PDF/ocr_dataset_1053.jsonl

{"citation_id": "19930082646", "source_url": "https://ntrs.nasa.gov/api/citations/19930082646/downloads/19930082646.pdf", "page_number": 2, "total_pages": 37, "image_filename": "19930082646_p2.jpg", "text": "```markdown\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nTECHNICAL NOTE 1980\n\nEFFECT OF FOREBODY WARP AND INCREASE IN AFTERBODY LENGTH\nON THE HYDRODYNAMIC QUALITIES OF A FLYING-BOAT\nHULL OF HIGH LENGTH-BEAM RATIO\n\nBy Walter J. Kapryan\n\nSUMMARY\n\nIn an attempt to achieve improved hydrodynamic qualities for flying boats, an investigation was made to determine the combined effect of a warped forebody and extended afterbody on the hydrodynamic characteristics of a hull having a basic length-beam ratio of 15. Each of these modifications has been found to result in improvements in hydrodynamic qualities and their combined effect is therefore of interest.\n\nThe stable range of trims available for take-off was increased by warping the forebody and extending the afterbody. For take-off at constant elevator deflection, the range of center-of-gravity positions for satisfactory take-off stability was increased. The smooth-water landing stability remained satisfactory. Spray characteristics were improved, no spray entering the propellers or striking the flaps at the design gross load. Definite improvements in rough-water landing behavior were obtained with the combination of warped forebody and extended afterbody. For landings in waves 4 feet high, the impact accelerations were reduced more than 50 percent. The maximum amplitudes of oscillation in trim and rise occurring during the landing run-out were also substantially decreased. In general, the changes in hydrodynamic characteristics effected by combining these modifications were in the direction that would be predicted on the basis of results obtained for the individual modifications. Although the effects were not directly additive, they were cumulative and changes in opposite directions were compensating.\n\nINTRODUCTION\n\nRecent towing-tank investigations of flying-boat hulls having high length-beam ratios included the effect of forebody warp and the effect of afterbody length on hydrodynamic characteristics. These\n```", "timestamp": "2026-07-22T05:59:15.081913+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 35, "total_pages": 98, "image_filename": "19930086073_p35.jpg", "text": "1.4\n1.2\n1.0\n.8\n.6\n.4\n.2\n0\nLift coefficient, $C_L$\n\n0 4 8 12 16 20 24 28 32 36\nAngle of attack, $\\alpha$, deg\n\n$\\circ$ $\\square$\n0.0 12.1\nAngle of sideslip, $\\beta$, deg\n\n(a) $C_L$ vs $\\alpha$.\nFigure 6.- Wing alone at two angles of sideslip with left aileron deflected 11.7° and right aileron deflected -11.3°.\n\nNACA RM A59H04\n33", "timestamp": "2026-07-22T05:59:17.024742+00:00"}
{"citation_id": "19930086097", "source_url": "https://ntrs.nasa.gov/api/citations/19930086097/downloads/19930086097.pdf", "page_number": 27, "total_pages": 36, "image_filename": "19930086097_p27.jpg", "text": "NACA RM A9H11 CONFIDENTIAL 25\n\n<!-- Image (229, 121, 829, 388) -->\n\nFigure 1.- Airfoils considered in theoretical analysis.\n\n<!-- Image (252, 474, 692, 869) -->\n\nFigure 2.- Limiting base pressure coefficient.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:59:17.960320+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 18, "total_pages": 78, "image_filename": "19930082483_p18.jpg", "text": "16\nNACA TN No. 1807\n\nThe turbine rotor (fig. 2) has 54 tapered blades $2\\frac{1}{8}$ inches long and has an over-all diameter of $16\\frac{1}{8}$ inches. The turbine is equipped with reaction-type blading; the amount of reaction, however, is slight. The turbine-nozzle assembly consists of 48 nozzle blades with a height of $2\\frac{1}{8}$ inches. The nozzle blades are held in position by an inner and an outer shroud band. The nozzle blades are precision cast and the rotor blades are machined. The surface finishes of each conform to the general standards for the particular methods of fabrication. Details of the rotor and nozzle blades are shown in figure 3.\n\nProvision was made to install gas baffles in the nozzle section to effect partial admission. Two views of the arrangement with $120^\\circ$ active nozzle arc are shown in figure 4.\n\nThe nozzle ring, having 48 nozzle blades, allows positioning of baffle segments so that for active nozzle-arc reduction in $7\\frac{1}{2}^\\circ$ increments, a number of whole nozzle passages is active. Thus the beginning and the cut-off of admission always occur along leading edges of, rather than between, nozzle blades. (See fig. 5.)\n\nSetup\n\nThe over-all test setup used for the investigation is shown in figure 6. Pressurized air supplied by the laboratory combustion-air system was introduced into an external hot-gas producer into which 62-octane gasoline was injected and burned. The combustion products provided the driving fluid, which then entered the turbine assembly normal to the turbine axis and flowed into the nozzle approach section in which a honeycomb-type flow straightener was installed. During this investigation, the engine burner lining was removed from the nozzle approach section and no burning was conducted at this location.\n\nFrom the nozzle approach section, the driving gases then passed through the nozzles and the turbine rotor and into the discharge section. The turbine-discharge section consists of an outer cylindrical shell with an internal diameter the same as that of the turbine-wheel stationary shroud, and an inner conical piece 30 inches long, which has a base diameter the same as the root diameter of the turbine wheel. Turbine back pressures were set by throttling to the laboratory low-pressure exhaust system.", "timestamp": "2026-07-22T05:59:19.532921+00:00"}
{"citation_id": "19930082546", "source_url": "https://ntrs.nasa.gov/api/citations/19930082546/downloads/19930082546.pdf", "page_number": 9, "total_pages": 65, "image_filename": "19930082546_p9.jpg", "text": "8\nNACA TN No. 1870\n\nThe absolute value of root mean square pressure $p$ is given by the following expression:\n\n$$\np = \\frac{1}{4\\sqrt{2}\\pi^2} \\left( \\left[ \\int_{0}^{2\\pi} \\left( Tx + \\frac{Qy \\sin \\theta}{R_o} \\right) \\frac{1}{S_\\theta^3} \\left[ \\cos(mB\\theta + kS_\\theta) + kS_\\theta \\sin(mB\\theta + kS_\\theta) \\right] d\\theta \\right]^2 \\right.\n$$\n$$\n\\left. + \\left[ \\int_{0}^{2\\pi} \\left( Tx + \\frac{Qy \\sin \\theta}{R_o} \\right) \\frac{1}{S_\\theta^3} \\left[ kS_\\theta \\cos(mB\\theta + kS_\\theta) - \\sin(mB\\theta + kS_\\theta) \\right] d\\theta \\right]^2 \\right)^{1/2}\n$$\n\nwhere\n\n$$\nk = \\frac{mB\\omega}{c}\n$$\n\nand\n\n$$\nS_\\theta = \\sqrt{x^2 + y^2 + R_o^2 - 2R_o y \\cos \\theta}\n$$\n\nwhich is the distance from the observer to the doublets of the effective propeller circle.", "timestamp": "2026-07-22T05:59:20.498015+00:00"}
{"citation_id": "19930091993", "source_url": "https://ntrs.nasa.gov/api/citations/19930091993/downloads/19930091993.pdf", "page_number": 18, "total_pages": 21, "image_filename": "19930091993_p18.jpg", "text": "```markdown\n14\nREPORT 928—NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nof 0.87. The results of these computations are plotted in figure 24, which shows that the power turbine can develop 1800 horsepower ($32.2 \\times 10^6$ ft-poundal/sec) with a specific fuel consumption of 0.56 pound fuel per horsepower-hour at an equivalent compressor speed of 290 rps and a temperature ratio of 4.0 for the gas generator.\n\nspeed of the power turbine in the same manner as the example.\n\nThe operation of the compressor in this gas turbine is shown in figure 25 by the operating curve of the compressor. The operating curve approaches the surge line of the compressor at the highest speed, so that the performance curves of the gas turbine will show a decrease in engine power and efficiency as the speed and the temperature ratio increase over approximately 300 rps and approximately 4.2, respectively. There is no means in this case for obtaining higher output from the power turbine; however, additional engine power might be obtained at higher engine speeds and temperature ratio by bleeding some of the gas-generator discharge gas for application in a power jet or by redesigning the power turbine for a higher flow capacity.\n\n<!-- Image (104, 178, 474, 603) -->\n\nFIGURE 24.—Over-all performance characteristics of gas turbine.\n\nThe ineffectiveness of the engine at low speeds is shown by the fact that at a speed of 180 rps, which is 62 percent of the full speed, the power developed is only 3 percent of full power. This condition indicates difficulty in starting the engine but engine power increases rapidly with fuel input for temperature ratios more than 2.3. A slight decrease in specific fuel consumption from the values in figure 24 is attained from the exhaust jet power because the pressure ratio of the nozzle is equal to the rampressure ratio. Inasmuch as the speed of the power turbine did not enter into the solution of this problem, the method of solution is always applicable for gas turbines whether or not the power turbine is connected with a gas-generator shaft, provided that the gas-flow pressure-ratio characteristics are independent of operating speed. If the turbine speed does influence the gas flow, the engine characteristics may be evaluated for each\n\n<!-- Image (546, 295, 936, 538) -->\n\nFIGURE 25.—Estimated compressor performance in gas-turbine engine.\n\nSUMMARY OF RESULTS\n\nThe principal results of the experimental study of operation and interaction of the components of the turbojet engine and the influence of the components on over-all engine performance are summarized as follows:\n\n1. The one-dimensional theory of combustion-chamber loss was inadequate to correlate the data obtained in the burner of the engine. This inadequacy was probably due to the simultaneous diffusion and burning in the combustion chamber of the engine.\n2. The method of correlating combustion efficiency of the burner was adequate because the variables used correlated the data to curves that followed the expected trend in the practical range of combustion temperatures.\n3. After corrections were made for tip leakage, the gas-flow pressure-ratio characteristics of the turbine obtained from cold-air component runs could be checked with the results obtained from the turbojet engine.\n4. As predicted from the component test with cold air, the range of turbine characteristics obtainable from turbojet-engine experiments was insufficient to establish the turbine-efficiency characteristics or to indicate how the turbine operating conditions fit into these characteristics. Resetting of the compressor stator blades satisfactorily extended the range.\n```", "timestamp": "2026-07-22T05:59:31.259592+00:00"}
{"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 75, "total_pages": 78, "image_filename": "19930082618_p75.jpg", "text": "```markdown\nNACA TN 1945\n73\n\n<!-- Image (163, 109, 874, 856) -->\n\nFigure 20.- Variation with Reynolds number of maximum section lift coefficient for four NACA 64-series airfoils of 0.12c thickness and various camber.\n```", "timestamp": "2026-07-22T05:59:31.514574+00:00"}
{"citation_id": "19930085972", "source_url": "https://ntrs.nasa.gov/api/citations/19930085972/downloads/19930085972.pdf", "page_number": 43, "total_pages": 46, "image_filename": "19930085972_p43.jpg", "text": "NACA RM L9B18\n41\n\nDownwash angle, $\\epsilon$, deg\nAngle of attack, $\\alpha$, deg\n\nTail-off lift-curve slope $(\\frac{dC_L}{d\\alpha})_o$\nHorizontal-tail effectiveness, $\\frac{dC_m}{dC_L}$\n\nNeutral-point and tail-off aerodynamic center location, $n_o$ and $n'_o$, percent $c$ ($\\Lambda=0$)\nLift coefficient, $C_L$\n\nDownwash gradient, $\\frac{d\\epsilon}{d\\alpha}$\nTail-off lift coefficient, $C_{L_o}$\n\nFigure 13.- Longitudinal-stability parameters of a variable-sweep model with and without faired wing cutout. $\\Lambda = 30^\\circ$.", "timestamp": "2026-07-22T05:59:36.241530+00:00"}
{"citation_id": "19930085962", "source_url": "https://ntrs.nasa.gov/api/citations/19930085962/downloads/19930085962.pdf", "page_number": 47, "total_pages": 51, "image_filename": "19930085962_p47.jpg", "text": "46\nCONFIDENTIAL\nNACA RM A9E05\n\nLift coefficient, $C_L$\nMach number, $M$\n(b) $\\alpha_{approx}$, $2^\\circ$.\n\n| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | 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{"citation_id": "19930085938", "source_url": "https://ntrs.nasa.gov/api/citations/19930085938/downloads/19930085938.pdf", "page_number": 42, "total_pages": 42, "image_filename": "19930085938_p42.jpg", "text": "```markdown\nNACA RM No. L9D04\n\n320,000\n\n280,000\n\n240,000\n\n200,000\n\n160,000\n\n120,000\n\n80,000\n\n40,000\n\n0\n\nRighting moment, lb - ft\n\nU. S. Navy requirement for righting moment with\nwing tip float submerged at 6° angle of heel\n\nOne boom leaves water\n\nWing tip would\nenter water\n\n0 2 4 6 8 10 12 14 16 18 20 22 24 26 28\nAngle of heel, deg\n\nNACA\n\nFigure 17.- Righting moment of twin-boom configuration without tip floats.\n\nNACA - Langley Field, Va.\n\n41\n```", "timestamp": "2026-07-22T05:59:47.482432+00:00"}
{"citation_id": "19930093769", "source_url": "https://ntrs.nasa.gov/api/citations/19930093769/downloads/19930093769.pdf", "page_number": 19, "total_pages": 39, "image_filename": "19930093769_p19.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:59:49.733469+00:00"}
{"citation_id": "19930086076", "source_url": "https://ntrs.nasa.gov/api/citations/19930086076/downloads/19930086076.pdf", "page_number": 37, "total_pages": 50, "image_filename": "19930086076_p37.jpg", "text": "NACA RM E9F09\n35\n\n[Figure: Cutaway view of a flame holder showing internal components with significant deposits or damage. A scale bar labeled \"INCHES\" is visible in the lower left corner.]\n\nNACA\nC-22314\n9-29-48\n\nFigure 13. - Cutaway view of flame holder 10 after 5 minutes of operation.", "timestamp": "2026-07-22T05:59:52.276090+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 70, "total_pages": 96, "image_filename": "19930085880_p70.jpg", "text": "68\nNACA RM No. 19C03\n\n[Figure: A graph plotting Trimming moment against Wetted area for various speeds.]\n\nTrimming moment, lb-ft\nSpeed (fps)\n30\n25\n20\n15\n10\n\nWetted area, sq ft\n(d) $\\tau = 160$.\nFigure 20.- Continued.\nNACA", "timestamp": "2026-07-22T05:59:58.934085+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 19, "total_pages": 78, "image_filename": "19930082483_p19.jpg", "text": "NACA TN No. 1807\n\nThe power produced by the turbine was absorbed by an eddy-current-type absorption dynamometer (fig. 6).\n\nMeasuring Apparatus\n\nWeight-flow measurements. - A standard A.S.M.E. flat-plate submerged orifice in the 12-inch combustion-air duct upstream of the hot-gas producer was used to measure the air flow. A calibrated rotameter was used to measure the quantity of fuel burned to heat the inlet air in the external hot-gas producer.\n\nThe turbine instrumentation used for these studies is shown in figure 7.\n\nNozzle-inlet section. - Four quadruple-shielded chromel-alumel thermocouples were equally spaced around the annular nozzle approach section 6 inches in front of the nozzle section to measure the turbine-inlet total temperatures. These thermocouples were sufficiently removed from the flame region to avoid radiation effects.\n\nFour static-pressure taps in the outer shell of the combustion chamber in the same cross-sectional plane as the inlet-temperature thermocouples measured the turbine-inlet static pressures. An arithmetic average of these readings was taken as the inlet static pressure. For the runs with reduced nozzle arc, only the static-pressure taps directly upstream of the active nozzles were used.\n\nA total-pressure probe located in the nozzle approach section about 2 inches upstream of the active nozzles was used to set inlet total pressures during turbine operation.\n\nNo instrumentation was installed in the flow region between the nozzles and the rotor blading.\n\nTurbine-rotor section. - A thermocouple and static-pressure tap each were located in the space on both sides of the disk in order to obtain measurements used to calculate the density of the gases surrounding the rotor disk. These observed quantities were considered representative of the stagnation conditions of the gases in these locations. This instrumentation was installed only for the runs with $120^\\circ$ admission and the data obtained were applied to the full-admission runs. Little change of the gas state at these locations is likely to occur because of varying the amount of gas admission.", "timestamp": "2026-07-22T06:00:00.989095+00:00"}
{"citation_id": "19930093789", "source_url": "https://ntrs.nasa.gov/api/citations/19930093789/downloads/19930093789.pdf", "page_number": 23, "total_pages": 29, "image_filename": "19930093789_p23.jpg", "text": "CONFIDENTIAL\n\nTo remove dust particles\nElectrostatic precipitator\n12-in. I.D. pipe\n7-in. flat-plate orifice\nAir at room temperature and pressure\nTemperature-control valves\nSteam air heater\nSurge tank, 24-in. I.D.\nTail pipe, 20-in. I.D.\nPlenum chamber\nPlenum chamber\nWater brake\nTurbine wheel\nPressure-ratio control valve\nSurge tank\nTo low-pressure exhaust system\nNACA\n\nFigure 9. - Diagrammatic sketch of experimental-equipment arrangement showing pressure-ratio control valve, air-flow path, and auxiliary equipment.\n\n22\nCONFIDENTIAL\nNACA RM No. E8I21", "timestamp": "2026-07-22T06:00:03.395885+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 32, "total_pages": 114, "image_filename": "19930086061_p32.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:00:03.773699+00:00"}
{"citation_id": "19930082472", "source_url": "https://ntrs.nasa.gov/api/citations/19930082472/downloads/19930082472.pdf", "page_number": 26, "total_pages": 34, "image_filename": "19930082472_p26.jpg", "text": "24\nNACA TN No. 1797\n\nUnflagged symbols indicate\nupper surface.\n\nFlagged symbols indicate\nlower surface.\n\nSpanwise\nstation, 2y/b\n20.9%\n28.1%\n41.7%\n57.4%\n71.4%\n85.0%\n92.5%\n96.2%\n\nPressure coefficient, P\n-12\n-8\n-4\n0\n4\n8\n\nChordwise station, x/c\n2\n4\n6\n8\n10\n\nNACA\n\n(c) $\\alpha=94^\\circ$\n\nFigure 5—Continued.", "timestamp": "2026-07-22T06:00:05.711844+00:00"}
{"citation_id": "19930082566", "source_url": "https://ntrs.nasa.gov/api/citations/19930082566/downloads/19930082566.pdf", "page_number": 6, "total_pages": 44, "image_filename": "19930082566_p6.jpg", "text": "4\nNACA TN No. 1889\n\nTEST PROCEDURE\n\nTest Specimen\n\nThe fatigue test specimens were machined from tubular sections about $16\\frac{1}{2}$ inches long, with an inside diameter of 2 inches and a wall thickness of $\\frac{1}{4}$ inch. The finished specimen is shown in figure 3 and has an over-all length of 16 inches. The other dimensions of the specimen are shown in figure 3. The inner walls of the specimens were left in the as-extruded form while the outer surfaces were polished circumferentially to a 9/0 finish with metallurgical abrasive paper. The wall thickness was measured to 0.0001 inch by a special apparatus as described in reference 2. The wall-thickness values were measured at five equal intervals along the tube length and six readings were taken around the circumference at each interval. The outside diameters at each interval in two perpendicular directions were also measured. The ratio of wall thickness to diameter of the specimen was 0.025, so that the stresses throughout the wall were essentially uniform. The circumferential elastic stress produced by internal pressure, as calculated assuming a thin wall and uniform stress distribution, is about 3 percent less than the exact value, while the axial stress, calculated assuming a thin wall, is about 2 percent more than the exact value. The ratio of diameter to reduced length of the specimen is about 0.25, thereby providing a sufficiently long section of the specimen free from the bending stresses produced by end restraints.\n\nFor a thin-walled tubular specimen subjected to an axial tensile load P and an internal pressure of p psi, the longitudinal and circumferential stresses are, respectively,\n\n$$\n\\sigma_1 = \\frac{P}{A} + \\frac{pd}{4t} = \\frac{P}{\\pi dt} + \\frac{pd}{4t} \\quad (1)\n$$\n\n$$\n\\sigma_2 = \\frac{pd}{2t} \\quad (2)\n$$\n\nwhere\n\n| | |\n| :--- | :--- |\n| A | cross-sectional area of tube |\n| d | internal diameter of specimen |\n| t | tube-wall thickness |", "timestamp": "2026-07-22T06:00:06.559135+00:00"}
{"citation_id": "19930082592", "source_url": "https://ntrs.nasa.gov/api/citations/19930082592/downloads/19930082592.pdf", "page_number": 5, "total_pages": 50, "image_filename": "19930082592_p5.jpg", "text": "4\nNACA TN 1914\n\nIn order to determine the oxide compositions, the oxidized surfaces were scraped to provide powders for use in X-ray-diffraction analysis. Analyses were made of oxide samples taken at different depths through the oxide layer in an attempt to determine changes in composition within the oxidation layer.\n\n## RESULTS AND DISCUSSION\n\n### Oxide Penetration\n\nThe results of the oxidation determinations are presented in figures 1 to 6 as plots of depth of oxidation penetration plotted against oxidation time, slope of the penetration-time curve (hereinafter called oxidation-rate constant) plotted against the reciprocal of the absolute temperature, and oxidation-rate constant plotted against percentage of molybdenum. Instead of fairing in the lines by the method of least squares, all experimental factors were considered and points judged most reliable were given the greatest weight. The contaminant, tungsten carbide, introduced during fabrication probably is a contributing factor, with other processing variables, to the scatter of data points evident in figures 1, 4, and 6. The small and variable percentage of tungsten carbide precludes any attempt to determine the effect of the tungsten carbide present on the oxidation penetration or oxidation mechanism.\n\nThe oxidation-penetration - time curves for molybdenum cermets are presented in figure 1. Figure 1(b) is an enlarged view of the area on figure 1(a) bounded by the origin, 0.02-inch penetration, and 18 hours. The depth of oxidation of these materials increases linearly with time, which indicates that the coating formed during oxidation does not slow the diffusion of oxygen to the oxidation interface. Increasing the percentage of molybdenum has a marked detrimental effect on the oxidation resistance.\n\nThe effect of change in temperature on the oxidation rate of molybdenum cermets agrees well with the exponential formula $K = e^{A/T}$ (where $K$ is the slope of the oxidation-penetration - time curve hereinafter designated the oxidation-rate constant, $T$ is the absolute temperature, and $A$ is a constant). For the molybdenum cermets $K = \\Delta p/\\Delta t$, for the tungsten cermets $K = \\Delta p/\\Delta \\log t$, and for the cobalt cermets $K = \\Delta \\log p/\\Delta \\log t$, (where $p$ is oxide penetration in in. and $t$ is time in hr). The numerical values of $K$ cannot be used as a basis for comparing molybdenum, tungsten, and cobalt cermets.", "timestamp": "2026-07-22T06:00:06.754415+00:00"}
{"citation_id": "19930082613", "source_url": "https://ntrs.nasa.gov/api/citations/19930082613/downloads/19930082613.pdf", "page_number": 3, "total_pages": 46, "image_filename": "19930082613_p3.jpg", "text": "2\nNACA TN 1958\n\nThe liners used in most American engines are made of Inconel, an alloy of the following nominal chemical composition (reference 1, p. 4):\n\n| Ni | Cr | Fe | Mn | Cu | Si | C | S |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| 78.5 | 14.0 | 6.5 | 0.25 | 0.2 | 0.25 | 0.08 | 0.015 |\n\nThis alloy is known to have good heat-resistant properties and is one of the best available for applications where flames directly contact the metal.\n\nAlthough the liners used in this investigation are not of the latest design, they are considered typical of liners made with louver flaps or fins protruding into the combustion zone to introduce air for cooling purposes. Determinations of temperature conditions in liners have previously been made at the NACA Lewis laboratory. Failures were attributed to high local temperature gradients, which were as great as $700^\\circ$ F per inch. Most cracks occurred in the area of maximum temperature.\n\nThe investigation reported herein was conducted to determine the factors that contribute to failures in combustion-chamber liners. Such metallurgical conditions as green-rot (reference 2), sulfide penetration (reference 3), oxide penetration, and carburization were considered. Analysis of such physical conditions as the thermal stresses present in the liner was found to be necessary in completing this research.\n\nThe effects of heat treatments and of removing stress raisers from punched edges prior to operation of the liners were determined by running liners in engines and examining cracks that formed.\n\nAPPARATUS AND PROCEDURE\n\nLiners. - Liners from two types of turbojet engine were investigated. They are designated type A and type B (fig. 1). The type-A liners had cracked in service and had been removed from the engine for this reason. The history of these liners was unavailable, but it was known that an unleaded fuel, probably AN-F-32, was used in their operation. An \"as-fabricated\" type-A liner was also examined. Type-B liners were run in an engine of the type shown in figure 2 in a cyclic \"accelerated-life\" run under the following conditions:", "timestamp": "2026-07-22T06:00:12.534259+00:00"}
{"citation_id": "19930085962", "source_url": "https://ntrs.nasa.gov/api/citations/19930085962/downloads/19930085962.pdf", "page_number": 48, "total_pages": 51, "image_filename": "19930085962_p48.jpg", "text": "```markdown\nNACA RM A9B05\nCONFIDENTIAL\n\n1.2\n1.0\n.8\n.6\n.4\n.2\n0\n-.2\n-.4\n-.6\n\nLift coefficient, $C_L$\n\n$\\delta_e$\n$30^\\circ$\n$20^\\circ$\n$10^\\circ$\n$6^\\circ$\n$4^\\circ$\n$2^\\circ$\n$0^\\circ$\n$-2^\\circ$\n$-4^\\circ$\n$-6^\\circ$\n$-10^\\circ$\n$-20^\\circ$\n$-30^\\circ$\n\n0 .1 .2 .3 .4 .5 .6 .7 .8 .9 1.0\nMach number, M\n(c) $\\alpha_{approx.}$, $4^\\circ$.\n\nFigure 14—Continued.\n\n[NACA logo]\n\nCONFIDENTIAL\n47\n```", "timestamp": "2026-07-22T06:00:14.611904+00:00"}
{"citation_id": "19930082703", "source_url": "https://ntrs.nasa.gov/api/citations/19930082703/downloads/19930082703.pdf", "page_number": 1, "total_pages": 28, "image_filename": "19930082703_p1.jpg", "text": "NACA TN 1983\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\n\nTECHNICAL NOTE 1983\n\nLONGITUDINAL FLYING QUALITIES OF SEVERAL SINGLE-ROTOR\nHELICOPTERS IN FORWARD FLIGHT\n\nBy F. B. Gustafson, Kenneth B. Amer, C. R. Haig,\nand J. P. Reeder\n\nLangley Aeronautical Laboratory\nLangley Air Force Base, Va.\n\n[Stamp: PROPERTY OF ARCHILD ENGINEERING LIBRARY]\n\n[Logo: NACA]\n\nWashington\nNovember 1949", "timestamp": "2026-07-22T06:00:18.672642+00:00"}
{"citation_id": "19930082646", "source_url": "https://ntrs.nasa.gov/api/citations/19930082646/downloads/19930082646.pdf", "page_number": 3, "total_pages": 37, "image_filename": "19930082646_p3.jpg", "text": "```markdown\n2\nNACA TN 1980\n\nmodifications produced improvements in the over-all hydrodynamic charac-\nteristics of a hull having a basic length-beam ratio of 15. Warping\nthe forebody reduced the bow spray and reduced the vertical accelera-\ntions encountered during landings in waves. See reference 1.\nExtending the afterbody increased the spray in the propellers but\nreduced the spray at the horizontal tail and reduced the vertical and\nangular accelerations during landings in waves. See reference 2.\nThese results indicated that further improvement might be achieved in a\nsingle configuration incorporating both the warped forebody and the\nextended afterbody. The hydrodynamic characteristics of such a hull\nconfiguration were accordingly investigated to determine the extent of\nthe improvement to be expected as a result of the combination. These\ncharacteristics were determined by the same procedures used in the\ninvestigations described in references 1 and 2 and are compared\nwith those of the basic hull (references 3 and 4).\n\nThe model was assumed to be a $\\frac{1}{10}$-size powered dynamic model of a\ntwin-engine propeller-driven flying boat having a design gross weight\nof 75,000 pounds, a gross load coefficient of 5.88, a wing loading\nof 41 pounds per square foot, and a power loading for take-off\nof 11.5 pounds per brake horsepower. The hull, which was one of the\nlength-beam-ratio series under investigation at Langley tank no. 1,\nhad a basic ratio of 15. The hydrodynamic characteristics of the basic\nhull are described in references 3 and 4, and its aerodynamic character-\nistics are described in reference 5.\n\nSYMBOLS\n\n| | |\n| :--- | :--- |\n| $g$ | acceleration due to gravity (32.2) feet per second per second |\n| $n_v$ | vertical acceleration, $g$ units |\n| $a$ | angular acceleration, radians per second per second |\n| $V$ | carriage speed (approx. 95 percent of airspeed), feet per second |\n| $V_v$ | sinking speed, feet per second |\n| $\\gamma$ | flight-path angle, degrees |\n| $\\delta_e$ | elevator deflection, degrees |\n| $\\tau$ | trim (angle between forebody keel at step and horizontal), degrees |\n| $\\tau_L$ | landing trim, degrees |\n```", "timestamp": "2026-07-22T06:00:19.392313+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 92, "total_pages": 99, "image_filename": "19930082511_p92.jpg", "text": "90\n\n$$\n\\frac{h}{H}\n$$\n\n$$\n\\frac{\\xi_1}{1.0}\n$$\n\n$$\n.5\n$$\n\n$$\n\\text{NACA}\n$$\n\n$$\n0 \\quad .4 \\quad .8 \\quad 1.2 \\quad 1.6 \\quad 2.0 \\quad 2.4 \\quad 2.8 \\quad 3.2\n$$\n\nDistance from entrance, $\\xi$, units of tunnel height\n\nFigure 20.- Tunnel-induced horizontal velocity on axis of two-dimensional closed-open-closed tunnel having one exit lip, with vortex at two locations along axis. Length of lower free surface is 1.5 times tunnel height. Ordinate is\n\n$$\n\\left( \\frac{\\text{Induced horizontal velocity}}{V} \\right) \\frac{hV}{\\Gamma}, \\text{ or}\n$$\n\n$$\n\\left( \\frac{\\text{Induced horizontal velocity}}{V} \\right) \\frac{2h}{cc_1}.\n$$\n\nNACA TN No. 1826", "timestamp": "2026-07-22T06:00:20.041873+00:00"}
{"citation_id": "19930085972", "source_url": "https://ntrs.nasa.gov/api/citations/19930085972/downloads/19930085972.pdf", "page_number": 44, "total_pages": 46, "image_filename": "19930085972_p44.jpg", "text": "42\nNACA RM L9B18\n\n[Figure: Four graphs showing aerodynamic parameters. Top-left: Downwash angle vs. Angle of attack. Top-right: Tail-off lift-curve slope and Horizontal-tail effectiveness vs. Tail-off lift coefficient. Bottom-left: Neutral point and tail-off aerodynamic-center location vs. Lift coefficient. Bottom-right: Downwash gradient vs. Tail-off lift coefficient. Legend indicates Basic and Alternate tail positions.]\n\nFigure 14.- Longitudinal-stability parameters of a variable-sweep model with two horizontal-tail positions. $\\Lambda = 45^\\circ$.", "timestamp": "2026-07-22T06:00:20.838526+00:00"}
{"citation_id": "19930086097", "source_url": "https://ntrs.nasa.gov/api/citations/19930086097/downloads/19930086097.pdf", "page_number": 28, "total_pages": 36, "image_filename": "19930086097_p28.jpg", "text": "26\nCONFIDENTIAL\nNACA RM A9H11\n\n$$ \\frac{(\\Delta C_d)_t}{C_{di}} = 1 + \\frac{\\text{Drag of blunt-trailing-edge airfoil}}{\\text{Drag of double-wedge airfoil}} $$\n\n[Graph: Plot of drag ratio vs. Mach number ($M_\\infty$) for various $\\eta$ values. The x-axis ranges from 1 to 5. The y-axis ranges from 0 to 1.0. Curves are labeled for conditions $0 \\le \\eta < 0.6$, $0 \\le \\eta \\le 1.0$, $0.2 < \\eta \\le 1.0$, $\\eta = 1.0$ (two curves). A secondary y-axis on the right is labeled $P_b/P_{bv}$ with values 1.0, 0.75, 0.50, 0.25, 0. The NACA logo is present.]\n\nFigure 3.- Estimated reductions in profile drag for equal thickness ratio; $\\frac{t}{c} = 0.10$.\n\n$$ \\frac{(\\Delta C_d)_s}{C_{di}} = 1 + \\frac{\\text{Drag of blunt-trailing-edge airfoil}}{\\text{Drag of double-wedge airfoil}} $$\n\n[Graph: Plot of drag ratio vs. Mach number ($M_\\infty$) for various $\\eta$ values. The x-axis ranges from 1 to 5. The y-axis ranges from 0 to 1.0. Curves are labeled for conditions $0 \\le \\eta < 0.6$, $0.1 < \\eta < 0.7$, $0.5 < \\eta < 0.8$, $0.7 < \\eta < 0.9$, $\\eta = 1.0$. A secondary y-axis on the right is labeled $P_b/P_{bv}$ with values 1.0, 0.75, 0.50, 0.25, 0. The NACA logo is present.]\n\nFigure 4.- Estimated reductions in profile drag for equal section modulus; $\\frac{t}{c} = 0.10$.\n\nCONFIDENTIAL\n\n[Handwritten notes in bottom left corner:\n12.8\n12.10\n12.2]", "timestamp": "2026-07-22T06:00:23.032876+00:00"}
{"citation_id": "19930086076", "source_url": "https://ntrs.nasa.gov/api/citations/19930086076/downloads/19930086076.pdf", "page_number": 38, "total_pages": 50, "image_filename": "19930086076_p38.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:00:24.153719+00:00"}
{"citation_id": "19930093769", "source_url": "https://ntrs.nasa.gov/api/citations/19930093769/downloads/19930093769.pdf", "page_number": 20, "total_pages": 39, "image_filename": "19930093769_p20.jpg", "text": "NACA RM No. E8L10a\n\nCONFIDENTIAL\n\nAir inlet\nHoney-\ncomb\nScreen\nDetail A - Flexible connection\nFlexible diaphragm\nSlip joint\nForward baffle\nDetail A\nRear baffle\nThrust-measur-\ning diaphragm\nThrust platform\nCooling-air duct\nNACA\nBlow-out\npatches\nExhaust\noutlet\n\nFigure 2. - General arrangement of turbojet-engine installation in altitude test chamber.\n\nCONFIDENTIAL\n\n19", "timestamp": "2026-07-22T06:00:24.573870+00:00"}
{"citation_id": "19930082546", "source_url": "https://ntrs.nasa.gov/api/citations/19930082546/downloads/19930082546.pdf", "page_number": 10, "total_pages": 65, "image_filename": "19930082546_p10.jpg", "text": "NACA TN No. 1870\n\nThis expression for $p$ may be written in nondimensional form as\n\n$$\n\\frac{p}{\\rho (UD)^2} = \\frac{1}{4\\sqrt{2}\\pi^2} \\left\\{ \\int_0^{2\\pi} \\left( \\frac{C_n p^2 x}{S_\\theta^3} + \\frac{C_Q p^3 y \\sin \\theta}{R_\\theta S_\\theta^3} \\right) \\left[ \\cos(mB\\theta + kS_\\theta) + kS_\\theta \\sin(mB\\theta + kS_\\theta) \\right] d\\theta \\right\\}^2\n$$\n\n$$\n+ \\left\\{ \\int_0^{2\\pi} \\left( \\frac{C_n p^2 x}{S_\\theta^3} + \\frac{C_Q p^3 y \\sin \\theta}{R_\\theta S_\\theta^3} \\right) \\left[ kS_\\theta \\cos(mB\\theta + kS_\\theta) - \\sin(mB\\theta + kS_\\theta) \\right] d\\theta \\right\\}^{1/2}\n\\tag{2}\n$$\n\nwhere $p$ is the magnitude of the root-mean-square oscillating pressure of a given harmonic.\n\nThe quantity $\\frac{p}{\\rho U^2 D^2}$ is defined as the free-space oscillating pressure coefficient and is designated $C_p$. The total free-space oscillating pressure is given by the expression $\\bar{p} = \\sqrt{\\sum_{m=1}^{m=\\infty} p_{mB}^2}$ where $p$ for any $mB$ value is given by equation (2) and the total free-space oscillating pressure coefficient is defined as $C_{\\bar{p}} = \\frac{\\bar{p}}{\\rho U^2 D^2}$.", "timestamp": "2026-07-22T06:00:29.663750+00:00"}
{"citation_id": "19930091993", "source_url": "https://ntrs.nasa.gov/api/citations/19930091993/downloads/19930091993.pdf", "page_number": 19, "total_pages": 21, "image_filename": "19930091993_p19.jpg", "text": "ANALYSIS OF PERFORMANCE OF JET ENGINE FROM CHARACTERISTICS OF COMPONENTS 15\n\n5. The efficiency contours obtained from engine data differed slightly from those obtained from component data. Although the peak efficiency in each case reached the same maximum value, in the component experiments this maximum occurred at a speed of approximately 180 rps, whereas in the engine experiments this maximum occurred at a speed of 155 rps.\n\n6. Estimates of engine performance with realizable improvements in the bearings and the burners showed that improved matching of the compressor and turbine high-efficiency regions was attainable chiefly because of reduced bearing torque. The engine modifications resulted in improvements in engine performance of 5 and 15 percent in equivalent thrust and specific thrust, respectively, at 290 rps and temperature ratio of 3.0. At an equivalent speed of 290 rps and a temperature ratio of 4.0, the improvements were 9 and 21 percent in equivalent thrust and specific thrust, respectively. In some engines, improvement in component performance may possibly impair the matching of the components and result in no gain in over-all engine performance.\n\n7. For engine acceleration, the operation points of the compressor for a fixed speed and temperature ratio were changed very little from steady-state operation but the turbine pressure ratio increased and the exhaust pressure ratio decreased from values for steady-state engine operation.\n\n8. If the exhaust gases from the turbojet engine are used to drive a power turbine, a unit can be designed that will develop 1800 horsepower with a specific fuel consumption of 0.56 pound fuel per horsepower-hour at an equivalent compressor speed of 290 rps and a temperature ratio of 4.0 for the gas generator. Higher speeds and temperature ratios would result in a reduction of engine power and efficiency because of the low compressor efficiency for operation in the region of lower flow beyond the surge line.\n\nCONCLUSIONS\n\nThe following generally valid conclusions can be drawn from this engine study:\n\n1. A method of analysis has been developed for application to jet engines and gas-turbine engines, which clearly shows the operation of the compressor and turbine components under all conditions of engine operation and permits the estimation of the effect on engine performance of modifications in the performance characteristics of the engine components.\n\n2. Approximate estimates of engine performance can be made from component calibration of the compressor and the turbine utilizing cold air. The accuracy of prediction of the gas-flow pressure-ratio characteristic with hot gas from cold-air calibration is quite good if proper clearance-flow corrections are made. If the turbine operates choked or nearly choked, the operation of the compressor in the engine may then be accurately predicted, but the value of the exhaust pressure ratio will reflect any inaccuracy in the turbine efficiency.\n\n3. Complete evaluation of the turbine characteristics may not always be possible from experiments of the complete engine and elaboration of the experimental procedure may be necessary to extend the range of turbine operation.\n\n4. The incidence of compressor surge during engine acceleration may be determined in the case of choked turbine flow from the operation point of the compressor for constant speed at the same temperature ratio and speed as for the beginning of acceleration. If the turbine is not choked, this criterion will be safe because the compressor in acceleration will operate at a condition slightly more removed from the surge line.\n\nFLIGHT PROPULSION RESEARCH LABORATORY,\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS,\nCLEVELAND, OHIO, June 2, 1948.", "timestamp": "2026-07-22T06:00:30.805523+00:00"}
{"citation_id": "19930093789", "source_url": "https://ntrs.nasa.gov/api/citations/19930093789/downloads/19930093789.pdf", "page_number": 24, "total_pages": 29, "image_filename": "19930093789_p24.jpg", "text": "312-8658\n1031\n\nNACA RM No. EB121\n\nCONFIDENTIAL\n\nEntrance total-temperature thermocouples, $T_1'$\nAir flow\nScreen\nEntrance total-pressure tube, $p_1'$\nLabyrinth air passage\nTail pipe\nExit static-pressure taps, $p_3$\nOuter guide shell\nAir flow\nExit total-temperature thermocouple, $T_3'$\nPlenum chamber\nStraightening grid\nTurbine disks\nExit static-pressure taps, $p_3$\nInner guide shell\n\nCONFIDENTIAL\n\nNACA\n\nFigure 10. - Cross section through turbine center line (configuration 1) showing turbine arrangement and instrumentation.\n\n23", "timestamp": "2026-07-22T06:00:35.243098+00:00"}
{"citation_id": "19930082472", "source_url": "https://ntrs.nasa.gov/api/citations/19930082472/downloads/19930082472.pdf", "page_number": 27, "total_pages": 34, "image_filename": "19930082472_p27.jpg", "text": "NACA TN No. 1797\n25\n\nUnflagged symbols indicate\nupper surface.\nFlagged symbols indicate\nlower surface.\n\nSpanwise\nstation, 2y/b\n20.9%\n28.1%\n41.7%\n57.4%\n71.4%\n85.0%\n92.5%\n96.2%\n\nPressure coefficient, P\n-1.2\n-.8\n-.4\n0\n.4\n.8\n\nChordwise station, x/c\n0\n4\n6\n8\n10\n\n(d) $\\alpha=12.5^\\circ$\n\nFigure 5—Continued.\n\nNACA", "timestamp": "2026-07-22T06:00:36.795589+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 36, "total_pages": 98, "image_filename": "19930086073_p36.jpg", "text": "34\nNACA RM A9H04\n\n<!-- Image (97, 175, 885, 833) -->\n\nLift coefficient, $C_L$\nDrag coefficient, $C_D$\n\n$\\circ$ 0.0\n$\\square$ 12.1\nAngle of sideslip, $\\beta$, deg\n\n(b) $C_L$ vs $C_D$.\n\nFigure 6.— Continued.", "timestamp": "2026-07-22T06:00:51.423929+00:00"}
{"citation_id": "19930082613", "source_url": "https://ntrs.nasa.gov/api/citations/19930082613/downloads/19930082613.pdf", "page_number": 4, "total_pages": 46, "image_filename": "19930082613_p4.jpg", "text": "NACA TN 1938\n3\n\n<!-- Table (203, 109, 791, 301) -->\n\\begin{tabular}{|c|c|c|c|}\n\\hline\n\\multicolumn{2}{|c|}{Duration} & Rotor speed & Gas temperature at \\\\\n\\cline{1-2}\n(min) & (sec) & (rpm) & exhaust-cone outlet \\\\\n& & & ($^\\circ$F) \\\\\n\\hline\n4 & 30 & 4000 $\\pm$50 & 1100 max. \\\\\n\\hline\n0 & 15 & Acceleration to & 1450 $\\pm$50 \\\\\n& & 11,500 & \\\\\n\\hline\n15 & 0 & 11,500 $\\pm$50 & 1240 $\\pm$20 \\\\\n\\hline\n0 & 15 & Deceleration to & 1260 max. \\\\\n& & 4000 & \\\\\n\\hline\n\\end{tabular}\n\nThe sulfur content of AN-F-32 fuel purchased for the investi-\ngation varied from 0.011 to 0.044 percent.\n\nThree groups of type-B liners were examined. The first group\nwas run for 66 hours and 57 minutes in an engine used in cyclic tests.\nThe liners became so severely cracked that four of them were removed\nfor examination. The other two groups of liners were modified by\nheat-treating and mechanical finishing in attempts to reduce failures\nand were run for 16 hours and 40 minutes.\n\nMetallurgical examination. - The metallurgical examination pri-\nmarily consisted of visual and metallographic studies of type-A\nliners that failed in service and type-B liners that had been run\nin the accelerated-life determinations for 66 hours and 57 minutes.\n\nThese liners were visually examined for size, shape, and origin\nof cracks, for scale patterns that indicate temperature gradients in\nthe metal, for carbonaceous deposits, and for buckling or warping.\nA large number of cracks were metallographically studied at high\nmagnifications. Small sections that contained cracks were removed\nfrom the liner, mounted in plastic, and polished with special pre-\ncautions to preserve edges and scale formations. In general, three\nareas of the microspecimens were examined both etched and unetched:\n(1) cracks and adjacent areas, (2) punched edges, and (3) areas away\nfrom surfaces exposed to the gaseous atmospheres. The cracks were\nexamined for intercrystalline and transcrystalline characteristics,\nscale formations, and possible oxidant penetration of the metal\nprior to crack formation. The three areas previously mentioned, par-\nticularly the punched edges, were checked for evidence of mechanical\nworking, the presence of stress raisers in the form of tears, or\nfissures, microstructural abnormalities, precipitation of carbides\nin grain boundaries and within grains, carburization, and grain\nsize.", "timestamp": "2026-07-22T06:00:54.350610+00:00"}
{"citation_id": "19930085962", "source_url": "https://ntrs.nasa.gov/api/citations/19930085962/downloads/19930085962.pdf", "page_number": 49, "total_pages": 51, "image_filename": "19930085962_p49.jpg", "text": "48\nCONFIDENTIAL\nNACA RM A9E05\n\nLift coefficient, $C_L$\nMach number, $M$\n(d) $\\alpha_{approx.}$, $6^\\circ$.\n\nFigure 14—Continued.\n\n[Figure: A graph plotting Lift coefficient ($C_L$) on the y-axis (ranging from -0.4 to 1.4) against Mach number ($M$) on the x-axis (ranging from 0 to 1.0). The graph contains multiple curves representing different angles of attack ($\\delta_e$), labeled as $30^\\circ$, $20^\\circ$, $10^\\circ$, $6^\\circ$, $4^\\circ$, $2^\\circ$, $0^\\circ$, $-2^\\circ$, $-4^\\circ$, $-6^\\circ$, $-10^\\circ$, $-20^\\circ$, and $-30^\\circ$. The curves show data points marked with various symbols (diamonds, squares, triangles, circles). The NACA logo is visible in the bottom right corner of the plot area.]\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:00:54.438111+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 93, "total_pages": 99, "image_filename": "19930082511_p93.jpg", "text": "1.0\n$\\frac{v}{V} = \\frac{\\sin V}{V} = \\frac{2\\epsilon h}{c \\sigma_1^2}$\n.8\n.6\n.4\n.2\n0\n-.2\n0 .4 .8 1.2 1.6 2.0 2.4 2.8 3.2\nDistance from entrance, $\\Sigma$, units of tunnel height\n\n$\\frac{\\xi}{h}$\n.5\n1.0\n.5\n1.0\n.5\n1.0\n\nClosed-open tunnel\nSymmetrical closed-open-closed tunnel\nClosed-open-closed tunnel with one exit lip\n\nNACA\n\nFigure 21.- Comparison of tunnel-induced angles on axis for three types of two-dimensional tunnels. Length of open sections for the closed-open-closed tunnels is 1.5 times tunnel height.\n\nNACA TN No. 1826\n76", "timestamp": "2026-07-22T06:00:56.630063+00:00"}
{"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 76, "total_pages": 78, "image_filename": "19930082618_p76.jpg", "text": "74\nNACA TN 1945\n\n3.0\nO NACA 63$_2$-415\n$\\square$ NACA 64$_2$-415\n$\\diamond$ NACA 65$_2$-415\n$\\triangle$ NACA 66$_2$-415\n\n2.8\n2.6\n2.4\n2.2\n2.0\n1.8\n1.6\n1.4\n1.2\n1.0\n.8\n\nMaximum section lift coefficient, $\\sigma_{l_{max}}$\n\nAirfoils with .20c split flaps deflected 60°\nPlain airfoils\n\nFlagged symbols and broken lines denote airfoils with standard leading-edge roughness\n\nNACA\n\n.5 1.0 2.0 3.0 4.0 5.0 10.0 $\\times$ 10$^6$\nReynolds number, R\n\nFigure 21.- Variation with Reynolds number of maximum section lift coefficient for four NACA 6-series airfoils of 0.4 design lift coefficient, 0.15c thickness, and various thickness distribution.", "timestamp": "2026-07-22T06:00:59.890153+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 20, "total_pages": 78, "image_filename": "19930082483_p20.jpg", "text": "18\nNACA TN No. 1807\n\n**Turbine-discharge section.** - Six static-pressure taps located in the plane normal to the turbine axis, $2\\frac{1}{4}$ inches downstream of the discharge side of the rotor wheel, indicated the discharge static pressures. Three of the taps were installed in the outer cylindrical wall at $120^\\circ$ intervals and three in the inner conical piece at the same angular positions as the outer wall taps. For the runs with reduced nozzle arc, only the static-pressure taps directly downstream of the active nozzles were used.\n\n**Tail-cone discharge section.** - Instrumentation was installed at the downstream end of the tail cone in an effort to establish the turbine-outlet static and total gas states after final mixing of the discharge driving fluid with the inactive gases downstream of the blocked nozzles. Two thermocouple probes, two total-pressure tubes, and three static-pressure taps spaced at $120^\\circ$ intervals were located as shown in figure 7. The measurements obtained, however, indicated that the mixing of the outlet gases was incomplete at this station.\n\n**Miscellaneous instrumentation.** - Thermocouples to measure the temperature rise of the oil across the journal and journal-thrust bearings were installed in the oil passages into and out of the bearings.\n\nThe rate of lubricant flow was not measured during turbine operation. However, a mock-up of the lubricating system was later set up to obtain this flow rate. This mock-up consisted of the shaft and bearing assembly together with the supply and scavenge pumps and a rotameter installed in the oil-supply line.\n\nA chronometric tachometer, driven by an electric generator geared to the dynamometer-rotor shaft was used for measuring turbine rotational speeds.\n\nAn NACA balanced-diaphragm dynamometer-torque indicator (reference 5) was used for torque measurements.\n\nOPERATIONAL PROCEDURE\n\nPerformance\n\nFor this investigation, the inlet total temperature was maintained constant at $800^\\circ$ R during operation with both full and reduced admission. The range of the variable operating conditions over which data were obtained includes inlet total pressure,", "timestamp": "2026-07-22T06:01:05.229278+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 33, "total_pages": 114, "image_filename": "19930086061_p33.jpg", "text": "NACA RM L9J07\n\n[Figure: Three-quarter rear view of wing 3 mounted in tunnel; $\\psi = 0^\\circ$, $\\alpha = 10^\\circ$.]\n\n(b) Three-quarter rear view of wing 3 mounted in tunnel; $\\psi = 0^\\circ$, $\\alpha = 10^\\circ$.\n\nFigure 3.- Concluded.\n\nNACA\nL-59367\n\n29", "timestamp": "2026-07-22T06:01:09.312800+00:00"}
{"citation_id": "19930093789", "source_url": "https://ntrs.nasa.gov/api/citations/19930093789/downloads/19930093789.pdf", "page_number": 25, "total_pages": 29, "image_filename": "19930093789_p25.jpg", "text": "```markdown\n24\n\nCONFIDENTIAL\n\nEquivalent mean blade speed $U(a_0/a_1)$, ft/sec\n\nEquivalent turbine shaft work, $E(a_0/a_1)^2$, Btu/lb\n\nBrake efficiency, $\\eta$\n\nTotal-pressure ratio, $q_1/q_3$\n\nEquivalent mean blade speed/equivalent weight flow, $U_0/\\dot{W}a_0$, ft/lb\n\nFigure 11. - Over-all performance chart with 70°-cone-angle stator and labyrinth shroud (configuration 1).\n\nNACA\n\nNACA RM No. E8121\n\n1031\n\n```", "timestamp": "2026-07-22T06:01:18.555780+00:00"}
{"citation_id": "19930082472", "source_url": "https://ntrs.nasa.gov/api/citations/19930082472/downloads/19930082472.pdf", "page_number": 28, "total_pages": 34, "image_filename": "19930082472_p28.jpg", "text": "26\nNACA TN No. 1797\n\nUnflagged symbols indicate\nupper surface.\nFlagged symbols indicate\nlower surface.\n\nSpanwise\nstation, 2y/b\n20.9%\n28.1%\n41.7%\n57.4%\n71.4%\n85.0%\n92.5%\n96.2%\n\nPressure coefficient, P\n-16\n-12\n-8\n-4\n0\n4\n8\n\nChordwise station, x/c\n4\n6\n8\n10\n\n(a) $\\alpha=14.6^\\circ$\n\nNACA\n\nFigure 5—Continued.", "timestamp": "2026-07-22T06:01:21.777009+00:00"}
{"citation_id": "19930082566", "source_url": "https://ntrs.nasa.gov/api/citations/19930082566/downloads/19930082566.pdf", "page_number": 7, "total_pages": 44, "image_filename": "19930082566_p7.jpg", "text": "```markdown\nNACA TN No. 1889\n5\n\nThe tubular specimens are subjected to synchronous variable loads P and p which have maximum values P' and p', minimum values P''' and p''', and mean values P'' and p''. The maximum, minimum, and mean values of the principal stresses $\\sigma_1$ and $\\sigma_2$ are\n\n$$\n\\left.\n\\begin{aligned}\n\\sigma_1' &= \\frac{P'}{\\pi d t} + \\frac{p'd}{4t} \\\\\n\\sigma_1''' &= \\frac{P'''}{\\pi d t} + \\frac{p'''d}{4t} \\\\\n\\sigma_1'' &= \\frac{P''}{\\pi d t} + \\frac{p''d}{4t}\n\\end{aligned}\n\\right\\} \\quad (3)\n$$\n\n$$\n\\left.\n\\begin{aligned}\n\\sigma_2' &= \\frac{p'd}{2t} \\\\\n\\sigma_2''' &= \\frac{p'''d}{2t} \\\\\n\\sigma_2'' &= \\frac{p''d}{2t}\n\\end{aligned}\n\\right\\} \\quad (4)\n$$\n\nThe fatigue strength of a material when subjected to the stresses in equations (3) and (4) depends upon both the ratio of the minimum to maximum stresses and to the ratio of the principal stresses. Since a very large number of tests would be required to cover completely all possible stress combinations, it was necessary to restrict the test program to a consideration of the influence of the principal stress ratio $\\sigma_2'/\\sigma_1'$ only.\n\nIn this investigation the ratios of minimum to maximum stresses $\\sigma_1'''/\\sigma_1'$ and $\\sigma_2'''/\\sigma_2'$ varied from about 0.10 to 0.20.\n\n### Testing Machine\n\nA special testing machine was designed and built for applying the fluctuating internal pressures and axial loads referred to in the foregoing section. Figures 4 and 5 show schematic drawings of the testing machine, while figure 6 is a photograph of the complete machine. Figures 7 to 11 are close-up photographs of various parts showing details of construction and operation of the machine.\n\nThe tubular specimen S is subjected to an axial fluctuating load by the lever K (fig. 6). This lever is subjected to a fluctuating\n```", "timestamp": "2026-07-22T06:01:23.680598+00:00"}
{"citation_id": "19930086097", "source_url": "https://ntrs.nasa.gov/api/citations/19930086097/downloads/19930086097.pdf", "page_number": 29, "total_pages": 36, "image_filename": "19930086097_p29.jpg", "text": "NACA RM A9H11 CONFIDENTIAL 27\n\n[Figure: Diagram showing an airfoil with coordinate axes X and Y, velocity vector $U_\\infty$, angle of attack $\\alpha$, and angles $\\theta_l = -\\left(\\frac{dy}{dx}\\right)_l + \\alpha$ and $\\theta_u = \\left(\\frac{dy}{dx}\\right)_u - \\alpha$. Note: $\\theta$ measured positive for elements facing the oncoming wind. NACA logo present.]\n\nFigure 5.- Coordinates used in theoretical calculations.\n\n[Figure: Graph plotting \"Increase in $\\frac{dC_l}{d\\alpha}$, percent\" (y-axis, 0 to 30) against \"$M_\\infty$\" (x-axis, 1 to 5). Two curves are shown: one labeled $\\frac{h}{c} = 0.10$ and another labeled $\\frac{h}{c} = 0.05$. NACA logo present.]\n\nFigure 6.- Theoretical increase in lift-curve slope of blunt-trailing-edge airfoils as compared to sharp-trailing-edge airfoils.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:01:26.653157+00:00"}
{"citation_id": "19930086076", "source_url": "https://ntrs.nasa.gov/api/citations/19930086076/downloads/19930086076.pdf", "page_number": 39, "total_pages": 50, "image_filename": "19930086076_p39.jpg", "text": "NACA RM E9F09\n\nMolybdenum prism\n\n$1\\frac{1}{2}''$\n\n$1''$\n\n$1''$\n\n$\\frac{1}{4}''$\n\n$1''$\n\n$1\\frac{3}{4}''$\n\n$1\\frac{1}{2}''$\n\n$1\\frac{3}{4}''$\n\n$1\\frac{1}{2}''$\n\n$\\frac{1}{4}''$\n\n$1''$\n\n$\\frac{1}{3}''$\n\n$1''$\n\n$17\\frac{1}{2}''$\n\n$\\frac{1}{16}''$\n\n$2''$\n\n$2''$\n\n$8''$\n\n$4''$\n\nFigure 14. - Schematic diagram of flame holder 11.\n\nNACA\n\n37", "timestamp": "2026-07-22T06:01:27.621457+00:00"}
{"citation_id": "19930093769", "source_url": "https://ntrs.nasa.gov/api/citations/19930093769/downloads/19930093769.pdf", "page_number": 21, "total_pages": 39, "image_filename": "19930093769_p21.jpg", "text": "CONFIDENTIAL\n\n(a) Standard spark plug.\n\n(b) Modified spark plug.\n\n(c) Extended-electrode spark plug.\n\nFigure 3. - Spark plugs used in starting.\n\nNACA\n\n20\n\nCONFIDENTIAL\n\nNACA RM No. E8L10a\n\n218 - 1461\n\n0/0", "timestamp": "2026-07-22T06:01:30.250222+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 37, "total_pages": 98, "image_filename": "19930086073_p37.jpg", "text": "NACA RM A9H04\n35\n\n[Figure: A graph plotting Lift coefficient, $C_L$ vs Pitching-moment coefficient, $C_m$. The vertical axis ranges from 0 to 1.4. The horizontal axis ranges from 0 to -0.24. Two curves are plotted: one with circle markers and one with square markers. A NACA logo is present in the bottom right corner of the plot area.]\n\nLift coefficient, $C_L$\nPitching-moment coefficient, $C_m$\n\n| | |\n| :--- | :--- |\n| $\\circ$ | $\\square$ |\n| 0.0 | 12.1 |\n\nAngle of sideslip, $\\beta$, deg\n(c) $C_L$ vs $C_m$.\nFigure 6.— Continued.", "timestamp": "2026-07-22T06:01:36.516809+00:00"}
{"citation_id": "19930082592", "source_url": "https://ntrs.nasa.gov/api/citations/19930082592/downloads/19930082592.pdf", "page_number": 6, "total_pages": 50, "image_filename": "19930082592_p6.jpg", "text": "```markdown\nNACA TN 1914\n5\n\nThe agreement is satisfactory for all the molybdenum ceramals at\n$2000^\\circ$ and $1785^\\circ$ F and for the 5- and 10-percent molybdenum ceramals\nat $1625^\\circ$ F. The oxidation rate of the 20- and 30-percent molybdenum\nceramals at $1625^\\circ$ F is lower than would exist if the rate formula\nheld consistently. The curves are assumed to be consistent with the\noxidation-rate formula at the higher temperatures because of the\nparallelism of the curves representing various molybdenum contents\n(fig. 2). Figure 2 indicates that at $1785^\\circ$ F and at $2000^\\circ$ F increasing\nthe percentage of molybdenum results in a drastic increase in\noxidation-rate constant. The variation in oxidation-rate constant\nat $2000^\\circ$ F caused by variation in the percentage of molybdenum is\nshown in figure 3.\n\nThe oxidation-penetration-time characteristics of the tungsten\nceramals are presented in figure 4. The data may be presented as a\nseries of straight lines if plotted with oxidation penetration on a\nlinear scale and the time of oxidation on a logarithmic scale.\nBecause the use of semilog paper results in straight lines, the\noxidation product formed on the tungsten ceramals evidently has a\nbetter resistance to oxygen diffusion than that found on the molyb-\ndenum ceramal.\n\nIn general, the ceramals containing the high percentages of\ntungsten had low resistances to oxidation (fig. 4(d)). At a temper-\nature of $1625^\\circ$ F, the 20- and 30-percent-tungsten ceramals oxidized\nmuch more rapidly than the 5- and 10-percent-tungsten ceramals but\nat $2000^\\circ$ F the oxidation-penetration rates of the tungsten ceramals\nwere of the same order of magnitude.\n\nAn appreciable increase in oxidation-rate constant at a temper-\nature of $1625^\\circ$ F when the tungsten percentage is increased from 10\nto 20 percent is shown in figure 5. At $1785^\\circ$ F, the oxidation-rate\nconstants for the 20- and 30-percent-tungsten ceramals are less than\nat $1625^\\circ$ F whereas the oxidation-rate constants of the 5- and\n10-percent tungsten ceramals remain practically constant. At $2000^\\circ$ F,\nthe oxidation-rate constants for all tungsten compositions are higher\nthan for the two low temperatures investigated.\n\nThe nonconformity of the oxidation-rate constants of the tungsten\nceramals to the exponential rate formula indicates a possible change\nin oxidation mechanism as temperature and percentage of tungsten vary.\n\nThe oxidation-penetration and time data for the cobalt ceramals\nare presented in figure 6. The generation of straight lines with\nslopes less than 1 on log-log coordinates indicates that the scale\n```", "timestamp": "2026-07-22T06:01:38.621679+00:00"}
{"citation_id": "19930082703", "source_url": "https://ntrs.nasa.gov/api/citations/19930082703/downloads/19930082703.pdf", "page_number": 2, "total_pages": 28, "image_filename": "19930082703_p2.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:01:40.579306+00:00"}
{"citation_id": "19930082546", "source_url": "https://ntrs.nasa.gov/api/citations/19930082546/downloads/19930082546.pdf", "page_number": 11, "total_pages": 65, "image_filename": "19930082546_p11.jpg", "text": "```markdown\n10\nNACA TN No. 1870\n\nAPPARATUS AND METHODS\n\nStatic tests were conducted for the measurement and analysis of the free-space pressures near the tips of five different propeller models. Tests were made in the tip Mach number range 0.45 to 1.00 for two two-blade 48-inch-diameter round-tip propellers, a four-blade 48-inch-diameter round-tip propeller, a two-blade 47-inch-diameter square-tip propeller, and a two-blade 85-inch-diameter round-tip propeller and for various blade angles. Comparative studies were also made to determine the effects on free-space pressures of a flat vertical wall and a curved surface which simulate the fuselage position in the pressure field.\n\nPropeller models used are shown in figure 2. These were mounted in adjustable hubs to allow the blade angles to be changed manually. The 85-inch-diameter Clark Y propeller, the NACA 4-(3)(06.3)-06 propeller, the NACA 4-(5)(08)-03 propeller, and the square-tip propeller were all tested as two-blade configurations. The NACA 4-(5)(08)-03 propeller was also tested as a four-blade configuration. The square-tip propeller blade shown has a (5)(08)-03 airfoil section and its diameter is 47 inches. The NACA designations are descriptive of the propeller. Numbers in the first group represent the propeller diameter in feet. Numbers in the first parentheses represent the design lift coefficient in tenths at the 0.7 radius. Numbers in the second parentheses give the blade thickness at the 0.7 radius in percent chord. The last group of numbers gives the blade solidity which is defined as the ratio of a single blade width at the 0.7 radius to the circumference of a circle with the same radius. Blade-form curves for the above models are given in figure 3.\n\nThe test propellers were driven by a 200 horsepower water-cooled variable-speed electric motor. Power to the motor was measured by means of a wattmeter, and motor-efficiency charts were used to determine power to the propellers.\n\nRoot mean square oscillating pressures were measured by a Massa Laboratories sound pressure measurement system calibrated to read directly in dynes per square centimeter. Figure 4 shows the test arrangement for measuring free-space pressures. Because ground reflection is considered negligible for this particular setup, the pressures measured are essentially free-space pressures except in the cases where reflecting surfaces were purposely placed in the pressure field. All pressure quantities in this paper are considered to be free-space oscillating pressures unless otherwise stated.\n\nMeasurements were made at several known distances from the propeller on lines parallel to the axis of rotation and at the same height above ground. At all times the microphone was doubly shock mounted and when reflecting surfaces were used the microphone was mounted separately to keep vibrations reaching it at a minimum.\n```", "timestamp": "2026-07-22T06:01:40.921765+00:00"}
{"citation_id": "19930085962", "source_url": "https://ntrs.nasa.gov/api/citations/19930085962/downloads/19930085962.pdf", "page_number": 50, "total_pages": 51, "image_filename": "19930085962_p50.jpg", "text": "NACA RM A9E05 CONFIDENTIAL 49\n\nLift coefficient, $C_L$\n\n6°\n30°\n20°\n10°\n6°\n4°\n2°\n0°\n-2°\n-4°\n-6°\n-10°\n-30°\n-20°\n\nMach number, M\n(e) $\\alpha_{approx.}$, 8°.\n\nFigure 14.—Concluded.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:01:41.307013+00:00"}