PDF/ocr_dataset_1047.jsonl

{"citation_id": "19930093769", "source_url": "https://ntrs.nasa.gov/api/citations/19930093769/downloads/19930093769.pdf", "page_number": 7, "total_pages": 39, "image_filename": "19930093769_p7.jpg", "text": "6 CONFIDENTIAL NACA RM No. E3L10a\n\nComparisons of the corrected values of net thrust, tail-pipe gas temperature, specific fuel consumption based on net thrust, and combustion efficiency for the two fuels are presented at the various flight conditions in figures 4, 5, 6, and 7, respectively. At all flight conditions investigated, the corrected net thrust and corrected tail-pipe gas temperature were approximately the same for the two fuels, as shown in figures 4 and 5, respectively.\n\nThe specific-fuel-consumption curves presented in figure 6 were obtained from the faired curves of corrected net thrust and corrected fuel flow for each flight condition. The data are presented in this manner to eliminate the effect of scatter on the specific-fuel-consumption curves. At an altitude of 5000 feet, the corrected specific fuel consumption is approximately the same for both fuels (fig. 6(a)). For an altitude of 20,000 feet and a flight Mach number of 0.60, the specific fuel consumption based on net thrust is slightly greater for AN-F-58 fuel than for gasoline over the range of engine speeds investigated (fig. 6(b)). At flight Mach numbers of 0.85 and 1.00 and an altitude of 20,000 feet, the specific fuel consumption for the two fuels is approximately the same at all engine speeds (figs. 6(c) and 6(d)). The heating value of AN-F-58 fuel is approximately 1 percent lower than that of gasoline (table I) and thus introduces a corresponding difference in the specific fuel consumption.\n\nThe corrected specific fuel consumption at an altitude of 35,000 feet and a flight Mach number of 1.00 is shown to be slightly higher for AN-F-58 fuel than for gasoline at all engine speeds (fig. 6(e)). At an altitude of 50,000 feet and a flight Mach number of 0.85, the corrected specific fuel consumption with AN-F-58 fuel is considerably greater than with gasoline at the low engine speeds and slightly greater at high engine speeds, as shown in figure 6(f). At this same altitude for the maximum operable engine speed, the specific fuel consumption for AN-F-53 fuel was about 9 percent higher than for gasoline.\n\nThe combustion efficiency, which is presented in figure 7, is approximately the same for both fuels at an altitude of 5000 feet (fig. 7(a)). At an altitude of 20,000 feet, the combustion efficiency is slightly lower for AN-F-58 fuel than for gasoline at flight Mach numbers of 0.60 and 0.85 (figs. 7(b) and 7(c), respectively) and slightly higher at a flight Mach number of 1.00 (fig. 7(d)) over the entire range of engine speeds investigated. Although the trend of the combustion-efficiency curves for the two fuels with flight Mach number is similar to the trend of the specific-fuel-consumption curves, some differences exist in the percentage of change in the performance with the two fuels and\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:47:34.934719+00:00"}
{"citation_id": "19930082472", "source_url": "https://ntrs.nasa.gov/api/citations/19930082472/downloads/19930082472.pdf", "page_number": 7, "total_pages": 34, "image_filename": "19930082472_p7.jpg", "text": "NACA TN No. 1797\n\nThe results of the force tests are shown in figure 4 in the form of conventional lift, drag, and pitching-moment curves. The results of the pressure-distribution measurements are shown in figure 5.\n\nTo determine whether or not the pressure-distribution measurements accurately showed all forces acting on the wing, the pressures were mechanically integrated to obtain the total values of lift, drag,$^{1}$ and pitching-moment coefficients. The values thus obtained are shown in figure 4 for comparison with the force data. As can be seen, good agreement was obtained and hence it was concluded that the pressure distributions accurately showed the forces acting on the wing.\n\nThe boundary-layer measurements consisted of total-head surveys over the wing at various angles of attack. Calibrations of the rakes had indicated that no measurable errors in the total head were incurred until a flow angle greater than $15^\\circ$ was experienced. Therefore, in surveying the boundary layers, data were obtained only at flow angles less than $15^\\circ$. When flow angles greater than this were encountered the rakes were realigned with the flow.\n\nThe results of the boundary-layer measurements are shown in figure 6. The boundary-layer thicknesses shown are the heights above the surface of the wing at which the free-stream value of total head was obtained regardless of the direction of flow.\n\nDISCUSSION\n\nThe major influence of separation on the characteristics of swept wings is typified by the longitudinal characteristics of the swept-forward wing which are shown in figure 4. At a moderate angle of attack (starting at about $10^\\circ$), drag began to rise rapidly and the pitching moments abruptly became more negative (the aerodynamic center shifted aft to 43 percent of the mean aerodynamic chord). At a higher angle of attack (starting at about $15^\\circ$), drag began to rise even more rapidly, the slope of the lift curve began to decrease, and pitching moments suddenly became positive. (The aerodynamic\n\n$^{1}$Skin-friction drag, naturally, was not indicated by the pressure-distribution measurements. Therefore, to obtain the total-drag coefficients shown in the figure the minimum profile-drag coefficient obtained by the force tests was added to the drag coefficients obtained by integrating the pressure distributions.", "timestamp": "2026-07-22T05:47:37.799731+00:00"}
{"citation_id": "19930091993", "source_url": "https://ntrs.nasa.gov/api/citations/19930091993/downloads/19930091993.pdf", "page_number": 6, "total_pages": 21, "image_filename": "19930091993_p6.jpg", "text": "2\nREPORT 928—NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nalso shown. Both inner and outer walls of the combustion chamber were provided with cooling shields designed to allow air from the compressor to flow between the shields and the wall. In each burner, the fuel (propane gas) was sprayed from two bars upstream of a perforated plate, which extended across the flow area. Ignition tubes connected the combustion zones of the various burners.\n\nThe entire compressor, the diffuser, and the bearing supports were made of aluminum. Turbine rotors, rotor blades, and the first set of nozzle blades were made of Nimonic 80, second and third sets of stator blades of Vitallium, and remaining portions of the machine, which were subject to high temperatures, were made of Inconel.\n\nThe engine was designed to operate at a compressor pressure ratio of 3.36, a gas flow of 4.18 pounds per second, and a rotor speed of 13,010 rpm for a compressor-inlet temperature of 440° R and a compressor-inlet pressure of 739 pounds per square foot.\n\nINSTALLATION AND INSTRUMENTATION\n\nThe engine inlet was connected to the laboratory refrigerated-air supply through a large stagnation chamber and the engine outlet was connected to the altitude exhaust system. The stagnation chamber, the compressor, and the turbine were completely lagged. Straighteners and screens were installed in the stagnation chamber to give uniform gas flow at the compressor inlet.\n\nAir flow was measured in the inlet ducting of the engine with a calibrated variable orifice of $\\pm 1\\frac{1}{2}$-percent accuracy. Fuel flow was measured with a standard A.S.M.E. thin-plate orifice.\n\nThe location of other instrument stations is shown in figure 2. Three iron-constantan thermocouples and two static-pressure taps were located at the compressor inlet (station 1)\n\nwere accurate to 0.5 percent of the differential from 32° F. When the data were averaged, the total-pressure readings were weighted for local mass flow.\n\n<!-- Image (518, 122, 935, 374) -->\nFIGURE 3.—Burners and combustion-chamber structure.\n\nAt the inlet to the turbine (station 3), six thermocouples and six total-pressure tubes were installed. The thermocouples used at the turbine inlet were shielded as shown in figure 4. This construction shielded the thermocouple and provided ventilation. Conduction of heat away from the junction was reduced by bathing a section of the thermocouple tip in flowing hot gases. Because the data obtained from these thermocouples often showed gas-temperature variations as great as 800° F at the turbine inlet, an average of the six temperature readings was not considered accurate enough for use in calculating performance. The principal use of these thermocouples was to determine the engine operating conditions. Chromel-alumel thermocouples were used at this station and the thermocouples were read within 2 percent of the differential from room temperature.\n\n<!-- Image (518, 609, 935, 732) -->\nFIGURE 4.—Shielded-type thermocouple composed of chromel-alumel thermocouple wire and Inconel shield and sheath material.\n\n<!-- Image (64, 581, 488, 825) -->\nFIGURE 2.—Schematic section of engine showing measuring stations.\n\nin the stagnation chamber. Inasmuch as the velocity in this chamber was very low, the static pressures and the observed temperatures were considered to be equivalent to the stagnation conditions. At the compressor outlet (station 2) were located three total-pressure tubes, three thermocouples, and two static-pressure taps. The iron-constantan thermocouples were equipped with recovery heads. Temperature readings\n\nAt the turbine outlet (station 4), readings of total pressures and temperatures were taken downstream of the straightening vanes and between the struts supporting the streamline tail cone. At this station, no tangential velocity component was assumed to exist. Six total-pressure tubes and nine chromel-alumel thermocouples were used at this station. Although six of the thermocouples had recovery heads and three were of the shielded type used in the turbine inlet, no significant differences in readings of the two types were noted. Variations in temperature readings between", "timestamp": "2026-07-22T05:47:38.786091+00:00"}
{"citation_id": "19930082474", "source_url": "https://ntrs.nasa.gov/api/citations/19930082474/downloads/19930082474.pdf", "page_number": 8, "total_pages": 21, "image_filename": "19930082474_p8.jpg", "text": "```markdown\n6\nNACA TN No. 1799\n\nmade to function satisfactorily at low speeds will offer greater difficulty or may even become inadequate at these higher speeds. Note that a peak is shown at about 30 miles per hour. In this range, if the controls are fixed, the helicopter will soon nose up, slow down, and slide backwards with resulting yawing motions and control difficulties.\n\nObservations Particularly Concerning Hovering\n\nThus far only the forward-flight characteristics have been discussed. Hovering, of course, precedes and follows all forward flight and is the outstanding reason for the existence of helicopters. At the present time, however, the problems associated with hovering are more indefinite than the problems in forward flight; they tend to disappear with a little flight practice; and they do not affect the general utility of the helicopter to the extent that limitations placed on night and instrument flying do.\n\nOne of the problems which the trainee must overcome in a helicopter of this type and size is the high control sensitivity in roll or, in other words, the high rate of roll per inch of stick displacement. This sensitivity can lead to overcontrolling which results in a short-period, pilot-induced, lateral oscillation. It is caused, apparently, by the pilot's lag in removing control following response of the machine. The result can be likened to what occurs with an autopilot having improper follow-up. A point to be remembered is that with constant ratio of control-stick displacement to cyclic feathering the steady rolling velocity obtained will vary inversely as the diameter of the rotor, or, the smaller machine will roll faster. Thus, sensitivity becomes less of a problem with larger machines.\n\nThe forces the pilot encounters in deflecting the stick can accentuate or minimize his impression of the sensitivity. The pilot should first be able to trim steady forces to zero. He should also have a force gradient, or spring constant, opposing displacement of the stick in order that he can properly judge the control being applied. The control-force gradient centers the stick when it is released; therefore, the lag in the pilot's follow-up process and the effort required are reduced. With one set of blades on the subject machine the lateral gradient was satisfactory, but with other blades peculiar characteristics appeared. In some cases the initial force change with deflection was proper, but the force returned to zero or even reversed as rolling velocity developed. This characteristic is considered very undesirable by the pilot. Figure 5 illustrates the character of the lateral forces immediately following stick displacement for two different rotors. Rotor A illustrates the type of transient force variation considered unsatisfactory, while the force variation for rotor B was considered acceptable.\n```", "timestamp": "2026-07-22T05:47:43.092702+00:00"}
{"citation_id": "19930086092", "source_url": "https://ntrs.nasa.gov/api/citations/19930086092/downloads/19930086092.pdf", "page_number": 14, "total_pages": 28, "image_filename": "19930086092_p14.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:47:45.877355+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 50, "total_pages": 96, "image_filename": "19930085880_p50.jpg", "text": "48\nNACA RM No. L9C03\n\n18\n16\n14\n12\n10\n8\n6\n4\n2\n0\nTrimming moment, lb-ft\n\nSpeed\n(fps)\n30\n\n0 .05 .10 .15 .20 .25 .30 .35\nWetted area, sq ft\n(d) $\\tau = 16^\\circ$.\nFigure 16.- Continued.\n\n[Figure: A graph plotting Trimming moment (lb-ft) against Wetted area (sq ft). The y-axis ranges from 0 to 18. The x-axis ranges from 0 to .35. There are four curves representing different speeds (10, 15, 20, 25 fps). The curves show an increasing trend. Data points are marked with circles, squares, diamonds, and triangles. A NACA logo is present at the bottom right of the graph area.]", "timestamp": "2026-07-22T05:47:50.161205+00:00"}
{"citation_id": "19930085972", "source_url": "https://ntrs.nasa.gov/api/citations/19930085972/downloads/19930085972.pdf", "page_number": 27, "total_pages": 46, "image_filename": "19930085972_p27.jpg", "text": "NACA RM L9B18\n25\n\nPitching-moment coefficient, $C_m$\n.2\n0\n-.2\n-.4\n-.6\n\n$t$\n(deg)\n$\\nabla$ 0\n$\\diamond$ 3\n$\\circ$ 0\n$\\square$ 3\n$\\triangle$ 0\n$\\triangle$ 3\n\nNo cutout\ntail off\nFaired cutout\ntail off\n\nLongitudinal-force coefficient, $C_x$\n-.4\n-.3\n-.2\n-.1\n0\n\nAngle of attack, $\\alpha$, deg\n16\n8\n0\n-8\n\nLift coefficient, $C_L$\n-4\n0\n4\n8\n12\n16\n\n[Figure: NACA logo]\n\n(c) External airfoil flaps.\nFigure 7.- Continued.", "timestamp": "2026-07-22T05:47:50.887773+00:00"}
{"citation_id": "19930086076", "source_url": "https://ntrs.nasa.gov/api/citations/19930086076/downloads/19930086076.pdf", "page_number": 18, "total_pages": 50, "image_filename": "19930086076_p18.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:47:54.850727+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 3, "total_pages": 78, "image_filename": "19930082483_p3.jpg", "text": "NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nTECHNICAL NOTE No. 1807\n\nEFFECTS OF PARTIAL ADMISSION ON PERFORMANCE\nOF A GAS TURBINE\n\nBy Robert C. Kohl, Howard Z. Herzig\nand Warren J. Whitney\n\nSUMMARY\n\nThe effects of partial admission of the driving fluid on the performance of a representative full-admission, production-type gas turbine were investigated.\n\nIn successive studies, the turbine was operated with full ($360^\\circ$) peripheral admission and with admission confined to $120^\\circ$ and $180^\\circ$ of the nozzle annulus. The corrected power output and over-all efficiency for these configurations are compared. Significance of the turbine losses is considered and the types and amounts of loss encountered with full and partial admission are tabulated. The occurrence of turbine-blade vibration induced by fractional-arc admission is discussed.\n\nMethods are presented for prediction of over-all turbine efficiency and power output for any degree of partial admission. These predictions of power output are accurate within $\\pm 1$ percent of corresponding observed values over most of the speed range and within $\\pm 3$ percent at the highest speeds. Efficiency predictions are within 1 point over most of the speed range and within $\\pm 1\\frac{1}{2}$ points at the lower and higher ends of the speed range. These predictions indicate marked losses in efficiency at admission of less than $180^\\circ$.\n\nINTRODUCTION\n\nAs applied to turbines, the term \"partial admission\" refers to a configuration in which the driving fluid is admitted to only a fraction of the nozzle annulus. Partial admission has long been used in steam-turbine practice; however, little application of this technique is known to have been made to gas turbines that have relatively high pressure ratios per stage. Partial admission is one of the possible methods that could be used for obtaining", "timestamp": "2026-07-22T05:47:57.048700+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 18, "total_pages": 98, "image_filename": "19930086073_p18.jpg", "text": "16\n\nMax. diam.\n4.49\n\n25.00\n\nLocation of split flap\n\nM.A.C. = 16.37\n\n63.03°\n53.94°\n\nMoment center\n0.25 M.A.C.\n\n1.80\n12.28\n\n2.67\n8.70\n10.83\n\n17.60\n28.08\n\n24.56\n5.72\n\n16.36\n8.00\n6.00\n\nAll dimensions in feet\nunless otherwise noted\n\n63.43°\n\n1.76\n\n8.18\n\n56.16\n\nNACA\n\n(a) General arrangement of model.\n\nFigure 2.- Geometric details of model investigated.\n\nNACA RM A52H04", "timestamp": "2026-07-22T05:48:00.494276+00:00"}
{"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 60, "total_pages": 78, "image_filename": "19930082618_p60.jpg", "text": "```markdown\n2.8\n2.4\n2.0\n1.6\n1.2\n.8\n.4\n0\n-.4\n-.8\n-1.2\n-1.6\n-2.0\n-2.4\n-2.8\n-3.2\n-3.6\n-4.0\n-16 -12 -8 -4 0 4 8 12 16 20 24 28 32 36 40 44 48 52 56 60 64 68 72 76 80 84 88 92 96 100 104 108 112 116 120 124 128 132 136 140 144 148 152 156 160 164 168 172 176 180 184 188 192 196 200 204 208 212 216 220 224 228 232 236 240 244 248 252 256 260 264 268 272 276 280 284 288 292 296 300 304 308 312 316 320 324 328 332 336 340 344 348 352 356 360 364 368 372 376 380 384 388 392 396 400 404 408 412 416 420 424 428 432 436 440 444 448 452 456 460 464 468 472 476 480 484 488 492 496 500 504 508 512 516 520 524 528 532 536 540 544 548 552 556 560 564 568 572 576 580 584 588 592 596 600 604 608 612 616 620 624 628 632 636 640 644 648 652 656 660 664 668 672 676 680 684 688 692 696 700 704 708 712 716 720 724 728 732 736 740 744 748 752 756 760 764 768 772 776 780 784 788 792 796 800 804 808 812 816 820 824 828 832 836 840 844 848 852 856 860 864 868 872 876 880 884 888 892 896 900 904 908 912 916 920 924 928 932 936 940 944 948 952 956 960 964 968 972 976 980 984 988 992 996 1000 1004 1008 1012 1016 1020 1024 1028 1032 1036 1040 1044 1048 1052 1056 1060 1064 1068 1072 1076 1080 1084 1088 1092 1096 1100 1104 1108 1112 1116 1120 1124 1128 1132 1136 1140 1144 1148 1152 1156 1160 1164 1168 1172 1176 1180 1184 1188 1192 1196 1200 1204 1208 1212 1216 1220 1224 1228 1232 1236 1240 1244 1248 1252 1256 1260 1264 1268 1272 1276 1280 1284 1288 1292 1296 1300 1304 1308 1312 1316 1320 1324 1328 1332 1336 1340 1344 1348 1352 1356 1360 1364 1368 1372 1376 1380 1384 1388 1392 1396 1400 1404 1408 1412 1416 1420 1424 1428 1432 1436 1440 1444 1448 1452 1456 1460 1464 1468 1472 1476 1480 1484 1488 1492 1496 1500 1504 1508 1512 1516 1520 1524 1528 1532 1536 1540 1544 1548 1552 1556 1560 1564 1568 1572 1576 1580 1584 1588 1592 1596 1600 1604 1608 1612 1616 1620 1624 1628 1632 1636 1640 1644 1648 1652 1656 1660 1664 1668 1672 1676 1680 1684 1688 1692 1696 1700 1704 1708 1712 1716 1720 1724 1728 1732 1736 1740 1744 1748 1752 1756 1760 1764 1768 1772 1776 1780 1784 1788 1792 1796 1800 1804 1808 1812 1816 1820 1824 1828 1832 1836 1840 1844 1848 1852 1856 1860 1864 1868 1872 1876 1880 1884 1888 1892 1896 1900 1904 1908 1912 1916 1920 1924 1928 1932 1936 1940 1944 1948 1952 1956 1960 1964 1968 1972 1976 1980 1984 1988 1992 1996 2000 2004 2008 2012 2016 2020 2024 2028 2032 2036 2040 2044 2048 2052 2056 2060 2064 2068 2072 2076 2080 2084 2088 2092 2096 2100 2104 2108 2112 2116 2120 2124 2128 2132 2136 2140 2144 2148 2152 2156 2160 2164 2168 2172 2176 2180 2184 2188 2192 2196 2200 2204 2208 2212 2216 2220 2224 2228 2232 2236 2240 2244 2248 2252 2256 2260 2264 2268 2272 2276 2280 2284 2288 2292 2296 2300 2304 2308 2312 2316 2320 2324 2328 2332 2336 2340 2344 2348 2352 2356 2360 2364 2368 2372 2376 2380 2384 2388 2392 2396 2400 2404 2408 2412 2416 2420 2424 2428 2432 2436 2440 2444 2448 2452 2456 2460 2464 2468 2472 2476 2480 2484 2488 2492 2496 2500 2504 2508 2512 2516 2520 2524 2528 2532 2536 2540 2544 2548 2552 2556 2560 2564 2568 2572 2576 2580 2584 2588 2592 2596 2600 2604 2608 2612 2616 2620 2624 2628 2632 2636 2640 2644 2648 2652 2656 2660 2664 2668 2672 2676 2680 2684 2688 2692 2696 2700 2704 2708 2712 2716 2720 2724 2728 2732 2736 2740 2744 2748 2752 2756 2760 2764 2768 2772 2776 2780 2784 2788 2792 2796 2800 2804 2808 2812 2816 2820 2824 2828 2832 2836 2840 2844 2848 2852 2856 2860 2864 2868 2872 2876 2880 2884 2888 2892 2896 2900 2904 2908 2912 2916 2920 2924 2928 2932 2936 2940 2944 2948 2952 2956 2960 2964 2968 2972 2976 2980 2984 2988 2992 2996 3000 3004 3008 3012 3016 3020 3024 3028 3032 3036 3040 3044 3048 3052 3056 3060 3064 3068 3072 3076 3080 3084 3088 3092 3096 3100 3104 3108 3112 3116 3120 3124 3128 3132 3136 3140 3144 3148 3152 3156 3160 3164 3168 3172 3176 3180 3184 3188 3192 3196 3200 3204 3208 3212 3216 3220 3224 3228 3232 3236 3240 3244 3248 3252 3256 3260 3264 3268 3272 3276 3280 3284 3288 3292 3296 3300 3304 3308 3312 3316 3320 3324 3328 3332 3336 3340 3344 3348 3352 3356 3360 3364 3368 3372 3376 3380 3384 3388 3392 3396 3400 3404 3408 3412 3416 3420 3424 3", "timestamp": "2026-07-22T05:48:07.834335+00:00"}
{"citation_id": "19930093789", "source_url": "https://ntrs.nasa.gov/api/citations/19930093789/downloads/19930093789.pdf", "page_number": 7, "total_pages": 29, "image_filename": "19930093789_p7.jpg", "text": "6 CONFIDENTIAL NACA RM No. E8I21\n\nthe labyrinth stationary shroud; configuration 2, the $0^\\circ$-cone-angle stator and the labyrinth stationary shroud; and configuration 3, the $0^\\circ$-cone-angle stator and the cylindrical stationary shroud.\n\nAPPARATUS AND PROCEDURE\n\nExperimental Equipment\n\nThe arrangement of the experimental equipment is diagrammatically shown in figure 9. Room air enters the electrostatic precipitator where dust particles are removed and passes through a submerged flat-plate orifice, an automatically controlled steam-supplied air heater, a surge tank, and into a pair of ducts leading to the plenum chamber immediately upstream of the turbine stator. The air-flow path from the plenum chamber through the turbine is shown in figure 10 with the labyrinth shroud and the $70^\\circ$-cone-angle stator of configuration 1 installed in the turbine. Air entering the plenum chamber passes through an 18-mesh, 27-gage screen, turns, and flows through a straightening grid. The air then passes through the stator and the rotor and leaves the turbine through the annular passage between the two exhaust-guide shells coaxially mounted in the tail pipe. The space between the outer guide shell and the tail pipe is provided to carry off the sealing air supplied to the labyrinth.\n\nDownstream of the turbine, the air flows through a large surge tank into a low-pressure exhaust system (fig. 9). The air pressure drop across the turbine is controlled by a butterfly valve downstream of the second surge tank. A 300-horsepower water brake, cradle-mounted for torque measurements, was used for power absorption.\n\nInstrumentation\n\nThe location of the turbine instrumentation is shown in figure 10. Entrance total pressure $p_1'$ was indicated by a total-pressure tube 3/4 inch upstream of the stator blades; the entrance total temperature $T_1'$ was measured with thermocouples at four stations within the pair of ducts leading to the plenum chamber. The exit static pressure $P_3$ was measured by six wall taps located in the exhaust-guide shells downstream of the rotor. The exit total temperature $T_3'$ was measured by three total-temperature thermocouples at the downstream end of the two exhaust-guide shells.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:48:09.694483+00:00"}
{"citation_id": "19930082472", "source_url": "https://ntrs.nasa.gov/api/citations/19930082472/downloads/19930082472.pdf", "page_number": 8, "total_pages": 34, "image_filename": "19930082472_p8.jpg", "text": "6\nNACA TN No. 1797\n\ncenter shifted forward to 5 percent of the mean aerodynamic chord\nforward of the leading edge.) These irregularities, typical of\nswept-wing characteristics, have generally been attributed to the\neffects of separation. More exact information of the effects of\nseparation is required, however, to intelligently approach the\nproblems of improving the unsatisfactory characteristics of swept\nwings operating in the moderate- and high-lift range. In the\nfollowing sections, pressure distributions, tuft studies, and\nboundary-layer measurements that were obtained on the 45° swept-\nforward wing are used to more precisely define the interrelations\nbetween separation phenomena and the longitudinal characteristics\nof the wing. For convenience, the characteristics in the moderate-\nlift range and the characteristics in the high-lift range are\nconsidered separately.\n\nThe Moderate-Lift Range\n\nThe effects of separation were first evidenced at a lift\ncoefficient of about 0.55 corresponding to an angle of attack of\nabout 10°. Drag began to rise rapidly while pitching moments became\nabruptly more negative. The reason for the change in the force\ncharacteristics can be seen in the pressure distributions over the\nstreamwise section at 20.9-percent semispan, summarized in the\nfollowing diagram:\n\n[Figure: Graph showing Chordwise pressure distributions at 20.9% semispan station. The vertical axis is Pressure coefficient, P, ranging from 0 to -2.4. The horizontal axis is Chordwise station, x/c, ranging from 0 to 1.0. The legend indicates $\\alpha$, angle of attack, with values 3.1°, 6.3°, 9.4°, 12.5°, and 16.6° corresponding to different line styles.]", "timestamp": "2026-07-22T05:48:09.975479+00:00"}
{"citation_id": "19930086097", "source_url": "https://ntrs.nasa.gov/api/citations/19930086097/downloads/19930086097.pdf", "page_number": 14, "total_pages": 36, "image_filename": "19930086097_p14.jpg", "text": "12 CONFIDENTIAL NACA RM A9H11\n\nthe increase in lift due to bluntness does not depend on the airfoil shape forward of the trailing edge.\n\nThis increase in lift may be quite appreciable, particularly at relatively high Mach numbers, as is indicated by the curves in figure 6. In this figure the increment $(C_2/C_1)(h/c)$ is plotted as a function of Mach number for two cases: $h/c=0.05$ and $h/c=0.10$, which represent, for example, fully blunt ($h/t=1$) airfoils of 5- and 10-percent thickness ratios, respectively. From the curves in figure 6, it can be seen that at a Mach number of 5, for example, the theoretical increase in lift-curve slope amounts to as much as 30 percent for a fully blunt airfoil of 10-percent-thickness ratio. Hence, at these relatively high Mach numbers the effect of trailing-edge bluntness on lift-curve slope can be of considerable practical importance.\n\nAs far as the section pitching-moment curve is concerned, an equally simple result is obtained using the second-order theory. The pitching moment about midchord is\n\n$$\nc_m = \\int_0^1 (P_u - P_l) \\left( \\frac{x}{c} - \\frac{1}{2} \\right) d\\left( \\frac{x}{c} \\right)\n\\tag{17}\n$$\n\nThe algebraic details of substituting equation (11) into (17) and integrating will be omitted, as they are the same as encountered in calculating the lift coefficient. The resulting expression, which applies to an arbitrary airfoil shape, is\n\n$$\n\\frac{dc_m}{d\\alpha} = \\frac{2C_2}{c^2} \\left( A - \\frac{hc}{2} \\right)\n\\tag{18}\n$$\n\nwhere $A$ is the cross-sectional area of the blunt-trailing-edge airfoil. Thus, the derivative $dc_m/d\\alpha$ is simply proportional to the difference between the cross-sectional area of the blunt-trailing-edge airfoil and the area of a simple wedge having the same trailing-edge thickness. Since most airfoils have a greater area than a simple wedge of the same base height, it follows from equation (18) that the effect of bluntness is to move the center of pressure closer to the midchord position. In the special case of zero thickness at the trailing edge the moment-curve slope is proportional to the airfoil cross-section area, as was pointed out in reference 7.\n\nMaximum Lift-Drag Ratio\n\nThe preceding analysis has shown that the minimum drag coefficient can be reduced by properly using bluntness at the trailing edge, and that by so doing the lift-curve slope always is slightly increased. Consequently, the accompanying change in maximum lift-drag ratio would\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:48:13.289753+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 19, "total_pages": 114, "image_filename": "19930086061_p19.jpg", "text": "NACA RM L9J07\n15\n\nThe effect of the flow on the over-all lift characteristics of the wings is illustrated by the increased section lift-curve slopes and the nonlinearity of the slopes along the span of each wing as given in the following table for $\\alpha = 0^\\circ$:\n\n| Station | $\\frac{y}{b/2}$ (percent) | Wing 1 | | Wing 2 | | Wing 3 | |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| | | R $\\times 10^6$ | $c_{l_\\alpha}$ | R $\\times 10^6$ | $c_{l_\\alpha}$ | R $\\times 10^6$ | $c_{l_\\alpha}$ |\n| 1 | 0 | 0.86 | 0.028 | 1.28 | 0.023 | 1.71 | 0.014 |\n| 2 | 16.7 | .71 | .030 | 1.06 | .026 | 1.43 | .015 |\n| 3 | 33.3 | .57 | .034 | .85 | .027 | 1.14 | .016 |\n| 4 | 50.0 | .43 | .044 | .64 | .028 | .86 | .017 |\n| 5 | 66.7 | .39 | .056 | .43 | .049 | .57 | .038 |\n| 6 | 83.3 | .14 | .094 | .21 | .081 | .29 | .071 |\n| 7 | 91.6 | .07 | .120 | .11 | .103 | .14 | .090 |\n\nThe scatter of the data and the insufficiency of low angle-of-attack data, except for wing 2, make the fairing of the curves of $c_l$ against $\\alpha$ in figures 34 to 36 and the determination of $c_{l_\\alpha}$ values somewhat questionable near zero lift. Nevertheless, the data of the preceding table are sufficiently reliable to show the trends of increased $c_{l_\\alpha}$ with increased aspect ratio of the related wings. Nonlinear lift curves would be expected from considerations of the varying three-dimensional vortex and boundary-layer flow.\n\nSPANWISE-LOAD DISTRIBUTIONS\n\nAt low angles of attack where the separated-flow region near the leading edge was small and sharp along the entire span, the lift over the wings was close to the theoretical lift and the spanwise-load distributions were approximately elliptical for the three wings. However, with increased angle of attack the distributions deviated from elliptical curves as an outboard dip and a hump farther inboard developed. The humps occurred at the spanwise locations where the region of separated flow covered a large extent of the chord and effectively gave the airfoil", "timestamp": "2026-07-22T05:48:15.767799+00:00"}
{"citation_id": "19930085977", "source_url": "https://ntrs.nasa.gov/api/citations/19930085977/downloads/19930085977.pdf", "page_number": 22, "total_pages": 33, "image_filename": "19930085977_p22.jpg", "text": "CONFIDENTIAL\n\nM\n1.18 Δ\n1.15 Δ\n1.10 ∇\n1.08 ∇\n1.05 ∇\n1.03 ∇\n1.00 ∇\n.98 Δ\n.95 Δ\n.93 Δ\n.90 Δ\n.88 ∇\n.85 Δ\n.80 Δ\n.70 □\n.60 □\n\nAngle of attack, $\\alpha$, deg\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n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"timestamp": "2026-07-22T05:48:18.855489+00:00"}
{"citation_id": "19930091987", "source_url": "https://ntrs.nasa.gov/api/citations/19930091987/downloads/19930091987.pdf", "page_number": 6, "total_pages": 12, "image_filename": "19930091987_p6.jpg", "text": "```markdown\n2\nREPORT 922—NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\n<!-- Image (93, 109, 483, 382) -->\n\nFIGURE 1.—Langley 24-inch high-speed tunnel. Schematic diagram illustrating tunnel shape and models.\n\nThe test section, originally circular (24-in. diameter), had been modified prior to the present investigation by the installation of flats on the tunnel walls. These flats reduced the width of the tunnel at the test section from 24 inches to 18 inches and changed the shape of the test section from circular to one more nearly approaching rectangular. The cross section of the tunnel at the model location is shown in figure 1.\n\nThe model for the infinite-aspect-ratio tests completely spanned the test section and passed through end plates fitted into the flat walls of the tunnel. The end plates, which accurately preserved the contours of the tunnel wall at the intersection of the tunnel wall and the model, provided clearances between model and tunnel wall and thereby permitted the forces acting on the model to be transmitted to and recorded by a 3-component balance to which the ends of the model were attached (fig. 2).\n\nThe models for tests of finite aspect ratios were installed and supported in the same manner as for tests of infinite aspect ratios except that the models extended one semispan into the air stream from the tunnel wall. This type of installation is satisfactory because the boundary layer on the tunnel wall, as previously discussed, is very thin. For the tests of wings having aspect ratios of 5 or less, two semispans, one from each wall, were installed for the purpose of doubling the magnitude of the forces to be measured by the standard balance of this tunnel, which was designed for larger force ranges than those encountered in these tests. For the wings of aspect ratios 5 and 7, tests were made with one and two semispan models mounted in the tunnel. The results of tests made with both one and two semispans mounted in the tunnel were in close agreement even without the tunnel-wall corrections.\n\nLift, drag, and pitching moment were measured on wings having rectangular plan forms and zero twist. The aspect\n\n<!-- Image (534, 73, 950, 268) -->\n\n(a) Over-all view with access door removed, showing model installation.\n\n<!-- Image (534, 305, 950, 523) -->\n\n(b) Downstream view with model in place.\n\nFIGURE 2.—Model mounting in test section of Langley 24-inch high-speed tunnel. Aspect ratio, 5; two-semispan installation.\n\nratios of the wings tested were $\\infty$, 9, 7, 5, 3, and 2. All the wings were of the NACA 0012 profile and had chords of 2 inches. Tests were made at angles of attack from $0^\\circ$ to $6^\\circ$ and at Mach numbers between 0.5 and the tunnel choked condition, that is, the maximum Mach number obtainable for a given model-tunnel combination. The Reynolds number range corresponding to this Mach number range is from $5.3 \\times 10^5$ to $7.6 \\times 10^5$.\n\nThe various factors affecting the accuracy of these data may, in general, be divided into two classes: accidental errors and systematic errors.\n\nThe accidental errors arose from inaccuracies in the calibrations of the balance and the static-pressure orifices and from design limitations on the maximum sensitivity of the balance. The maximum sensitivity appears to be the primary source of accidental errors and is a maximum for the small-area wing of aspect ratio 2 at low Mach numbers. At a Mach number of 0.50 for the wing of aspect ratio 2, the accidental errors in coefficients appear to be of the following order:\n\nWing lift coefficient, $C_L$ . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .", "timestamp": "2026-07-22T05:48:19.341856+00:00"}
{"citation_id": "19930085972", "source_url": "https://ntrs.nasa.gov/api/citations/19930085972/downloads/19930085972.pdf", "page_number": 28, "total_pages": 46, "image_filename": "19930085972_p28.jpg", "text": "26\nNACA RM L9B18\n\nPitching-moment coefficient, $C_m$\n0\n-.2\n-.4\n-.6\n-.8\n\n$i_t$ (deg)\n$\\Delta$ 0 No cutout\n$\\nabla$ 3\n$\\diamondsuit$ 0 Tail off\n$\\circ$ 0 Faired cutout\n$\\square$ 3\n$\\triangle$ 3 Tail off\n\nLongitudinal-force coefficient, $C_x$\n-.4\n-.3\n-.2\n-.1\n0\n\nAngle of attack, $\\alpha$, deg\n16\n8\n0\n-8\n\n-4 0 4 .8 1.2\nLift coefficient, $C_L$\n\n[NACA logo]\n\n(d) External airfoil flaps, alternate tail position.\nFigure 7.- Continued.", "timestamp": "2026-07-22T05:48:20.401289+00:00"}
{"citation_id": "19930085938", "source_url": "https://ntrs.nasa.gov/api/citations/19930085938/downloads/19930085938.pdf", "page_number": 26, "total_pages": 42, "image_filename": "19930085938_p26.jpg", "text": "NACA RM No. L9B04\n25\n\n[Figure: Bottom view of a twin-engine aircraft model mounted on a support structure in a wind tunnel.]\n\nNACA\nL-55733.1\n\n(D) Bottom view.\nFigure 5.- Concluded.", "timestamp": "2026-07-22T05:48:23.627863+00:00"}
{"citation_id": "19930086092", "source_url": "https://ntrs.nasa.gov/api/citations/19930086092/downloads/19930086092.pdf", "page_number": 15, "total_pages": 28, "image_filename": "19930086092_p15.jpg", "text": "NACA RM A9F14 CONFIDENTIAL 13\n\n[Figure: Diagram of an aircraft showing forces and moments. Labels include: $\\alpha$, $\\beta$, $V$, $C_n$, $C_L$, $C_Y$, $C_D$, $C_m$, $C_{H\\alpha}$, $\\delta_r$, and \"moment center\". A note states: \"Note: All forces, moments, and angles are shown positive.\"]\n\nFigure 1.—Standard NACA sign convention.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:48:23.777163+00:00"}
{"citation_id": "19930085914", "source_url": "https://ntrs.nasa.gov/api/citations/19930085914/downloads/19930085914.pdf", "page_number": 36, "total_pages": 42, "image_filename": "19930085914_p36.jpg", "text": "```markdown\nNACA RM A9D25\n35\n\n<!-- Image (167, 101, 814, 410) -->\n\n<!-- Image (167, 453, 814, 793) -->\n\nFigure 14.-The effects of wing roughness on the variation\nof lift-drag ratio with lift coefficient of the wing-\nfuselage combination at a Reynolds number of 2,000,000.\n```", "timestamp": "2026-07-22T05:48:25.398070+00:00"}
{"citation_id": "19930085962", "source_url": "https://ntrs.nasa.gov/api/citations/19930085962/downloads/19930085962.pdf", "page_number": 37, "total_pages": 51, "image_filename": "19930085962_p37.jpg", "text": "36 CONFIDENTIAL NACA RM A9E05\n\nLift coefficient, $C_L$\n\nElevator deflection, $\\delta_e$, deg\n\n(f) $M, 0.87$.\n\nFigure 12. — Continued.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:48:29.310513+00:00"}
{"citation_id": "19930086076", "source_url": "https://ntrs.nasa.gov/api/citations/19930086076/downloads/19930086076.pdf", "page_number": 19, "total_pages": 50, "image_filename": "19930086076_p19.jpg", "text": "NACA RM E9F09\n17\n\n<!-- Image (142, 201, 896, 409) -->\n\n<!-- Table (335, 446, 663, 697) -->\n\\begin{tabular}{|c|c|c|c|c|}\n\\hline\nFlame- & \\multicolumn{3}{c|}{Dimensions} & Total \\\\\nholder & a & b & c & gutters \\\\\n\\hline\n1 & $1\\frac{1}{2}$\" & $\\frac{3}{4}$\" & $2\\frac{1}{2}$\" & 22 \\\\\n\\hline\n2 & $2\\frac{1}{2}$\" & $\\frac{3}{4}$\" & $2\\frac{1}{2}$\" & 14 \\\\\n\\hline\n3 & $1\\frac{1}{16}$\" & $\\frac{3}{4}$\" & $2\\frac{1}{2}$\" & 30 \\\\\n\\hline\n4 & $1\\frac{1}{2}$\" & $\\frac{3}{4}$\" & 2\" & 22 \\\\\n\\hline\n5 & $1\\frac{1}{2}$\" & 1\" & $2\\frac{1}{2}$\" & 22 \\\\\n\\hline\n6 & \\multicolumn{3}{c|}{Same as flame} & \\\\\n& \\multicolumn{3}{c|}{holder 1 except for} & \\\\\n& \\multicolumn{3}{c|}{addition of third} & \\\\\n& \\multicolumn{3}{c|}{row of gutters in} & \\\\\n& \\multicolumn{3}{c|}{center} & 33 \\\\\n\\hline\n\\end{tabular}\n\nFigure 3. - Schematic diagram of various gutter-type flame holders.", "timestamp": "2026-07-22T05:48:30.324163+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 51, "total_pages": 96, "image_filename": "19930085880_p51.jpg", "text": "NACA RM No. L9C03\n49\n\nTrimming moment, lb-ft\nSpeed\n(fps)\n25\n20\n15\n10\n\nWetted area, sq ft\n(e) $\\tau = 20^\\circ$.\nFigure 16.- Concluded.\nNACA", "timestamp": "2026-07-22T05:48:31.365259+00:00"}
{"citation_id": "19930093769", "source_url": "https://ntrs.nasa.gov/api/citations/19930093769/downloads/19930093769.pdf", "page_number": 8, "total_pages": 39, "image_filename": "19930093769_p8.jpg", "text": "NACA RM No. E9L10a CONFIDENTIAL 7\n\nin the changes with engine speed. The combustion efficiency at an altitude of 35,000 feet and a Mach number of 1.00 (fig. 7(e)) is slightly higher for gasoline than for AN-F-58 over the entire range of engine speeds, although the difference is small at the maximum and minimum engine speeds. The combustion efficiency at an altitude of 50,000 feet and a flight Mach number of 0.85 (fig. 7(f)) is considerably lower with AN-F-58 fuel than for gasoline over the entire range of engine speeds investigated. The difference in combustion efficiency at maximum engine speed for the two fuels is about the same as the differences in specific fuel consumption previously noted.\n\nLow-Engine-Speed Blow-Out with AN-F-58 Fuel and Gasoline\n\nThe conditions at which combustion blow-out occurred were determined by gradually reducing the engine speed while holding the altitude and flight Mach number conditions constant. While the engine speed was being reduced, attempts were occasionally made to accelerate the engine in order to assure that dead-band operation was not being encountered. The occurrence of combustor blow-out was noted by a rapid reduction in engine speed and tail-pipe gas temperature. These blow-out experiments were conducted with AN-F-58 fuel (NACA fuel number 48-210) in engine A and, for comparison, with gasoline in engine B. The two different engines used for this investigation might result in different normal performance for the two fuels but no significant difference is expected in the altitude blow-out limits because both engines were of the same basic design.\n\nThe altitude blow-out limits of both fuels are presented in figure 8 in which the altitude is plotted against the engine speed at which blow-out occurred for flight Mach numbers of 0.25 and 0.60. The altitude blow-out limits of the two fuels are the same at a flight Mach number of 0.60. For a flight Mach number of 0.25 at altitudes above 42,500 feet, a lower engine speed could be attained before blow-out occurred with AN-F-58 fuel than with gasoline. For the altitudes investigated below 42,500 feet, the opposite results were obtained. The differences in the altitude limits are generally within the normal reproducibility of this type of data; from the limited range of these data, no significant differences in altitude limits for the two fuels could therefore be established.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:48:33.034177+00:00"}
{"citation_id": "19930082474", "source_url": "https://ntrs.nasa.gov/api/citations/19930082474/downloads/19930082474.pdf", "page_number": 9, "total_pages": 21, "image_filename": "19930082474_p9.jpg", "text": "NACA TN No. 1799\n\nThe longitudinal forces immediately following abrupt stick displacement differ in character from the lateral forces (fig. 6). In this case neither rotor A nor rotor B showed acceptable characteristics, although the pilot reported the characteristics of rotor A noticeably inferior to those of rotor B.\n\nIn another case abrupt stick motions were found to cause forces perpendicular to the direction of motion which tended to whirl the stick in the direction of rotor rotation. The stick would go to full deflection in a spiral motion if released. The forces for restraint of the stick became higher as the stick was moved more rapidly. Overcontrol results in this case because the pilot fights the forces.\n\nNo less important in promoting overcontrol is high control friction. Friction prevents accurate positioning of the control because of the extremely nonlinear force gradient it provides for small deflections and because the control tends to jump as static friction is broken. Friction also prevents self-centering of the control and consequently causes poor follow-up and an increase in the required pilot effort. The control difficulties imposed by high sensitivity and undesirable forces fortunately can be greatly lessened with relatively little practice. The control difficulties imposed by friction, however, always increase the demands on pilot effort and are hardest for the pilot to overcome in avoiding overcontrol.\n\nThe extrapolation of roll measurements to full control deflection indicates that the maximum rate of roll for this aircraft is as great as those of some modern fighter airplanes at the speeds for their maximum rates of roll. The high rate of roll achieved with the helicopter is apparently due to low damping and not to high control power, because the moments developed about the center of gravity are always relatively small. Computations of the damping indicate that it is a fraction of that for airplanes and could be expected to result in large amounts of continued roll following the centering of the controls from high rates of roll. In observations made at 40 miles per hour, however, where experiments with large rates of roll were convenient, no tendency to overshoot could be detected by the pilot. In hovering, both pilot observations and instrument measurements have indicated that the tendency to overshoot, while presumably present, is secondary to the effects of the stability with speed which results from the lateral motion acquired. Apparently, the lateral velocity can, in accordance with the details of the maneuver, either cancel or add to the tendency to overshoot.\n\nMany descriptions of the control response of this and similar helicopters in terms of lag have been made. The control lag, as defined by the time necessary for the rotor to reach a position corresponding to any specified stick position during steady motion of the controls, has been", "timestamp": "2026-07-22T05:48:45.492159+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 4, "total_pages": 78, "image_filename": "19930082483_p4.jpg", "text": "2\nNACA TN No. 1807\n\npart-load operation or power regulation. This method could utilize a system of individual gas generators, each generator discharging into a section of the turbine-nozzle periphery, the number of generators fired depending on the general turbine power output desired.\n\nAn analytical study of the manner in which normal turbine operating losses arise was made at the NACA Lewis laboratory and is presented herein. The additional losses introduced by use of partial admission are also discussed. This study further shows that the turbine performance at full and partial admission can be correlated on the basis of the loss differentials.\n\nIn order to determine quantitatively the effect of partial admission on turbine power output and turbine efficiency, a representative single-stage turbine was modified to restrict the driving fluid to $120^\\circ$ and $180^\\circ$ of the nozzle annulus. At $120^\\circ$ admission, turbine performance was obtained for three values of total-pressure ratio, a range of nozzle-inlet pressures from 20 to 45 inches of mercury absolute, an inlet temperature of $800^\\circ$ R, and over a range of rotor speeds that includes the design rotor speed as corrected to standard conditions. At $180^\\circ$ admission, turbine performance was obtained at a total-pressure ratio of 2.0, an inlet total pressure of 45 inches of mercury absolute, and an inlet temperature of $800^\\circ$ R over the complete range of corrected rotor speeds. Performance with full and partial admission at equivalent operating conditions is compared on the basis of corrected power output and over-all efficiencies. Because no provision was made for motoring the turbine, rotational losses were calculated using applicable loss formulas. Validity of these formulas was established by using actual loss measurements on a smaller, though similar, turbine.\n\nPartial admission is considered as a power-reduction device. The methods developed for estimating the losses were used to make a study of the effects of partial admission at various degrees of admission; the results of the study are summarized in a curve of power reduction plotted against efficiency. This curve is then superimposed on composite-type plots representing the individual effects of the normally interacting turbine operating variables. The procedure permits ready comparison of partial admission as a method of power reduction with other methods such as, inlet pressure, pressure ratio, and rotor speed on the basis of flexibility, range of effectiveness, and over-all efficiency at reduced power levels.\n\nThe causes of induced turbine-rotor-blade vibration are briefly discussed and means are examined for minimizing the effect of vibrational stresses at resonance conditions.", "timestamp": "2026-07-22T05:48:45.703987+00:00"}
{"citation_id": "19930085977", "source_url": "https://ntrs.nasa.gov/api/citations/19930085977/downloads/19930085977.pdf", "page_number": 23, "total_pages": 33, "image_filename": "19930085977_p23.jpg", "text": "```markdown\n22\nNACA RM L9H22\n\nCONFIDENTIAL\n\n<!-- Image (100, 110, 852, 876) -->\n\nFigure 8.- Concluded.\nCONFIDENTIAL\n```", "timestamp": "2026-07-22T05:48:49.841188+00:00"}
{"citation_id": "19930082472", "source_url": "https://ntrs.nasa.gov/api/citations/19930082472/downloads/19930082472.pdf", "page_number": 9, "total_pages": 34, "image_filename": "19930082472_p9.jpg", "text": "NACA TN No. 1797\n\nThese pressure distributions are typical of those obtained over the inboard 41.7 percent of the wing. Up to about $9.4^\\circ$ angle of attack, normal pressure distributions were obtained. At $12.5^\\circ$, however, pressures failed to recover over the rear portion of the section even though the rate of growth of the forward pressures was little affected.$^2$ This change in the pressure distributions caused little change in the section lift curves but shifted the center of pressure of the sections rearward. The shift of center of pressure can be seen in figure 7 in which are shown section lift and center-of-pressure curves obtained by mechanically integrating the pressure distributions. The rearward shift of center of pressure occurred at about $10^\\circ$ over the section at 20.9-percent semispan. As angle of attack was increased, sections further outboard exhibited this movement until at about $16^\\circ$, sections out to 41.7-percent semispan were so affected. This chordwise redistribution of load caused the negative trend in the wing pitching-moment curve between $10^\\circ$ and $16^\\circ$ angle of attack.\n\nThe previously mentioned change in the pressure distribution is comparable to changes that occur in pressure distributions over two-dimensional airfoils during the initial stages of turbulent boundary-layer separation. In the two-dimensional case, the failure of the pressures to recover is associated with the formation of a large wake, in effect, an abnormally thick boundary layer following the reversal of flow over the rear portion of airfoil. In the case of the swept-forward wing, however, no reversal of flow was indicated.\n\nAn understanding of the phenomena that caused the change in the pressure distributions at about $12.5^\\circ$ angle of attack can be obtained by following the reasoning in reference 2. Separation was taken to mean that the fluid in the boundary layer had lost the component of momentum that carried it across the surface in a direction perpendicular to the long axis of the wing. Therefore, when separation occurs over an oblique wing it was reasoned that the boundary layer would flow in a direction parallel to the long axis of the wing and on this basis would not necessarily be expected to be accompanied by a reversal of the boundary-layer flow. In this respect, separation over an oblique wing would differ from separation over a two-dimensional section. A rapid increase in boundary-layer thickness would, however, be expected due to the combined effects of chordwise and spanwise flow. In this respect, separation over an oblique wing would be similar to separation over a two-dimensional section. Boundary-layer measurements substantiated the above\n\n---\n\n$^2$Since pressure distributions were not obtained between $9.4^\\circ$ and $12.5^\\circ$ angle of attack, this change in the pressure distributions could have occurred at any point between those two angles.", "timestamp": "2026-07-22T05:48:51.730573+00:00"}
{"citation_id": "19930085938", "source_url": "https://ntrs.nasa.gov/api/citations/19930085938/downloads/19930085938.pdf", "page_number": 27, "total_pages": 42, "image_filename": "19930085938_p27.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:48:53.506430+00:00"}
{"citation_id": "19930085972", "source_url": "https://ntrs.nasa.gov/api/citations/19930085972/downloads/19930085972.pdf", "page_number": 29, "total_pages": 46, "image_filename": "19930085972_p29.jpg", "text": "NACA RM L9B18\n27\n\nPitching-moment coefficient, $C_m$\n4\n.2\n0\n-.2\n-.4\n\n$i$ (deg)\n$\\nabla$ 0\n$\\nabla$ 3\n$\\diamond$ 0 tail off\n$\\square$ 0 Unfaired cutout\n$\\circ$ 3\n$\\triangle$ tail off\nNo cutout\n\nLongitudinal-force coefficient, $C_x$\n.3\n.2\n.1\n0\n\nAngle of attack, $\\alpha$, deg\n16\n8\n0\n-8\n\nLift coefficient, $C_L$\n-4\n0\n4\n8\n12\n\nNACA\n\n(e) Sharp-leading-edge wing section.\nFigure 7.- Continued.", "timestamp": "2026-07-22T05:48:53.809508+00:00"}
{"citation_id": "19930086092", "source_url": "https://ntrs.nasa.gov/api/citations/19930086092/downloads/19930086092.pdf", "page_number": 16, "total_pages": 28, "image_filename": "19930086092_p16.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:48:54.605055+00:00"}
{"citation_id": "19930085914", "source_url": "https://ntrs.nasa.gov/api/citations/19930085914/downloads/19930085914.pdf", "page_number": 37, "total_pages": 42, "image_filename": "19930085914_p37.jpg", "text": "36\nNACA RM A9D25\n\n<!-- Image (171, 90, 823, 800) -->\n\n(a) $C_D$ vs $\\alpha$.\nFigure 15.- The effect of Mach number on the aerodynamic characteristics of the fuselage at a Reynolds number of 2,000,000.", "timestamp": "2026-07-22T05:48:59.579242+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 20, "total_pages": 114, "image_filename": "19930086061_p20.jpg", "text": "16\nNACA RM L9J07\n\nlarger camber, thereby producing a higher loading. Farther outboard where\nthe separated region covered the entire chord and the boundary layer\nthickened, the sections effectively became stalled, thereby causing the\naforementioned dip in the span-loading curves. Since the humps and dips\nwere a function of the vortex flow, they shifted progressively inboard\nwith increased angle of attack. Yawing the wings moved more of the span-\nwise loading to the leading semispan, especially at low angles of attack\n(figs. 37 to 40).\n\nThe spanwise-loading humps and dips for wing 2 at $\\psi = 0^\\circ$ are shown\nin figure 41. The spanwise loading at $\\alpha = 4.1^\\circ$ agreed fairly well with\nthe Weissinger theoretical loading obtained by use of reference 13,\nalthough the experimental curve had a hump above the theoretical curve\noutboard from station 4. Study of the basic pressure distributions\nreveals that the hump resulted from the weakening or loss of the negative-\nchordwise-pressure dip behind the vortex. The humps located at approxi-\nmately 65, 60, 30, 15, 10, and 0 percent of the semispan as the vortex\nswept back at angles of attack of $8.1^\\circ$, $14.1^\\circ$, $24.1^\\circ$, $32.1^\\circ$, $36.1^\\circ$,\nand $44.1^\\circ$ may be attributed mainly to additional camber effects. Yawing\nwing 2 reduced or removed the humps in the loading curves on the trailing\nsemispan (figs. 42 to 44) and increased the magnitude of the humps on\nthe leading semispan. The loading difference between the two semispans\nwas most pronounced for low angles of attack and decreased as the angle\nof attack was increased.\n\nThe variation of the spanwise loading with angle of attack and yaw\nfollowed the same trends for all three wings except that the humps and\ndips tended to be more pronounced for wing 1 than for the other two wings.\nThe comparison of the experimental and theoretical loadings for $\\alpha = 4.1^\\circ$\nof wing 3 was much poorer than for the other wings with the experimental\nloading being considerably higher for the center station 1.\n\nThe discrepancy noted between the loading of station 1 in positive\nand negative yaw generally increased with angle of yaw. The increase in\nthe pitch angularity of the air stream in the negative-yaw direction, as\nfound by the survey, undoubtedly had an appreciable effect in causing\nthe loading of station 1 to be generally higher at negative yaw than at\npositive yaw. However, the large discrepancy in the variation among\nwings, especially between wing 3 and the other two wings, seems to indi-\ncate that the effect of the orifices of station 1 not being located on\nthe exact center of the rounded ridge had a greater effect in yaw than\nthe flow irregularity as discussed in the section entitled \"Air-Stream\nFlow Analysis.\"", "timestamp": "2026-07-22T05:49:01.005931+00:00"}
{"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 61, "total_pages": 78, "image_filename": "19930082618_p61.jpg", "text": "```markdown\nNACA TN 1945\n\n<!-- Image (79, 96, 875, 794) -->\n\n(c) Section drag characteristics and section pitching-moment characteristics about the aerodynamic center of the plain NACA 4415 airfoil section.\n\nFigure 13.- Concluded.\n\n59\n```", "timestamp": "2026-07-22T05:49:02.343945+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 19, "total_pages": 98, "image_filename": "19930086073_p19.jpg", "text": "NACA RM A9H04\n\nPoints of tangency\nNose radius, 0.00250c\nin terms of basic chord,\n0.00254c in terms of\nmodified chord\n\nBasic section\n\nArc of circle with center of curvature 0.62c above or\nbelow chord line at station 0.23c, basic chord\n\nMaximum thickness of basic section, 0.050c\n\nMaximum thickness of modified section,\n0.0475c in terms of basic chord,\n0.0483c in terms of modified chord\n\nPoints of tangency\n\nTrailing-edge angle,\n3.58°\n\nChord line\n\nStation\nBasic chord\n0 .0176 .0201 .150 .200 .230 .250 1.00\n\nStation\nModified chord\n0 .00254 .135 .186 .216 .236 1.00\n\nLeading edge\n\nMaximum thickness\n\nTrailing edge\n\n(b) Modified double-wedge airfoil section.\n\nFigure 2.- Concluded.\n\n17", "timestamp": "2026-07-22T05:49:03.644514+00:00"}
{"citation_id": "19930086097", "source_url": "https://ntrs.nasa.gov/api/citations/19930086097/downloads/19930086097.pdf", "page_number": 15, "total_pages": 36, "image_filename": "19930086097_p15.jpg", "text": "NACA RM A9H11 CONFIDENTIAL 13\n\nalso be expected to be of importance. Since the lift-curve slope of sharp- and blunt-trailing-edge airfoils differs only when second-order terms are considered, the analysis which follows must consider terms of equal order throughout. Only profiles symmetric about the chord line will be considered here, as the algebra would otherwise become unduly involved without significantly affecting the final result.\n\nThe section drag coefficient to second order in angular deflection terms is\n\n$$\nc_d = \\int_0^1 (P_l \\theta_l + P_u \\theta_u) \\, d\\left(\\frac{x}{c}\\right) + c_{db} + c_{df}\n\\tag{19}\n$$\n\nSubstituting equation (11) and noting that for symmetrical airfoils\n\n$$\n\\left(\\frac{dy}{dx}\\right)_u = -\\left(\\frac{dy}{dx}\\right)_l = \\frac{dy}{dx}\n$$\n\n$$\nc_d = 2C_1 \\int_0^1 \\left[\\left(\\frac{dy}{dx}\\right)^2 + \\alpha^2\\right] d\\left(\\frac{x}{c}\\right) + 6C_2 \\int_0^1 \\left[\\alpha^2 \\frac{dy}{dx} + \\frac{1}{3}\\left(\\frac{dy}{dx}\\right)^3\\right] d\\left(\\frac{x}{c}\\right) + c_{db} + c_{df}\n$$\n\n$$\n= 2C_1 \\alpha^2 + 6C_2 \\alpha^2 \\frac{h}{2c} + 2C_1 \\int_0^1 \\left[\\left(\\frac{dy}{dx}\\right)^2 + \\frac{C_2}{C_1} \\left(\\frac{dy}{dx}\\right)^3\\right] d\\left(\\frac{x}{c}\\right) + c_{db} + c_{df}\n$$\n\nFor simplicity, the base drag coefficient will be taken as being approximately independent of $\\alpha$, then\n\n$$\nc_d = 2C_1 \\alpha^2 \\left(1 + \\frac{3C_2 h}{2C_1 c}\\right) + c_{dmin}\n\\tag{20}\n$$\n\nSince\n\n$$\nc_l = 2C_1 \\alpha \\left(1 + \\frac{C_2}{C_1} \\frac{h}{c}\\right)\n$$\n\nthe drag-lift ratio is approximately\n\n$$\n\\frac{c_d}{c_l} = \\frac{1}{1 + \\frac{C_2}{C_1} \\frac{h}{c}} \\left[ \\alpha \\left(1 + \\frac{3C_2 h}{2C_1 c}\\right) + \\frac{c_{dmin}}{2C_1 \\alpha} \\right]\n\\tag{21}\n$$\n\nThe minimum of this function occurs when\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:49:07.488738+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 52, "total_pages": 96, "image_filename": "19930085880_p52.jpg", "text": "50\nNACA RM No. L9C03\n\nDraft, ft\n.64\n.56\n.48\n.40\n.32\n.24\n.16\n.08\n0\n\nSpeed\n(fps)\n10 O\n15 □\n20 ◇\n25 △\n30 ▽\n\n0 .05 .10 .15 .20 .25 .30 .35\nWetted area, sq ft\n(a) $\\tau = 40^\\circ$.\n\n[Figure: A graph plotting Draft (ft) against Wetted area (sq ft) for various speeds. Data points are represented by different symbols corresponding to speeds of 10, 15, 20, 25, and 30 fps. A trend line is drawn through the data points.]\n\nNACA\n\nFigure 17.- Variation of draft with wetted area. Model 250A.", "timestamp": "2026-07-22T05:49:12.088742+00:00"}
{"citation_id": "19930091987", "source_url": "https://ntrs.nasa.gov/api/citations/19930091987/downloads/19930091987.pdf", "page_number": 7, "total_pages": 12, "image_filename": "19930091987_p7.jpg", "text": "CHARACTERISTICS OF LOW-ASPECT-RATIO WINGS AT SUPERCRITICAL MACH NUMBERS\n\nThe systematic errors arose from end interference and tunnel-wall interference. The data have been corrected for end interference resulting from the small leakage through the clearance gap at the juncture of the model and the tunnel wall by corrections determined experimentally (reference 4). With the wings of low aspect ratio, a large part of the whole span is affected by the gap leakage, and the effects of the leakage are therefore expected to be relatively greater. It is believed that the drag coefficients for the low-aspect-ratio wings are higher and the lift-curve slopes are lower than would be found in the absence of leakage. These effects are being studied experimentally in the Langley 24-inch high-speed tunnel.\n\nTunnel-wall interference has been investigated theoretically (reference 5) and the existence of constriction effects in high-speed tunnels has been shown experimentally (references 6 and 7). The theoretically derived corrections, however, have not been experimentally verified at supercritical Mach numbers. The errors indicated theoretically by the method of reference 5 increase as model size, Mach number, drag coefficient, and lift coefficient increase. The theoretically indicated errors for the infinite-aspect-ratio wing at an angle of attack of $6^\\circ$ and a Mach number $M$ of 0.84 are:\n\nCorrected $M = M \\times 1.014$\n\nCorrected $C_L = C_L \\times 0.982$\n\nCorrected $C_D = C_D \\times 0.980$\n\nThese values indicate that the effect of tunnel-wall interference on these data is small, 2 percent or less, and therefore no correction has been applied.\n\nThe choking phenomenon is an additional factor that enters into the problem of testing at high Mach numbers. At the choked Mach number sonic velocities extend from model to tunnel wall and the static pressure is lower behind the model than it is ahead; thus large gradients in the pressure are produced (reference 7). The resulting flow past the model is unlike any free-air condition. Data obtained at the choked Mach number are therefore of questionable value and are not presented herein.\n\nRESULTS\n\nThe basic results are presented in figures 3 to 5. Figure 3 shows the lift coefficient $C_L$ plotted against the angle of attack $\\alpha$ for each of nine values of the Mach number from 0.5 to 0.9. Similarly, the drag results are shown by polar diagrams in figure 4. The moment coefficients are given in figure 5 plotted against lift coefficient for six values of the Mach number in the range from 0.5 to 0.9. The minimum drag coefficients for the various aspect ratios are shown in figure 6. All the results are for the actual aspect ratios tested and are not corrected to infinite aspect ratio. Thus, the induced effects are included. Since the wings of higher aspect ratio have the lower choking Mach numbers, data for these wings are presented for somewhat lower Mach numbers than the data for the lower aspect ratios. As noted previously, no data at the choked condition are presented. The highest Mach number for which data are presented for each wing is approximately 0.025 less than the corresponding choking Mach number.\n\n[Figure: Nine subplots labeled (a) through (i), each showing $C_L$ vs. $\\alpha$, deg, with multiple curves representing different aspect ratios. Vertical axis labeled $C_L$ ranging from 0 to .6. Horizontal axis labeled $\\alpha$, deg ranging from 0 to 4. A legend on the right indicates aspect ratios: $\\infty$, 9, 7, 5, 3, 2. Below the plots, a key lists Mach numbers for each subplot: (a) $M=0.50$, (b) $M=0.60$, (c) $M=0.70$, (d) $M=0.75$, (e) $M=0.80$, (f) $M=0.825$, (g) $M=0.85$, (h) $M=0.875$, (i) $M=0.90$. Caption reads: FIGURE 3.—Lift curves for various aspect ratios and Mach numbers. NACA 0312 section, rectangular plan form and tip.]", "timestamp": "2026-07-22T05:49:15.466842+00:00"}
{"citation_id": "19930093769", "source_url": "https://ntrs.nasa.gov/api/citations/19930093769/downloads/19930093769.pdf", "page_number": 9, "total_pages": 39, "image_filename": "19930093769_p9.jpg", "text": "8 CONFIDENTIAL NACA RM No. E8L10a\n\nAltitude Starting Characteristics\n\nThe altitude starting characteristics of AN-F-58 fuel were investigated with NACA fuel 48-210 in engine A and, for comparison, with gasoline in engine B. The procedure used to start the engine consisted in obtaining the windmilling speed of the engine at the particular flight conditions, turning on the ignition, and advancing the fuel throttle. The ignition and the fuel flow were maintained for a period of 30 seconds before the attempted start was considered unsatisfactory. At each flight condition investigated, a series of three attempts was made before starts at the particular condition were considered impossible. If ignition of the fuel was obtained at windmilling speed, an attempt was also made to accelerate the engine. Immediately after ignition, the pressure in the exhaust end of the test chamber increased from 4 to 10 inches of mercury and this pressure surge prevented these experiments from being a true simulation of flight conditions. The altitude pressure was restored to its correct value in about 10 seconds and the engine was accelerated after the correct conditions had become stabilized.\n\nThe first spark plug used in the starting experiments was the standard spark plug (fig. 3(a)). The experimental results with this spark plug are shown in figure 9(a), which includes data for satisfactory starts, burner ignition without engine acceleration, and no ignition. The spark plug was cleaned for each experiment. If an attempted start was unsuccessful, the spark plug filled with fuel, which prevented satisfactory ignition at any operating condition. Although starts were obtained at altitudes up to 30,000 feet at low flight Mach numbers (fig. 9(a)), practically the use of the standard spark plug would not provide satisfactory engine starting at any flight condition.\n\nThe second spark plug used was the modified spark plug (fig. 3(b)). Several attempts to start the engine with this spark plug at an altitude of 8000 feet and a flight Mach number of 0.30 were unsuccessful and further use of this spark plug was therefore discontinued.\n\nThe third spark plug used was the extended-electrode type (fig. 3(c)). Results of starting experiments with this spark plug are presented in figure 9(b). For all these experiments, the spark plug was used without cleaning between runs. During the experiment in the region of unsatisfactory starts, occasional starts were made at satisfactory starting conditions to assure that the ignition system was functioning properly. These data are insufficient to establish definitely the boundaries of the\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:49:18.096550+00:00"}
{"citation_id": "19930086076", "source_url": "https://ntrs.nasa.gov/api/citations/19930086076/downloads/19930086076.pdf", "page_number": 20, "total_pages": 50, "image_filename": "19930086076_p20.jpg", "text": "18\nNACA RM E9F09\n\n<!-- Image (127, 108, 816, 876) -->\n\nFigure 4. - Lean combustion limits for flame holders 1 to 6. Inlet-air pressure, 55 inches mercury absolute; inlet-air temperature, 200° F.", "timestamp": "2026-07-22T05:49:21.047825+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 5, "total_pages": 78, "image_filename": "19930082483_p5.jpg", "text": "NACA TN No. 1807\n3\n\nANALYSIS\n\nThe ideal power in the driving fluid of any turbine may be expressed by\n\n$$ \\text{ideal power} = W \\Delta_s h' \\quad (1) $$\n\nwhere\n\nW weight flow through turbine\n\n$\\Delta_s h'$ isentropic enthalpy drop per pound of driving fluid based on total-pressure ratio\n\nPrimed symbols indicate stagnation state. (All symbols used in this analysis are defined in appendix A.)\n\nUse of the isentropic enthalpy drop based on total-pressure ratio assumes that useful work can be recovered from the turbine-discharge gas velocity.\n\nNot all the weight flow expressed in equation (1) is available to the turbine for doing work. Some fraction of the gas flow passes through the clearance space between the rotor-blade tips and the stationary turbine outer casing. The presence of this rotor-tip leakage constitutes a turbine power loss. Friction, separation, and other viscous effects give rise to losses in the nozzle and rotor flow passages, which further prevent realization of the ideal power in equation (1).\n\nGross power is defined as ideal power minus these rotor-tip leakage and nozzle and rotor-blading losses.\n\nThe mechanism by which power is derived from the driving fluid in a turbine is described by\n\n$$ \\text{blade power} = W(1-K_I) u_m (\\bar{c}_{u,1} - \\bar{c}_{u,2}) \\quad (2) $$\n\nwhere\n\n$K_I$ percentage of active gas through rotor-tip clearance space, constant for any particular turbine\n\nu velocity, (ft/sec)\n\nc absolute velocity, (ft/sec)", "timestamp": "2026-07-22T05:49:21.231745+00:00"}
{"citation_id": "19930085938", "source_url": "https://ntrs.nasa.gov/api/citations/19930085938/downloads/19930085938.pdf", "page_number": 28, "total_pages": 42, "image_filename": "19930085938_p28.jpg", "text": "NACA RM No. L9B04\n27\n\n<!-- Image (109, 109, 926, 904) -->\n\nFigure 6.- General arrangement of model 237-7TB. (All dimensions are in inches.)", "timestamp": "2026-07-22T05:49:25.454134+00:00"}
{"citation_id": "19930093789", "source_url": "https://ntrs.nasa.gov/api/citations/19930093789/downloads/19930093789.pdf", "page_number": 8, "total_pages": 29, "image_filename": "19930093789_p8.jpg", "text": "NACA RM No. E9I21 CONFIDENTIAL 7\n\nWater-brake torque was indicated by an NACA balanced-diaphragm dynamometer-torque indicator (reference 1). Air flow w was measured by the 7-inch submerged flat-plate orifice upstream of the air heater (fig. 9). Turbine speed was indicated by a calibrated electric tachometer.\n\nThe accuracy in reading the data may be summarized as follows:\n\n(1) Absolute pressure, ±0.05 inch of mercury  \n(2) Temperatures, ±1° R  \n(3) Orifice pressure drop, ±0.05 inch in a range with a minimum of 15 inches  \n(4) Torque, ±0.05 inch in a range with a minimum of 17 inches  \n(5) Rotative speed, ±13 rpm in a range with a minimum of 3700 rpm  \n\nAn estimate of the reproducibility of the measurements indicates that the probable variation in the brake efficiency for a single data point is less than 0.01 for total-pressure ratios $p_1'/p_3'$ greater than 2.00.\n\n---\n\n**Experimental Procedure**\n\nData were taken at turbine speeds from 3700 to 16,800 rpm (equivalent mean blade speeds from 188 to 855 ft/sec) at total-pressure ratios $p_1'/p_3'$ from 1.25 to 3.70. At each of ten pressure ratios, the turbine was operated at six to ten different speeds. For all runs the entrance total temperature was held between 658° and 662° R; the entrance total pressure varied with the air flow and the barometric pressure between $26\\frac{1}{2}$ and 28 inches of mercury.\n\n[Handwritten annotation in red ink near bottom left: “198 lb/sec”]\n\n[Handwritten annotation in red ink near bottom right: “= p₁’”]\n\n---\n\n**PERFORMANCE CALCULATIONS**\n\nAll turbine performance data were reduced to NACA sea-level air conditions at the turbine entrance. The performance was determined in terms of the following variables:\n\n(1) Brake efficiency, $\\eta$  \n(2) Total-pressure ratio, $p_1'/p_3'$  \n(3) Equivalent turbine-shaft work, $E \\left(\\frac{a_0}{a_1}\\right)^2$  \n(4) Equivalent mean rotor-blade speed, $U \\left(\\frac{a_0}{a_1}\\right)$\n\n[Handwritten calculations in black ink on right side:]\n\n$$\n= U \\sqrt{\\frac{p_1'}{p_1}} = U \\sqrt{\\frac{519 \\times 14}{1960 \\times 13}} = .534 U\n$$\n\n[Below, handwritten:]\n\nfor design:  \nfor test $U \\sqrt{\\frac{p_3'}{660}} = .961 U$\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:49:28.890717+00:00"}
{"citation_id": "19930091993", "source_url": "https://ntrs.nasa.gov/api/citations/19930091993/downloads/19930091993.pdf", "page_number": 7, "total_pages": 21, "image_filename": "19930091993_p7.jpg", "text": "```markdown\nANALYSIS OF PERFORMANCE OF JET ENGINE FROM CHARACTERISTICS OF COMPONENTS 3\n\nvarious stations indicated a probable error of the mean value of the exhaust temperature varying from about $2^\\circ$ to $33^\\circ$ F. This error caused no error in the enthalpy drop in the turbine $\\Delta H_t$, equivalent isentropic enthalpy drop in the turbine $\\Delta H_{t,s}(\\theta_{T,3}, \\text{equivalent turbine torque } W_2\\Delta H_t/(n\\sigma_{T,3}\\theta_{T,3}))$, and equivalent turbine gas flow $W_3/(\\sigma_{T,3}\\sqrt{\\theta_{T,3}})$, but did result in a maximum probable error of 1 percent in equivalent turbine gas flow $W_3/(\\sigma_{T,3}\\sqrt{\\theta_{T,3}})$ and equivalent turbine speed $n/\\sqrt{\\theta_{T,3}}$ and a maximum probable error of 2 percent in equivalent turbine enthalpy drop $\\Delta H_t/\\theta_{T,3}$ and turbine efficiency. (All symbols are defined in the appendix.)\n\nShaft speed was measured by an alternating-current generator built into the compressor. The rotative frequency was determined by comparison of the output frequency of the built-in alternator with that of a calibrated variable-frequency generator. These measurements were accurate to within 1 percent.\n\nThermocouples were also installed in the oil lines and on the bearings in order to determine the bearing power losses and oil viscosity.\n\n### PROCEDURE AND CALCULATIONS\n\nInaccuracy of the measured turbine-inlet temperature and the unknown heat losses made it impossible to determine from the inlet- and outlet-gas temperatures the mechanical-power input from the gas to the turbine rotor with the desired accuracy. This power was computed by adding the power output of the compressor to the gas and the power absorbed by the bearings.\n\nThe bearing-power consumptions were determined from the heat absorbed by the lubricating oil. The front compressor bearing had a separate oil supply from the remainder of the bearing system, which consisted of rear compressor and turbine bearings contained between the compressor and the turbine. In order to reduce the heat-transfer error to a minimum while the bearing power was calibrated, the inlet-air temperature was matched with the mean oil temperature while the front-bearing power consumption of the compressor was measured over a range of speeds. In calibrating the rear-bearing system, the compressor-outlet-air temperature was matched with the mean oil temperature for the system. The rear bearing was then calibrated over a range of engine air pressures in order to vary the thrust but no effect of load on the thrust bearing could be found. When the bearing-power consumptions were reduced to equivalent values for a mean oil temperature of $100^\\circ$ F, both sets of bearings showed a power absorption exactly proportional to the square of the speed.\n\nThe range of conditions for operation and study of the engine was limited by several factors. Simulated high-altitude operation had to be avoided because of the failure of the burners to operate satisfactorily with low pressure. There were two limitations on the temperature of the intake air: (1) Inasmuch as the compressor was made of aluminum, the inlet-air temperature was so limited that the compressor-outlet-air temperature did not exceed $300^\\circ$ F because of the reduction in the yield stress of this metal above $300^\\circ$ F; and (2) because investigation of the performance over a wide range of the ratio of turbine-inlet-gas temperature to compressor-inlet-air temperature $T_3/T_1$ was desirable, very cold inlet air had to be used to explore engine operation at high values of this parameter without exceeding the maximum safe metal temperature of the turbine. Periodic checks during the course of the experiments revealed that no creep of the wheel had taken place after 76 hours of engine operation at or below a gas temperature of $1250^\\circ$ F. The turbine rotor blades rubbed on the casing at the conclusion of a 4-hour period of operation at $1350^\\circ$ F with speeds varying from 110 to 215 rps. Subsequent measurements showed that the first rotor had stretched 0.067 inch and the second rotor, which was about $150^\\circ$ F cooler, stretched 0.005 inch. The radial stress at the blade roots was estimated to be 16,000 pounds per square inch. For this stress, static-stress data on Nimonic 80 showed that the critical metal temperature was in the region of $1200^\\circ$ to $1250^\\circ$ F. Succeeding operating gas temperatures were consequently limited to $1250^\\circ$ F and no further difficulties were noted.\n\nEngine data were obtained over a variable range of the equivalent compressor speed $n/\\sqrt{\\theta_{T,1}}$ and the equivalent temperature ratio $\\theta_{T,3}/\\theta_{T,1}$ where\n\n$n$ rotative speed of engine, (rps)\n$\\theta$ ratio of square of sonic speed of gas to square of sonic speed for normal air\n\nSubscript $T$ denotes stagnation condition. Numerical subscripts indicate conditions at stations shown in figure 2. Because the measured turbine-inlet temperatures were not considered exact enough for computation of $\\theta_{T,3}/\\theta_{T,1}$, these temperatures were computed from the turbine-outlet temperatures and from the bearing and compressor power. Operating conditions of the engine were approximately set for desired values of $\\theta_{T,3}/\\theta_{T,1}$ from the readings of the turbine-inlet thermocouples by using\n\n$$ \\frac{\\theta_{T,3}}{\\theta_{T,1}} = \\frac{\\gamma_3 R_3 T_{T,3}}{\\gamma_1 R_1 T_{T,1}} = \\frac{T_{T,3}}{T_{T,1}} $$\n\nwhere\n\n$R$ gas constant, (ft-pound)/(lb) ($^\\circ$R)\n$\\gamma$ ratio of specific heats\n\nThe ratio $T_{T,3}/T_{T,1}$ was set at values from 4.43 to values for fuel input of zero. Corrected compressor speeds varied from 90 to 290 rps.\n\nThe enthalpy rise of the air in the compressor was determined from the stagnation inlet- and outlet-air temperatures and the tables of reference 5. The isentropic enthalpy rise was computed from the stagnation pressures $p_{T,1}$ and $p_{T,2}$ and the same tables. For the turbine, the enthalpy of the exhaust gas was determined from the exhaust temperature $T_{T,4}$, the fuel-air ratio $f$, and the charts of reference 6. The enthalpy drop in the turbine was then determined from the equation\n\n$$ W_1(1+f)(H_{T,3}-H_{T,4}) = W_1(H_{T,2}-H_{T,1}) + P_s $$\n\nwhere\n\n$H$ enthalpy, (ft-pound)/lb\n$P_s$ power absorbed by bearings, (ft-pound)/sec\n$W_1$ mass flow through compressor, (lb/sec)\n```", "timestamp": "2026-07-22T05:49:30.570701+00:00"}
{"citation_id": "19930082472", "source_url": "https://ntrs.nasa.gov/api/citations/19930082472/downloads/19930082472.pdf", "page_number": 10, "total_pages": 34, "image_filename": "19930082472_p10.jpg", "text": "```markdown\n8\nNACA TN No. 1797\n\nreasoning. In figure 6, it can be seen that over the streamwise\nsection at 20.9-percent semispan the boundary layer thickened rapidly\nat $12.5^\\circ$ angle of attack. This is the angle of attack at which the\npressures first failed to recover and at which tufts began to oscil-\nlate but did not reverse direction. Furthermore, the portion of the\nchord over which the boundary layer thickened was the same as that\nover which the pressures failed to recover.\n\nThe effects of separation of the turbulent boundary layer were\nevidenced both in thickened boundary layers and in a failure of\npressures to recover over the rear portion of the sections even\nthough reversal of flow did not occur. Thus it apparently is a form\nof turbulent separation that caused the initial increase in drag and\nthe rearward shift of aerodynamic center.\n\nHigh-Lift Range\n\nAbove $16.6^\\circ$ angle of attack, the force tests showed that the\nnegative trend of pitching moments was rapidly and completely\nreversed, with the result that marked longitudinal instability was\nindicated. This was accompanied by greatly increased drag and a\ngradual decrease in lift-curve slope. Corresponding changes that\ntook place in the pressure distributions over the streamwise section\nat 20.9-percent semispan can be seen in the following diagram:\n\n[Figure: Graph showing Chordwise pressure distributions at 20.9% semispan station.\nY-axis: Pressure coefficient, $P$ (-70 to 0).\nX-axis: Chordwise station, $x/c$ (0 to 1.0).\nLegend:\n$P$ at $14.6^\\circ$ at peak = 90\n$P$ at $16.6^\\circ$ at peak = 6.4\n$\\alpha$, angle of attack\n$14.6^\\circ$ ---\n$16.6^\\circ$ - - -\n$20.7^\\circ$ — —]\n```", "timestamp": "2026-07-22T05:49:32.246436+00:00"}
{"citation_id": "19930085972", "source_url": "https://ntrs.nasa.gov/api/citations/19930085972/downloads/19930085972.pdf", "page_number": 30, "total_pages": 46, "image_filename": "19930085972_p30.jpg", "text": "28\nNACA RM L9B18\n\nPitching-moment coefficient, $C_m$\n.2\n0\n-.2\n-.4\n-.6\n\n4\n(deg)\n$\\nabla$ $O$ No cutout\n$\\diamond$ $O$ tail off\n$\\square$ $O$ Unfaired cutout\n$\\triangle$ tail off\n\nLongitudinal-force coefficient, $C_x$\n.3\n.2\n.1\n0\n\nAngle of attack, $\\alpha$, deg\n16\n8\n0\n-8\n\nLift coefficient, $C_L$\n-4 0 4 8 12\n\nNACA\n\n(f) Wing vane.\nFigure 7.- Concluded.", "timestamp": "2026-07-22T05:49:33.795349+00:00"}
{"citation_id": "19930086092", "source_url": "https://ntrs.nasa.gov/api/citations/19930086092/downloads/19930086092.pdf", "page_number": 17, "total_pages": 28, "image_filename": "19930086092_p17.jpg", "text": "NACA RM A9F14 CONFIDENTIAL 15\n\n[Figure: A large model of a swept-wing aircraft mounted on supports inside a wind tunnel, with a person standing nearby for scale. The NACA logo and identifier \"A-12457\" are visible in the lower right corner of the image.]\n\n(a) Three-quarter front view.\n\nFigure 2.— The $63^\\circ$ swept-back wing-fuselage combination with $63^\\circ$ swept-back vertical tail.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:49:33.973998+00:00"}
{"citation_id": "19930091987", "source_url": "https://ntrs.nasa.gov/api/citations/19930091987/downloads/19930091987.pdf", "page_number": 8, "total_pages": 12, "image_filename": "19930091987_p8.jpg", "text": "4\nREPORT 922—NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\n<!-- Image (108, 78, 901, 874) -->\n\n(a) $M=0.30$.\n(d) $M=0.75$.\n(g) $M=0.85$.\n(b) $M=0.60$.\n(e) $M=0.80$.\n(h) $M=0.875$.\n(c) $M=0.70$.\n(f) $M=0.825$.\n(i) $M=0.90$.\n\nFIGURE 4.—Polars for various aspect ratios and Mach numbers. NACA 0012 section; rectangular plan form and tip.", "timestamp": "2026-07-22T05:49:46.240894+00:00"}
{"citation_id": "19930086076", "source_url": "https://ntrs.nasa.gov/api/citations/19930086076/downloads/19930086076.pdf", "page_number": 21, "total_pages": 50, "image_filename": "19930086076_p21.jpg", "text": "NACA RM E9F09\n19\n\n[Figure: Three damaged metal gutter structures with a scale labeled \"INCHES\" showing 0 and 1. A NACA logo with \"C-21562\" and \"6-7-48\" is in the bottom right corner of the image.]\n\nFigure 5. - Three gutters located at downstream end of two rows of flame holder 6 after 10 minutes of operation.", "timestamp": "2026-07-22T05:49:49.458469+00:00"}