PDF/ocr_dataset_1022.jsonl

{"citation_id": "19930085487", "source_url": "https://ntrs.nasa.gov/api/citations/19930085487/downloads/19930085487.pdf", "page_number": 3, "total_pages": 36, "image_filename": "19930085487_p3.jpg", "text": "NACA RM No. E9J22\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nRESEARCH MEMORANDUM\n\nVIBRATION SURVEY OF BLADES IN 10-STAGE  \nAXIAL-FLOW COMPRESSOR  \n\nI - STATIC INVESTIGATION  \n\nBy André J. Meyer, Jr. and Howard F. Calvert\n\nSUMMARY\n\nAn investigation was conducted to determine the cause of failures in the seventh- and tenth-stage blades of an axial-flow compressor. The natural frequencies of all rotor blades were measured and critical-speed diagrams were plotted. These data show that the failures were possibly caused by resonance of a first bending-mode vibration excited by a fourth order of the rotor speed in the seventh stage and a sixth order in the tenth stage.\n\nINTRODUCTION\n\nSeveral experimental jet-propulsion engines were constructed in which the six-stage axial-flow compressor of the engine was replaced with a ten-stage axial-flow compressor to increase thrust output. The other components of the engine were relatively unchanged. During thrust-stand tests conducted by the manufacturer, several blade failures occurred in early experimental 10-stage compressors.\n\nThe NACA Lewis laboratory conducted an investigation to determine the cause and to find, if possible, a means of preventing the blade failures in this engine. The first part of the investigation consisted in statically measuring the natural frequencies of each blade in the compressor and comparing these frequencies with possible exciting forces. First bending-, second bending-, and first torsional-mode frequencies were measured in all stages; second torsional- and third and fourth bending-mode frequencies were determined only in the first stage. The node shapes of higher modes of vibration were determined in an attempt to correlate the position of high-stress points for the various modes with the location of actual failures.", "timestamp": "2026-07-22T04:53:25.322980+00:00"}
{"citation_id": "19930082476", "source_url": "https://ntrs.nasa.gov/api/citations/19930082476/downloads/19930082476.pdf", "page_number": 25, "total_pages": 41, "image_filename": "19930082476_p25.jpg", "text": "CHART 7.- EFFECT OF INDIVIDUAL AND COMBINED RUDDER DEFLECTIONS ON THE SPIN CHARACTERISTICS OF MODEL\n(RUDDERS AND AILERONS UNLINKED)\n[Right erect spins; elevator set to 13° up, ailerons neutral, rudders set as indicated]\n\nA. Loading 2 ($\\frac{I_X - I_Y}{mb^2} = -4.9 \\times 10^{-4}$; $\\mu = 3.89$; loading 2 in table II and point 2 in fig. 4)\n\nB. Loading 3 ($\\frac{I_X - I_Y}{mb^2} = 165 \\times 10^{-4}$; $\\mu = 5.32$; loading 3 in table II and point 3 in fig. 4)\n\nC. Loading 4 ($\\frac{I_X - I_Y}{mb^2} = -1.8 \\times 10^{-4}$; $\\mu = 10.39$; loading 4 in table II and point 4 in fig. 4)\n\n| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | 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| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | 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{"citation_id": "19930082542", "source_url": "https://ntrs.nasa.gov/api/citations/19930082542/downloads/19930082542.pdf", "page_number": 13, "total_pages": 53, "image_filename": "19930082542_p13.jpg", "text": "12\nNACA TN No. 1867\n\ndepending on the amount of deformation and the time period considered. It is also quite certain that optimum treatments would be very different at higher temperatures than $1200^\\circ$ F.\n\nIn all the test work the number of tests was limited to the minimum necessary to indicate the level of properties for each treatment. Some of the reported strengths may be somewhat in error because of variations in test material. Precise determination of the properties for design purposes would require a more thorough testing program.\n\nRecommended Treatments for Gas-Turbine Discs\n\nThe problem of determining which treatments will produce the best material for the disc of a gas turbine is not clear even from the data available. The choice of any particular treatment would depend on the yield strength necessary at room temperature, the time period for rupture at $1200^\\circ$ F considered, and the ductility restrictions. The as-rolled stock when hot-cold-worked produced the highest yield strengths and rupture strength, except for 1000 hours for fracture, with the best ductility in the rupture test. In other words, for applications involving service of only a few hundred hours the hot-rolled and hot-cold-worked material had the best all-round properties. The difficulty with the finding is that in production it would probably be very difficult to control the hot-working conditions satisfactorily so that consistently uniform properties could be obtained, particularly long-time rupture ductility at $1200^\\circ$ F.\n\nThe best solution treatment alone appears to be about $2100^\\circ$ F. The yield strength at room temperature, however, will be below 40,000 psi. The stresses for rupture in 100 and 1000 hours should be about 51,000 and 40,000 psi.\n\nThe relative effectiveness of aging depends on prior treatment and on the rupture time at $1200^\\circ$ F considered. The major benefit appears to be increased ductility in the rupture test. The data indicate that an aging time of either 2 or 24 hours at $1350^\\circ$ to $1400^\\circ$ F is the best treatment although there is surprisingly little difference in properties when the aging temperature is as high as $1750^\\circ$ F. Aging at these temperatures should produce a yield strength at room temperature between 40,000 and 50,000 psi, a 100-hour rupture strength at $1200^\\circ$ F of about 51,000 psi, and a 1000-hour strength between 35,000 and 43,000 psi.\n\nThe highest possible strength at room temperature and in rupture tests at $1200^\\circ$ F is dependent on the presence of hot-cold-work. The hot-cold-working should be carried out at temperatures below $1400^\\circ$ F in order to develop maximum properties. A reduction of 10 to 15 percent in this temperature range is all that is required. Material processed", "timestamp": "2026-07-22T04:53:28.337269+00:00"}
{"citation_id": "19930082498", "source_url": "https://ntrs.nasa.gov/api/citations/19930082498/downloads/19930082498.pdf", "page_number": 17, "total_pages": 49, "image_filename": "19930082498_p17.jpg", "text": "```markdown\n16\n\nTABLE II.- RESULTS OF MUFFLER INVESTIGATION MADE WITH PROPELLER REMOVED - Continued\n\n| Muffler configuration | Engine speed (rpm) | Fundamental firing frequency, F (cps) | Sound-pressure level (db) | | | | | | | | | | | | | Other sounds | | | | Back pressure |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| (a) | | | Over-all | 0.5F | 1.0F | 1.5F | 2.0F | 2.5F | 3.0F | 3.5F | 4.0F | 4.5F | 5.0F | 5.5F | 6.0F | 6.5F | 7.0F | cps | db | cps | db | |\n| 21 [Figure: Muffler configuration 21] | 2000 | 100.0 | 87.0 | 45 | 72 | 61 | 87 | 60 | 63 | 56 | 67 | 58 | 58 | 53 | 57 | 52 | (b) | 435 | 66 | ---- | Medium |\n| 22 [Figure: Muffler configuration 22] | 1650 | 82.5 | 86.0 | 53 | 86 | 58 | 55 | 55 | 70 | 61 | 58 | 68 | 72 | (b) | (b) | (b) | (b) | ---- | ---- | ---- | Medium |\n| | 2000 | 100.0 | 84.5 | 60 | 80 | 60 | 50 | 58 | 68 | 68 | 80 | 65 | 60 | 55 | 66 | (b) | (b) | 330 | 77 | 265 | 65 | |\n| 23 Same as 22 except 16 1/2-inch tail pipe | 1650 | 82.5 | 85.5 | 53 | 85 | 56 | 60 | 62 | 81 | 70 | 69 | (b) | 68 | (b) | (b) | (b) | (b) | 330 | 82 | 750 | 62 | Medium |\n| | 2000 | 100.0 | 85.0 | 60 | 78 | 55 | 75 | 74 | 76 | 72 | 75 | 60 | 57 | 60 | 60 | (b) | (b) | | | | | |\n| 24 Same as 22 except 32 1/2-inch tail pipe | 1650 | 82.5 | 84.0 | (b) | 72 | 43 | 57 | 68 | 66 | 55 | 52 | 64 | 70 | (b) | (b) | (b) | (b) | ---- | ---- | ---- | Medium |\n| | 2000 | 100.0 | 86.0 | 56 | 73 | 56 | 61 | 69 | 73 | 70 | 83 | 67 | 60 | 61 | 68 | (b) | (b) | | | | | |\n| 25 Same as 22 except 50-inch tail pipe | 2000 | 100.0 | 83.5 | 56 | 71 | 55 | 73 | 64 | 77 | 70 | 65 | 68 | 57 | 57 | 57 | (b) | (b) | 335 | 77 | ---- | Medium |\n| 26 Same as 22 except 60-inch tail pipe | 2000 | 100.0 | 83.0 | 55 | 73 | 55 | 72 | 66 | 76 | 69 | 77 | 63 | 58 | 61 | 58 | (b) | (b) | 267 | 77 | ---- | Medium |\n| 27 Same as 22 except 70-inch tail pipe | 2000 | 100.0 | 84.5 | 57 | 74 | 55 | 73 | 62 | 73 | 74 | 65 | 60 | 60 | 58 | 65 | (b) | (b) | 368 | 81 | ---- | Medium |\n| 28 Same as 22 except 92 1/2-inch tail pipe | 2000 | 100.0 | 85.5 | 57 | 82 | 57 | 76 | 63 | 79 | 72 | 81 | 65 | 60 | 57 | 70 | (b) | (b) | 330 | 72 | ---- | Medium |\n| 29 Muffler 22 reversed with 32-inch tail pipe | 2000 | 100.0 | 83.0 | 60 | 72 | 63 | 65 | 67 | 77 | 83 | 70 | 60 | 58 | 55 | 57 | (b) | (b) | ---- | ---- | ---- | Medium |\n| 30 [Figure: Muffler configuration 30] | 2000 | 100.0 | 86.0 | 56 | 75 | 62 | 86 | 60 | 73 | 69 | 67 | 61 | 58 | 55 | 55 | (b) | 60 | 33.5 | 65 | ---- | High |\n| 31 [Figure: Muffler configuration 31] | 1650 | 82.5 | 82.0 | (b) | 81 | 55 | 63 | 53 | 55 | 52 | (b) | (b) | (b) | (b) | (b) | (b) | (b) | ---- | ---- | ---- | High |\n| | 2000 | 100.0 | 82.0 | (b) | 77 | 53 | 58 | 55 | 67 | 53 | (b) | (b) | (b) | (b) | (b) | (b) | (b) | ---- | ---- | ---- | |\n\n$^a$In the sketches of the configurations, all dimensions are in inches and all cross sections are circular, except where otherwise indicated.\n$^b$The sound-pressure level was below the range of the analyzer.\n\nNACA\nNACA TN No. 1858\n```", "timestamp": "2026-07-22T04:53:29.794266+00:00"}
{"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 11, "total_pages": 78, "image_filename": "19930082618_p11.jpg", "text": "NACA TN 1945\n\nThe drag polars for the airfoils tested with roughened leading edges do not generally have a range of lift coefficient over which the drag is essentially constant but, rather, are of parabolic form. The lower portions of the parabolas, over which the drag variation with lift coefficient is the least, appear to become narrower for most of the airfoils as the Reynolds number is reduced, except at the lowest Reynolds number ($0.7 \\times 10^6$). The behavior of the drag polars at a Reynolds number of $0.7 \\times 10^6$ probably results from the fact that the leading-edge roughness employed was not sufficiently large to cause fully developed turbulent boundary layers at this low value of the Reynolds number. In most cases, airfoil thickness and camber do not appear to have a very pronounced or consistent effect upon the manner in which the lift-coefficient range corresponding to the lower portion of the drag polars of the rough NACA 6-series sections varies with the Reynolds number, or upon the actual width of the range itself. Movement of the position of minimum pressure on the basic thickness form at zero lift from 40 percent to 60 percent of the chord does, however, seem to reduce the width of the range at most Reynolds numbers. The data for the NACA 4- and 5-digit-series sections (figs. 11 to 15, part (c)) show that the range of lift coefficient corresponding to the lower portion of the drag polars for these airfoils in the rough condition does not differ greatly at most Reynolds numbers from that shown by most of the NACA 6-series sections.\n\nMinimum drag.- The Reynolds number has a very important effect upon the minimum drag (figs. 1 to 15, part (c)) of the airfoils, both in the smooth condition and with roughened leading edges. In order to show more clearly the magnitude and trend of the effect, the drag coefficient corresponding to the measured design lift coefficient (designated minimum drag coefficient) has been plotted in figure 16 as a function of Reynolds number for each of the 15 airfoils tested. For convenience in comparing the drag variation of the different airfoils, the data for the NACA 6-series airfoils are arranged in this plot in three groups according to systematic variations of thickness, camber, and thickness distribution. The data for the NACA 4- and 5-digit-series sections are plotted in one group. The drag coefficient at the experimental design lift coefficient is seen to increase with decreasing Reynolds number for all the airfoils in the smooth condition (fig. 16(a)) and, except at the lowest Reynolds number, for the airfoils with roughened leading edges (fig. 16(b)). The previously mentioned effect of roughness size is probably responsible for the drag reduction shown by the results for the rough airfoils at a Reynolds number of $0.7 \\times 10^6$.\n\nFor the smooth NACA 6-series airfoils, the amount by which the minimum drag coefficient increases as the Reynolds number is lowered appears to become larger as the thickness ratio of the sections is increased and as the position of minimum pressure is moved rearward along", "timestamp": "2026-07-22T04:53:35.287771+00:00"}
{"citation_id": "19930082617", "source_url": "https://ntrs.nasa.gov/api/citations/19930082617/downloads/19930082617.pdf", "page_number": 12, "total_pages": 58, "image_filename": "19930082617_p12.jpg", "text": "NACA TN 1962\n11\n\n**Cylinder 72**\n*   **Main Diagram:**\n    *   Top dimension: 57.9\"\n    *   Side dimension: 20\"\n    *   Bottom dimensions: 34.74\", 3.86\"\n    *   Vertical bands labeled: A B C D E F G H I J K L M N O\n*   **Section through band H:**\n    *   Circular diagram with points numbered 1 through 16.\n    *   Angle marked: $45^\\circ$\n*   **Ring section:**\n    *   Dimensions: $\\frac{1}{2}\"$, $\\frac{1}{8}\"$\n\n**Cylinder 73**\n*   **Main Diagram:**\n    *   Top dimension: 57.9\"\n    *   Side dimension: 20\"\n    *   Bottom dimensions: 34.74\", 3.86\"\n    *   Vertical bands labeled: A B C D E F G H I J K L M N O\n*   **Section through band H:**\n    *   Circular diagram with points numbered 1 through 16.\n    *   Angle marked: $45^\\circ$\n*   **Ring section:**\n    *   Dimensions: $\\frac{1}{2}\"$, $\\frac{1}{4}\"$\n\n**Cylinder 74**\n*   **Main Diagram:**\n    *   Top dimension: 59.11\"\n    *   Side dimension: 20\"\n    *   Bottom dimensions: 33.41\", 2.57\"\n    *   Vertical bands labeled: A B C D E F G H I J K L M N O P Q R S T U V W\n*   **Section through band L:**\n    *   Circular diagram with points numbered 1 through 16.\n    *   Angle marked: $45^\\circ$\n*   **Ring section:**\n    *   Dimensions: $\\frac{1}{2}\"$, $\\frac{1}{4}\"$\n\n**Cylinder 75**\n*   **Main Diagram:**\n    *   Top dimension: 59.11\"\n    *   Side dimension: 20\"\n    *   Bottom dimensions: 33.41\", 2.57\"\n    *   Vertical bands labeled: A B C D E F G H I J K L M N O P Q R S T U V W\n*   **Section through band L:**\n    *   Circular diagram with points numbered 1 through 16.\n    *   Angle marked: $45^\\circ$\n*   **Ring section:**\n    *   Dimensions: $\\frac{3}{8}\"$, $\\frac{3}{8}\"$\n\n[NACA Logo]\n\nPosition of strain gages\n\n(a) Cylinders 72 to 75.\n\nFigure 1.- Schematic drawings of test specimens.", "timestamp": "2026-07-22T04:53:35.592201+00:00"}
{"citation_id": "19930082245", "source_url": "https://ntrs.nasa.gov/api/citations/19930082245/downloads/19930082245.pdf", "page_number": 37, "total_pages": 66, "image_filename": "19930082245_p37.jpg", "text": "```markdown\n1.6\n1.4\n1.2\n1.0\n.8\n.6\n.4\n.2\n0\n-.2\n-.4\n-.6\n-.8\n.1 .2 .3 .4 .5 .6 .7 .8 .9\nMach number, M\nAileron section normal-force coefficient, $c_{n_a}$\n$\\delta_a$ (deg)\n30\n18\n12\n4\n2\n0\n-2\n-4\n-6\n-12\n\n.16\n.12\n.08\n.04\n0\n-.04\n-.08\n-.12\n-.16\n-.20\n-.24\n-.56\n-.60\n.1 .2 .3 .4 .5 .6 .7 .8 .9\nMach number, M\nAileron section hinge-moment coefficient, $c_h$\n$\\delta_a$ (deg)\n-12\n-6\n-4\n-2\n0\n2\n4\n12\n18\n30\nNACA\n\n(d) $c_n = 0.2$.\nFigure 7 .-Continued.\n\n36\nNACA TN No. 1596\n```", "timestamp": "2026-07-22T04:53:36.545871+00:00"}
{"citation_id": "19930085838", "source_url": "https://ntrs.nasa.gov/api/citations/19930085838/downloads/19930085838.pdf", "page_number": 118, "total_pages": 118, "image_filename": "19930085838_p118.jpg", "text": "116\nNACA RM No. L9B23\n\n<!-- Image (116, 110, 839, 999) -->\n\n(a) 0.177c thick airfoil.\n\n(b) 0.154c thick airfoil.\n\nFigure 23.- Hinge-moment characteristics of a flap on two NACA 7-series-type airfoils with double slotted flap and straight-sided Frise aileron. $\\alpha_0 = 0^\\circ$; $\\delta_a = 0^\\circ$; $R = 6.0 \\times 10^6$ (approx.).", "timestamp": "2026-07-22T04:53:37.327468+00:00"}
{"citation_id": "19930085519", "source_url": "https://ntrs.nasa.gov/api/citations/19930085519/downloads/19930085519.pdf", "page_number": 2, "total_pages": 46, "image_filename": "19930085519_p2.jpg", "text": "NACA RM No. L8K19\nRESTRICTED\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nRESEARCH MEMORANDUM\n\nWIND-TUNNEL INVESTIGATION AT LOW SPEEDS OF VARIOUS\nPLUG-AILERON AND LIFT-FLAP CONFIGURATIONS\nON A 42° SWEPTBACK SEMISPAN WING\n\nBy Leslie E. Schneiter and James M. Watson\n\nSUMMARY\n\nA wind-tunnel investigation has been performed on a 42° sweptback-wing model to determine the lateral control characteristics of a plug-aileron configuration consisting of six segments extending from the wing 20-percent-span to the wing 80-percent-span stations and placed perpendicular to the free-stream flow with the center of each plug segment on the wing 70-percent-chord line. The basic plug aileron and several modifications thereof were investigated for a range of plug projections through a large angle-of-attack range. In addition, several types of lift flaps were investigated and a full-span slotted-flap configuration was developed. The plug-aileron characteristics were determined with the full-span slotted flap in the optimum location. The lateral control characteristics of a partial-span plain aileron were also determined for comparison with the plug-aileron results.\n\nOf the various flap configurations investigated on this wing (full-span slotted flap at deflections of 30°, 40°, and 50°, a half-span slotted flap at 50° deflection, and a half-span split flap and a half-span Zap flap both at 60° deflection), the full-span slotted flap at 30° deflection gave the most satisfactory calculated trimmed-gliding characteristics for an airplane with an assumed wing loading of 40 pounds per square foot and a tail length of 3.0 mean aerodynamic chords.\n\nThe results show that the plug aileron investigated with the faired plug-slot lower lip gave positive rolling-moment coefficients at all projections throughout the angle-of-attack range investigated, although there was a large reduction in rolling-moment coefficient at all projections at angles of attack above the wing-tip stall angle. The maximum values of rolling-moment coefficient produced by the plug aileron with the faired lower lip were about 130 percent larger with the full-span slotted flap deflected than with the flap neutral.\n\nRESTRICTED", "timestamp": "2026-07-22T04:53:37.576959+00:00"}
{"citation_id": "19930082450", "source_url": "https://ntrs.nasa.gov/api/citations/19930082450/downloads/19930082450.pdf", "page_number": 20, "total_pages": 37, "image_filename": "19930082450_p20.jpg", "text": "NACA TN No. 1778\n19\n\n[Figure: A diagram showing a cross-section of a panel with multiple stiffeners. A magnified view of a single stiffener section is shown below, with various dimensions labeled.]\n\n[Diagram labels:]\n$b_A$\n$S$\n$\\bar{h}$\n$r_A$\n$t_S$\nAxis of c.g.\n$t_W$\n$b_W$\n$H$\n$r_F$\n$b_S$\n$b_F$\nNACA\n\nFigure 1. - Symbols for panel dimensions.", "timestamp": "2026-07-22T04:53:38.502594+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 22, "total_pages": 99, "image_filename": "19930082511_p22.jpg", "text": "20\nNACA TN No. 1826\n\nCASE 1 - CLOSED-OPEN TUNNEL\n\nTotal perturbation velocity. - By the transformation\n$$z = e^{\\pi \\zeta} \\quad (1)$$\nthe tunnel in the $\\zeta$-plane, represented by an infinitely long strip of unit height, is transformed to the upper half of the $z$-plane. The correspondence between points is shown in figure 13.\n\nThe complex velocity (rather than the more usual complex potential) is considered to be retained in the transformation, and the problem is thus to find a function $Q(z)$, where\n$$u - iv = Q(z) = q(\\zeta)$$\nsuch that\n(1) On the closed sections of the boundary, that is, for $z$ real and $|z| < 1$,\n$$I.P. Q(z) = 0$$\n(2) On the open sections of the boundary, that is, for $z$ real and $|z| > 1$,\n$$R.P. Q(z) = 0$$\n(3) For $z = \\pm 1$,\n$$Q(z) = 0$$\n(4) $Q(z)$ is finite at infinity\n\nConsider the complex velocity $G(z)$ corresponding to a vortex at $z_1$ and its reflection at $\\bar{z}_1$:\n$$G(z) = i \\left( \\frac{1}{z - z_1} - \\frac{1}{z - \\bar{z}_1} \\right) \\quad (2)$$\nThis function, which is of order $1/z^2$ at infinity, satisfies conditions (1) and (4) but not (2) and (3). Functions of the form $\\sqrt{1 - z^2}$, $z\\sqrt{1 - z^2}$, ... satisfy conditions (1), (2), and (3). At infinity, $\\sqrt{1 - z^2}$ is of order $z$ and $z\\sqrt{1 - z^2}$ is of order $z^2$; therefore,", "timestamp": "2026-07-22T04:53:39.066552+00:00"}
{"citation_id": "19930082485", "source_url": "https://ntrs.nasa.gov/api/citations/19930082485/downloads/19930082485.pdf", "page_number": 29, "total_pages": 62, "image_filename": "19930082485_p29.jpg", "text": "28\nNACA TN No. 1810\n\nTABLE II - POSITION OF BLADE PRESSURE TAPS\n\n[Figure: Diagram showing an airfoil cross-section with coordinate axes X and h. Points $X_1$ and $X_2$ are indicated on the chord line.]\n\n| Root section | | | Pitch section | | | Tip section | | |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| h | $X_1$ | $X_2$ | h | $X_1$ | $X_2$ | h | $X_1$ | $X_2$ |\n| (in.) | (in.) | (in.) | (in.) | (in.) | (in.) | (in.) | (in.) | (in.) |\n| -0.050 | 0 | 0 | -0.050 | 0 | 0 | -0.050 | 0 | 0 |\n| .200 | .150 | | .200 | .149 | | .200 | .148 | |\n| .400 | .178 | | .400 | .173 | | .400 | .168 | |\n| .700 | .119 | | .700 | .108 | | .700 | .097 | |\n| 1.000 | -.114 | | 1.000 | -.115 | | 1.000 | -.114 | |\n| 1.300 | -.614 | | 1.300 | -.543 | | 1.300 | -.472 | |\n| 1.435 | -.970 | | 1.435 | -.934 | | 1.435 | -.890 | |\n| 1.300 | | -.835 | 1.300 | | -.770 | 1.300 | | -.706 |\n| 1.100 | | -.570 | 1.100 | | -.532 | 1.100 | | -.497 |\n| .800 | | -.316 | .800 | | -.299 | .800 | | -.281 |\n| .500 | | -.172 | .500 | | -.162 | .500 | | -.152 |\n| .300 | | -.110 | .300 | | -.105 | .300 | | -.100 |\n\nNACA\n1026", "timestamp": "2026-07-22T04:53:40.359634+00:00"}
{"citation_id": "19930082585", "source_url": "https://ntrs.nasa.gov/api/citations/19930082585/downloads/19930082585.pdf", "page_number": 13, "total_pages": 30, "image_filename": "19930082585_p13.jpg", "text": "12\nNACA TN 1907\n\nAt zero time (the instant of power failure) all the quantities in the above equations (12a), (13a), and (14) are known. The step-by-step process is started by computing from these the values of $V_0$, $\\dot{\\Omega}_0$, and $\\ddot{\\Omega}_0$. After a short interval of time, say $\\Delta t$ seconds, the new value of rotor angular velocity $\\Omega_{\\Delta t}$ may be found approximately from the abbreviated Taylor series:\n\n$$\n\\Omega_{\\Delta t} = \\Omega_0 + (\\Delta t)\\dot{\\Omega}_0 + \\frac{(\\Delta t)^2}{2}\\ddot{\\Omega}_0 \\tag{15}\n$$\n\nIt is assumed that the change in $V$ in the period $\\Delta t$ is the time $\\Delta t$ times the average rate of change of $V$ in the time interval $\\Delta t$. Thus,\n\n$$\nV_{\\Delta t} = V_0 + \\frac{\\Delta t}{2}(\\dot{V}_0 + \\dot{V}_{\\Delta t}) \\tag{16}\n$$\n\nFrom equation (12a),\n\n$$\n\\dot{V}_{\\Delta t} = g - \\frac{\\rho abcR^3}{2W/g} \\left[ \\frac{\\dot{\\Omega}_{\\Delta t}}{2R} (V_{\\Delta t} - V_{\\Delta t}) + \\frac{\\theta_{\\Delta t}}{3} \\Omega_{\\Delta t}^2 \\right] \\tag{17}\n$$\n\nSolving equations (16) and (17) simultaneously for $\\dot{V}_{\\Delta t}$ (the unknowns being $\\dot{V}_{\\Delta t}$ and $V_{\\Delta t}$),\n\n$$\n\\dot{V}_{\\Delta t} = \\frac{g - \\frac{\\rho abcR^3}{2W/g} \\left[ \\frac{\\dot{\\Omega}_{\\Delta t}}{2R} \\left( V_0 + \\frac{\\Delta t}{2}\\dot{V}_0 - V_{\\Delta t} \\right) + \\frac{\\theta_{\\Delta t}}{3} \\Omega_{\\Delta t}^2 \\right]}{1 + \\Delta t \\frac{\\rho abcR^2}{8W/g} \\dot{\\Omega}_{\\Delta t}} \\tag{18}\n$$\n\nwhence $V_{\\Delta t}$ may be found from equation (16).", "timestamp": "2026-07-22T04:53:46.437729+00:00"}
{"citation_id": "19930082496", "source_url": "https://ntrs.nasa.gov/api/citations/19930082496/downloads/19930082496.pdf", "page_number": 21, "total_pages": 50, "image_filename": "19930082496_p21.jpg", "text": "20\nNACA TN No. 1856\n\nNACA\n\nTABLE IV - PHASE 2 OF QUASI-SERVICE TURBINE-BLADE EVALUATION\n\n<!-- Table (172, 109, 930, 937) -->\n\\begin{tabular}{|c|c|c|c|c|c|c|c|c|c|c|c|c|c|}\n\\hline\n\\multirow{3}{*}{Run} & \\multicolumn{2}{c|}{Operating} & \\multicolumn{2}{c|}{Cumulative} & \\multicolumn{2}{c|}{Conditions} & \\multirow{3}{*}{\\shortstack{Metal \\\\ blades \\\\ cracked \\\\ and replaced}} & \\multirow{3}{*}{\\shortstack{Metal \\\\ blades \\\\ completely \\\\ fractured \\\\ and re- \\\\ placed}} & \\multirow{3}{*}{\\shortstack{Total \\\\ metal \\\\ blades \\\\ re- \\\\ placed}} & \\multirow{3}{*}{\\shortstack{Cumula- \\\\ tive \\\\ total \\\\ metal \\\\ blades \\\\ re- \\\\ placed}} & \\multirow{3}{*}{\\shortstack{Cumula- \\\\ tive \\\\ total \\\\ control- \\\\ blades \\\\ frac- \\\\ tured}} & \\multirow{3}{*}{\\shortstack{Ceramal \\\\ blades \\\\ frac- \\\\ tured}} \\\\\n\\cline{2-7}\n& \\multicolumn{2}{c|}{time} & \\multicolumn{2}{c|}{operating} & \\multirow{2}{*}{\\shortstack{Tip \\\\ speed \\\\ (rpm)}} & \\multirow{2}{*}{\\shortstack{Indi- \\\\ cated \\\\ inlet \\\\ gas \\\\ tem- \\\\ pera- \\\\ ture \\\\ ($^\\circ$F)}} & & & & & & \\\\\n\\cline{2-5}\n& (hr) & (min) & (hr) & (min) & & & & & & & & \\\\\n\\hline\n\\multirow{4}{*}{1} & & 30 & & 30 & 10,000 & 478 & & & & & & \\\\\n\\cline{2-7}\n& & 30 & 1 & 1 & 10,000 & 478 & & & & & & \\\\\n\\cline{2-7}\n& & 30 & 1 & 2 & 10,000 & 478 & & & & & & \\\\\n\\cline{2-7}\n& & 30 & 2 & 4 & 10,000 & 478 & & & & & & \\\\\n\\hline\n\\multirow{4}{*}{2} & 2 & 30 & 4 & 34 & 10,000 & 2000 & & 2 control & 6 & 6 & 2 & \\\\\n\\cline{2-7}\n& 1 & 30 & 5 & 4 & 12,500 & 596 & & 4 original & & & & \\\\\n\\cline{2-7}\n& 1 & 20 & 6 & 24 & 12,500 & 596 & & & & & & \\\\\n\\cline{2-7}\n& & 20 & 7 & 44 & 15,000 & 716 & & 4 original & 13 & 19 & 3 & \\\\\n\\hline\n\\multirow{2}{*}{3} & & 46 & 8 & 30 & 15,000 & 716 & \\shortstack{8 original \\\\ 1 control} & 4 original & 17 & 36 & 3 & \\\\\n\\cline{2-7}\n& & 20 & 8 & 50 & 15,000 & 716 & 6 original & 2 original & 8 & 44 & 3 & \\\\\n\\hline\n\\multirow{2}{*}{4} & & 46 & 9 & 36 & 15,000 & 716 & \\shortstack{9 original \\\\ 3 control} & 7 original & 19 & 63 & 6 & 1 broken \\\\ during \\\\ overhaul \\\\\n\\cline{2-7}\n& & 55 & 10 & 31 & 15,000 & 716 & \\shortstack{12 original \\\\ 2 control} & 4 original & 18 & 81 & 8 & \\\\\n\\hline\n\\multirow{2}{*}{5} & 1 & 3 & 11 & 34 & 15,000 & 716 & \\shortstack{1 replacement \\\\ (installed \\\\ after run 5)} & 11 original & 12 & 93 & 8 & \\\\\n\\cline{2-7}\n& & 33 & 12 & 7 & 15,000 & 716 & \\shortstack{8 original \\\\ 7 replacement \\\\ (installed \\\\ after runs 1, \\\\ 3, and 6)} & 4 original & 15 & 108 & 8 & \\shortstack{1 service \\\\ failure \\\\ 1 destroyed \\\\ by dovetail \\\\ failure \\\\ in wheel} \\\\\n\\hline\n\\end{tabular}", "timestamp": "2026-07-22T04:53:50.646779+00:00"}
{"citation_id": "19930085471", "source_url": "https://ntrs.nasa.gov/api/citations/19930085471/downloads/19930085471.pdf", "page_number": 5, "total_pages": 28, "image_filename": "19930085471_p5.jpg", "text": "NACA RM No. L8J11\nUNCLASSIFIED\nRESTRICTED\nCONFIDENTIAL\n3\n\nA brief discussion is given of the effects of concentrated masses placed at the wing tip, the center of gravity of the masses coinciding with the center of gravity of the wing, and the effects of sharp and blunt leading edges on the wings.\n\nAPPARATUS AND TEST METHODS\n\nThe tests were made at a Mach number of 1.3 in an \"intermittent\" two-dimensional supersonic tunnel having a 9.24-inch by 18.23-inch test section (Figs. 1 and 2). The apparatus operates from atmospheric pressure to a vacuum. A quick-operating valve allows a steady-flow condition to be reached in approximately 0.15 second and this condition of steady flow can be maintained for 5 to 7 seconds. To prevent condensation in the test section, a room was constructed at the tunnel entrance in which the air could be heated. Variation of the air temperature from $170^\\circ$ F to $220^\\circ$ F results in a velocity range at the test section from 950 miles per hour to 990 miles per hour at a Mach number of 1.3. The test-section Mach number determined by optical means varied from 1.29 to 1.31. The test-section Mach number determined by a pressure survey showed a variation from 1.31 to 1.33 (fig. 3).\n\nThe models were mounted cantilever fashion from the side of the tunnel (fig. 4). In order to avoid possible model failure during the tunnel transient conditions, and since supersonic flutter characteristics were being determined, it was desirable to withhold the model from the flow during the periods of acceleration and deceleration. A pneumatic-cylinder arrangement was installed for this purpose (fig. 2). By using this device the model could be held outside the tunnel wall until stable flow conditions were reached. Then, by means of electrically operated valves, controlled by an observer, the model could be allowed to enter the tunnel slowly and to withdraw quickly if necessary.\n\nThe flutter models having rectangular plan forms were constructed of laminated Sitka spruce and also of duralumin. The wing dimensions and parameters are given in table I. The chords ranged from 3.03 to 4.22 inches and the lengths ranged from 6 to 9.12 inches. Both thick and thin sections with blunt and sharp leading edges were used. The airfoil sections used were 3-, 5-, and 8-percent-thick circular arcs, 3-percent-thick double wedge, NACA 16-010, and NACA 65-007. The mass-density parameter $1/\\kappa$ ranged from 52 to 268, the center of gravity ranged from 46 to 63 percent chord, and the elastic axis ranged from 34 to 52 percent chord.\n\nBefore each model was installed in the tunnel, its weight, moment of inertia, and section center-of-gravity position were determined. After installation in the tunnel the elastic axis was located by observing the\n\nCONFIDENTIAL\nUNCLASSIFIED", "timestamp": "2026-07-22T04:53:51.390130+00:00"}
{"citation_id": "19930082614", "source_url": "https://ntrs.nasa.gov/api/citations/19930082614/downloads/19930082614.pdf", "page_number": 12, "total_pages": 36, "image_filename": "19930082614_p12.jpg", "text": "```markdown\n10\nNACA TN 1939\n\nGraphical Analysis\n\nAs the number of variables in the equation of motion is increased,\nit becomes impractical to obtain an accurate solution by direct inte-\ngration of the equation of motion along the flight path. Because of\nthe arbitrary manner in which some of the quantities involved can be\nvaried, a method of calculation in which specific variations of these\nquantities are assumed would place untenable restrictions on the use-\nfulness of the solution. Among the quantities which cannot be defined\nmathematically by expressions suitable for all problems are the follow-\ning:\n\n(a) Variation of engine output with time\n\n(b) Variations of drag and thrust with Mach number\n\n(c) Variation of flight-path angle with time\n\n(d) Variation of drag with lift\n\n(e) Variation of drag with time when the air brakes are\noperated\n\nThese variations, which are either arbitrary or empirical, and the\nvariation of atmospheric density with time can all be taken into\naccount if the calculations are made by graphical means. The calcula-\ntions can be performed by a step-by-step method in which the longitudi-\nnal acceleration in each step is obtained from graphs that have been\nprepared. The change in velocity can then be ascertained as the inte-\ngral of the acceleration with respect to time. This integral may be\nevaluated graphically, but it has been found that the integration is\nsufficiently accurate if trapezoidal elements of area are assumed and\nan arithmetical evaluation is used.\n\nConstant flight-path inclination.— The longitudinal acceleration\n(or deceleration) at the beginning of the interval of time being\nstudied is determined by the known or assumed initial flight conditions.\nThe acceleration at any later time, however, cannot be computed\ndirectly, since it is a function of speed and altitude, the values of\nwhich are not known. This acceleration may be found by dividing the\ntotal time interval into a series of increments and calculating, step\nby step, the conditions at the end of each increment. The step-by-step\ncalculation is made by estimating the speed and altitude at the end of\neach increment upon the basis of the known conditions at the beginning\nof that increment. With these estimated values, the longitudinal accel-\neration at the end of the increment may be computed. The speed and\naltitude at this time may then be calculated to greater accuracy and\ncompared with the estimated values. Although this procedure amounts to\na method of successive approximations, it has been found that the\n```", "timestamp": "2026-07-22T04:53:52.954128+00:00"}
{"citation_id": "19930082914", "source_url": "https://ntrs.nasa.gov/api/citations/19930082914/downloads/19930082914.pdf", "page_number": 8, "total_pages": 66, "image_filename": "19930082914_p8.jpg", "text": "NACA TN No. 1857\n\nInterferometers are quite delicate instruments, as far as adjustment is concerned, and are quite subject to being thrown out of adjustment by mechanical vibrations and by temperature changes. When the present instrument was designed it was hoped to produce an instrument that would not cause an inordinate waste of time in keeping it in adjustment. The result has been quite satisfactory. Apparently because of its symmetrical design (and probably also because cast iron has a low coefficient of thermal expansion), no effects of temperature changes on the adjustment of the interferometer have ever been noticed. Ordinarily the interferometer is enclosed in a sheet-aluminum case. Even when it is used with one side of the case removed and the temperature in the room is changed by opening the windows, or when the operators stand close to the interferometer, no change in the fringes is noticed.\n\nIn an effort to keep the interferometer in adjustment despite vibrations, it was at first suspended from a framework by springs that gave the system a natural frequency of about 1 cycle per second. It was soon found, when studying the boundary layer on a flat plate, that the beam of light could not be kept lined up parallel with the flat plate. Moreover, at that time the glass plates and mirrors were supported at four points, and they could not be clamped very tightly without causing strains in the glass and distortion of the fringes. As a result, the interferometer would not stay long in adjustment, despite the fact that it was suspended by springs. Accordingly, the supports of the plates and mirrors were changed to the three-point type (see fig. 7) and were clamped down very tightly without distorting the fringes. The spring support was discarded and the interferometer table was placed on the floor as just described. As a result, the interferometer has stayed in adjustment for some length of time. In fact, in the past 9 months no adjustments of any kind have been made, and the interferometer has stayed in perfect adjustment.\n\nEach of the two plates and two mirrors is rotatable about two axes, one in a plane parallel to that of the interferometer base (fig. 6) and the other perpendicular to the plane of the interferometer base. For rotation about the first of these axes, the mirror holder is mounted in journal bearings. (See fig. 7.) These bearings are clamped tight, just short of binding. A lever is attached to the axle. The ends of micrometers press against the lever. Rotation of the plate or mirror is effected by advancing one micrometer while retracting the other. When the proper position is reached, both micrometers are screwed very tightly against the lever. For rotation about the other axis, the plate or mirror housing rotates on ball bearings. Again a lever is moved by means of two micrometers. (The micrometers provide an economical method of obtaining precision screw threads. The scales on the micrometers are not used.)\n\nFor adjustment of the white-light fringes; that is, for adjusting the difference in optical-path length of the two beams, where optical-path length is defined as the integral of the product of index of refraction and the differential of geometric path length, an adjustment", "timestamp": "2026-07-22T04:53:57.425748+00:00"}
{"citation_id": "19930082617", "source_url": "https://ntrs.nasa.gov/api/citations/19930082617/downloads/19930082617.pdf", "page_number": 13, "total_pages": 58, "image_filename": "19930082617_p13.jpg", "text": "```markdown\n12\nNACA TN 1962\n\n<!-- Image (100, 80, 898, 899) -->\n\n(b) Cylinders 76 to 79.\nFigure 1.- Concluded.\n```", "timestamp": "2026-07-22T04:53:59.478229+00:00"}
{"citation_id": "19930082712", "source_url": "https://ntrs.nasa.gov/api/citations/19930082712/downloads/19930082712.pdf", "page_number": 8, "total_pages": 14, "image_filename": "19930082712_p8.jpg", "text": "6\nNACA TN 1998\n\nairfoil with roughened leading edge there is, of course, a complete absence of a region of low drag corresponding to extensive laminar layers. In the region where the drag rises rapidly with increase in lift coefficient, increasing the Reynolds number causes some slight decrease in the drag of the airfoil with roughened leading edge. The apparent adverse scale effect between Reynolds numbers of $2.6 \\times 10^6$ and $3.0 \\times 10^6$ may be attributed to the fact that the data at Reynolds numbers of $1.8 \\times 10^6$ and $2.6 \\times 10^6$ (reference 2) were obtained by employing a smaller extent of roughness than was used in the present investigation.\n\nThe drag coefficient at the experimental design lift coefficient, the value at the center of the low-drag region of the drag-lift curve, is plotted in figures 2(a) and 2(b) as a function of Reynolds number for the NACA 8-H-12 section. The values of the experimental design lift coefficient selected for the smooth and rough surface conditions of the NACA 8-H-12 airfoil are 0.52 and 0.22, respectively. The data show that the drag of the smooth section at design lift, although it remains relatively constant between Reynolds numbers of $3.0 \\times 10^6$ and $6.0 \\times 10^6$, decreases in general as the Reynolds number is increased from $1.8 \\times 10^6$ to $11.0 \\times 10^6$. In the rough surface condition the drag remains nearly constant up to a Reynolds number of $3.0 \\times 10^6$ and then decreases progressively as the Reynolds number is increased from $3.0 \\times 10^6$ to $11.0 \\times 10^6$ (fig. 2(b)).\n\nIn figure 2 the section drag coefficient at the experimental section design lift coefficient is also shown plotted against Reynolds number for the NACA 0012 and NACA 23012 airfoils. The drag coefficient at design lift coefficient for the NACA 8-H-12 airfoil section is less than that for either the NACA 0012 or the NACA 23012 section in the smooth condition (fig. 2(a)). This can be attributed to the larger region of laminar flow prevailing on the NACA 8-H-12 airfoil in the smooth condition. When roughness is applied, however, this advantage of lower drag for the NACA 8-H-12 section is retained only at the lower Reynolds numbers and is lost at a Reynolds number of $6.0 \\times 10^6$, where the drag coefficients for the three airfoil sections are equal (fig. 2(b)).\n\n**Pitching moment.** - Pitching moments were measured about the horizontal axis on which the model was pivoted, and from these values, the position of the aerodynamic center and the pitching-moment coefficients about the aerodynamic center were calculated. The section moment coefficients about the pivot position $C_{mp}$ and about the aerodynamic center $C_{mac}$ are plotted in figures 1(a) and 1(b), respectively, for the airfoil in both the smooth and rough conditions. The positions of the aerodynamic center are tabulated in fig. 1(b). The positions of the aerodynamic center for the two lower Reynolds numbers (reference 2) have been recalculated and are somewhat different from the values given in reference 2.", "timestamp": "2026-07-22T04:54:00.111759+00:00"}
{"citation_id": "19930082476", "source_url": "https://ntrs.nasa.gov/api/citations/19930082476/downloads/19930082476.pdf", "page_number": 26, "total_pages": 41, "image_filename": "19930082476_p26.jpg", "text": "CHART 8.- EFFECT OF COMBINED RUDDER DEFLECTIONS ON THE SPIN CHARACTERISTICS OF MODEL (RUDDERS AND AILERONS UNLINKED)\n\n[Right erect spins; elevator set to 13° up, rudders and ailerons set as indicated]\n\nA. Loading 3 $\\left(\\frac{I_x - I_y}{mb^2} = 165 \\times 10^{-4}; \\mu = 5.32; \\text{loading 3 in table II and point 3 in fig. 4}\\right)$; ailerons $\\frac{1}{2}$ with the spin; right aileron $22\\frac{1}{2}^\\circ$ up, left aileron $30^\\circ$ down\n\nRight rudder setting against the spin, degrees\n\nLeft rudder setting against the spin, degrees\n\nB. Loading 4 $\\left(\\frac{I_x - I_y}{mb^2} = -18 \\times 10^{-4}; \\mu = 10.39; \\text{loading 4 in table II and point 4 in fig. 4}\\right)$; right aileron $30^\\circ$ up, left aileron neutral\n\nRight rudder setting against the spin, degrees\n\nLeft rudder setting against the spin, degrees\n\naOscillatory spin, too difficult to control in tunnel to permit obtaining data.\nbOscillatory spin, range of values or average value given.\n\nNACA\n\nModel values converted to corresponding full-scale values:\nU inner wing up\nD inner wing down\n\n| $\\alpha$ (deg) | $g$ (deg) |\n| :--- | :--- |\n| $V$ (fps) | $\\Omega$ (rps) |\n\nNACA TN No. 1801", "timestamp": "2026-07-22T04:54:00.547509+00:00"}
{"citation_id": "19930085487", "source_url": "https://ntrs.nasa.gov/api/citations/19930085487/downloads/19930085487.pdf", "page_number": 4, "total_pages": 36, "image_filename": "19930085487_p4.jpg", "text": "2\nNACA RM No. E5J22\n\nAPPARATUS AND PROCEDURE\n\nThe apparatus used to obtain the natural frequencies of the\ncompressor blades is shown in figure 1. The blades, made from a\nmagnetic material, are susceptible to excitation by an electromag-\nnetic coil. Power for the coil used for this purpose was supplied\nby amplifying an oscillator signal.\n\nThe coil was held near the end of the blade, perpendicular to\nthe blade surface, and at a distance sufficient to prevent the blade\nfrom striking the core during resonance. The frequency of excita-\ntion was varied until high amplitudes were observed. The appearance\nof a high amplitude indicated that resonance had been obtained, and\nthe natural frequency of the blade was then read directly from the\noscillator dial. First bending-, second bending-, and first\ntorsional-mode frequencies were measured in all stages; second\ntorsional- and third and fourth bending-mode frequencies were deter-\nmined only in the first stage. The total error involved in making\nthe frequency measurements was less than 2 percent.\n\nSmall granules of ordinary table salt were sprinkled on a\nblade vibrating at one of its natural frequencies to outline the\nsand pattern of the nodes (fig. 2). A violin bow and the magnetic\ncoil were used to excite the blades in obtaining the node patterns.\n\nDISCUSSION AND RESULTS\n\nAccording to the engine manufacturer, blade failures had\noccurred in six different engines operating within the speed range\nof 16,000 to 17,000 rpm at high pressure ratios. The exact speeds\nof the various engines at the time of failure are unknown except\nin one case in which the engine was running at 16,600 rpm. In each\ncase only one blade broke and the fracture always occurred in either\nthe seventh or tenth stage of the compressor. The fractures occurred\n3/16 to 1/4 inch from the periphery of the rotor disk; fatigue had\nstarted at the point of maximum thickness on the convex side of the\nblade.\n\nBlade failure in only the seventh and tenth stages of the com-\npressors definitely indicates that the failures are caused by vibra-\ntory stresses inasmuch as all the blades of the sixth, seventh, and\neighth stages are the same in shape and size except for the amount\ntrimmed from the end of the blade. The ninth- and tenth-stage blades\nare also similar to each other except in length. The engines were\noperated at the rated speed of 17,000 rpm, approximately 400 rpm above", "timestamp": "2026-07-22T04:54:01.942135+00:00"}
{"citation_id": "19930093773", "source_url": "https://ntrs.nasa.gov/api/citations/19930093773/downloads/19930093773.pdf", "page_number": 34, "total_pages": 47, "image_filename": "19930093773_p34.jpg", "text": "```markdown\nNACA RM E9G09\n33\n\n1159\n\nFlight\nMach\nnumber\nO 0.21\n□ .53\n◇ .72\n● .85\n△ .97\n\nFuel-air ratio, f/a\n.020\n.016\n.012\n.008\n.004\n0\n\nEngine speed, N, rpm\n2 3 4 5 6 7 8 x 10³\n\n[NACA logo]\n\n(e) Fuel-air ratio.\nFigure 5. - Continued. Effect of flight Mach number on variation of\nengine performance with engine speed at altitude of 25,000 feet.\n```", "timestamp": "2026-07-22T04:54:04.871238+00:00"}
{"citation_id": "19930082542", "source_url": "https://ntrs.nasa.gov/api/citations/19930082542/downloads/19930082542.pdf", "page_number": 14, "total_pages": 53, "image_filename": "19930082542_p14.jpg", "text": "NACA TN No. 1867\n\nin this manner can be expected to have yield strengths from 90,000 to 110,000 psi at room temperature and rupture strengths for fracture in 100 hours from 60,000 to 63,000 psi and in 1000 hours from 52,000 to 55,000 psi. Very low ductility in the rupture test will be associated with any effectively solution-treated and hot-cold-worked materials unless the working temperature is above $1400^\\circ$ F.\n\nThe upper limit of temperature of solution-treating prior to hot-cold-working lies between $2050^\\circ$ and $2200^\\circ$ F. A solution temperature of $2200^\\circ$ F will result in somewhat poorer properties than a $2050^\\circ$ F treatment. The major objection to the higher solution temperature, however, is the excessive brittleness under slow strain rates. The lower limit of solution-treating lies between $1950^\\circ$ and $1800^\\circ$ F. The properties obtained with a $1950^\\circ$ F treatment were quite similar to those with a $2050^\\circ$ F treatment with somewhat better ductility in the rupture test. The rupture strengths after treating at $1800^\\circ$ F were low.\n\nThe limited data available indicate that an aging treatment prior to hot-cold-working will not affect the room-temperature properties but will have a pronounced effect on the rupture properties. Rupture strengths apparently will be substantially lower and the ductility much better. This procedure, however, avoids the stress-concentration brittleness of the plain solution-treated and hot-cold-worked material.\n\nProper aging after hot-cold-working apparently will substantially improve material solution-treated at $2200^\\circ$ F prior to hot-cold-working. When solution-treated at $2050^\\circ$ F, however, aging will reduce all properties except ductility in the rupture test. The sensitivity to stress-concentration brittleness will also be removed. The potential improvements from aging after hot-cold-working are sufficient to warrant further work to establish the results of such treatments more completely.\n\nIt then appears that if high yield strength at room temperature and high rupture strength at $1200^\\circ$ F are adequate criterions of gas-turbine service the best procedure is to solution-treat in the temperature range from $1950^\\circ$ to $2050^\\circ$ F and then to hot-cold-work to 10- to 15-percent reduction at temperatures below $1400^\\circ$ F. Hot-cold-working the alloy in the hot-worked condition is not recommended, even though equal or better properties were obtained by that procedure in this investigation, because hot-worked material will probably be quite variable.\n\nTreatments for Prolonged Service\n\nHot-cold-work produced the highest rupture strengths for prolonged service at $1200^\\circ$ F as is shown by the typical curves of stress against rupture time in figure 15.", "timestamp": "2026-07-22T04:54:07.112619+00:00"}
{"citation_id": "19930082485", "source_url": "https://ntrs.nasa.gov/api/citations/19930082485/downloads/19930082485.pdf", "page_number": 30, "total_pages": 62, "image_filename": "19930082485_p30.jpg", "text": "NACA TN No. 1810\n29\n\n1026\n\nVelocity-\npotential\nlines\n\nAssumed\nstreamline\n\nNACA\n\nFigure 1.- Blade section showing potential lines used\nfor calculating the surface velocities.", "timestamp": "2026-07-22T04:54:07.210483+00:00"}
{"citation_id": "19930082450", "source_url": "https://ntrs.nasa.gov/api/citations/19930082450/downloads/19930082450.pdf", "page_number": 21, "total_pages": 37, "image_filename": "19930082450_p21.jpg", "text": "20\nNACA TN No. 1778\n\n$$\n\\frac{H}{t_W} = 21\n$$\n$$\n\\left(\\frac{b_W}{t_W} = 20\\right)\n$$\n\n$$\n\\bar{\\sigma}_f, \\text{ksi} \\quad \\frac{S}{t_S} \\text{ or } \\frac{b_S}{t_S}\n$$\n\n$$\n\\sigma_{cr}, \\text{ksi}\n$$\n7.3\n\nColors indicate minimum-weight proportions for $\\frac{t_W}{t_S} = 0.51$.\nRed means some other, blue means no other value of $\\frac{t_W}{t_S}$ gives less weight.\n\n$$\n\\frac{P_1}{t_S}, \\text{ksi}\n$$\n\n31\n(30)\n\n$$\n\\sigma_{cr}, \\text{ksi}\n$$\n11.4\n7.3\n\n41\n(40)\n\n$$\n\\sigma_{cr}, \\text{ksi}\n$$\n18.2\n11.4\n7.3\n\n[Figure: NACA logo]\n\n$$\n\\frac{P_1}{L\\sqrt{C}}, \\text{ksi}\n$$\n\nFigure 2.-Direct-reading design charts for 24S-T aluminum-alloy Z-stiffened panels, $\\frac{t_W}{t_S}=0.51$.", "timestamp": "2026-07-22T04:54:10.256260+00:00"}
{"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 12, "total_pages": 78, "image_filename": "19930082618_p12.jpg", "text": "10\nNACA TN 1945\n\nthe chord (fig. 16(a)). Variation of camber seems to produce only slight, inconsistent changes in slope of the curve of drag against Reynolds number. These trends indicate that the advantage in drag reduction to be derived from the use of thin airfoil sections increases as the Reynolds number is lowered and that relatively far forward positions of minimum pressure are desirable at low values of the Reynolds number. It is interesting to note that, although the purpose of moving the position of minimum pressure rearward to 60 percent of the chord is to decrease the drag by increasing the relative extent of laminar flow, the section with minimum pressure farthest forward actually has more favorable drag characteristics at the two lowest Reynolds numbers. The fact that regions of laminar separation probably exist behind the position of minimum pressure and increase in extent as the airfoil thickness is increased and as the position of minimum pressure is moved rearward is believed to be responsible for the observed effect of airfoil thickness and thickness distribution on the drag at the lower Reynolds numbers.\n\nThe drag data for the NACA 4-digit-series and 5-digit-series airfoils in the smooth surface condition (fig. 16(a)) generally do not show as much variation with Reynolds number as do those for the NACA 6-series sections. Because of the differences in scale effect, the advantage in drag reduction derived from employing a smooth NACA 6-series section as compared with one of the smooth NACA 4- or 5-digit-series sections diminishes as the Reynolds number is lowered. At a Reynolds number of $0.7 \\times 10^6$, the drag values for the smooth condition of the NACA 6-series airfoils and the NACA 4- and 5-digit-series sections are of about the same order of magnitude. The fact that the scale effect on the minimum drag of the NACA 4-digit- and 5-digit-series sections is smaller than that shown by the NACA 6-series sections may possibly be attributed to the following two effects: first, from some preliminary studies there is reason to believe that there exists on the NACA 6-series sections a region of laminar separation behind the position of minimum pressure which increases in extent as the Reynolds number is lowered and causes the drag to increase rapidly. Within the same range of Reynolds number, however, the transition point on the NACA 4-digit- and 5-digit-series sections is ahead of the incipient laminar separation point so that no regions of separated flow exist. Second, the transition point on the NACA 4- and 5-digit-series sections probably moves rearward as the Reynolds number is reduced so that the relative extent of laminar flow increases as the Reynolds number is decreased. The extent of laminar flow on the NACA 6-series sections is limited at the position of laminar separation which, of course, does not vary with Reynolds number. If the Reynolds number were sufficiently low so that the transition point on the NACA 4-digit- and 5-digit-series sections were to occur behind the incipient separation point, regions of laminar separation", "timestamp": "2026-07-22T04:54:11.372911+00:00"}
{"citation_id": "19930082498", "source_url": "https://ntrs.nasa.gov/api/citations/19930082498/downloads/19930082498.pdf", "page_number": 18, "total_pages": 49, "image_filename": "19930082498_p18.jpg", "text": "```markdown\nNACA TN No. 1838\n\nTABLE II.- RESULTS OF MUFFLER INVESTIGATION MADE WITH PROPELLER REMOVED - Continued\n\n| Muffler configuration | Engine speed (rpm) | Fundamental firing frequency, F (cps) | Sound-pressure level (db) | | | | | | | | | | | Other sounds | | | Back pressure |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| (a) | | | Over-all | 0.5F | 1.0F | 2.0F | 3.0F | 4.0F | 5.0F | 6.0F | 7.0F | cps | db | cps | db | |\n| 32 [Figure: 8 Holes-1/2 diam. 8 Holes-1/4 diam. 10 Holes-1 diam. 2 1/2 1/2 1/2 4 65 10 10 24 12 20 20 20 10 10] | 1690 | 82.5 | 79.5 | 50 | 77 | 50 | 62 | 55 | 60 | 50 | 50 | 50 | (b) | (b) | (b) | (b) | --- | --- | --- | High |\n| | 2000 | 100.0 | 82.0 | --- | --- | --- | --- | --- | --- | --- | --- | --- | --- | --- | --- | --- | --- | --- | --- | |\n| | 2790 | 135.5 | 88.0 | --- | --- | --- | --- | --- | --- | --- | --- | --- | --- | --- | --- | --- | --- | --- | --- | |\n| 33 [Figure: 8 Holes-1/2 diam. 8 Holes-1/4 diam. 2 1/2 1/2 1/2 1/2 6 10 10 24 20 20 20 10] | 1690 | 82.5 | 79.0 | --- | (b) | 77 | 50 | 59 | 50 | 60 | 50 | 50 | 50 | (b) | (b) | (b) | (b) | --- | --- | --- | High |\n| | 2000 | 100.0 | 82.0 | --- | --- | --- | --- | --- | --- | --- | --- | --- | --- | --- | --- | --- | --- | --- | --- | |\n| 34 [Figure: 8 Holes-1/2 diam. 8 Holes-1/4 diam. 2 1/2 1/2 1/2 1/2 10 20 20 12 10] | 2000 | 100.0 | 86.0 | 54 | 86 | 60 | 78 | 58 | 71 | 59 | 62 | 57 | 57 | 55 | 59 | (b) | (b) | 133 | 60 | --- | --- | High |\n| 35 [Figure: 8 Holes-1/2 diam. 8 Holes-1/4 diam. 2 1/2 1/2 1/2 1/2 10 20 20 15 12] | 2000 | 100.0 | 87.0 | 57 | 87 | 64 | 76 | 64 | 68 | 64 | 73 | 62 | 64 | 59 | 59 | 59 | 57 | 369 | 67 | 800 | 57 | High |\n| 36 [Figure: 8 Holes-1/2 diam. 8 Holes-1/4 diam. 2 1/2 1/2 1/2 1/2 8 20 20] | 2000 | 100.0 | 87.5 | 48 | 87 | 57 | 77 | 53 | 68 | 58 | 60 | 58 | 48 | 48 | 48 | 53 | 60 | --- | --- | --- | --- | High |\n\n$^a$In the sketches of the configurations, all dimensions are in inches and all cross sections are circular, except where otherwise indicated.\n$^b$The sound-pressure level was below the range of the analyzer.\n\nNACA\n17\n```", "timestamp": "2026-07-22T04:54:12.779238+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 23, "total_pages": 99, "image_filename": "19930082511_p23.jpg", "text": "NACA TN No. 1826\n21\n\neither of the products $G \\sqrt{1 - z^2}$, $Gz \\sqrt{1 - z^2}$, or a linear combination of the two satisfies the four conditions and has a pole of the first order at $z_1$. These facts suggest that the desired velocity function is of the form\n\n$$Q(z) = i \\left( \\frac{1}{z - z_1} - \\frac{1}{z - \\bar{z}_1} \\right) (A + Bz) \\sqrt{1 - z^2} \\quad (3)$$\n\nwhere A and B are as yet undetermined real constants.\n\nThe values of A and B are to be determined such that the pole at $z_1$ represents a vortex of strength $\\Gamma'$ in the $\\zeta$-plane. Thus\n\n$$\\Gamma' = \\oint q(\\zeta) \\, d\\zeta = \\oint Q(z) \\, \\frac{d\\zeta}{dz} \\, dz = \\frac{1}{\\pi} \\oint Q(z) \\, \\frac{dz}{z}$$\n\nwhere the integral is taken about the point $z_1$. By Cauchy's integral formula\n\n$$\\Gamma' = -2(A + Bz_1) \\frac{\\sqrt{1 - z_1^2}}{z_1} \\quad (4)$$\n\nThe values of $\\Gamma'$ and $z_1$ are known, so that this complex equation can be solved for the two real constants A and B. Substituting these values in equation (3) will thus give the desired complex velocity function.\n\n**Tunnel-interference velocity.**- The tunnel-interference velocity is defined as the difference between the total perturbation velocity $q(\\zeta)$ due to the presence of the vortex in the tunnel and the velocity due to a vortex in an unbounded medium. That is, the tunnel-interference velocity $q_i(\\zeta, \\zeta_1)$ is\n\n$$q_i(\\zeta, \\zeta_1) = i \\left( \\frac{1}{z - z_1} - \\frac{1}{z - \\bar{z}_1} \\right) (A + Bz) \\sqrt{1 - z^2} + \\frac{i\\Gamma'}{2\\pi(\\zeta - \\zeta_1)} \\quad (5)$$\n\nIf the vortex is on the axis of the tunnel, that is, if $\\zeta_1 = \\xi_1 + \\frac{i}{2}$, then from equation (1), $z_1 = iy_1$, and equation (4) gives", "timestamp": "2026-07-22T04:54:16.155335+00:00"}
{"citation_id": "19930082487", "source_url": "https://ntrs.nasa.gov/api/citations/19930082487/downloads/19930082487.pdf", "page_number": 29, "total_pages": 33, "image_filename": "19930082487_p29.jpg", "text": "NACA TN No. 1813\n\n27\n\nSection drag coefficient, $c_d$\n\n| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | |", "timestamp": "2026-07-22T04:54:16.399579+00:00"}
{"citation_id": "19930082245", "source_url": "https://ntrs.nasa.gov/api/citations/19930082245/downloads/19930082245.pdf", "page_number": 38, "total_pages": 66, "image_filename": "19930082245_p38.jpg", "text": "```markdown\nNACA TN NO. 1596\n\n1.6\n1.4\n1.2\n1.0\n.8\n.6\n.4\n.2\n0\n-.2\n-.4\n-.6\n-.8\n.1 .2 .3 .4 .5 .6 .7 .8 .9\nMach number, M\nAileron section normal-force coefficient, $C_{n_a}$\n\n$\\delta_a$\n(deg)\n30\n18\n12\n4\n2\n0\n-2\n-4\n-6\n-12\n\n1.6\n.12\n.08\n.04\n0\n-.04\n-.08\n-.12\n-.16\n-.20\n-.24\n-.56\n-.60\n.1 .2 .3 .4 .5 .6 .7 .8 .9\nMach number, M\nAileron section hinge-moment coefficient, $C_h$\n\n$\\delta_a$\n(deg)\n-12\n-6\n-4\n-2\n0\n2\n4\n18\n12\n30\n\n(a) $c_n = 0.4$.\nFigure 7. - Continued.\n\nNACA\n37\n```", "timestamp": "2026-07-22T04:54:16.587941+00:00"}
{"citation_id": "19930085536", "source_url": "https://ntrs.nasa.gov/api/citations/19930085536/downloads/19930085536.pdf", "page_number": 1, "total_pages": 20, "image_filename": "19930085536_p1.jpg", "text": "NACA-FM No. E8K05\nRM No. E8K05\n\n[Figure: NACA logo with wings]\n\nNACA\n\nRESEARCH MEMORANDUM\n\nPRESSURE DISTRIBUTIONS ON THIN CONICAL BODY OF ELLIPTIC\nCROSS SECTION AT MACH NUMBER 1.89\n\nBy Stephen H. Maslen\n\nLewis Flight Propulsion Laboratory\nCleveland, Ohio\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\n\nWASHINGTON\nJanuary 20, 1949\nDeclassified June 11, 1953", "timestamp": "2026-07-22T04:54:16.775285+00:00"}
{"citation_id": "19930082585", "source_url": "https://ntrs.nasa.gov/api/citations/19930082585/downloads/19930082585.pdf", "page_number": 14, "total_pages": 30, "image_filename": "19930082585_p14.jpg", "text": "NACA TN 1907\n\nNow $\\dot{\\Omega}_{\\Delta t}$ may be computed from equation (13a) and the approximate value of $\\Omega_{\\Delta t}$ from equation (15) checked by\n\n$\\Omega_{\\Delta t} = \\Omega_0 + \\frac{\\Delta t}{2}(\\dot{\\Omega}_0 + \\dot{\\Omega}_{\\Delta t})$\n\nIf this does not check quite closely the value first obtained from equation (15), and used in equation (18), a trial-and-error process can be used to find $\\Omega_{\\Delta t}$. In this case, a value of $\\Omega_{\\Delta t}$ would be assumed, in place of equation (15), used in equation (18), and then checked by equation (19). Good agreement should be achieved before the process is repeated for the next interval. Experience has shown that for $\\Delta t = \\frac{1}{2}$ second, the first estimate furnished by equation (15) is usually checked quite closely (within 0.05 radian/sec) by equation (19).\n\nThis process is repeated for as many intervals of time as are necessary for $\\Omega$ and $V$ to reach their final steady values in steady autorotation.\n\nA running check can be kept on the approach to the instability caused by excessive blade stalling (see reference 1) by computing the value of the inflow ratio $\\lambda$ at every interval of time, and spotting $\\lambda$ and $\\theta$ on curves of $Q$ against $\\lambda$ and $\\theta$, in which the effect of blade stall has been included, as in figure 3 of reference 1. The value of $\\lambda$ is given by\n\n$\\lambda = \\frac{(V - v)}{\\Omega R}$\n\nLong before the \"second, unstable trim point\" (described in reference 1) is approached, however, it would be necessary to use modified equations (12a) and (13a), altered to include the effects of blade stalling. In this case, the trick represented by equation (15) to improve the convergence of the step-by-step process would be excessively complex to apply. The trial-and-error process necessary to carry out the numerical integration of the new equations (12a) and (13a) would be arduous at each interval of time, and, for adequate accuracy, the time intervals would have to be much more frequent.", "timestamp": "2026-07-22T04:54:17.617293+00:00"}
{"citation_id": "19930085471", "source_url": "https://ntrs.nasa.gov/api/citations/19930085471/downloads/19930085471.pdf", "page_number": 6, "total_pages": 28, "image_filename": "19930085471_p6.jpg", "text": "UNCLASSIFIED\nRECONFIDENTIAL\nNACA RM No. L8J11\n\naxis of zero twist optically. The first bending frequency and the\ndamping were obtained from a free-vibration record of the model. Since\nthe wings were uniform, the stiffness parameter GJ could be computed\nfrom a torsional-vibration record obtained with a mass of large known\nmoment of inertia attached to the wing tip. The uncoupled first torsion\nfrequency could then be calculated. The torsional damping was determined\nfrom the torsional-vibration-decay curve.\n\nThe models were designed not to flutter on the first run. The\nmodels were progressively modified by adding lead to shift their centers\nof gravity and by slotting to shift their elastic axes and change their\nbending and torsion frequencies, until flutter occurred. If a model was\nsaved, its parameters were changed to yield another flutter point. The\nvalues of the radius-of-gyration parameter $r_{\\alpha}^2$ and center of gravity\nwere determined from the unmodified wing and the added masses.\n\nDuring each test run, the following data were recorded simultaneously\nby means of a recording oscillograph: flutter frequency, position of the\nmodel in the tunnel, and static pressure which indicates the Mach number.\n\nA sample record of the flutter of model B-5 is given in figure 5.\n\nMETHOD OF ANALYSIS\n\nThe flutter data for the particular wings tested are put in coef-\nficient form and compared with the analytic solution of the supersonic\nbending-torsion flutter problem in a two-dimensional flow given in\nreference 1. The data of this paper were obtained at a Mach number\nof 1.31 and, since aerodynamic coefficients at this Mach number are not\nincluded in reference 1, a linear interpolation was made between the\ncomputed values of the flutter-speed coefficient at Mach numbers of 1.25\nand 1.43, for which the aerodynamic coefficients are available. First\nbending and uncoupled first torsion frequencies and damping\ncoefficients $\\xi_h$ and $\\xi_{\\alpha}$ were used in the theoretical analyses.\n\nIt is desirable to examine the possible errors introduced into the\nresults by the method of interpolation. It is known that the error may\nbe very large; for example, in the single-degree-of-torsional-instability\ncase for the elastic-axis position at 50 percent chord the interpolation\nwas made directly between the aerodynamic coefficients at Mach\nnumbers 1.25 and 1.43. This was necessary since the wing is stable at\na Mach number of 1.43 and the theory yields no solution for the flutter-\nspeed coefficient under these conditions.\n\nRECONFIDENTIAL\nUNCLASSIFIED", "timestamp": "2026-07-22T04:54:21.095079+00:00"}
{"citation_id": "19930085519", "source_url": "https://ntrs.nasa.gov/api/citations/19930085519/downloads/19930085519.pdf", "page_number": 3, "total_pages": 46, "image_filename": "19930085519_p3.jpg", "text": "2\nNACA RM No. L8K19\n\nThe total maximum rolling-moment coefficient resulting from 40° total deflection of a 49-percent-span by 20-percent-chord aileron was about the same as that produced by the plug aileron with the full-span slotted flap deflected. The aileron rolling-moment coefficients with the partial-span slotted flap deflected were equal to or only slightly greater than those with the flap neutral.\n\nINTRODUCTION\n\nThe spoiler type of control device has been proposed in reference 1 as a means of lateral control for sweptback wings. All of the spoiler configurations reported in reference 1, however, had some of the objectionable characteristics normally associated with spoilers on unswept wings; namely, a large reduction in rolling-moment coefficient at high angles of attack and low or reversed effectiveness at small spoiler projections for all angles of attack. Unpublished data showing the favorable rolling-moment characteristics in the transonic speed range obtained by one of the more satisfactory spoiler configurations of reference 1 indicated that further work toward improving the low-speed characteristics of spoilers on swept wings would be desirable. References 2 and 3 reported that the plug aileron (formed by the installation of a slot through the wing behind the spoiler) on unswept wings eliminated the objectionable rolling-moment characteristics exhibited by plain spoilers. References 2 and 3 further showed that the installation of a full-span slotted flap, in addition to giving high maximum lift coefficients for landing, greatly improved the rolling-moment effectiveness of the plug aileron.\n\nReported herein are the results of a high-lift and lateral-control investigation performed on a 42° sweptback semispan-wing model in the Langley 300 MPH 7- by 10-foot tunnel. The high-lift characteristics of a full-span slotted flap were determined on this model for a range of flap deflections and positions, and an attempt was made to improve the maximum lift characteristics of the full-span slotted flap by the installation of flap-slot flow-control vanes. The high-lift characteristics of a half-span slotted flap at one deflection and position and of a half-span Zap flap were also determined. A comparison of the calculated trimmed-gliding characteristics of the 42° sweptback wing under an assumed set of airplane conditions and equipped with several types and spans of lift flaps was made. Included in the lateral-control part of the investigation were the determination of the lateral control characteristics of a basic plug aileron and several revisions thereof. In addition, the characteristics of a 49-percent-span by 20-percent-chord plain aileron were determined for comparison with the plug-aileron results. Also determined were the lateral control characteristics of the most satisfactory plug-aileron configuration with the full-span slotted flap deflected to its optimum deflection and position and of the partial-span aileron with the partial-span slotted flap deflected.", "timestamp": "2026-07-22T04:54:24.703610+00:00"}
{"citation_id": "19930082614", "source_url": "https://ntrs.nasa.gov/api/citations/19930082614/downloads/19930082614.pdf", "page_number": 13, "total_pages": 36, "image_filename": "19930082614_p13.jpg", "text": "NACA TN 1939\n\nvelocity and altitude can usually be estimated so accurately that the first approximation is sufficient. If it is not, a second approximation may be made, or else smaller increments of time can be chosen. The time increments should be small enough so that all important changes in acceleration are taken into account, but large enough to keep the number of steps to a minimum.\n\nThe change in altitude during the time $\\Delta t$ for flight with a constant flight-path angle is\n\n$\\Delta h = \\overline{V} (\\Delta t) \\sin \\gamma$\n\nwhere $\\overline{V}$ is the average velocity. The estimated altitude change for each $\\Delta t$ is given by this relation in which $\\overline{V}$ is the average velocity estimated for that incremental time.\n\nThe longitudinal acceleration at the end of the time increment is given by figure 1, which includes the effects of altitude and wing loading (fig. 1(a)), of drag coefficient (fig. 1(b)), of velocity (fig. 1(c)), and of flight-path inclination (fig. 1(d)). When the speed variation of an airplane at a constant flight-path inclination is being calculated, the values of flight-path angle and wing loading are known. The altitude and speed are estimated, as indicated in the preceding paragraphs. The net drag coefficient, which may be a function of the flight Mach number and the airplane lift coefficient, is obtained from experimental or theoretical data using values of Mach number and lift coefficient computed from the estimated speed and altitude.\n\nThe increase or decrease in velocity during the time $\\Delta t$, assumed to be given by the area of the trapezoidal element under the curve of acceleration against time, is\n\n$\\Delta V = \\frac{1}{2} (a_1 + a) \\Delta t$\n\nin which $a_1$ and $a$ are the longitudinal acceleration at the beginning and of the end of the time increment. The velocity and Mach number which result from this velocity change can now be compared with the estimated values to ascertain whether the estimates are accurate.\n\nVariable flight-path inclination.— Because proposed aerodynamic brakes are to be designed for use in flight at any time, a method for calculating the speed of an airplane during any flight evolution is needed. The calculations for a dive at constant angle discussed in the preceding section would not accurately represent a steep dive of a", "timestamp": "2026-07-22T04:54:26.172701+00:00"}
{"citation_id": "19930082476", "source_url": "https://ntrs.nasa.gov/api/citations/19930082476/downloads/19930082476.pdf", "page_number": 27, "total_pages": 41, "image_filename": "19930082476_p27.jpg", "text": "NACA TN No. 1801\n25\n\n<!-- Image (179, 70, 906, 854) -->\n\nFigure 1.- Drawing of the $\\frac{1}{11}$-scale model of the twin-tail low-wing personal-owner-type airplane as tested in the Langley 20-foot free-spinning tunnel. Center of gravity indicated for normal loading.", "timestamp": "2026-07-22T04:54:29.371089+00:00"}
{"citation_id": "19930093773", "source_url": "https://ntrs.nasa.gov/api/citations/19930093773/downloads/19930093773.pdf", "page_number": 35, "total_pages": 47, "image_filename": "19930093773_p35.jpg", "text": "34\nNACA RM E9G09\n\n| | |\n| :--- | :--- |\n| Flight Mach number | |\n| $\\circ$ | 0.21 |\n| $\\square$ | .53 |\n| $\\diamond$ | .72 |\n| $\\triangle$ | .85 |\n| $\\nabla$ | .97 |\n\nExhaust-gas total temperature, $T_7$, $^\\circ R$\n2000\nLimiting temperature\n1600\n1200\n800\n400\n2 3 4 5 6 7 8x$10^3$\nEngine speed, N, rpm\n\nNACA\n\n(f) Exhaust-gas total temperature.\nFigure 5. - Concluded. Effect of flight Mach number on variation of engine performance with engine speed at altitude of 25,000 feet.\n\n6511", "timestamp": "2026-07-22T04:54:29.598119+00:00"}
{"citation_id": "19930082914", "source_url": "https://ntrs.nasa.gov/api/citations/19930082914/downloads/19930082914.pdf", "page_number": 9, "total_pages": 66, "image_filename": "19930082914_p9.jpg", "text": "8\nNACA TN No. 1857\n\nwas provided whereby the entire housing of mirror $M_1$ could be translated in a direction perpendicular to the plane of the mirror. The ways on which the housing moved were found, however, to be not sufficiently smooth, and when the mirror was moved, the plane of the mirror changed, and thus the orientation of the fringes changed. The screw was also found to be too coarse. Therefore, for easier adjustment of the white-light fringes, a compensating plate was installed in each beam ($C_1$ and $C_2$, fig. 8).\n\nEach plate is rotatable about two mutually perpendicular horizontal axes. One plate is placed in a horizontal position, and the white-light fringes are adjusted by rotating the other plate. The fact that the path in glass may be different for the two beams does not matter, because the light is too nearly monochromatic for dispersion to have an appreciable effect.\n\nThe splitter plates and the mirrors are 4 inches square and 1/2 inch thick. The plates are polished flat on each side to a tolerance of 1/10 wavelength. The angle of the wedge formed by the two sides of a plate does not exceed 2 seconds of arc. One side of each plate is coated with zinc sulphide for increased reflection. Use of this coating is a departure from the conventional method. This coating has the advantage over the conventional one, which is a thin coating of silver or aluminum, in that no light is absorbed by the coating. The ideal coating would transmit 50 percent of the light and reflect 50 percent. Since each beam is transmitted once and reflected once at splitter plates, the net result would be two beams going into the camera, each of which had 25 percent of the intensity of the original beam. For the zinc sulphide coating that was used, approximately 40 percent is reflected and 60 percent transmitted. This is satisfactory, however, because each beam into the camera then has 24 percent of the original intensity (neglecting losses at the mirrors).\n\nThe two mirrors are flat to a tolerance of 1/10 wavelength and are front-surfaced with rhodium. Because of the low reflectance of rhodium (about 75 percent), it is expected that aluminum would be more satisfactory.\n\nDescription of light source and camera.- A satisfactory light source for application of interferometry to the study of flow phenomena must provide nearly monochromatic light in a nearly parallel beam. An additional requirement for taking interferograms of high-speed flow is that the duration of the light be sufficiently short. For taking interferograms that are free from blurring of phenomena in an open, or free, supersonic jet, exposure times of the order of three microseconds or less must be used. (Inherent in free supersonic jets are high-frequency vibrations.) Because meeting each of the requirements of being monochromatic, parallel, and of short duration tends to reduce the intensity, one must add the fourth and obvious requirement of sufficient intensity.\n\nThe light-source problem has been solved by the same general method that was used by the Princeton group (reference 10), that is, by use of a high-voltage magnesium spark and a monochromator. A diagram of the complete light-source optical system is shown in figure 8. The two", "timestamp": "2026-07-22T04:54:30.138345+00:00"}
{"citation_id": "19930082712", "source_url": "https://ntrs.nasa.gov/api/citations/19930082712/downloads/19930082712.pdf", "page_number": 9, "total_pages": 14, "image_filename": "19930082712_p9.jpg", "text": "NACA TN 1998\n\nPitching moments about the aerodynamic center for the smooth airfoil are slightly positive for all six Reynolds numbers. Generally, the value of the pitching moment seems to become somewhat more positive as the Reynolds number is increased. Roughness has the effect of decreasing in magnitude the value of the moment coefficient about the aerodynamic center and of shortening the range of lift coefficient over which the moment-curve slope is constant.\n\nThe position of the aerodynamic center shows an appreciable forward shift as the Reynolds number is increased from $1.8 \\times 10^6$ to $2.6 \\times 10^6$ for the smooth airfoil and from $1.8 \\times 10^6$ to $3.0 \\times 10^6$ for the airfoil with leading-edge roughness; increases in the Reynolds number above these values had little effect. Within the range of Reynolds number from $3.0 \\times 10^6$ to $11.0 \\times 10^6$, the position of the aerodynamic center is farther forward for the rough than for the smooth surface condition.\n\nCONCLUSIONS\n\nThe results of the present investigation of the NACA 8-H-12 airfoil section through a range of Reynolds number from $3.0 \\times 10^6$ to $11.0 \\times 10^6$, together with those obtained from a previous investigation of this airfoil at Reynolds numbers of $1.8 \\times 10^6$ and $2.6 \\times 10^6$, indicate the following conclusions:\n\n1. No unusual scale effects on lift, drag, or pitching moment were present for the smooth NACA 8-H-12 airfoil over the range of Reynolds number from $1.8 \\times 10^6$ to $11.0 \\times 10^6$. This was also true for the airfoil with roughened leading edge except for an apparent adverse scale effect on the drag between Reynolds numbers of $2.6 \\times 10^6$ and $3.0 \\times 10^6$, which may be attributed to a difference in the extent of roughness employed at these Reynolds numbers.\n\n2. The values of the pitching-moment coefficient about the aerodynamic center were somewhat positive and increased in magnitude with increasing Reynolds number for the smooth NACA 8-H-12 airfoil. Roughening the leading edge caused the value of the pitching moment about the aerodynamic center to decrease in magnitude.\n\n3. The position of the aerodynamic center had a pronounced forward movement between Reynolds numbers of $1.8 \\times 10^6$ and $2.6 \\times 10^6$ for the smooth section, and between $1.8 \\times 10^6$ and $3.0 \\times 10^6$ for the section with roughened leading edge.", "timestamp": "2026-07-22T04:54:32.840570+00:00"}
{"citation_id": "19930082485", "source_url": "https://ntrs.nasa.gov/api/citations/19930082485/downloads/19930082485.pdf", "page_number": 31, "total_pages": 62, "image_filename": "19930082485_p31.jpg", "text": "30\n\nSurface critical velocity ratio, $\\frac{V}{V_{cr}}$\n\nPressure side, $S_2$, in.\n\nSuction side, $S_1$, in.\n\nCalculated\nExperimental\n\nBlade-surface length, S, in.\n(a) Root section.\n\nFigure 2.- Blade-surface velocities at design conditions.\n\nNACA TN No. 1810\n\nT206", "timestamp": "2026-07-22T04:54:33.430678+00:00"}
{"citation_id": "19930082496", "source_url": "https://ntrs.nasa.gov/api/citations/19930082496/downloads/19930082496.pdf", "page_number": 22, "total_pages": 50, "image_filename": "19930082496_p22.jpg", "text": "NACA TN No. 1836\n21\n\n[Figure: Diagram of a tensile-strength evaluation setup with labeled components]\n\nBall and socket\nUpper cross head\nHelium inlet\nAdapter rod\nAsbestos sheet\nUpper collar\nFurnace\nSpecimen grip\nTensile specimen\nLower collar\nSpecimen thermocouple\nFurnace-control thermocouple\nLower cross head\n\nNACA\n\nFigure 1. - Setup for tensile-strength evaluation.", "timestamp": "2026-07-22T04:54:34.399937+00:00"}
{"citation_id": "19930082617", "source_url": "https://ntrs.nasa.gov/api/citations/19930082617/downloads/19930082617.pdf", "page_number": 14, "total_pages": 58, "image_filename": "19930082617_p14.jpg", "text": "NACA TN 1962\n13\n\nStringers\nO 1 to 9\nX 10 to 16\n\nMoment\n(in. - lb)\n1 28.8 X $10^3$\n2 57.6 X $10^3$\n3 86.4 X $10^3$\n4 115.2 X $10^3$\n5 144.0 X $10^3$\n6 201.6 X $10^3$\n\n3.86\"\nBand B\nA\nA\nA-A\n\nDistance from horizontal diameter, in.\n10\n9\n8\n7\n6\n5\n4\n3\n2\n1\n0\n1\n2\n3\n4\n5\n6\n7\n8\n9\n10\n\n16 . 12 . 8 . 4 . 0 . 4 . 8 . 12 . 16 . 20 X $10^{-4}$\nStrain\n\n1 2 3 4 5 6\n\nNACA\n\nFigure 2.- Strain diagram of cylinder 72. Band B.", "timestamp": "2026-07-22T04:54:34.581722+00:00"}
{"citation_id": "19930085487", "source_url": "https://ntrs.nasa.gov/api/citations/19930085487/downloads/19930085487.pdf", "page_number": 5, "total_pages": 36, "image_filename": "19930085487_p5.jpg", "text": "NACA RM No. E3J22\n\nthe speed at which the blades failed; hence, the failures cannot be attributed to stress rupture caused by centrifugal force.\n\nNatural-Frequency Measurements\n\nBecause all the failures occurred within the narrow speed range of 16,000 to 17,000 rpm, the blades were possibly excited at their resonant point. The fundamental natural frequency was therefore measured for each blade; these data as well as the highest, lowest, and average frequencies observed in each stage are listed in table I. A 20-percent variation in blade frequencies existed in any single stage; if the rotor spins at rated speed, the frequencies of all blades might approach the highest frequency in the stage because the centrifugal force and the thermal expansion raise the frequency by tightening the bulbous blade root. The centrifugal force also has a stiffening effect on the blades, which increases the frequency from 72 percent in the first stage to 16 percent in the tenth stage above the static measurements. Several blades in the ninth and tenth stages were loose but would act like firmly clamped blades when the compressor was in operation.\n\nSecond bending- and first torsional-mode vibrations were easily excited in all stages; their frequencies were measured and are recorded in table II. Only in the eighth stage was the frequency of the second bending mode higher than the frequency of the first torsional mode. No reason for this exception was determined.\n\nNode Locations\n\nThe location of failure and the starting point of fatigue made it necessary to know the stress patterns for all modes of vibration in order to determine the modes that could cause the failure. High-stress points are located at positions of greatest change in slope of the deflection curve for the vibrating blade; the greatest slope changes generally occur at the antinode points. In addition to the antinode at the tip of the blades, the other antinodes are located approximately midway between nodes. The high-stress points were estimated from sand patterns shown in figures 2 to 8. All of the nodes, except the one at the root and the one closest to the tip of the blade, are points of zero stress. Application of this analysis to the sand pattern shown in figure 4, for example, indicates that a high-stress point would be located (1) at the root, (2) about two-thirds the distance from the root to the nearest node because the root is fixed, not hinged, and (3) at some point between the node nearest the root and the tip of the blade. Low stresses occur at a", "timestamp": "2026-07-22T04:54:36.867907+00:00"}
{"citation_id": "19930082450", "source_url": "https://ntrs.nasa.gov/api/citations/19930082450/downloads/19930082450.pdf", "page_number": 22, "total_pages": 37, "image_filename": "19930082450_p22.jpg", "text": "NACA TN No. 1778\n21\n\n$$\n\\frac{P_L}{l_S}, \\text{ ksi}\n$$\n\n$$\n\\frac{H}{t_W} = 26\n$$\n$$\n\\left(\\frac{b_W}{t_W} = 25\\right)\n$$\n\n$$\n\\frac{S}{l_S} \\text{ or } \\frac{b_S}{l_S}\n$$\n\n$$\n\\overline{\\sigma}_f, \\text{ ksi}\n$$\n\n$$\n\\sigma_{cr}, \\text{ ksi}\n$$\n11.4\n7.3\n\n[Figure: Cross-section diagram showing dimensions $H$, $b_W$, $t_W$, $l_S$, $b_S$ or $S$]\n\n$$\n\\frac{P_L}{l_S}, \\text{ ksi}\n$$\n\n$$\n\\sigma_{cr}, \\text{ ksi}\n$$\n11.4\n7.3\n\n$$\n\\frac{P_L}{L t_S}, \\text{ ksi}\n$$\n\n$$\n\\frac{t_W}{l_S} = 0.51\n$$\n\n$$\n\\sigma_{cr}, \\text{ ksi}\n$$\n15.9\n11.4\n7.3\n\n24S-T\n$$\n\\sigma_{cy} = 44 \\text{ ksi}\n$$\n\nNACA\n\nFigure 2.-Concluded.", "timestamp": "2026-07-22T04:54:40.968947+00:00"}
{"citation_id": "19930082542", "source_url": "https://ntrs.nasa.gov/api/citations/19930082542/downloads/19930082542.pdf", "page_number": 15, "total_pages": 53, "image_filename": "19930082542_p15.jpg", "text": "14\nNACA TN No. 1867\n\nThis finding is contrary to the current metallurgical belief that the hot-cold-worked condition of all alloys is so unstable that it will lose its strength superiority at only slightly longer time periods than 1000 hours. This belief is probably correct for lower-alloyed materials but is evidently not true for more highly alloyed materials such as low-carbon N-155 alloy.\n\nThe curves in figure 15 show that the slope of the curves of stress against rupture time for properly solution-treated and hot-cold-worked low-carbon N-155 alloy is not much greater and is at a higher stress level than for the best solution-treated material. On the basis of rupture strength, therefore, the best preparation for prolonged service is to hot-cold-work properly. The highest rupture strengths were obtained by solution-treating at $2050^\\circ$ F prior to 15-percent reduction at $1200^\\circ$ F and then aging at $1400^\\circ$ F or by solution-treating at $1950^\\circ$ F followed by hot-cold-working to 15-percent reduction at $1200^\\circ$ F.\n\nOnly intermediate rupture strengths at prolonged time periods result from the best solution temperature of $2100^\\circ$ F. Aging treatments apparently result in the lowest strengths at the longer time periods.\n\nData Correlation\n\nOne of the objectives of this investigation is to develop relationships which will permit checking the relative properties of any lot of low-carbon N-155 alloy by means of a few simple tests. This phase has not yet been properly investigated. The relationships between Brinell hardness and properties is, however, shown in figures 16, 17, and 18. At any hardness level the yield strengths at room temperature may vary as much as 30,000 psi. Hardness is indicative of rupture strength to within about 10,000 psi.\n\nThe relationships between hardness and properties do not give consideration to the individual treatments. The data curves for various treatments in previous figures indicate that hardness would correlate much closer with properties for any one treatment. More data on several heats are needed before the degree of reproducibility of hardness for any one treatment can be determined. It is believed, however, that the general curves of figures 16, 17, and 18 are indicative of the range in properties which might be expected at any hardness level.\n\nTwo reports have been issued for low-carbon N-155 discs. (See references 4 and 5.) The degree to which the properties of these discs correlated with the results of this investigation is summarized in table IV. The average test data for the discs are compared with the data for similarly treated bar stock in this investigation. Estimated values are also included which are based on the Brinell hardness correlations and the experience gained from the trends of the effects", "timestamp": "2026-07-22T04:54:43.381473+00:00"}
{"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 13, "total_pages": 78, "image_filename": "19930082618_p13.jpg", "text": "NACA TN 1945\n\nand, consequently, higher drags and more pronounced scale effect would be expected for these airfoils.\n\nIn the rough surface condition, the minimum drag coefficients of the NACA 6-series and 4- and 5-digit-series airfoil sections vary with Reynolds number in about the same manner. The values of the drag of comparable NACA 6-series, 00-series, and 230-series sections are also about the same at most Reynolds numbers; whereas the drag values of the NACA 44-series sections are, comparatively, appreciably higher. In general, increases in the airfoil thickness ratio and camber cause rather consistent increases in the drag throughout the Reynolds number range; whereas variations in thickness form seem to have a relatively small effect.\n\nDrag outside the low-drag range.- From an inspection of the data of figures 1 to 15, it can be seen that the drag outside of the relatively flat portion of the polar increases for all the airfoils in the smooth and rough surface conditions as the Reynolds number is lowered from $9.0 \\times 10^6$ to $0.7 \\times 10^6$. The magnitude of the scale effect is generally largest for the smooth surface condition. Variations in airfoil-design parameters have some influence upon the magnitude and character of the scale effect; however, consistent trends are difficult to distinguish.\n\nLift\n\nThe lift parameters which are usually considered to be of most importance are the lift-curve slope, the angle of zero lift, and the maximum lift coefficient. From the lift data presented in figures 1 to 15, the values of these parameters have been determined at each Reynolds number for the airfoils tested and are plotted as functions of Reynolds number in figures 17 to 22.\n\nLift-curve slope.- According to reference 1, the slope of the lift curve is considered to be defined by a straight line tangent to the lift curve at the design lift coefficient. The lift-curve slopes for a Reynolds number of $6.0 \\times 10^6$, presented in reference 1, could be determined quite easily in accordance with this definition since the lift data corresponding to the higher Reynolds numbers generally show only a small amount of dispersion and are characterized by a nearly linear variation with angle of attack within the low lift-coefficient range. In the present experiments at Reynolds numbers below $3.0 \\times 10^6$, however, the necessarily low dynamic pressures reduced the accuracy of the measuring apparatus so that some scatter is present in the lift data. For this reason, and because some of the lift curves tended to have slight jogs and variations in slope near the design lift coefficient,", "timestamp": "2026-07-22T04:54:46.332842+00:00"}
{"citation_id": "19930082245", "source_url": "https://ntrs.nasa.gov/api/citations/19930082245/downloads/19930082245.pdf", "page_number": 39, "total_pages": 66, "image_filename": "19930082245_p39.jpg", "text": "```markdown\n1.6\n$\\delta_a$\n(deg)\n30\n1.4\n1.2\n1.0\n.8\n.6\n.4\n.2\n0\n-.2\n-.4\n-.6\n-.8\n.1 .2 .3 .4 .5 .6 .7 .8 .9\nMach number, M\n\nAileron section normal-force coefficient, $c_{n_a}$\n\n18\n12\n4\n2\n0\n-2\n-4\n-6\n-12\n\n.16\n.12\n.08\n.04\n0\n-.04\n-.08\n-.12\n-.16\n-.20\n-.24\n-.28\n-.32\n-.36\n-.40\n-.44\n-.48\n-.52\n-.56\n-.60\n.1 .2 .3 .4 .5 .6 .7 .8 .9\nMach number, M\n\nAileron section hinge-moment coefficient, $c_{h_a}$\n\n$\\delta_a$\n(deg)\n-12\n-6\n-4\n-2\n0\n2\n4\n18\n12\n30\n\n(f) $c_n = 0.6$.\nFigure 7. - Continued.\n\n38\nNACA CW No. 1596\nNACA\n```", "timestamp": "2026-07-22T04:54:49.509061+00:00"}
{"citation_id": "19930082476", "source_url": "https://ntrs.nasa.gov/api/citations/19930082476/downloads/19930082476.pdf", "page_number": 28, "total_pages": 41, "image_filename": "19930082476_p28.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T04:54:51.293466+00:00"}
{"citation_id": "19930082585", "source_url": "https://ntrs.nasa.gov/api/citations/19930082585/downloads/19930082585.pdf", "page_number": 15, "total_pages": 30, "image_filename": "19930082585_p15.jpg", "text": "14\nNACA TN 1907\n\nA SAMPLE DESIGN STUDIED\n\nUsing the methods described in the preceding section, the transitions from hovering to steady autorotation have been computed for a helicopter with the following physical characteristics:\n\nW = 2700 pounds\nb = 3\nR = 20 feet\n\nc = 1.25 feet\na = 5.6 per radian\n$c_{d_0}' = 0.0087 - 0.0216\\alpha_r + 0.400\\alpha_r^2$\n\n$I_1 = 100, 200, 400 \\text{ slug-feet}^2$\n\nThis sample design is the same as was considered in reference 1, for which the steady autorotative characteristics were computed.\n\nThe initial hovering state considered is\n\n$V_0 = 0$\n$\\Omega_0 = 25.1 \\text{ radians per second}$\n\n$v_0 = 21.25 \\text{ feet per second}$\n$\\theta_0 = 7.25^\\circ = 0.1265 \\text{ radian}$\n\nFor transitions in which the pitch is reduced after power failure, the final state considered is\n\n$V_F = 31.2 \\text{ feet per second}$\n$\\Omega_F = 20.8 \\text{ radians per second}$\n\n$v_F = 25.1 \\text{ feet per second}$\n$\\theta_F = 4^\\circ = 0.0698 \\text{ radian}$\n\nFor transitions in which the pitch is not reduced, the final steady state is taken as\n\n$V_F = 30.5 \\text{ feet per second}$\n$\\Omega_F = 16.5 \\text{ radians per second}$\n\n$v_F = 26.4 \\text{ feet per second}$\n$\\theta_F = 7.25^\\circ = 0.1265 \\text{ radian}$", "timestamp": "2026-07-22T04:54:55.142658+00:00"}
{"citation_id": "19930085519", "source_url": "https://ntrs.nasa.gov/api/citations/19930085519/downloads/19930085519.pdf", "page_number": 4, "total_pages": 46, "image_filename": "19930085519_p4.jpg", "text": "NACA RM No. L8K19\n\nSYMBOLS AND CORRECTIONS\n\nThe forces and moments on the wing are presented about the wind axes. The X-axis is in the plane of symmetry of the model and is parallel to the tunnel air flow. The Z-axis is in the plane of symmetry of the model and is perpendicular to the X-axis. The Y-axis is perpendicular to both the X-axis and Z-axis. All three axes intersect at a point 37.22 inches rearward of the leading edge of the wing root on the line of intersection of the plane of symmetry and the chord plane of the model, as shown in figure 1.\n\n| Symbol | Definition |\n| :--- | :--- |\n| $C_L$ | lift coefficient $\\left( \\frac{\\text{Twice lift of semispan model}}{qS} \\right)$ |\n| $C_{L_t}$ | trimmed lift coefficient |\n| $C_D$ | drag coefficient $(D/qS)$ |\n| $C_m$ | pitching-moment coefficient about Y-axis $(M/qS\\overline{c})$ |\n| $C_l$ | rolling-moment coefficient about X-axis $(L/qSb)$ |\n| $C_n$ | yawing-moment coefficient about Z-axis $(N/qSb)$ |\n| $D$ | twice drag of semispan model, pounds |\n| $M$ | twice pitching moment of semispan model about Y-axis, foot-pounds |\n| $L$ | rolling moment due to plug projection or aileron deflection about X-axis, foot-pounds |\n| $N$ | yawing moment due to plug projection or aileron deflection about Z-axis, foot-pounds |\n| $q$ | dynamic pressure, pounds per square foot $\\left( \\frac{1}{2}\\rho V^2 \\right)$ |\n| $S$ | twice area of semispan model, 32.24 square feet |\n| $\\overline{c}$ | wing mean aerodynamic chord (M.A.C.), 2.89 feet $\\left( \\frac{2}{S} \\int_{0}^{b/2} c^2 \\, dy \\right)$ |\n| $b$ | twice span of semispan model measured along Y-axis, 11.36 feet |\n| $c'$ | local wing chord measured along lines perpendicular to wing trailing edge |", "timestamp": "2026-07-22T04:54:58.332034+00:00"}