PDF/ocr_dataset_1072.jsonl

{"citation_id": "19930085928", "source_url": "https://ntrs.nasa.gov/api/citations/19930085928/downloads/19930085928.pdf", "page_number": 4, "total_pages": 22, "image_filename": "19930085928_p4.jpg", "text": "NACA RM No. A9A31 CONFIDENTIAL 3\n\nL length of subsonic diffusor\n\nm rate of mass flow\n\nM Mach number\n\np static pressure\n\nx distance between the duct throat and a station in the diffusor\n(considered positive in the downstream direction)\n\nγ ratio of the specific heat of air at constant pressure to the\nspecific heat at constant volume, 1.400\n\nSubscripts\n\no free stream\n\n1 duct entrance\n\n2 duct throat\n\n3 settling chamber\n\n4 exit throat\n\nx any station in the duct at the distance x from the duct throat\n(The subscripts designate the station of the measured quantity. See\nfig. 1.)\n\nAPPARATUS AND TESTS\n\nTwo models having different contraction ratios in the inlet\npassage were tested in the Ames 8- by 8-inch supersonic wind tunnel.\nThe tests were performed through a free-stream Mach number range of\n1.36 to 2.01 and at Reynolds numbers, based upon the length of the\nbody ahead of the inlet, between 2.21 and 3.10 million. A descrip-\ntion of the wind-tunnel equipment and the test procedure is given\nin reference 2.\n\nThe external shape of the models was the same as that of the\nmodel of reference 1. The forebody consisted of a 10-caliber ogival\nnose followed by a cylindrical section. The twin scoops enclosed\n37.2 percent of the forebody circumference, and the height-width\nratio of each scoop was 0.75. A $12^\\circ$ ramp was used ahead of each\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:38:31.894188+00:00"}
{"citation_id": "19930085862", "source_url": "https://ntrs.nasa.gov/api/citations/19930085862/downloads/19930085862.pdf", "page_number": 36, "total_pages": 60, "image_filename": "19930085862_p36.jpg", "text": "34\nNACA RM No. L9A07\n\n$P_R$\n$\\delta_a$\n(deg)\n25\n-20\n-15\n-10\n-5\n0\n5\n10\n15\n20\n25\n\n$C_{ha}$\nNACA\n\n$\\alpha$, deg\n\n(b) $C_{ha}$ and $P_R$ against $\\alpha$.\nFigure 9.— Continued.", "timestamp": "2026-07-22T06:38:32.332597+00:00"}
{"citation_id": "19930085951", "source_url": "https://ntrs.nasa.gov/api/citations/19930085951/downloads/19930085951.pdf", "page_number": 4, "total_pages": 92, "image_filename": "19930085951_p4.jpg", "text": "2\nNACA RM L9D29\n\nINTRODUCTION\n\nA general investigation of the aerodynamic characteristics of a series of full-scale 10-foot-diameter propellers at airspeeds up to 500 miles per hour has been made in the Langley 16-foot high-speed tunnel. The purpose of this general investigation was to determine the combined influence of propeller-design parameters and air compressibility upon propeller performance. The blade designs embody variations in shank form, blade airfoil section, design lift coefficient or camber, blade width, and blade thickness ratio. Most of the blades have the high-critical-speed NACA 16-series airfoil sections (reference 1) and have been designed neglecting compressibility effects for a minimum induced-energy loss when operating as four-blade propellers at an advance ratio of 2.1 and a blade angle of $45^\\circ$ at the 0.7 radius.\n\nThe primary effects of blade-section camber on propeller performance have been presented in reference 2, and the data showing the characteristics of other related propellers in the series have been presented in references 3 to 9. This paper presents the aerodynamic characteristics of two propellers and extends the investigation of related propellers to include those having thickness ratios as low as 0.05 at the 0.7 radius. The purpose of the paper is to make a comparison of the performance data for these two propellers with the data contained in references 3, 5, 7, and 8 to afford an evaluation of the effects of blade-section thickness ratio on propeller aerodynamic characteristics.\n\nThe thickness ratio of propeller blade sections is of increasing importance in the design of propellers for high speeds because of the compromise which must be made between structural requirements and the requirements for thin high-critical-speed sections necessary to avoid excessive compressibility losses. Compressibility effects have long been known to cause radical changes in the characteristics of the sections along a propeller blade, and a lack of suitable airfoil section characteristics at the present time has made it necessary to evaluate by propeller tests the effect of blade-section thickness ratios upon propeller performance.\n\nSYMBOLS\n\n| | | |\n| :--- | :--- | :--- |\n| B | number of blades | |\n| b | blade width, feet | |\n| $C_P$ | propeller power coefficient | $\\left(\\frac{P}{\\rho n^3 D^5}\\right)$ |\n| $C_T$ | propeller thrust coefficient | $\\left(\\frac{T}{\\rho n^2 D^4}\\right)$ |", "timestamp": "2026-07-22T06:38:33.068490+00:00"}
{"citation_id": "19930085899", "source_url": "https://ntrs.nasa.gov/api/citations/19930085899/downloads/19930085899.pdf", "page_number": 21, "total_pages": 29, "image_filename": "19930085899_p21.jpg", "text": "20\nNACA RM No. L9A21\n\nBending-moment coefficient, $C_B$\nLift coefficient, $C_L$\n\n| M | |\n| :--- | :--- |\n| 1.15 | $\\triangleleft$ |\n| 1.10 | $\\triangleright$ |\n| 1.08 | $\\circ$ |\n| 1.05 | $\\triangle$ |\n| 1.03 | $\\diamond$ |\n| 1.00 | $\\square$ |\n| .98 | $\\triangle$ |\n| .95 | $\\nabla$ |\n| .93 | $\\square$ |\n| .90 | $\\triangleright$ |\n| .88 | $\\nabla$ |\n| .85 | $\\nabla$ |\n| .80 | $\\diamond$ |\n| .70 | $\\square$ |\n| .60 | $\\circ$ |\n\n[Figure: A graph plotting Bending-moment coefficient ($C_B$) against Lift coefficient ($C_L$) for various Mach numbers (M). The y-axis ranges from -0.16 to 0.32. The x-axis ranges from -0.2 to 0.8. Multiple curves are plotted, each corresponding to a specific M value indicated in the legend on the right with unique symbols.]\n\nFigure 8.- Concluded.", "timestamp": "2026-07-22T06:38:34.362407+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 61, "total_pages": 78, "image_filename": "19930082483_p61.jpg", "text": "NACA TN No. 1807\n59\n\n1032\n\n[Figure: A circular mechanical assembly with various labeled parts. Labels include: \"Active nozzle arc\" pointing to an arc at the top; \"Active nozzles\" pointing to the outer ring of vanes; \"Nozzle baffles\" pointing to segments in the lower left; \"Shielded-type inlet thermocouples\" pointing to sensors in the lower right. A ruler marked in inches is at the bottom of the assembly.]\n\n(a) Upstream view.\n\nNACA\nC. 19004\n6.16.47\n\nFigure 4. - Nozzle-inlet section of turbine with baffle segments installed for runs with gas admission over $120^\\circ$ of nozzle periphery.", "timestamp": "2026-07-22T06:38:35.675473+00:00"}
{"citation_id": "19930082918", "source_url": "https://ntrs.nasa.gov/api/citations/19930082918/downloads/19930082918.pdf", "page_number": 46, "total_pages": 62, "image_filename": "19930082918_p46.jpg", "text": "NACA TN 1940\n53\n\n[Figure: Micrograph showing a textured surface with scattered dark features and a crack-like line near the top edge.]\n\n(d) Aged 100 hours.\n\n[Figure: Micrograph showing a textured surface with more numerous and larger dark features compared to the previous image, including a prominent elongated feature at the bottom left.]\n\n(e) Aged 1000 hours.\n\nFigure 7.- Concluded.\nNACA", "timestamp": "2026-07-22T06:38:36.732824+00:00"}
{"citation_id": "19930085900", "source_url": "https://ntrs.nasa.gov/api/citations/19930085900/downloads/19930085900.pdf", "page_number": 25, "total_pages": 33, "image_filename": "19930085900_p25.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:38:37.840597+00:00"}
{"citation_id": "19930083221", "source_url": "https://ntrs.nasa.gov/api/citations/19930083221/downloads/19930083221.pdf", "page_number": 42, "total_pages": 47, "image_filename": "19930083221_p42.jpg", "text": "40\nNACA TN No. 1824\n\nWhen the span of the swept-back wing becomes very large, the\nslope of the trailing edge\napproaches asymptotically the\nslope of the leading edge. It\nfollows that for infinitely\nlarge aspect ratio the limiting\nvalue of the load distribution\non the outboard sections should\napproach the value given by\nsimple sweep theory for an\ninfinitely long swept-back\nlifting surface with constant\nchord. This result is, in\nfact, a consequence of equa-\ntion (71).\n\n<!-- Image (139, 147, 465, 391) -->\n\nFigure 19.- Variation of lift\nand drag with aspect ratio for\na swept-back wing at $M_o = 1$.\n\nAmes Aeronautical Laboratory,\nNational Advisory Committee for Aeronautics,\nMoffett Field, Calif.\n\nAPPENDIX\n\nLIST OF IMPORTANT SYMBOLS\n\n| | |\n| :--- | :--- |\n| $a_1$ | y coordinate of trailing edge, $y = a_1(x)$ |\n| $a_o$ | free-stream speed of sound |\n| $a$ | local speed of sound |\n| $a^*$ | critical speed of sound |\n| $A$ | aspect ratio $\\left[ \\frac{(\\text{span})^2}{(\\text{wing area})} \\right]$ |", "timestamp": "2026-07-22T06:38:38.614913+00:00"}
{"citation_id": "19930085870", "source_url": "https://ntrs.nasa.gov/api/citations/19930085870/downloads/19930085870.pdf", "page_number": 24, "total_pages": 92, "image_filename": "19930085870_p24.jpg", "text": "NACA RM No. L9D07 CONFIDENTIAL 23\n\nTABLE 2.- SUMMARY OF RESULTS FOR 8-PERCENT-\nTHICK TRIANGULAR WINGS\n\n| Wing | Wedge leading edge | | | | Elliptical leading edge | | | |\n| :--- | :---: | :---: | :---: | :---: | :---: | :---: | :---: | :---: |\n| | $\\left(\\frac{dC_L}{d\\alpha}\\right)_{L=0}$ | $\\left(\\frac{dC_m}{d\\alpha}\\right)_{L=0}$ | $\\left(\\frac{L}{D}\\right)_{max}$ | $C_{D_{min}}$ | $\\left(\\frac{dC_L}{d\\alpha}\\right)_{L=0}$ | $\\left(\\frac{dC_m}{d\\alpha}\\right)_{L=0}$ | $\\left(\\frac{L}{D}\\right)_{max}$ | $C_{D_{min}}$ |\n| | | | | M = 1.62 | | | | |\n| 1 | 0.0232 | -0.00075 | 5.0 | 0.0133 | 0.0221 | -0.00004 | 5.3 | 0.0121 |\n| 2 | .0337 | -.00113 | 5.3 | .0184 | .0347 | -.00008 | 5.8 | .0159 |\n| 3 | .0366 | -.00133 | 5.1 | .0220 | .0393 | -.00003 | 5.6 | .0184 |\n| 4 | .0382 | -.00127 | 5.0 | .0255 | .0407 | -.00011 | 5.5 | .0207 |\n| 5 | .0388 | -.00118 | 4.8 | .0288 | .0423 | .00013 | 5.3 | .0233 |\n| 6 | .0388 | -.00106 | 4.7 | .0309 | .0422 | .00013 | 5.0 | .0261 |\n| 7 | .0384 | -.00100 | 4.6 | .0313 | .0426 | .00039 | 5.0 | .0259 |\n| 8 | .0385 | -.00081 | 4.4 | .0352 | .0421 | .00047 | 4.6 | .0300 |\n| 9 | .0387 | -.00056 | 4.2 | .0372 | .0429 | .00070 | 4.4 | .0324 |\n| 10 | .0396 | -.00054 | 4.2 | .0388 | .0431 | .00075 | 4.4 | .0337 |\n| 11 | .0416 | .00007 | 4.1 | .0433 | .0449 | .00146 | 4.1 | .0399 |\n| | | | | M = 1.92 | | | | |\n| 1 | 0.0216 | -0.00078 | 5.2 | 0.0125 | 0.0215 | -0.00036 | 5.3 | 0.0119 |\n| 2 | .0287 | -.00013 | 4.9 | .0186 | .0298 | -.00023 | 5.3 | .0154 |\n| 3 | .0294 | -.00095 | 4.7 | .0219 | .0317 | -.00010 | 5.0 | .0185 |\n| 4 | .0295 | -.00083 | 4.4 | .0256 | .0328 | -.00005 | 4.9 | .0205 |\n| 5 | .0300 | -.00056 | 4.3 | .0277 | .0335 | .00029 | 4.6 | .0228 |\n| 6 | .0296 | -.00048 | 4.1 | .0291 | .0334 | .00027 | 4.5 | .0258 |\n| 7 | .0299 | -.00034 | 4.1 | .0292 | .0332 | .00057 | 4.3 | .0253 |\n| 8 | .0308 | -.00018 | 4.0 | .0333 | .0330 | .00073 | 4.0 | .0293 |\n| 9 | .0316 | .00000 | 3.8 | .0342 | .0337 | .00077 | 3.9 | .0323 |\n| 10 | .0324 | .00010 | 3.9 | .0355 | .0340 | .00110 | 3.8 | .0331 |\n| 11 | .0346 | .00057 | 3.7 | .0396 | .0353 | .00180 | 3.5 | .0397 |\n| | | | | M = 2.40 | | | | |\n| 1 | 0.0189 | -0.00070 | 4.7 | 0.0127 | 0.0192 | -0.00009 | 5.2 | 0.0109 |\n| 2 | .0223 | -.00058 | 4.4 | .0179 | .0236 | .00008 | 4.8 | .0148 |\n| 3 | .0225 | -.00027 | 4.2 | .0211 | .0244 | .00035 | 4.5 | .0179 |\n| 4 | .0229 | -.00003 | 4.1 | .0238 | .0246 | .00047 | 4.3 | .0196 |\n| 5 | .0237 | .00009 | 4.0 | .0260 | .0247 | .00055 | 4.0 | .0224 |\n| 6 | .0243 | .00016 | 3.8 | .0272 | .0249 | .00059 | 3.8 | .0248 |\n| 7 | .0245 | .00022 | 3.8 | .0272 | .0254 | .00064 | 3.7 | .0252 |\n| 8 | .0256 | .00037 | 3.7 | .0301 | .0254 | .00077 | 3.3 | .0291 |\n| 9 | .0266 | .00053 | 3.7 | .0293 | .0258 | .00086 | 3.2 | .0319 |\n| 10 | .0281 | .00053 | 3.5 | .0325 | .0262 | .00095 | 3.2 | .0333 |\n| 11 | .0317 | .00100 | 3.4 | .0350 | .0270 | .00145 | 3.0 | .0397 |\n\n$^a$See table 1(a).\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:38:39.366966+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 39, "total_pages": 149, "image_filename": "19930083192_p39.jpg", "text": "NACA TN 1976\n35\n\nthe airplane as a whole since the motions of the airplane have an\nimportant bearing on gust-load calculation. The significant parameters\nof airplane behavior as determined from analytical and experimental\ninvestigations are subsequently summarized and possible simplifications\ninvestigated.\n\nAnalytical and Experimental Studies\n\nThe material presented is the result of analytical and experimental\nstudy of arbitrary and special configurations. The analytical studies\nhave been of a general nature and about two-thirds of the experimental\nstudies can be so classed. The rest of the experimental research has\nbeen done in connection with specific problems or airplane designs. The\nscope and results of the analytical studies are presented first, followed\nby the experimental studies. In addition to the behavior of and the\nloads on airplanes caused by vertical gusts uniform across the span,\nlimited research and studies on unsymmetrical and lateral gusts are\ndescribed and reported in appropriate sections.\n\nAnalytical studies.- The analytical studies of the loads on and\nbehavior of an airplane traversing a gust have been performed for a\nuniform upward-acting vertical gust. Current theory indicates that the\nincremental loadings for a downward-acting gust are equal in magnitude\nbut opposite in sign. Further assumptions are that:\n\n(1) The airplane maintains a constant forward speed during the\ntraverse of the gust\n\n(2) The airplane is in equilibrium prior to entry into the gust\n\n(3) The aerodynamic center of an aerodynamic surface is at the\nquarter-chord point\n\n(4) The moment coefficient at zero lift is a constant for transient\nconditions and has the same value as for steady flow conditions\n\n(5) The control surfaces of the airplane are locked\n\nExtended analyses in which the airplane is considered free to pitch\nas well as to rise under the action of a gust have been made by the NACA\nfor conventional, canard, and tailless airplanes. In the case of the\nconventional airplane, the analysis was rather wide in scope; whereas\nfor the other two types of aircraft, the analyses were limited to specific\ncases. The procedure used was an iteration process.\n\nThe calculations were aimed at covering all configurations and values\nof stability which would be reasonable for conventional airplanes and", "timestamp": "2026-07-22T06:38:39.480490+00:00"}
{"citation_id": "19930085890", "source_url": "https://ntrs.nasa.gov/api/citations/19930085890/downloads/19930085890.pdf", "page_number": 24, "total_pages": 26, "image_filename": "19930085890_p24.jpg", "text": "NACA RM No. E9C11\n\n[Figure: Three circular metal components labeled (a), (b), and (c), with a scale bar marked “INCHES” below them. Component (a) is a clean injector plate; (b) is a heavily damaged exhaust nozzle; (c) is a corroded or burned injector plate. A NACA logo with code “C-21842” and date “7-16-48” appears in the lower right of the figure.]\n\n(a) Injector plate  \n(b) Exhaust nozzle  \n(c) Injector plate  \n\nFigure 8. - Typical burned-out exhaust nozzle and injector plates.\n\n23", "timestamp": "2026-07-22T06:38:47.648880+00:00"}
{"citation_id": "19930085862", "source_url": "https://ntrs.nasa.gov/api/citations/19930085862/downloads/19930085862.pdf", "page_number": 37, "total_pages": 60, "image_filename": "19930085862_p37.jpg", "text": "NACA RM No. L9A07\n35\n\n<!-- Image (256, 159, 753, 809) -->\n\n(c) $C_L$ and $C_m$ against $\\alpha$.\nFigure 9.- Concluded.", "timestamp": "2026-07-22T06:38:54.218218+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 62, "total_pages": 78, "image_filename": "19930082483_p62.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:38:58.754556+00:00"}
{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 43, "total_pages": 72, "image_filename": "19930085491_p43.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:39:01.849401+00:00"}
{"citation_id": "19930082546", "source_url": "https://ntrs.nasa.gov/api/citations/19930082546/downloads/19930082546.pdf", "page_number": 60, "total_pages": 65, "image_filename": "19930082546_p60.jpg", "text": "NACA TN No. 1870\n59\n\n<!-- Image (175, 110, 842, 936) -->\n\nFigure 19.- Concluded.", "timestamp": "2026-07-22T06:39:02.062883+00:00"}
{"citation_id": "19930085911", "source_url": "https://ntrs.nasa.gov/api/citations/19930085911/downloads/19930085911.pdf", "page_number": 13, "total_pages": 52, "image_filename": "19930085911_p13.jpg", "text": "12\nCONFIDENTIAL\nNACA RM E9F22\n\n1. A maximum combustion efficiency of 91 percent was obtained\nin unit A-5 at a free-stream Mach number of 1.70, which sustained a\ndiffuser total-pressure recovery of 0.90. The gas total-temperature\nratio of 5.1 was equivalent to an exhaust-gas total temperature of\n4050° R. A maximum net acceleration of 2.0 g's and a thrust coef-\nficient of 0.56 were produced.\n\n2. For the flight conditions encountered, combustion effi-\nciencies from 40 to 91 percent were obtained with unit A-5 during\noperation within a fuel-air-ratio range of 0.043 to 0.070. At the\nleanest fuel-air ratio, 0.043, a low combustion efficiency of about\n40 percent occurred in unit A-5. As the fuel-air ratio suddenly\nincreased to 0.065, the combustion efficiency increased to 52 per-\ncent. Fuel-air ratios above 0.084 produced sporadic combustion\naccompanied by extremely low combustion efficiency in units A-3 and\nA-4. For unit A-5, at constant values of fuel-air ratio (0.062 to\n0.065) and combustion-chamber-inlet velocity (150 to 160 ft/sec),\nincreases in combustion-chamber-inlet static pressure from 2800 to\n7200 pounds per square foot and static temperature from 635° to\n790° R increased the combustion efficiency from 52 to 91 percent.\n\n3. As expected, a decrease in gas total-temperature ratio was\naccompanied by a decrease in diffuser total-pressure recovery\nlargely due to increasing shock losses within the diffuser. At a\nfree-stream Mach number of 1.20, the diffuser total-pressure recov-\nery decreased from 0.94 to 0.52 with a decrease in gas total-\ntemperature ratio from 5.0 to 1.0.\n\n4. Thrust coefficients increased with an increase in gas total-\ntemperature ratio and flight Mach numbers within the range of the\ndata obtained. At a free-stream Mach number of 1.00, the thrust\ncoefficient rose from -0.2 to 0.48 as the gas total-temperature\nratio increased from 1.0 to 6.0.\n\n5. For a given gas total-temperature ratio, the minimum value\nof external drag coefficient (approximately 0.115) occurred at free-\nstream Mach numbers of 0.90 to 1.00 and the maximum values occurred\nat free-stream Mach numbers of 1.10 to 1.20. At a free-stream Mach\nnumber of 1.70, increasing the gas total-temperature ratio from\n4.0 to 5.0 increased the external drag coefficient from 0.17 to\n0.32 as the flow conditions ahead of the inlet were altered.\n\nLewis Flight Propulsion Laboratory,\nNational Advisory Committee for Aeronautics,\nCleveland, Ohio.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:39:02.428801+00:00"}
{"citation_id": "19930085951", "source_url": "https://ntrs.nasa.gov/api/citations/19930085951/downloads/19930085951.pdf", "page_number": 5, "total_pages": 92, "image_filename": "19930085951_p5.jpg", "text": "NACA RM L9D29\nUNCLASSIFIED\nCONFIDENTIAL\n3\n\n$C_{l_d}$ blade-section design lift coefficient\n\nD propeller diameter, feet\n\nh blade-section maximum thickness, feet\n\nJ propeller advance ratio $\\left( \\frac{V}{nD} \\right)$\n\nM air-stream Mach number\n\n$M_t$ helical-tip Mach number $\\left( M \\sqrt{1 + \\left( \\frac{\\pi}{J} \\right)^2} \\right)$\n\nn propeller rotational speed, rps\n\nP power absorbed by propeller, foot-pounds per second\n\nT propeller thrust, pounds\n\nV airspeed, feet per second\n\nx fraction of propeller-tip radius\n\n$\\beta$ blade angle at any radius, degrees\n\n$\\beta_{0.75R}$ blade angle at 0.75 tip radius, degrees\n\n$\\eta$ propeller efficiency $\\left( \\frac{C_T}{J C_P} \\right)$\n\n$\\eta_i$ induced efficiency\n\n$\\rho$ mass density of air, slugs per cubic foot\n\n$\\sigma$ solidity $\\left( \\frac{Bb}{\\pi D} \\right)$\n\nAPPARATUS\n\nPropeller dynamometer.— Photographs of the 2000-horsepower dynamometer are shown in figures 1 and 2, and a diagram showing the important dimensions of the propeller dynamometer and its location with respect to the\n\nUNCLASSIFIED\nCONFIDENTIAL", "timestamp": "2026-07-22T06:39:03.498928+00:00"}
{"citation_id": "19930082918", "source_url": "https://ntrs.nasa.gov/api/citations/19930082918/downloads/19930082918.pdf", "page_number": 47, "total_pages": 62, "image_filename": "19930082918_p47.jpg", "text": "NACA TN 1940\n55\n\n[Figure: (a) Unaged.]\n\n[Figure: (b) Aged 0.5 hour.]\n\n[Figure: (c) Aged 1.0 hour.]\n\n[Figure: (d) Aged 3.0 hours.]\n\nNACA\nFigure 8.- Effect of aging at 1600° F on microstructure of low-carbon\nN-155 alloy solution-treated 10 hours at 2200° F and water-quenched.\nCross section of bar X1000. Electrolytically etched in 10 percent\nchromic acid.", "timestamp": "2026-07-22T06:39:03.792012+00:00"}
{"citation_id": "19930085900", "source_url": "https://ntrs.nasa.gov/api/citations/19930085900/downloads/19930085900.pdf", "page_number": 26, "total_pages": 33, "image_filename": "19930085900_p26.jpg", "text": "NACA RM L9D20\nCONFIDENTIAL\n25\n\n[Figure: (a) Normal jets; trim, 6.3°.]\n\n[Figure: (b) Slanted jets; trim, 7.9°.]\n\n[Figure: (c) Slanted jets; trim, 7.9°. NACA L-59857]\n\nFigure 11.- Jets spaced $\\frac{1}{4}$-inch apart; station 10 to 42; 35 feet per second.\nCONFIDENTIAL", "timestamp": "2026-07-22T06:39:04.074738+00:00"}
{"citation_id": "19930085928", "source_url": "https://ntrs.nasa.gov/api/citations/19930085928/downloads/19930085928.pdf", "page_number": 5, "total_pages": 22, "image_filename": "19930085928_p5.jpg", "text": "4\nCONFIDENTIAL\nNACA RM No. A9A31\n\nduct entrance.\n\nThe model dimensions and the internal duct shapes are shown in figures 1 and 2. The contraction ratios $A_2/A_1$ were selected for two different inlet Mach numbers. It was originally believed that the flow through a twin-scoop inlet having the proper slot area and dimensions would be similar to that through the perforated inlet of reference 3. If so, there would be no difficulty in causing the normal wave to move into the inlet at the design Mach number, and the scoops could be made to operate with a weak normal shock wave in the throat of the constricted passage. Model A of figure 1 had an inlet-contraction ratio of 0.914, the value for isentropic compression to sonic velocity from a uniform inlet Mach number of 1.36. With the model tested, an average inlet Mach number of 1.36 would occur at a free-stream Mach number of approximately 1.6. If there were no slots in the duct walls and if the flow were unidimensional and inviscid, this contraction ratio would permit a normal shock wave to enter the inlet when the intake Mach number was greater than 1.5 (reference 4) or when the free-stream Mach number was greater than 1.8. Model B had an inlet contraction ratio of 0.748, the value for isentropic compression to sonic velocity from a uniform inlet Mach number of 1.70, a value which occurred at a free-stream Mach number of approximately 2.0. If there were no slots in the duct walls of this model, a normal shock wave could not theoretically enter the inlet at even the maximum test Mach number.\n\nIn the subsonic diffuser of the model of reference 1, the rate of change of cross-sectional area with longitudinal position in the diffuser increased slowly from zero at the inlet to a constant value of 0.080 square inch per inch at 25 percent of the diffuser length. The data of reference 5 show that a large adverse pressure gradient exists in the upstream section of such diffusers. Since there was an initial boundary layer on one wall of the scoops being tested, this adverse pressure gradient probably caused the retarded air to separate in the high-velocity section of the diffuser and created excessive pressure losses. To reduce the adverse pressure gradient and the probability of this separation, a diffuser was designed to change the internal pressure distribution. The shape was calculated according to unidimensional theory to produce a pressure gradient proportional to the local static pressure. In other words, as the pressure increased in the diffuser, the pressure gradient increased correspondingly; thus, the smallest gradient would occur immediately downstream of the inlet and, the largest, just ahead of the settling chamber. The resulting diffuser was trumpet shaped; it diverged at a very small angle immediately downstream of the channel throat\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:39:08.763970+00:00"}
{"citation_id": "19930083221", "source_url": "https://ntrs.nasa.gov/api/citations/19930083221/downloads/19930083221.pdf", "page_number": 43, "total_pages": 47, "image_filename": "19930083221_p43.jpg", "text": "NACA TN No. 1824\n41\n\n| Symbol | Definition |\n| :--- | :--- |\n| $C(u)$ | Fresnel's cosine integral $\\left( \\int_0^u \\cos \\frac{\\pi}{2} x^2 \\, dx \\right)$ |\n| $c_o$ | wing root chord |\n| $C_L$ | lift coefficient $\\left[ \\frac{\\text{lift}}{q(\\text{wing area})} \\right]$ |\n| $C_{L_\\alpha}$ | $\\frac{dC_L}{d\\alpha}$ |\n| $C_{L_\\alpha}(t)$ | indicial lift coefficient |\n| $C_{D_i}$ | induced drag coefficient $\\left[ \\frac{D_i}{q(\\text{wing area})} \\right]$ |\n| $D$ | drag |\n| $D_i$ | induced drag |\n| $\\text{erf}(x)$ | error function of $x \\left( \\frac{2}{\\sqrt{\\pi}} \\int_0^x e^{-\\lambda^2} \\, d\\lambda \\right)$ |\n| $E(\\psi_n, k_n)$ | elliptic integral of second kind $\\left( \\int_0^{\\psi_n} \\sqrt{1 - k_n^2 \\sin^2 \\varphi} \\, d\\varphi \\right)$ |\n| $E_n$ | $E(\\frac{\\pi}{2}, k_n)$ |\n| $F(\\psi_n, k_n)$ | elliptic integral of first kind $\\left( \\int_0^{\\psi_n} \\frac{d\\varphi}{\\sqrt{1 - k_n^2 \\sin^2 \\varphi}} \\right)$ |\n| $K_n$ | $F(\\frac{\\pi}{2}, k_n)$ |\n| $K_n', E_n'$ | elliptic integrals with moduli $k_n'$ |\n| $k_n$ | modulus of elliptic functions |\n| $k_n'$ | $\\sqrt{1 - k_n^2}$ |\n| $l$ | length of body |", "timestamp": "2026-07-22T06:39:08.967611+00:00"}
{"citation_id": "19930085913", "source_url": "https://ntrs.nasa.gov/api/citations/19930085913/downloads/19930085913.pdf", "page_number": 10, "total_pages": 34, "image_filename": "19930085913_p10.jpg", "text": "NACA RM L9F24\n\nThe variation of the flutter velocity with spanwise weight position for the configurations weighted at the leading edge and those weighted at the midchord are shown in figures 4(a) and 4(b), respectively.\n\nDISCUSSION OF RESULTS\n\nThe models used in the series of tests were solid metal cantilever wings with thin rectangular cross sections and could easily be reproduced in case flutter proved to be destructive. These models were of such a nature that they fluttered at low Mach numbers. The lifting characteristics of the airfoil section used are approximately the same as those of a conventional airfoil with the exception that flow separation associated with the stalling condition occurs at a lower angle of attack. Since the model was mounted at essentially a zero angle of attack, it is very unlikely that the flutter speed was appreciably influenced by this separation effect.\n\nThe first three natural frequencies, the flutter frequency, and the flutter velocity of each model configuration tested are given in table I. The quantities have been plotted in figures 3 and 4. The figures show that, in general, a marked change in flutter frequency and a large increase in the flutter speed occurred when the weight was located between 40 and 80 percent of the wing length.\n\nThe variation in flutter velocity due to a variation of sweepback for a given chordwise and spanwise position of the weight is shown in figure 4. The second and third natural modes of vibration of the models weighted at the leading edge (fig. 4(a)) were of a highly coupled nature as shown by figure 2. Apparently this large amount of coupling had a greater effect on the flutter speed than did sweepback. In general, for this leading-edge weight position, variation in sweepback did not cause a large difference in flutter speed. The major effect on the flutter speed of the models weighted at the leading edge was due to the spanwise location of the weight.\n\nIn figure 4(b) the flutter velocities of the models weighted at the midchord are presented. In this case the mass coupling was relatively small. As is noted in figure 4(b), the effect of sweepback was more pronounced. The flutter velocity of the unswept wing was not greatly affected by spanwise weight position. When the wing was swept back, however, spanwise weight position did have an effect, probably because sweepback induced an amount of coupling which was further increased by the addition of the weight to the wing.\n\nAn unswept wing carrying a single weight on the leading edge was experimentally investigated in reference 3 and analytically investigated", "timestamp": "2026-07-22T06:39:09.932753+00:00"}
{"citation_id": "19930085862", "source_url": "https://ntrs.nasa.gov/api/citations/19930085862/downloads/19930085862.pdf", "page_number": 38, "total_pages": 60, "image_filename": "19930085862_p38.jpg", "text": "```markdown\n36\nNACA RM No. L9A07\n\n<!-- Image (111, 109, 779, 849) -->\n\n(a) $C_l$, $C_n$, and $C_{N_a}$ against $\\alpha$.\nFigure 10.- Aileron characteristics of wing with extensible leading-edge\nflaps and fences.\n```", "timestamp": "2026-07-22T06:39:15.155706+00:00"}
{"citation_id": "19930085548", "source_url": "https://ntrs.nasa.gov/api/citations/19930085548/downloads/19930085548.pdf", "page_number": 40, "total_pages": 46, "image_filename": "19930085548_p40.jpg", "text": "NACA RM No. E8L30\n39\n\n1077\n\n| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | 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| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | |", "timestamp": "2026-07-22T06:39:15.480024+00:00"}
{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 44, "total_pages": 72, "image_filename": "19930085491_p44.jpg", "text": "NACA RM No. A8J04 CONFIDENTIAL 43\n\n[Figure: A photograph of a model installed in a wind tunnel. The model is a slender, pointed object mounted on a sting. A ruler is visible for scale. The image is labeled \"NACA A-12271\".]\n\nFigure 3.- Model installed in wind tunnel.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:39:18.240974+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 63, "total_pages": 78, "image_filename": "19930082483_p63.jpg", "text": "NACA TN No. 1807\n61\n\n[Figure: A circular mechanical assembly with concentric rings and bolts. Labels point to \"Active nozzle arc\" at the top and \"Nozzle baffles\" at the bottom right. A ruler marked in inches is visible at the bottom left of the image.]\n\n(b) Downstream view.\n\nFigure 4. - Concluded. Nozzle-inlet section of turbine with baffle segments installed for runs with gas admission over $120^\\circ$ of nozzle periphery.\n\nNACA\nC.20548\n1-29-48", "timestamp": "2026-07-22T06:39:18.407120+00:00"}
{"citation_id": "19930085922", "source_url": "https://ntrs.nasa.gov/api/citations/19930085922/downloads/19930085922.pdf", "page_number": 8, "total_pages": 22, "image_filename": "19930085922_p8.jpg", "text": "NACA RM No. L9C23\n\n7\n\nTABLE I.- ORDINATES OF SYMMETRICAL\nAIRFOIL SECTION\n\n[All dimensions in percent of wing chord\nparallel to plane of symmetry of wing]\n\n| Station, x | Ordinate, $\\pm y$ |\n| :--- | :--- |\n| 0 | 0 |\n| .5871 | 1.0958 |\n| .8803 | 1.3226 |\n| 1.4661 | 1.6687 |\n| 2.9264 | 2.2597 |\n| 5.8297 | 2.9981 |\n| 8.7103 | 3.4923 |\n| 11.5680 | 3.8626 |\n| 17.2154 | 4.3929 |\n| 22.7728 | 4.7516 |\n| 28.2409 | 4.9951 |\n| 33.6203 | 5.1488 |\n| 38.9118 | 5.2322 |\n| 44.1160 | 5.2200 |\n| 49.2336 | 5.1300 |\n| 54.2654 | 4.9088 |\n| 59.2118 | 4.5506 |\n| 64.0736 | 4.0784 |\n| 68.9587 | 3.5320 |\n| 73.5461 | 2.9550 |\n| 78.1583 | 2.3821 |\n| 82.6881 | 1.8395 |\n| 87.1366 | 1.3383 |\n| 91.5043 | .8757 |\n| 95.7921 | .4408 |\n| 100.0000 | .0206 |\n\n[Figure: Diagram of airfoil section with x and y axes]\nNACA", "timestamp": "2026-07-22T06:39:19.492852+00:00"}
{"citation_id": "19930082546", "source_url": "https://ntrs.nasa.gov/api/citations/19930082546/downloads/19930082546.pdf", "page_number": 61, "total_pages": 65, "image_filename": "19930082546_p61.jpg", "text": "60\nNACA TN No. 1870\n\n$$\n\\frac{p}{p_{\\text{measured at } \\theta=0.087}}\n$$\n\n| NACA Propeller | B | $\\beta_{0.75}$ (deg) |\n| :--- | :--- | :--- |\n| $\\diamond$ 4-(5)(08)-03 | 2 | 15 |\n| $\\triangle$ 4-(5)(08)-03 | 2 | 20 |\n| $\\triangle$ 4-(5)(08)-03 | 2 | 30 |\n| $\\square$ 4-(3)(06.3)-06 | 2 | 10 |\n| $\\square$ 4-(3)(06.3)-06 | 2 | 15 |\n| $\\square$ Square tip | | |\n\n[Figure: Graph plotting $\\frac{p}{p_{\\text{measured at } \\theta=0.087}}$ against $\\frac{d}{D}$. The x-axis ranges from 0 to .36. The y-axis ranges from 0 to 2.00. A curve is plotted with data points marked by symbols corresponding to the legend. A small inset diagram labeled \"NACA\" is near the top right of the plot area.]\n\nFigure 20.- Free-space pressure attenuation curve used in calculating the values of figure 19.", "timestamp": "2026-07-22T06:39:23.084738+00:00"}
{"citation_id": "19930085911", "source_url": "https://ntrs.nasa.gov/api/citations/19930085911/downloads/19930085911.pdf", "page_number": 14, "total_pages": 52, "image_filename": "19930085911_p14.jpg", "text": "NACA RM E9F22 CONFIDENTIAL 13\n\nAPPENDIX - METHODS OF CALCULATION\n\nSymbols\n\nThe following symbols are used in this report:\n\nA cross-sectional area, sq ft\n\nAcr critical area necessary to bring local Mach number isentropically to unity, sq ft\n\nAmax maximum cross-sectional area, sq ft\n\nan axial acceleration component due to difference between net thrust and drag, g's\n\nCD external drag coefficient\n\nCF net-thrust coefficient\n\nD external drag, lb\n\nFn net thrust, lb\n\ng acceleration due to gravity, ft/sec²\n\nHa enthalpy of air and fuel before combustion, Btu/lb air\n\nHg enthalpy of burned gases at exhaust-gas temperature, Btu/lb exhaust gas\n\nh lower heating value of fuel, 18,500 Btu/lb\n\nM Mach number\n\nP total pressure, lb/sq ft\n\np static pressure, lb/sq ft\n\nQ rate of heat release of flares, Btu/sec\n\nq dynamic pressure, lb/sq ft\n\nR gas constant, ft-lb/(°R)(lb)\n\nT total temperature, °R\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:39:23.952741+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 74, "total_pages": 98, "image_filename": "19930086073_p74.jpg", "text": "72\n\nLift coefficient, $C_L$\n\n| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | 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| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | |", "timestamp": "2026-07-22T06:39:26.767512+00:00"}
{"citation_id": "19930085900", "source_url": "https://ntrs.nasa.gov/api/citations/19930085900/downloads/19930085900.pdf", "page_number": 27, "total_pages": 33, "image_filename": "19930085900_p27.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:39:30.357815+00:00"}
{"citation_id": "19930082918", "source_url": "https://ntrs.nasa.gov/api/citations/19930082918/downloads/19930082918.pdf", "page_number": 48, "total_pages": 62, "image_filename": "19930082918_p48.jpg", "text": "NACA TN 1940\n57\n\n[Figure: Micrograph showing grain structure with precipitates and boundaries]\n(e) Aged 10 hours.\n\n[Figure: Micrograph showing grain structure with precipitates and boundaries]\n(f) Aged 100 hours.\n\n[Figure: Micrograph showing grain structure with precipitates and boundaries]\n(g) Aged 1000 hours.\n\nFigure 8.— Concluded.\nNACA", "timestamp": "2026-07-22T06:39:35.224488+00:00"}
{"citation_id": "19930085919", "source_url": "https://ntrs.nasa.gov/api/citations/19930085919/downloads/19930085919.pdf", "page_number": 8, "total_pages": 47, "image_filename": "19930085919_p8.jpg", "text": "NACA RM No. A9C21 CONFIDENTIAL 7\n\nof attack decreased, (2) the wing efficiency factor, $\\frac{\\mathrm{d} C_{\\mathrm{L}}^{2}}{\\mathrm{~d} C_{\\mathrm{D}_{1}}} \\times \\frac{1}{\\pi A}$,\n\ndecreased, and (3) the aerodynamic center shifted forward. The surface tufts indicated a complete breakdown of flow near the wing tip at a lift coefficient of about 0.4.\n\nThe addition of either fuselage increased the lift-curve slope $(\\partial C_{\\mathrm{L}} / \\partial \\alpha)$ from 0.042 to 0.046 per degree and increased the drag at low lift coefficients. The same increase of the lift-curve slope was measured for a geometrically similar model, having a full-span wing, in the Ames 40- by 80-foot wind tunnel (reference 4). The wing in combination with the short fuselage had the same general characteristics as the wing in combination with the long fuselage except for slight differences in the pitching moments. During the investigation of the various control devices, the short fuselage was used in combination with the wing to permit testing up to an angle of attack of $38^{\\circ}$.\n\nReynolds Number\n\nMost of the data in this report were obtained at a Reynolds number of 4.2 million; however, to investigate possible dynamic-scale effects the data presented in figure 8 were obtained throughout a Reynolds number range of 2.5 to 7.2 million. Increasing the Reynolds number from 2.5 to 4.2 million increased the lift coefficient attained before the occurrence of longitudinal instability of the wing with the long fuselage from about 0.4 to 0.5, but had a negligible effect on this lift coefficient of the plain wing. However, a further increase of Reynolds number to 7.2 million resulted in no improvement of this lift coefficient. The drag coefficients were reduced slightly for all lift coefficients between 0.1 and 0.8, but the lift-curve slope was not greatly affected by increasing the Reynolds number from 2.5 to 7.2 million.\n\nSplit Flaps\n\nThe effect of the 0.25-chord split flap in several chordwise positions on the characteristics of the model is shown in figure 9. The split flap with its hinge line at the trailing edge of the wing yielded the largest increment of lift coefficient for all angles of\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:39:36.378574+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 64, "total_pages": 78, "image_filename": "19930082483_p64.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:39:37.846912+00:00"}
{"citation_id": "19930085951", "source_url": "https://ntrs.nasa.gov/api/citations/19930085951/downloads/19930085951.pdf", "page_number": 6, "total_pages": 92, "image_filename": "19930085951_p6.jpg", "text": "4\nUNCLASSIFIED\nNACA RM L9D29\n\nLangley 16-foot-tunnel test section is shown in figure 3. A detailed description of all the test apparatus and the methods of measuring thrust and torque are presented in reference 3. The fairing profile was calculated from a distribution of sources and sinks to produce a body of revolution with uniform axial velocity in the plane of the propeller. This axial-velocity distribution has been checked experimentally and found to be uniform within 1 percent. The gap between the propeller blade and the spinner surface at the propeller blade-spinner juncture is very small (fig. 1) but is not sealed.\n\nPropeller blades.- The two propellers for which data are presented in this paper are the NACA 10-(3)(062)-045A and NACA 10-(3)(05)-045. The NACA design numbers are descriptive of the shape, size, and aerodynamic characteristics of the blades used in this investigation. The digits of the first group of numbers represent the propeller diameter in feet, and the remaining digits indicate the design lift coefficient, thickness ratio, and solidity per blade at the 0.7 radius. The following table shows the blade design numbers of the various propellers discussed in this paper, and also shows the significance of the groups of digits in the number designation:\n\n| NACA design number | $c_{l_d}$ at 0.7R | h/b at 0.7R | $\\sigma$/B at 0.7R |\n| :--- | :---: | :---: | :---: |\n| 10-(3)(08)-03 | 0.3 | 0.08 | 0.03 |\n| 10-(3)(08)-03R | .3 | .08 | .03 |\n| 10-(3)(12)-03 | .3 | .12 | .03 |\n| 10-(3)(05)-045 | .3 | .05 | .045 |\n| 10-(3)(062)-045 | .3 | .062 | .045 |\n| 10-(3)(062)-045A | .3 | .062 | .045 |\n| 10-(3)(08)-045 | .3 | .08 | .045 |\n\nThe suffix R indicates a blade having conventional round shank sections, and the suffix A indicates a blade with modified shank sections. The NACA 16-series blade sections were used for all the propellers listed in the table, and with the exception of the NACA 10-(3)(08)-03R blade, wide airfoil sections extend to the spinner. The spinner has a diameter 21.7 percent of the diameter of a 10-foot propeller.\n\nFigure 4 shows the blade-form curves for the NACA propellers having a solidity of 0.03 per blade at the 0.7 radius, and figure 5 shows a comparison of the blade sections at two radii for the same group of propellers. The blade designs are closely related, but two of the propellers of this group differ not only in thickness but also in distribution of section design lift coefficient, blade width, and pitch distribution. These differences between the NACA 10-(3)(08)-03 and NACA 10-(3)(08)-03R blades are the result of an effort to maintain the minimum induced-energy-loss loading\n\nUNCLASSIFIED", "timestamp": "2026-07-22T06:39:39.103304+00:00"}
{"citation_id": "19930085870", "source_url": "https://ntrs.nasa.gov/api/citations/19930085870/downloads/19930085870.pdf", "page_number": 25, "total_pages": 92, "image_filename": "19930085870_p25.jpg", "text": "24\nCONFIDENTIAL\nTABLE 3.- SUMMARY OF RESULTS FOR THIN-PLATE\nTRIANGULAR WINGS WITH M = 1.92\n\n| Wing | $\\left(\\frac{dC_L}{d\\alpha}\\right)_{L=0}$ | $\\left(\\frac{dC_m}{d\\alpha}\\right)_{L=0}$ | $\\left(\\frac{L}{D}\\right)_{max}$ | $C_{Dmin}$ | R | $\\left(\\frac{dC_L}{d\\alpha}\\right)_{L=0}$ | $\\left(\\frac{dC_m}{d\\alpha}\\right)_{L=0}$ | $\\left(\\frac{L}{D}\\right)_{max}$ | $C_{Dmin}$ | R | $\\in$ | $A_{l.e.}$ |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| | | | | | | | | | | | | |\n| 13 | 0.0388 | 0.0022 | 8.00 | 0.0070 | $1.08 \\times 10^6$ | 0.0379 | 0.0011 | 8.10 | 0.0071 | $1.06 \\times 10^6$ | 25.25 | 61.97 |\n| 14 | 0.0387 | .0014 | 7.80 | .0079 | .87 | .0380 | .0013 | 7.65 | .0090 | .85 | 55.57 |\n| 17 | 0.0386 | .0017 | 7.80 | .0083 | .82 | .0386 | .0018 | 7.35 | .0121 | .77 | 55.50 |\n| 18 | 0.0395 | .0019 | 7.85 | .0092 | .76 | .0395 | .0015 | 7.15 | .0121 | .75 | 54.83 |\n\nNACA RM No. L9D07\n[Figure: NACA logo]\nCONFIDENTIAL\n\n30% decrease in $\\frac{L}{D}$ from 1.5 to 1.3.\n24% \" \" $C_{Dmin}$ \" \" \"", "timestamp": "2026-07-22T06:39:44.177523+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 40, "total_pages": 149, "image_filename": "19930083192_p40.jpg", "text": "36 NACA TN 1976\n\nsuch values of the mass parameter and moment of inertia which would fall within the capacity of the gust tunnel. The pertinent airplane characteristics assumed in the analysis are given in table XI for all the combinations considered. As can be seen from the table, the static margins range from 0.009 to -0.503 and the mass parameter ranged from 10.25 to 30.75. The calculations were made for three gust-gradient distances - 0, 8, and 16 chords. The flight conditions assumed were a gust velocity of 6 feet per second and a forward speed of about 40 miles per hour. The calculations were not made for all combinations of variables noted in table XI, but some intermediate values were interpolated.\n\nThe results of the calculations are shown in tables XII(a), XII(b), and XII(c) as the acceleration increments at maximum acceleration for the airplane with the center of gravity at 15, 25, and 35 percent of the mean aerodynamic chord, respectively. Each table shows the results for the three gust-gradient distances and for various combinations of tail area and tail length. The incremental values of acceleration which are contributed by each motion or source are tabulated and then totaled to obtain the wing acceleration, the acceleration increment resulting from the lift on the tail, and finally the total acceleration increment impressed on the airplane.\n\nIn addition to the results of the analysis being presented in tabular form, figures 34 and 35 indicate the effect of pitch on the total wing load. In figure 34, the ratio of the total wing load increment, including the effect of pitch, divided by the total wing load increment, assuming the pitch equal to zero, is shown. A value of 1 indicates that the effect of pitch was negligible and values greater than 1 indicate that the effect of pitch increased the total wing load. The results are shown as a function of the static margin for three gust-gradient distances. The individual curves shown in the figure represent given configurations but different center-of-gravity positions. Figure 35 is a similar plot of data from tables XII(a), XII(b), and XII(c) of the total acceleration increment as a function of the static margin for comparison with the experimental data from the gust tunnel.\n\nFigure 36 shows the wing lift increment resulting from pitch of the airplane in terms of the total wing lift as a function of mass parameter. The three solid curves represent the three center-of-gravity positions. The curves represent the \"average\" airplane with medium tail length and medium tail area traversing a flat-top gust with a gradient distance of 8 chords. The dash lines included in the figure represent the pitch-increment ratio $\\frac{\\Delta \\theta}{U/V}$. This ratio has been used to correct for pitch effects on the basis that the change in acceleration due to pitch is", "timestamp": "2026-07-22T06:39:44.340284+00:00"}
{"citation_id": "19930082546", "source_url": "https://ntrs.nasa.gov/api/citations/19930082546/downloads/19930082546.pdf", "page_number": 62, "total_pages": 65, "image_filename": "19930082546_p62.jpg", "text": "NACA TN No. 1870\n61\n\n$$\\frac{C_p}{C_F M_t}$$\n\n| mB | |\n| :--- | :--- |\n| 1 | $\\circ$ |\n| 2 | $\\square$ |\n| 3 | $\\diamond$ |\n| 4 | $\\triangle$ |\n| 5 | $\\nabla$ |\n| 6 | $\\square$ |\n| 7 | $\\triangle$ |\n| 8 | $\\nabla$ |\n\n[Figure: A graph plotting $\\frac{C_p}{C_F M_t}$ against $M_t$. The x-axis ranges from 0.4 to 1.0. The y-axis ranges from 0 to 0.48. There are eight curves corresponding to mB values from 1 to 8, distinguished by different symbols and line styles as indicated in the legend.]\n\nNACA\n\n$M_t$\n\nFigure 21.— Effect of tip Mach number at constant power on the pressure amplitudes of the fundamental frequencies of various propellers.\n$\\frac{d}{D} = 0.10$.", "timestamp": "2026-07-22T06:39:46.201336+00:00"}
{"citation_id": "19930085911", "source_url": "https://ntrs.nasa.gov/api/citations/19930085911/downloads/19930085911.pdf", "page_number": 15, "total_pages": 52, "image_filename": "19930085911_p15.jpg", "text": "```markdown\n14 CONFIDENTIAL NACA RM E9F22\n\nt static temperature, $^\\circ$R\nV velocity, ft/sec\nW gross weight of ram jet, lb\n$W_a$ air flow, lb/sec\n$W_f$ fuel flow, lb/sec\n$W_i$ initial gross weight of ram jet, lb\n$\\gamma$ ratio of specific heat at constant pressure to specific heat\nat constant volume\n$\\eta_b$ combustion efficiency, percent\n$\\tau$ time, sec\n\nSubscripts:\n\n0 free stream\n0,B free stream behind normal shock\n1 diffuser inlet\n2 $4\\frac{5}{8}$ inches downstream of air inlet (at static orifice)\n3 65 inches downstream of air inlet (at dynamic-pressure rake)\n4 diffuser outlet or combustion-chamber inlet (upstream side of\nflame holder)\n5 combustion-chamber inlet (downstream side of flame holder)\n6 combustion-chamber outlet\n7 exhaust-nozzle outlet\n\nCalculations\n\nAn atmospheric survey is made after the drop of a ram-jet\nunit. From the survey, the free-stream static temperature $t_0$ and\npressure $p_0$ are determined as a function of true altitude, as\n\nCONFIDENTIAL\n\n1152\n```", "timestamp": "2026-07-22T06:39:47.779826+00:00"}
{"citation_id": "19930082918", "source_url": "https://ntrs.nasa.gov/api/citations/19930082918/downloads/19930082918.pdf", "page_number": 49, "total_pages": 62, "image_filename": "19930082918_p49.jpg", "text": "NACA TN 1940\n59\n\n[Figure: (a) Aged 1.0 hour.]\n\n[Figure: (b) Aged 10 hours.]\n\n[Figure: (c) Aged 100 hours.]\n\n[Figure: (d) Aged 1000 hours.]\n\nFigure 9.- Electron micrographs of replicas (X8500) prepared from low-carbon N-155 alloy solution-treated 10 hours at 2200° F, water-quenched, and aged at 1600° F.", "timestamp": "2026-07-22T06:39:59.625166+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 65, "total_pages": 78, "image_filename": "19930082483_p65.jpg", "text": "1032\n\nNACA TN NO. 1807\n\nStart of active\nnozzles\n\nNozzle blade\n\nActive nozzle cut-off\n\nGas baffles\n\nINCHES\n0 1 2 3 4 5\n\nNACA\nC.17903\n2-14-47\n\nFigure 5. - Detailed view of nozzle inlet showing beginning and cut-off of active nozzle arc occurring along leading edges\nof nozzle blades.\n\n63", "timestamp": "2026-07-22T06:40:06.234493+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 75, "total_pages": 98, "image_filename": "19930086073_p75.jpg", "text": "NACA RM A9H04\n73\n\nLift coefficient, $C_L$\nDrag coefficient, $C_D$\n\n$\\circ$ $\\delta_{a_L} = +10.8$, $\\delta_{a_R} = -10.8$\n$\\square$ Aileron deflection, $\\delta_a$, deg\n\n(b) $C_L$ vs $C_D$.\n\nFigure 16.— Continued.\n\n[Figure: Graph showing Lift coefficient vs Drag coefficient with two curves marked by circles and squares]", "timestamp": "2026-07-22T06:40:11.060946+00:00"}
{"citation_id": "19930085870", "source_url": "https://ntrs.nasa.gov/api/citations/19930085870/downloads/19930085870.pdf", "page_number": 26, "total_pages": 92, "image_filename": "19930085870_p26.jpg", "text": "NACA RM No. L9D07\n25\n\nCONFIDENTIAL\n\nModel sting support\nModel sting\nFixed windshield\nMoveable sting windshield\nAngle-of-attack mirror\n\nFigure 1.- Drawing of model and support installation.\nL-57800\nCONFIDENTIAL", "timestamp": "2026-07-22T06:40:12.370288+00:00"}
{"citation_id": "19930085922", "source_url": "https://ntrs.nasa.gov/api/citations/19930085922/downloads/19930085922.pdf", "page_number": 9, "total_pages": 22, "image_filename": "19930085922_p9.jpg", "text": "8\n\n37.12\n17.74\n8.72\n8.00\nC\n4.50\nSting support\n24.44°\n12.56\n(MAC)\n15.36\n25% c\n35°\n9.83\n6.69\n3.84\n38.03°\n17.67\n0 5 10\nScale, inches\n5.80\n2.80\n2.78\n9.28\n2.32\n\nTABULATED DATA\nWing\nArea 3.17 sq ft\nAspect ratio 3.0\nMean aerodynamic chord 1.05 ft\nDihedral 0°\nTaper ratio 0.6\nAirfoil (Table I) symmetrical\nLocation of max. thickness 0.39c\nMaximum thickness 0.105c\n\nVertical fins\nArea (two) 0.82 sq ft\nAspect ratio 1.75\nMoment reference point 0.25 M.A.C.\nAilerons\nArea (two) 0.348 sq ft\nSweep, hinge axis 24.44°\n\nNACA\n\nFigure 1.- Drawing of test wing and vertical fins.\n\nNACA RM No. L9G23", "timestamp": "2026-07-22T06:40:12.950816+00:00"}
{"citation_id": "19930082546", "source_url": "https://ntrs.nasa.gov/api/citations/19930082546/downloads/19930082546.pdf", "page_number": 63, "total_pages": 65, "image_filename": "19930082546_p63.jpg", "text": "62\nNACA TN No. 1870\n\nVibration amplitude, in.\nFundamental frequency of panel excitation, cps\n\n| | |\n| :--- | :--- |\n| $\\circ$ | Experiment (without reinforcement) |\n| $\\square$ | Experiment (with reinforcement) |\n| — | Calculated by equation 3 |\n| - - - | Calculated by equation 7(b) |\n\n[Figure: Graph showing vibration amplitude vs. fundamental frequency with multiple curves and data points. A NACA logo is visible in the bottom right corner of the plot area.]\n\n(a) Flat vertical wooden wall with two-blade propeller.\nFigure 22.- Panel frequency-response curves.", "timestamp": "2026-07-22T06:40:15.814591+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 41, "total_pages": 149, "image_filename": "19930083192_p41.jpg", "text": "NACA TN 1976\n37\n\nproportional to the ratio of the pitch increment to the gust angle. The distance between the dash and solid lines is a measure of the error of this assumption.\n\nThe data in tables XII(a), XII(b), and XII(c) were also utilized to obtain the total tail load increment divided by the tail load if the downwash, the vertical motion, and the pitch are assumed to be zero. The results are shown in figure 37 as a function of the static margin $dC_{m_{CG}}/dC_L$ for the three gust-gradient distances. The total tail load shown is that at the time of maximum total airplane load. The lines connecting the points represent the same configuration and the movement along the line to the right indicates increasing static stability.\n\nDetailed calculations have also been made for a flying wing with about 25° of sweepback and for a canard airplane. The characteristics of the airplanes are shown in table XI. The calculations for the flying wing were made for two center-of-gravity positions and for three gradient distances - 0, 8, and 17.5 chords. Table XIII shows the results of the calculations and includes the contributions from the different sources for the two center-of-gravity positions. The total wing load increments are shown for each center-of-gravity position. Calculations were made by using finite-aspect-ratio unsteady-lift functions on the basis that no fuselage was present. The total acceleration increment has been plotted in figure 38(a) for each center-of-gravity position. The canard airplane had the same general characteristics as the conventional Boeing B-247 transport airplane. The calculations for the canard airplane for net wing area and for the two unsteady-lift functions have been obtained from reference 15 and are presented in table XIV. The calculated total acceleration increments for the canard are shown in figure 38(b) as a function of gradient distance.\n\nFigure 39 indicates the effect on the calculated loads of substituting the wing loading for the mass parameter. This figure is a plot of the acceleration based on wing loading divided by that based on the mass parameter as a function of wing loading. The calculations were made for three classes of airplanes - transports, personal airplanes, and flying boats. The results are shown as a function of the wing loading. Points lying above the line indicate overestimation of the acceleration increment, and points below the line represent underestimation of the acceleration increment.\n\nAs mentioned in the section about gust structure, the suggestion was made of either a triangular or sinusoidal gust shape in the direction of flight. As a check on the interchangeability of the two gust shapes, figure 40 shows the ratio of the acceleration due to a sinusoidal gust to that for a triangular gust as a function of the mass parameter when the unsteady-lift functions for infinite aspect ratio and aspect ratio 6", "timestamp": "2026-07-22T06:40:23.684476+00:00"}
{"citation_id": "19930085928", "source_url": "https://ntrs.nasa.gov/api/citations/19930085928/downloads/19930085928.pdf", "page_number": 6, "total_pages": 22, "image_filename": "19930085928_p6.jpg", "text": "NACA RM No. A9A31 CONFIDENTIAL 5\n\nwhere the local Mach number was assumed to be 1.0, and at relatively large angles near the settling chamber where the local Mach number was assumed to be 0.2. The equation that relates the area at a given station to the desired pressure variation is derived in the appendix. The distance between the duct entrances and the settling chamber of the two models was about 14 percent less than that of the model of reference 1. This length was reduced because the computed area variation was very small in the high-velocity section of the diffuser if the original length were used. It was believed that the growth of the boundary layer in such a channel would compensate for the slight increase in divergence.\n\nThe tests were made with each model set at an angle of attack of $0^{\\circ}$. The effects of slots were investigated by testing first without slots and then with the slots that were found to produce the greatest recovery for the inlet form of reference 1. To study the effect of slot area upon total-pressure recovery, the various combinations of slot height and length shown in figure 1 were tested. The slot heights were approximately 28 percent and 14 percent of the scoop height, and the slot lengths were 75 percent and 110 percent of the distance from the scoop entrance to the duct throat. Measurements of the total pressure in the settling chamber of the models were made at three equally spaced circumferential positions. At pressure ratios near the maximum, the differences in the measurements were no greater than 2 percent of the total pressure, a fact which indicates a relatively uniform velocity distribution. However, differences up to 15 percent of the total pressure were observed when the flow into the scoops was unsteady or when the variation of pressure recovery with mass-flow ratio was large. The total-pressure ratios presented in this report are based upon the average of the three pressure measurements.\n\nRESULTS AND DISCUSSION\n\nSeveral features have been incorporated in the present models to determine if large improvements in total-pressure recovery could be attained. These features were the inlet contractions, slots, and subsonic diffusers designed to reduce the adverse pressure gradient in the high-velocity section. The results, therefore, include the combined effects of these variables. Since the improvement in recovery was found to be relatively large, it is desirable to evaluate the magnitude of the contribution of each variable and to determine the reason for its favorable effect. A subsequent report will discuss tests of models designed to provide this information.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:40:24.273852+00:00"}
{"citation_id": "19930085951", "source_url": "https://ntrs.nasa.gov/api/citations/19930085951/downloads/19930085951.pdf", "page_number": 7, "total_pages": 92, "image_filename": "19930085951_p7.jpg", "text": "NACA RM L9D29\nUNCLASSIFIED\nCONFIDENTIAL\n5\n\nas far inboard as possible on the round-shank blade. The NACA 10-(3)(12)-03 blade has the same radial distribution of blade-section design lift coefficient, the same blade width, and approximately the same pitch distribution as the NACA 10-(3)(08)-03 blade, but its thickness is greater at all radii.\n\nFigure 6 shows the blade-form curves and figure 7 shows the section comparisons for the group of propellers having a solidity of 0.045 per blade at the 0.7 radius. The propellers of this group have the same radial distribution of blade-section design lift coefficient, the same blade width, and approximately the same pitch distribution. The NACA 10-(3)(08)-045 design has the thickest shank sections of this group, although its outboard sections are considerably thinner than those of the NACA 10-(3)(062)-045 design. The NACA 10-(3)(062)-045A design was made by simply thinning the shank sections of the NACA 10-(3)(062)-045 blade until they had the same thickness as the NACA 10-(3)(08)-045 design at the spinner; the thickness of the sections between the spinner and the 0.7 radius was obtained by a faired line between these two radii. The NACA 10-(3)(05)-045 design was made by thinning the sections of the NACA 10-(3)(062)-045 blade until the sections at the spinner and tip had the same thickness as the NACA 10-(3)(062)-045A blade, but the sections between these two radii were made thinner.\n\nTESTS AND REDUCTION OF DATA\n\nThrust, torque, and rotational speed were measured during tests at fixed blade angles of $20^\\circ$, $25^\\circ$, $30^\\circ$, $35^\\circ$, $40^\\circ$, $45^\\circ$, $50^\\circ$, and $55^\\circ$ at the three-quarter (45-in.) radius. A constant rotational speed was used for most of the tests, and a range of advance ratio was covered by changing the tunnel airspeed, which could be varied from about 60 to 500 miles per hour. The range of blade angles covered at the various rotational speeds used in the tests of the NACA 10-(3)(062)-045A and NACA 10-(3)(05)-045 propellers is shown in table I. Similar information, together with figure numbers, is also shown in table I for the other propellers as taken from references 3, 5, 7, and 8. At the higher blade angles, the complete range of advance ratio could not be covered at the higher rotational speeds because of power limitations. In order to obtain propeller characteristics at maximum tunnel airspeeds, a blade angle ($45^\\circ$) was chosen for which the peak-efficiency operating condition could be attained when the tunnel airspeed was at or near the maximum and the dynamometer was operating at its maximum power and rotational speed. For these tests at a blade angle of $45^\\circ$, the rotational speed was varied to obtain data from the peak-efficiency condition to the zero-torque operating condition.\n\nThe test data have been corrected for tunnel-wall interference and for forces acting on the spinner by the methods described in reference 3 and are presented in the form of the usual thrust and power coefficients and propeller efficiency. Propeller thrust, as used herein, is defined as\n\nUNCLASSIFIED\nCONFIDENTIAL", "timestamp": "2026-07-22T06:40:31.405104+00:00"}
{"citation_id": "19930083221", "source_url": "https://ntrs.nasa.gov/api/citations/19930083221/downloads/19930083221.pdf", "page_number": 44, "total_pages": 47, "image_filename": "19930083221_p44.jpg", "text": "42\nNACA TN No. 1824\n\n$M_o$\nfree-stream Mach number $\\left( \\frac{V_o}{a_o} \\right)$\n\n$N_1(\\omega t, \\nu)$\n$$ \\int_{\\sqrt{\\nu}}^{\\sqrt{\\omega t}} \\left[ \\cos x^2 C(\\omega t - x^2) - \\sin x^2 S(\\omega t - x^2) \\right] dx $$\n\n$N_2(\\omega t, \\nu)$\n$$ \\int_{\\sqrt{\\nu}}^{\\sqrt{\\omega t}} \\left[ \\cos x^2 S(\\omega t - x^2) + \\sin x^2 C(\\omega t - x^2) \\right] dx $$\n\n$\\frac{\\Delta p}{q}$\nloading coefficient (pressure on lower surface minus pressure on upper surface divided by free-stream dynamic pressure)\n\n$q$\nfree-stream dynamic pressure $\\left( \\frac{1}{2} \\rho_o V_o^2 \\right)$\n\n$r, \\theta$\npolar coordinates in $yz$ plane ($y = r \\cos \\theta$, $z = r \\sin \\theta$)\n\n$S$\nFresnel's sine integral $\\left( \\int_o^u \\sin \\frac{\\pi}{2} x^2 dx \\right)$\n\n$s$\noperational equivalent of $t$\n\n$s_o$\nwing semispan\n\n$t'$\ntime\n\n$t$\n$a_o t'$\n\n$t_o$\nmaximum distance measured parallel to $y$ axis from $x$ axis to trailing edge (fig. 15)\n\n$u, v, w$\nperturbation velocity components in $x, y, z$ directions, respectively\n\n$V$\nlocal velocity\n\n$V_o$\nfree-stream velocity", "timestamp": "2026-07-22T06:40:34.275449+00:00"}
{"citation_id": "19930085911", "source_url": "https://ntrs.nasa.gov/api/citations/19930085911/downloads/19930085911.pdf", "page_number": 16, "total_pages": 52, "image_filename": "19930085911_p16.jpg", "text": "NACA RM E9F22 CONFIDENTIAL 15\n\nobtained by radar-tracking the descending airplane. From the radar data, the position of the ram-jet unit as a function of time after release is determined. By differentiation, the velocity and the trajectory angle of the unit are obtained. The velocity of the ram-jet unit is calculated by integrating the total acceleration, which is the sum of the telemetered net acceleration and the acceleration component due to gravity. An average velocity curve is drawn and corrected by applying wind-velocity corrections obtained from radar-tracking a weather balloon. The corrected relative air velocity is defined as the free-stream velocity $V_0$ of the ram jet.\n\nFree-stream conditions are determined in accordance with the following general equations for compressible flow:\n\n$$\nM = \\frac{V}{\\sqrt{\\gamma g R t}} \\tag{1}\n$$\n\n$$\nT = t \\left(1 + \\frac{\\gamma - 1}{2} M^2\\right) \\tag{2}\n$$\n\n$$\nP = p \\left(1 + \\frac{\\gamma - 1}{2} M^2\\right)^{\\frac{\\gamma}{\\gamma - 1}} \\tag{3}\n$$\n\nThe free-stream total pressure measured by the telemetering equipment is actually the total pressure as measured behind a normal shock when $M_0 > 1.0$ and is designated $P_{0,B}$. In order to obtain $P_0$ from the telemetered $P_{0,B}$, it is corrected for the normal-shock loss in accordance with\n\n$$\nP_0 = P_{0,B} \\left[ \\frac{(\\gamma - 1) M_0^2 + 2}{(\\gamma + 1) M_0^2} \\right]^{\\frac{\\gamma}{\\gamma - 1}} \\left( \\frac{2\\gamma}{\\gamma + 1} M_0^2 - \\frac{\\gamma - 1}{\\gamma + 1} \\right)^{\\frac{1}{\\gamma - 1}} \\tag{4}\n$$\n\nThese telemetered values of $P_0$ are used only in the performance calculation when values cannot be calculated from the radar data.\n\nThe Mach number at station 2 is determined from the following general equation. The static pressure is measured at station 2\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:40:34.475386+00:00"}