PDF/ocr_dataset_1007.jsonl

{"citation_id": "19930085936", "source_url": "https://ntrs.nasa.gov/api/citations/19930085936/downloads/19930085936.pdf", "page_number": 47, "total_pages": 59, "image_filename": "19930085936_p47.jpg", "text": "46\nNACA RM No. E9B03\n\n<!-- Image (148, 129, 875, 864) -->\n\n(e) $\\theta = 270^\\circ$ longitudinal plane.\nFigure 8. - Concluded. Pressure distributions along longitudinal\nplanes at $10^\\circ$ angle of attack for range of yaw angles.", "timestamp": "2026-07-22T04:23:38.785949+00:00"}
{"citation_id": "19930086078", "source_url": "https://ntrs.nasa.gov/api/citations/19930086078/downloads/19930086078.pdf", "page_number": 14, "total_pages": 42, "image_filename": "19930086078_p14.jpg", "text": "12 CONFIDENTIAL NACA RM L9H04\n\n1. Sufficient aileron effectiveness was generally obtained at moderate and high lift coefficients with the extensible ailerons investigated. However, the control effectiveness at low lift coefficients appears to be inadequate for satisfactory application to an airplane. It is thought that the extensible ailerons may be sufficiently effective for some types of missiles.\n\n2. Yawing moments produced by the extensible ailerons investigated were comparable to those produced by conventional flap-type ailerons.\n\nLangley Aeronautical Laboratory\nNational Advisory Committee for Aeronautics\nLangley Air Force Base, Va.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T04:23:40.248133+00:00"}
{"citation_id": "19930085952", "source_url": "https://ntrs.nasa.gov/api/citations/19930085952/downloads/19930085952.pdf", "page_number": 44, "total_pages": 50, "image_filename": "19930085952_p44.jpg", "text": "NACA RM L9C24\n43\n\n$C_{h_a}$\n0\n-1\n-2\n\n$C_L$\n10.2\n9.8\n9.4\n9.0\n.2\n.1\n0\n\n$C_m$\n-1\n-2\n-3\n\n$\\delta_a, deg$\n-48 -40 -32 -24 -16 -8 0\n\nO Articulated propellers; V/nD, 0.20\n$\\square$ Rigid propellers; V/nD, 0.19\n\n[Figure: NACA logo]\n\n(a) $\\alpha \\approx 69^\\circ$.\n\nFigure 20.- Concluded.", "timestamp": "2026-07-22T04:23:40.927756+00:00"}
{"citation_id": "19930085842", "source_url": "https://ntrs.nasa.gov/api/citations/19930085842/downloads/19930085842.pdf", "page_number": 59, "total_pages": 104, "image_filename": "19930085842_p59.jpg", "text": "NACA RM L9C29\n\nIncrement of pitching-moment coefficient due to tail, $\\Delta C_{m_t}$\n\nAngle of attack, $\\alpha$, deg\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nFigure 26.- Variation of increment of pitching-moment coefficient due to horizontal tail with angle of attack. Basic model configuration; propellers removed.\n\n55", "timestamp": "2026-07-22T04:23:45.641012+00:00"}
{"citation_id": "19930086015", "source_url": "https://ntrs.nasa.gov/api/citations/19930086015/downloads/19930086015.pdf", "page_number": 29, "total_pages": 54, "image_filename": "19930086015_p29.jpg", "text": "```markdown\n28\n\nCONFIDENTIAL\n\nNACA RM A9E24\n\nFlagged symbols\ndenote survey\noff center line\n\nVertical distance from tunnel center line, z, in.\n\n| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | 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| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | 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{"citation_id": "19930085930", "source_url": "https://ntrs.nasa.gov/api/citations/19930085930/downloads/19930085930.pdf", "page_number": 61, "total_pages": 92, "image_filename": "19930085930_p61.jpg", "text": "60\nNACA RM L9G07\n\n[Figure: Graph showing Average exit Mach number, $M_{2,av}$ vs Area ratio, $\\frac{A_2}{A_1}$]\n\nCONFIDENTIAL\n\n$\\circ \\quad \\frac{p_1}{p_A} = 1.28$\n\n$\\square \\quad \\frac{p_1}{p_A} = 1.00$\n\n$\\diamond \\quad \\frac{p_1}{p_A} = .85$\n\nCONFIDENTIAL\n\nNACA\n\n(a) 50-percent-span station.\n\nFigure 25.- The variation of the average Mach number with area ratio for three static-pressure ratios for model 2.", "timestamp": "2026-07-22T04:23:46.551449+00:00"}
{"citation_id": "19930086022", "source_url": "https://ntrs.nasa.gov/api/citations/19930086022/downloads/19930086022.pdf", "page_number": 26, "total_pages": 34, "image_filename": "19930086022_p26.jpg", "text": "24\nNACA RM L9E24\n\n[Figure: Graph of $P_R$ vs. $\\delta_a$ (deg)]\n$P_R$\n.8\n.4\n0\n-.4\n-.8\n$\\delta_a$\n(deg)\n25\n20\n15\n10\n5\n0\n-5\n-10\n-15\n-20\n-25\n\n[Figure: Graph of $C_{h_a}$ vs. $\\alpha$, deg]\n.16\n.12\n.08\n.04\n0\n-.04\n-.08\n-.12\n-.16\n-.20\n-4 0 4 8 12 16 20\n$\\alpha$, deg\nNACA\n\n(b) $P_R$ and $C_{h_a}$ plotted against $\\alpha$.\nFigure 6.— Continued.", "timestamp": "2026-07-22T04:23:50.056644+00:00"}
{"citation_id": "19930085965", "source_url": "https://ntrs.nasa.gov/api/citations/19930085965/downloads/19930085965.pdf", "page_number": 32, "total_pages": 67, "image_filename": "19930085965_p32.jpg", "text": "NACA RM E9E06\n31\n\nFrom the vector diagram, the following relation is true:\n$$I_{max} d\\beta = \\gamma E_{max} dl \\cos (45^\\circ+\\beta-\\delta) \\tag{A6}$$\nor, when equations (A5) and (A6) are combined,\n$$d\\beta = \\frac{\\cos (45^\\circ+\\beta-\\delta)}{\\sin (45^\\circ+\\beta-\\delta)} \\frac{\\frac{0.4 \\pi \\gamma E_{max}}{H_{max} l} dl \\sin (45^\\circ+\\beta-\\delta)}{1 + \\int \\frac{0.4 \\pi \\gamma E_{max}}{H_{max} l} dl \\sin (45^\\circ+\\beta-\\delta)} \\tag{A7}$$\nIntegrating from 0 to $\\beta_2$ and from 0 to $l_2$, where $\\alpha_2$ is the average of $\\beta_2-\\delta_2$ as before, yields\n$$\\beta_2 = \\cot (45^\\circ+\\alpha_2) \\log_e \\left[ 1 + \\int_0^{l_2} \\frac{0.4 \\pi \\gamma E_{max}}{H_{max} l} dl \\sin(45^\\circ+\\beta-\\delta) \\right] \\tag{A8}$$\nThe first step in the evaluation of the integral is to replace $E_{max}$ by using equations (10) and (12).\n$$\\int_0^{l_2} \\frac{0.4 \\pi \\gamma E_{max}}{H_{max} l} dl \\sin (45^\\circ+\\beta-\\delta) =$$\n$$\\int_0^{l_2} \\frac{0.4 \\pi \\gamma 2\\pi f 10^{-8}}{H_{max} l} \\left[ \\frac{E_{max} l}{\\sqrt{2} c} + E_{max} l \\cos (45^\\circ+\\alpha) \\right] dl \\sin (45^\\circ+\\beta-\\delta)$$\nor\n$$= \\int_0^{l_2} \\left[ \\sqrt{2} c + 2c^2 l \\cos (45^\\circ+\\alpha) \\right] dl \\sin (45^\\circ+\\beta-\\delta)$$\nor where $\\alpha$ is again the average value of $(\\beta-\\delta)$,\n$$\\int_0^{l_2} \\frac{0.4 \\pi \\gamma E_{max}}{H_{max} l} dl \\sin (45^\\circ+\\beta-\\delta) = \\left[ \\sqrt{2} c l_2 + c^2 l_2^2 \\cos (45^\\circ+\\alpha_2) \\right] \\sin (45^\\circ+\\alpha_2) \\tag{A9}$$", "timestamp": "2026-07-22T04:23:50.418060+00:00"}
{"citation_id": "19930085551", "source_url": "https://ntrs.nasa.gov/api/citations/19930085551/downloads/19930085551.pdf", "page_number": 77, "total_pages": 82, "image_filename": "19930085551_p77.jpg", "text": "76\nNACA RM No. L8K30\n\n<!-- Image (143, 110, 740, 869) -->\n\nFigure 21.- Lateral and longitudinal trim characteristics with asymmetric power; trim with trim tabs; C-54D airplane. Clean condition; power; No. 1 engine idling; No. 2, 3, and 4 normal rated power.", "timestamp": "2026-07-22T04:23:51.174015+00:00"}
{"citation_id": "19930086060", "source_url": "https://ntrs.nasa.gov/api/citations/19930086060/downloads/19930086060.pdf", "page_number": 15, "total_pages": 20, "image_filename": "19930086060_p15.jpg", "text": "NACA RM L9F02\n13\n\nCONFIDENTIAL\n\n<!-- Image (115, 109, 834, 883) -->\n\n(a) Maximum diameter at 20-percent station.\nFigure 5.- Theoretical pressure distribution at zero lift and M = 1.40.\nCONFIDENTIAL", "timestamp": "2026-07-22T04:23:56.237151+00:00"}
{"citation_id": "19930085529", "source_url": "https://ntrs.nasa.gov/api/citations/19930085529/downloads/19930085529.pdf", "page_number": 80, "total_pages": 85, "image_filename": "19930085529_p80.jpg", "text": "```markdown\nNACA RM No. L8A30a\n79\n\nTABLE T3\n$$\n\\left[ \\Lambda = -45^\\circ, \\delta_{N_1} = -10.0^\\circ, \\alpha = -2^\\circ \\right]\n$$\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| A1 | 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 | -0.048 | -0.037 | -0.019 | 0.000 | -0.040 |\n| 7 | 59.0 | -0.082 | -0.093 | -0.095 | -0.090 | -0.181 | 92 | 72.5 | -0.016 | -0.011 | -0.001 | -0.015 | -0.007 |\n| 8 | 67.5 | -0.091 | -0.095 | -0.060 | -0.096 | -0.083 | 93 | 84.0 | -- | -- | -- | -- | -- |\n| 9 | 77.5 | -- | -- | -- | -- | -- | 94 | 94.0 | -- | -- | -- | -- | -- |\n| 10 | 87.5 | -- | -- | -- | -- | -- | | | | | | | |\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 | -0.141 | -0.211 | -0.441 | -0.928 | -0.612 |\n| 15 | 27.5 | -0.135 | -0.176 | -0.225 | -0.277 | -0.297 | 98 | 41.0 | -0.128 | -0.140 | -0.199 | -0.277 | -0.346 |\n| 16 | 40.0 | -0.154 | -0.194 | -0.239 | -0.304 | -0.362 | 99 | 52.5 | -0.110 | -0.115 | -0.115 | -0.136 | -0.285 |\n| 17 | 50.0 | -0.151 | -0.185 | -0.220 | -0.267 | -0.378 | 100 | 62.5 | -0.085 | -0.082 | -0.075 | -0.075 | -0.172 |\n| 18 | 59.0 | -0.129 | -0.156 | -0.184 | -0.207 | -0.368 | 101 | 72.5 | -0.032 | -0.031 | -0.026 | -0.021 | -0.060 |\n| 19 | 67.5 | -0.086 | -0.110 | -0.125 | -0.140 | -0.292 | 102 | 86.3 | -0.014 | -0.015 | -0.018 | -0.024 | -0.023 |\n| 20 | 77.5 | -0.046 | -0.059 | -0.065 | -0.092 | -0.198 | 103 | 94.5 | -0.004 | -0.002 | -0.003 | -0.008 | -0.019 |\n| 21 | 88.0 | -0.001 | -0.015 | -0.014 | -0.026 | -0.051 | | | | | | | |\n| 22 | 95.3 | -- | -- | -- | -- | -- | | | | | | | |\n| C23 | 2.0 | -0.418 | -0.432 | -0.463 | -0.473 | -0.478 | 104 | 3.0 | -- | -- | -- | -- | -0.882 |\n| 24 | 6.0 | -0.175 | -0.201 | -0.201 | -0.206 | -0.230 | 105 | 10.0 | -- | -- | -- | -0.460 | -0.711 |\n| 25 | 15.0 | -0.061 | -0.065 | -0.063 | -0.066 | -0.117 | 106 | 25.0 | -0.293 | -0.397 | -0.427 | -0.406 | -0.460 |\n| 26 | 27.5 | -0.103 | -0.130 | -0.155 | -0.156 | -0.153 | 107 | 41.0 | -- | -- | -- | -- | -- |\n| 27 | 40.0 | -0.146 | -0.186 | -0.213 | -0.255 | -0.251 | 108 | 52.5 | -- | -- | -- | -- | -- |\n| 28 | 50.0 | -0.141 | -0.176 | -0.232 | -0.303 | -0.303 | 109 | 62.5 | -0.105 | -0.128 | -0.165 | -0.289 | -0.323 |\n| 29 | 59.0 | -0.132 | -0.164 | -0.207 | -0.270 | -0.296 | 110 | 72.5 | -0.055 | -0.068 | -0.083 | -0.154 | -0.255 |\n| 30 | 67.5 | -0.091 | -0.115 | -0.147 | -0.200 | -0.251 | 111 | 85.1 | -0.006 | -0.015 | -0.024 | -0.040 | -0.172 |\n| 31 | 77.5 | -0.055 | -0.075 | -0.100 | -0.113 | -0.201 | 112 | 94.6 | -0.043 | -0.008 | -0.018 | -0.006 | -0.117 |\n| 32 | 88.0 | -0.008 | -0.009 | -0.014 | -0.036 | -0.138 | | | | | | | |\n| 33 | 95.3 | -- | -- | -- | -- | -- | | | | | | | |\n| D34 | 2.0 | -0.325 | -0.344 | -0.347 | -0.348 | -0.342 | 113 | 3.0 | -0.598 | -0.766 | -0.798 | -0.791 | -0.780 |\n| 35 | 15.0 | -0.061 | -0.061 | -0.061 | -0.066 | -0.067 | 114 | 10.0 | -0.316 | -0.363 | -0.390 | -0.610 | -0.621 |\n| 36 | 27.5 | -0.078 | -0.088 | -0.094 | -0.099 | -0.101 | 115 | 25.0 | -0.260 | -0.318 | -0.345 | -0.401 | -0.423 |\n| 37 | 40.0 | -0.113 | -0.136 | -0.156 | -0.176 | -0.181 | 116 | 41.0 | -0.176 | -0.210 | -0.231 | -0.253 | -0.314 |\n| 38 | 50.0 | -0.139 | -0.163 | -0.181 | -0.191 | -0.191 | 117 | 52.5 | -0.181 | -0.222 | -0.258 | -0.290 | -0.271 |\n| 39 | 59.0 | -0.134 | -0.156 | -0.163 | -0.163 | -0.163 | 118 | 62.5 | -- | -- | -- | -- | -- |\n| 40 | 67.5 | -- | -- | -- | -- | -- | 119 | 72.5 | -0.065 | -0.089 | -0.119 | -0.129 | -0.132 |\n| 41 | 77.5 | -0.069 | -0.066 | -0.070 | -0.071 | -0.101 | 120 | 87.4 | -0.017 | -0.021 | -0.030 | -0.038 | -0.053 |\n| 42 | 87.5 | -0.019 | -0.019 | -0.041 | -0.061 | -0.068 | 121 | 94.2 | -0.048 | -0.009 | -0.017 | -0.034 | -0.035 |\n| 43 | 94.2 | -0.038 | -0.017 | -0.004 | -0.008 | -0.077 | | | | | | | |\n| E44 | 2.0 | -0.333 | -0.347 | -0.347 | -0.347 | -0.321 | 122 | 3.0 | -0.478 | -0.678 | -0.697 | -0.691 | -0.667 |\n| 45 | 6.0 | -0.140 | -0.152 | -0.147 | -0.150 | -0.135 | 123 | 10.0 | -0.295 | -0.342 | -0.361 | -0.421 | -0.462 |\n| 46 | 15.0 | -0.060 | -0.061 | -0.060 | -0.065 | -0.053 | 124 | 25.0 | -0.250 | -0.300 | -0.321 | -0.370 | -0.401 |\n| 47 | 27.5 | -0.072 | -0.080 | -0.089 | -0.089 | -0.104 | 125 | 41.0 | -0.214 | -0.254 | -0.300 | -0.333 | -0.402 |\n| 48 | 40.0 | -0.118 | -0.138 | -0.157 | -0.167 | -0.181 | 126 | 52.5 | -0.167 | -0.203 | -0.233 | -0.271 | -0.357 |\n| 49 | 50.0 | -0.132 | -0.138 | -0.155 | -0.158 | -0.168 | 127 | 62.5 | -0.129 | -0.152 | -0.169 | -0.167 | -0.205 |\n| 50 | 59.0 | -0.108 | -0.125 | -0.144 | -0.159 | -0.168 | 128 | 72.5 | -0.080 | -0.101 | -0.108 | -0.101 | -0.101 |\n| 51 | 67.5 | -0.085 | -0.101 | -0.118 | -0.136 | -0.164 | 129 | 78.0 | -0.042 | -0.057 | -0.072 | -0.076 | -0.089 |\n| 52 | 77.5 | -0.049 | -0.061 | -0.070 | -0.111 | -0.111 | 130 | 85.3 | -0.017 | -0.021 | -0.029 | -0.035 | -0.037 |\n| 53 | 88.5 | -0.089 | -0.078 |", "timestamp": "2026-07-22T04:23:56.415053+00:00"}
{"citation_id": "19930082090", "source_url": "https://ntrs.nasa.gov/api/citations/19930082090/downloads/19930082090.pdf", "page_number": 111, "total_pages": 122, "image_filename": "19930082090_p111.jpg", "text": "NACA TN No. 1455\n109\n\n[Figure: A cylindrical heat exchanger with a copper ring at one end and a dense array of stainless-steel fins extending along its length.]\n\nFigure 57.- Copper and stainless-steel heat exchanger O.", "timestamp": "2026-07-22T04:23:59.185421+00:00"}
{"citation_id": "19930086003", "source_url": "https://ntrs.nasa.gov/api/citations/19930086003/downloads/19930086003.pdf", "page_number": 25, "total_pages": 36, "image_filename": "19930086003_p25.jpg", "text": "```markdown\nNACA RM L9J08\n\nCONFIDENTIAL\n\nM\n1.15 ◇\n1.10 □\n1.08 ▽\n1.05 △\n1.03 ▽\n1.00 △\n.98 ◇\n.95 ◇\n.93 ◇\n.90 △\n.88 ▽\n.85 △\n.80 ◇\n.70 □\n.60 ◇\n\nAngle of attack, $\\alpha$, deg\n0\n0\n0\n0\n0\n0\n0\n0\n12\n8\n4\n0\n0\n-4\n\nLift coefficient, $C_L$\n-.2 0 .2 .4 .6 .8\n\nPitching-moment coefficient, $C_m$\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0.1\n0\n-.1\n\nLift coefficient, $C_L$\n-.2 0 .2 .4 .6 .8\n\nCONFIDENTIAL\n\nM\n1.15 ◇\n1.10 □\n1.08 ▽\n1.05 △\n1.03 ▽\n1.00 △\n.98 ◇\n.95 ◇\n.93 ◇\n.90 △\n.88 ▽\n.85 △\n.80 ◇\n.70 □\n.60 ◇\n\nDrag coefficient, $C_D$\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0.04\n0\n0.08\n0.12\n\nLift coefficient, $C_L$\n-.2 0 .2 .4 .6 .8\n\n[Figure: Three graphs showing aerodynamic characteristics. Left graph: Angle of attack vs. Lift coefficient for various Mach numbers. Middle graph: Pitching-moment coefficient vs. Lift coefficient for various Mach numbers. Right graph: Drag coefficient vs. Lift coefficient for various Mach numbers. All graphs have a grid background and use different symbols for different Mach numbers as indicated in the legends.]\n\nFigure 9.-- Aerodynamic characteristics for a model with 45° sweptback wing, aspect ratio 6, taper ratio 0.6, and NACA 65A006 airfoil section. Wing fuselage.\n\n23\n```", "timestamp": "2026-07-22T04:24:06.410657+00:00"}
{"citation_id": "19930085936", "source_url": "https://ntrs.nasa.gov/api/citations/19930085936/downloads/19930085936.pdf", "page_number": 48, "total_pages": 59, "image_filename": "19930085936_p48.jpg", "text": "NACA RM No. E9B03\n47\n\nPressure\ncoefficient\n$C_p$\n.4\n.2\n0\n-.2\n-.2\n0\n.2\n.4\n.6\n\nAngle of\nattack\n(deg)\n-15\n-6\n0\n12\n24\n\n[Figure: Polar plot showing radial pressure distributions with various line styles corresponding to different angles of attack.]\n\nNACA\n\n(a) x/L = 0.148.\n\nFigure 9. - Radial pressure distributions at $0^\\circ$ yaw angle for various\nangles of attack.", "timestamp": "2026-07-22T04:24:07.479048+00:00"}
{"citation_id": "19930085958", "source_url": "https://ntrs.nasa.gov/api/citations/19930085958/downloads/19930085958.pdf", "page_number": 37, "total_pages": 43, "image_filename": "19930085958_p37.jpg", "text": "36\nNACA RM No. L9B11\n\n1.6\n1.2\n.8\n.4\n0\n-.08\n0 8 16 24 32\n$\\alpha$, deg\n$C_L$\n$C_m$\n\n1.6\n1.2\n.8\n.4\n0\n-.08\n0 8 16 24 32\n$\\alpha$, deg\n$C_L$\n$C_m$\n\n1.6\n1.2\n.8\n.4\n0\n-.08\n0 8 16 24 32\n$\\alpha$, deg\n$C_L$\n$C_m$\n\nCross flow\nRough\nIntermittently stalled\nCompletely stalled\n\n0.25$\\bar{c}$\n$\\alpha=11.0^\\circ$\n$C_L=0.96$\n\n$\\alpha=6.8^\\circ$\n$C_L=0.76$\n\n$\\alpha=15.3^\\circ$\n$C_L=1.26$\n\n$\\alpha=15.2^\\circ$\n$C_L=1.17$\n\n$\\alpha=11.1^\\circ$\n$C_L=1.01$\n\n$\\alpha=17.4^\\circ$\n$C_L=1.33$\n\n$\\alpha=17.3^\\circ$\n$C_L=1.27$\n\n$\\alpha=15.3^\\circ$\n$C_L=1.26$\n\n$\\alpha=19.5^\\circ$\n$C_L=1.45$\n\n$\\alpha=21.3^\\circ$\n$C_L=1.23$\n\n$\\alpha=17.4^\\circ$\n$C_L=1.32$\n\n$\\alpha=21.6^\\circ$\n$C_L=1.51$\n\n$\\alpha=25.4^\\circ$\n$C_L=1.29$\n\n$\\alpha=20.5^\\circ$\n$C_L=1.43$\n\n$\\alpha=22.6^\\circ$\n$C_L=1.51$\n\nNACA\n\n(a) Low wing.\n(b) Midwing.\n(c) High wing.\n\nFigure 17.- Stalling characteristics of a $42^\\circ$ sweptback wing-fuselage combination with $0.55\\frac{b}{2}$ extensible leading-edge flaps and split flaps.", "timestamp": "2026-07-22T04:24:08.080636+00:00"}
{"citation_id": "19930085952", "source_url": "https://ntrs.nasa.gov/api/citations/19930085952/downloads/19930085952.pdf", "page_number": 45, "total_pages": 50, "image_filename": "19930085952_p45.jpg", "text": "44\nNACA RM L9C24\n\n<!-- Image (168, 109, 739, 836) -->\n\n(a) Variation of M and H with $\\delta_B$.\nFigure 21.- Effect of propeller operation and forward velocity on\nallavator effectiveness. Articulated propellers; $\\alpha_u = 90^\\circ$;\n$\\beta = 11.5^\\circ$; 2500 rpm; $\\delta_F = 0^\\circ$.", "timestamp": "2026-07-22T04:24:08.446197+00:00"}
{"citation_id": "19930085838", "source_url": "https://ntrs.nasa.gov/api/citations/19930085838/downloads/19930085838.pdf", "page_number": 75, "total_pages": 118, "image_filename": "19930085838_p75.jpg", "text": "NACA RM No. L9B23\n73\n\n<!-- Image (175, 111, 891, 935) -->\n\n(j) $\\delta_f = 40^\\circ$.\nFigure 9.- Continued.", "timestamp": "2026-07-22T04:24:13.863880+00:00"}
{"citation_id": "19930085842", "source_url": "https://ntrs.nasa.gov/api/citations/19930085842/downloads/19930085842.pdf", "page_number": 60, "total_pages": 104, "image_filename": "19930085842_p60.jpg", "text": "```markdown\n56\n\n$\\alpha = 23.2^\\circ$\n$\\alpha = 11.3^\\circ$\n$\\alpha = -0.6^\\circ$\n\n$\\delta_{a_{TR}}, deg$\n$\\delta_{a_{TR}}, deg$\n$\\delta_{a_{TR}}, deg$\n\nRight-Allavator hinge-moment coefficient, $C_{h_a}$\n-30\n-20\n-10\n10\n20\n-2\n-4\n-48 -40 -32 -24 -16 -8 0 8 16\nAllavator deflection, $\\delta_a$, deg\n\nNATIONAL ADVISORY\nCOMMITTEE FOR AERONAUTICS\n\nFigure 27.- The effect of right-allavator tab setting on the variation of $C_{h_a}$ with $\\delta_a$. Model in basic configuration; $\\delta_{a_{TL}} = 0^\\circ$; $\\delta_r = 0^\\circ$; propellers removed; $\\delta_F = 0^\\circ$.\n\nNACA RM 19C29\n```", "timestamp": "2026-07-22T04:24:18.515208+00:00"}
{"citation_id": "19930086078", "source_url": "https://ntrs.nasa.gov/api/citations/19930086078/downloads/19930086078.pdf", "page_number": 15, "total_pages": 42, "image_filename": "19930086078_p15.jpg", "text": "NACA RM L9H04 CONFIDENTIAL 13\n\nREFERENCES\n\n1. Regenscheit, B.: Versuche mit einem Flügelendquerruder an einem Rechteckflügel. UM Nr. 3219, Aerodynamische Versuchsanstalt Göttingen, Feb. 2, 1945.\n\n2. 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\n3. 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\n4. Swanson, Robert S., and Toll, Thomas A.: Jet-Boundary Corrections for Reflection-Plane Models in Rectangular Wind Tunnels. NACA Rep. 770, 1943.\n\n5. Wood, K. D.: Aspect Ratio Corrections. Jour. Aero. Sci., vol. 10, no. 8, Oct. 1943, pp. 270-272.\n\n6. Polhamus, Edward C.: A Simple Method of Estimating the Subsonic Lift and Damping in Roll of Sweptback Wings. NACA TN 1862, 1949.\n\n7. Loftin, Laurence K., Jr.: Theoretical and Experimental Data for a Number of NACA 6A-Series Airfoil Sections. NACA TN 1368, 1947.\n\n8. Abbott, Ira H., Von Doenhoff, Albert E., and Stivers, Louis S., Jr.: Summary of Airfoil Data. NACA Rep. 824, 1945.\n\n9. Goodman, Alex, and Brewer, Jack D.: Investigation at Low Speeds of the Effect of Aspect Ratio and Sweep on Static and Yawing Stability Derivatives of Untapered Wings. NACA TN 1669, 1948.\n\n10. Anon: Wind-Tunnel Tests of the KAQ-1 Airplane. Part IV - Aileron Effectiveness for Various Partial Extensions. TED No. TMB 2328. David Taylor Model Basin, U.S.N., Washington, D. C., July 29, 1946.\n\n11. Toll, Thomas H., and Queijo, M. J.: Approximate Relations and Charts for Low-Speed Stability Derivatives of Swept Wings. NACA TN 1581, 1948.\n\n12. Anon: Flying Qualities of Piloted Airplanes. U. S. Air Force Specification No. 1815-B, June 1, 1948.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T04:24:19.990529+00:00"}
{"citation_id": "19930086081", "source_url": "https://ntrs.nasa.gov/api/citations/19930086081/downloads/19930086081.pdf", "page_number": 12, "total_pages": 44, "image_filename": "19930086081_p12.jpg", "text": "10 CONFIDENTIAL NACA RM L9H05\n\ncalculated values of 0.0045 and -0.0015, respectively. With regard to rolling moment, the curves of the three fence configurations (3-percent-thick control, fig. 10(a)), agreed within the experimental accuracy, and the average value of $C_{l\\delta}$ was 0.00076 for the small fuselage as compared with a calculated value of 0.00090. This experimental value was used, as described in the section entitled \"Results,\" to obtain a value of $\\frac{pb}{2V}\\delta$ of 0.0058 radian per degree for M = 1.9 (fig. 11). The value of $\\frac{pb}{2V}\\delta$, which is the ratio of $C_{l\\delta}$ to $C_{l_p}$ (both based on the same area and span), was in good agreement with theory. The agreement was somewhat fortuitous, however, since both experimental values ($C_{l\\delta}$ and $C_{l_p}$) were about 15 percent lower than the corresponding calculated values. The data of figure 11 show agreement between the rocket test results and theory at supersonic speeds.\n\nEffect of fence.- Adding a fence to the wing at the juncture of the wing panel and the control surface did not appreciably change the aerodynamic characteristics of the wing. There was no change in the drag characteristics or in the rolling effectiveness of the control (within the experimental accuracy), and the use of a fence of the dimensions investigated at this Mach number appears unwarranted.\n\nEffect of airfoil-section modifications.- Increasing the control-surface thickness from 3 to 7 percent and moving the position of maximum thickness far forward increased the minimum drag coefficient less than 0.001 (fig. 12). Lift effectiveness of the control was decreased slightly and was accompanied by a 15-percent decrease in control rolling effectiveness.\n\nWhen the airfoil sections comprising the outer one-third of the exposed wing panel were increased from 3- to 9-percent thickness (with the thickness distribution unchanged), the minimum drag coefficient was increased about 0.002 (fig. 12). In the initial configuration of the wing with 9-percent-thick tip sections, the wing airfoil sections were composed of flat-side elements with unfaired intersections. Rounding the nose of the airfoil increased the minimum drag, whereas incorporating fairing to eliminate sharp breaks in contour decreased the minimum drag. As the lift coefficient was increased, the differences in drag for the various configurations decreased. These results are for tests where the Mach line lies just ahead of the wing leading edge.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T04:24:21.049019+00:00"}
{"citation_id": "19930086022", "source_url": "https://ntrs.nasa.gov/api/citations/19930086022/downloads/19930086022.pdf", "page_number": 27, "total_pages": 34, "image_filename": "19930086022_p27.jpg", "text": "NACA RM L9B24\n25\n\n<!-- Image (263, 172, 753, 362) -->\n\n<!-- Image (263, 420, 753, 778) -->\n\n(c) $C_m$ and $C_L$ plotted against $\\alpha$.\nFigure 6.- Concluded.", "timestamp": "2026-07-22T04:24:22.104865+00:00"}
{"citation_id": "19930085930", "source_url": "https://ntrs.nasa.gov/api/citations/19930085930/downloads/19930085930.pdf", "page_number": 62, "total_pages": 92, "image_filename": "19930085930_p62.jpg", "text": "NACA RM L9G07\n61\n\n[Figure: Graph of Average exit Mach number, $M_{2_{av}}$ vs Area ratio, $\\frac{A_2}{A_1}$]\n\nCONFIDENTIAL\n\n$\\circ \\frac{p_1}{p_A} = 1.28$\n$\\square \\frac{p_1}{p_A} = 1.00$\n$\\diamond \\frac{p_1}{p_A} = .85$\n\nCONFIDENTIAL\n\nNACA\n\n(b) 10.15-percent-span station.\n\nFigure 25.- Concluded.", "timestamp": "2026-07-22T04:24:25.088126+00:00"}
{"citation_id": "19930085965", "source_url": "https://ntrs.nasa.gov/api/citations/19930085965/downloads/19930085965.pdf", "page_number": 33, "total_pages": 67, "image_filename": "19930085965_p33.jpg", "text": "32\nNACA RM ESE06\n\nWith substitution of equation (A9), equation (A8) becomes\n$$\n\\beta_2 = \\cot(45^\\circ+\\alpha_2) \\log_e \\left\\{ 1 + \\left[ \\sqrt{2} \\text{ c } l_2 + \\text{c}^2 l_2^2 \\cos(45^\\circ+\\alpha_2) \\right] \\sin(45^\\circ+\\alpha_2) \\right\\} \\quad \\text{(A10)}\n$$\n\nSubstitution of equation (15b) into equation (A10) gives the solution for $\\beta_2$\n$$\n\\beta_2 = \\cot(45^\\circ+\\alpha_2) \\log_e \\left[ 1 + \\sin(45^\\circ+\\alpha_2) \\left( \\frac{H_{\\text{max } 2}}{H_{\\text{max } 1}} - 1 \\right) \\right] \\quad \\text{(A11)}\n$$", "timestamp": "2026-07-22T04:24:25.311631+00:00"}
{"citation_id": "19930086105", "source_url": "https://ntrs.nasa.gov/api/citations/19930086105/downloads/19930086105.pdf", "page_number": 7, "total_pages": 22, "image_filename": "19930086105_p7.jpg", "text": "NACA RM E9H12 CONFIDENTIAL 5\n\nResults of the instantaneous pressure measurements are presented in figure 3 for values of outlet-inlet area ratio $A_4/A_1$ of 0.7, 0.8, and 1.0. The ratio of combustion-chamber static pressure to free-stream total pressure $p_3/P_0$ is shown as a function of fuel-air ratio f/a. For each outlet area, the mean static pressure gradually increased to an optimum value and then dropped off with increasing fuel-air ratio. The instantaneous data indicated that pressure fluctuations occurred before the optimum mean static pressure. For values of f/a greater than that at the optimum mean static pressure, the pulsations became intense and the maximum instantaneous pressures in some cases exceeded the free-stream stagnation pressure and thus indicated the occurrence of explosive combustion.\n\nWith increasing values of outlet-inlet area ratio, the optimum mean static pressure decreased and occurred at a higher fuel-air ratio. The pressure fluctuations before the optimum mean static pressure increased in amplitude with increasing values of $A_4/A_1$ and extended over a larger range of f/a.\n\nValues of optimum mean $p_3/P_0$ with the corresponding maximum and minimum data from figure 3 are plotted as a function of $A_4/A_1$ in figure 4. Within experimental accuracy, the maximum static pressure data coincided with the static pressure at optimum cold recovery, and the minimum static pressure data fell along the curve obtained without combustion. This correlation seemed to substantiate further the theory in reference 5 that at optimum recovery the limits of the pressure variation would correspond to conditions at blow-out and optimum heat release. It now appears that, for the perforated conical flame holder, the maximum diffuser pressure recovery in the presence of rough combustion will occur when the maximum instantaneous pressure ratio equals the optimum cold $p_3/P_0$.\n\nFrequency measurements (figs. 3(a) and 3(b)) appear to be somewhat inadequate but do indicate the order of magnitude. The fundamental frequency did not remain constant but increased from 6 to 35 cycles per second as the fuel-air ratio increased beyond the value at the optimum mean static pressure. High-speed schlieren photographs of the oscillating shock indicated much the same order of frequency. For illustration, two sequences of motion pictures are shown in figure 5. The first set reveals the steady-shock pattern upstream of the diffuser inlet when the engine was operating under conditions of optimum total-pressure recovery. The second sequence (fig. 5(b)) shows that portion of the cycle where the shock was\n\nCONFIDENTIAL", "timestamp": "2026-07-22T04:24:28.457868+00:00"}
{"citation_id": "19930085551", "source_url": "https://ntrs.nasa.gov/api/citations/19930085551/downloads/19930085551.pdf", "page_number": 78, "total_pages": 82, "image_filename": "19930085551_p78.jpg", "text": "NACA RM No. L8K30\n77\n\nSideslip angle, deg\nRight\n10\n0\n\nRolling velocity, rad/sec\nRight\n.2\n.1\n0\n\nControl force, lb\nRight\n40\n0\nAileron\nRudder\n\nControl angle, deg\nRight\n20\n10\n0\nTotal aileron\nRudder\n0 2 4\nTime, sec\n\nSideslip angle, deg\nLeft\n0\n10\n\nRolling velocity, rad/sec\nLeft\n0\n.1\n\nControl force, lb\nRight\n40\n0\nLeft\n40\nRudder\nAileron\n\nControl angle, deg\nRight\n10\n0\nLeft\n10\n20\nRudder\nTotal aileron\n0 2 4 6 8\nTime, sec\n\nNACA\n\nFigure 22.- Time histories of rudder fixed aileron rolls. C-54D airplane; clean condition; power for level flight; 120 miles per hour.", "timestamp": "2026-07-22T04:24:29.300040+00:00"}
{"citation_id": "19930090382", "source_url": "https://ntrs.nasa.gov/api/citations/19930090382/downloads/19930090382.pdf", "page_number": 5, "total_pages": 37, "image_filename": "19930090382_p5.jpg", "text": "4\nCONFIDENTIAL\nNACA RM L9I07\n\nwas varied. The range of blade angle covered for each forward Mach number is given in the following table:\n\n| Forward Mach number, M | Blade angle at 0.7-radius station, $\\beta_{0.7R}$ (deg) | | | | | | | | | |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| 0.175 | 20 | 25 | 30 | | | | | | | |\n| .23 | 20 | 25 | 30 | 35 | | 45 | | | | |\n| .35 | | | 30 | 35 | 40 | | 50 | 60 | | |\n| .43 | | | | 35 | 40 | 45 | | 55 | 65 | |\n| .53 | | | | | 40 | 45 | 50 | 55 | 60 | |\n| .60 | | | | | | 45 | 50 | 55 | | 65 |\n| .65 | | | | | | 45 | 50 | 55 | 60 | 65 |\n| .70 | | | | | | | 50 | 55 | | 65 |\n| .75 | | | | | | | 50 | 55 | 60 | 65 | 70 |\n| .80 | | | | | | | | 55 | 60 | 65 | 70 |\n| .85 | | | | | | | | | 60 | 65 | 70 |\n| .90 | | | | | | | | | 60 | 65 | |\n| .925 | | | | | | | | | | 65 | |\n\nREDUCTION OF DATA\n\nPropeller thrust.- Propeller thrust, as used herein, is defined as the shaft tension produced by the spinner-to-tip portion of the blades. The method used in determining thrust tares and in evaluating the propeller thrust is described in detail in reference 1.\n\nPropeller torque.- Torque-tare corrections were found to be small and depended only on spinner rotational speed. The indicated torque reading was corrected for the spinner tare (a maximum of 1.2 foot-pounds at 6000 rpm).\n\nCONFIDENTIAL", "timestamp": "2026-07-22T04:24:29.533966+00:00"}
{"citation_id": "19930086060", "source_url": "https://ntrs.nasa.gov/api/citations/19930086060/downloads/19930086060.pdf", "page_number": 16, "total_pages": 20, "image_filename": "19930086060_p16.jpg", "text": "14\nNACA RM L9F02\n\nCONFIDENTIAL\n\n.4\n.3\n.2\n.1\n$C_p$ 0\n-.1\n-.2\n\n$K = 40 \\text{ percent}$\n\nNACA\n\n(b) Maximum diameter at 40-percent station.\nFigure 5.- Continued.\nCONFIDENTIAL", "timestamp": "2026-07-22T04:24:30.186898+00:00"}
{"citation_id": "19930082090", "source_url": "https://ntrs.nasa.gov/api/citations/19930082090/downloads/19930082090.pdf", "page_number": 112, "total_pages": 122, "image_filename": "19930082090_p112.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T04:24:34.900394+00:00"}
{"citation_id": "19930086003", "source_url": "https://ntrs.nasa.gov/api/citations/19930086003/downloads/19930086003.pdf", "page_number": 26, "total_pages": 36, "image_filename": "19930086003_p26.jpg", "text": "24\nNACA RM L9I08\n\nCONFIDENTIAL\n\n<!-- Image (137, 109, 773, 853) -->\n\nM\n1.15 $\\diamond$\n1.10 $\\square$\n1.08 $\\nabla$\n1.05 $\\nabla$\n1.03 $\\nabla$\n1.00 $\\triangle$\n.98 $\\diamond$\n.95 $\\diamond$\n.93 $\\square$\n.90 $\\square$\n.88 $\\triangle$\n.85 $\\triangle$\n.80 $\\diamond$\n.70 $\\square$\n.60 $\\circ$\n\nNACA\n\nLift coefficient, $C_L$\nCONFIDENTIAL\nFigure 9.- Concluded.", "timestamp": "2026-07-22T04:24:38.801292+00:00"}
{"citation_id": "19930085936", "source_url": "https://ntrs.nasa.gov/api/citations/19930085936/downloads/19930085936.pdf", "page_number": 49, "total_pages": 59, "image_filename": "19930085936_p49.jpg", "text": "48\nNACA RM No. E9B03\n\nAngle of\nattack\n(deg)\nPressure\ncoefficient\n$C_p$\n-15\n-6\n0\n12\n24\n\n<!-- Image (153, 119, 846, 872) -->\n\n(b) x/L = 0.898.\nFigure 9. - Concluded. Radial pressure distributions at $0^\\circ$ yaw angle\nfor various angles of attack.", "timestamp": "2026-07-22T04:24:39.562687+00:00"}
{"citation_id": "19930085958", "source_url": "https://ntrs.nasa.gov/api/citations/19930085958/downloads/19930085958.pdf", "page_number": 38, "total_pages": 43, "image_filename": "19930085958_p38.jpg", "text": "```markdown\nNACA RM No. L9E11\n\nGliding speed\n(mph)\nSinking speed\n(fps)\n\n<!-- Image (100, 100, 890, 780) -->\n\nFigure 18.- Glide characteristics of a 42° sweptback wing for a wing loading of 40 pounds per square foot; standard sea-level conditions.\n\n37\n```", "timestamp": "2026-07-22T04:24:41.541065+00:00"}
{"citation_id": "19930085952", "source_url": "https://ntrs.nasa.gov/api/citations/19930085952/downloads/19930085952.pdf", "page_number": 46, "total_pages": 50, "image_filename": "19930085952_p46.jpg", "text": "NACA RM L9C24\n45\n\n<!-- Image (259, 172, 821, 738) -->\n\n(b) Variation of lift with $\\delta_a$.\nFigure 21.- Concluded.", "timestamp": "2026-07-22T04:24:42.579212+00:00"}
{"citation_id": "19930085838", "source_url": "https://ntrs.nasa.gov/api/citations/19930085838/downloads/19930085838.pdf", "page_number": 76, "total_pages": 118, "image_filename": "19930085838_p76.jpg", "text": "74\nNACA RM No. L9B23\n\n<!-- Image (113, 100, 850, 898) -->\n\n(k) $\\delta_f = 40^\\circ$.\nFigure 9.- Continued.", "timestamp": "2026-07-22T04:24:43.311762+00:00"}
{"citation_id": "19930085982", "source_url": "https://ntrs.nasa.gov/api/citations/19930085982/downloads/19930085982.pdf", "page_number": 29, "total_pages": 32, "image_filename": "19930085982_p29.jpg", "text": "NACA RM E9E13\n27\n\n1131\n\n| | | | | | | | | | |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| | | | | | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | 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| | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | 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| | | | | | |\n| | | | | | | | | | |\n| | | | | | | | | | |\n| | |", "timestamp": "2026-07-22T04:24:51.292634+00:00"}
{"citation_id": "19930085842", "source_url": "https://ntrs.nasa.gov/api/citations/19930085842/downloads/19930085842.pdf", "page_number": 61, "total_pages": 104, "image_filename": "19930085842_p61.jpg", "text": "NACA RM L9C29\n\nPitching-moment coefficient, $C_m$\n\n$\\alpha = 23.2^\\circ$\n$\\delta_{e_{TR}}, deg$\n-30\n-20\n-10\n10\n20\n\n$\\alpha = 11.3^\\circ$\n$\\delta_{e_{TR}}, deg$\n-30\n-20\n-10\n10\n20\n\n$\\alpha = 0.6^\\circ$\n$\\delta_{e_{TR}}, deg$\n-30\n-20\n-10\n10\n20\n\nNATIONAL ADVISORY\nCOMMITTEE FOR AERONAUTICS\n\nAllavator deflection, $\\delta_a$, deg\n\nFigure 23.- The effect of right-elevator tab setting on the variation of $C_m$ with $\\delta_a$. Model in basic configuration; $\\delta_{e_{TL}} = 0^\\circ$; $\\delta_r = 0^\\circ$; propellers removed; $\\delta_f = 0^\\circ$.\n\n57", "timestamp": "2026-07-22T04:24:56.372053+00:00"}
{"citation_id": "19930086078", "source_url": "https://ntrs.nasa.gov/api/citations/19930086078/downloads/19930086078.pdf", "page_number": 16, "total_pages": 42, "image_filename": "19930086078_p16.jpg", "text": "14 CONFIDENTIAL NACA RM L9H04\n\n13. Goodman, Alex, and Fisher, Lewis R.: Investigation at Low Speeds of the Effect of Aspect Ratio and Sweep on Rolling Stability Derivatives of Untapered Wings. NACA TN 1835, 1949.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T04:24:58.604982+00:00"}
{"citation_id": "19930086022", "source_url": "https://ntrs.nasa.gov/api/citations/19930086022/downloads/19930086022.pdf", "page_number": 28, "total_pages": 34, "image_filename": "19930086022_p28.jpg", "text": "26\nNACA RM L9E24\n\n.01\n$C_n$ 0\n-.01\n$\\delta_{a_{total}}$\n50°\n\nLeading and\ntrailing-edge flaps\nFences\noff\nOff\nOn\nOff\nOn\nOn\n\n.06\n.05\n.04\n$C_l$ .03\n.02\n.01\n0\n50°\n30°\n12°\nNACA\n0\n4\n8\n12\n16\n20\n$\\alpha$, deg\n\nFigure 7.- Rolling-moment and yawing-moment characteristics for several\nmodel configurations and total aileron deflections.", "timestamp": "2026-07-22T04:25:03.331568+00:00"}
{"citation_id": "19930085930", "source_url": "https://ntrs.nasa.gov/api/citations/19930085930/downloads/19930085930.pdf", "page_number": 63, "total_pages": 92, "image_filename": "19930085930_p63.jpg", "text": "62\nNACA RM L9G07\n\n[Figure: Graph showing Average stagnation-pressure recovery vs Area ratio]\n\nCONFIDENTIAL\nO $\\frac{P_1}{P_A} = 1.28$\n$\\square$ $\\frac{P_1}{P_A} = 1.00$\n$\\diamond$ $\\frac{P_1}{P_A} = .85$\n\nAverage stagnation-pressure recovery, $\\frac{P_{dav}}{P_o}$\n\nArea ratio, $\\frac{A_2}{A_1}$\n\n(a) 50-percent-span station.\n\nCONFIDENTIAL\nNACA\n\nFigure 26.- The variation of the average stagnation-pressure recovery with area ratio for three static-pressure ratios for model 2.", "timestamp": "2026-07-22T04:25:07.840393+00:00"}
{"citation_id": "19930086151", "source_url": "https://ntrs.nasa.gov/api/citations/19930086151/downloads/19930086151.pdf", "page_number": 7, "total_pages": 34, "image_filename": "19930086151_p7.jpg", "text": "```markdown\nNACA RM L9J28 CONFIDENTIAL 5\n\nto the ceiling (fig. 2), the ceiling thereby acting as a reflection plane. The wing, exclusive of ailerons, was constructed of steel and mahogany to the plan-form dimensions shown in figure 1. The wing had NACA 64A010 airfoil sections normal to the wing leading edge and had neither twist nor dihedral. The wing tip was a body of revolution.\n\nA vertical end plate which roughly approximated a vertical tail surface was mounted on the main part of the wing, inboard of the wing-tip body of revolution, for a portion of the investigation. This end plate was a $\\frac{1}{2}$-inch-thick sheet of plywood with rounded edges and was cut to the plan-form dimensions and mounted on the wing as shown in figure 1.\n\nTwo plan forms of wing-tip ailerons were used in the present investigation; one aileron had a parallelogram plan form, and the other a triangular plan form. Both ailerons had root chords of 0.625c and equal areas (fig. 1). The ailerons were constructed of $\\frac{1}{4}$-inch-sheet duralumin with a rounded leading edge and a 12° beveled trailing edge along the entire span of each aileron. The trailing edges of both ailerons were swept back 45°. The ailerons were deflected about a spanwise axis passing through the 0.5-tip-chord station of the wing and the 0.5-root-chord station of the aileron.\n\nAlthough the ailerons investigated did not have a conventional airfoil section, as would probably be the case in a practical application, the ailerons are believed to simulate an actual airplane arrangement sufficiently well to supply representative data.\n\nTESTS\n\nAll tests of the 45° sweptback wing with the parallelogram- and triangular-plan-form wing-tip ailerons were performed in the Langley 300 MPH 7- by 10-foot tunnel at a dynamic pressure of approximately 50.5 pounds per square foot, which corresponds to a Mach number of 0.19 and a Reynolds number of about $4.4 \\times 10^6$ based on the wing mean aerodynamic chord.\n\nThe aerodynamic characteristics in pitch were determined for the wing-aileron configurations with and without the end plate through an angle-of-attack range from positive to negative wing stall. The lateral control characteristics of each wing-aileron configuration (with and without the end plate) were also determined through a similar angle-of-attack range at various aileron deflections between 0° and approximately 30°.\n\nCONFIDENTIAL\n```", "timestamp": "2026-07-22T04:25:11.011486+00:00"}
{"citation_id": "19930086081", "source_url": "https://ntrs.nasa.gov/api/citations/19930086081/downloads/19930086081.pdf", "page_number": 13, "total_pages": 44, "image_filename": "19930086081_p13.jpg", "text": "NACA RM L9H05 CONFIDENTIAL 11\n\nControl-Surface Characteristics\n\nDivision of loading between the control surface and the inboard panel of the wing.- A comparison of the slope values for the curves of figures 8 and 14 gives a measure of the part of the angle-of-attack loading carried on the control. The control surface covered 11 percent of the wing area and carried 18 percent of the lift load due to changing angle of attack $C_{L_\\alpha}$. High tip loading would be expected, because the highest loading on a sweptback wing is carried on the rays originating from the wing apex which lay in the region of the leading edge. The ratio of the value of $C_{l_\\alpha}$ for the control surface to the value of $C_{l_\\alpha}$ for the complete wing indicates that the lift load on the control surface was responsible for about one-fourth of the wing-panel rolling moment.\n\nA part of the additional loading caused by control-surface deflection was carried on the inner panel of the wing (fig. 15), substantiating the carry-over loading indicated in reference 5. The experimental value of $C_{l_\\delta}$ was 0.003 for the control surface and 0.004 for the complete wing. These values compare with calculated values of 0.0036 and 0.0045, respectively. The measured $C_{l_\\delta}$ value of 0.00068 for the control surface was less than the previously mentioned value of 0.00078 for the complete wing with large fuselage.\n\nControl-surface loading.- The results of the initial series of tests in which loads were measured on the control alone (figs. 13 to 15), though limited in scope, indicated that the fence had little effect on the loading of the control surface. The results of the second series of tests (figs. 16 to 18) permit a more detailed analysis of control-surface loading.\n\nWith the control surface undeflected, the value of $C_{N_{f_\\alpha}}$ was about 0.065 (value calculated by linear theory was 0.078). As the control was deflected, $C_{N_{f_\\alpha}}$ decreased in value especially at the highest deflections. The moment coefficient about the hinge line $C_{M_f}$, which corresponds to the control-surface hinge moment, varied nonlinearly with angle of attack as a result of a rearward shift in center of pressure which occurred when the wing was rotated from a streamwise direction. This effect was not defined at high control deflections since the angle-of-attack range did not include zero incidence.\n\nThe hinge-moment coefficient also varied nonlinearly with control deflection for high control deflections. (See fig. 18.) Increasing the angle of attack aggravated this condition at least in the well-defined range of negative deflections and appeared to decrease the linear\n\nCONFIDENTIAL", "timestamp": "2026-07-22T04:25:11.856112+00:00"}
{"citation_id": "19930085551", "source_url": "https://ntrs.nasa.gov/api/citations/19930085551/downloads/19930085551.pdf", "page_number": 79, "total_pages": 82, "image_filename": "19930085551_p79.jpg", "text": "78\nNACA RM No. L8K30\n\n<!-- Image (257, 123, 680, 867) -->\n\n(a) Clean condition; power for level flight.\n\nFigure 23.- Variation of aileron wheel force and aileron effectiveness parameter with change in total aileron angle at various speeds. C-54D airplane.", "timestamp": "2026-07-22T04:25:12.612560+00:00"}
{"citation_id": "19930086060", "source_url": "https://ntrs.nasa.gov/api/citations/19930086060/downloads/19930086060.pdf", "page_number": 17, "total_pages": 20, "image_filename": "19930086060_p17.jpg", "text": "NACA RM L9F02\n15\n\nCONFIDENTIAL\n\n.3\n.2\n.1\n$C_p$ 0\n-.1\n-.2\n\n$K = 60 \\text{ percent}$\n\n[Figure: Graph showing $C_p$ distribution over an airfoil shape with a NACA logo]\n\n(c) Maximum diameter at 60-percent station.\nFigure 5.- Continued.\nCONFIDENTIAL", "timestamp": "2026-07-22T04:25:14.998332+00:00"}
{"citation_id": "19930086105", "source_url": "https://ntrs.nasa.gov/api/citations/19930086105/downloads/19930086105.pdf", "page_number": 8, "total_pages": 22, "image_filename": "19930086105_p8.jpg", "text": "6 CONFIDENTIAL NACA RM E5H12\n\noscillating upstream of the inlet at a frequency of approximately 28 cycles per second. This pronounced change in operating conditions was effected by only a slight increase in the rate of fuel flow (from 76.5 to 78.0 lb/hr).\n\nRegenerative-Type Burner\n\nThis investigation also included experiments on a regenerative-type burner with propylene oxide as the main fuel. High combustion efficiencies and smooth-burning characteristics were evidenced in previous experiments with this burner-fuel combination. Because of a breakdown of the variable-inductance pressure pickup, a balanced diaphragm gage was employed. This gage measured only the maximum and minimum pressures.\n\nThe resultant data are presented in figure 6 with $p_3/p_0$ plotted as a function of $f/a$. The data obtained with the regenerative-type burner show pressure fluctuations again occurring before the optimum mean $p_3/p_0$. Sporadic pulsing in the vicinity of and beyond the optimum mean $p_3/p_0$ prevented accurate measurements of the minimum pressure due to an inability to weight the effect of these intermittent fluctuations. The dashed lines indicate the conditions at which these stray pulses began to increase in frequency and intensity. The maximum static pressures were better defined than the minimum static pressures.\n\nAt the peak indicated by the static tubes, the maximum instantaneous static pressure agreed within experimental accuracy with the mean static pressure at optimum cold recovery. At the same conditions, the minimum static pressure dropped off rapidly, as shown by the dashed lines, but it could not be established whether this minimum value would coincide with the $p_3/p_0$ value obtained without combustion. The general trend, however, was in agreement with that presented in figure 4 for the perforated conical flame holder.\n\nDiffuser Total-Pressure Recovery\n\nThe regenerative-type burner using propylene oxide as the main fuel gave a marked improvement in diffuser total-pressure recovery with combustion as compared with that obtained with the perforated conical flame holder using liquid-fuel injection. The relative diffuser performance with the two burners is shown in figure 7, where the optimum total-pressure recovery $p_3/p_0$ with combustion\n\nCONFIDENTIAL", "timestamp": "2026-07-22T04:25:16.179264+00:00"}
{"citation_id": "19930082090", "source_url": "https://ntrs.nasa.gov/api/citations/19930082090/downloads/19930082090.pdf", "page_number": 113, "total_pages": 122, "image_filename": "19930082090_p113.jpg", "text": "NACA TN No. 1455\n111\n\nVentilating air\n5.0\"\nI.D.\n30\"\nApprox. 10'\n7.83\"\nI.D.\nExhaust gas\n22\"\n26\"\n7.83\"\nI.D.\nExhaust\ngas\n13.5\"\n17.5\"\n40\"\n60\"\n5.0\"\nI.D.\nVentilating air\n\nx Temperature traverse\no Static-pressure tap\n\nNACA\n\nFigure 58.- Schematic diagram of test setup of heat exchanger O and air\nshroud, showing location of static-pressure and temperature measuring\nstations.", "timestamp": "2026-07-22T04:25:17.553044+00:00"}
{"citation_id": "19930085936", "source_url": "https://ntrs.nasa.gov/api/citations/19930085936/downloads/19930085936.pdf", "page_number": 50, "total_pages": 59, "image_filename": "19930085936_p50.jpg", "text": "NACA RM No. E9B03\n49\n\nPressure\ncoefficient\n$C_p$\n.2\n0\n-.2\n-.2\n0\n.2\n\nAngle of\nyaw\n(deg)\n-12\n-6\n0\n\n[Figure: Polar plot showing radial pressure distributions with three curves corresponding to different yaw angles.]\n\nNACA\n\n(a) x/L = 0.148.\nFigure 10. - Radial pressure distributions at $0^\\circ$ angle of attack for three yaw angles.", "timestamp": "2026-07-22T04:25:22.009008+00:00"}
{"citation_id": "19930090382", "source_url": "https://ntrs.nasa.gov/api/citations/19930090382/downloads/19930090382.pdf", "page_number": 6, "total_pages": 37, "image_filename": "19930090382_p6.jpg", "text": "NACA RM L9I07 CONFIDENTIAL 5\n\nTunnel-wall correction.- The force-test data have been corrected for the effect of tunnel-wall constraint on velocity at the propeller test plane by using the method described in reference 1. These results are presented in figure 4 as the ratio of free-air velocity to the tunnel-datum velocity as a function of thrust disk-loading coefficient and tunnel-datum Mach number.\n\nAccuracy of results.- Analysis of the accuracy of the separate measurements required to define fully the propeller characteristics has indicated that errors in the results presented herein are probably less than ±1 percent. Repeat runs have shown that the data can be reproduced to within ±1 percent.\n\nRESULTS AND DISCUSSION\n\nThe basic propeller characteristics are presented in figure 5. For each value of tunnel-datum Mach number M the propeller thrust and power coefficients and efficiency are plotted against advance ratio. The variation of tip Mach number with advance ratio is also included. As used herein, the tunnel-datum Mach number M is not corrected for tunnel-wall constraint. The free-air Mach number, however, can be obtained by applying the tunnel-wall corrections, presented in figure 4, to the tunnel-datum Mach number. At the high Mach numbers the tunnel-wall correction is generally less than 1 percent but, in the exact use of the propeller characteristics presented in figure 5 wherever small changes in Mach number produce large changes in propeller characteristics, the tunnel-datum Mach number should be corrected to free-air Mach number.\n\nAttention is called to the use of the blade angle $\\beta_{0.7R}$ at the 0.7-radius station to designate the pitch setting of the blades for this investigation.\n\nEffect of forward Mach number on maximum efficiency.- The variation of maximum efficiency with forward Mach number is presented in figure 6 for all the blade angles investigated. Results for a blade angle of $70^\\circ$ were obtained only for forward Mach numbers of 0.75, 0.80, and 0.85 because a crack developed in one of the blades and thereby prohibited further investigation of this propeller. In general, the results are similar to those presented in reference 1 for the NACA 4-(5)(08)-03 propeller.\n\nThe maximum delay in the onset of adverse compressibility effects was obtained for a blade angle of $70^\\circ$ ($\\beta_{0.75R} \\approx 68^\\circ$) in comparison to a blade angle ($\\beta_{0.75R}$) of $65^\\circ$ for the NACA 4-(5)(08)-03 propeller. This result occurs primarily because of the difference in design pitch.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T04:25:23.223484+00:00"}
{"citation_id": "19930086003", "source_url": "https://ntrs.nasa.gov/api/citations/19930086003/downloads/19930086003.pdf", "page_number": 27, "total_pages": 36, "image_filename": "19930086003_p27.jpg", "text": "CONFIDENTIAL\n\nDownwash angle, $\\epsilon$, deg\n\nM = 0.60\n\nM = 0.70\n\nM = 0.80\n\nM = 0.85\n\n$\\alpha$, deg -2, -1, 0, 1, 2, 3, 4, 6, 8, 10\n\n$\\square$ $\\diamond$ $\\circ$ $\\triangle$ $\\blacktriangle$ $\\square$ $\\diamond$ $\\nabla$ $\\blacktriangledown$\n\nDownwash angle, $\\epsilon$, deg\n\nM = 0.88\n\nM = 0.90\n\nM = 0.93\n\nM = 0.95\n\nTail-height, $h_t$, percent semispan\n\nCONFIDENTIAL\n\nFigure 10.- Effective downwash angles in region of tail plane for a model with 45° sweptback wing, aspect ratio 6, taper ratio 0.6, and NACA 65A006 airfoil section. Wing alone.\n\nNACA RM L9108\n\nNACA\n\n25", "timestamp": "2026-07-22T04:25:23.321665+00:00"}
{"citation_id": "19930085838", "source_url": "https://ntrs.nasa.gov/api/citations/19930085838/downloads/19930085838.pdf", "page_number": 77, "total_pages": 118, "image_filename": "19930085838_p77.jpg", "text": "NACA RM No. L9B23\n75\n\n<!-- Image (162, 110, 895, 999) -->\n\nSection angle of attack, $\\alpha_o$, deg\n(1) $\\delta_f = 40^\\circ$.\nFigure 9.- Concluded.", "timestamp": "2026-07-22T04:25:24.922485+00:00"}
{"citation_id": "19930085952", "source_url": "https://ntrs.nasa.gov/api/citations/19930085952/downloads/19930085952.pdf", "page_number": 47, "total_pages": 50, "image_filename": "19930085952_p47.jpg", "text": "46\nNACA RM L9C24\n\n<!-- Image (191, 116, 676, 783) -->\n\n(a) $\\alpha \\approx 41^\\circ$; $\\frac{V}{nD} = 0.40$.\n\nFigure 22.— Variation of $C_m$, $C_L$, $C_{h_f}$, and $C_{h_a}$ with $\\delta_f$ for conditions of $C_{D_R} = 0$. Articulated propellers; $\\beta = 11.5^\\circ$; $\\delta_a = -48^\\circ$.", "timestamp": "2026-07-22T04:25:25.750679+00:00"}
{"citation_id": "19930085958", "source_url": "https://ntrs.nasa.gov/api/citations/19930085958/downloads/19930085958.pdf", "page_number": 39, "total_pages": 43, "image_filename": "19930085958_p39.jpg", "text": "```markdown\n1.2\n.8\n.4\n$q_1/q$\n\n24\n16\n8\n0\n-8\n$\\epsilon, deg$\n\n38\n\n.12\n.08\n.04\n0\n-.04\n-.08\n-.12\n-.16\n-.20\n-.24\n$C_m$\n\n1.0\n.8\n.6\n.4\n.2\n0\n-.2\n-.4\n$C_L$\n\nTail height\n(percent $\\frac{b}{2}$)\n46.6\n33.9\n21.1\n-1.1\n\nTail height\n(percent $\\frac{b}{2}$) $l_t$\noff\n$\\circ$ 46.6 -1.1\n$\\square$ 33.9 -1.4\n$\\diamond$ 21.1 -1.2\n$\\triangle$ -1.1 -2.1\n$\\nabla$\n\n-4 0 4 8 12 16 20 24\n$\\alpha, deg$\n\n-4 0 4 8 12 16 20 24 28\n$\\alpha, deg$\n\nNACA\n\nFigure 19.- Characteristics of a $42^\\circ$ sweptback wing-fuselage combination with a horizontal tail.\nFlaps off; low wing.\n\nNACA RM No. 19B11\n```", "timestamp": "2026-07-22T04:25:27.967484+00:00"}