PDF/ocr_dataset_1011.jsonl

{"citation_id": "19930085842", "source_url": "https://ntrs.nasa.gov/api/citations/19930085842/downloads/19930085842.pdf", "page_number": 70, "total_pages": 104, "image_filename": "19930085842_p70.jpg", "text": "66\nNACA RM L9C29\n\n<!-- Image (108, 92, 867, 851) -->\n\nFigure 37.- Variation of $C_{hr}$, $C_n$, and $C_l$ with $\\delta_r$ on a $\\frac{1}{3}$-scale model of the airplane. Propellers removed; basic model configuration; V = 100 miles per hour.", "timestamp": "2026-07-22T04:31:34.479614+00:00"}
{"citation_id": "19930082447", "source_url": "https://ntrs.nasa.gov/api/citations/19930082447/downloads/19930082447.pdf", "page_number": 5, "total_pages": 24, "image_filename": "19930082447_p5.jpg", "text": "```markdown\nNACA TN No. 1775\n3\n\nLoad-factor coefficient\n\n$$C_l = \\frac{n_1 w g}{\\dot{y}_o^2} \\left( \\frac{W}{g} \\left\\{ \\frac{6 \\sin \\tau \\cos^2 \\tau}{[f(\\beta)]^2 \\phi(A) \\rho \\pi} \\right\\} \\right)^{1/3}$$\n\nDraft coefficient\n\n$$C_d = y \\sqrt{\\frac{g}{W}} \\left( \\frac{[f(\\beta)]^2 \\phi(A) \\rho \\pi}{6 \\sin \\tau \\cos^2 \\tau} \\right)^{1/3}$$\n\nTime coefficient\n\n$$C_t = t \\dot{y}_o \\sqrt{\\frac{g}{W}} \\left( \\frac{[f(\\beta)]^2 \\phi(A) \\rho \\pi}{6 \\sin \\tau \\cos^2 \\tau} \\right)^{1/3}$$\n\nAPPARATUS\n\nThe Langley impact basin and standard equipment are described in reference 2.\n\nThe model tested was the forebody of a prismatic float having a dead-rise angle of 40° designated the Langley impact basin model M-3. The model was essentially the same as that used in the tests reported in references 3 and 4, except for the angle of dead rise. The size and shape of the model are defined by the lines and dimensions shown in figure 1. The offsets are given in table I. The model mounted on the carriage boom is shown in figure 2.\n\nThe instrumentation used to measure horizontal displacement and velocity and vertical displacement and velocity was described in reference 2. Accelerations in the vertical direction were measured by a standard NACA accelerometer having a natural frequency of 16.5 cycles per second with approximately 0.67 critical damping and a range of -1g to 6g. The contact and exit of the model were determined by means of an electrical circuit completed by the water.\n\nPRECISION\n\nThe instrumentation used in the tests gives measurements that are believed accurate within the following limits:\n```", "timestamp": "2026-07-22T04:31:39.299487+00:00"}
{"citation_id": "19930086151", "source_url": "https://ntrs.nasa.gov/api/citations/19930086151/downloads/19930086151.pdf", "page_number": 14, "total_pages": 34, "image_filename": "19930086151_p14.jpg", "text": "12\nCONFIDENTIAL\nNACA RM L9J28\n\n3. Adding the end plate to the wing increased the wing lift-curve slope and the drag, but decreased the wing maximum lift and the lift-drag ratios appreciably and also decreased the aileron effectiveness.\n\n4. Aileron plan form generally had little effect on the values of rolling-moment coefficient and yawing-moment coefficient produced by aileron deflection, or on the lift, drag, and pitching-moment characteristics of the wing model.\n\nLangley Aeronautical Laboratory\nNational Advisory Committee for Aeronautics\nLangley Air Force Base, Va.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T04:31:41.564713+00:00"}
{"citation_id": "19930086105", "source_url": "https://ntrs.nasa.gov/api/citations/19930086105/downloads/19930086105.pdf", "page_number": 15, "total_pages": 22, "image_filename": "19930086105_p15.jpg", "text": "NACA RM E9H12\nCONFIDENTIAL\n\n$$ \\frac{P_3}{P_0} $$\nCombustion-chamber static pressure\nFree-stream total pressure\n\nMean pressure ratio as indicated on manometer\n\n[Figure: Graph showing pressure ratio fluctuations over time]\n\nTime, sec\n\nFigure 2. - Typical signal from variable-inductance pressure pickup showing wave form of pressure fluctuations at combustion-chamber inlet when ram jet is operating with shock oscillating in and out of diffuser inlet. Shock diffuser with triple-shock projecting cone; fundamental frequency, approximately 22 cycles per second.\n\nNACA\n13", "timestamp": "2026-07-22T04:31:42.158102+00:00"}
{"citation_id": "19930086015", "source_url": "https://ntrs.nasa.gov/api/citations/19930086015/downloads/19930086015.pdf", "page_number": 37, "total_pages": 54, "image_filename": "19930086015_p37.jpg", "text": "36\n\nCONFIDENTIAL\n\nCONFIDENTIAL\n\nNACA RM A9E24\n\nFlow inclination, $\\epsilon$, deg\n\nHorizontal distance from window center line, x, in.\n\nO Experimental survey\n— Angle variation from pressure data\n\nz = 0\n\nNACA\n\n(a) D=165.12; M=1.23.\n\nFigure 11.— The variation of stream angle axially in the Ames 6- by 6-foot supersonic wind tunnel. y=0; stagnation pressure = 9 lb/sq in. abs.", "timestamp": "2026-07-22T04:31:43.300895+00:00"}
{"citation_id": "19930085529", "source_url": "https://ntrs.nasa.gov/api/citations/19930085529/downloads/19930085529.pdf", "page_number": 84, "total_pages": 85, "image_filename": "19930085529_p84.jpg", "text": "```markdown\nNACA RM No. L8A30a\n83\n\nTABLE 77\n$$\n\\left[ \\Lambda = -45^\\circ, \\delta_{LE} = 9.8^\\circ, \\alpha = 2^\\circ \\right]\n$$\n\nCONFIDENTIAL\n\n| UPPER SURFACE | | | | | | | LOWER SURFACE | | | | | | |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| **Tube** | **Per-cent chord** | **Mach Number** | | | | | **Tube** | **Per-cent chord** | **Mach Number** | | | | |\n| | | **0.60** | **0.80** | **0.89** | **0.925** | **0.96** | | | **0.60** | **0.80** | **0.89** | **0.925** | **0.96** |\n| A 1 | 2.0 | -- | -- | -- | -- | -- | 86 | 3.0 | -- | -- | -- | -- | -- |\n| 2 | 6.0 | -- | -- | -- | -- | -- | 87 | 10.0 | -- | -- | -- | -- | -- |\n| 3 | 15.0 | -- | -- | -- | -- | -- | 88 | 25.0 | -- | -- | -- | -- | -- |\n| 4 | 27.5 | -- | -- | -- | -- | -- | 89 | 41.0 | -- | -- | -- | -- | -- |\n| 5 | 40.0 | -- | -- | -- | -- | -- | 90 | 52.5 | -- | -- | -- | -- | -- |\n| 6 | 50.0 | -- | -- | -- | -- | -- | 91 | 62.5 | -.068 | -.087 | -.039 | -.028 | -.086 |\n| 7 | 59.0 | -.107 | -.094 | -.069 | -.044 | -.115 | 92 | 72.5 | .000 | -.001 | -.003 | .000 | -.023 |\n| 8 | 67.5 | -.074 | -.067 | -.046 | -.021 | -.078 | 93 | 84.0 | -- | -- | -- | -- | -- |\n| 9 | 77.5 | -- | -- | -- | -- | -- | 94 | 94.0 | -- | -- | -- | -- | -- |\n| 10 | 89.5 | -- | -- | -- | -- | -- | | | | | | | |\n| 11 | 96.0 | -- | -- | -- | -- | -- | | | | | | | |\n| B12 | 2.0 | -- | -- | -- | -- | -- | 95 | 3.0 | -- | -- | -- | -- | -- |\n| 13 | 6.0 | -- | -- | -- | -- | -- | 96 | 10.0 | -- | -- | -- | -- | -- |\n| 14 | 15.0 | -- | -- | -- | -- | -- | 97 | 25.0 | -.058 | -.085 | -.123 | -.156 | -.202 |\n| 15 | 27.5 | -.054 | -.023 | -.429 | -.648 | -.709 | 98 | 41.0 | -.081 | -.100 | -.126 | -.154 | -.251 |\n| 16 | 40.0 | -.238 | -.211 | -.210 | -.448 | -.608 | 99 | 52.5 | -.061 | -.077 | -.109 | -.121 | -.215 |\n| 17 | 50.0 | -.215 | -.213 | -.152 | -.175 | -.614 | 100 | 62.5 | -.048 | -.059 | -.172 | -.082 | -.194 |\n| 18 | 59.0 | -.173 | -.170 | -.123 | -.046 | -.434 | 101 | 72.5 | -.011 | -.002 | -.005 | -.009 | -.121 |\n| 19 | 67.5 | -.110 | -.130 | -.069 | -.018 | -.253 | 102 | 86.3 | .040 | .033 | .031 | .035 | -.056 |\n| 20 | 77.5 | -.068 | -.068 | -.068 | -.004 | -.095 | 103 | 94.5 | .074 | .071 | .073 | .084 | -.020 |\n| 21 | 88.0 | .011 | .014 | .020 | .043 | -.023 | | | | | | | |\n| 22 | 95.5 | -- | -- | -- | -- | -- | | | | | | | |\n| C23 | 2.0 | -.815 | -.681 | -.738 | -.638 | -.524 | 104 | 3.0 | .321 | .335 | .328 | .317 | .298 |\n| 24 | 6.0 | -.281 | -.694 | -.678 | -.635 | -.561 | 105 | 10.0 | .193 | .110 | .100 | .091 | .095 |\n| 25 | 15.0 | -.444 | -.599 | -.621 | -.590 | -.547 | 106 | 25.0 | -.022 | -.035 | -.063 | -.079 | -.093 |\n| 26 | 27.8 | -.359 | -.306 | -.626 | -.631 | -.608 | 107 | 41.0 | -- | -- | -- | -- | -- |\n| 27 | 40.0 | -.313 | -.301 | -.624 | -.642 | -.642 | 108 | 52.5 | -- | -- | -- | -- | -- |\n| 28 | 50.0 | -.253 | -.261 | -.563 | -.547 | -.576 | 109 | 62.5 | -.052 | -.073 | -.102 | -.136 | -.187 |\n| 29 | 59.0 | -.215 | -.213 | -.454 | -.454 | -.453 | 110 | 72.5 | -.028 | -.033 | -.033 | -.033 | -.142 |\n| 30 | 67.5 | -.150 | -.162 | -.193 | -.447 | -.401 | 111 | 85.1 | .027 | .017 | .006 | -.028 | -.107 |\n| 31 | 77.5 | -.093 | -.104 | -.077 | -.115 | -.173 | 112 | 94.6 | .063 | .056 | .047 | .019 | -.080 |\n| 32 | 88.0 | -.003 | -.008 | .001 | -.063 | -.252 | | | | | | | |\n| 33 | 95.5 | -- | -- | -- | -- | -- | | | | | | | |\n| D34 | 2.0 | -.551 | -.600 | -.564 | -.504 | -.430 | 113 | 3.0 | .239 | .243 | .229 | .221 | .198 |\n| 35 | 15.0 | -.380 | -.552 | -.495 | -.486 | -.466 | 114 | 10.0 | .095 | .096 | .087 | .089 | .068 |\n| 36 | 27.5 | -.335 | -.412 | -.451 | -.455 | -.504 | 115 | 25.0 | -.029 | -.049 | -.069 | -.088 | -.094 |\n| 37 | 40.0 | -.323 | -.399 | -.458 | -.514 | -.524 | 116 | 41.0 | -.047 | -.060 | -.078 | -.085 | -.096 |\n| 38 | 50.0 | -.282 | -.312 | -.412 | -.452 | -.478 | 117 | 52.5 | -.054 | -.067 | -.087 | -.097 | -.107 |\n| 39 | 59.0 | -.208 | -.246 | -.301 | -.337 | -.396 | 118 | 62.5 | -- | -- | -- | -- | -- |\n| 40 | 67.5 | -- | -- | -- | -- | -- | 119 | 72.5 | -.009 | -.019 | -.038 | -.051 | -.065 |\n| 41 | 77.5 | -.104 | -.128 | -.178 | -.170 | -.176 | 120 | 87.4 | .057 | .043 | .024 | .006 | .000 |\n| 42 | 87.5 | -.034 | -.046 | -.097 | -.097 | -.099 | 121 | 94.6 | .047 | .034 | .006 | -.009 | -.022 |\n| 43 | 94.2 | .012 | .000 | -.041 | -.050 | -.051 | | | | | | | |\n| E44 | 2.0 | -.295 | -.658 | -.620 | -.570 | -.473 | 122 | 3.0 | .271 | .292 | .273 | .273 | .241 |\n| 45 | 6.0 | -.437 | -.527 | -.544 | -.548 | -.505 | 123 | 10.0 | .127 | .124 | .110 | .110 | .089 |\n| 46 | 15.0 | -.373 | -.455 | -.492 | -.488 | -.476 | 124 | 25.0 | .007 | .009 | .004 | -.005 | -.022 |\n| 47 | 27.5 | -.337 | -.377 | -.477 | -.495 | -.505 | 125 | 41.0 | -.033 | -.039 | -.048 | -.054 | -.067 |\n| 48 | 40.0 | -.323 | -.391 | -.474 | -.505 | -.537 | 126 | 52.5 | -.025 | -.029 | -.043 | -.045 | -.057 |\n| 49 | 50.0 | -.281 | -.319 | -.419 | -.466 | -.496 | 127 | 62.5 | -.022 | -.026 | -.040 | -.045 | -.058 |\n| 50 | 59.0 | -.232 | -.265 | -.357 | -.409 | -.454 | 128 | 72.5 | -.009 | -.020 | -.044 | -.057 | -.102 |\n| 51 | 67.5 | -.167 | -.198 | -.268 | -.309 | -.354 | 129 | 70.0 | .015 | .008 | -.012 | -.018 | -.042 |\n| 52 | 77.5 | -.101 | -.117 | -.130 | -.132 | -.146 | 130 | 85.3 | .047 | .043 | .030 | .027 | .009 |\n| 53 | 88.5 | .033 | .022 | .016 | .003 | -.023 | 131 | 94.1 | .050 | .047 | .040 | .041 | .024 |\n| 54 | 95.5 | .006 | .001 | .009 | -.007 | -.013 | | | | | | | |\n| F55 | 2.0 | -- | -- | -- | -- | -- | 132 | 3.0 | -- | -- | -- | -- | -- |\n| 56 | 6.0 | -.417 | -.495 | -.499 | -.485 | -.420 | 133 | 10.0 | .121 | .129 | .116 | .115 | .093 |\n| 57 | 15.0 | -.356 | -.423 | -.456 | -.455 | -.439 | 134 | 25.0 | .018 | .023 | .009 | .009 | -.012 |\n| 58 | 27.5 | -.310 | -.373 | -.473 | -.495 | -.505 | 135 | 41.0 | -.023 | -.023 | -.022 | -.012 | -.009 |\n| 59 | 49.0 | -.319 | -.395 | -.443 | -.473 | -.482 | 136 | 52.5 | -.006 | -.004 | -.012 | -.012 | -.024 |\n| 60 | 50.0 | -.294 | -.322 | -.366 | -.375 | -.402 | 137 | 62.5 | -- | -- | -- | -- | -.008 |\n| 61 | 59.0 | -.264 | -.322 | -.384 | -.427 | -.433 | 138 | 72.5 | .071 | .054 | .058 | .100 | -.053 |\n| 62 | 67.5 | -.212 | -.222 | -.261 | -.314 | -.344 | 139 | 83.4 | .147 | .124 | .158 | .163 | .157 |\n| 63 | 86.5 | -.040 | -.056 | -.068 | -.082 | -.129", "timestamp": "2026-07-22T04:31:44.956770+00:00"}
{"citation_id": "19930086003", "source_url": "https://ntrs.nasa.gov/api/citations/19930086003/downloads/19930086003.pdf", "page_number": 36, "total_pages": 36, "image_filename": "19930086003_p36.jpg", "text": "```markdown\n34\nNACA RM L9I08\n\n.08\nWing alone\nCONFIDENTIAL\n$(C_D)_{L=0}$ .04\n0\nAspect ratio 6\nAspect ratio 4\n\n.08\nWing fuselage\n$(C_D)_{L=0}$ .04\n0\n.6 .7 .8 .9 1.0 1.1 1.2\nMach number, M\nCONFIDENTIAL\n\nFigure 15.— Effect of aspect ratio on the minimum drag characteristics\nobtained from tests using a sponge-wiper seal for wings with\n45° sweepback, taper ratio 0.6, and NACA 65A006 airfoil section.\n\nNACA-Langley - 11-1-49 - 300\n```", "timestamp": "2026-07-22T04:31:45.704562+00:00"}
{"citation_id": "19930085930", "source_url": "https://ntrs.nasa.gov/api/citations/19930085930/downloads/19930085930.pdf", "page_number": 74, "total_pages": 92, "image_filename": "19930085930_p74.jpg", "text": "NACA RM L9907\n\nCONFIDENTIAL\nUNCLASSIFIED\n\n○ 50-percent-span station\n□ 25-percent-span station\n◇ 10.15-percent-span station\n\nStagnation-pressure recovery, $\\frac{P_2}{P_0}$\n\n1.0\n.8\n.6\n.4\n.2\n0\n\n0 .2 .4 .6 .8 1.0 1.2 1.4 1.6\nDistance from convex surface\n\n← Concave surface\n\nUNCLASSIFIED\nCONFIDENTIAL\n\nNACA\n\nFigure 33.- The variation of the stagnation-pressure recovery with distance from convex surface for model 3.\n\n73", "timestamp": "2026-07-22T04:31:48.287370+00:00"}
{"citation_id": "19930082450", "source_url": "https://ntrs.nasa.gov/api/citations/19930082450/downloads/19930082450.pdf", "page_number": 4, "total_pages": 37, "image_filename": "19930082450_p4.jpg", "text": "NACA TN No. 1778\n3\n\nthe first form of design chart (figs. 2 to 5) dashed lines are used to\nindicate values of average stress at failure $\\overline{\\sigma}_F$; whereas, on the alternate\nform of design chart (figs. 6 to 9) dashed lines are used to indicate\nvalues of $\\frac{P_i}{L/\\sqrt{c}}$. In both forms the value of $\\overline{\\sigma}_F$ corresponding to the\npoint at which each curve is cut by a short heavy line is the value of\nthe stress for local buckling $\\sigma_{cr}$ for the proportions represented by the\ncurves. For example, the value of $\\sigma_{cr}$ for $\\frac{H}{t_W} = 21$ and $\\frac{S}{t_S} = 35$ in\nfigure 2 is approximately 29 ksi. (Only a short panel of these proportions\nwould buckle before failure - one having a value of $\\frac{P_i}{L/\\sqrt{c}} \\ge 0.27$.) If\nthe value of $\\sigma_{cr}$ is so low that the short heavy line would fall outside\nthe boundaries of the chart, a numerical value of $\\sigma_{cr}$ is given and is\nassociated with the proper proportions by a leader to the curve. The panel\nproportions which have minimum weight are indicated on both forms of\nthese charts by the use of colors as follows:\n\n(1) If the proportions correspond to a blue region, they are the\nproportions which give the lightest possible 24S-T Z-stiffened panel\nwhich will meet the design conditions\n\n(2) If the proportions correspond to a red region, they are the\nlightest possible at the ratio of stiffener thickness to skin thickness\ngiven by that particular chart, but some other thickness ratio would\ngive a lighter design\n\n(3) If the proportions correspond to a white region, the proportions\nmeet the design conditions, but they are not the lightest which will\nmeet the conditions\n\nBecause in many cases the proportions may be varied somewhat from\nthose indicated by the red and blue regions with little change in the\nvalue of the stress that can be carried, too much importance should not\nbe attached to the exact proportions indicated by the colors to have\nminimum weight. In any particular case for which a deviation from the\nminimum-weight proportions is made, however, caution dictates that the\nweight penalty associated with this deviation be determined.\n\nThe direct-reading design charts presented herein were developed in\nthe manner described in reference 3 from the test data and resulting\ncurves given in reference 2.", "timestamp": "2026-07-22T04:31:48.840841+00:00"}
{"citation_id": "19930085965", "source_url": "https://ntrs.nasa.gov/api/citations/19930085965/downloads/19930085965.pdf", "page_number": 40, "total_pages": 67, "image_filename": "19930085965_p40.jpg", "text": "NACA RM ESE06\n\n$$\nM_{S} = 3600 \\, m \\, V_{C} \\, \\sin \\alpha\n$$\n\n(B12)\n\nDissipation of heat as result of raising temperature of impinging water. - At the leading edge and for the concave blade surface, the rate of heat dissipation was computed with the assumed collection efficiency. The heat required to raise the temperature of the impinging water to that of the blade was therefore equal to\n\n$$\nH_{3} = M_{C} \\, p_{W} \\, (t_{s} \\, d - t_{c})\n$$\n\n(B13)\n\nneglecting the small kinetic temperature rise in the impinging water.\n\nTotal heat dissipation. - The total heat dissipation for the leading edge and the concave blade surface is the summation of the three types of heat loss\n\n$$\nH = (H_{1} + H_{2}) + H_{3}\n$$\n\n(B14)\n\nwhereas the total heat dissipation of the convex blade surface was assumed to be\n\n$$\nH = H_{1} + H_{2}\n$$\n\n(B15)\n\nThe average heat dissipation presented per square foot of vane surface (fig. 18(a)) is the integrated average of the dissipation rates for both sides of the vane. The heat dissipation due to evaporation, $H_{2}$ in equation (B15), is assumed to apply because some moisture will probably run back on the convex face. Also, the heat requirement for the convex face must be nearly as high as that for the concave face to prevent the development of regions of subfreezing temperatures on the blades. The increase in heating requirement that results from this conservative assumption should help to safeguard the leading edge by action of conduction in the blade metal.\n\n---\n\n**REFERENCES**\n\n1. von Glahn, Uwe: Ice Protection of Turbojet Engines by Inertia Separation of Water. I - Alternate-Duct System. NACA RM E9A27, 1948.\n\n2. Fleming, William A., and Saari, Martin J.: Inlet Icing and Effectiveness of Hot-Gas Bleedback for Ice Protection of Turbojet Engine. NACA RM E8J25c, 1948.", "timestamp": "2026-07-22T04:31:49.554768+00:00"}
{"citation_id": "19930090382", "source_url": "https://ntrs.nasa.gov/api/citations/19930090382/downloads/19930090382.pdf", "page_number": 15, "total_pages": 37, "image_filename": "19930090382_p15.jpg", "text": "NACA RM L9I07\n17\n\nCONFIDENTIAL\n\nTip Mach number, $M_t$\n1.5 1.0 .5 0\n\nEfficiency, $\\eta$\n1.00 .75 .50 .25 0\n\nPower coefficient, $C_P$\n.300 .275 .250 .225 .200 .175 .150 .125 .100 .075 .050 .025 0\n\nThrust coefficient, $C_T$\n.300 .275 .250 .225 .200 .175 .150 .125 .100 .075 .050 .025 0\n\nAdvance ratio, J\n0 .5 1.0 1.5 2.0 2.5 3.0 3.5 4.0 4.5\n\n$\\beta_{0.75}$\n0° 15° 20° 25° 30° 35° 45°\n\n$C_T$\n$C_P$\n$\\eta$\n$M_t$\n\nCONFIDENTIAL\n(b) M=0.23.\nFigure 5 - Continued.\n\nNACA", "timestamp": "2026-07-22T04:31:54.281906+00:00"}
{"citation_id": "19930086078", "source_url": "https://ntrs.nasa.gov/api/citations/19930086078/downloads/19930086078.pdf", "page_number": 26, "total_pages": 42, "image_filename": "19930086078_p26.jpg", "text": "24\nNACA RM L9H04\n\nCONFIDENTIAL\n\n<!-- Image (90, 116, 838, 824) -->\n\n(a) $\\delta_a = 4^\\circ$.\n\nFigure 7.- Lateral control characteristics of unswept wing with large-chord wing-tip aileron at various extensions.", "timestamp": "2026-07-22T04:31:56.369305+00:00"}
{"citation_id": "19930092013", "source_url": "https://ntrs.nasa.gov/api/citations/19930092013/downloads/19930092013.pdf", "page_number": 4, "total_pages": 21, "image_filename": "19930092013_p4.jpg", "text": "# National Advisory Committee for Aeronautics\n\n*Headquarters, 1724 F Street NW., Washington 25, D. C.*\n\nCreated by act of Congress approved March 3, 1915, for the supervision and direction of the scientific study of the problems of flight (U. S. Code, title 50, sec. 151). Its membership was increased from 12 to 15 by act approved March 2, 1929, and to 17 by act approved May 25, 1948. The members are appointed by the President, and serve as such without compensation.\n\nJEROME C. HUNSAKER, Sc. D., Massachusetts Institute of Technology, *Chairman*\n\nALEXANDER WETMORE, Sc. D., Secretary, Smithsonian Institution, *Vice Chairman*\n\nHON. JOHN R. ALISON, Assistant Secretary of Commerce.\nDETLEY W. BRONK, Ph. D., President, Johns Hopkins University.\nKARL T. COMPTON, Ph. D., Chairman, Research and Development Board, Department of Defense.\nEDWARD U. CONDON, Ph. D., Director, National Bureau of Standards.\nJAMES H. DOOLITTLE, Sc. D., Vice President, Shell Union Oil Corp.\nR. M. HAZEN, B. S., Director of Engineering, Allison Division, General Motors Corp.\nWILLIAM LITTLEWOOD, M. E., Vice President, Engineering, American Airlines, Inc.\nTHEODORE C. LONNQUEST, Rear Admiral, United States Navy, Deputy and Assistant Chief of the Bureau of Aeronautics.\n\nDONALD L. PUTT, Major General, United States Air Force, Director of Research and Development, Office of the Chief of Staff, Matériel.\nJOHN D. PRICE, Vice Admiral, United States Navy, Vice Chief of Naval Operations.\nARTHUR E. RAYMOND, Sc. D., Vice President, Engineering, Douglas Aircraft Co., Inc.\nFRANCIS W. REICHELDERFER, Sc. D., Chief, United States Weather Bureau.\nHON. DELOS W. RENTZEL, Administrator of Civil Aeronautics, Department of Commerce.\nHOYT S. VANDENBERG, General, Chief of Staff, United States Air Force.\nTHEODORE P. WRIGHT, Sc. D., Vice President for Research, Cornell University.\n\nHUGH L. DRYDEN, Ph. D., *Director*\nJOHN W. CROWLEY, JR., B. S., *Associate Director for Research*\n\nJOHN F. VICTORY, LL. M., *Executive Secretary*\nE. H. CHAMBERLIN, *Executive Officer*\n\nHENRY J. REID, D. Eng., Director, Langley Aeronautical Laboratory, Langley Field, Va.\nSMITH J. DEFRANCE, B. S., Director, Ames Aeronautical Laboratory, Moffett Field, Calif.\nEDWARD R. SHARP, Sc. D., Director, Lewis Flight Propulsion Laboratory, Cleveland Airport, Cleveland, Ohio\n\n## TECHNICAL COMMITTEES\n\nAERODYNAMICS\nPOWER PLANTS FOR AIRCRAFT\nAIRCRAFT CONSTRUCTION\n\nOPERATING PROBLEMS\nINDUSTRY CONSULTING\n\n*Coordination of Research Needs of Military and Civil Aviation*\n*Preparation of Research Programs*\n*Allocation of Problems*\n*Prevention of Duplication*\n*Consideration of Inventions*\n\nLANGLEY AERONAUTICAL LABORATORY\nLangley Field, Va.\n\nLEWIS FLIGHT PROPULSION LABORATORY\nCleveland Airport, Cleveland, Ohio\n\nAMES AERONAUTICAL LABORATORY\nMoffett Field, Calif.\n\n*Conduct, under unified control, for all agencies of scientific research on the fundamental problems of flight*\n\n## OFFICE OF AERONAUTICAL INTELLIGENCE\nWashington, D. C.\n\n*Collection, classification, compilation, and dissemination of scientific and technical information on aeronautics*\n\nII", "timestamp": "2026-07-22T04:31:56.714028+00:00"}
{"citation_id": "19930085838", "source_url": "https://ntrs.nasa.gov/api/citations/19930085838/downloads/19930085838.pdf", "page_number": 86, "total_pages": 118, "image_filename": "19930085838_p86.jpg", "text": "84\nNACA RM No. L9B23\n\n<!-- Image (101, 99, 840, 763) -->\n\n(a) $\\delta_f = 25^\\circ$.\nFigure 11.- Hinge-moment characteristics of a flap on the approximately 17.7-percent-chord thick NACA 7-series-type airfoil with double slotted flap and straight-sided Frise aileron. Aileron balance, $0.331c_a$; $R = 6.0 \\times 10^6$ (approx.)", "timestamp": "2026-07-22T04:31:56.956289+00:00"}
{"citation_id": "19930093773", "source_url": "https://ntrs.nasa.gov/api/citations/19930093773/downloads/19930093773.pdf", "page_number": 6, "total_pages": 47, "image_filename": "19930093773_p6.jpg", "text": "NACA RM E9G09\n\nengine speeds above 5750 rpm. In the engine-speed range between 4500 to 6600 rpm, the highest specific fuel consumption occurred at an altitude of 5000 feet; at engine speeds above 6600 rpm, the highest specific fuel consumption occurred at an altitude of 50,000 feet. The data indicated no consistent altitude effect at engine speeds below 5750 rpm, probably because of large variations in component efficiencies in the low engine-speed range. The minimum specific fuel consumption of 1.05 pounds per hour per pound of net thrust was obtained at an engine speed of approximately 6400 rpm at altitudes from 15,000 to 45,000 feet. The specific fuel consumption at temperature-limited engine speed varied from 1.20 to 1.30 over the range of altitudes investigated.\n\nThe engine fuel-air ratio (fig. 4(e)) increased with altitude at engine speeds above 4500 rpm. Data obtained at lower engine speeds indicated no consistent altitude effect.\n\nThe exhaust-gas temperature (fig. 4(f)) decreased with an increase in altitude at low engine speeds and increased with altitude at high engine speeds. A change in altitude from 5000 to 25,000 feet resulted in a decrease in temperature-limited engine speed from 7880 to 7550 rpm. The trend of the data indicates that an increase in altitude beyond 25,000 feet would further reduce the maximum permissible engine speed. Inasmuch as maximum thrust is obtained at full engine speed (7900 rpm), and maximum exhaust-gas total temperature, the desirability of using a variable-area exhaust nozzle to permit operation at full engine speed at all altitudes is evident.\n\nEffect of flight Mach number. - Performance data obtained at an altitude of 25,000 feet and flight Mach numbers of 0.21 to 0.97 are presented in figure 5 to show the effect of flight Mach number on net thrust, air flow, fuel flow, specific fuel consumption, fuel-air ratio, and exhaust-gas total temperature.\n\nAs the flight Mach number was raised, the net thrust decreased at engine speeds below 6800 rpm and increased at higher engine speeds for flight Mach numbers above 0.53 (fig. 5(a)). An increase in Mach number from 0.21 to 0.53 at engine speeds above 7000 rpm had no appreciable effect on the net thrust. The engine air flow (fig. 5(b)) increased consistently with an increase in flight Mach number. As the flight Mach number was increased, the engine fuel consumption (fig. 5(c)) decreased at engine speeds below 6150 rpm and increased at higher engine speeds. At temperature-limited engine speed, the specific fuel consumption based on net thrust (fig. 5(d)) increased from 1.21 to 1.43 as", "timestamp": "2026-07-22T04:32:08.190745+00:00"}
{"citation_id": "19930086081", "source_url": "https://ntrs.nasa.gov/api/citations/19930086081/downloads/19930086081.pdf", "page_number": 23, "total_pages": 44, "image_filename": "19930086081_p23.jpg", "text": ".2\n.1\n$C_L$ 0\n-.1\n0\n$C_m$\n-.04\n.02\n$C_D$\n0\n0\n-.004\n$C_n$\n-.008\n-3 -2 -1 0 1 2 3\n$\\alpha$, deg\n\nCONFIDENTIAL\n.024\n$\\delta$\n(deg)\n.020\n$\\circ$ 0.3\n$\\square$ 2.5\n$\\diamond$ 4.1\n$\\triangle$ 7.5\n$\\nabla$ 10.0\n.016\n.012\n.008\n$C_l$\n.004\n0\n-.004\n-.008\n-3 -2 -1 0 1 2 3\n$\\alpha$, deg\nCONFIDENTIAL\n\nNACA RM L9H05\n\nFigure 6.- Aerodynamic characteristics of a semispan delta wing with a half-delta tip control surface tested in the presence of a small fuselage. Large fence on. Three-percent-thick control; R = 4.0 $\\times$ 10$^6$; M = 1.90. Flagged symbols denote repeat tests.\n\n21", "timestamp": "2026-07-22T04:32:10.294853+00:00"}
{"citation_id": "19930085936", "source_url": "https://ntrs.nasa.gov/api/citations/19930085936/downloads/19930085936.pdf", "page_number": 59, "total_pages": 59, "image_filename": "19930085936_p59.jpg", "text": "58\nNACA RM No. E9B03\n\n<!-- Image (119, 89, 812, 279) -->\n\n(a) Angle of yaw, $-12^\\circ$.\n\n<!-- Image (119, 364, 812, 554) -->\n\n(b) Angle of yaw, $0^\\circ$.\n\n<!-- Image (119, 639, 812, 829) -->\n\n(c) Angle of yaw, $12^\\circ$.\n\nFigure 16. - Pressure coefficients on wedge surface at $10^\\circ$ angle of attack for three angles of yaw.\n\nNACA - Langley Field, Va.", "timestamp": "2026-07-22T04:32:10.481305+00:00"}
{"citation_id": "19930082485", "source_url": "https://ntrs.nasa.gov/api/citations/19930082485/downloads/19930082485.pdf", "page_number": 1, "total_pages": 62, "image_filename": "19930082485_p1.jpg", "text": "Y3.N2V5:6/1810\nGOVT. DOC.\n\nNACA TN No. 1810\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\n\nTECHNICAL NOTE\n\nNo. 1810\n\nCOMPARISON BETWEEN PREDICTED AND OBSERVED PERFORMANCE\nOF GAS-TURBINE STATOR BLADE DESIGNED\nFOR FREE-VORTEX FLOW\n\nBy M. C. Huppert and Charles MacGregor\n\nLewis Flight Propulsion Laboratory\nCleveland, Ohio\n\n[Figure: NACA logo]\n\nWashington\nApril 1949\n\nBUSINESS, SCIENCE\n& TECHNOLOGY DEP'T.\nAPR 22 1949", "timestamp": "2026-07-22T04:32:10.782756+00:00"}
{"citation_id": "19930082245", "source_url": "https://ntrs.nasa.gov/api/citations/19930082245/downloads/19930082245.pdf", "page_number": 6, "total_pages": 66, "image_filename": "19930082245_p6.jpg", "text": "NACA TN No. 1596\n\ntunnel or that of free air. At the time of the tests, the maximum Mach number of the tunnel was approximately 0.75.\n\nThe model used in this investigation was a 24-inch-chord airfoil with a 20-percent-chord plain aileron, was of uniform cross section, and spanned the tunnel test section. The main portion of the airfoil passed through the walls of the tunnel with a small clearance gap and was attached to the balance frame of the tunnel in a manner typical of model installation in the Langley 8-foot high-speed tunnel for tests of this type (fig. 1). A gap of 1/16 inch was maintained between the ends of the aileron and the tunnel walls to permit deflection of the aileron. The main portion of the model was of two-steel-spar construction with $\\frac{3}{4}$-inch steel ribs and $\\frac{1}{4}$-inch cold-finished-steel skin built to conform to the ordinates of the NACA 66,1-115 airfoil section as given in table I.\n\nTwo aileron shapes were tested, and these shapes are designated as the true-contour aileron and the beveled-trailing-edge aileron. The general dimensions of the airfoil section are given in figure 2, and of the aileron sections, in figure 3. The profile of the true-contour aileron corresponded to the ordinates of the rear part of the NACA 66,1-115 airfoil section. The profile of the beveled-trailing-edge aileron was formed by a $30^\\circ$ trailing-edge angle and straight lines as shown in figure 3; thus a profile which was thicker than that of the true-contour aileron resulted. The ailerons were constructed of solid dural and were interchangeably attached to the same main portion of the model by six clamp-type hinges. The ailerons had no aerodynamic nose balance. No seal was used between the main portion of the airfoil and the aileron at any time during the tests. The gap between the aileron and the aileron cover plates on both upper and lower surfaces was 0.002c (fig. 2). The cover plates were made of $\\frac{1}{8}$-inch steel.\n\nSufficient static-pressure orifices were installed on the main portion of the model and on the ailerons to determine the complete pressure distribution over the airfoil (fig. 2). The orifices were located in the region of the midspan of the model. The tests were made with a model having aerodynamically smooth surfaces.\n\nTest procedure.— Airfoil normal-force, airfoil pitching-moment, aileron normal-force, and aileron hinge-moment data were determined from static-pressure-distribution measurements which were obtained by photographing a multiple-tube liquid manometer. The tests were made at various angles of attack and aileron deflections. Data were obtained for both aileron configurations at moderate deflections at Mach numbers of 0.25, 0.35, 0.457, 0.55, 0.60, 0.65, 0.70, 0.725, and approximately 0.75; data were obtained also at Mach numbers of 0.45, 0.50, and 0.675 for the true-contour aileron. Data for the larger aileron deflections were", "timestamp": "2026-07-22T04:32:11.930634+00:00"}
{"citation_id": "19930082476", "source_url": "https://ntrs.nasa.gov/api/citations/19930082476/downloads/19930082476.pdf", "page_number": 4, "total_pages": 41, "image_filename": "19930082476_p4.jpg", "text": "```markdown\n2\nNACA TN No. 1801\n\nhas been presented in reference 1. This criterion was based on available\ntest results from the Langley 20-foot free-spinning tunnel of models of\napproximately 60 military designs considered to have proportions of mass\nand dimensional characteristics similar to those of light-airplane designs.\nThis work is now being extended to cover spinproofing as well as spin\nrecovery for a range of model configurations and loadings typical of\npersonal-type aircraft. The results presented herein are for a particular\nmodel having interconnected aileron and rudder controls and limited elevator\ndeflection.\n\nIn addition to determining the effect of simulated two-control operation\nwith the rudders and ailerons linked, the individual effects of the rudders,\nailerons, and elevators in producing a spin for the model were also deter-\nmined in the present investigation. The model was tested for two different\nwing loadings and for three different mass distributions. In the present\nstudy, requirements for spinproofing this particular model were determined\nand an estimate of the probable recovery characteristics was made from a\nstudy of the spin behavior for different control deflections.\n\nThe model used was of such size as to be considered a $\\frac{1}{11}$-scale model of\nan airplane of the personal-owner type. The results are given, therefore,\nin terms of a full-scale airplane on the basis of a $\\frac{1}{11}$-scale model.\n\nSYMBOLS\n\n| Symbol | Definition |\n| :--- | :--- |\n| S | wing area, square feet |\n| b | wing span, feet |\n| m | mass of airplane, slugs |\n| $\\bar{c}$ | mean aerodynamic chord, feet |\n| $x/\\bar{c}$ | ratio of the distance of center of gravity rearward of leading edges of mean aerodynamic chord to the mean aerodynamic chord |\n| $z/\\bar{c}$ | ratio of the perpendicular distance between center of gravity and fuselage reference line to the mean aerodynamic chord (positive when center of gravity is below fuselage reference line) |\n| $I_X, I_Y, I_Z$ | moments of inertia about X, Y, and Z body axes, respectively, slug-feet$^2$ |\n```", "timestamp": "2026-07-22T04:32:13.839203+00:00"}
{"citation_id": "19930085842", "source_url": "https://ntrs.nasa.gov/api/citations/19930085842/downloads/19930085842.pdf", "page_number": 71, "total_pages": 104, "image_filename": "19930085842_p71.jpg", "text": "```markdown\nNACA RM L9C29\n67\n\n<!-- Image (146, 110, 903, 841) -->\n\nFigure 38.- Variation of $C_L$, $C_D$, and $C_Y$ with $\\delta_r$ on a $\\frac{1}{3}$-scale model of the airplane. Propellers removed; basic model configuration; V = 100 miles per hour.\n```", "timestamp": "2026-07-22T04:32:18.853045+00:00"}
{"citation_id": "19930086105", "source_url": "https://ntrs.nasa.gov/api/citations/19930086105/downloads/19930086105.pdf", "page_number": 16, "total_pages": 22, "image_filename": "19930086105_p16.jpg", "text": "```markdown\n14\nCONFIDENTIAL\nNACA RM E9H12\n\no Average of 40 static orifices\n□ Maximum and minimum indicated\nby variable-inductance\npressure pickup\n\n<!-- Image (106, 93, 909, 821) -->\n\n(a) Outlet-inlet area ratio, $A_4/A_1$, 0.7.\n(b) Outlet-inlet area ratio, $A_4/A_1$, 0.8.\n\nFigure 3. - Effect of fuel-air ratio on combustion-chamber-inlet static pressure. Perforated conical flame holder.\n```", "timestamp": "2026-07-22T04:32:27.372442+00:00"}
{"citation_id": "19930082487", "source_url": "https://ntrs.nasa.gov/api/citations/19930082487/downloads/19930082487.pdf", "page_number": 1, "total_pages": 33, "image_filename": "19930082487_p1.jpg", "text": "629.1309\nY3.N21/5:4/1813\nGOVT. DOC.\nNACA TN No. 1813\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\nTECHNICAL NOTE\nNo. 1813\nA STUDY OF FLOW CHANGES ASSOCIATED WITH AIRFOIL\nSECTION DRAG RISE AT SUPERCRITICAL SPEEDS\nBy Gerald E. Nitzberg and Stewart Crandall\nAmes Aeronautical Laboratory\nMoffett Field, Calif.\nNACA\nWashington\nFebruary 1949\nCONN. STATE LIBRARY\nFEB 23 1949\nBUSINESS, SCIENCE\n& TECHNOLOGY DEP'T.", "timestamp": "2026-07-22T04:32:33.980105+00:00"}
{"citation_id": "19930086151", "source_url": "https://ntrs.nasa.gov/api/citations/19930086151/downloads/19930086151.pdf", "page_number": 15, "total_pages": 34, "image_filename": "19930086151_p15.jpg", "text": "NACA RM L9J28 CONFIDENTIAL 13\n\nREFERENCES\n\n1. Knight, Montgomery, and Bamber, Millard J.: Wind Tunnel Tests on a Model of a Monoplane Wing with Floating Ailerons. NACA TN 316, 1929.\n\n2. Weick, Fred E., and Harris, Thomas A.: Wind-Tunnel Research Comparing Lateral Control Devices; Particularly at High Angles of Attack. IV - Floating Tip Ailerons on Rectangular Wings. NACA Rep. 424, 1932.\n\n3. Weick, Fred E., and Harris, Thomas A.: Wind-Tunnel Research Comparing Lateral Control Devices, Particularly at High Angles of Attack. XI. Various Floating Tip Ailerons on Both Rectangular and Tapered Wings. NACA TN 458, 1933.\n\n4. Soulé, H. A., and Gracey, W.: A Flight Comparison of Conventional Ailerons on a Rectangular Wing and of Conventional and Floating Wing-Tip Ailerons on a Tapered Wing. NACA Rep. 630, 1938.\n\n5. Turner, Thomas R., Lockwood, Vernard E., and Vogler, Raymond D.: Preliminary Investigation of Various Ailerons on a $42^\\circ$ Sweptback Wing for Lateral Control at Transonic Speeds. NACA RM L8D21, 1948.\n\n6. Hagerman, John R., and O'Hare, William M.: Investigation of Extensible Wing-Tip Ailerons on an Untapered Semispan Wing at $0^\\circ$ and $45^\\circ$ Sweepback. NACA RM L9H04, 1949.\n\n7. Polhamus, Edward C.: Jet-Boundary-Induced-Upwash Velocities for Swept Reflection-Plane Models Mounted Vertically in 7- by 10-Foot, Closed, Rectangular Wind Tunnels. NACA TN 1752, 1948.\n\n8. Herriot, John G.: Blockage Corrections for Three-Dimensional-Flow Closed-Throat Wind Tunnels, with Consideration of the Effect of Compressibility. NACA RM A7B28, 1947.\n\n9. Swanson, Robert S., and Toll, Thomas A.: Jet-Boundary Corrections for Reflection-Plane Models in Rectangular Wind Tunnels. NACA Rep. 770, 1943.\n\n10. Bates, William R.: Collection and Analysis of Wind-Tunnel Data on the Characteristics of Isolated Tail Surfaces with and without End Plates. NACA TN 1291, 1947.\n\n11. Hemke, Paul E.: Drag of Wings with End Plates. NACA Rep. 267, 1927.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T04:32:37.867928+00:00"}
{"citation_id": "19930090382", "source_url": "https://ntrs.nasa.gov/api/citations/19930090382/downloads/19930090382.pdf", "page_number": 16, "total_pages": 37, "image_filename": "19930090382_p16.jpg", "text": "18\nNACA RM L9I07\n\nCONFIDENTIAL\n\nTip Mach number, $M_t$\n1.5\n1.0\n.5\n0\n\nEfficiency, $\\eta$\n1.00\n.75\n.50\n.25\n0\n\nThrust coefficient, $C_T$\n.300\n.275\n.250\n.225\n.200\n.175\n.150\n.125\n.100\n.075\n.050\n.025\n0\n\nPower coefficient, $C_P$\n.60\n.55\n.50\n.45\n.40\n.35\n.30\n.25\n.20\n.15\n.10\n.05\n0\n\nAdvance ratio, J\n(c) M=0.35\nFigure 5 - Continued.\n\nCONFIDENTIAL\n\n[Graph showing curves for $C_T$, $C_P$, $\\eta$, and $M_t$ versus Advance ratio, J, with $\\beta_0$ at 30°, 35°, 40°, 50°, 60°]\n\nNACA", "timestamp": "2026-07-22T04:32:40.114754+00:00"}
{"citation_id": "19930085930", "source_url": "https://ntrs.nasa.gov/api/citations/19930085930/downloads/19930085930.pdf", "page_number": 75, "total_pages": 92, "image_filename": "19930085930_p75.jpg", "text": "```markdown\nCONFIDENTIAL\nUNCLASSIFIED\n\n74\n\n| | |\n| :--- | :--- |\n| $\\circ$ 50-percent-span station | |\n| $\\square$ 25-percent-span station | |\n| $\\diamond$ 10.15-percent-span station | |\n\nLocal static-pressure ratio, $\\frac{p_2}{p_0}$\n\n.26\n\n.22\n\n.18\n\n.14\n\n.10\n\n.06\n\n0 .2 .4 .6 .8 1.0 1.2 1.4 1.6\n\nDistance from convex surface\n\nConcave surface\n\nUNCLASSIFIED\nCONFIDENTIAL\n\nNACA\n\nFigure 34.- The variation of the local static-pressure ratio with distance from convex surface for model 3.\n\nNACA RM L9907\n```", "timestamp": "2026-07-22T04:32:40.460541+00:00"}
{"citation_id": "19930082450", "source_url": "https://ntrs.nasa.gov/api/citations/19930082450/downloads/19930082450.pdf", "page_number": 5, "total_pages": 37, "image_filename": "19930082450_p5.jpg", "text": "4\nNACA TN No. 1778\n\nUSE OF THE DIRECT-READING DESIGN CHARTS\n\nThe manner of using the direct-reading design charts depends in some measure on the desired degree of precision of interpolation among the curves. For many purposes, interpolation by inspection is of adequate accuracy, and the use of the charts requires only the calculation of the values of the design parameters $P_1/t_S$ and $\\frac{P_1}{L/\\sqrt{C}}$ to permit the desired proportions to be read directly from the curves. The proportions for minimum weight, moreover, may be found directly as those corresponding to the blue region on the curves.\n\nIf more accurate interpolation is desired, a plot can readily be made of $H/t_W$, $\\bar{\\sigma}_F$, and $\\sigma_{cr}$ against $S/t_S$ at the given values of $P_1/t_S$ and $\\frac{P_1}{L/\\sqrt{C}}$ and the proportions can be picked from it. (This plot is similar to that which results from the use of the minimum-weight design procedure with the previously available design charts as illustrated in reference 2.) On a plot of this type, the proportions for minimum weight correspond to those associated with the highest value of $\\bar{\\sigma}_F$.\n\nAs a check on the accuracy of interpolation, the cross-sectional area per inch of width of the design may be determined from the values of $\\bar{t}/t_S$ given in tables 2 to 5 and the value of the intensity of loading $P_1$ that can be carried on this cross-sectional area per inch at the value of $\\bar{\\sigma}_F$ given by the charts may then be compared with the design value of $P_1$.\n\nILLUSTRATIVE EXAMPLE\n\nIn order to illustrate the use of the direct reading design charts and the simplicity of the computations associated with them, a panel will be designed for minimum weight to meet the same principal design conditions used to illustrate the design procedures in reference 2, namely:\n\n(1) Intensity of loading $P_1 = 3.0$ kips per inch\n(2) Skin thickness $t_S = 0.064$ inch\n(3) Effective length $L/\\sqrt{C} = 20$ inches", "timestamp": "2026-07-22T04:32:40.915193+00:00"}
{"citation_id": "19930086078", "source_url": "https://ntrs.nasa.gov/api/citations/19930086078/downloads/19930086078.pdf", "page_number": 27, "total_pages": 42, "image_filename": "19930086078_p27.jpg", "text": "NACA RM L9H04\n25\n\nCONFIDENTIAL\n\n$$\n\\frac{S_a}{S} \\quad \\text{Nominal extension}\n$$\n$$\n\\begin{array}{ll}\n\\Box & 0.010 \\quad \\frac{1}{4} \\\\\n\\circ & .020 \\quad \\frac{1}{2} \\\\\n\\Diamond & .030 \\quad \\frac{3}{4} \\\\\n\\triangle & .040 \\quad \\text{Full}\n\\end{array}\n$$\n\nYawing-moment coefficient, $C_n$\nRolling-moment coefficient, $C_l$\n\nAngle of attack, $\\alpha$, deg\n\nCONFIDENTIAL\nNACA\n\n(b) $\\delta_a = 6^\\circ$.\nFigure 7.- Concluded.", "timestamp": "2026-07-22T04:32:41.621957+00:00"}
{"citation_id": "19930085965", "source_url": "https://ntrs.nasa.gov/api/citations/19930085965/downloads/19930085965.pdf", "page_number": 41, "total_pages": 67, "image_filename": "19930085965_p41.jpg", "text": "40\nNACA RM E9E06\n\n3. Gray, V. H., and Campbell, R. G.: A Method for Estimating Heat\nRequirements for Ice Prevention on Gas-Heated Hollow Propeller\nBlades. NACA TN 1494, 1947.\n\n4. Steinmetz, Charles Proteus: Theory and Calculation of Transient\nElectric Phenomena and Oscillations. McGraw-Hill Book Co.,\nInc., 3d ed., 1920, pp. 361-374.\n\n5. Anon.: Electrical Steel Sheets. Bull. No. 2, Carnegie-Illinois\nSteel Corp. (Pittsburgh, Pa.), 1941, p. 112.\n\n6. Roters, Herbert C.: Electromagnetic Devices. John Wiley & Sons,\nInc., 1941, pp. 116-150.\n\n7. Anon.: Armco Magnetic Ingot Iron. The American Rolling Mill\nCo. (Middletown, Ohio), 1945.\n\n8. Anon.: Westinghouse Metals and Alloys for Communications and\nElectronics Equipment. Westinghouse Electric Corp. (East\nPittsburgh, Pa.).\n\n9. Lewis, William: A Flight Investigation of the Meteorological\nConditions Conducive to the Formation of Ice on Airplanes.\nNACA TN 1393, 1947.\n\n10. Hardy, J. K.: Protection of Aircraft against Ice. Rep. No.\nS.M.E.3380, British R.A.E., July 1946.\n\n11. Martinelli, R. C., Guibert, A. G., Morrin, E. H., and Boelter,\nR. M. K.; An Investigation of Aircraft Heaters. VIII -\nA Simplified Method for the Calculation of the Unit over\nWings. NACA ARR, March 1943.", "timestamp": "2026-07-22T04:32:41.907272+00:00"}
{"citation_id": "19930085838", "source_url": "https://ntrs.nasa.gov/api/citations/19930085838/downloads/19930085838.pdf", "page_number": 87, "total_pages": 118, "image_filename": "19930085838_p87.jpg", "text": "NACA RM No. L9B23\n\n85\n\nFlap section hinge-moment coefficient, $c_{h_f}$\n\n$\\delta_a = 5^\\circ$\n\n$\\delta_1$ (deg)\n\n-3\n\n-2\n\n0\n\n$\\delta_a = 10^\\circ$\n\n$\\delta_1$ (deg)\n\n-2\n\n-1\n\n0\n\n$\\delta_a = 15^\\circ$\n\n$\\delta_1$ (deg)\n\n-2\n\n-1\n\n0\n\nSection angle of attack, $\\alpha_o$, deg\n\n(b) $\\delta_f = 25^\\circ$.\n\nFigure 11.- Continued.", "timestamp": "2026-07-22T04:32:44.046078+00:00"}
{"citation_id": "19930086081", "source_url": "https://ntrs.nasa.gov/api/citations/19930086081/downloads/19930086081.pdf", "page_number": 24, "total_pages": 44, "image_filename": "19930086081_p24.jpg", "text": "22\nNACA RM L9H05\n\nCONFIDENTIAL\n\n<!-- Image (117, 109, 737, 874) -->\n\nFigure 7.- Aerodynamic characteristics of a semispan delta wing with half-delta tip control surface tested in the presence of a small fuselage. Fence off. Seven-percent-thick control; R = 4.0 x 10^6; M = 1.90. Flagged symbols denote repeat runs.", "timestamp": "2026-07-22T04:32:48.898610+00:00"}
{"citation_id": "19930082496", "source_url": "https://ntrs.nasa.gov/api/citations/19930082496/downloads/19930082496.pdf", "page_number": 1, "total_pages": 50, "image_filename": "19930082496_p1.jpg", "text": "NATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\n\nTECHNICAL NOTE\n\nNo. 1836\n\nINITIAL INVESTIGATION OF CARBIDE-TYPE CERAMAL OF\n80-PERCENT TITANIUM CARBIDE PLUS 20-PERCENT\nCOBALT FOR USE AS GAS-TURBINE-BLADE MATERIAL\nBy Charles A. Hoffman, G. Mervin Ault, and James J. Gangler\n\nLewis Flight Propulsion Laboratory\nCleveland, Ohio\n\n[Figure: NACA logo]\n\nWashington\nMarch 1949", "timestamp": "2026-07-22T04:32:51.300568+00:00"}
{"citation_id": "19930082485", "source_url": "https://ntrs.nasa.gov/api/citations/19930082485/downloads/19930082485.pdf", "page_number": 2, "total_pages": 62, "image_filename": "19930082485_p2.jpg", "text": "NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nTECHNICAL NOTE NO. 1810\n\nCOMPARISON BETWEEN PREDICTED AND OBSERVED PERFORMANCE  \nOF GAS-TURBINE STATOR BLADE DESIGNED  \nFOR FREE-VORTEX FLOW  \n\nBy M. C. Huppert and Charles MacGregor\n\nSUMMARY\n\nA comparison was made between the calculated design performance of a gas-turbine stator blade and its performance in a sector of an annular cascade tunnel. Information was obtained with regard to the three-dimensional effects that occur; the influence these effects have upon various performance parameters are presented.\n\nThe gas velocities on the blade surfaces were computed by the stream-filament method and compared with the experimental values. The calculated values satisfactorily agree with the experimental values.\n\nThe design conditions of free-vortex flow were not obtained. The experimental tangential velocities were greater than the design values at the tip of the blade but lower than the design values at the root of the blade.\n\nINTRODUCTION\n\nIn the design of gas-turbine stator blades, compromises are generally made between aerodynamic considerations and such practical considerations as service life and cost of manufacture. In order for a turbine to perform efficiently and to meet design conditions of power and weight flow of gas, however, the blades must be designed by a procedure that accurately predicts blade performance.\n\nThe blade profiles are generally designed according to two-dimensional-flow theory. The velocity distribution along the blade surface must be such that flow separation or excessively high local velocities will be avoided. This velocity distribution may be determined by an incompressible potential-flow theory, which is applicable for any solidity or camber of the blades (references 1 to 4).", "timestamp": "2026-07-22T04:32:52.028296+00:00"}
{"citation_id": "19930085842", "source_url": "https://ntrs.nasa.gov/api/citations/19930085842/downloads/19930085842.pdf", "page_number": 72, "total_pages": 104, "image_filename": "19930085842_p72.jpg", "text": "68\nNACA RM L9C29\n\n[Figure: Graph showing Power coefficient, $C_p$ vs Propeller advance-diameter ratio, $V/nD$. The y-axis ranges from 0 to .24. The x-axis ranges from 0 to .14. There are two sets of curves. The upper set is labeled with $\\alpha, deg$ and $\\beta, deg$ values: 11.3, 30; 5.3, 30; 2.6, 30. The lower set is labeled with $\\alpha, deg$ and $\\beta, deg$ values: 11.4, 20; 3.4, 20; 2.5, 20. The National Advisory Committee for Aeronautics logo is present in the bottom right corner of the graph.]\n\n(a) Variation of $C_p$ with $V/nD$.\n\nFigure 39.- Typical curves showing propulsive characteristics. Basic model configuration; all control surfaces neutral.", "timestamp": "2026-07-22T04:32:56.874593+00:00"}
{"citation_id": "19930082476", "source_url": "https://ntrs.nasa.gov/api/citations/19930082476/downloads/19930082476.pdf", "page_number": 5, "total_pages": 41, "image_filename": "19930082476_p5.jpg", "text": "```markdown\nNACA TN No. 1801\n3\n\n$\\frac{I_X - I_Y}{mb^2}$ inertia yawing-moment parameter\n\n$\\frac{I_Y - I_Z}{mb^2}$ inertia rolling-moment parameter\n\n$\\frac{I_Z - I_X}{mb^2}$ inertia pitching-moment parameter\n\n$\\rho$ air density, slugs per cubic foot\n\n$\\mu$ airplane relative density $\\left(\\frac{m}{\\rho Sb}\\right)$\n\n$\\alpha$ angle between fuselage reference line and vertical (approximately equal to absolute value of angle of attack at plane of symmetry), degrees\n\n$\\phi$ angle between span axis and horizontal, degrees\n\n$V$ full-scale true rate of descent, feet per second\n\n$\\Omega$ full-scale angular velocity about spin axis, revolutions per second\n\nURVC unshielded rudder volume coefficient (see reference 1)\n\nTDR tail damping ratio (see reference 1)\n\nTDPF tail-damping power factor (see reference 1)\n\nFor this model, the helix angle, the angle between the flight path and the vertical, was approximately $7^\\circ$.\n\nSideslip at the center of gravity of the model in the spin is considered inward when the inner wing is down by an amount greater than the helix angle. (Angle of sideslip equals the angle between span axis and horizontal minus the helix angle.)\n\nAPPARATUS AND METHODS\n\nModel\n\nThe $\\frac{1}{11}$-scale model used for the tests corresponded to an airplane of the dimensional characteristics presented in table I. A three-view drawing of the model is given in figure 1 and a photograph of the model is presented in figure 2. The model was tested without a propeller.\n```", "timestamp": "2026-07-22T04:32:59.786899+00:00"}
{"citation_id": "19930093773", "source_url": "https://ntrs.nasa.gov/api/citations/19930093773/downloads/19930093773.pdf", "page_number": 7, "total_pages": 47, "image_filename": "19930093773_p7.jpg", "text": "6\nNACA RM E9G09\n\nthe flight Mach number increased from 0.21 to 0.97. This variation\nof specific fuel consumption based on net thrust with flight Mach\nnumber increased at the low engine speeds. The minimum specific\nfuel consumption of 1.05 pounds per hour per pound of net thrust\noccurred at a flight Mach number of 0.21 and an engine speed of\napproximately 6400 rpm.\n\nThe engine fuel-air ratio (fig. 5(e)) decreased at all engine\nspeeds as the flight Mach number was raised. The exhaust-gas total\ntemperature (fig. 5(f)) was, in general, reduced by an increase in\nflight Mach number at all engine speeds except between 7000 and\n7500 rpm, where a change in flight Mach number had no appreciable\neffect. Maximum engine speed was limited by exhaust-gas total\ntemperature at flight Mach numbers below 0.72.\n\nGeneralized performance. - Altitude performance data for a\nflight Mach number of 0.21 have been generalized to standard sea-\nlevel conditions by use of the correction factors $\\delta$ and $\\theta$\n(reference 1). In the development of this method of generalization,\nit was shown that these correction factors alone were insufficient\nto reduce the results completely to a single curve. The use of\nadditional parameters, such as flight Mach number and Reynolds\nnumber, may be necessary for a complete generalized description of\nengine characteristics. Changes in flight Mach number or changes\nin component efficiency associated with changes in Reynolds number\ntherefore lessen the possibility of reducing data obtained at\nvarious altitudes to a single curve.\n\nPerformance data obtained at a flight Mach number of 0.21\nat altitudes from 5000 to 50,000 feet are presented in figure 6\nto show the effect of altitude on the corrected values of net\nthrust, air flow, fuel flow, specific fuel consumption, fuel-air\nratio, and exhaust-gas total temperature.\n\nThe variation of corrected net thrust with altitude was\nsufficiently small that data obtained at all altitudes from 5000 to\n50,000 feet could be represented by a single curve (fig. 6(a)).\nThe corrected engine air flow (fig. 6(b)) decreased as the altitude\nwas increased at corrected engine speeds above 5400 rpm. For\ncorrected engine speeds below 5400 rpm, the data appear to reduce\nto a single curve.\n\nGeneralized performance parameters depending on fuel consumption\nformed a single curve only near maximum engine speed and at altitudes\nbelow 35,000 feet. Above 35,000 feet and at reduced engine speeds,\nthe corrected fuel consumption (fig. 6(c)) increased as the altitude\nwas raised. Near maximum engine speed, the corrected specific fuel", "timestamp": "2026-07-22T04:33:09.089622+00:00"}
{"citation_id": "19930086151", "source_url": "https://ntrs.nasa.gov/api/citations/19930086151/downloads/19930086151.pdf", "page_number": 16, "total_pages": 34, "image_filename": "19930086151_p16.jpg", "text": "14 CONFIDENTIAL NACA RM L9J28\n\n12. Polhamus, Edward C.: A Simple Method of Estimating the Subsonic Lift and Damping in Roll of Sweptback Wings. NACA TN 1862, 1949.\n\n13. Lange and Wacke: Test Report on Three- and Six-Component Measurements on a Series of Tapered Wings of Small Aspect Ratio (Partial Report: Triangular Wing). NACA TM 1176, 1948.\n\n14. Winter, H.: Flow Phenomena on Plates and Airfoils of Short Span. NACA TM 798, 1936.\n\n15. Goodman, Alex, and Adair, Glenn H.: Estimation of the Damping in Roll of Wings through the Normal Flight Range of Lift Coefficient. NACA TN 1924, 1949.\n\n16. Anon: U. S. Air Force Specification for Flying Qualities of Piloted Airplanes. No. 1815-B, June 1948.\n\n17. Fischel, Jack, and Schneiter, Leslie E.: An Investigation at Low Speed of a $51.3^\\circ$ Sweptback Semispan Wing Equipped with 16.7-Percent-Chord Plain Flaps and Ailerons Having Various Spans and Three Trailing-Edge Angles. NACA RM L8H20, 1948.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T04:33:11.610964+00:00"}
{"citation_id": "19930082487", "source_url": "https://ntrs.nasa.gov/api/citations/19930082487/downloads/19930082487.pdf", "page_number": 2, "total_pages": 33, "image_filename": "19930082487_p2.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T04:33:11.956478+00:00"}
{"citation_id": "19930086105", "source_url": "https://ntrs.nasa.gov/api/citations/19930086105/downloads/19930086105.pdf", "page_number": 17, "total_pages": 22, "image_filename": "19930086105_p17.jpg", "text": "1179\n\nCONFIDENTIAL\n\no Average of 40 static orifices\n□ Maximum and minimum indicated\nby variable-inductance\npressure pickup\n\nCombustion-chamber-inlet static pressure, $P_2/P_0$\nFree-stream total pressure\n\n1.2\n\n1.0\n\n.8\n\n.6\n\n.4\n\n.2\n\n$P_2/P_0$ at optimum\ncold recovery\n\n.01 .02 .03 .04\nFuel-air ratio, f/a\n(c) Outlet-inlet area ratio, $A_4/A_1$, 1.0.\n\nFigure 3. - Concluded. Effect of fuel-air ratio on\ncombustion-chamber-inlet static pressure. Perfo-\nrated conical flame holder.\n\nNACA RM E9H12\n\n□ Maximum and minimum experimental\ninstantaneous pressures at\noptimum recovery (from faired\ncurves of fig. 3)\n\nCombustion-chamber-inlet static pressure, $P_2/P_0$\nFree-stream total pressure\n\n1.0\n\n.9\n\n.8\n\n.7\n\n.6\n\n.5\n\n$P_2/P_0$ at optimum\ncold recovery\n\nWith combustion at\noptimum recovery\n(static tubes)\n\nWithout combustion\n(static tubes)\n\nNACA\n\n.4 .6 .8 1.0 1.2\nOutlet-inlet area ratio, $A_4/A_1$\n\nFigure 4. - Variation of maximum and minimum combustion-\nchamber-inlet static pressures with outlet-inlet area\nratio at optimum mean static pressure. Perforated\nconical flame holder.\n\nCONFIDENTIAL\n\n15", "timestamp": "2026-07-22T04:33:13.115315+00:00"}
{"citation_id": "19930082245", "source_url": "https://ntrs.nasa.gov/api/citations/19930082245/downloads/19930082245.pdf", "page_number": 7, "total_pages": 66, "image_filename": "19930082245_p7.jpg", "text": "```markdown\n6\nNACA TN No. 1596\n\nobtained at maximum test Mach numbers which were lower than those for\nthe moderate deflections. The test procedure consisted of setting the\naileron at a given deflection and then making tests through the angle-\nof-attack range at each of the test Mach numbers. The Reynolds number\nrange of the tests is shown in figure 4.\n\nPRECISION\n\nThe discrepancies between the characteristics of a section at the\nmidspan region of the airfoil as measured in the tunnel and the charac-\nteristics of the airfoil in free air are caused principally by the\neffects of tunnel-wall interference, air leakage through the clearance\ngap between the model and the tunnel walls, and tunnel air-stream\nturbulence.\n\nAn estimate of the tunnel-wall interference corrections which\nincludes the effects of model constriction, wake blockage, and stream-\nline curvature was made for the airfoil with the aileron undeflected\nby the methods of references 5 and 6, which are based on the assump-\ntion that the camber of the airfoil is small. For the value of the\nratio of model airfoil chord to tunnel diameter (0.25) used in the present\ntests, the magnitude of the tunnel-wall-interference corrections is quite\nsmall. These corrections have not been applied to the data. At a Mach\nnumber of 0.70, the Mach number as presented is too low by 2 percent.\nAt a Mach number of 0.70 and an airfoil section normal-force coefficient\nof 0.7, the airfoil section normal-force coefficient is too high by an\nincrement of 0.05, the section pitching-moment coefficient is too high\nby 0.008, and the angle of attack is too low by $0.1^\\circ$. At a Mach number\nof 0.70 and an airfoil section normal-force coefficient of 0.2, the air-\nfoil section normal-force coefficient is too high by an increment\nof 0.015, the section pitching-moment coefficient is too high by 0.003,\nand the angle of attack is too low by $0.02^\\circ$. At lower Mach numbers, the\ncorrections are less than those given at a Mach number of 0.70.\n\nThe apparent choking Mach number of the model in the tunnel, based\non the ratio of the projected thickness of the model to the tunnel\ndiameter, was estimated to be 0.77 (reference 6). The data presented\nherein at a Mach number of 0.75, which was close to the estimated choking\nMach number, should be considered to be of doubtful validity inasmuch as\nthe air flow may have been influenced by the incipient choking restriction.\n\nThe air leakage through the clearance gap between the model and the\ntunnel walls had an insignificant effect since the pressure-distribution\nmeasurements were made near the midspan of the model and the span-chord\nratio (4.0) was large. The numerical effects of air-stream turbulence\nare not known; however, the turbulence level of the Langley 8-foot high-\nspeed tunnel is low.\n```", "timestamp": "2026-07-22T04:33:14.185723+00:00"}
{"citation_id": "19930085838", "source_url": "https://ntrs.nasa.gov/api/citations/19930085838/downloads/19930085838.pdf", "page_number": 88, "total_pages": 118, "image_filename": "19930085838_p88.jpg", "text": "86\nNACA RM No. L9B23\n\n<!-- Image (107, 109, 836, 813) -->\n\n(c) $\\delta_f = 25^\\circ$.\nFigure 11.- Continued.", "timestamp": "2026-07-22T04:33:14.756410+00:00"}
{"citation_id": "19930086078", "source_url": "https://ntrs.nasa.gov/api/citations/19930086078/downloads/19930086078.pdf", "page_number": 28, "total_pages": 42, "image_filename": "19930086078_p28.jpg", "text": "26\nNACA RM L9H04\n\nCONFIDENTIAL\n\n<!-- Image (109, 110, 832, 809) -->\n\nFigure 8.- Lateral control characteristics of unswept wing with triangular wing-tip aileron at various deflections, fully extended.", "timestamp": "2026-07-22T04:33:14.972098+00:00"}
{"citation_id": "19930085930", "source_url": "https://ntrs.nasa.gov/api/citations/19930085930/downloads/19930085930.pdf", "page_number": 76, "total_pages": 92, "image_filename": "19930085930_p76.jpg", "text": "UNCLASSIFIED\nCONFIDENTIAL\n\nNACA RM L9907\n\n[Figure: A shadowgraph of the flow in the passage at an area ratio of 1.208 for model 3.]\n\nFigure 35.- A shadowgraph of the flow in the passage at an area ratio of 1.208 for model 3.\n\nNACA\n\nCONFIDENTIAL\n\n75", "timestamp": "2026-07-22T04:33:15.168090+00:00"}
{"citation_id": "19930092013", "source_url": "https://ntrs.nasa.gov/api/citations/19930092013/downloads/19930092013.pdf", "page_number": 5, "total_pages": 21, "image_filename": "19930092013_p5.jpg", "text": "# REPORT 948\n\n## AN APPARATUS FOR VARYING EFFECTIVE DIHEDRAL IN FLIGHT WITH APPLICATION TO A STUDY OF TOLERABLE DIHEDRAL ON A CONVENTIONAL FIGHTER AIRPLANE\n\nBy WILLIAM M. KAUFFMAN, CHARLES J. LIDDELL, JR., ALLAN SMITH, and RUDOLPH D. VANDYKE, JR.\n\n---\n\n### SUMMARY\n\nAn apparatus for varying effective dihedral in flight by means of servo actuation of the ailerons in response to sideslip angle is described. The results of brief flight tests of the apparatus on a conventional fighter airplane are presented and discussed. The apparatus is shown to have satisfactorily simulated a wide range of effective dihedral under static and dynamic conditions. The effects of a small amount of servo lag are shown to be measurable when the apparatus is simulating small negative values of dihedral. However, these effects were not considered by the pilots to give the airplane an artificial feel.\n\nThe results of an investigation employing the apparatus to determine the tolerable (safe for normal fighter operation) range of effective dihedral on the test airplane are presented. A survey of pilots’ opinions was made to determine which values of effective dihedral were intolerable. It was found that small amounts of negative dihedral (of the order of $-5^\\circ$) as well as values of positive dihedral greater than $20^\\circ$ could be tolerated by the pilots. It was found, in fact, that at landing-approach speeds an effective dihedral high enough ($28.4^\\circ$) to produce oscillatory instability could be tolerated. The occurrence of rolling-velocity reversals during rudder-fixed aileron rolls with high positive values of effective dihedral did not adversely affect the pilots’ opinions of the over-all lateral handling characteristics. The relation between the findings of this investigation and the present Air Force-Navy stability and control specifications is discussed.\n\nAs part of this program, flight tests were planned to determine the effects of changes in effective dihedral on the dynamic-stability characteristics of a conventional fighter airplane. An apparatus for varying the effective dihedral in flight was developed, since this procedure was considered necessary to isolate the effects of stability changes on the airplane behavior from those due to other influences such as air gustiness. This apparatus consists essentially of a servomechanism which deflects the ailerons through a differential linkage in proportion to the movement of a sideslip vane. The ability of the apparatus to simulate satisfactorily changes in dihedral of about $\\pm 9^\\circ$ at high speeds without great practical difficulty led to minor modifications in order to extend the range to approximately twice this value. The revised apparatus thus permitted simulation of the large dihedral range characteristic of swept, triangular, and low-aspect-ratio plan forms. In order to determine some of the difficulties likely to be encountered with extreme effective dihedral, the apparatus then was employed in a flight investigation to evaluate, from measurements and pilots’ opinions, the tolerable limits of effective dihedral for the test airplane.\n\nThe description and flight evaluation of the dihedral-effect control apparatus and flight determination of the tolerable range of effective dihedral of the test airplane have been reported previously in references 4 and 5. This report combines this information into one report and includes some additional information.\n\n---\n\n### INTRODUCTION\n\nFor many years the NACA has been carrying on research in the field of flying qualities of piloted airplanes. A set of preliminary flying-qualities requirements was published in reference 1. This work and the work of other organizations have led to the formulation of flying-qualities requirements by the military services (references 2 and 3). However, these specifications have, in general, been based on experience with airplanes of conventional configuration and, hence, with more or less conventional handling characteristics. The use of highly swept-back wings, triangular wings, and wings of low aspect ratio for airplanes to be operated at very high speeds and altitudes has introduced stability and control characteristics which hitherto had not been considered in flying-qualities work. A reexamination of certain aspects of the present flying-qualities requirements has, therefore, been initiated. Particular emphasis is being given to the dynamic lateral and directional motions.\n\n---\n\n### NOTATION\n\n| Symbol | Description |\n|--------|-------------|\n| $\\Gamma_e$ | effective dihedral, degrees |\n| $V$ | true airspeed, feet per second |\n| $V_i$ | indicated airspeed, knots |\n| $q$ | dynamic pressure, pounds per square foot |\n| $S$ | wing area, square feet |\n| $b$ | wing span, feet |\n| $\\delta_a$ | total aileron deflection (sum of left and right aileron deflections, left when left aileron is up), degrees |\n| $(\\delta_a)_s$ | $\\delta_a$ due to servo action, degrees |\n| $\\delta_r$ | rudder deflection, degrees |\n| $\\delta_t$ | aileron tab deflection (positive when tab located on left aileron is up), degrees |\n| $\\theta$ | lateral stick deflection, degrees |\n\n--- \n\n[Page number: 1]", "timestamp": "2026-07-22T04:33:16.828430+00:00"}
{"citation_id": "19930086081", "source_url": "https://ntrs.nasa.gov/api/citations/19930086081/downloads/19930086081.pdf", "page_number": 25, "total_pages": 44, "image_filename": "19930086081_p25.jpg", "text": "NACA RM L9H05\n23\n\nCONFIDENTIAL\n\n| $\\delta$ (deg) |\n| :--- |\n| $\\circ$ 0 |\n| $\\square$ 2.1 |\n| $\\diamond$ 4.0 |\n| $\\triangleright$ 6.2 |\n| $\\triangleleft$ 8.2 |\n| $\\triangledown$ 10.2 |\n| $\\nabla$ 12.1 |\n\n$C_l$\n.036\n.032\n.028\n.024\n.020\n.016\n.012\n.008\n.004\n0\n-.004\n-.008\n-.012\n-.016\n-.020\n-.024\n-.028\n-.032\n-.036\n\n-7 -6 -5 -4 -3 -2 -1 0 1 2 3 4 5 6 7 8\n$\\alpha$, deg\n\nCONFIDENTIAL\nNACA\n\n(b) Variation of $C_l$ with $\\alpha$.\nFigure 7.- Concluded.", "timestamp": "2026-07-22T04:33:21.634126+00:00"}
{"citation_id": "19930085842", "source_url": "https://ntrs.nasa.gov/api/citations/19930085842/downloads/19930085842.pdf", "page_number": 73, "total_pages": 104, "image_filename": "19930085842_p73.jpg", "text": "NACA RM L9C29\n69\n\n[Figure: A graph plotting Thrust coefficient, $C_{Te}$ against Propeller advance-diameter ratio, $V/nD$. The vertical axis ranges from 0 to .24. The horizontal axis ranges from 0 to 14. The graph contains multiple curves with data labels. One set of labels reads: $\\beta, deg$ $\\alpha, deg$ / 20 44 / 20 42 / 20 43. Another set of labels reads: $\\alpha, deg$ $\\beta, deg$ / 11.3 30 / 5.3 30 / -0.6 30. A box in the lower right corner contains the text: NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS.]\n\n(b) Variation of $C_{Te}$ with $V/nD$.\n\nFigure 39.— Continued.", "timestamp": "2026-07-22T04:33:24.957783+00:00"}
{"citation_id": "19930082447", "source_url": "https://ntrs.nasa.gov/api/citations/19930082447/downloads/19930082447.pdf", "page_number": 6, "total_pages": 24, "image_filename": "19930082447_p6.jpg", "text": "4\nNACA TN No. 1775\n\nHorizontal velocity, feet per second . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .", "timestamp": "2026-07-22T04:33:27.865202+00:00"}
{"citation_id": "19930082476", "source_url": "https://ntrs.nasa.gov/api/citations/19930082476/downloads/19930082476.pdf", "page_number": 6, "total_pages": 41, "image_filename": "19930082476_p6.jpg", "text": "4\nNACA TN No. 1801\n\nFor the tests, the model was ballasted with lead weights to represent an airplane at an altitude of 5000 feet ($\\rho = 0.002049$ slug/cu ft). The normal weight, moments of inertia, and center of gravity of the airplane were selected on the basis of dimensions of an airplane typical of this type.\n\nWind Tunnel and Testing Technique\n\nThe tests were performed in the Langley 20-foot free-spinning tunnel, the operation of which is generally similar to that for the Langley 15-foot free-spinning tunnel described in reference 2 except that the model launching technique has been changed. With the controls set in the desired position, the model is now launched by hand with rotation into the vertically rising air stream. After the model assumes a fairly constant spin attitude, the spin parameters $\\alpha$, $\\Omega$, $\\beta$, and V are measured and recorded. The model values are converted to full-scale values by methods described in reference 2. For the spins which have a rate of descent in excess of that which can readily be obtained in the tunnel, either the rate of descent is recorded as greater than the velocity at the time the model hits the safety net or the spin is referred to in a footnote on the chart as merely a \"steep spin.\" When the model after being launched with forced rotation into a spin stopped rotating without movement of the controls, the result is recorded as a \"no spin\" condition. A photograph of the model during a spin in the tunnel is shown in figure 3.\n\nRecoveries from steady spins were not attempted for this model because it appeared that recovery characteristics could be estimated with sufficient accuracy. The turns required for recovery are normally considered from the time the controls are moved until the time the spin rotation ceases.\n\nThe term \"linked controls\" used throughout this paper indicates that the rudders and ailerons were set in such a manner as to simulate an inter-connection between them for two-control operation of the airplane. Thus, when rudders were set with the spin (right wheel in a right spin), the ailerons were also with the spin (right aileron up and left aileron down in a right spin). The term \"wheel setting\" refers to the control wheel of the airplane and indicates the deflection of the ailerons and rudders; \"wheel with the spin\" indicates that for a right spin the right aileron is up, the left aileron is down, and both rudders are deflected to the right.\n\nPRECISION\n\nThe model test results presented are believed to be the true values given by the model within the following limits:", "timestamp": "2026-07-22T04:33:30.037796+00:00"}
{"citation_id": "19930090382", "source_url": "https://ntrs.nasa.gov/api/citations/19930090382/downloads/19930090382.pdf", "page_number": 17, "total_pages": 37, "image_filename": "19930090382_p17.jpg", "text": "NACA RM L9I07\n19\n\nCONFIDENTIAL\n\nTip Mach number, $M_t$\nEfficiency, $\\eta$\n\nPower coefficient, $C_P$\nThrust coefficient, $C_T$\n\nAdvance ratio, J\n(d) M=0.43.\nFigure 5 - Continued.", "timestamp": "2026-07-22T04:33:30.230710+00:00"}
{"citation_id": "19930086105", "source_url": "https://ntrs.nasa.gov/api/citations/19930086105/downloads/19930086105.pdf", "page_number": 18, "total_pages": 22, "image_filename": "19930086105_p18.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T04:33:31.421896+00:00"}