PDF/ocr_dataset_1010.jsonl

{"citation_id": "19930086078", "source_url": "https://ntrs.nasa.gov/api/citations/19930086078/downloads/19930086078.pdf", "page_number": 22, "total_pages": 42, "image_filename": "19930086078_p22.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T04:29:30.688311+00:00"}
{"citation_id": "19930085930", "source_url": "https://ntrs.nasa.gov/api/citations/19930085930/downloads/19930085930.pdf", "page_number": 70, "total_pages": 92, "image_filename": "19930085930_p70.jpg", "text": "CONFIDENTIAL\n\nNACA RM L9907\n\n[Figure: A shadowgraph image showing a curved aerodynamic passage with flow visualization. On the left, a cylindrical structure with perforated holes is visible. The curved duct extends to the right, terminating in a series of vertical fins or vanes at the exit. The background is dark, indicating schlieren or shadowgraph imaging technique.]\n\nFigure 30.— A shadowgraph of the flow in the passage at an area ratio of 1.175 for model 2.\n\nCONFIDENTIAL\n\nNACA\n\n69", "timestamp": "2026-07-22T04:29:31.586775+00:00"}
{"citation_id": "19930092013", "source_url": "https://ntrs.nasa.gov/api/citations/19930092013/downloads/19930092013.pdf", "page_number": 2, "total_pages": 21, "image_filename": "19930092013_p2.jpg", "text": "# AERONAUTIC SYMBOLS\n\n## 1. FUNDAMENTAL AND DERIVED UNITS\n\n| Symbol | Metric | | English | |\n| :--- | :--- | :--- | :--- | :--- |\n| | **Unit** | **Abbreviation** | **Unit** | **Abbreviation** |\n| Length....... | $l$ | meter.................. | m | foot (or mile)........... | ft (or mi) |\n| Time......... | $t$ | second................. | s | second (or hour)......... | sec (or hr) |\n| Force........ | $F$ | weight of 1 kilogram... | kg | weight of 1 pound........ | lb |\n| Power........ | $P$ | horsepower (metric).... | | horsepower............... | hp |\n| Speed........ | $V$ | kilometers per hour.... | kph | miles per hour........... | mph |\n| | | meters per second...... | mps | feet per second.......... | fps |\n\n## 2. GENERAL SYMBOLS\n\n$W$ Weight=$mg$\n$g$ Standard acceleration of gravity=9.80665 m/s$^2$ or 32.1740 ft/sec$^2$\n$m$ Mass=$\\frac{W}{g}$\n$I$ Moment of inertia=$mk^2$. (Indicate axis of radius of gyration $k$ by proper subscript.)\n$\\mu$ Coefficient of viscosity\n$\\nu$ Kinematic viscosity\n$\\rho$ Density (mass per unit volume)\nStandard density of dry air, 0.12497 kg-m$^{-4}$s$^2$ at 15° C and 760 mm; or 0.002378 lb-ft$^{-4}$ sec$^2$\nSpecific weight of \"standard\" air, 1.2255 kg/m$^3$ or 0.07651 lb/cu ft\n\n## 3. AERODYNAMIC SYMBOLS\n\n$S$ Area\n$S_w$ Area of wing\n$G$ Gap\n$b$ Span\n$c$ Chord\n$A$ Aspect ratio, $\\frac{b^2}{S}$\n$V$ True air speed\n$q$ Dynamic pressure, $\\frac{1}{2}\\rho V^2$\n$L$ Lift, absolute coefficient $C_L=\\frac{L}{qS}$\n$D$ Drag, absolute coefficient $C_D=\\frac{D}{qS}$\n$D_0$ Profile drag, absolute coefficient $C_{D_0}=\\frac{D_0}{qS}$\n$D_i$ Induced drag, absolute coefficient $C_{D_i}=\\frac{D_i}{qS}$\n$D_p$ Parasite drag, absolute coefficient $C_{D_p}=\\frac{D_p}{qS}$\n$C$ Cross-wind force, absolute coefficient $C_c=\\frac{C}{qS}$\n$i_w$ Angle of setting of wings (relative to thrust line)\n$i_t$ Angle of stabilizer setting (relative to thrust line)\n$Q$ Resultant moment\n$\\Omega$ Resultant angular velocity\n$R$ Reynolds number, $\\rho \\frac{Vl}{\\mu}$ where $l$ is a linear dimension (e.g., for an airfoil of 1.0 ft chord, 100 mph, standard pressure at 15° C, the corresponding Reynolds number is 935,400; or for an airfoil of 1.0 m chord, 100 mps, the corresponding Reynolds number is 6,865,000)\n$\\alpha$ Angle of attack\n$\\epsilon$ Angle of downwash\n$\\alpha_0$ Angle of attack, infinite aspect ratio\n$\\alpha_i$ Angle of attack, induced\n$\\alpha_a$ Angle of attack, absolute (measured from zero-lift position)\n$\\gamma$ Flight-path angle", "timestamp": "2026-07-22T04:29:32.989484+00:00"}
{"citation_id": "19930086105", "source_url": "https://ntrs.nasa.gov/api/citations/19930086105/downloads/19930086105.pdf", "page_number": 14, "total_pages": 22, "image_filename": "19930086105_p14.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T04:29:33.541880+00:00"}
{"citation_id": "19930082450", "source_url": "https://ntrs.nasa.gov/api/citations/19930082450/downloads/19930082450.pdf", "page_number": 1, "total_pages": 37, "image_filename": "19930082450_p1.jpg", "text": "V3,N21/5.6/1778\n34134\nGOVT. DOC.\n\nNACA TN No. 1778\n\n# NATIONAL ADVISORY COMMITTEE\n# FOR AERONAUTICS\n\n## TECHNICAL NOTE\nNo. 1778\n\nDIRECT-READING DESIGN CHARTS FOR 24S-T ALUMINUM-ALLOY\nFLAT COMPRESSION PANELS HAVING LONGITUDINAL\nFORMED Z-SECTION STIFFENERS\n\nBy Norris F. Dow and Albert S. Keevil, Jr.\n\nLangley Aeronautical Laboratory\nLangley Field, Va.\n\n[Figure: NACA logo]\n\nWashington\nJanuary 1949\n\nCONN. STATE LIBRARY\n\nJAN 20 1949\n\nBUSINESS, SCIENCE\n& TECHNOLOGY DEPT.", "timestamp": "2026-07-22T04:29:36.519933+00:00"}
{"citation_id": "19930082447", "source_url": "https://ntrs.nasa.gov/api/citations/19930082447/downloads/19930082447.pdf", "page_number": 2, "total_pages": 24, "image_filename": "19930082447_p2.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T04:29:37.807580+00:00"}
{"citation_id": "19930082090", "source_url": "https://ntrs.nasa.gov/api/citations/19930082090/downloads/19930082090.pdf", "page_number": 120, "total_pages": 122, "image_filename": "19930082090_p120.jpg", "text": "118\n\n38\"\n4.5\"\n15\"\n3\"\nVentilating\nair\n2\"\n24.25\"\n7.83\"\nExhaust gas\nExhaust gas\n7.83\"\nI.D.\nI.D.\nVentilating air\n3.5\"\n27\"\n5.0\"\nI.D.\nNACA\n\no Static-pressure tap\nx Temperature traverse\n\nFigure 63.- Schematic diagram of test setup of heat exchanger P and air shroud, showing location of static-pressure and temperature measuring stations.\n\nNACA TN No. 1455", "timestamp": "2026-07-22T04:29:39.425623+00:00"}
{"citation_id": "19930093773", "source_url": "https://ntrs.nasa.gov/api/citations/19930093773/downloads/19930093773.pdf", "page_number": 3, "total_pages": 47, "image_filename": "19930093773_p3.jpg", "text": "2\nNACA RM E9G09\n\nAt an altitude of 25,000 feet, the internal drag of a wind-milling engine varied from 2 percent of the available net thrust at a true airspeed of 200 miles per hour to 15 percent at a true airspeed of 650 miles per hour.\n\nINTRODUCTION\n\nAn investigation has been conducted in the NACA Lewis altitude wind tunnel to determine the over-all performance, component performance, and operational characteristics of a J47 turbojet engine over a wide range of simulated flight conditions.\n\nData are presented in graphical form to show the engine performance over a range of altitudes from 5000 to 50,000 feet and flight Mach numbers from 0.21 to 0.97. The effect of altitude is shown at a flight Mach number of 0.21 and the effect of flight Mach number is shown at an altitude of 25,000 feet. Performance data are generalized by two methods to determine the range of flight conditions for which engine performance may be predicted from performance data obtained at a given flight condition. Curves are presented to show the windmilling characteristics of the engine. All engine performance data obtained in the investigation are also presented in tabular form.\n\nDESCRIPTION OF ENGINE\n\nThe J47 turbojet engine used in the altitude-wind-tunnel investigation (fig. 1) has a sea-level static thrust rating of 5000 pounds at an engine speed of 7900 rpm and a turbine-outlet gas temperature of $1275^\\circ$ F. At this rating the air flow is approximately 94 pounds per second. The engine has a 12-stage axial-flow compressor with a pressure ratio of approximately 5.1 at rated engine speed, eight cylindrical direct-flow-type combustion chambers, a single-stage impulse turbine, and a fixed-area exhaust nozzle. The exhaust nozzle, which was used in this investigation and was designated standard, had an outlet area of 280 square inches. This exhaust nozzle produced a turbine-outlet temperature of approximately $1275^\\circ$ F at a flight Mach number of 0.21, an altitude of 5000 feet, and an engine speed of 7900 rpm. The over-all length of the engine excluding the exhaust nozzle is 143 inches, the maximum diameter is approximately 37 inches, and the total weight is 2475 pounds.", "timestamp": "2026-07-22T04:29:40.345616+00:00"}
{"citation_id": "19930086003", "source_url": "https://ntrs.nasa.gov/api/citations/19930086003/downloads/19930086003.pdf", "page_number": 33, "total_pages": 36, "image_filename": "19930086003_p33.jpg", "text": "```markdown\nCONFIDENTIAL\n\nM = 0.90\n$\\alpha = 10^\\circ$\nM = 0.95\n$\\alpha = 10^\\circ$\nM = 1.00\n\n$\\frac{q_{wake}}{q}$\n1.2\n.8\n\nWing alone\nWing fuselage\n\n$\\alpha = 4^\\circ$\n$\\alpha = 4^\\circ$\n\n$\\frac{q_{wake}}{q}$\n1.2\n.8\n\n$\\alpha = 0^\\circ$\n$\\alpha = 0^\\circ$\n\n$\\frac{q_{wake}}{q}$\n1.2\n.8\n\n- 80 - 40 0 40 80\n- 80 - 40 0 40 80\n- 80 - 40 0 40 80\n\nTail-height, $h_t$, percent semispan\nCONFIDENTIAL\n\nFigure 13.- Continued.\n\nNACA RM L9J08\n31\n```", "timestamp": "2026-07-22T04:29:41.975349+00:00"}
{"citation_id": "19930085842", "source_url": "https://ntrs.nasa.gov/api/citations/19930085842/downloads/19930085842.pdf", "page_number": 67, "total_pages": 104, "image_filename": "19930085842_p67.jpg", "text": "```markdown\nNACA RM L5029\n\n$\\alpha, deg$ $\\delta_a, deg$\n35.2 -48\n29.3 -48\n23.2 -33\n11.3 -15\n0.6 0\n\nDrag coefficient, $C_D$\n5\n4\n3\n2\n1\n0\n\n$\\cong 7mm$\n\n-16 -8 0 8 16 24 32\nFlap deflection, $\\delta_F, deg$\n\nNATIONAL ADVISORY\nCOMMITTEE FOR AERONAUTICS\n\nFigure 34.- Variation of drag coefficient with flap deflection. Model in basic configuration;\n$\\delta_{a_{T_L}} = \\delta_{a_{T_R}} = 0^\\circ$; $\\delta_r = 0^\\circ$; propellers removed.\n\n63\n```", "timestamp": "2026-07-22T04:29:46.206887+00:00"}
{"citation_id": "19930082245", "source_url": "https://ntrs.nasa.gov/api/citations/19930082245/downloads/19930082245.pdf", "page_number": 3, "total_pages": 66, "image_filename": "19930082245_p3.jpg", "text": "```markdown\n2\nNACA TN No. 1596\n\nanalysis and publication of the results of these tests were deferred owing to the exigencies of other problems more closely connected with the war effort. The results are being published at this time to add to existing information on the effects of compressibility on wings with various types of control surfaces.\n\nThe specific purpose of the present tests was to determine the high-speed characteristics of an NACA 66,1-115 low-drag airfoil section equipped with 20-percent-chord plain ailerons. A true-contour aileron having ordinates the same as those of the rear part of the NACA 66,1-115 airfoil section was one of the ailerons tested. The results of reference 4 had shown, both theoretically and experimentally at low speeds, that the hinge moments of a control surface could be reduced by thickening and beveling the trailing edge. Additional tests were made, therefore, to determine the effects of compressibility on an NACA 66,1-115 airfoil section equipped with a beveled-trailing-edge aileron.\n\nSection characteristics were determined from complete pressure distributions over the main portion of the airfoil and the aileron. The tests were made for Mach numbers up to 0.75 and included various wing angles of attack and aileron deflections.\n\nSYMBOLS\n\nThe term \"main portion of the airfoil\" is used herein to mean that part of the airfoil excluding the aileron. The aerodynamic coefficients and other symbols used in this paper are as follows:\n\n| | |\n| :--- | :--- |\n| $a$ | speed of sound in undisturbed stream. |\n| $c$ | section chord of airfoil with aileron neutral (fig. 1)<br>(2.000 ft on model) |\n| $c_M$ | chord of main portion of airfoil |\n| $c_a$ | section chord of aileron measured along chord from hinge axis of aileron to trailing edge of aileron (0.400 ft on model) |\n| $c_h$ | section hinge-moment coefficient of aileron about hinge axis determined from pressure-distribution data; component due to aileron chord force neglected $$ \\left(\\frac{1}{c_a}\\right)^2 \\int_0^{c_a} (P_U - P_L)x \\, dx $$ |\n```", "timestamp": "2026-07-22T04:29:46.738906+00:00"}
{"citation_id": "19930090382", "source_url": "https://ntrs.nasa.gov/api/citations/19930090382/downloads/19930090382.pdf", "page_number": 12, "total_pages": 37, "image_filename": "19930090382_p12.jpg", "text": "NACA RM L9I07\n13\n\nCONFIDENTIAL\n\n[Figure: Two views of a propeller blade, one showing the broad face and one showing the thin edge. A ruler is placed at the base for scale. A label in the bottom left corner reads \"NACA L-61571.1\".]\n\nFigure 3.- NACA 4-(4)(06)-04 propeller.\nCONFIDENTIAL", "timestamp": "2026-07-22T04:29:49.918649+00:00"}
{"citation_id": "19930085529", "source_url": "https://ntrs.nasa.gov/api/citations/19930085529/downloads/19930085529.pdf", "page_number": 83, "total_pages": 85, "image_filename": "19930085529_p83.jpg", "text": "82\nNACA RM No. L8A30a\n\nTABLE 76\n$$\n[A = -45^\\circ, \\theta_{FL} = 9.8^\\circ, \\alpha = -4^\\circ]\n$$\n\n**CONFIDENTIAL**\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** | **0.905** | **0.96** | | | **0.60** | **0.80** | **0.89** | **0.905** | **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 | +0.042 | +0.038 | -0.031 | -0.063 | -0.056 |\n| 7 | 59.0 | -0.096 | -0.101 | -0.108 | -0.090 | -0.231 | 92 | 72.5 | -0.013 | -0.008 | .003 | .000 | -0.019 |\n| 8 | 67.5 | -.062 | -.064 | -.065 | -.057 | -.107 | 93 | 84.0 | -- | -- | -- | -- | -- |\n| 9 | 77.5 | -- | -- | -- | -- | -- | 94 | 94.0 | -- | -- | -- | -- | -- |\n| 10 | 87.5 | -- | -- | -- | -- | -- | | | | | | | |\n| 11 | 96.0 | -- | -- | -- | -- | -- | | | | | | | |\n| | | | | | | | 95 | 3.0 | -- | -- | -- | -- | -- |\n| B12 | 2.0 | -- | -- | -- | -- | -- | 96 | 10.0 | -- | -- | -- | -- | -- |\n| 13 | 6.0 | -- | -- | -- | -- | -- | 97 | 25.0 | -.160 | -.137 | -.170 | -.268 | -.366 |\n| 14 | 15.0 | -- | -- | -- | -- | -- | 98 | 41.0 | -.150 | -.153 | -.136 | -.118 | -.349 |\n| 15 | 27.5 | -.160 | -.203 | -.261 | -.341 | -.343 | 99 | 52.5 | -.114 | -.119 | -.110 | -.080 | -.266 |\n| 16 | 40.0 | -.176 | -.210 | -.255 | -.322 | -.404 | 100 | 62.5 | -.080 | -.083 | -.079 | -.051 | -.160 |\n| 17 | 50.0 | -.169 | -.197 | -.228 | -.304 | -.416 | 101 | 72.5 | -.024 | -.025 | -.023 | -.017 | -.077 |\n| 18 | 59.0 | -.149 | -.166 | -.189 | -.213 | -.340 | 102 | 84.3 | .017 | .019 | .021 | .029 | .069 |\n| 19 | 67.5 | -.098 | -.118 | -.124 | -.101 | -.335 | 103 | 94.5 | .057 | .058 | .059 | .063 | .007 |\n| 20 | 77.5 | -.052 | -.064 | -.066 | -.050 | -.236 | | | | | | | |\n| 21 | 88.0 | .001 | .016 | .016 | .021 | -.076 | | | | | | | |\n| 22 | 95.5 | -- | -- | -- | -- | -- | | | | | | | |\n| | | | | | | | 104 | 3.0 | -.535 | -.694 | -.647 | -.849 | -.639 |\n| C23 | 2.0 | .311 | .336 | .345 | .356 | .368 | 105 | 10.0 | -.336 | +.436 | +.451 | +.453 | +.453 |\n| 24 | 6.0 | .095 | .098 | .106 | .118 | .130 | 106 | 25.0 | -.302 | -.436 | -.406 | -.407 | |\n| 25 | 15.0 | -.065 | -.071 | -.083 | -.074 | -.080 | 107 | 41.0 | -- | -- | -- | -- | -- |\n| 26 | 27.8 | -.145 | -.179 | -.225 | -.234 | -.212 | 108 | 52.5 | -- | -- | -- | -- | -- |\n| 27 | 40.0 | -.180 | -.223 | -.289 | -.323 | -.296 | 109 | 62.5 | -.113 | -.128 | -.131 | -.282 | -.386 |\n| 28 | 50.0 | -.152 | -.205 | -.267 | -.322 | -.358 | 110 | 72.5 | -.045 | -.056 | -.058 | -.098 | -.320 |\n| 29 | 59.0 | -.156 | -.187 | -.212 | -.223 | -.319 | 111 | 85.1 | .000 | -.006 | -.011 | -.008 | -.221 |\n| 30 | 67.5 | -.108 | -.129 | -.160 | -.240 | -.287 | 112 | 94.6 | .049 | .019 | .030 | .026 | -.135 |\n| 31 | 77.5 | -.061 | -.081 | -.106 | -.150 | -.258 | | | | | | | |\n| 32 | 88.0 | .007 | .012 | -.009 | -.041 | -.160 | | | | | | | |\n| 33 | 95.5 | -- | -- | -- | -- | -- | | | | | | | |\n| | | | | | | | 113 | 3.0 | -.278 | -.449 | -.395 | -.522 | -.541 |\n| D34 | 2.0 | .258 | .250 | .254 | .259 | .250 | 114 | 10.0 | -.251 | -.301 | -.327 | -.340 | -.371 |\n| 35 | 15.0 | -.064 | -.068 | -.077 | -.081 | -.093 | 115 | 25.0 | -.208 | -.208 | -.238 | -.279 | -.279 |\n| 36 | 27.5 | -.127 | -.146 | -.171 | -.171 | -.201 | 116 | 41.0 | -.179 | -.221 | -.254 | -.259 | -.261 |\n| 37 | 40.0 | -.173 | -.200 | -.228 | -.236 | -.238 | 117 | 52.5 | -.130 | -.130 | -.130 | -.130 | -.130 |\n| 38 | 50.0 | -.118 | -.165 | -.199 | -.232 | -.281 | 118 | 62.5 | -- | -- | -- | -- | -- |\n| 39 | 59.0 | -.119 | -.165 | -.199 | -.232 | -.281 | 119 | 72.5 | -.080 | -.089 | -.106 | -.126 | -.157 |\n| 40 | 67.5 | -- | -- | -- | -- | -- | 120 | 87.4 | .020 | .004 | -.024 | -.084 | -.035 |\n| 41 | 77.5 | -.097 | -.083 | -.103 | -.126 | -.115 | 121 | 94.2 | .027 | .014 | -.012 | -.095 | -.041 |\n| 42 | 87.5 | -.007 | -.003 | -.001 | .077 | .077 | | | | | | | |\n| 43 | 94.0 | .031 | .016 | -.006 | -.036 | -.036 | | | | | | | |\n| | | | | | | | 122 | 3.0 | -- | -.293 | -.319 | -.336 | -.370 |\n| E44 | 2.0 | .212 | .217 | .214 | .211 | | 123 | 10.0 | -- | -.204 | -.243 | -.266 | -.237 |\n| 45 | 6.0 | .035 | .034 | .024 | .024 | | 124 | 25.0 | -- | -.227 | -.227 | -.229 | -.232 |\n| 46 | 15.0 | -.096 | -.115 | -.121 | -.134 | | 125 | 41.0 | -- | -.197 | -.197 | -.197 | -.197 |\n| 47 | 27.5 | -.159 | -.188 | -.219 | -.219 | | 126 | 52.5 | -- | -.134 | -.152 | -.148 | -.138 |\n| 48 | 40.0 | -.210 | -.249 | -.269 | -.300 | | 127 | 62.5 | -- | -.114 | -.114 | -.114 | -.114 |\n| 49 | 50.0 | -.227 | -.265 | -.296 | -.296 | | 128 | 72.5 | -- | -.083 | -.113 | -.140 | -.129 |\n| 50 | 59.0 | -.180 | -.211 | -.229 | -.274 | | 129 | 78.0 | -- | -.064 | -.066 | -.078 | -.123 |\n| 51 | 67.5 | -.135 | -.154 | -.153 | -.189 | | 130 | 85.3 | -- | -.004 | -.020 | -.025 | -.041 |\n| 52 | 77.5 | -.076 | -.092 | -.092 | -.055 | | 131 | 94.1 | -- | .010 | .007 | .008 | .004 |\n| 53 | 88.5 | .051 | .036 | .035 | .025 | | | | | | | | |\n| 54 | 95.5 | .011 | .008 | .008 | .002 | | | | | | | | |\n| | | | | | | | 132 | 3.0 | -- | -- | -- | -- | -- |\n| F55 | 2.0 | .086 | .080 | .089 | -- | -- | 133 | 10.0 | -- | -.126 | -.171 | -.241 | -.233 |\n| 56 | 6.0 | -.066 | -.080 | -.089 | -.114 | | 134 | 25.0 | -- | -.181 | -.202 | -.207 | -.211 |\n| 57 | 15.0 | -.085 | -.097 | -.123 | -.114 | | 135 | 41.0 | -- | -.113 | -.167 | -.128 | -.187 |\n| 58 | 27.5 | -.161 | -.188 | -.224 | -.222 | | 136 | 52.5 | -- | -.113 | -.128 | -.128 | -.141 |\n| 59 | 40.0 | -.213 | -.245 | -.276 | -.309 | | 137 | 62.5 | -- | -.033 | -.039 | -.031 | -.023 |\n| 60 | 50.0 | -.223 | -.265 | -.296 | -.280 | | 138 | 72.5 | -- | .016 | .016 | .040 | .043 |\n| 61 | 59.0 | -.197 | -.218 | -.245 | -.263 | | 139 | 83.4 | -- | .056 | .059 | .073 | .113 |\n| 62 | 67.5 | -.120 | -.110 | -.122 | -.122 | | 140 | 94.0 | -- | .033 | .033 | .033 | .034 |\n| 63 | 77.5 | -.080 | -.0", "timestamp": "2026-07-22T04:29:52.371205+00:00"}
{"citation_id": "19930086015", "source_url": "https://ntrs.nasa.gov/api/citations/19930086015/downloads/19930086015.pdf", "page_number": 35, "total_pages": 54, "image_filename": "19930086015_p35.jpg", "text": "34\n\nCONFIDENTIAL\n\nD\nD\n○ - 175.25\n□ - 165.12\n◇ - 147.20\n△ - 129.32\n△ - 113.43\n□ - 100.47\n○ - 88.52\n\nMach number, M\n1.8\n1.7\n1.6\n1.5\n1.4\n1.3\n1.2\n1.1\n1.0\n\n-36 -32 -28 -24 -20 -16 -12 -8 -4 0 4 8 12 16 20 24 28\nHorizontal distance from window center line, x, in\n\n[Figure: Graph showing the variation of Mach number along the horizontal center line of the Ames 6- by 6-foot supersonic wind tunnel. Stagnation pressure = 9 lb/sq in. abs.]\n\nFigure 9.- The variation of Mach number along the horizontal center line of the Ames 6- by 6-foot supersonic wind tunnel. Stagnation pressure = 9 lb/sq in. abs.\n\nNACA\n\nCONFIDENTIAL\n\nNACA RM A57E24", "timestamp": "2026-07-22T04:29:54.406244+00:00"}
{"citation_id": "19930086151", "source_url": "https://ntrs.nasa.gov/api/citations/19930086151/downloads/19930086151.pdf", "page_number": 13, "total_pages": 34, "image_filename": "19930086151_p13.jpg", "text": "NACA RM L9J28 CONFIDENTIAL 11\n\nwing with the triangular aileron accounts for the larger values of $pb/2V$ usually obtained with the parallelogram ailerons. In general, because of the differences in the variation of the values of $C_{l_p}$ with $\\alpha$ for the plain wing and the wing with end plate, the rolling effectiveness of the ailerons on the wing with end plate exhibited large increases with increase in $\\alpha$ as contrasted to the smaller increases in rolling effectiveness with increase in $\\alpha$ (up to $\\alpha = 10^\\circ$) exhibited by the ailerons on the plain wing. As a result of these trends, the ailerons on the wing with end plate produced larger values of $pb/2V$ at large values of $\\alpha$ than did the ailerons on the plain wing; however, if the true variation of $C_{l_p}$ with $\\alpha$ for the wing with end plate were known, the results may differ somewhat from those shown by the present data. The data of figures 14 to 16 also show that the aileron differential generally had a negligible effect on the rolling performance of any wing-aileron configuration, except possibly at very large angles of attack, for which an increased rolling effectiveness is usually exhibited by employing the 2:1 or 3:1 differential as compared with the 1:1 differential.\n\nAs previously discussed, the effects of adverse aileron yaw on the estimated rolling-performance characteristics shown in figures 14 to 16 have not been considered in the calculations. These adverse yawing moments would tend to reduce the rolling effectiveness of the ailerons by inducing sideslip – particularly in the high-lift-coefficient range. In some instances, a sizeable deflection of the rudder may be required to perform a coordinated roll. It is well to note, however, that these adverse yawing moments are comparable to those produced by conventional flap-type ailerons (reference 17).\n\nCONCLUSIONS\n\nAn investigation of triangular- and parallelogram-plan-form deflectable wing-tip ailerons on an untapered $45^\\circ$ sweptback semispan wing with and without an end plate (simulating a vertical fin) was performed in the Langley 300 MPH 7- by 10-foot tunnel. The rectangular end plate was mounted on the wing just inboard of the ailerons. The results of the investigation led to the following conclusions:\n\n1. Each of the aileron configurations investigated should provide adequate lateral control over the entire angle-of-attack range investigated.\n\n2. The yawing moments resulting from aileron deflection were generally adverse – particularly at large angles of attack and aileron deflections.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T04:29:56.562572+00:00"}
{"citation_id": "19930085965", "source_url": "https://ntrs.nasa.gov/api/citations/19930085965/downloads/19930085965.pdf", "page_number": 39, "total_pages": 67, "image_filename": "19930085965_p39.jpg", "text": "38\nNACA RM ESE06\n1185\n\nHeat required. - The heat dissipation from a surface in conditions of icing may be divided into its four components:\n\n(1) Dissipation of heat due to convection from the surface to the air\n\n(2) Dissipation of heat by evaporation of water from surface because of difference in vapor pressure\n\n(3) Dissipation of heat as a result of raising temperature of impinging water\n\n(4) Dissipation of heat due to radiation from relatively warmer surface\n\nFor this analysis, the convex surface of the blade was assumed to dissipate heat only by convection and evaporation, because very little, if any, water is assumed to impinge upon the convex surface of the blade. Heat losses due to radiation were small and were neglected.\n\nDissipation of heat due to convection and evaporation. - The heat dissipated as a result of convection and evaporation was computed by the following simplified equation, as derived from reference 10:\n\n$$H_1 + H_c = h_x \\left[ (t_s - t_{sx}) + \\frac{0.622 L}{c_p} \\left( \\frac{e_s - e_x}{P_x} \\right) \\right] \\quad \\text{(B10)}$$\n\nWater impingement. - The water impingement per unit area of leading edge and vane surface was computed assuming a collection efficiency of 100 percent. The areas affected were evaluated from the geometry of the guide vane with straight-line impingement assumed on all exposed surfaces.\n\nThe average water impingement per unit area of leading edge was calculated as follows:\n\n$$M_e = \\frac{2}{\\pi} V_c m (3600) \\quad \\text{(B11)}$$\n\nThe water impingement per unit area of vane surface was calculated by the following equation, which represents the concave surface as a plane surface parallel to the chord line:", "timestamp": "2026-07-22T04:30:01.655644+00:00"}
{"citation_id": "19930085930", "source_url": "https://ntrs.nasa.gov/api/citations/19930085930/downloads/19930085930.pdf", "page_number": 71, "total_pages": 92, "image_filename": "19930085930_p71.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T04:30:02.307330+00:00"}
{"citation_id": "19930082476", "source_url": "https://ntrs.nasa.gov/api/citations/19930082476/downloads/19930082476.pdf", "page_number": 1, "total_pages": 41, "image_filename": "19930082476_p1.jpg", "text": "Y3,N21/5: c/1801\nGOVT. DOC.\n\nNACA TN No. 1801\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\n\nTECHNICAL NOTE\nNo. 1801\n\nWIND-TUNNEL INVESTIGATION OF THE SPINNING CHARACTERISTICS\nOF A MODEL OF A TWIN-TAIL LOW-WING PERSONAL-OWNER-TYPE\nAIRPLANE WITH LINKED AND UNLINKED RUDDER AND\nAILERON CONTROLS\n\nBy Walter J. Klinar and Lawrence J. Gale\n\nLangley Aeronautical Laboratory\nLangley Field, Va.\n\n[Figure: NACA logo]\n\nWashington\nJanuary 1949\n\n[Stamp: PUBLIC DOCUMENTS DEPT. CONN. STATE LIBRARY]\nFEB 4 1949\n\nBUSINESS, SCIENCE\n& TECHNOLOGY DEP'T.", "timestamp": "2026-07-22T04:30:07.404802+00:00"}
{"citation_id": "19930086081", "source_url": "https://ntrs.nasa.gov/api/citations/19930086081/downloads/19930086081.pdf", "page_number": 20, "total_pages": 44, "image_filename": "19930086081_p20.jpg", "text": "```markdown\n18\n\nCONFIDENTIAL\n0.024 i.e. radius at\nroot — tapers to\n0 at tip.\n\n<!-- Image (66, 79, 880, 772) -->\n\n(b) Fences.\n\n(c) Control surface; mean aerodynamic\nchord, 1.83; span, 1.57.\n\nFigure 3.- Concluded.\n\nNACA RM L9H05\n```", "timestamp": "2026-07-22T04:30:14.572613+00:00"}
{"citation_id": "19930086078", "source_url": "https://ntrs.nasa.gov/api/citations/19930086078/downloads/19930086078.pdf", "page_number": 23, "total_pages": 42, "image_filename": "19930086078_p23.jpg", "text": "NACA RM L9H04\n21\n\nCONFIDENTIAL\n\nPitching-moment coefficient, $C_m$\n1\n0\n-1\n\nM\no 0.19\n□ .27\n△ .37\n\nDrag coefficient, $C_D$\n3\n2\n1\n0\n\nAngle of attack, $\\alpha$, deg\n20\n16\n12\n8\n4\n0\n-4\n-8\n\n-4 -2 0 2 4 6 8\nLift coefficient, $C_L$\n\nCONFIDENTIAL\nNACA\n\nFigure 4.- Plain-wing aerodynamic characteristics of the unswept wing configuration for several Mach numbers.", "timestamp": "2026-07-22T04:30:18.097438+00:00"}
{"citation_id": "19930085838", "source_url": "https://ntrs.nasa.gov/api/citations/19930085838/downloads/19930085838.pdf", "page_number": 84, "total_pages": 118, "image_filename": "19930085838_p84.jpg", "text": "82\nNACA RM No. L9B23\n\n[Figure: A graph plotting Aileron section hinge-moment coefficient, $C_{h_a}$, against Section angle of attack, $\\alpha_o$, deg. The graph contains multiple curves with data points. There are two legends within the graph area. One legend in the upper right quadrant lists $\\delta_a$ (deg) and $\\delta_t$ (deg) with values 0, -10 and -5, -10. Another legend in the lower right quadrant lists $\\delta_t = -15^\\circ$ and $\\delta_a$ (deg) with values -10, -5, 0. The NACA logo is present in the lower right corner of the graph.]\n\n(g) $\\delta_f = 40^\\circ$.\nFigure 10.- Continued.", "timestamp": "2026-07-22T04:30:18.711466+00:00"}
{"citation_id": "19930082450", "source_url": "https://ntrs.nasa.gov/api/citations/19930082450/downloads/19930082450.pdf", "page_number": 2, "total_pages": 37, "image_filename": "19930082450_p2.jpg", "text": "NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nTECHNICAL NOTE NO. 1778\n\nDIRECT-READING DESIGN CHARTS FOR 24S-T ALUMINUM-ALLOY\nFLAT COMPRESSION PANELS HAVING LONGITUDINAL\nFORMED Z-SECTION STIFFENERS\n\nBy Norris F. Dow and Albert S. Keevil, Jr.\n\nSUMMARY\n\nDirect-reading design charts are presented for 24S-T aluminum-alloy flat compression panels having longitudinal formed Z-section stiffeners. These charts make possible the direct determination of the stress and all the panel proportions required to carry a given intensity of loading with a given skin thickness and effective length of panel.\n\nINTRODUCTION\n\nDesign charts for wing compression panels have been presented in several different forms. (See references 1 and 2.) In reference 3, a form was developed which permitted the direct selection of proportions for given values of the principal design conditions - intensity of loading, skin thickness, and effective length of panel. This form also made possible the ready determination of the proportions having minimum weight to meet these conditions. The charts presented in reference 3 covered 75S-T aluminum-alloy flat compression panels having longitudinal straight-web Y-section stiffeners. Similar charts for 24S-T aluminum-alloy panels with extruded, straight-web Y-section stiffeners are presented in reference 4, and direct-reading design charts for 24S-T aluminum-alloy panels with formed Z-section stiffeners are presented herein.\n\nSYMBOLS\n\nThe symbols used for the panel dimensions are given in figure 1. In addition, the following symbols are used:\n\nc coefficient of end fixity as used in Euler column formula\n\nd rivet diameter, inches\n\nL length of panel, inches", "timestamp": "2026-07-22T04:30:23.424202+00:00"}
{"citation_id": "19930082447", "source_url": "https://ntrs.nasa.gov/api/citations/19930082447/downloads/19930082447.pdf", "page_number": 3, "total_pages": 24, "image_filename": "19930082447_p3.jpg", "text": "NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nTECHNICAL NOTE NO. 1775\n\nHYDRODYNAMIC IMPACT LOADS IN SMOOTH WATER FOR A PRISMATIC FLOAT HAVING AN ANGLE OF DEAD RISE OF $40^\\circ$\n\nBy Philip M. Edge, Jr.\n\nSUMMARY\n\nA prismatic-float forebody with an angle of dead rise of $40^\\circ$ was subjected to smooth-water impacts in the Langley impact basin. The tests were made at fixed trims of $3^\\circ$, $6^\\circ$, $9^\\circ$, and $12^\\circ$ for a range of flight-path angles from approximately $2^\\circ$ to $22^\\circ$.\n\nThe data are presented, and converted into dimensionless variables for correlation of the experimental results with hydrodynamic impact theory and for comparison of the runs among themselves. The average value of the dead-rise function for an angle of dead rise of $40^\\circ$ is evaluated and compared with similar values for angles of dead rise of $30^\\circ$ and $22\\frac{1}{2}^\\circ$ and with the theoretical dead-rise function. The experimental data are shown to be in good agreement with values predicted by theory.\n\nINTRODUCTION\n\nThe development of seaplanes having high aerodynamic performance accompanied by high stalling speeds and high wing loadings has resulted in increased impact loads. The designer of the modern seaplane is confronted with the dual problem of predicting the water loads and of devising means of reducing these loads.\n\nIn order to provide a more rational basis for the prediction of impact loads, reference 1 presented an analysis which showed that the motion and time characteristics of an impact may be represented by means of generalized variables. The variation of the generalized variables is governed solely by the magnitude of the approach parameter $\\kappa$ which may be considered a criterion of impact similarity.\n\nOne possible means of reducing the water loads on seaplanes is the use of sharper angles of dead rise. A program undertaken at the Langley impact basin to determine the variation of impact loads with angle of dead rise has therefore been expanded to include tests of a seaplane float having a $40^\\circ$ angle of dead rise. Data were obtained at fixed trim with a V-bottom prismatic-float forebody of $40^\\circ$ dead-rise", "timestamp": "2026-07-22T04:30:26.114825+00:00"}
{"citation_id": "19930085936", "source_url": "https://ntrs.nasa.gov/api/citations/19930085936/downloads/19930085936.pdf", "page_number": 57, "total_pages": 59, "image_filename": "19930085936_p57.jpg", "text": "56\nNACA RM No. E9B03\n\n<!-- Image (127, 99, 820, 283) -->\n\n(a) Angle of yaw, $-12^\\circ$.\n\n<!-- Image (127, 368, 840, 554) -->\n\n(b) Angle of yaw, $0^\\circ$.\n\n<!-- Image (127, 636, 840, 822) -->\n\n(c) Angle of yaw, $12^\\circ$.\n\nFigure 14. - Pressure coefficients on wedge surface at $0^\\circ$ angle of attack for three angles of yaw.", "timestamp": "2026-07-22T04:30:26.712893+00:00"}
{"citation_id": "19930093773", "source_url": "https://ntrs.nasa.gov/api/citations/19930093773/downloads/19930093773.pdf", "page_number": 4, "total_pages": 47, "image_filename": "19930093773_p4.jpg", "text": "NACA RM E9G09\n\nAir enters the engine through an annular inlet (fig. 2) and passes into the compressor through a row of inlet guide vanes. The air is discharged from the compressor through two rows of straightening vanes. From the annular outlet of the compressor, the air flows into the combustion chambers where it is mixed with fuel injected through duplex fuel nozzles. The mixture is burned and the hot gases of combustion flow through the turbine-inlet stator blades, the turbine, and into the atmosphere through the tail pipe and the exhaust nozzle.\n\nINSTALLATION\n\nThe engine was mounted on a wing in the test section of the altitude wind tunnel (fig. 1). Dry refrigerated air was supplied to the engine from the tunnel make-up air system through a duct connected to the engine inlet. A frictionless slip joint in the duct made possible the measurement of engine thrust and drag by the tunnel balance scales. The air flow through the duct was throttled from approximately sea-level pressure to a total pressure at the engine inlet corresponding to the desired flight Mach number at a given altitude.\n\nInstrumentation for measuring pressures and temperatures was installed at various stations in the engine (fig. 2). Instrumentation for measuring air flow was installed at the inlet-air-duct venturi throat (station r), the engine inlet (station 1), and the exhaust-nozzle outlet (station 7).\n\nPROCEDURE\n\nThrust values were calculated from tunnel balance-scale measurements and also from values of gas flow and jet velocity obtained from measurements with the exhaust-nozzle survey rake. The exhaust-nozzle jet-velocity coefficient, defined as the ratio of scale jet thrust to rake jet thrust, is shown as a function of exhaust-nozzle pressure ratio in figure 3. Engine performance is based on thrust values obtained from the balance scales because this method includes the losses resulting from the inefficiency of the exhaust nozzle.\n\nSymbols and methods of calculation are given in the appendix.\n\nPerformance data were obtained at the following altitudes and flight Mach numbers:", "timestamp": "2026-07-22T04:30:27.880624+00:00"}
{"citation_id": "19930085842", "source_url": "https://ntrs.nasa.gov/api/citations/19930085842/downloads/19930085842.pdf", "page_number": 68, "total_pages": 104, "image_filename": "19930085842_p68.jpg", "text": "64\nNACA RM L9C29\n\n<!-- Image (101, 109, 833, 840) -->\n\nFigure 35.- Variation of $C_{n_a}$, $C_l$, and $C_n$ with right-ailavator deflection on a $\\frac{1}{3}$-scale model of the airplane. Left-ailavator fixed; propellers removed; V = 100 miles per hour; $\\delta_r = 0^\\circ$; basic model configuration.", "timestamp": "2026-07-22T04:30:29.885465+00:00"}
{"citation_id": "19930086003", "source_url": "https://ntrs.nasa.gov/api/citations/19930086003/downloads/19930086003.pdf", "page_number": 34, "total_pages": 36, "image_filename": "19930086003_p34.jpg", "text": "CONFIDENTIAL\n32\n\nM = 1.05\n$\\alpha = 10^\\circ$\nM = 1.10\n$\\alpha = 10^\\circ$\nM = 1.15\n\n$\\frac{q_{wake}}{q}$\n1.2\n.8\n\nWing alone\nWing fuselage\n\n$\\alpha = 4^\\circ$\n$\\alpha = 4^\\circ$\n\n$\\frac{q_{wake}}{q}$\n1.2\n.8\n\n$\\alpha = 0^\\circ$\n$\\alpha = 0^\\circ$\n\n$\\frac{q_{wake}}{q}$\n1.2\n.8\n- 80 - 40 0 40 80\n- 80 - 40 0 40 80\n- 80 - 40 0 40 80\n\nTail-height, $h_t$, percent semispan\nCONFIDENTIAL\n\nFigure 13.— Concluded.\n\nNACA RM L51C08", "timestamp": "2026-07-22T04:30:34.312040+00:00"}
{"citation_id": "19930090382", "source_url": "https://ntrs.nasa.gov/api/citations/19930090382/downloads/19930090382.pdf", "page_number": 13, "total_pages": 37, "image_filename": "19930090382_p13.jpg", "text": "NACA RM L9I07\n15\n\nCONFIDENTIAL\n\nVelocity ratio, $V_0/V$\nThrust coefficient, $\\frac{T_c}{1-M^2}$\n\nFigure 4.- Tunnel-wall-interference correction for 4-foot-diameter propeller in Langley 8-foot high-speed tunnel.", "timestamp": "2026-07-22T04:30:35.312051+00:00"}
{"citation_id": "19930085930", "source_url": "https://ntrs.nasa.gov/api/citations/19930085930/downloads/19930085930.pdf", "page_number": 72, "total_pages": 92, "image_filename": "19930085930_p72.jpg", "text": "~~UNCLASSIFIED~~\nCONFIDENTIAL\n\nConcave surface: high static pressure;\nlow velocity\n\nNACA RM L9G07\n\nNegative\npressure gradient\n\nNegative\npressure gradient\n\nSide wall\n\nSide wall\n\nCentrifugal force\n\nConvex surface: low static pressure;\nhigh velocity\n\nNACA\n\n~~UNCLASSIFIED~~\nCONFIDENTIAL\n\nFigure 31.- A schematic diagram of the boundary layer and the secondary flow.\n\n71", "timestamp": "2026-07-22T04:30:38.466900+00:00"}
{"citation_id": "19930092013", "source_url": "https://ntrs.nasa.gov/api/citations/19930092013/downloads/19930092013.pdf", "page_number": 3, "total_pages": 21, "image_filename": "19930092013_p3.jpg", "text": "REPORT 948\n\nAN APPARATUS FOR VARYING EFFECTIVE DIHEDRAL  \nIN FLIGHT WITH APPLICATION TO A STUDY OF  \nTOLERABLE DIHEDRAL ON A COVENTIONAL  \nFIGHTER AIRPLANE\n\nBy WILLIAM M. KAUFFMAN, CHARLES J. LIDDELL, Jr.  \nALLAN SMITH, and RUDOLPH D. VAN DYKE, Jr.\n\nAmes Aeronautical Laboratory  \nMoffett Field, California", "timestamp": "2026-07-22T04:30:39.463671+00:00"}
{"citation_id": "19930086015", "source_url": "https://ntrs.nasa.gov/api/citations/19930086015/downloads/19930086015.pdf", "page_number": 36, "total_pages": 54, "image_filename": "19930086015_p36.jpg", "text": "NACA RM A9E24 CONFIDENTIAL 35\n\n[Figure: A line graph plotting \"Vertical distance from tunnel center line, z, in.\" on the y-axis against \"Mach number, M\" on the x-axis. The y-axis ranges from -24 to 24. The x-axis ranges from 1.0 to 1.8. The graph contains seven vertical series of data points connected by lines, each series corresponding to a specific value of D as indicated in the legend to the right. The legend lists D values: -175.25 (circle), -165.12 (square), -147.20 (diamond), -129.32 (upward triangle), -113.43 (right-pointing triangle), -100.47 (downward triangle), and -88.52 (left-pointing triangle). A NACA logo is present below the x-axis label.]\n\nFigure 10.- The variation of Mach number vertically at the center line of the test section of the Ames 6- by 6-foot supersonic wind tunnel. Stagnation pressure = 9 lb/sq in. abs.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T04:30:39.842368+00:00"}
{"citation_id": "19930082476", "source_url": "https://ntrs.nasa.gov/api/citations/19930082476/downloads/19930082476.pdf", "page_number": 2, "total_pages": 41, "image_filename": "19930082476_p2.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T04:30:41.846640+00:00"}
{"citation_id": "19930082245", "source_url": "https://ntrs.nasa.gov/api/citations/19930082245/downloads/19930082245.pdf", "page_number": 4, "total_pages": 66, "image_filename": "19930082245_p4.jpg", "text": "NACA TN No. 1596\n3\n\n$c_m$\nsection pitching-moment coefficient of airfoil about quarter-chord point of airfoil due to normal forces on main portion of airfoil and aileron; components due to chord forces on main portion of wing and aileron neglected\n\n$$\n\\left( \\left( \\frac{1}{c} \\right)^2 \\left[ \\int_0^{c_M} (P_U - P_L) \\left( x - \\frac{c}{4} \\right) dx + \\int_0^{c_a} (P_U - P_L) x \\ dx - c_{n_a} c_a \\left( x_h - \\frac{c}{4} \\right) \\cos \\delta_a + c_{n_a} c_a y_h \\sin \\delta_a \\right] \\right)\n$$\n\n$c_n$\nsection normal-force coefficient of airfoil determined from pressure-distribution data; component due to aileron chord force neglected $\\left( \\frac{1}{c} (c_M c_{n_M} + c_a c_{n_a} \\cos \\delta_a) \\right)$\n\n$c_{n_M}$\nsection normal-force coefficient of main portion of airfoil determined from pressure-distribution data $\\left( \\frac{1}{c_M} \\int_0^{c_M} (P_L - P_U) dx \\right)$\n\n$c_{n_a}$\nsection normal-force coefficient of aileron determined from pressure-distribution data $\\left( \\frac{1}{c_a} \\int_{-r}^{c_a} (P_L - P_U) dx \\right)$\n\nM\nMach number (V/a)\n\n$M_{cr}$\ncritical Mach number; that is, Mach number in undisturbed stream at which local velocity first reaches local velocity of sound at any point on airfoil surface\n\nP\npressure coefficient $\\left( \\frac{p_l - p}{q} \\right)$", "timestamp": "2026-07-22T04:30:44.448162+00:00"}
{"citation_id": "19930086078", "source_url": "https://ntrs.nasa.gov/api/citations/19930086078/downloads/19930086078.pdf", "page_number": 24, "total_pages": 42, "image_filename": "19930086078_p24.jpg", "text": "22\nNACA RM L9H04\n\nCONFIDENTIAL\n\nPitching-moment coefficient, $C_m$\nAngle of attack, $\\alpha$, deg\nLift coefficient, $C_l$\nDrag coefficient, $C_D$\n\nM\n$\\circ$ 0.19\n$\\square$ .27\n$\\diamond$ .37\n\nCONFIDENTIAL\nNACA\n\nFigure 5.- Plain-wing aerodynamic characteristics of the 45° sweptback wing configuration for several Mach numbers.", "timestamp": "2026-07-22T04:30:47.065391+00:00"}
{"citation_id": "19930086081", "source_url": "https://ntrs.nasa.gov/api/citations/19930086081/downloads/19930086081.pdf", "page_number": 21, "total_pages": 44, "image_filename": "19930086081_p21.jpg", "text": "CONFIDENTIAL\n\n$C_L$\n.2\n.1\n0\n-.1\n\n$\\delta$\n(deg)\n○ 0.3\n△ 2.4\n◇ 4.3\n□ 7.2\n▽ 10.3\n\n$C_m$\n.04\n0\n-.04\n\n$C_D$\n.04\n.02\n0\n\n-2 -1 0 1 2 3\n$\\alpha$, deg\n\nCONFIDENTIAL\n\n$C_t$\n.020\n.016\n.012\n.008\n.004\n0\n-.004\n\n$C_n$\n0\n-.004\n-.008\n\n-1 0 1 2 3\n$\\alpha$, deg\n\nNACA RM L9H05\n\nFigure 4.- Aerodynamic characteristics of a semispan delta wing with a half-delta tip control surface tested in the presence of a small fuselage. Fence off. Three-percent-thick control; R = 4.0 x $10^6$; M = 1.90. Flagged symbols denote repeat tests.\n\n19", "timestamp": "2026-07-22T04:30:48.621125+00:00"}
{"citation_id": "19930082090", "source_url": "https://ntrs.nasa.gov/api/citations/19930082090/downloads/19930082090.pdf", "page_number": 121, "total_pages": 122, "image_filename": "19930082090_p121.jpg", "text": "```markdown\nNACA TN No. 1455\n119\n\n<!-- Image (185, 168, 831, 897) -->\n\nFigure 64.- Thermal output and isothermal frictional pressure drops of slotted-fin heat exchanger P.\n\nNACA-Langley - 7-6-49 - 1150\n```", "timestamp": "2026-07-22T04:30:57.071742+00:00"}
{"citation_id": "19930082447", "source_url": "https://ntrs.nasa.gov/api/citations/19930082447/downloads/19930082447.pdf", "page_number": 4, "total_pages": 24, "image_filename": "19930082447_p4.jpg", "text": "2\nNACA TN No. 1775\n\nangle. The data were obtained at the Langley impact basin in smooth water for a wide range of trim angles, velocities, and flight-path angles. The test simulated flight conditions in which the effects of the presence of the afterbody is small. The data are compared with the generalized theoretical results previously mentioned and the effect of dead-rise angle on hydrodynamic loads is analyzed.\n\nSYMBOLS\n\n| | |\n| :--- | :--- |\n| $g$ | acceleration due to gravity, 32.2 feet per second$^2$ |\n| $n_{1_w}$ | impact load factor, measured normal to water surface, $g$ units |\n| $t$ | time after contact, seconds |\n| $W$ | dropping weight, pounds |\n| $\\dot{x}$ | velocity of model parallel to water surface, feet per second |\n| $y$ | draft of model normal to water surface, feet |\n| $\\dot{y}$ | velocity of model normal to water surface, feet per second |\n| $\\beta$ | angle of dead rise, degrees |\n| $\\gamma$ | flight-path angle relative to water surface, degrees |\n| $\\rho$ | mass density of water, slugs per cubic foot |\n| $\\tau$ | trim angle, degrees |\n| $f(\\beta)$ | dead-rise function |\n| $\\phi(A)$ | aspect-ratio correction factor |\n\nSubscripts:\n\n| | |\n| :--- | :--- |\n| $o$ | at water contact |\n| max | maximum |\n\nDimensionless variables:\n\nApproach parameter\n\n$$ \\kappa = \\frac{\\sin \\tau}{\\sin \\gamma_o} \\cos (\\tau + \\gamma_o) $$", "timestamp": "2026-07-22T04:31:05.677091+00:00"}
{"citation_id": "19930085842", "source_url": "https://ntrs.nasa.gov/api/citations/19930085842/downloads/19930085842.pdf", "page_number": 69, "total_pages": 104, "image_filename": "19930085842_p69.jpg", "text": "NACA RM L9C29\n65\n\nPitching-moment coefficient, $C_m$\nDrag coefficient, $C_D$\nLift coefficient, $C_L$\n\n$\\alpha, deg$ $\\delta_{a_L}, deg$\n0.6 0\n11.3 -15\n23.2 -32\n\n$\\alpha, deg$ $\\delta_{a_L}, deg$\n0.6 0\n11.3 -15\n23.2 -32\n\n$\\alpha, deg$ $\\delta_{a_L}, deg$\n0.6 0\n11.3 -15\n23.2 -32\n\nRight-ailavator deflection, $\\delta_{a_R}, deg$\n\nNATIONAL ADVISORY\nCOMMITTEE FOR AERONAUTICS\n\nFigure 36.- Variation of $C_L$, $C_D$, and $C_m$ with right-ailavator deflection on a $\\frac{1}{3}$-scale model of the airplane. Left-ailavator fixed; propellers removed; V = 100 miles per hour; $\\delta_T = 0^\\circ$; basic model configuration.", "timestamp": "2026-07-22T04:31:06.838759+00:00"}
{"citation_id": "19930082450", "source_url": "https://ntrs.nasa.gov/api/citations/19930082450/downloads/19930082450.pdf", "page_number": 3, "total_pages": 37, "image_filename": "19930082450_p3.jpg", "text": "```markdown\n2\nNACA TN No. 1778\n\np\nrivet pitch, inches\n\n$P_i$\ncompressive load per inch of panel width, kips per inch\n\n$\\bar{t}$\ncross-sectional area per inch of panel width, expressed as an\nequivalent or average thickness, inches\n\n$\\rho$\nradius of gyration, inches\n\n$\\bar{\\sigma}_F$\naverage stress at failing load, ksi\n\n$\\sigma_{cr}$\nstress for local buckling of sheet, ksi\n\n$\\sigma_{cy}$\ncompressive yield stress, ksi\n\nDIRECT-READING DESIGN CHARTS\n\nDirect-reading design charts for 24S-T aluminum-alloy flat compression\npanels with longitudinal formed Z-section stiffeners having the properties\nand proportions given in tables 1 to 5 are presented in two forms in\nfigures 2 to 9. In the first form (figs. 2 to 5), the design conditions\nof intensity of loading, effective length of panel, and skin thickness\nare incorporated in the ordinate $P_i/t_S$ and the abscissa $\\frac{P_i}{L/\\sqrt{C}}$. This\nform, having the design conditions incorporated in the ordinate and\nabscissa, is the more useful for most design purposes because the curves\nare more widely spaced and interpolation is more straightforward. In\nthe second (alternate) form (figs. 6 to 9), the average stress at\nfailure $\\bar{\\sigma}_F$ is plotted against $P_i/t_S$ as was done in the summary plots\nof reference 5. This alternate form, having the stress - an inverse\nmeasure of weight for a given load - as ordinate, is the more useful\nfor making generalizations and comparisons of structural efficiency\nbecause it shows how nearly the stress actually carried approaches the\nupper limit corresponding to the stress that would be achieved by a pure\nshell construction if a pure shell could carry the load without failure.\nThis upper limit of stress is represented by the lines for $\\bar{\\sigma}_F = \\frac{P_i}{t_S}$\n(infinite stiffener spacing) in figures 6 to 9.\n\nValues of the ratios of stiffener thickness to skin thickness $t_W/t_S$,\nspacing of rivet lines to skin thickness $S/t_S$ (because there is one\nrivet line associated with each Z-section, the stiffener spacing $b_S$ is\nequal to S, the spacing of rivet lines), and height of stiffener to\nstiffener thickness $H/t_W$, which will satisfy the design conditions, may\nbe found directly from these charts, and the corresponding section\nproperties $\\bar{t}/t_S$, $h/t_S$, and $\\rho/t_S$ may be found from tables 2 to 5. In\n```", "timestamp": "2026-07-22T04:31:07.055304+00:00"}
{"citation_id": "19930086003", "source_url": "https://ntrs.nasa.gov/api/citations/19930086003/downloads/19930086003.pdf", "page_number": 35, "total_pages": 36, "image_filename": "19930086003_p35.jpg", "text": "CONFIDENTIAL\n\nWing alone\nWing fuselage\n\nNACA RM L9J08\n\n$C_{D_{L=0}}$\n.08\n.04\n0\n\n$y_{c.p.}$\n60\n$C_L = 0.1$\n40\n\n$\\left(\\frac{\\partial C_L}{\\partial \\alpha}\\right)_M$\n.08\n$C_L = 0$\n.04\n$C_L = 0.4$\n\n$\\left(\\frac{\\partial \\epsilon}{\\partial \\alpha}\\right)_M$\n4\n$C_L = 0$\n$h_t$\n0\n-30\n30\n0\n\n$\\left(\\frac{\\partial C_m}{\\partial C_L}\\right)_M$\n4\n$C_L = 0.4$\n2\n0\n-2\n$C_L = 0$\n-4\n.6 .7 .8 .9 1.0 1.1 1.2\nMach number, M\n\n$\\left(\\frac{\\partial \\epsilon}{\\partial \\alpha}\\right)_M$\n4\n$C_L = 0$\n$h_t$\n-30\n30\n0\n\n$\\frac{q_{wake}}{q}$\n1.2\n$\\pm 30$\n.8\n$C_L = 0$\n0\n.6 .7 .8 .9 1.0 1.1 1.2\nMach number, M\n\nCONFIDENTIAL\n\nFigure 14.- Summary of aerodynamic characteristics for a model with 45° sweptback wing, aspect ratio 6, taper ratio 0.6, and NACA 65A006 airfoil section.\n\n33", "timestamp": "2026-07-22T04:31:11.014497+00:00"}
{"citation_id": "19930085930", "source_url": "https://ntrs.nasa.gov/api/citations/19930085930/downloads/19930085930.pdf", "page_number": 73, "total_pages": 92, "image_filename": "19930085930_p73.jpg", "text": "```markdown\nCONFIDENTIAL\n\nAverage exit Mach number, $M_{2av}$\n\n| | | |\n| :--- | :--- | :--- |\n| $\\bigcirc$ | 50-percent-span station | |\n| $\\square$ | 25-percent-span station | |\n| $\\diamondsuit$ | 10.15-percent-span station | |\n\n2.2\n2.0\n1.8\n1.6\n1.4\n1.2\n\n0 .2 .4 .6 .8 1.0 1.2 1.4 1.6\nDistance from convex surface\n\nConcave surface\n\nUNCLASSIFIED\nCONFIDENTIAL\n\nNACA\n\nFigure 32.- The variation of the average exit Mach number with distance from convex surface for model 3.\n\nNACA RM L9G07\n72\n```", "timestamp": "2026-07-22T04:31:13.717512+00:00"}
{"citation_id": "19930090382", "source_url": "https://ntrs.nasa.gov/api/citations/19930090382/downloads/19930090382.pdf", "page_number": 14, "total_pages": 37, "image_filename": "19930090382_p14.jpg", "text": "```markdown\n16\nNACA RM L9I07\n\nCONFIDENTIAL\n\nTip Mach number, $M_t$\n1.5\n1.0\n0.5\n0\n\nEfficiency, $\\eta$\n1.00\n.75\n.50\n.25\n0\n\nPower coefficient, $C_P$\n.300\n.275\n.250\n.225\n.200\n.175\n.150\n.125\n.100\n\nThrust coefficient, $C_T$\n.075\n.050\n.025\n0\n\nAdvance ratio, J\n(c) M=0.175.\n\n[Figure: Graph showing curves labeled $C_T$, $C_P$, and $\\eta$ plotted against Advance ratio, J. The curves for $C_T$ and $C_P$ are grouped near the bottom left, with lines indicating $\\beta_{0.75R}$ values of 15°, 20°, 25°, 30°, and 35°. The efficiency curve $\\eta$ peaks around J=1.5. A NACA logo is present near the top right of the graph area.]\n\nFigure 5 - Characteristics of NACA 4-(4)(06)-04 propeller.\n```", "timestamp": "2026-07-22T04:31:14.637480+00:00"}
{"citation_id": "19930085838", "source_url": "https://ntrs.nasa.gov/api/citations/19930085838/downloads/19930085838.pdf", "page_number": 85, "total_pages": 118, "image_filename": "19930085838_p85.jpg", "text": "NACA RM No. L9B23\n83\n\nAileron section hinge-moment coefficient, $C_{h_a}$\n\n$\\delta_t = -20^\\circ$\n$\\delta_a$ (deg)\n-10\n\n0\n-5\n\n$\\delta_t = -25^\\circ$\n$\\delta_a$ (deg)\n-15\n-10\n-10 check run\n\n0\n-25\n\nNACA\n\nSection angle of attack, $\\alpha_o$, deg\n\n(h) $\\delta_F = 40^\\circ$.\nFigure 10.- Concluded.", "timestamp": "2026-07-22T04:31:20.520937+00:00"}
{"citation_id": "19930086078", "source_url": "https://ntrs.nasa.gov/api/citations/19930086078/downloads/19930086078.pdf", "page_number": 25, "total_pages": 42, "image_filename": "19930086078_p25.jpg", "text": "```markdown\nNACA RM L9H04\n23\n\nCONFIDENTIAL\n\n<!-- Image (158, 126, 896, 826) -->\n\nFigure 6.- Lateral control characteristics of unswept wing with large-chord wing-tip aileron at various deflections, fully extended.\n```", "timestamp": "2026-07-22T04:31:20.727386+00:00"}
{"citation_id": "19930093773", "source_url": "https://ntrs.nasa.gov/api/citations/19930093773/downloads/19930093773.pdf", "page_number": 5, "total_pages": 47, "image_filename": "19930093773_p5.jpg", "text": "4\nNACA RM E9G09\n\n| Altitude (ft) | Flight Mach number |\n| :--- | :--- |\n| 5,000 | 0.21 |\n| 15,000 | 0.21, 0.53 |\n| 25,000 | 0.21, 0.53, 0.72, 0.85, 0.97 |\n| 35,000 | 0.21, 0.53, 0.72 |\n| 45,000 | 0.21, 0.53 |\n| 50,000 | 0.21 |\n\nComplete ram-pressure recovery at the compressor inlet was assumed in the calculation of flight Mach number. The fuel used was AN-F-32 with a lower heating value of 18,550 Btu per pound. The engine-inlet air temperature was held at approximately NACA standard values for each simulated flight condition except those of high altitude and low Mach number. Engine-inlet air temperatures below 439° R were unobtainable.\n\nRESULTS AND DISCUSSION\n\nAll the data obtained in the performance investigation of the engine with a standard exhaust nozzle are compiled in table 1. The engine-inlet pressures and temperatures deviated slightly from the desired inlet conditions. The data presented graphically in non-generalized form have therefore been adjusted to NACA standard altitude conditions by means of the factors $\\delta_a$ and $\\theta_a$ (appendix A).\n\nEngine Performance\n\nEffect of altitude. - Engine-performance data obtained at a flight Mach number of 0.21 at altitudes from 5000 to 50,000 feet are presented to show the effects of altitude on net thrust, air flow, fuel flow, specific fuel consumption, fuel-air ratio, and exhaust-gas total temperature in figure 4. Engine net thrust, air flow, and fuel consumption decreased consistently as the altitude increased (figs. 4(a) to 4(c)). Data obtained at high engine speeds are not shown for an altitude of 15,000 feet because the flight Mach number was inconsistent with other altitudes. At altitudes above 15,000 feet, the maximum engine speed was limited by turbine-outlet temperature.\n\nThe specific fuel consumption (fig. 4(d)) was essentially constant for altitudes from 5000 to 45,000 feet at engine speeds above 7200 rpm and for altitudes from 15,000 to 45,000 feet at", "timestamp": "2026-07-22T04:31:20.913217+00:00"}
{"citation_id": "19930082245", "source_url": "https://ntrs.nasa.gov/api/citations/19930082245/downloads/19930082245.pdf", "page_number": 5, "total_pages": 66, "image_filename": "19930082245_p5.jpg", "text": "4\nNACA TN No. 1596\n\n$P_{cr}$ critical pressure coefficient; that is, pressure coefficient\nat any point on airfoil surface where local velocity is\nequal to local velocity of sound\n\np static pressure in undisturbed stream\n\n$P_l$ local static pressure at a point on airfoil section\n\nq dynamic pressure in undisturbed stream $\\left(\\frac{1}{2}\\rho V^2\\right)$\n\nR Reynolds number $(\\rho Vc/\\mu)$\n\nr radius of round nose of aileron (0.0710 ft on model)\n\nV velocity in undisturbed stream\n\nx distance along chord from leading edge of airfoil or from\nhinge axis of aileron\n\n$x_h$ hinge-axis location along airfoil chord from leading edge of\nairfoil (1.600 ft on model)\n\n$y_h$ hinge-axis location normal to chord (0.0075 ft above chord\non model)\n\n$\\alpha$ angle of attack\n\n$\\delta_a$ aileron deflection; positive when trailing edge is down\n\n$\\rho$ mass density in undisturbed stream\n\n$\\mu$ coefficient of viscosity in undisturbed stream\n\nSubscripts:\n\nU upper surface\n\nL lower surface\n\nAPPARATUS AND METHODS\n\nApparatus.- The tests were made in the Langley 8-foot high-speed\ntunnel, which is of the single-return, circular-cross-section, closed-\nthroat type. The air-stream turbulence, as indicated by comparative\nairfoil measurements and by hot-wire measurements, is small but slightly\nhigher than that of the Langley two-dimensional low-turbulence pressure", "timestamp": "2026-07-22T04:31:22.443345+00:00"}
{"citation_id": "19930082090", "source_url": "https://ntrs.nasa.gov/api/citations/19930082090/downloads/19930082090.pdf", "page_number": 122, "total_pages": 122, "image_filename": "19930082090_p122.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T04:31:29.144775+00:00"}
{"citation_id": "19930082476", "source_url": "https://ntrs.nasa.gov/api/citations/19930082476/downloads/19930082476.pdf", "page_number": 3, "total_pages": 41, "image_filename": "19930082476_p3.jpg", "text": "NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nTECHNICAL NOTE NO. 1801\n\nWIND-TUNNEL INVESTIGATION OF THE SPINNING CHARACTERISTICS  \nOF A MODEL OF A TWIN-TAIL LOW-WING PERSONAL-OWNER-TYPE  \nAIRPLANE WITH LINKED AND UNLINKED RUDDER AND  \nAILERON CONTROLS\n\nBy Walter J. Klinar and Lawrence J. Gale\n\nSUMMARY\n\nA spin investigation has been conducted in the Langley 20-foot free-spinning tunnel of a model of a twin-tail low-wing personal-owner-type airplane with linked and unlinked rudder and aileron controls. The model was tested for two wing loadings and three mass distributions.\n\nThe results obtained when the rudders and ailerons were linked for two-control operation indicated that the model generally would not spin. The spins that were obtained were steep, and the test results indicated that full reversal of the controls from any spinning condition would result in satisfactory recovery.\n\nA study of the individual effects of rudders and ailerons at the various loadings showed that when a spin was obtained the inboard aileron (right aileron in a right spin) when deflected up was largely responsible for maintaining the spin. The results indicated that a reverse differential aileron system having the up aileron movement limited to a very small deflection would be effective in preventing the spin. The outboard rudder (left rudder in a right spin) was the more effective rudder in terminating or maintaining the spin, and differential rudder deflections which maintained the outboard rudder at or near neutral were particularly effective in preventing the attainment of spinning equilibrium.\n\nINTRODUCTION\n\nThe Langley Laboratory of the NACA is conducting an investigation to provide data that will be helpful in proportioning the mass and dimensional characteristics of light airplanes to eliminate the spin or to provide good spin-recovery characteristics. An approximate criterion for designing the tail of a light airplane for good spin recovery from fully developed spins", "timestamp": "2026-07-22T04:31:29.450801+00:00"}
{"citation_id": "19930086081", "source_url": "https://ntrs.nasa.gov/api/citations/19930086081/downloads/19930086081.pdf", "page_number": 22, "total_pages": 44, "image_filename": "19930086081_p22.jpg", "text": "CONFIDENTIAL\n\n.2\n.1\n$C_L$ 0\n-.1\n\n$\\delta$\n(deg)\n○ 0.3\n□ 4.0\n◇ 10.0\n\n.04\n$C_m$ 0\n-.04\n\n.04\n$C_D$ .02\n0\n\n-2 -1 0 1 2 3 4\n$\\alpha$, deg\n\n.020\n.016\n.012\n.008\n$C_l$ .004\n0\n-.004\n\n0\n$C_n$ -.004\n-.008\n\n-2 -1 0 1 2 3\n$\\alpha$, deg\n\nCONFIDENTIAL\n\nNACA\n\nFigure 5.- Aerodynamic characteristics of a semispan delta wing with a half-delta tip control surface tested in the presence of a small fuselage. Small fence on. Three-percent-thick control; R = $4.0 \\times 10^6$; M = 1.90. Flagged symbols denote repeat tests.\n\n20\nNACA RM L9E05", "timestamp": "2026-07-22T04:31:30.626487+00:00"}
{"citation_id": "19930085936", "source_url": "https://ntrs.nasa.gov/api/citations/19930085936/downloads/19930085936.pdf", "page_number": 58, "total_pages": 59, "image_filename": "19930085936_p58.jpg", "text": "NACA RM No. E9B03\n57\n\n(a) Angle of yaw, $-12^\\circ$.\n$$\n\\begin{array}{cccccccc}\n\\bigcirc & & & & & & & \\\\\n& \\bigcirc & \\bigcirc & \\bigcirc & & & & \\\\\n& -.133 & -.148 & -.143 & \\bigcirc & & & \\\\\n& & \\bigcirc & \\bigcirc & -.118 & & & \\\\\n& & -.046 & -.083 & & & & \\\\\n& & & \\bigcirc & -.053 & & & \\\\\n& & & \\bigcirc & -.024 & & & \\\\\n& & & & \\bigcirc & \\bigcirc & \\bigcirc & \\bigcirc \\\\\n& & & & -.004 & -.050 & -.084 & -.089 \\\\\n& & & & \\bigcirc & \\bigcirc & \\bigcirc & \\bigcirc \\\\\n& & & & .063 & .044 & -.012 & -.039 \\\\\n& & & & .034 & & & \\\\\n& & & & \\bigcirc & \\bigcirc & \\bigcirc & \\bigcirc \\\\\n& & & & .072 & .102 & .085 & .051 \\\\\n\\end{array}\n$$\n\n(b) Angle of yaw, $0^\\circ$.\n$$\n\\begin{array}{cccccccc}\n\\bigcirc & & & & & & & \\\\\n& \\bigcirc & \\bigcirc & \\bigcirc & & & & \\\\\n& .056 & .066 & .067 & \\bigcirc & & & \\\\\n& & \\bigcirc & \\bigcirc & .076 & & & \\\\\n& & .076 & .080 & & & & \\\\\n& & & \\bigcirc & .099 & & & \\\\\n& & & \\bigcirc & .098 & & & \\\\\n& & & & \\bigcirc & \\bigcirc & \\bigcirc & \\bigcirc \\\\\n& & & & .124 & .128 & .110 & .096 \\\\\n& & & & \\bigcirc & \\bigcirc & \\bigcirc & \\bigcirc \\\\\n& & & & .154 & .138 & .121 & .102 \\\\\n& & & & .151 & & & \\\\\n& & & & \\bigcirc & \\bigcirc & \\bigcirc & \\bigcirc \\\\\n& & & & .165 & .147 & .122 & .109 \\\\\n\\end{array}\n$$\n\n(c) Angle of yaw, $12^\\circ$.\n$$\n\\begin{array}{cccccccc}\n\\bigcirc & & & & & & & \\\\\n& \\bigcirc & \\bigcirc & \\bigcirc & & & & \\\\\n& .111 & .166 & .149 & \\bigcirc & & & \\\\\n& & \\bigcirc & \\bigcirc & .170 & & & \\\\\n& & .201 & .195 & & & & \\\\\n& & & \\bigcirc & .227 & & & \\\\\n& & & \\bigcirc & .235 & & & \\\\\n& & & & \\bigcirc & \\bigcirc & \\bigcirc & \\bigcirc \\\\\n& & & & .269 & .249 & .232 & .197 \\\\\n& & & & \\bigcirc & \\bigcirc & \\bigcirc & \\bigcirc \\\\\n& & & & .306 & .273 & .250 & .231 \\\\\n& & & & .287 & & & \\\\\n& & & & \\bigcirc & \\bigcirc & \\bigcirc & \\bigcirc \\\\\n& & & & .194 & .274 & .248 & .251 \\\\\n\\end{array}\n$$\n\n[Figure: NACA logo]\n\nFigure 15. - Pressure coefficients on wedge surface at $5^\\circ$ angle of attack for three angles of yaw.", "timestamp": "2026-07-22T04:31:32.777831+00:00"}