flat dolly. The sheets are neither countersunk nor\ndimpled but forced into the equally deforming rivet head\nby flat driving of the upsetting mushroom or barrel-shaped\nclosing head. Properly executed, the strength of this\nrivet joint will be sufficient, but since it is impossible\nto check the quality of workmanship on the finished piece\n(closing head too flat!) this method should be abandoned\nin favor of P or F S x on all highly stressed struc-\ntural parts.\n\nIf a contoured rivet head is desired, the DIN L 177\nflat, round-head rivets can be used in similar manner\n(F R/F S z). Then the flat dolly is replaced by a dolly\nwith snap.\n\nIII. AUTOMATIC STRIP INSERTION FOR\nDRAWING-SHEET SECTIONS (PATENTED)\n\nThe sections made of strip commonly used in airplane\nconstruction can be rolled through separately driven sets\nof rolls or else be fabricated by drawing. The rolling", "timestamp": "2026-07-19T18:18:02.037273+00:00"}
{"citation_id": "19930091655", "source_url": "https://ntrs.nasa.gov/api/citations/19930091655/downloads/19930091655.pdf", "page_number": 1, "total_pages": 22, "image_filename": "19930091655_p1.jpg", "text": "Library. Mass. Inst. of Tech.\nAERO. & ASTRO. LIBRARY\nMASS. INST. TECH.\n17 FEB 1937\nLIBRARY\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\n\nREPORT No. 580 C.3\n\nHEAT TRANSFER TO FUEL SPRAYS INJECTED\nINTO HEATED GASES\n\nBy ROBERT F. SELDEN and ROBERT C. SPENCER\n\n[Figure: Seal of the National Advisory Committee for Aeronautics]\n\n1937\n\nFor sale by the Superintendent of Documents, Washington, D. C.\nSubscription price, $3.00 per year\nPrice 10 cents\n\n623.742\nN58n", "timestamp": "2026-07-19T18:18:02.037515+00:00"}
{"citation_id": "19930094538", "source_url": "https://ntrs.nasa.gov/api/citations/19930094538/downloads/19930094538.pdf", "page_number": 26, "total_pages": 43, "image_filename": "19930094538_p26.jpg", "text": "24 N.A.C.A. Technical Memorandum No. 878\n\nvertical tangent, passes through a minimum, and slowly increases again. For complete tension field, the angle $\\alpha$ jumps at $\\tau = \\tau_0$ to an angle $\\alpha_0$ dependent on the structural quantities and the buckling stress $\\tau_0$, and rises evenly. In both cases the angles would, at very high values of $T/T_0$ approach one and the same value, affected only by $\\gamma$, $\\kappa$, and $\\delta$, provided no change in material behavior occurred in the meantime. The computed principal axes' angles do not agree with the measured angles of wrinkles. In principle elastic displacements in the directions inclined under the angle of wrinkling are definitely feasible. By the Ebner-Heck method, angle $\\alpha$ jumps from $45^\\circ$ to $0^\\circ$ when $T = T_0$, only to rise again immediately with vertical tangent, and then follow a course similar to that for complete tension field.\n\nIn the solution of the angles of twist (fig. 20), on the assumption of incomplete tension field, the correct choice of $R_0$ and $\\sigma_z$ affords a satisfactory agreement between the computed curve $\\psi(T)$ and the test points. The theoretical values are somewhat too high on cylinder No. III, but a more favorable choice of $R_0$ and $\\sigma_z$ would result in an excessively high theoretical failing strength. (Of. section IV.) For $\\tau = \\tau_0$, the calculation involves a minor jump in angle of twist, since the share of $\\epsilon$ in equations (2a) and (4), due to the second principal stress $\\sigma_2$ and the radial displacement, is not allowed for. Assuming a complete tension field, the sudden wrinkling by at $T = T_0$, results in a great jump of angle of twist; the computed values are much too high throughout the entire range. By the Ebner-Heck method, it gives $\\psi = 0$ and $\\frac{d\\psi}{dT} = \\infty$ at $T = T_0$; in the further course the computed angles of twist themselves are too great, although the error is somewhat less than with the complete tension-field method.\n\nd) Review of the calculating methods.- Under the assumption of incomplete tension field, the calculation for cylinders Nos. III and IV can, by proper selection of $R_0$ and $\\sigma_z$ be brought into satisfactory agreement with the experimental results, both as regards buckling stiffness and ultimate buckling strength (fig. 20, and section IV). The provisional assumptions, equations (26) and (27), are therefore practical to a certain extent for defining the \"stretching\" of the sheet panels. With the chosen $R_0$ and $\\sigma_z$, the compressive stresses $\\sigma_x$ of the stringers", "timestamp": "2026-07-19T18:18:05.275909+00:00"}
{"citation_id": "19930091714", "source_url": "https://ntrs.nasa.gov/api/citations/19930091714/downloads/19930091714.pdf", "page_number": 31, "total_pages": 36, "image_filename": "19930091714_p31.jpg", "text": "EFFECT OF COMPRESSIBILITY ON PROPELLERS IN TAKE-OFF AND CLIMBING RANGE 27\n\ncomputations but the thrust is given in figure 65 together with those for design B and the uncorrected thrust.\n\nIt may be noted from figure 65 that little, if any, loss in thrust due to compressibility is evident for this example. The explanation lies in the fact that the tip speed drops to about 0.7 the speed of sound in the take-off range owing to the decrease in engine speed. It may be noted from figure 58 that a maximum of only 4 percent in efficiency is lost for this tip speed. It appears from this example that computations for correcting the thrust of fixed-pitch propellers may not be worth while in many instances. A preliminary estimate of the tip speed in the take-off range together with a reference to the correction factors would indicate the importance of further computations. It probably is desirable in any case to make allowances in the design of propellers for differences in the power coefficient for test data from low and high tip speeds in order to determine the diameter and the blade angle.\n\nREFERENCES\n\n1. Briggs, L. J., Hull, G. F., and Dryden, H. L.: Aerodynamic Characteristics of Airfoils at High Speeds. T. R. No. 207, N. A. C. A., 1925.\n2. Stack, John: The N. A. C. A. High-Speed Wind Tunnel and Tests of Six Propeller Sections. T. R. No. 463, N. A. C. A., 1933.\n3. Briggs, L. J., and Dryden, H. L.: Aerodynamic Characteristics of Twenty-Four Airfoils at High Speeds. T. R. No. 319, N. A. C. A., 1929.\n4. Stack, John: The Compressibility Burble. T. N. No. 543, N. A. C. A., 1935.\n5. Douglas, G. P., and Perring, W. G. A.: Wind-Tunnel Tests with High Tip Speed Airscrews.\n   (a) The Characteristics of the Aerofoil Section R. A. F. 31a at High Tip Speeds. R. & M. No. 1086, British A. R. C., 1927.\n   (b) The Characteristics of a Bi-Convex Aerofoil at High Speeds. R. & M. No. 1091, British A. R. C., 1927.\n   (c) The Characteristics of Bi-Convex No. 2 Aerofoil Section at High Speeds. R. & M. No. 1123, British A. R. C., 1928.\n   (d) The Characteristics of a Conventional Airscrew Section, Aerofoil R. & M. No. 322, No. 3, at High Speeds. R. & M. No. 1124, British A. R. C., 1928.\n   (e) Some Experiments upon an Airscrew of Conventional Blade Section, Aerofoil R. & M. No. 322, No. 3, at High Speeds. R. & M. No. 1174, British A. R. C., 1928.\n   (f) The Characteristics of a Conventional Airscrew Section 0.082c Thick and of R. A. F. 27 and R. A. F. 28. R. & M. No. 1198, British A. R. C., 1929.\n6. Hartshorn, A. S., and Douglas, G. P.: Wind Tunnel Tests on High Tip Speed Airscrews. Further Experiments on Scale Effect. R. & M. No. 1417, British A. R. C., 1932.\n7. Jennings, W. G., and Ormerod, A.: Full Scale Experiments on High Tip Speed Airscrews. The Effect of Thickness of Section on Airscrew Performance. R. & M. No. 1339, British, A. R. C., 1931.\n8. Wood, Donald H.: Full-Scale Tests of Metal Propellers at High Tip Speeds. T. R. No. 375, N. A. C. A., 1931.\n9. Weick, Fred E., and Wood, Donald H.: The Twenty-Foot Propeller Research Tunnel of the National Advisory Committee for Aeronautics. T. R. No. 300, N. A. C. A., 1928.\n10. Biermann, David, and Hartman, Edwin P.: Tests of Five Full-Scale Propellers in the Presence of a Radial and a Liquid-Cooled Engine Nacelle Including Tests of Two Spinners. T. R. No. 642, N. A. C. A., 1938.\n11. Biermann, David, and Hartman, Edwin P.: The Aerodynamic Characteristics of Six Full-Scale Propellers Having Different Airfoil Sections. T. R. No. 650, N. A. C. A., 1939.", "timestamp": "2026-07-19T18:18:24.255538+00:00"}
{"citation_id": "19930094551", "source_url": "https://ntrs.nasa.gov/api/citations/19930094551/downloads/19930094551.pdf", "page_number": 4, "total_pages": 18, "image_filename": "19930094551_p4.jpg", "text": "N.A.C.A. Technical Memorandum No. 855 3\n\nbelow a limit of error of ±0.5 m in the computation of the pressure-head position relative to the airplane. The static pressure tubing consisted of an all-metal hose of 7 mm outside diameter made by the Berlin-Karlsruher Industrie-Werke. The all-metal hose was preferred after previous tests had disclosed that the customary metal-wound rubber hose is occasionally squeezed off at the point of emergence from the airplane*. The total length of available hose was 24 m. The hose lead from a spool of 200 mm diameter over a roller to the pressure head. A dynamometer of 20 kg test range was provided to be installed in the hose portion between roller and spool and indicated the tension in the hose. Flight performance losses due to the drag component in flight direction were accounted for in measurements of the slope of the upper end of the hose relative to the longitudinal axis of the airplane. Agreement of the measurements with respect to time for the reading of the bearing-loop settings by two observers was insured by a light signal.\n\nThe flight tests were made at speeds of from 200 to 350 km/h. No readings were possible above this speed, because the pressure head was no longer within range of the forward bearing loop.\n\nThe speedometer employed in these flights was flight-tested over a square course. All measurements were made in horizontal flight as much as possible in order to preserve the condition of flow direction perpendicular to the direction of gravity. Compliance with this condition makes direct comparison of the free-flight data with the subsequently described calculation method for defining the tubing curve, possible. Horizontal flight was possible up to speeds of 320 km/h; increased to 350 km/h with full throttle, the airplane already exhibited a sinking speed of 2 m/s. Even so, the resulting errors remain within the bounds of measuring accuracy. In the first flight tests an experimental model consisting of a streamlined brass pipe of 40 mm diameter filled with hard lead and four plain fins was employed. The total weight of the model amounted to 5.4 kg. The same plain stabilizing fins were used on the final design (fig. 2) rather than the stabilizing cone with 6° angle of incidence (fig. 2a) used on the low-speed pressure head, which had proved impracticable for high speeds.\n\n*Owing to the relatively short life of metal hose, various synthetic products are being investigated by the Institute for Aeronautics.", "timestamp": "2026-07-19T18:18:39.447845+00:00"}
{"citation_id": "19930094538", "source_url": "https://ntrs.nasa.gov/api/citations/19930094538/downloads/19930094538.pdf", "page_number": 27, "total_pages": 43, "image_filename": "19930094538_p27.jpg", "text": "N.A.C.A. Technical Memorandum No. 878 25\n\non cylinder No. III are correct within the experimental range, but too high for cylinder No. IV. The latter is probably due to the stronger than linear decrease of curve R($\\phi$) against $\\phi = \\frac{T}{T_0} = 1$. The ring stresses on cylinder IV are obviously computed correctly since the theoretical and experimental values for bending moments and bulkhead stresses are approximately the same. The theoretical values for bending moment and bulkhead stresses are too high for cylinder No. III. A plausible explanation for this is that the second principal stress, even in the center, between stringers, is somewhat dependent on the shearing stress T. But a complete explanation is impossible until the legitimacy of R($\\phi$) and $\\xi_y(\\phi)$ has been proved by experiments.\n\nWith the use of the complete tension field, the angles of twist are much too high. This divergence is largely attributable to the assumption of wrinkling $\\xi_y$ on buckling. The compressive stresses in the stringers are tolerably correct within the experimental range for cylinder No. III, since R ~ 1; but the fact that R < 1 results in markedly excessive theoretical values for $\\sigma_x$ on cylinder No. IV. As to bulkhead stresses and bending moments, the same statement made for incomplete tension field holds true.\n\nWith the compressive stresses in the stiffeners as a result of the omitted effective skin strip and (on cylinder No. IV, especially) the different kind of allowance for the skin-buckling strength, the Ebner-Heck method yields even greater errors than the calculation for complete tension field; and the results are similar with the Wagner-Ballerstedt method.\n\nThe solution of the twisting stiffness is therefore contingent upon accurate experiments on the behavior of the sheet panels in the buckling stage; for this, account of the \"incomplete tension field\" is absolutely essential. On the contrary, calculation on the basis of complete tension field is practical for estimating the expected stresses, since it comprises their order of magnitude and always remains on the safe side.", "timestamp": "2026-07-19T18:18:47.857574+00:00"}
{"citation_id": "19930094564", "source_url": "https://ntrs.nasa.gov/api/citations/19930094564/downloads/19930094564.pdf", "page_number": 4, "total_pages": 16, "image_filename": "19930094564_p4.jpg", "text": "N.A.C.A. Technical Memorandum No. 852 3\n\na maximum critical Reynolds Number of $4.05 \\times 10^6$ in flight in still air manifested the value $3.7 \\times 10^5$ in tunnel center.\n\nIn the meantime the Americans fortunately succeeded in proving by comparison of sphere and maximum-lift measurements in the N.A.C.A. variable-density tunnel and in the N.A.C.A. full-scale tunnel, that the turbulence lowers the critical Reynolds Number of the sphere in the same ratio as it does for the maximum-lift measurements, (reference 5), or in other words, that for maximum-lift investigations the Reynolds Number of the test must be multiplied by the ratio of the critical Reynolds Number of the sphere in nonturbulent air to that in the tunnel in order to obtain the Reynolds Number of the maximum lift measurement applicable in flight. The ratio of the critical Reynolds Number of the sphere in nonturbulent air stream to the critical Reynolds Number in the tunnel is called the \"turbulence factor.\"\n\n$$\n\\text{T.F.} = \\frac{\\text{Reynolds Number(still air)}}{\\text{Reynolds Number(tunnel)}}\n$$\n\n(Reynolds Number of sphere for drag ($c_w$) = 0.3.) The Reynolds Number which is valid for maximum lift in flight and which is obtained by multiplying the turbulence factor with the Reynolds Number of the maximum-lift measurement is called \"effective\" Reynolds Number.\n\n$$\nR_{\\text{effective}} = R_{\\text{test}} \\text{ T.F.}\n$$\n\nThe effective Reynolds Number has proved satisfactory in the comparison of $c_{a_{\\text{max}}}$ measurements effected in several different tunnels as well as in free flight (reference 5). It constitutes a definite advance in the elucidation of the $c_{a_{\\text{max}}}$ question and removes the existing uncertainty. It is used hereinafter for comparing the $c_{a_{\\text{max}}}$ measurements of the DVL with those of other tunnels.\n\nThe turbulence factors of various tunnels are listed in table I.", "timestamp": "2026-07-19T18:18:54.112640+00:00"}
{"citation_id": "19930094542", "source_url": "https://ntrs.nasa.gov/api/citations/19930094542/downloads/19930094542.pdf", "page_number": 66, "total_pages": 102, "image_filename": "19930094542_p66.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-19T18:18:56.667406+00:00"}
{"citation_id": "19930093641", "source_url": "https://ntrs.nasa.gov/api/citations/19930093641/downloads/19930093641.pdf", "page_number": 10, "total_pages": 47, "image_filename": "19930093641_p10.jpg", "text": "8\n\nparisons and drag increments given in the report, the bare-wing data that were obtained immediately following the tests of the wing-nacelle condition have been used as a reference for the results of the wing-nacelle tests, and the bare-wing data obtained after the enclosed-engine tests have been used as their references. Fortunately the slope of the scale-effect curves, although differing between the two test groups, showed good agreement within each group of test conditions.\n\nThe principal drag comparisons are made between data obtained at a tunnel speed of 100 miles per hour, corresponding to a Reynolds Number of 4,300,000. These comparisons show that the wing nacelles increase the high-speed drag coefficient by an increment of 0.0015, or 8.7 percent. The underslung Prestone radiators and leading-edge oil coolers add 0.0035, or 20.4 percent, so that the total increase in drag due to the exposed power-plant installation is 0.0050, or 29.1 percent.\n\nDrag increments for the extension-shaft installations were small, being in most cases within the experimental accuracy. The shortest extension shaft gave the highest drag increment, as shown by the 4-percent increase in the high-speed drag for position 3 (table I); this result may possibly be attributed to the disturbed flow from the end of the extension shaft as it passes over the wing.\n\nThe propeller spinners shown in figure 7 do not appreciably affect either the high-speed or the minimum drag coefficients. The results for the 8-inch cowlings, chosen to represent a 56-inch-diameter air-cooled engine nacelle on the leading edge of a 100-ton airplane, show about a 4- to 5-percent increase in the high-speed drag coefficient.\n\nMaximum lift.- The maximum lift coefficients for all the arrangements tested are summarized in table I. The extension shafts for the enclosed-engine arrangements are apparently not detrimental to the maximum lift; in fact, the pusher arrangement shows an unexplainable higher value of maximum lift coefficient than the bare-wing condition. The lower maximum lift coefficients for the conditions with nacelles on the wing loading edge are caused by nacelle interference; the effect is clearly demonstrated by the tuft observations shown in figures 24 and 25. For the wing-alone condition (fig. 24), the stall progresses uniformly inward from the tips with increasing lift coefficient; whereas, for the wing-nacelle condition (fig. 25),", "timestamp": "2026-07-19T18:19:09.027665+00:00"}
{"citation_id": "19930094544", "source_url": "https://ntrs.nasa.gov/api/citations/19930094544/downloads/19930094544.pdf", "page_number": 17, "total_pages": 43, "image_filename": "19930094544_p17.jpg", "text": "N.A.C.A. Technical Memorandum No. 872 15\n\nhave been attached directly to the outer framing, relying upon the inherently stiff main rings for rigidity. In the English airships the passenger and crew spaces are located in the interior of the airship in the forward half of the airship, likewise the living spaces for the complement of the \"Akron.\" The latter are located adjacent to the side corridors; between them a free space is bridged over, which serves for the accommodation of five airplanes. Figures 19 to 24 can serve further to clarify the frame structures of the various rigid airships. Further figures are found in references 10 to 16.\n\n2. Structural Elements\n\nJust as the five newer airships differ in general arrangement of framing, they also differ from one another in girder design. The LZ 127 has girders similar to those which were usual in earlier Zeppelin airships. The longitudinal and ring girders are of triangular form, their channel-shaped corner members being joined by means of corrugated lattices (fig. 35). For the LZ 129, entirely new kinds of girders have been developed, which likewise are shown in figure 35. The corner members are joined by means of oppositely set U-shaped struts, extensively provided with lightening holes. The pot-shaped corner members used for the new girders are especially shown in figure 35. The upper sections are used in the more lightly stressed, the lower in the more heavily stressed girders. Figure 36 shows a truss member of a main ring of LZ 127. The kind of latticing for the various girders is clearly recognized in this. Figure 37 shows the girders newly developed by the Luftschiffbau Zeppelin and having the oppositely set strut bracing, and shows also the attachment of the latter to the outer and inner legs of the corner members.\n\nIn the \"Akron\" a departure has been made from the triangular type of girder and rectangular box girders (fig. 35) have been developed for the ring members. These girders have no real corner members. Rather, the wall plates of the girders grip over one another at the corners and have stiffening grooves there. Merely by the setting-in of a corner piece the corners are transformed into closed sections. The wall plates have extensive lightening holes. In like manner this construction is also applicable to triangular box girders. The ring girders used in the R 101 have an appearance similar to that of the ring girders in", "timestamp": "2026-07-19T18:19:09.571469+00:00"}
{"citation_id": "19930094552", "source_url": "https://ntrs.nasa.gov/api/citations/19930094552/downloads/19930094552.pdf", "page_number": 20, "total_pages": 32, "image_filename": "19930094552_p20.jpg", "text": "18 N.A.C.A. Technical Memorandum No. 864\n\ndrical shell loaded at four points. Two loading conditions were investigated, namely, a bending load and an \"arching\" load. From the measured longitudinal stresses the shears and the transverse stresses in the skin were obtained by integration of the fundamental equations of the two-dimensional stress field.\n\nThe measured stresses in the longitudinal stiffeners (stringers) are in satisfactory agreement with the computed values of Ebner and Köller for both loading conditions. The longitudinal stresses in the stiffeners between the \"main spars\" increase steadily, in the bending case, in the external fiber region from the loading end toward the fixed end. In the neighborhood of the neutral axis these stresses, after a rapid initial rise, attain a maximum and then decrease again to a value which corresponds to the linear stress distribution due to bending as a beam. In the arching type of loading condition, the stresses in the stiffeners between the main spars remain small.\n\nThe experimentally determined shears for the bending loading condition and arching loading condition with the two front bulkheads not attached to the skin were continuously distributed over the cylinder shell. In the arching loading condition with the two forward bulkheads riveted to the sheet, however, discontinuities arise in these bulkheads because the shears are transmitted to the bulkheads. In this case, comparison with the computed values gives a stronger concentration of the experimentally determined shears at the load application end. Within a bay the shears vary over the length. In the peripheral direction, however, they vary strongly under the bending load in the region of the extreme fibers. Near the neutral axis, however, they are practically constant along a strip. The same is true for the arching loading condition in all walls between the main spars.\n\nThe transverse stresses are small under the arching load if the first two bulkheads take up the shear differences as a result of their being riveted to the sheet. If the attachment to the skin is loosened, however, very high peripheral stresses are set up in the bay directly behind the force application (panel V) and these fluctuate greatly and change their signs within the bay. In the bending case, the peripheral stresses remain small.\n\nThe loading of the stiffener system depends essen-", "timestamp": "2026-07-19T18:19:13.837615+00:00"}