Buckets:
| {"citation_id": "19930090755", "source_url": "https://ntrs.nasa.gov/api/citations/19930090755/downloads/19930090755.pdf", "page_number": 23, "total_pages": 23, "image_filename": "19930090755_p23.jpg", "text": "N.A.C.A. Technical Memorandum No.445\nFigs.4,5,11.\n\n[Figure: Diagram of a device for welding tubes of like diameter, showing two views: a side cross-section and a top-down view.]\n\nFig.4 Device for welding tubes of like diameter.\n\n[Figure: Diagram of a mounting for tensile test, showing a central specimen held in a fixture with numbered parts 1 through 6.]\n\nFig.5 Mounting for tensile test.\n\n[Figure: Diagram illustrating annealed zones of welds, showing three configurations: \"Not annealed\", \"Annealed through welded laps\", and \"Annealed by butt welding\", each with corresponding cross-sectional views.]\n\nFig.11 Annealed zone of a weld.", "timestamp": "2026-07-19T11:20:49.967470+00:00"} | |
| {"citation_id": "19930094828", "source_url": "https://ntrs.nasa.gov/api/citations/19930094828/downloads/19930094828.pdf", "page_number": 1, "total_pages": 24, "image_filename": "19930094828_p1.jpg", "text": "FILE COPY\nNO 1-W\n\nN 62 57588\n\nTECHNICAL MEMORANDUMS\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nNo. 588\n\nTHE BEHM ACOUSTIC SOUNDER FOR AIRPLANES\nWITH REFERENCE TO ITS ACCURACY\nBy Ernest Schreiber\n\nJahrbuch 1930\nder Deutschen Versuchsanstalt für Luftfahrt\n\nWashington\nOctober, 1930\n\nFILE COPY\nTo be returned to\nthe files of the National\nAdvisory Committee\nfor Aeronautics\nWashington, D. C.", "timestamp": "2026-07-19T11:20:51.575139+00:00"} | |
| {"citation_id": "19930094849", "source_url": "https://ntrs.nasa.gov/api/citations/19930094849/downloads/19930094849.pdf", "page_number": 5, "total_pages": 26, "image_filename": "19930094849_p5.jpg", "text": "N.A.C.A. Technical Memorandum No. 565\n\nin performance followed the adoption of the heavy-oil aircraft engine. The automobile engine is also shown for comparison. The data are based on endurance tests for all three engines.\n\nThe first line gives the horsepower per liter of stroke volume or piston displacement. This shows the fundamental superiority of the two-stroke cycle, since this value of 21 hp per liter already reaches and even excels the best carburetor engines and the piston speed is no lower. Line three gives the weight-power ratio in kilograms per horsepower. The weight was reduced to 1/20 of that of the stationary engine.\n\nThe steps taken in 1924 and 1925 were not applied to an experimental engine of small or medium power but, on the contrary, to a 700–800 hp oil engine, the \"FO 3.\"\n\nThe most important further constructive step was the change to an integral engine-and-drive casting (Fig. 5). This change resulted from considerations of the strength and rigidity of a complete connection between the cylinders and the driving gear.\n\nThe large casing which unites 5 cylinders into one block, imposes great demands upon the technique of casting light metals. It must be a very strong, bubble-free, and flawless casting which is also corrosion-proof from the flow of water through it.\n\nFrom the very beginning, the material chosen was silumin, a special alloy developed at the suggestion of Junkers. This is remarkably easy to cast and has good strength, density, and", "timestamp": "2026-07-19T11:20:56.813507+00:00"} | |
| {"citation_id": "19930094822", "source_url": "https://ntrs.nasa.gov/api/citations/19930094822/downloads/19930094822.pdf", "page_number": 2, "total_pages": 40, "image_filename": "19930094822_p2.jpg", "text": "NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS.\n\nTECHNICAL MEMORANDUM NO. 594\n\nPRACTICAL TESTS WITH THE \"AUTO CONTROL SLOT\"*\n\nBy G. Lachmann\n\nPART II: Discussion\n\nDr. W. Pleines.— For some time the D.V.L. (Deutsche Versuchsanstalt für Luftfahrt) (German Aeronautic Institute) has been investigating the question of the applicability of Handley Page slotted wings to German airplanes. Comparative gliding tests were made with open and closed slots on an Albatros L 75 airplane equipped with the Handley Page \"auto control slot.\" This investigation served to determine the effect of the auto control slot on the properties and performances of airplanes at large angles of attack. The most important problems were as to whether the angle of glide at small angles of attack can be increased by the adoption of the auto control slot and, in particular, as to whether the flight characteristics at large angles of attack are improved\n\n---\n\n*\"Praktische Erfahrungen mit dem automatischen Spaltflügel,\" a paper read before the Wissenschaftlichen Gesellschaft für Luftfahrt, March 17, 1930. From Zeitschrift für Flugtechnik und Motorluftschiffahrt, September 15, 1930, pp. 440-448. For Part I (Dr. Lachmann's lecture) see N.A.C.A. Technical Memorandum No. 593.", "timestamp": "2026-07-19T11:20:59.323544+00:00"} | |
| {"citation_id": "19930090741", "source_url": "https://ntrs.nasa.gov/api/citations/19930090741/downloads/19930090741.pdf", "page_number": 28, "total_pages": 40, "image_filename": "19930090741_p28.jpg", "text": "N.A.C.A. Technical Memorandum No. 359\n28\n\nvery satisfactory and enables quick and reliable results. The\ncoordinated values from the individual drawings can, for ex-\nample, be transferred every minute to a numerical table.\n\nA good dynamic-pressure gauge is of special importance\nfor all test flights. Care must also be taken to reduce retar-\ndation through inertia as much as possible. The tubes should\nnot be too small, although no air is transmitted. Any number\nof indicating or recording instruments can, with the aid of\nT-pieces, be connected with the same measuring instrument. In\nusing Venturi tubes, special attention must be given to the\nstatic-pressure disturbances. If several indicators are used,\ncare must be taken to have them all under the same static pres-\nsure. The task of the airplane pilot is simply to fly accord-\ning to the dynamic pressure. During a series of tests, he must\ntherefore keep the dynamic pressure constant by a very cautious\nmanipulation of the elevator control. It is entirely indiffer-\nent as to how the scale on the indicator is divided, but it is\nbetter to indicate the dynamic pressure in millimeters of water\ncolumn rather than in the often misleading speed scale of kilo-\nmeters per hour. It is only important for the pilot to deter-\nmine experimentally at what position of the pointer the air-\nplane climbs best and for him to understand the danger of fall-\ning below this value (i.e., of \"stalled flight\"). It is very\nhelpful to mark this dynamic-pressure value in red. Exceeding\nan upper limit of the dynamic pressure may also become danger-", "timestamp": "2026-07-19T11:21:02.081124+00:00"} | |
| {"citation_id": "19930094847", "source_url": "https://ntrs.nasa.gov/api/citations/19930094847/downloads/19930094847.pdf", "page_number": 20, "total_pages": 46, "image_filename": "19930094847_p20.jpg", "text": "N.A.C.A. Technical Memorandum No. 567\n\nleft to right.\n\nThe comparison of this field with the natural field of force of the Magnus flow, as drawn from a photograph (Fig. 11), shows only a slight similarity between the two. Only in the hatched upper partial field do the lines of force return to the surface of the cylinder. Behind this, however, on the whole rear side, they are strengthened by the shearing forces of the friction and deflected over the lower side of the cylinder, so that, together with the lines of force emanating from the front side, they now encircle the cylinder spirally and finally escape into the fluid.\n\nThus the natural lines of force are described by the forces derived from the rotating cylinder by friction and maintain the peculiar circulation in its vicinity which, according to Magnus and Lord Rayleigh, is the immediate cause of the lateral force. Accordingly the assumption is also disproved experimentally, that the circulation is produced automatically without the direct effect of the friction, and that the natural Magnus flow is correctly represented by the theoretical potential flow of Lord Rayleigh.\n\nThe same result is obtained by comparing the theoretical and natural flows as represented by streamlines. Figure 12 represents an initial stage of the Magnus flow after three or four revolutions of the cylinder. It shows the initial series of vortices, which very soon develop into the starting vortex of", "timestamp": "2026-07-19T11:21:10.775093+00:00"} | |
| {"citation_id": "19930091325", "source_url": "https://ntrs.nasa.gov/api/citations/19930091325/downloads/19930091325.pdf", "page_number": 34, "total_pages": 40, "image_filename": "19930091325_p34.jpg", "text": "```markdown\nTABLE I—Continued\nPRESSURE IN POUNDS PER SQUARE FOOT ON THE TS AIRPLANE\n\n| Station No. | Upper wing | | | | | | | | | | | Lower wing | | | | | | | | | | | Tail surface | | | | | | | |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| | 1 | 2 | 3 | 4 | 5 | 6 | 7 | 8 | 9 | 10 | 11 | 1 | 2 | 3 | 4 | 5 | 6 | 7 | 8 | 9 | 10 | 11 | 1 | 2 | 3 | 4 | 5 | 6 | 7 | 8 |\n\n**Level flight**\n\n| Air speed (M. P. H.): | | | | | | | | | | | | | | | | | | | | | | | | | | | | | |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| 58.5 | 15 | 21 | 23 | 18 | 15 | 12 | 8 | 6.5 | 4 | 2 | 0.5 | 15 | 20.5 | 20 | 14 | 13.5 | 12 | 9 | 5.5 | 3.5 | 2.5 | 1.5 | | | | | | | | |\n| 62.8 | 19 | 25 | 27 | 23.5 | 19 | 13.5 | 9 | 6.5 | 8 | 2 | 2 | 15 | 19 | 18 | 14.5 | 13 | 12 | 10 | 6 | 4 | 3 | 2 | | | | | | | | |\n| 71.5 | 15 | 23 | 26.5 | 23 | 21 | 12.5 | 9 | 6 | 3 | 1 | 8 | 12.5 | 13 | 13 | 13.5 | 14 | 11.5 | 8 | 5 | 3 | 1.5 | | | | | | | | | |\n| 79.0 | 15 | 21 | 24 | 22 | 21 | 18 | 13.5 | 10.5 | 7.5 | 4.5 | 1.5 | 0 | 11 | 11 | 9 | 14 | 16 | 13 | 9 | 6 | 4 | 1.5 | | | | | | | | |\n| 85.5 | 15 | 21 | 23 | 19 | 21.5 | 22 | 18 | 12.5 | 9 | 4 | 2 | -1.5 | 1 | 8 | 8 | 9.5 | 14 | 18.5 | 15.5 | 11 | 6 | 2 | | | | | | | | |\n| 93.5 | -12 | 2 | 16 | 17.5 | 24.5 | 22 | 18 | 13.5 | 10 | 6 | 2 | -23.5 | -4 | 2.5 | 5.5 | 13 | 19.5 | 17 | 11 | 19 | 6 | 2 | | | | | | | | |\n| 106.0 | -15 | 3 | 11 | 13 | 24.5 | 23 | 21 | 16.5 | 12 | 8 | 5 | 0 | -23 | -3 | 2 | 13 | 19 | 20 | 14 | 19 | 5 | 2 | | | | | | | | |\n| 117.0 | -31 | -6 | 4 | 9.5 | 22 | 21.5 | 17 | 11.5 | 6 | 1.5 | 0 | -39 | -3 | -16.5 | -12 | 14 | 19 | 18 | 12 | 7.5 | 5 | 3.5 | | | | | | | | |\n| 127.0 | -47 | -26 | -3 | 5 | 26 | 25.5 | 22 | 15.5 | 8 | 3.5 | 2 | -72 | -50 | -24 | -18 | 17.5 | 23 | 21 | 14.5 | 10 | 6.5 | 3.5 | | | | | | | | |\n\n**Pull-up—power on**\n\n| Time (seconds): | | | | | | | | | | | | | | | | | | | | | | | | | | | | | |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| 0 | 0 | 10 | 28 | 31 | 22.5 | 18 | 12.5 | 11.8 | 3.5 | -70 | -40 | | -19 | 22.5 | 26 | 26 | 14.5 | 13.5 | 9.3 | 5.5 | -24 | -19 | -3.5 | -3.4 | -4 | 2 | 0 | 0 | |\n| 0.50 | -48 | 0 | 10 | 28 | 31 | 22.5 | 18 | 14 | 7 | 3.5 | -70 | -42 | | -19 | 22.5 | 26 | 26 | 14.5 | 13.5 | 9.3 | 5.5 | -25 | -19 | -11 | -15 | -12 | -37 | -13 | -7.5 | |\n| 0.75 | 0 | 38 | 58 | 59 | 43 | 34 | 28 | 19 | 8 | 0 | 0 | 74 | | 45 | 60 | 34 | 34 | 21.6 | 18 | 10.2 | 5.5 | 0 | -14 | 0 | -5 | -10.2 | -27 | -12 | -5 | |\n| 1.00 | 82 | 144 | 122 | 109 | 87.5 | 64 | 31.5 | 17.3 | 7.6 | 0 | 78 | 93 | | 71 | 64 | 49 | 42.5 | 26 | 20 | 11 | 5.5 | 30 | 0 | 0 | -0.4 | -7 | -23 | -10.5 | -4 | |\n| 1.25 | 110 | 129 | 128 | 97 | 91 | 59 | 43.5 | 29 | 16.5 | 6.5 | 0 | 69 | 84 | | 61 | 57 | 43 | 36 | 21.6 | 18 | 11 | | 20 | 0 | 0 | 0 | -7 | -20 | -5.5 | -3.5 | |\n| 1.50 | 82 | 104 | 109 | 81 | 73.5 | 55.5 | 37 | 25 | 15 | 8.5 | 0 | 56 | 63 | | 38 | 48 | 38 | 32.5 | 17.5 | 16 | 9.5 | | 13 | 0 | 0 | 0 | -5 | -14 | -3 | -2.5 | |\n| 2.50 | 46 | 62 | 58 | 48 | 43 | 35 | 22.5 | 16 | 10 | 4 | 0 | 25 | 39 | | 33 | 31 | 25 | 22.5 | 11.4 | 10 | 3.5 | 3 | 5 | 0 | 0 | 0 | -2 | -10.2 | -2 | 0 | |\n| 3.50 | 22 | 38 | 33 | 29 | 24 | 14 | 13.5 | 9 | 7.7 | 2 | 0 | 16 | 27 | | 18 | 21 | 15 | 14 | 6 | 5 | 2 | 3 | 3 | 0 | 0 | 0 | 0 | -7 | 0 | 0 | |\n\n**Pull-up—power off**\n\n| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| 0 | -40 | -12 | 0 | 29 | 30 | 24 | 19.5 | | 7.5 | 5.5 | -68 | -28 | -16 | 19 | 26.5 | 22.5 | 12.2 | 12.5 | 5.5 | 0 | -31.5 | -16.5 | -4.5 | -5 | -5 | -4 | 0 | | |\n| 0.25 | 0 | 0 | 0 | 28 | 30 | 28 | 18.4 | | 7 | 5.5 | -94 | -40 | | -28 | 19 | 27 | 22.5 | 11.5 | 12.5 | 5.5 | 0 | -30 | -16.5 | -3.5 | -9 | -13 | -35 | -13 | -4.5 | |\n| 0.50 | 0 | 88 | | 64.5 | 55 | 39.5 | 28 | | 9.5 | 4 | 42 | 45 | 38 | 36.5 | 47 | 34 | 17.5 | 18 | 8.5 | 0 | -36 | -20 | -5.5 | 0 | -12 | -31 | -15.5 | | |\n| 0.75 | 92 | 164 | 104 | 86 | 64 | 46 | 32.5 | | 10 | 8.5 | 4 | 94 | 78 | 76 | 76 | 66 | 54 | 41 | 20.5 | 21 | 10 | | 0 | -2 | 0 | 0 | -9 | -23 | -4 | |\n| 1.00 | 116 | 152 | | 88 | 64 | 44.8 | 35.5 | | 9.5 | 4 | 89 | 102 | 100 | 76 | 66 | 49 | 41 | 23.5 | 20 | 17.5 | | 33 | 9 | 0 | 0 | -9 | -18 | -12.5 | | |\n| 1.25 | 109 | 128 | 143 | 88 | 64 | 44 | 32 | | 8.5 | 4 | | 78 | 69 | 54 | 55 | 39 | 37.5 | 20.5 | 18 | 10 | | 28 | 7.5 | 0 | 0 | -9 | -16 | -10.5 | | |\n| 1.50 | 72 | 98 | 111 | 67 | 50 | 33 | 26.5 | | 8 | 4 | | 65 | 80 | 54 | 55 | 39 | 32.5 | 17.5 | 17 | 7.5 | 0 | 23 | -4 | 0 | 0 | -9 | -12 | -9 | | |\n| 1.75 | 58 | 94 | | 67 | 50 | 33 | 26.5 | 17 | 8 | 4 | | 56 | 68 | 50 | 51 | 39 | 29 | 19.5 | 14.5 | 7.5 | 0 | 14 | 0 | 0 | 0 | -7 | -11 | -9 | | |\n\n**Pull-down**\n\n| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| 0 | -56 | 0 | 28 | 31 | 21 | 19.2 | | 8.5 | | -74 | -43 | -28 | -14 | 19 | 23 | 22.3 | 13.4 | 13.8 | 8.3 | | -30 | -16.5 | -5.5 | -3.2 | -3.5 | 6 | 0 | 0 | |\n| 0.50 | -56 | -20 | -5 | 29 | 30 | 23 | 17 | 7 | 5.5 | | -94 | -43 | -28 | -14 | 19 | 23 | 22.3 | 11.8 | 12.3 | 8.3 | | -30 | -16.5 | -5.5 | -3.2 | -3.5 | 40 | 13.5 | 4.4 | |\n| 1.00 | -96 | -59 | 20 | 21 | 15.5 | 11.7 | 7 | 5.5 | | -129 | -65 | -65 | -38 | 7 | 12 | 17 | 10.2 | 11 | 8.3 | | -63 | -22.5 | -9.8 | -4.7 | -5.5 |", "timestamp": "2026-07-19T11:21:18.318015+00:00"} | |
| {"citation_id": "19930090744", "source_url": "https://ntrs.nasa.gov/api/citations/19930090744/downloads/19930090744.pdf", "page_number": 21, "total_pages": 25, "image_filename": "19930090744_p21.jpg", "text": "N.A.C.A. Technical Memorandum No.361\n\nFig.2\n\n[Figure: Graph showing variation in mechanical properties of Duralumin against tempering temperatures, with curves labeled Breaking strength,R; Elastic limit,E; Elongation,A; Resilience,ρ. Axes: vertical R E A (0 to 32), horizontal Tempering temperatures,°C (0 to 400°). Right-side axis for ρ (0 to 5).]\n\nFig.2 Variation in the mechanical properties \nof Duralumin plotted against the temp- \nering temperatures.", "timestamp": "2026-07-19T11:21:19.058276+00:00"} | |
| {"citation_id": "19930094857", "source_url": "https://ntrs.nasa.gov/api/citations/19930094857/downloads/19930094857.pdf", "page_number": 17, "total_pages": 68, "image_filename": "19930094857_p17.jpg", "text": "N.A.C.A. Technical Memorandum No. 557 16\n\nHence a tube had to be used, the readings of which vary as the square of the angle of inclination.\n\nThis does not apply to the Brabbée tube, since up to about $20^\\circ$, it is nearly insensitive to angular deflections (Fig. 6).* On the other hand, according to Figure 6, the normal Prandtl tube with a bore/diameter ratio of 0.3 fully satisfies this requirement up to about $20^\\circ$. Hence a Prandtl tube was used instead of a Brabbée tube for certain measurements (flights 25, 28, 30, 31).\n\nThese measurements were expected to produce a slightly greater drag than those made with a Brabbée tube. Hence, according to the above consideration and owing to the lack of angular sensitivity of the Brabbée tube, its momentum measurements in the vortex trail were too large, and the difference in comparison with the potential flow was therefore too small. The result, however, was negative. The measurements made with the Prandtl tube did not yield uniformly greater values than those obtained with other tubes. The smallest drag coefficients ever determined were, in fact, measured with the Prandtl tube. We might thus be led to believe that the angular deflections in the vortex trail never attain values worth mentioning. This conclusion, however, should be adopted with caution, since the actual conditions have not yet been sufficiently explained.\n\n*The Brabbée tube was originally adopted because of its lack of sensitivity to errors of inclination.", "timestamp": "2026-07-19T11:21:19.268738+00:00"} | |
| {"citation_id": "19930090695", "source_url": "https://ntrs.nasa.gov/api/citations/19930090695/downloads/19930090695.pdf", "page_number": 29, "total_pages": 45, "image_filename": "19930090695_p29.jpg", "text": "N.A.C.A. Technical Memorandum No. 367 28\n\noratory under my charge and then undertook to apply the principle of his rudder to the sail of a ship. Sailing vessels had gradually come into a very disadvantageous position in comparison with steamships and Diesel-engine ships, due principally to the necessity of a large crew for handling the sails and making the frequent repairs in the rigging. Flettner wished to introduce metal sails, made like the wings of metal airplanes, which would automatically assume the correct position with reference to the wind by means of wind vanes and auxiliary rudders. The storm problem was very difficult, however. The metal sails could not be reefed, though they could always be brought, by means of an auxiliary rudder, /into the exact direction of the wind, so that they received no lateral pressure. What if, however, the auxiliary rudder should be damaged in a storm and remain in such a position that the sails would receive the full pressure? There was the further disillusionment that the old type of sail, when correctly adjusted to the wind, was not so bad as we had been inclined to assume, but generated forces equal to about 80% of those produced by metal sails of the same size. The metal sails had to be very large, therefore, in order to fully replace the old sails.\n\nMr. Flettner therefore began to search for some other substitute. When informed of the Göttingen experiments with the rotating cylinder, he quickly decided to have the availability of such cylinders for his sail ship investigated and made arrangements with us for this purpose. Our previous experiments enabled us to suggest immediately, as the most favorable, the very form which was subsequently installed on his ship. For the reasons already mentioned, each cylinder had to be long,", "timestamp": "2026-07-19T11:21:30.905739+00:00"} | |
| {"citation_id": "19930094849", "source_url": "https://ntrs.nasa.gov/api/citations/19930094849/downloads/19930094849.pdf", "page_number": 6, "total_pages": 26, "image_filename": "19930094849_p6.jpg", "text": "N.A.C.A. Technical Memorandum No. 565 5\n\ncorrosion resistance with a specific gravity of only 2.65.\n\nThe first \"FO 3\" engine came to the test stand for trial in the middle of 1926. The first tests promised great advantages for aircraft.\n\nThree months later this engine reached a peak performance of 830 b.hp at 1200 r.p.m. with a mean effective pressure of 8.3 atmospheres during a half-hour run. The empty weight of the complete engine was about 930 kg (2050 lb.). The duration of the run may appear short but the results were important, in that they showed that the goal could be reached by the path upon which a start had been made. The next problem was to design the parts of the engine so they would be thoroughly capable of withstanding the working stresses. For every new engine has its defects, and an 800 hp aviation engine at this stage of its development may be very troublesome.\n\nIn the \"FO 3\" engine a certain fundamental defect in the 5-cylinder type evidenced quite an undesirable effect in the imperfect balancing of the centrifugal moments. On this account a change was made about a year later to a 6-cylinder engine, the present \"SL 1,\" whose dimensions were proportional but somewhat smaller, corresponding to a cylinder bore of 120 mm (4.72 in.) instead of 140 mm (5.51 in.) in the 5-cylinder engine.\n\nWhile this engine was being built, numerous experiments were made on the separate problems of the working process and of light construction for which two special single-cylinder experimental engines were used (Fig. 6).", "timestamp": "2026-07-19T11:21:32.630575+00:00"} | |
| {"citation_id": "19930090739", "source_url": "https://ntrs.nasa.gov/api/citations/19930090739/downloads/19930090739.pdf", "page_number": 24, "total_pages": 33, "image_filename": "19930090739_p24.jpg", "text": "N.A.C.A. Technical Memorandum No. 357\n23\n\nthey possess a small downward velocity component in the cylinder axis, they meet the air stream generated by the diffuser during the compression stroke and mix with it principally in the marginal region produced by the cylindrical neck. It seems to be worthy of note that this neck is short and angular and that the working piston, at the end of the compression stroke comes very close to the surface of the heat shield, so that all but a small fraction of the air charge is driven through the neck and strong eddies are generated in the air stream on both its inner and outer surfaces.\n\nAt the beginning of the combustion, the burning gases press from both sides toward the cylinder axis and, since they suffer many changes in their direction of motion and mean cross sections, on the way to the working cylinder, thus generate, between the neighboring stream filaments, strong displacements and eddies, which assist the combustion. The fuel consumption of 190 g (6.7 oz.) per $HP_e/hr$. (converted to a lower heat value of 10,000 heat units per kilogram at an engine power of 100 HP.) seems very favorable. The combustion takes place, as shown by the original indicator diagrams, in the form of a simple explosion. At a compression pressure of 11 atm. the combustion pressure was 27 atm. (Fig. 25, right).\n\nThis engine also resembles an explosion engine in two other respects. The air charge introduced during the suction stroke, which, moreover, helps to expel the exhaust gases from the com-", "timestamp": "2026-07-19T11:21:36.968484+00:00"} | |
| {"citation_id": "19930094836", "source_url": "https://ntrs.nasa.gov/api/citations/19930094836/downloads/19930094836.pdf", "page_number": 8, "total_pages": 42, "image_filename": "19930094836_p8.jpg", "text": "N.A.C.A. Technical Memorandum No. 580\n\n$$\nu = \\frac{8 \\, J}{\\rho \\, \\pi \\, b^2 \\, \\Delta \\, a}\n$$\n\nThis applies to a plate element of width $b$ and length $\\Delta a$ when $J = P \\, \\Delta t$. Hence the mass of water to be accelerated is\n\n$$\nM_W = \\rho \\, \\frac{\\pi}{8} \\, b^2 \\, \\Delta \\, a \\tag{1}\n$$\n\nThe distribution of the impact pressures over $b$ is elliptical. It is assumed that the plate is absolutely rigid. In practice the flow and the pressure distribution are subject to variation.\n\n5. The assumption of an infinite length does not apply to the actual float bottom. In fact, the length of the bottom portion which strikes the water is of the same order of magnitude as the width. Since, on the assumption of an infinite plate length, the bottom width goes into the second power, while it has a smaller power in the case of a finite bottom length, the latter must be taken into consideration, in order to avoid wrong conclusions regarding the effect of the width of the hull on the impact. In this case the bottom portion concerned can also be considered as a plate in an infinite liquid, the one-sidedness of the process being taken into consideration. The water mass accelerated by such plates of a finite chine ratio was determined experimentally by means of small vibrations. When a body vibrates in a nonviscous, incompressible, infinite fluid at rest, the mass of the body is increased by the flow which de-", "timestamp": "2026-07-19T11:21:37.165709+00:00"} | |
| {"citation_id": "19930094824", "source_url": "https://ntrs.nasa.gov/api/citations/19930094824/downloads/19930094824.pdf", "page_number": 9, "total_pages": 21, "image_filename": "19930094824_p9.jpg", "text": "N.A.C.A. Technical Memorandum No. 592\n\nweight 12%. In connection herewith we determined the stresses and deformations in a beam panel as well as the stress over all struts. The position of the measured panels a, b, c, and d is seen in Figure 11. The obtained test figures are shown at an enlarged scale in Figure 19.\n\nUpon closer examination of the panel deformations in Figure 19, we find: The top strut is stretched, and its mean stress over b, c is $120 \\text{ kg/cm}^2$ on the C.G. line of the angles. The stress is naturally lower near the four corners, because the gusset plates increase the cross section. In the center of both gusset plates the stress was found to be $136 \\text{ kg/cm}^2$ while, according to Professor Wagner, it amounted to $120 \\text{ kg/cm}^2$. The stress in the compression strut was $147 \\text{ kg/cm}^2$ in the center of both intersection points, as measured on the C.G. line of the angles. The calculated compression stress is $156 \\text{ kg/cm}^2$, according to Professor Wagner's report. Upper and lower struts being rigidly connected to the upright members by the angle plates, the deflection curves from corner to corner are S-shaped. The upright members are under $18 \\text{ kg/cm}^2$ compression in contrast to $22.5 \\text{ kg/cm}^2$, according to the calculation mentioned.\n\nInasmuch as it was impossible to place the measuring instruments other than on the struts and vertical members, we obtained a mean reduced stress in the plate, whose value is given at the four corners with respect to the angle position. The", "timestamp": "2026-07-19T11:21:37.357194+00:00"} | |
| {"citation_id": "19930094853", "source_url": "https://ntrs.nasa.gov/api/citations/19930094853/downloads/19930094853.pdf", "page_number": 14, "total_pages": 25, "image_filename": "19930094853_p14.jpg", "text": "N.A.C.A. Technical Memorandum No. 561\n\nGain in Useful Load and in Radius of Action\n\nThese reductions in the braking effect are very appreciable, but it is interesting also to examine the consequences of this improvement in the fineness of the airplane, that is, (for example), to determine the additional load that can be carried in flight.\n\nCalling $C_X/C_Z$ the fineness with propeller locked and $\\frac{C_X}{C_Z} - \\Delta \\frac{C_X}{C_Z}$ the fineness with a free-wheel propeller, the equations of flight are:\n\n$$\n75 \\rho P_m = \\pi \\frac{C_X}{C_Z} V_1 \\quad (\\pi = \\text{load carried})\n$$\n\nand\n\n$$\n75 \\rho P_m = \\pi' \\left( \\frac{C_X}{C_Z} - \\Delta \\frac{C_X}{C_Z} \\right) V_1',\n$$\n\nfrom which we derive\n\n$$\n\\pi' = \\pi \\frac{\\frac{C_X}{C_Z}}{\\left( \\frac{C_X}{C_Z} - \\frac{R}{\\pi} \\right)^{2/3}}\n$$\n\nExample 1.- We can assume:\n\n$$\n\\pi = 6000 \\text{ kg} \\quad \\text{and} \\quad \\frac{C_X}{C_Z} = 0.111\n$$\n\nwhence\n\n$$\n\\pi' = 6000 \\left( \\frac{0.111}{0.111 - \\frac{76.5}{6000}} \\right)^{2/3}\n$$\n\n$$\n= 6000 \\left( \\frac{0.111}{0.0983} \\right)^{2/3}\n$$\n\n$$\n\\pi' = 6000 \\times 1.085 = 6510 \\text{ kg}, \\quad \\text{or a gain in useful load of}\n$$\n\n$$\n\\pi' - \\pi = 510 \\text{ kg},\n$$", "timestamp": "2026-07-19T11:21:37.823989+00:00"} | |
| {"citation_id": "19930094822", "source_url": "https://ntrs.nasa.gov/api/citations/19930094822/downloads/19930094822.pdf", "page_number": 3, "total_pages": 40, "image_filename": "19930094822_p3.jpg", "text": "N.A.C.A. Technical Memorandum No. 594\n2\n\nthereby and equilibrium in gliding flight is guaranteed even at\nlarger than ordinary angles of attack.\n\nThe arrangement and form of the auto control slot, whose\nfunctioning has just been described in detail by Lachmann, is\nshown in Figure 26. The auxiliary wings or flaps were placed\nsimply on the leading edge of the upper wing opposite the aile-\nrons. When closed, they did not exactly reproduce the original\nprofile shape for constructional reasons. They could be freed\nfor automatic functioning simply by removing a few bolts.\n\nThe angle of attack, air speed, sinking speed, longitudinal\ninclination, angle of glide and elevator deflection were meas-\nured in unaccelerated gliding flight. The most important results\nof these measurements, whose execution and calculation will not\nhere be described in detail, are presented in Figures 27-36.\n\nFigure 27 shows the course of the sinking speed $v_s$ as\nplotted against the actual dynamic pressure $q_w$ with open and\nclosed slots. The actual dynamic pressure $q_w$ was determined\nfrom speed flights over a quadrangular course by a suitable\ncalibration of the indications of the dynamic-pressure gauge.\n\nWith the slots closed $v_s$ increases greatly for $q_w$\nvalues smaller than 50 kg/m² (10.24 lb./sq.ft.) and greater\nthan 40 kg/m² (8.19 lb./sq.ft.), while, with the slots open,\nthe previous course of the $v_s$ curve is preserved up to $q_w$\nvalues of about 45 kg/m² (9.22 lb./sq.ft.). Above this point,\nhowever, the $v_s$ values increase much more rapidly. Through", "timestamp": "2026-07-19T11:21:44.596385+00:00"} | |
| {"citation_id": "19930094847", "source_url": "https://ntrs.nasa.gov/api/citations/19930094847/downloads/19930094847.pdf", "page_number": 21, "total_pages": 46, "image_filename": "19930094847_p21.jpg", "text": "N.A.C.A. Technical Memorandum No. 567\n21\n\nPrandtl. In the earlier stages there is no accumulation of\nboundary-layer material from which, according to Prandtl, this\nvortex develops; neither is there any thickening of the bounda-\nry layer, which, according to the last modification of the\ntheory, is the cause of this vortex. On the other hand, it is\nshown how the fluid layers from the upper side of the cylinder\nare thrust wedge-shaped against the lower side. At the tip\nof the wedge, thrust far forward, these layers bend sharply\nbackward toward the counterflow and enclose the rotational line\nR, from which the vortices of the layer successively proceed.\n\nBigger photographs show, after the disappearance of the\nlarge initial vortex, small vortices on the division line,\nwhere the flow from the upper side of the cylinder reunites\nwith the flow on the lower side (Fig. 13). These constantly\ndiminishing small vortices take their position according to the\nimmediate continuation of the initial vortex sheet. Figure 13\nshows the finished Magnus flow. In contrast with the symmet-\nrical pattern of the theory, it is turned by the effect of\nfriction through an angle $\\alpha$, so that the free Magnus force\nCM is now composed of a purely lateral force $C_a$ and a re-\nsistance or drag $C_w$. One is here again referred to the above-\nmentioned beautiful picture of the Magnus flow taken by O.\nTietjen.\n\nThe velocity difference between the surface of the rotat-\ning cylinder and the surrounding fluid, and hence also the", "timestamp": "2026-07-19T11:21:46.003547+00:00"} | |
| {"citation_id": "19930094852", "source_url": "https://ntrs.nasa.gov/api/citations/19930094852/downloads/19930094852.pdf", "page_number": 4, "total_pages": 32, "image_filename": "19930094852_p4.jpg", "text": "N.A.C.A. Technical Memorandum No. 562\n3\n\nwere recently made at the Junkers factory.* These experiments\ncover the angles of flow in curving motion (case 1). Of course\nthey also cover the effect of the oblique flow in those cases\nin which the relative flow is generally assumed with sufficient\naccuracy to be axial. These cases include the straightaway\nflight of an airplane whose propeller axis has a direction dif-\nferent from the direction of flight, inverted flight, the flow\nrelations in certain arrangements of the propeller above or be-\nlow the wings and similar cases.\n\nIn order to determine the behavior of propellers under the\noperating conditions of case 2, Durand and Lesley investigated,\nin 1921, a series of propellers of different H/D ratios (0.3,\n0.5, 0.7) at angles of flow of $\\alpha = 60$ to $90^\\circ$ in a wind tunnel\nand determined the axial thrust and the torque.** Margoulis\nhas recnely investigated helicopter-propeller models in a lat-\neral wind.***\n\nSince the former of the above-mentioned cases (oblique flow\nagainst airplane propellers) relates to small angles of flow and\nthe latter (oblique flow against helicopter propellers) relates\nto large angles of flow and hence practically cover the range\nfrom $\\alpha = 0$ to $90^\\circ$, and since, moreover, the quantities measured\n\n*Book, \"Ueber die Einheit von Triebwerk und Flugwerk,\" Year-\nbook of the W.G.L., 1928, p.66.\n**Durand and Lesley, \"Tests on Air Propellers in Yaw.\" N.A.C.A.\nTechnical Report No. 113, 1921.\n***Margoulis, \"Les hélicoptères,\" Paris, 1922; \"Nouvelles re-\ncherches expérimentales sur les hélices d'hélicoptères,\" Comptes\nRendus 184, 1927, p.735.", "timestamp": "2026-07-19T11:21:47.586748+00:00"} | |
| {"citation_id": "19930094779", "source_url": "https://ntrs.nasa.gov/api/citations/19930094779/downloads/19930094779.pdf", "page_number": 1, "total_pages": 57, "image_filename": "19930094779_p1.jpg", "text": "FILE COPY\nNO. 1-\nFILE COPY\nNO. 2-W\nN 62 57637\n\nTECHNICAL MEMORANDUMS\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nNo. 637\n\nTHE DEVELOPMENT, DESIGN AND CONSTRUCTION\nOF GLIDERS AND SAILPLANES\nBy A. Lippisch\n\nThe Journal of The Royal Aeronautical Society\nJuly, 1931\n\nREPRODUCED BY\nNATIONAL TECHNICAL\nINFORMATION SERVICE\nU.S. DEPARTMENT OF COMMERCE\nSPRINGFIELD, VA. 22151\n\nWashington\nSeptember, 1931\n\n57", "timestamp": "2026-07-19T11:21:49.361289+00:00"} | |
| {"citation_id": "19930094849", "source_url": "https://ntrs.nasa.gov/api/citations/19930094849/downloads/19930094849.pdf", "page_number": 7, "total_pages": 26, "image_filename": "19930094849_p7.jpg", "text": "N.A.C.A. Technical Memorandum No. 565\n\nIn the following, the working process is first investigated more fully. Fresh air must be supplied to the cylinders not only with extreme rapidity, but with a minimum expenditure of energy. This requires short, large pipes between the blower and the cylinders and, above all, control ports with large cross sections. In the two-stroke cycle these can be enlarged at will over a wide range, but only at the expense of the volume of fresh air contained in the cylinder. In the interest of the greatest possible cylinder performance, one endeavors therefore to obtain the greatest possible charging volume with the most favorable scavenging.\n\nFigure 7 shows the indicator diagram of the two-stroke-cycle engine, i.e., the pressure plotted against the displacement. This working process is controlled by the simultaneous motion of both pistons. Thorough expansion of the combustion gases through properly timed opening of the exhaust ports before the scavenging piston opens the fresh-air inlet is especially important for high-speed engines. The two-shaft engine possesses a very practical and simple means of enlarging the exhaust opening, in that the crank is adjusted to run a little ahead on the exhaust side. In this manner the outlet ports not only open earlier, but the scavenging ports close correspondingly later, even after the closing of the exhaust ports, and thus fill the cylinder with fresh air at the scavenging pressure.\n\nIn order to give an illustration of the proportions of the", "timestamp": "2026-07-19T11:21:57.803172+00:00"} | |
| {"citation_id": "19930094823", "source_url": "https://ntrs.nasa.gov/api/citations/19930094823/downloads/19930094823.pdf", "page_number": 16, "total_pages": 29, "image_filename": "19930094823_p16.jpg", "text": "N.A.C.A. Technical Memorandum No. 593\n15\n\nically efficient airplane in a field surrounded by obstacles.\nA good fineness ratio (L/D) is an advantage for the economy\nof flight and a disadvantage in landing.\n\nFigure 18 shows the phases of an average landing over an\nobstacle. Let $\\epsilon$ represent the fineness ratio. Then\n$$S_1 = (H - h') \\frac{1}{\\epsilon},$$\nand for two airplanes, differing simply in their angle of glide,\nthe difference in the total landing distance is\n$$\\Delta L = (H - h') \\left( \\frac{1}{\\epsilon_1} - \\frac{1}{\\epsilon_2} \\right).$$\nFigure 19 shows the reduction in the length of glide and\nalso the reduction in the total landing distance, as made possi-\nble by the impairment of the fineness ratio, the assumptions\nbeing $\\epsilon_1 = \\frac{1}{9}$ and $h'_1$ (levelling-off altitude) = 3 m\n(9.84 ft.).\n\nFor heavily loaded and aerodynamically efficient airplanes,\nthe best fineness ratio occurs at large lift coefficients, so\nthat further increase in the angle of attack in the subcritical\nregion of the polar causes no considerable impairment of the\nangle of glide. The $c_a$ value corresponding to the best fine-\nness ratio results from the condition that the induced drag\nequals the remaining drag. If we put $c_{wg} = 0.04$ (coefficient\nof the remaining drag) for a modern commercial airplane and\nadopt an aspect ratio of $\\lambda = 7$, we obtain from\n$$c_{wi} = \\frac{c_a^2}{\\pi \\cdot \\lambda} = 0.04$$", "timestamp": "2026-07-19T11:22:02.000743+00:00"} | |
| {"citation_id": "19930094857", "source_url": "https://ntrs.nasa.gov/api/citations/19930094857/downloads/19930094857.pdf", "page_number": 18, "total_pages": 68, "image_filename": "19930094857_p18.jpg", "text": "N.A.C.A. Technical Memorandum No. 557 17\n\n5. q, the dynamic pressure in the after control plane, is measured with a Prandtl tube (tube III, Fig. 5). The dynamic-pressure readings of such a tube are independent of the angle of attack of the flow within quite a wide range (Reference 4). Logically, the considerations of No. 3 also apply to this case. This influence is negligible, however, owing to the small numerical value of the second integral.\n\n6. As stated in Section II, b, 3, $q_1$ is derived from q by addition of the total pressure drop\n\n$$\nq_1 = q + (g_0 - g).\n$$\n\n7. The exact position of the two Pitot tubes, located behind the wing, is shown in Figure 7. Their distance from the trailing edge is about 30% of the wing chord. This accounts for the comparatively large part (1/5 to 1/8) of the second integral in the total result. It might be objected that the wing chord opposite the two tubes is not the same. Owing to the fact, however, that the drag coefficient was calculated for the portion of the wing located in front of the Brabbée tube, this small error affects the correction member only and can be practically disregarded. The distance of the measuring instruments from their supports is sufficient to prevent the latter from disturbing the flow.", "timestamp": "2026-07-19T11:22:02.596265+00:00"} | |
| {"citation_id": "19930094828", "source_url": "https://ntrs.nasa.gov/api/citations/19930094828/downloads/19930094828.pdf", "page_number": 2, "total_pages": 24, "image_filename": "19930094828_p2.jpg", "text": "NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS.\n\nTECHNICAL MEMORANDUM NO. 588.\n\nTHE BEHM ACOUSTIC SOUNDER FOR AIRPLANES\nWITH REFERENCE TO ITS ACCURACY.*\n\nBy Ernest Schreiber.\n\nRelative altimetry is of great importance for increasing\nthe safety in aerial transportation, because it makes possible\nsafe flying at night, by poor visibility, and notably, when\nlanding. Among the instruments of this type is the Behm sounder,\nwhich operates on an acoustic principle.**\n\nPrinciples of Acoustic Altimetry\n\nAcoustic altimetry is based upon the measurement of the\ntime interval required for the sound to travel from the aircraft\nto the ground and back to the aircraft as an echo. From this\ninterval it is possible to arrive at the height at which the\nairplane flies. With c as the velocity of sound in the free\natmosphere and $\\Delta t$ as the total elapsed interval of the sound,\nthe sound path covered is $s_0 = c \\Delta t$, and the corresponding\nheight is $h_0 = \\frac{1}{2} c \\Delta t$, the velocity of sound being measured\nin meters per second; and $\\Delta t$ in seconds. According to the ex-\n\n*\"Das Behmlot für Flugzeuge und die mit ihm erzielte Genauigkeit.\"\nJahrbuch 1930 der Deutschen Versuchsanstalt für Luftfahrt, pp.\n483-490.\n\n**Behm-Luftlot, Zeitschrift für Flugtechnik und Motorluftschiff-\nfahrt, Vol. 15, 1924, p. 285; and A. Behm, \"Das Behmlot und seine\nEntwicklung als akustischer Höhenmesser für Luftfahrzeuge.\"\nBerichte und Abhandlungen der Wissenschaftlichen Gesellschaft für\nLuftfahrt, No. 13, 1928, p. 58.", "timestamp": "2026-07-19T11:22:16.432182+00:00"} | |
| {"citation_id": "19930094836", "source_url": "https://ntrs.nasa.gov/api/citations/19930094836/downloads/19930094836.pdf", "page_number": 9, "total_pages": 42, "image_filename": "19930094836_p9.jpg", "text": "N.A.C.A. Technical Memorandum No. 580\n\nvelops during the motion (Stokes*, Green**). In an ideal non-viscous fluid this flow is a potential flow. In viscous fluids the potential flow can be maintained with a good approximation, provided the body makes very short and quick vibrations.*** Under the above assumptions of disregarded friction and wave formation, this fact permits of easily determining the accelerated water mass as already suggested by Föttinger for other purposes but, so far as I know, never put into practice.\n\nFigure 2 shows the test installation. Plate 1, stiffened by a longitudinal rib, is secured to a duralumin tube 2, which connected with two steel springs 3 and 4, can vibrate along its longitudinal axis. This system, which is capable of vibrating, is deflected approximately 0.2 mm (0.008 in.) and then suddenly released by severing a wire. The resulting damped vibration was plotted by means of a scratch recording device 5,**** directly and without lever transmission, with a diamond on a glass plate moved laterally by the electric motor 6. The resulting diagram was estimated under a microscope using the simultaneously recorded time marks. This estimation, made on the assumption of proportional damping, showed that the influence of the damping on the period of vibration was negligibly small.\n\n---\n\n*Stokes, \"On Some Cases of Fluid Motion.\" Camb. Trans., 8, 1843, Math. and Phys. Papers I, p. 17.\n\n**Green, \"Researches on the Vibration of Pendulums in Fluid Media,\" Trans. R. S. Edin., 1883, Math. Papers, p. 315.\n\n***Föttinger, Jahrbuch d. Schiffbautechn. Gesellschaft, 1924.\n\n****Pabst, W., \"Aufzeichnungen schneller Schwingungen nach dem Ritzverfahren,\" Zeitschrift des Vereines deutscher Ingenieure, 1929. No. 46.", "timestamp": "2026-07-19T11:22:26.318226+00:00"} | |
| {"citation_id": "19930094853", "source_url": "https://ntrs.nasa.gov/api/citations/19930094853/downloads/19930094853.pdf", "page_number": 15, "total_pages": 25, "image_filename": "19930094853_p15.jpg", "text": "N.A.C.A. Technical Memorandum No. 581 14\n\nrepresenting, in fuel saved, an addition of 660 km (410 miles)\nto the radius of action (for a two-engine airplane with one engine stopped).\n\nExample 2.- The additional load for the second case, determined\nin the same fashion, would be\n\n$$\n\\pi' = 6000 \\times 1.08 = 6480 \\text{ kg, or a gain}\n$$\n\nof\n\n$$\n\\pi' - \\pi = 480 \\text{ kg,}\n$$\n\nrepresenting, in fuel saved, a supplementary radius of action\nof 621 km (for a two-engine airplane with one engine stopped).\n\nExample 3.- This example is based on the assumption of 1800 kg\nload.\n\n$$\n\\pi' = 1800 \\left( \\frac{0.111}{0.111 - 0.0124} \\right)^{2/3} = 1800 \\times 1.08\n$$\n\n$$\n= 1940 \\text{ kg,}\n$$\n\n$$\n\\pi' - \\pi = 140 \\text{ kg, representing an additional radius of action}\n$$\n\nof 625 km (for a two-engine airplane with one engine stopped).\n\nThis represents a gain, due simply to the use of the free-wheel propeller, of one kilogram raised or 1300 meters radius of action per horsepower of the stopped engine.", "timestamp": "2026-07-19T11:22:28.521616+00:00"} | |
| {"citation_id": "19930094849", "source_url": "https://ntrs.nasa.gov/api/citations/19930094849/downloads/19930094849.pdf", "page_number": 8, "total_pages": 26, "image_filename": "19930094849_p8.jpg", "text": "N.A.C.A. Technical Memorandum No. 565\n\ncontrol members in a high-speed engine, a comparison is made (Fig. 8) between the characteristics of the well-known four-stroke-cycle carburetor engine, Junkers \"L 5\" and the oil engine. All values in this figure are computed on the basis of equal piston displacement and r.p.m.\n\nThe amount of port opening of both engines is plotted in the diagram against the time, i.e., the crank angle. Whereas in the four-stroke cycle, 440 degrees of crank angle are available for exhaust and intake, only 150 degrees are available in the two-stroke cycle (about 1/3). Therefore, the sectional area of the port opening is several times greater in the two-stroke cycle and the opening and closing take place much more rapidly. One can clearly recognize, in the two-stroke cycle, the influence of the advanced angle on the displacement of the exhaust and scavenging processes.\n\nThe product of control cross-sectional area by the time, i.e., the area in the control diagram, gives a measure of the charging of the cylinder. These values are given in horizontal bars in the next row, separated for intake and exhaust (Fig. 8). The values give sufficient information concerning the proportions, above all, the large exhaust cross-sectional area of the two-stroke cycle.\n\nThe volume ratios between the control points (at the bottom in the figure) refer to equal piston displacements. It will be seen that, in the two-stroke cycle, for the sake of the best", "timestamp": "2026-07-19T11:22:37.887156+00:00"} | |
| {"citation_id": "19930094779", "source_url": "https://ntrs.nasa.gov/api/citations/19930094779/downloads/19930094779.pdf", "page_number": 2, "total_pages": 57, "image_filename": "19930094779_p2.jpg", "text": "NOTICE\n\nTHIS DOCUMENT HAS BEEN REPRODUCED\nFROM THE BEST COPY FURNISHED US BY\nTHE SPONSORING AGENCY. ALTHOUGH IT\nIS RECOGNIZED THAT CERTAIN PORTIONS\nARE ILLEGIBLE, IT IS BEING RELEASED\nIN THE INTEREST OF MAKING AVAILABLE\nAS MUCH INFORMATION AS POSSIBLE.", "timestamp": "2026-07-19T11:22:40.512741+00:00"} | |
| {"citation_id": "19930090695", "source_url": "https://ntrs.nasa.gov/api/citations/19930090695/downloads/19930090695.pdf", "page_number": 30, "total_pages": 45, "image_filename": "19930090695_p30.jpg", "text": "N.A.C.A. Technical Memorandum No. 367 29\n\nand projecting disks were added to both ends. The free upper disk had somewhat different functions than the previously described disks next to the walls (Fig. 23). Without them, air would have been drawn from the rear side of the cylinder into the negative-pressure region and would have dissipated the circulatory flow throughout a considerable portion of the length of the cylinder and indeed all the more, the greater the negative pressure would have otherwise been. Of course the disk had to rotate also, so that the already mentioned separation would not take place. The disks afforded the further advantage, which was clearly demonstrated in the experiments, of decreasing the induced drag, by dividing the marginal eddy into two eddies flowing away from the walls of the disks, thus producing an effect similar to the transition from a monoplane to a biplane.*\n\nFirst a cylinder with a built-in electric motor was tested (Fig. 24), once without disks and then with two pairs of disks of different diameters. Fig. 25 shows the combined lift coefficients $c_a$ and drag coefficients $c_w$ (resistance in the direction of the wind divided by Fq) in the form of polar curves, the dashed line (in the lower left corner) being the polar curve of an airplane wing. In Fig. 26, $c_a$ is plotted against u/V (ratio of the peripheral velocity of the cylinder to the velocity of the wind). It is obvious that the cyl-\n\n*Mr. Flettner has established his claim that he learned of this action of the disks from other sources and that he would have used disks on the cylinder even without our suggestion.", "timestamp": "2026-07-19T11:22:43.919053+00:00"} | |
| {"citation_id": "19930094824", "source_url": "https://ntrs.nasa.gov/api/citations/19930094824/downloads/19930094824.pdf", "page_number": 10, "total_pages": 21, "image_filename": "19930094824_p10.jpg", "text": "N.A.C.A. Technical Memorandum No. 532\n\nmaximum stress occurs at point 6 for this panel and amounts to \n$172 \\text{ kg/cm}^2$ at $\\varphi = 36^\\circ$. The stress, parallel to the struts \nis lower, although it can become maximum in the panel of the \nmaximum moment. The computed value of the plate stress, according \nto the report mentioned, is $180 \\text{ kg/cm}^2$.\n\nThe principal result of the defined stresses in the tension and compression strut prove that the measured stresses differ less than $10\\%$ from the computed stresses of Professor Wagner.\n\nThen for comparison, we compared the deflections of the beam with those of a lattice beam obtained by cutting out the web plate. The deflections of course are incomparably higher, as, for instance, a plate wall girder with 7 vertical members is 6 times more rigid than a beam from which the plate has been removed.\n\nIn conclusion, we compared the deflections and stresses of the plate wall girder with those of a lattice beam. To be sure, we used the same strut cross sections in both. The proportions of the vertical members and of the diagonals set at $45^\\circ$, were made with the intention of ensuring simultaneous weight in plate and lattice beam. Thus the diagonals and uprights had the same cross sections as the struts. Of course, these proportions do not agree with rational structural methods, but may be resorted to for comparing similarly constructed plate wall beams.", "timestamp": "2026-07-19T11:22:45.400275+00:00"} | |
| {"citation_id": "19930090739", "source_url": "https://ntrs.nasa.gov/api/citations/19930090739/downloads/19930090739.pdf", "page_number": 25, "total_pages": 33, "image_filename": "19930090739_p25.jpg", "text": "N.A.C.A. Technical Memorandum No. 357\n24\n\nbustion chambers, can be regulated by an air shutter. Then,\nhowever, there is installed in the axis of the cylinder head\nan auxiliary spark plug, which is set in operation when the\ncompression pressure, in starting from the cold condition, is\nnot sufficient to effect automatic combustion of the fuel mix-\nture, and which remains in operation until the engine has be-\ncome sufficiently heated. This spark-plug is a proof that an\nexplosive mixture is formed by the two spraying nozzles in con-\njunction with the diffuser and that the air flow produced by\nthe neck forces this mixture against the cylinder head and the\nspark plug.\n\nAs follows from the original indicator diagrams, spontane-\nous combustion often begins at a considerable distance from\nthe dead center, which might give rise to heavy shocks in the\ndriving gear, a disadvantage which also occurs occasionally in\nother solid-injection Diesel engines, when an explosive drive,\ndependent on spontaneous combustion, is produced in too pure\na form.\n\nBanner's engine, built by the Falk Corporation in Milwau-\nkee (Fig. 24) forms a fine companion piece to Price's engine.\nBanner's engine has a mixed indicator diagram, half Otto and\nhalf Diesel. Viewed externally, it is similar to Price's en-\ngine, in that it has the same arrangement of the inlet and ex-\nhaust valves in the cylinder head, a heat shield, two opposite\ncombustion chambers, two injection nozzles and a diffuser com-", "timestamp": "2026-07-19T11:22:45.741279+00:00"} | |
| {"citation_id": "19930094847", "source_url": "https://ntrs.nasa.gov/api/citations/19930094847/downloads/19930094847.pdf", "page_number": 22, "total_pages": 46, "image_filename": "19930094847_p22.jpg", "text": "N.A.C.A. Technical Memorandum No. 567\n\namount of the friction, is greatest in the stream saddle on the lower side and reaches a minimum at a point A (Fig. 14) in the right upper quadrant. This is represented by a crescent whose greatest width is at the saddle S. This illustrates the unequal effect of the friction in contrast with its theoretical replacement by an all-round uniform circulation.\n\nThe mechanism of the Magnus force is accordingly as follows. If no rotation and friction were present, the two lateral currents flowing around the cylinder at the beginning of the translatory motion would unite again in the middle of the rear side. Here in the field of the increasing friction, the friction layers are thrust toward the lower side by the rotation. By the resistance of the opposed lateral current its kinetic energy is partially transformed into pressure. This pressure reaches its maximum value in the stream saddle forward under the cylinder where the two motions, diagonal to each other, maintain the equilibrium and come to rest at a point. Thereby the whole stream, meeting the cylinder in front of the saddle, is forced to flow over the cylinder. On this path of the diminishing friction there is no obstructing counterflow. The frictional forces therefore have only an accelerating effect and thus increase the effect of the overpressure in the stream saddle, which dominates this portion of the circulation. The resultant effect of the friction therefore appears to be an increase in the velocity of the flow on the upper side of the cylinder.", "timestamp": "2026-07-19T11:22:48.265761+00:00"} | |
| {"citation_id": "19930094852", "source_url": "https://ntrs.nasa.gov/api/citations/19930094852/downloads/19930094852.pdf", "page_number": 5, "total_pages": 32, "image_filename": "19930094852_p5.jpg", "text": "N.A.C.A. Technical Memorandum No. 563\n4\n\nby the earlier experimenters are limited in all cases to a portion of the thrust components, and moments, the authors generalized the problem according to the behavior of a propeller in a side wind and undertook the task of investigating screw propellers of different pitch in the wind tunnel at angles of flow of 0 to 90°. In addition to the propulsive efficiency, all forces and moments in a wind-fast system and in a propeller-fast system of coordinates were to be investigated.\n\nThe Propellers Tested\n\nTwo propeller models of the same diameter D, but of different pitch H, were tested.\n\n1. An airplane propeller model of uniform pitch H = 0.5 D throughout the whole radius (Fig. 1). The propeller was of the type S₁ F₂ A₁ P₁ of the N.A.C.A. reports.*\n\n2. A helicopter propeller with a uniform pitch of H = 0.3 D over the greater (outer) part of the radius (Fig. 2).\n\nThe diameter of both propellers was D = 0.32 m (12.6 in.). They were made from a block of glued strips of wood, like full-size propellers. In reporting the results, the propellers are simply designated as I and II.\n\n*Durand and Lesley, \"Experimental Research on Air Propellers-V.\" N.A.C.A. Technical Report No. 141, 1925.", "timestamp": "2026-07-19T11:22:48.443287+00:00"} | |
| {"citation_id": "19930094857", "source_url": "https://ntrs.nasa.gov/api/citations/19930094857/downloads/19930094857.pdf", "page_number": 19, "total_pages": 68, "image_filename": "19930094857_p19.jpg", "text": "N.A.C.A. Technical Memorandum No. 557\n18\n\nd) The Testing Installation\n\nOn the whole, the original arrangement (Reference 2) was satisfactory. Its main part consists of a vertical graduated bar located behind the wing. A slide carrying the Pitot tubes for the measurement of the pressures in the turbulent zone can be moved along this bar from the observer's cockpit. The pressures are recorded by U-shaped alcohol pressure gauges of which motion pictures are taken. Before and during the present tests, a number of changes and improvements were made. They are briefly recorded below, since they embody, in part, the results of measuring tests.\n\n1. The new position of the Pitot tube for the measurement of the undisturbed flow (tube I) has already been mentioned in c, 2. The Pitot tube (Fig. 8) is mounted on a braced mast above and slightly to the right of the fuselage. When entering a hangar, the mast can be let down, after loosening a turnbuckle. The original arrangement of the forward Pitot tubes on the left wing, which showed a tendency to vibrate in harmony with the wing, was abandoned.\n\n2. In order to avoid the influence of the rail and of the tube holder, the support of the rear Pitot tubes (Fig. 9) was extended and reinforced. A new slide with a wide copper guide was provided and fitted without play, in order to fix the position of the Pitot tubes more accurately. The notches in the", "timestamp": "2026-07-19T11:22:49.735520+00:00"} | |
| {"citation_id": "19930094823", "source_url": "https://ntrs.nasa.gov/api/citations/19930094823/downloads/19930094823.pdf", "page_number": 17, "total_pages": 29, "image_filename": "19930094823_p17.jpg", "text": "N.A.C.A. Technical Memorandum No. 593\n16\n\nthe $c_n$ value 0.94 and\n\n$$\n\\left(\\frac{c_n}{c_w}\\right)_{\\text{max}} = \\frac{0.94}{0.08} = 10.65.\n$$\n\nSeparation of the boundary layer begins between $c_n = 1.2$ and 1.3 on all thick profiles and the lateral damping begins to diminish at that point. Flight is therefore unsafe in this region, if it is possible at all through pulling. This example shows how narrowly restricted is the range of glide with safe lateral stability after the best fineness ratio is exceeded.\n\nIf, on the other hand, a wing equipped with stabilizing slots at its tips can retain its lateral stability at angles of attack of 25 to 30°, the profile drag is then immensely increased by the separation of the flow, and gliding angles of 1/4 to 1/3 can be attained, whereby the gliding speed is approximately of the order of magnitude of the minimum speed. Stalling to such a large angle of attack can be effected by gradual throttling and simultaneously gradual pulling, without serious changes in the attitude of the airplane.\n\nOf course it is not generally possible to glide so near the ground, because the sinking speed is too great and there is no reserve lift for levelling off and resuming horizontal flight. The landing must take place as shown in Figure 19. The airplane glides at full stall to the point A. Then the angle of attack is greatly reduced by pushing (and possibly", "timestamp": "2026-07-19T11:22:49.943415+00:00"} | |
| {"citation_id": "19930094849", "source_url": "https://ntrs.nasa.gov/api/citations/19930094849/downloads/19930094849.pdf", "page_number": 9, "total_pages": 26, "image_filename": "19930094849_p9.jpg", "text": "N.A.C.A. Technical Memorandum No. 565\n\nscavenging, a lower volumetric efficiency (72% against 92% in the four-stroke cycle) must be taken.\n\nWe must not forget, however, that, in the two-stroke cycle, these quantities must be multiplied by 2 in determining the cylinder performance for a definite mean pressure. In our particular case the two-stroke cycle gives 55% greater performance on the basis of volumetric efficiency.\n\nIn the one-cylinder experimental engine, only the size of the charging manifold, the advance angle of the crank shaft of the exhaust pistons, the inclination of the scavenging channels and the scavenging pressure were systematically altered, and the influence on fuel consumption and maximum attainable performance were observed and measured. Figure 9 shows the effect of varying the rotation of the scavenging air on the combustion simply by varying the angle of incidence; a gain of 10% in performance is obtained.\n\nBy the proper dimensioning of the admission ports and correct conduction of air, it was possible to lower the scavenging pressure to 0.2 atmosphere at full power without impairing the combustion. Hand in hand with the expedients taken to improve the admission of air went the improvement of fuel injection.\n\nFrom a physical standpoint the task consists in bringing each fuel particle into contact with the requisite amount of combustion air in the shortest possible time. The difficulty of this lies in the shortness of the time available. In an oil en-", "timestamp": "2026-07-19T11:23:25.103243+00:00"} | |
| {"citation_id": "19930094855", "source_url": "https://ntrs.nasa.gov/api/citations/19930094855/downloads/19930094855.pdf", "page_number": 12, "total_pages": 13, "image_filename": "19930094855_p12.jpg", "text": "N.A.C.A. Technical Memorandum No.559\nFig.6\n\nFig.6 Vertical section through the axis of the engine shaft and the axis of the propeller. The etched portion of the mechanism (with or without crosshatching) does not revolve with the propeller shaft. 1, packing gland; 2, felt packing; 3, collar for regulating the play by acting on the conical end of the hub which has saw cuts 5; 4, upper spring for stopping the balls; 5, sleeve made of nitrided steel and provided with a helicoidal thread; 7, key; 8, metal blade; 9, balls on the helicoidal ramp; 10, sleeve for holding the root of the blade; 11, lower spring for stopping the balls; 12, nitrided-steel hub with helicoidal thread for receiving root of blade; 13, socket for head of pin 16; 14, centering piece for the blade root mounted on rollers; 15, hub bushing; 17, connecting rod articulated on pin 16 and on the collar 18 which turns with the propeller shaft; 19, slide block receiving, in three bronze grooves 20, the ends of the shafts 28 forming jacks 21; 22, ball bearings with deep grooves; 33, thrust bearing absorbing the reaction of the jack 21; 24, small toothed wheel keyed to shaft 25; 25, large transmission gear driven by 24 and in turn actuating the two small control wheels of the other two jacks not shown in the cut; 26, cover of the fixed casing; 27, felt packing between fixed and revolving portions; 28, shaft for controlling the pitch. At the bottom, cuts perpendicular to the axis of a blade, made at different distances from the axis of the propeller shaft. Just above the latter, ball-and-socket joint between connecting rod 17 and pin 18.\n\n[Figure: Technical drawing showing a vertical section of an engine shaft and propeller mechanism, with numbered parts and cross-sections labeled \"Section ½ Coupe BB\" and \"½ Coupe AA\". A small inset diagram is labeled \"Attache de la biellette sur la broche\".]", "timestamp": "2026-07-19T11:23:41.416446+00:00"} | |
| {"citation_id": "19930094824", "source_url": "https://ntrs.nasa.gov/api/citations/19930094824/downloads/19930094824.pdf", "page_number": 11, "total_pages": 21, "image_filename": "19930094824_p11.jpg", "text": "N.A.C.A. Technical Memorandum No. 592\n\nThe experiment shows briefly then, that the deflection of the lattice beam is about $15\\%$ less than that of the most rigid plate wall beam, and that the difference in strut tension in the plate wall beam amounts to less than $5\\%$ of that in the lattice beam.\n\nDiscussion\n\nProfessor H. Wagner: Dr. Mathar's report was enjoyable for two reasons: one, because it indicates the importance which the Aachen Institute lays on problems of stresses in buckled plates, and again, because the test data, particularly the behavior of stresses, agree pretty closely with my theoretical deliberations. However, I wish to make a few remarks about certain salient points. I noticed that the loading of the plate wall in these tests was always very low in comparison to its strength. And inasmuch as the limit case of the field of tension diagonal agrees so much more closely with actual conditions, as the stress becomes higher, one might suppose that the discrepancies between the calculated and the measured stiffness, as established by Dr. Mathar, would be still lower under higher loading.\n\nOne diagram of Dr. Mathar shows the effect of the spacing and the type of upright members on the beam stiffness. There the stiffness has a tendency to reach a limit value very quickly if the vertical members are not spaced too far, in which", "timestamp": "2026-07-19T11:23:45.195273+00:00"} | |
| {"citation_id": "19930090739", "source_url": "https://ntrs.nasa.gov/api/citations/19930090739/downloads/19930090739.pdf", "page_number": 26, "total_pages": 33, "image_filename": "19930090739_p26.jpg", "text": "N.A.C.A. Technical Memorandum No. 357 25\n\nmon to both combustion chambers, but the shapes of the combustion chambers, fuel jets and diffuser neck are logically adapted to one another. The combustion chambers are shallow with rectangular cross sections; the fuel jets are flat and fan-shaped; the neck is narrow and rectangular. Not only has more pains been taken to prevent the fuel from coming in contact with the walls of the combustion chambers, but the pure explosion drive is intentionally left, for the uniform distribution of the fuel in the combustion air, which Price thought desirable, can no longer be the question. The flat fuel jets and their enveloping vapors are, instead, bedded in air on all sides until the combustion begins and are first dispersed and mixed with the air as a result of the combustion and during their passage through the diffuser. This is shown in the mixed form of the indicator diagram. For an engine of 19\" cylinder bore, 23\" stroke and 300 R.P.M., the indicator diagram shows a pressure of 27 atm. at the end of the compression (Fig. 25). The first partial combustion, during which about 24% of the heat contained in the total fuel charge is liberated, raises the pressure to 39 atm. (1.45-fold), which is maintained, during a piston stroke of 6%, to the beginning of the expansion, while the remaining 70% of the total heat in the fuel is being liberated.\n\nIn Banner's engine, the injection begins 8.5° before the dead center. At this instant the air in the combustion cham-", "timestamp": "2026-07-19T11:23:46.220022+00:00"} | |
| {"citation_id": "19930094847", "source_url": "https://ntrs.nasa.gov/api/citations/19930094847/downloads/19930094847.pdf", "page_number": 23, "total_pages": 46, "image_filename": "19930094847_p23.jpg", "text": "N.A.C.A. Technical Memorandum No. 567 23\n\nThe centrifugal force, which is proportional to the square of this velocity, finally produces the low negative pressure on the upper side, which, together with the overpressure in the stream saddle, produces the Magnus force.\n\nRegarding the problem of the ratio of the positive and negative pressures, Prandtl first found the magnitude of the positive pressure on the under side of the cylinder to be equal to the dynamic pressure $(p = \\frac{\\rho}{2} \\times V^2)$ of the simple wind velocity $V$, where $\\rho$ represents the density of the fluid. The negative pressure depends on the velocity of the flow on the upper side of the cylinder. On the nonrotating cylinder this is theoretically $u = 2 V$ both above and below. Now, in order that this velocity on the under side may be zero at the center of dynamic pressure, it is assumed that the counter-circulation must also have the velocity $2V$. Therefore the velocity on the upper side is $4V$. From this it follows, on the assumption that the friction does not here come into consideration, that there must be at this point a pressure decrease of $\\frac{\\rho}{2} (4V)^2 = 16 \\frac{\\rho}{2} V^2$, equal to 16 times the amount of the simple dynamic pressure on the lower side. This produces a negative pressure 15 times the dynamic pressure.\n\nAgainst this method of calculation, it may be first objected that the peripheral velocity of the supplementary circulation does not need to be $u = 2V$, in order to produce the dynamic pressure of the simple wind velocity at the center of dy-", "timestamp": "2026-07-19T11:23:48.321328+00:00"} | |
| {"citation_id": "19930094779", "source_url": "https://ntrs.nasa.gov/api/citations/19930094779/downloads/19930094779.pdf", "page_number": 3, "total_pages": 57, "image_filename": "19930094779_p3.jpg", "text": "NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nTECHNICAL MEMORANDUM NO. 637\n\nTHE DEVELOPMENT, DESIGN AND CONSTRUCTION\nOF GLIDERS AND SAILPLANES*\n\nBy A. Lippisch\n\nINTRODUCTION\n\nThe Gliding and Soaring Flight movement is as old as history. The observation of bird flight must have shown the men of former centuries in the same way as our own generation, that apart from power flight with wing beats there must be another flight possibility, enabling the use of the energy in the movement of air masses for flight without the expenditure of other power. The experiments undertaken in those times on the solution of the flight problem have come down to us only through myths and sagas, and we cannot differentiate between truth and imagination in these stories.\n\nExperimental research, which can certainly be considered as the foundation of modern physics, has also in the realm of aerodynamics laid the basis for modern aeronautics. In this gliding and soaring flight plays the role of full-scale experiments, not as an end in itself, but as a proving ground and last station before the invention of power flight.\n\nThe actual gliding and sailing flying had its beginning in the sailing flight movement which took place after the war and which was a result of the Rhön Sailplane Contests.\n\nPART I\n\nSo if I commence my lecture with a few remarks about the development of the gliding and sailing flight movement, I will begin this outline with the successful sailplane of the first Rhön Sailplane Contest.\n\n*The Journal of The Royal Aeronautical Society, July, 1931, pp. 532-578.", "timestamp": "2026-07-19T11:23:49.878326+00:00"} | |
| {"citation_id": "19930094822", "source_url": "https://ntrs.nasa.gov/api/citations/19930094822/downloads/19930094822.pdf", "page_number": 4, "total_pages": 40, "image_filename": "19930094822_p4.jpg", "text": "N.A.C.A. Technical Memorandum No. 594\n3\n\nthe course of the $v_s$ curve, which is reversed for an airplane\nwith the slots open, at small dynamic pressures down to 36 kg/m²\n(7.37 lb./sq.ft.), in such a way that the $v_s$ values increase\nconsiderably, with increasing dynamic pressure from 36 to\n42 kg/m² (7.37 to 8.6 lb./sq.ft.), it is found that the maxi-\nmum lift of the airplane is reached and even exceeded.\n\nWith the slots closed, gliding flights can not be made at\n$q_w < 41$ kg/m² (8.4 lb./sq.ft.), because the increased difficulty\nof operating the airplane, due to the lessened damping and\ncontrollability about the longitudinal axis, renders impossible\nthe taking of measurements and the longer maintenance of a\ngiven state of equilibrium. With the slots open, however, the\ndamping about the longitudinal axis and the controllability,\neven in a state of equilibrium with the control stick pulled\nclear back, were still adequate for long glides at constant\ndynamic pressure.\n\nIn Figure 28, the angle of glide $\\varphi$, as determined from\nthe air speed and the sinking speed, is plotted against the\nactual dynamic pressure $q_w$. The difference in the size of\nthe gliding angle calls for the auto control slot on this\nairplane, because greater gliding angles can be attained at\nlower speeds with the slots open. The maximum angle of glide\nwith the slots closed was about 10.5°, while it was 13.5°\nwith the slots open. The airplane was found to be perfectly\ncontrollable with the slots open and to be able to maintain", "timestamp": "2026-07-19T11:23:51.162018+00:00"} | |
| {"citation_id": "19930094853", "source_url": "https://ntrs.nasa.gov/api/citations/19930094853/downloads/19930094853.pdf", "page_number": 16, "total_pages": 25, "image_filename": "19930094853_p16.jpg", "text": "N.A.C.A. Technical Memorandum No. 561\n15\n\nPractical Construction of the Free-Wheel Propeller\n\nThe preceding calculations are intended to show what can be expected from the use of a free-wheel propeller. We will now consider briefly how such a device can be made.\n\nThe following description relates to one method of making such a device. It may be possible to find ways for improving this device still further. It is interesting to note, however, that the first one made has already proved very satisfactory. It weighs only 10 to 12 kilograms additional for 500 hp at 1000 r.p.m., and 7 to 8 kg for the same power at 2000 r.p.m.\n\nThis device for releasing the propeller when the engine stops consists of interposing, between the engine shaft and the propeller, a so-called \"free wheel\" which enables the transmission of force in only one direction, namely, from engine to propeller and not in the opposite direction, that is, from propeller to engine.\n\nThe free wheel can be interposed between the crank shaft and the propeller hub, or even in the reduction gear (if the engine has one). We prefer the solution involving the interposition of a free wheel between the engine shaft and the propeller hub, thus forming a free-wheel hub which can be mounted on the end of the engine shaft like an ordinary hub.\n\nThe sectional drawing and photographs show the construction of the free-wheel hub with sufficient clearness.*\n\n*For a detailed description of this mechanism as applied to an automobile, see \"The de Lavaud Automatic Transmission\" in S.A.E. Journal, Vol. XXIII, pp.571-582, December, 1928.", "timestamp": "2026-07-19T11:23:51.823182+00:00"} | |
| {"citation_id": "19930094836", "source_url": "https://ntrs.nasa.gov/api/citations/19930094836/downloads/19930094836.pdf", "page_number": 10, "total_pages": 42, "image_filename": "19930094836_p10.jpg", "text": "N.A.C.A. Technical Memorandum No. 580\n9\n\nFigure 3 is a microphotograph of such a vibration and of the corresponding time marks. Before the tests, the spring constant was determined by loading the device with known weights and recording the resulting deflection (Table I).\n\nTABLE I. Determination of Spring Constant\n\n| No. | Load kg | Distance from base line 1/100 mm | $K = \\frac{P}{f}$ kg/cm | Mean K value kg/cm |\n| :--- | :--- | :--- | :--- | :--- |\n| 1 | 2.5 | 14.3 | 175.0 | |\n| 2 | 3.0 | 17.0 | 176.5 | |\n| 3 | 3.5 | 20.0 | 175.0 | |\n| 4 | 4.0 | 23.0 | 174.0 | |\n| 5 | 4.5 | 25.5 | 176.5 | |\n| 6 | 5.0 | 28.0 | 178.5 | |\n| 7 | 5.5 | 31.0 | 177.5 | 176.0 |\n| 8 | 6.0 | 34.0 | 176.2 | |\n| 9 | 6.5 | 37.0 | 175.9 | |\n| 10 | 7.0 | 40.0 | 175.0 | |\n| 11 | 7.5 | 42.5 | 176.5 | |\n\nThe mass of the instrument was then determined by causing it to vibrate in air. A comparison of the mass determined by vibration with that obtained by weighing showed that the steel springs participated in the vibrating mass of the device to an extent of 35.8% of their total mass. The vibration of the plates (the dimensions and weights of which are given in Table II) against water was then tested by placing the device over the water-filled tank shown in the background of Figure 2. The water surface was approximately 45 cm (18 in.) above the plate and did not seem to be affected by its vibration. At the point of immersion of the tube, a concentrically progressing undulatory motion of very small amplitude was observed. It was merely due", "timestamp": "2026-07-19T11:23:52.538666+00:00"} | |
| {"citation_id": "19930094823", "source_url": "https://ntrs.nasa.gov/api/citations/19930094823/downloads/19930094823.pdf", "page_number": 18, "total_pages": 29, "image_filename": "19930094823_p18.jpg", "text": "N.A.C.A. Technical Memorandum No. 593\n17\n\nby giving more gas), so that the point B is reached at normal gliding speed which enables the conventional levelling off and continued flight at gradually diminishing speed.\n\nThe requisite acceleration altitude and time are easily underestimated. The following example will show how to estimate them.\n\nG = gross weight.\n$c_w$ = drag coefficient with the subcritical $c_a$ value and angle of attack corresponding to the fineness ratio $\\epsilon_o$.\n$\\gamma$ = air density (assumed to be constant).\nF = wing area.\n$v_o$ = speed at the beginning of the acceleration at B.\n$v_1$ = speed at normal gliding flight.\ng = gravity constant.\n\nThe differential equation for the acceleration process is then\n\n$$ \\frac{G}{g} \\cdot \\frac{dv}{dt} = G \\cdot \\epsilon - c_w \\cdot F \\cdot \\frac{\\gamma}{2g} v^2 $$\n\nAfter introducing the constants $a = \\epsilon \\cdot g$ and $b = \\frac{c_w \\cdot \\gamma}{2 G/F}$, the integration yields the following results:\n\na) Time\n$$ t = \\sqrt{\\frac{1}{a \\cdot b}} \\left[ \\text{arc tan} \\left( v \\sqrt{\\frac{b}{a}} \\right) \\right]_{v_o}^{v_1} \\quad (1) $$", "timestamp": "2026-07-19T11:24:31.823557+00:00"} | |
| {"citation_id": "19930090740", "source_url": "https://ntrs.nasa.gov/api/citations/19930090740/downloads/19930090740.pdf", "page_number": 28, "total_pages": 33, "image_filename": "19930090740_p28.jpg", "text": "N.A.C.A. Technical Memorandum No. 358\nFigs. 26,28a,28b & 29\n\nDeutz\n\"Displacer\" engine\n[Figure: Cross-section diagram of an engine]\nFig. 26\n\n[Figure: Diagram of a cylinder component]\nFig. 28a\nRuston and Hornsby\n\nTurbais-Peugeot\n\nRuston and Hornsby\n[Figure: Diagram of a mechanical component]\nFig. 28b\n\n[Figure: Cross-section diagram of an engine]\nFig. 29", "timestamp": "2026-07-19T11:24:36.290801+00:00"} | |
| {"citation_id": "19930094779", "source_url": "https://ntrs.nasa.gov/api/citations/19930094779/downloads/19930094779.pdf", "page_number": 4, "total_pages": 57, "image_filename": "19930094779_p4.jpg", "text": "2 N.A.C.A. Technical Memorandum No. 637.\n\nFigure 1 shows the aircraft \"Schwarzer Teufel\" (Black Devil) of the Aachen Flying Club, designed by W. Klemperer.\n\nThe underlying idea of the design is on one side low structural weight and on the other side the greatest possible reduction of the parasite drag. Especially notable in this and other designs by Klemperer is the unusually carefully carried out structure. By this means Klemperer was able to attain an empty weight for this machine of 133 lb. (61 kg.) or a wing loading of 1.86 lb./sq.ft.\n\nThis type was not copied in the years following due to the fact that although the sinking speed was satisfactory, the gliding angle was too great. This matter depends to a considerable extent on the low-wing construction.\n\nThe sailplane which has to-day almost become a classic is the \"Vampyr\" (fig. 2) of the Flying Club of the Hannover Engineering School, which was designed by G. Madelung.\n\nIn this design we find for the first time the essential lines of thought clearly worked out in the design of a sailplane.\n\nThe problem is to build an aircraft with low sinking speed, good gliding angle, sufficient strength and good maneuverability.\n\nThe solution is: A cantilever high-wing type with a thick highly cambered wing section with large span and aspect ratio.\n\nIn this type a single-spar wing was used for the first time in which the torsional forces were taken by the leading edge. This was built up as a thin-walled tube, closed at the rear by the spar proper. This \"torsion-nose\" allowed at the same time of the possibility of keeping the most sensitive part of the wing section, the leading edge, the exactly correct shape. Plywood is the essential material which first made this construction possible.\n\nFurther notable characteristics of the \"Vampyr\" are: The three-part wing of the fuselage completely built of plywood and with a landing gear consisting of three football wheels, and the pendulum-type elevator.", "timestamp": "2026-07-19T11:24:36.888387+00:00"} | |
| {"citation_id": "19930094822", "source_url": "https://ntrs.nasa.gov/api/citations/19930094822/downloads/19930094822.pdf", "page_number": 5, "total_pages": 40, "image_filename": "19930094822_p5.jpg", "text": "N.A.C.A. Technical Memorandum No. 594\n\nits equilibrium for a long time. At a gliding angle of $10.5^\\circ$ with the slots closed, the damping about the longitudinal axis was appreciably diminished even for slight disturbances of the condition of equilibrium. A long gliding flight could be made only with difficulty.\n\nThe course of the elevator deflection $\\delta_{\\text{HR}}$, as plotted against the actual dynamic pressure in Figure 29, shows that the maximum elevator deflection with the slots open ($>15^\\circ$) is greater than with the slots closed (about $10.5^\\circ$) and at much lower $q_w$ values. The maximum elevator deflection in gliding flight with the slots open corresponded therefore with the maximum deflection of the control stick in the pulling direction.\n\nFor determining the maximum angle of attack, the inclination of the direction of flow to the longitudinal axis of the airplane, at a given point in front of the airplane, was taken as the criterion for the angle of attack. The measurements were made with a new instrument developed by the D.V.L. The angle of attack $\\alpha'$ (Fig. 30) is here designated as the angle between the direction of flow and the line of symmetry of the instrument at the given point. In Figure 30 the $\\alpha'$ values, thus obtained, are plotted against the actual dynamic pressure $q_w$ both with the slots open and with them closed. As was expected, all the measuring points indicate a common course throughout the whole angle-of-attack region, the only differ-", "timestamp": "2026-07-19T11:24:38.719207+00:00"} | |
| {"citation_id": "19930094855", "source_url": "https://ntrs.nasa.gov/api/citations/19930094855/downloads/19930094855.pdf", "page_number": 13, "total_pages": 13, "image_filename": "19930094855_p13.jpg", "text": "N.A.C.A. Technical Memorandum No. 559\n\nFigs. 7,8,9,10\n\nFig. 7\n\nFig. 8\n\nFig. 9\n\nView of rear side of hub showing the fixed housing of the mechanism for controlling the pitch. Note the three jack bearings with the control shaft projecting from one of them.\n\nFig. 10\n\nFigs. 7,8,9,10 Photographs. Assembly and parts of the Ratier metal propeller with pitch variable in flight.", "timestamp": "2026-07-19T11:24:40.595151+00:00"} | |
| {"citation_id": "19930094828", "source_url": "https://ntrs.nasa.gov/api/citations/19930094828/downloads/19930094828.pdf", "page_number": 4, "total_pages": 24, "image_filename": "19930094828_p4.jpg", "text": "N.A.C.A. Technical Memorandum No. 588 3\n\nmust be corrected, which, according to Figure 1, is quite simple. The true altitude is\n\n$$\nh = \\frac{\\Delta t}{2} \\sqrt{c^2 - v^2}.\n$$\n\nIf $v_0$ is the speed of the airplane upon which the design of the altitude scale of the sounder is based, the indicated altitude for a certain $\\Delta t$ becomes\n\n$$\nh = \\frac{1}{2} \\Delta t \\sqrt{c^2 - v_0^2},\n$$\n\nwhile it should be\n\n$$\nh_{soll} = \\frac{1}{2} \\Delta t \\sqrt{c^2 - v^2}.\n$$\n\nNow we have\n\n$$\n\\frac{h}{h_{soll}} = \\frac{\\sqrt{c^2 - v_0^2}}{\\sqrt{c^2 - v^2}}\n$$\n\nso that\n\n$$\nh - h_{soll} = \\frac{h_{soll} \\left( \\sqrt{c^2 - v_0^2} - \\sqrt{c^2 - v^2} \\right)}{\\sqrt{c^2 - v^2}}.\n$$\n\nThe difference of $h - h_{soll}$ yields the error $f$, for which the altitude, indicated on the sounder, must be corrected. Figure 2 is a graph which shows these speed errors from 20 to 60 m/s, and for the $h_{soll}$ values up to 100 m. This graph likewise shows that the air speed may be ignored when using the sounder. Its effect can still further be reduced by calibrating for a slightly higher speed ($v_0 = 20$ m/s).\n\nLastly, we examine the effect of the distance $a$ of both microphones on the height measurement.", "timestamp": "2026-07-19T11:24:40.919909+00:00"} | |
| {"citation_id": "19930094847", "source_url": "https://ntrs.nasa.gov/api/citations/19930094847/downloads/19930094847.pdf", "page_number": 24, "total_pages": 46, "image_filename": "19930094847_p24.jpg", "text": "N.A.C.A. Technical Memorandum No. 567 24\n\nnamic pressure, For this $u = V$ suffices, since the circulation is expected to counterbalance, at the center of dynamic pressure, only the simple wind strength, not its double, which is here no longer the case, due to the displacement of the center of dynamic pressure. However, if $u = V$, there is then, on the upper side, a resultant velocity of $U = 3V$ and consequently a theoretical pressure diminution of ninefold the amount of the simple dynamic pressure, so that the negative pressure would produce eight times, instead of fifteen times the dynamic pressure. Simultaneously the lift coefficient $(c_a)_{\\text{max}} = 4\\pi = 12.57$ calculated by Prandtl would drop to $2\\pi = 6.28$.\n\nIn reality, as we have seen, the circulatory motion produced by the decreasing friction cannot have the simple form like the theoretically uniform supplemental circulation. It was found, however, that it produces a motion over the after symmetrical half of the cylinder which is equal and opposite to the wind force $V$ and which produces the overpressure in the stream saddle. Since the same acceleration from the friction must be assumed over the forward symmetrical half of the cylinder, there is produced, together with the velocity of the simple potential flow on the upper side of the cylinder, the velocity $3V$ and therefrom, according to Prandtl's calculation, as above, a negative pressure eight times as large as the positive pressure on the lower side. Along with this summary of the results, however, the following observations should not be overlooked.", "timestamp": "2026-07-19T11:24:41.109092+00:00"} | |
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