PDF/ocr_dataset_1067.jsonl

{"citation_id": "19930085548", "source_url": "https://ntrs.nasa.gov/api/citations/19930085548/downloads/19930085548.pdf", "page_number": 30, "total_pages": 46, "image_filename": "19930085548_p30.jpg", "text": "```markdown\nNACA RM No. E8L30\n\n2.0\nNACA\n\nProduct of discharge coefficient\nand port area, sq in.\n\n1.6\n\n1.2\n\n.8\n\n.4\n\n0\n100\nExhaust port opens\nInlet port opens\n120\n140\n160\n180\n200\n220\nInlet port closes\nExhaust port closes\n240\n260\nCrank angle, deg A.T.C.\n\nO Inlet\n□ Exhaust\n\n(b) Products of inlet- and exhaust-port areas and corresponding\ndischarge coefficients at each crank angle. Average values:\ninlet, 0.653 square inch; exhaust, 0.923 square inch.\n\nFigure 7. - Concluded. Flow coefficients of $3\\frac{1}{4}$- by $4\\frac{1}{2}$-inch ported cylinder.\n\n29\n```", "timestamp": "2026-07-22T06:29:43.833770+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 48, "total_pages": 78, "image_filename": "19930082483_p48.jpg", "text": "46\nNACA TN No. 1807\n\ntherefore\n\n$$W = \\frac{514}{\\left(\\frac{23.76}{0.707}\\right)}$$\n\n$$= 15.29 \\text{ lb/sec}$$\n\n3. Discharge total pressure by equation (44)\n\n$$p_e' = \\frac{p_1'}{\\left(\\frac{p_1'}{p_e'}\\right)}$$\n\n$$= \\frac{29.92}{2.1}$$\n\n$$= 14.25 \\text{ in. Hg absolute}$$\n\n4. Discharge static pressure from equation (45)\n\n$$\\frac{p_e'}{p_e} = \\left[ \\frac{1}{2} + \\sqrt{\\frac{1}{4} + \\frac{1}{2g} \\frac{\\gamma-1}{\\gamma} \\left( RT_1'' - \\frac{\\gamma-1}{\\gamma} 550 \\frac{P}{W} \\right) \\left( \\frac{W}{70.73 \\ p_e \\ A_e} \\right)^2} \\right]^{\\frac{\\gamma}{\\gamma-1}}$$\n\nwhere\n\n$p_e'$ 14.25 in. Hg absolute\n\n$g$ 32.174 ft/sec$^2$\n\n$\\gamma$ 1.4\n\n$R$ 53.345 ft-lb/lb/$^\\circ$F\n\n$T_1'$ 518.6$^\\circ$ R\n\n$P$ 388 hp\n\n$W$ 15.29 lb/sec\n\n$A_e$ 0.686 sq ft", "timestamp": "2026-07-22T06:29:51.062181+00:00"}
{"citation_id": "19930085900", "source_url": "https://ntrs.nasa.gov/api/citations/19930085900/downloads/19930085900.pdf", "page_number": 7, "total_pages": 33, "image_filename": "19930085900_p7.jpg", "text": "6\nCONFIDENTIAL\nNACA RM L9D20\n\nconstant. An average air flow per jet of $11 \\times 10^{-5}$ pounds per second through the rows of $\\frac{1}{4}$-inch-spaced jets would amount to a total of about 14 pounds per second in the full-size hypothetical airplane (neglecting scale effect).\n\nComparison of Rows of Jets and Strips Simulating Chines\n\nThe strips (less than 2 percent of the maximum fuselage diameter) like the jets were intended as spoilers to cause separation. A comparison between the results for rows of jets simulating chines and the results for strips placed at the same location as the rows of jets is given in figure 13. The trim, resistance, and lift curves for the two modifications were practically the same. The strips also gave results substantially the same as the jets for other chine lengths. Figure 14 shows that the spray off the strips was much cleaner and did not rise as high as that for the jets.\n\nComparison of Rows of Jets and Strips Simulating Multiple Steps\n\nIn contrast to the results obtained when strips were substituted for rows of jets simulating chines there was no correlation between the results obtained when strips were substituted for rows of jets simulating steps. As shown in figure 15, the resistance and trim for the strips with V-steps pointed forward were very much higher than for $\\frac{1}{4}$-inch-spaced jets in the same configuration; they were even higher than for the basic model. This jet configuration was the best of the three jet configurations simulating multiple steps reported in reference 1.\n\nFigure 16 shows the results obtained with strips arranged in all three of the multiple-step configurations described in reference 1. The forward side of each individual strip is the hypotenuse of the $45^\\circ$ right triangle forming its cross section. The very high resistance for the V-steps pointed forward was not obtained with the other two step configurations. The maximum resistance for the V-steps pointed aft was about 2.5 pounds at 15 feet per second and the resistance never exceeded 2 pounds at the higher speeds. This configuration was a considerable improvement over the strips simulating chines for which the maximum resistance was about 4 pounds at 50 feet per second. No readings for the transverse steps were taken at speeds above 40 feet per second because the model became unstable.\n\nThe trim for the V-steps pointed aft reached a maximum of about $8^\\circ$ and dropped rapidly above 25 feet per second reaching a minimum of about $1^\\circ$ at 55 feet per second. The trim track for the transverse strips was similar.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:29:51.254947+00:00"}
{"citation_id": "19930082592", "source_url": "https://ntrs.nasa.gov/api/citations/19930082592/downloads/19930082592.pdf", "page_number": 48, "total_pages": 50, "image_filename": "19930082592_p48.jpg", "text": "NACA TN 1914\n47\n\n[Figure: Micrograph showing cross-section of material with labeled layers]\n- Bakelite\n- Outer oxide layer\n- Outer transition zone\n- Outer-oxide layer\n- Inner transition zone\n- Inner oxide layer\n\nNACA\nC-22917\n2-7-49\n\nFigure 18. - Oxidation zone of 30-percent-cobalt - titanium carbide ceramal. Oxide coating appears as two layers separated at oxidation interface and between layers by transition zone. Temperature, $2000^\\circ$ F; time at temperature, 30 hours; unetched; magnification, X50.", "timestamp": "2026-07-22T06:29:51.454908+00:00"}
{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 31, "total_pages": 72, "image_filename": "19930085491_p31.jpg", "text": "30 CONFIDENTIAL NACA RM No. A8J04\n\nA general consideration of the change in liquid-film patterns of figure 14(b) agrees with the observed variation of $k_a$ with sweep. The value of $k_a$ will be influenced by both the leading-edge suction pressures and the amount of pressure recovery over the rear of the wing. Therefore, although the area of leading-edge attached flow and high suction pressures is reduced as the sweep angle increases up to $67.0^\\circ$, the pressure recovery resulting from the increased area of reattached turbulent flow results in a nearly constant value of $k_a$ with sweep. At the highest angle of sweep, however, all of the leading-edge suction force is lost since laminar separation occurs along the entire leading-edge length. The change in pressure distribution associated with this loss of leading-edge suction would therefore be expected to increase $k_a$ as is shown by figure 10(c) between $67.0^\\circ$ and $69.9^\\circ$ sweep.\n\nEffect of sweep on maximum lift-drag ratio.— Figure 10(d) shows the variation of maximum lift-drag ratio with the factor $m$. This curve shows that the angle of leading-edge sweep for maximum lift-drag ratio at this Mach number is near $67.0^\\circ$ which corresponds to a value of $m$ equal to 0.49. The limitations of the linear theory when used with the present wings prevent a determination of the complete theoretical variation of maximum lift-drag ratio with sweep but it is noteworthy that the trend indicated by the four lowest sweep angles is similar to that obtained experimentally.\n\nTo give an indication of the relative proportions of the difference between experiment and theory due to the differences in minimum drag coefficient and drag-rise factor, an additional calculated curve is included in figure 10(d). This curve was determined using the experimental minimum drag coefficient at a Reynolds number of 0.62 million and the theoretical drag-rise factor. Thus, the difference between this curve and the experimental maximum lift-drag curve is a direct reflection of the differences in drag-rise factor. The differences between the two calculated curves is then the result of the higher experimental minimum drag-coefficient values since the drag-rise factor in both cases was taken as the theoretical value. The probable reasons for the differences between theory and experiment were discussed in the preceding sections which considered the effects of Reynolds numbers on minimum drag coefficient and drag-rise factor.\n\nThe value of $m$ of 0.49 at which the maximum experimental lift-drag ratio occurs is close to that indicated by the theory of reference 1 for a comparable Mach number with wings having trailing edges coincident with the Mach lines. It is interesting to note\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:29:51.972650+00:00"}
{"citation_id": "19930082646", "source_url": "https://ntrs.nasa.gov/api/citations/19930082646/downloads/19930082646.pdf", "page_number": 32, "total_pages": 37, "image_filename": "19930082646_p32.jpg", "text": "NACA TN 1980\n31\n\n$\\tau = 9.4^\\circ$\n$V = 53.9$ mph\n$\\tau = 11.9^\\circ$\n\n$\\tau = 7.8^\\circ$\n$V = 47.4$ mph\n$\\tau = 12.5^\\circ$\n\n$\\tau = 5.5^\\circ$\n$V = 43.1$ mph\n$\\tau = 12.4^\\circ$\n\n$\\tau = 5.0^\\circ$\n$V = 38.8$ mph\n$\\tau = 11.8^\\circ$\nNACA\nL-59841\n\n(a) Warped forebody and\nextended afterbody.\n(b) Basic forebody and\nbasic afterbody.\n\nFigure 15.- Spray on tail surfaces during landing at design gross load.\n$\\delta_e = -10^\\circ$.", "timestamp": "2026-07-22T06:29:55.896967+00:00"}
{"citation_id": "19930082918", "source_url": "https://ntrs.nasa.gov/api/citations/19930082918/downloads/19930082918.pdf", "page_number": 31, "total_pages": 62, "image_filename": "19930082918_p31.jpg", "text": "30\nNACA TN 1940\n\nTABLE 2\n\nEFFECT OF AGING ON WIDTH OF (111) DIFFRACTION LINE OF LOW-CARBON\nN-155 ALLOY SOLUTION-TREATED 10 HOURS AT 2200° F\nAND WATER-QUENCHED\n[Copper K$\\alpha_1\\alpha_2$ radiation]\n\n| Aging temperature ($^\\circ$F) | Aging time (hr) | Line width, $W/W_o$ (1) | Average width | Mean deviation from average |\n| :--- | :--- | :--- | :--- | :--- |\n| 1400 | 1.0 | 0.97<br>1.04<br>1.01 | 1.01 | $\\pm$0.02 |\n| | 3.0 | 1.04<br>.97 | 1.00 | $\\pm$.02 |\n| | 10.0 | .99<br>.96 | .98 | $\\pm$.02 |\n| | 30.0 | 1.06<br>.97<br>1.06 | 1.03 | $\\pm$.04 |\n| | 100.0 | 1.00<br>.91<br>.99 | .97 | $\\pm$.04 |\n| 1600 | .5 | .95<br>1.04 | .99 | $\\pm$.04 |\n| | 1.0 | 1.03<br>1.02 | 1.02 | $\\pm$.02 |\n| | 3.0 | 1.03<br>1.15<br>1.04 | 1.07 | $\\pm$.05 |\n| | 10.0 | 1.00<br>1.01<br>1.00 | 1.00 | $\\pm$.0 |\n| | 100.0 | 1.00<br>1.02<br>.91 | .96 | $\\pm$.05 |\n\n$^1$W width of (111) line of specimen aged at indicated time and temperature.\n$W_o$ width of (111) line of unaged material.\nNACA", "timestamp": "2026-07-22T06:29:59.406866+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 60, "total_pages": 114, "image_filename": "19930086061_p60.jpg", "text": "56\nNACA RM L9J07\n\nLeft semispan\nRight semispan\nUpper\nLower\n-3\n-2\n-1\nP\n0\n1\n-3\n-2\n-1\n0\n1\n\n(a) $\\psi = 0^\\circ$\n\nLeft semispan\nRight semispan\nUpper\nLower\n-3\n-2\n-1\nP\n0\n1\n-3\n-2\n-1\n0\n1\n10°\n\n(b) $\\psi = 10^\\circ$\n\nNACA\n\nFigure 17.- Pressure distribution about wing 2 at various angles of yaw;\n$\\alpha = 4.1^\\circ$.", "timestamp": "2026-07-22T06:30:00.790750+00:00"}
{"citation_id": "19930085859", "source_url": "https://ntrs.nasa.gov/api/citations/19930085859/downloads/19930085859.pdf", "page_number": 23, "total_pages": 31, "image_filename": "19930085859_p23.jpg", "text": "```markdown\nNACA RM No. L9B25\n\nDownwash angle, $\\epsilon$, deg\nM = 0.70\n4\n2\n0\n-2\n\nM = 0.80\n4\n2\n0\n-2\n\nM = 0.85\n4\n2\n0\n-2\n\n$\\alpha$, deg -2, -1, 0, 1, 2, 3, 4, 5, 6, 7, 8, 10\n$\\circ$ $\\square$ $\\diamond$ $\\triangle$ $\\triangleright$ $\\nabla$ $\\circ$ $\\diamond$ $\\triangle$ $\\triangleright$ $\\nabla$\n\nDownwash angle, $\\epsilon$, deg\nM = 0.90\n4\n2\n0\n-2\n-80 -40 0 40 80\n\nM = 0.93\n4\n2\n0\n-2\n-80 -40 0 40 80\n\nM = 0.95\n4\n2\n0\n-2\n-80 -40 0 40 80\n\nTail height, $h_t$, percent semispan\nNACA\n\nFigure 9.- Effective downwash angles in region of tail plane for a model with $35^\\circ$ sweptback wing,\naspect ratio 4, taper ratio 0.6, and NACA 65A006 airfoil. Wing alone.\n\n21\n```", "timestamp": "2026-07-22T06:30:04.390175+00:00"}
{"citation_id": "19930083221", "source_url": "https://ntrs.nasa.gov/api/citations/19930083221/downloads/19930083221.pdf", "page_number": 30, "total_pages": 47, "image_filename": "19930083221_p30.jpg", "text": "28\nNACA TN No. 1824\n\nIf equation (46) had been derived for a body of revolution, then\nf(x) would have been independent of the angle $\\mu$ and in that case\nthe expression for drag would reduce to the form\n\n$$D = - \\frac{\\rho_0}{4\\pi} \\int_0^l \\int_0^l f'(x_1)f'(x_2) \\ln|x_1-x_2| dx_1 dx_2 \\quad (46a)$$\n\nThis expression was given by von Kármán in reference 15.\n\nFor the study of the drag of a lifting surface, consider now the\ngeneral expression for the velocity potential given by equation (30).\nThe doublet distribution occupies in this case both the wing plan form\nand the wake since the jump in $\\Phi$ exists also in the vortex wake.\nBy use of the transformations in equation (39), equation (30) becomes\n\n$$\\varphi(x,y,z) = \\frac{\\beta^2 z}{2\\pi} \\sqrt{\\int_\\tau \\int \\frac{\\cos \\mu \\Delta\\varphi(\\xi,\\eta) d\\xi d\\eta}{[(x-x_1)^2 - \\beta^2(y-y_1)^2 - \\beta^2 z^2]^{3/2}}}$$\n\nand, exactly as in the case of the source distribution, this can be\nreduced to\n\n$$\\varphi(x,y,z) = \\frac{\\beta^2 z}{2\\pi} \\sqrt{\\int \\frac{\\cos \\mu dx_1}{[(x-x_1)^2 - \\beta^2 y^2 - \\beta^2 z^2]^{3/2}} \\int \\Delta\\varphi(x_1,\\eta) d\\eta}$$\n\nSetting\n\n$$g(x_1,\\mu) = \\cos \\mu \\int \\Delta\\varphi(x_1,\\eta) d\\eta$$\n\nit follows that\n\n$$\\varphi(x,r,\\theta) = \\frac{\\beta^2 r \\sin \\theta}{2\\pi} \\sqrt{\\int_0^{x-\\beta r} \\frac{g(x_1,\\mu) dx_1}{[(x-x_1)^2 - \\beta^2 r^2]^{3/2}}}$$", "timestamp": "2026-07-22T06:30:05.675892+00:00"}
{"citation_id": "19930082613", "source_url": "https://ntrs.nasa.gov/api/citations/19930082613/downloads/19930082613.pdf", "page_number": 41, "total_pages": 46, "image_filename": "19930082613_p41.jpg", "text": "Page intentionally left blank\n\nPage intentionally left blank", "timestamp": "2026-07-22T06:30:07.554486+00:00"}
{"citation_id": "19930085862", "source_url": "https://ntrs.nasa.gov/api/citations/19930085862/downloads/19930085862.pdf", "page_number": 22, "total_pages": 60, "image_filename": "19930085862_p22.jpg", "text": "```markdown\n20\nNACA RM No. L9A07\n\n[Figure: Diagram of an aircraft wing section showing the geometry of an aileron and spoilers. The top part shows the aileron with dimensions and a cross-section A-A. The bottom part shows the spoilers with dimensions and a cross-section B-B.]\n\nApproximate\nlocation of strain\ngage beams\n\nAileron\n\nHinge axis\n\nChord line\n\nFlexible\nseal\n\nSection A-A\n(enlarged)\n\nSpoilers\n\nSection B-B\n(enlarged)\n\nNACA\n\nFigure 4.- Geometry of aileron and spoilers.\n```", "timestamp": "2026-07-22T06:30:08.877003+00:00"}
{"citation_id": "19930082546", "source_url": "https://ntrs.nasa.gov/api/citations/19930082546/downloads/19930082546.pdf", "page_number": 47, "total_pages": 65, "image_filename": "19930082546_p47.jpg", "text": "46\nNACA TN No. 1870\n\nPressure, p, dynes/cm²\n2000\n1600\n1200\n800\n400\n0\n\nmB\n2\n4\n6\n8\n\n0 .02 .04 .06 .08 .10\n$C_p$\n\n(a) $M_t = 0.75$.\n\nFigure 11.- Effect of power coefficient on the relative pressure amplitudes of the first four harmonics of the NACA 4-(5)(08)-03 two-blade propeller in the plane of rotation. $\\frac{d}{D} = 0.083$.", "timestamp": "2026-07-22T06:30:09.723923+00:00"}
{"citation_id": "19930085847", "source_url": "https://ntrs.nasa.gov/api/citations/19930085847/downloads/19930085847.pdf", "page_number": 21, "total_pages": 32, "image_filename": "19930085847_p21.jpg", "text": "NACA RM A9D04\nCONFIDENTIAL\n19\n\nDrag-producing area\nThrust-producing area\n\n<!-- Image (282, 100, 759, 847) -->\n\n(b) M = 0.73 ($C_L = 0.20$).\nFigure 7.- Continued.\nCONFIDENTIAL", "timestamp": "2026-07-22T06:30:10.584903+00:00"}
{"citation_id": "19930082592", "source_url": "https://ntrs.nasa.gov/api/citations/19930082592/downloads/19930082592.pdf", "page_number": 49, "total_pages": 50, "image_filename": "19930082592_p49.jpg", "text": "**Page intentionally left blank**\n\n**Page intentionally left blank**", "timestamp": "2026-07-22T06:30:16.007038+00:00"}
{"citation_id": "19930085890", "source_url": "https://ntrs.nasa.gov/api/citations/19930085890/downloads/19930085890.pdf", "page_number": 11, "total_pages": 26, "image_filename": "19930085890_p11.jpg", "text": "10\nNACA RM No. E9C11\n\nSUMMARY OF RESULTS\n\nExperiments were conducted to determine the performance of a 100-pound-thrust rocket engine at a combustion-chamber pressure of 300 pounds per square inch absolute using liquid diborane and liquid oxygen as propellants. The engine used for most of the experiments had an injection system of four pairs of intersecting jets and a ratio of combustion-chamber volume to exhaust-nozzle-throat area of 325 inches. The experiments produced the following results:\n\n1. The shape of the experimental specific-impulse curve was similar to that of the theoretical curve. Both curves reached a maximum near a ratio of fuel weight to total propellant weight of 0.36. The maximum uncorrected experimental specific impulse, as indicated by a faired curve through the data, was 249 pound-seconds per pound or approximately 83 percent of the theoretical value for equilibrium expansion and the fuel and the nozzle used.\n\n2. When corrections for heat loss to the engine walls and small deviations of combustion-chamber pressure were applied to the experimental data, the maximum performance value was about 274 pound-seconds per pound or 92 percent of the theoretical value for equilibrium expansion and the fuel and the nozzle used.\n\n3. The maximum uncorrected experimental volume specific impulse was $192 \\times 62.4$ pound-seconds per cubic foot and the corresponding value when corrected for heat loss and pressure deviations was $199 \\times 62.4$ pound-seconds per cubic foot. The maximum volume specific impulse occurred near a ratio of fuel weight to total propellant weight of 0.25. The theoretical curve had a maximum value of $217 \\times 62.4$ pound-seconds per cubic foot at a mixture ratio of 0.22 for the fuel and the nozzle used.\n\n4. When the injection system was changed from the eight-hole, solid jet system to the four-hole system, there was no apparent change in specific impulse. When the characteristic length was changed from 325 to 159 inches (using the eight-hole system), there was a small decrease in performance. The study of the effects of the injection system and combustion volume were not sufficiently extensive to determine optimum configurations. Several experiments made to determine the sensitivity of diborane to temperature and to detonation produced no explosions nor detonations.\n\nNational Advisory Committee for Aeronautics,\nLewis Flight Propulsion Laboratory,\nCleveland, Ohio.", "timestamp": "2026-07-22T06:30:16.332159+00:00"}
{"citation_id": "19930082646", "source_url": "https://ntrs.nasa.gov/api/citations/19930082646/downloads/19930082646.pdf", "page_number": 33, "total_pages": 37, "image_filename": "19930082646_p33.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:30:19.812136+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 49, "total_pages": 78, "image_filename": "19930082483_p49.jpg", "text": "NACA TN No. 1807\n\nTherefore\n\n$$\np_e = 10.80 \\text{ in. Hg absolute}\n$$\n\n5. Ratio of inlet total pressure to discharge static pressure by equation (46)\n\n$$\n\\frac{p_1'}{p_e} = \\frac{29.92}{10.80}\n$$\n\n$$\n= 2.770\n$$\n\n6. Efficiency based on ratio of inlet total pressure to discharge static pressure by equation (47)\n\n$$\n\\eta = \\frac{P}{\\left( \\frac{W \\Delta_s h}{0.707} \\right)}\n$$\n\nFrom tables in reference 8,\n\n$$\n\\Delta_s h = 32.01 \\text{ Btu/lb}\n$$\n\nTherefore\n\n$$\n\\eta = \\frac{388}{\\left[ \\frac{(15.29)(32.01)}{0.707} \\right]}\n$$\n\n$$\n= 0.560\n$$\n\n7. Blade pitch-line velocity $u_m$ from equation (48)\n\n$$\nu_m = \\pi D \\left( \\frac{N}{60} \\right)\n$$\n\nwhere\n\nD 1.167 ft  \nN 8650 rpm  \n\nTherefore\n\n$$\nu_m = 528 \\text{ ft/sec}\n$$", "timestamp": "2026-07-22T06:30:20.623661+00:00"}
{"citation_id": "19930085900", "source_url": "https://ntrs.nasa.gov/api/citations/19930085900/downloads/19930085900.pdf", "page_number": 8, "total_pages": 33, "image_filename": "19930085900_p8.jpg", "text": "NACA RM L9D20 CONFIDENTIAL 7\n\nWhen the strips were used in the form of V-steps pointed forward, the after half of the model was sucked under and a large amount of spray was thrown out to either side. The spray characteristics for the V-steps pointed aft were about the same as for the chine strips; only a small amount of spray was thrown out in an almost horizontal direction.\n\nWhen the hypotenuse formed the after side of each strip and the forward side was perpendicular to the fuselage bottom, the results obtained were nearly the same as shown in figure 16 although the resistance was generally slightly higher.\n\nResults reported in reference 1 showed similar hydrodynamic characteristics among the three multiple-step configurations when jets were used. When strips were substituted for jets in these multiple-step configurations the V-steps pointed forward gave results entirely different from the other two. It appears that the effect of strips on the hydrodynamic characteristics of the fuselage was more dependent on the configuration used than was the effect of jets.\n\nCONCLUSIONS\n\nThe results of model tests to determine the effect of various jet and strip modifications on the hydrodynamic characteristics of a streamline fuselage indicate the following conclusions:\n\n1. The resistance was decreased as the jet spacing was decreased or the length of the jet rows simulating chines was increased.\n\n2. Substitution of chine jets slanted $45^\\circ$ aft for jets normal to the center line increased the trim but had little effect on the resistance.\n\n3. As the average air flow per jet was increased, the resistance was reduced at a decreasing rate until at flows greater than $11 \\times 10^{-5}$ pounds per second the resistance remained practically constant.\n\n4. In the chine configuration, strips protruding less than 2 percent of the maximum fuselage diameter gave about the same resistance and trim as rows of $\\frac{1}{4}$-inch-spaced jets, but the spray characteristics for the strips were better.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:30:22.994867+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 61, "total_pages": 114, "image_filename": "19930086061_p61.jpg", "text": "NACA RM L9J07\n57\n\nLeft semispan\nRight semispan\nUpper\nLower\nP\n120°\n(c) $\\psi = 20^\\circ$\n\nLeft semispan\nRight semispan\nUpper\nLower\nP\n135°\n(d) $\\psi = 35^\\circ$\n\nFigure 17.- Concluded.", "timestamp": "2026-07-22T06:30:27.902686+00:00"}
{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 32, "total_pages": 72, "image_filename": "19930085491_p32.jpg", "text": "NACA RM No. A8J04 CONFIDENTIAL 31\n\nthat the maximum value will occur where the rate of decrease of minimum drag coefficient is equal to the rate of increase in drag-rise factor, since at this sweep angle the rate of change in $(L/D)_{\\text{max}}$ (equation (3)) is then zero. At the optimum leading-edge sweep angle ($67^\\circ$) for the wings of the present study, additional data were obtained at the highest possible test Reynolds number, 0.95 million, and a value of maximum lift-drag ratio of 7.4 was obtained and is indicated in figure 10(d). This increase from 7.1 at a Reynolds number of 0.62 million was found to result from decreases in both minimum drag coefficient and drag-rise factor as was noted in the discussion of the Reynolds number effect on WF-63.\n\nThe optimum lift coefficient decreases as the sweep angle increases as shown in figure 10(d). The reason for this variation is apparent from a consideration of equation (4) and the variations of $C_{D_{\\text{min}}}$ and $\\Delta C_D / (\\Delta C_L)^2$ previously discussed.\n\nEffect of sweep on pitching moment.— The variation of pitching-moment coefficient and center-of-pressure location are plotted in figure 8 for the positive range of lift coefficients for all configurations in the investigation. It will be noted that the variations of moment coefficient and center-of-pressure position with lift coefficient for all configurations is similar to that for WF-63 which has been previously discussed. The effect of sweep on the center-of-pressure travel is shown by a comparison of figure 8(a) for WF-57 and figure 8(f) for WF-70. For WF-57, the maximum percent travel was about 8 percent of the mean aerodynamic chord in a lift-coefficient range of 0.18, while WF-70 shows 21-percent travel in a lift-coefficient range of only 0.13. Since the actual mean aerodynamic chord length (table I) of WF-70 is greater than that of WF-57, the absolute center-of-pressure travel is even larger than that indicated by the difference in percent travel. The effect of increasing sweep on the center-of-pressure travel is thus unfavorable. Although it was expected that there might be some change in the pitching-moment characteristics as the trailing edge passed through the Mach cone for $M_0 = 1.53$, no such effect was noted.\n\nSCHLIEREN PHOTOGRAPHS\n\nAs might be expected there is a correlation between the shock-wave pattern behind the wing of each configuration and the boundary-layer flow on the wing surfaces. The location of the compression wave that exists behind the trailing edge was found to be dependent upon the area of separated flow and, therefore, also is affected by\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:30:31.386385+00:00"}
{"citation_id": "19930085899", "source_url": "https://ntrs.nasa.gov/api/citations/19930085899/downloads/19930085899.pdf", "page_number": 9, "total_pages": 29, "image_filename": "19930085899_p9.jpg", "text": "8\nNACA RM No. L9A21\n\nTABLE I.- FUSELAGE ORDINATES\n[Basic fineness ratio 12; actual fineness ratio 10\nachieved by cutting off the rear one-sixth of\nthe body; 5/4 located at 1/2]\n\n[Figure: Diagram of a fuselage shape with dimensions labeled: $l=14.14$, $\\frac{5}{6}l$, $\\frac{2}{2}$, $x$, $r$, $D(Max)$]\n\nOrdinates\n| $x/l$ | $r/l$ | $x/l$ | $r/l$ |\n| :--- | :--- | :--- | :--- |\n| 0 | 0 | 0 | 0 |\n| .005 | .00231 | .4500 | .04143 |\n| .0075 | .00298 | .5000 | .04167 |\n| .0125 | .00428 | .5500 | .04130 |\n| .0250 | .00722 | .6000 | .04024 |\n| .0500 | .01205 | .6500 | .03842 |\n| .0750 | .01613 | .7000 | .03562 |\n| .1000 | .01971 | .7500 | .03128 |\n| .1500 | .02593 | .8000 | .02526 |\n| .2000 | .03090 | .8338 | .02000 |\n| .2500 | .03465 | .8500 | .01852 |\n| .3000 | .03741 | .9000 | .01125 |\n| .3500 | .03933 | .9500 | .00439 |\n| .4000 | .04063 | 1.0000 | 0 |\n\nL. E. radius = 0.00051\n\nNACA", "timestamp": "2026-07-22T06:30:31.849759+00:00"}
{"citation_id": "19930085859", "source_url": "https://ntrs.nasa.gov/api/citations/19930085859/downloads/19930085859.pdf", "page_number": 24, "total_pages": 31, "image_filename": "19930085859_p24.jpg", "text": "```markdown\n22\n\nDownwash angle, $\\epsilon$, deg\n\nM = 0.98\nM = 1.00\nM = 1.03\n\n$\\alpha$, deg -2, -1, 0, 1, 2, 3, 4, 5, 6, 7, 8, 10\n\n$\\circ$ $\\square$ $\\diamond$ $\\triangle$ $\\nabla$ $\\triangleright$ $\\triangleleft$ $\\circ$ $\\circ$ $\\diamond$ $\\triangle$ $\\nabla$ $\\triangleright$\n\nDownwash angle, $\\epsilon$, deg\n\nM = 1.05\nM = 1.10\nM = 1.15\n\n- 80 - 40 0 40 80\n- 80 - 40 0 40 80\n- 80 - 40 0 40 80\n\nTail height, $h_t$, percent semispan\n\n[Figure: NACA logo]\n\nFigure 9.— Concluded.\n\nNACA RM No. 19B25\n```", "timestamp": "2026-07-22T06:30:32.211893+00:00"}
{"citation_id": "19930085862", "source_url": "https://ntrs.nasa.gov/api/citations/19930085862/downloads/19930085862.pdf", "page_number": 23, "total_pages": 60, "image_filename": "19930085862_p23.jpg", "text": "```markdown\nNACA RM No. L9A07\n21\n\n<!-- Image (217, 112, 827, 863) -->\n\n(a) $C_l$, $C_n$, and $C_{Na}$ against $\\alpha$.\nFigure 5.- Aileron characteristics of plain wing.\n```", "timestamp": "2026-07-22T06:30:34.072600+00:00"}
{"citation_id": "19930082546", "source_url": "https://ntrs.nasa.gov/api/citations/19930082546/downloads/19930082546.pdf", "page_number": 48, "total_pages": 65, "image_filename": "19930082546_p48.jpg", "text": "NACA TN No. 1870\n47\n\nPressure, p, dynes/cm²\n2000\n1600\n1200\n800\n400\n0\n0 .02 .04 .06 .08 .10\n$\\sigma_p$\n\nmB\nO ———— 2\n□ - - - - - 4\n◇ ———— 6\n△ - - - - - 8\n\n(b) $M_t = 0.90$.\nFigure 11.— Continued.", "timestamp": "2026-07-22T06:30:38.117638+00:00"}
{"citation_id": "19930085870", "source_url": "https://ntrs.nasa.gov/api/citations/19930085870/downloads/19930085870.pdf", "page_number": 14, "total_pages": 92, "image_filename": "19930085870_p14.jpg", "text": "NACA RM No. L9D07 CONFIDENTIAL 13\n\nIf these shocks be traced forward, the apparent point of origin will be found at a point between the apex of the ridge line and the forward tip of the sting, being nearer the former. As the wings are given angle of attack, these shocks separate into two distinct shocks, neither of which occupies the position in relation to the wing tips that occurred for the $\\alpha = 0^\\circ$ condition. One shock has moved inboard and the other outboard. The rate of outward travel with angle of attack for the outboard shock is much greater than the rate of inward travel for the inner shock. For wedge-leading-edge wing 5, tracing the shocks forward places the apparent point of origin aft of the ridge-line apex and well ahead of the forward tip of the sting. For wedge-leading-edge wing 11, tracing the inboard shock at $\\alpha = 4^\\circ$ produces a point of origin aft of the sting tip while the outer shock continues to maintain a point of origin between the sting tip and the ridge-line apex. Thus the sting may be eliminated as a source of these shocks. Comparison of the photographs of the elliptical-leading-edge wings (fig. 20) and the corresponding photographs of the wedge-leading-edge wings (fig. 16(b)) shows that the shocks leave the trailing edge of the elliptical-leading-edge wing slightly further inboard than on the wedge-leading-edge wing. This would seem to indicate that the shock origin for the elliptical leading edge was behind that for the wedge-leading-edge wing. Tracing of the shock on elliptical-leading-edge wing 11 at $\\alpha = 0^\\circ$ yields the apparent point of origin well aft of the sting tip; whereas, for the same condition of the wedge-leading-edge wing, the apparent point of origin lies ahead of the sting tip. The shocks are evidently produced by second-order compressibility effects similar to those observed on unswept wings at transonic speeds. It is possible that thickness distribution, leading-edge shape, and ridge-line angularity are predominant factors in formation and location of the shocks. The easy curvature of the ridge line of the elliptical-leading-edge wings would probably favor a delay in formation of the shocks. As stated previously, a relatively large exposure time was necessary for the schlieren photographs of the elliptical-leading-edge wings. This probably explains the appearance of the shed vortices in these photographs.\n\nThe liquid-film patterns for wings 5 and 11 are shown in figures 17(a), 17(b), and 17(d). In contrast to wing 5, wing 11 shows the area of large shear intensity near the leading edge to extend even aft of the ridge line for both the wedge- and elliptical-leading-edge configurations. This is probably associated with the higher component of free-stream velocity normal to the leading edge of wing 11. The sequence of liquid-film photographs presented in figure 17(d) shows the progressive shifting of the transition line on both upper and lower surfaces with angle of attack for wing 5. The difference in absolute location of the transition lines on upper and lower surfaces at other than zero angle of attack is practically the same as the difference in location of the two shocks observed in the schlieren photographs. In addition, the location and curvature of the transition line shown on each surface at angle of attack may be\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:30:42.524294+00:00"}
{"citation_id": "19930085548", "source_url": "https://ntrs.nasa.gov/api/citations/19930085548/downloads/19930085548.pdf", "page_number": 31, "total_pages": 46, "image_filename": "19930085548_p31.jpg", "text": "30\nNACA RM No. E8L30\n\nPressure drop across cylinder, lb/sq in.\nInlet-manifold\npressure\n(lb/sq in. abs.)\n$\\circ$ 80\n$\\diamond$ 100\n$\\diamond$ 120\n$\\Delta$ 135\n\nFuel-air ratio\n\n[Figure: Graph showing pressure drop across cylinder vs fuel-air ratio for different inlet-manifold pressures]\n\nFigure 8. - Effect of inlet-manifold pressure and fuel-air ratio on cylinder pressure drop. Compression ratio, 5.25; inlet-manifold temperature, 400° F.", "timestamp": "2026-07-22T06:30:44.934673+00:00"}
{"citation_id": "19930085847", "source_url": "https://ntrs.nasa.gov/api/citations/19930085847/downloads/19930085847.pdf", "page_number": 22, "total_pages": 32, "image_filename": "19930085847_p22.jpg", "text": "20\nCONFIDENTIAL\nNACA RM A9D04\n\nDrag-producing area\nThrust-producing area\n\n<!-- Image (231, 101, 702, 873) -->\n\n(c) $M = 0.76$ ($C_L = 0.19$).\n\nFigure 7.-Continued.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:30:44.946285+00:00"}
{"citation_id": "19930082592", "source_url": "https://ntrs.nasa.gov/api/citations/19930082592/downloads/19930082592.pdf", "page_number": 50, "total_pages": 50, "image_filename": "19930082592_p50.jpg", "text": "NACA TN 1914\n49\n\nScale fracture\nSingle-phase\nsurface layer\nInner-\ntransition\nzone\nTwo-phase\ninner layer\nUnoxidized body\n\nNACA\nC-22918\n2-7-49\n\nFigure 19. - Oxidation zone of 30-percent-cobalt - titanium carbide ceramal. This view\nillustrates complex nature of oxide coatings found on cobalt-bearing ceramals. Tem-\nperature, 1625° F; time at temperature, 100 hours; etchant, potassium hydroxide plus\npotassium ferricyanide KOH+K$_3$Fe(CN)$_6$; magnification, X250.\n\nNACA-Langley - 7-18-49 - 850", "timestamp": "2026-07-22T06:30:49.085095+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 30, "total_pages": 149, "image_filename": "19930083192_p30.jpg", "text": "26\nNACA TN 1976\n\nThe values of $C_{L_\\alpha}$ and $C_{L_g}$ for the infinite-aspect-ratio wing are presented in figure 26. If the pitching motion of the airplane is not considered, the exact shapes of the unsteady-lift functions are unimportant and the discrepancies noted in figure 25 do not appear to be significant.\n\nFinite aspect ratio.- The effect of finite aspect ratio on the unsteady-lift functions has been investigated by Jones in reference 14 and these results are shown in figure 26. This figure indicates that, as the aspect ratio decreases, the lift develops more rapidly, particularly at the start, until for the limiting case of zero aspect ratio the variation in lift might be expected to correspond to that for steady flow conditions. The estimation of the unsteady-lift functions for lower aspect ratios than those shown in figure 26 has been made by extrapolating the data by assuming that at zero aspect ratio, the unsteady-lift function disappears. Unpublished tests made with a flying wing having an aspect ratio of about 1.27 showed that this method led to reasonable values. The calculated value of the acceleration ratio was 0.92 as compared to an average experimental value of 0.93. These results indicate, therefore, that for flying wings finite theory applies. Other results, shown in figure 27 and given by Keuthe in reference 17 and by Sears and Keuthe in reference 26, indicate that the unsteady lift developed on a low-aspect-ratio wing subjected to a gust more closely approximates the unsteady-lift function for the two-dimensional case than that for aspect ratio 3.\n\nAn indirect verification of unsteady-lift functions has been obtained by comparing the calculated and actual response of airplane models to a sharp gust. The curves in figure 28 show the calculated acceleration ratio $\\Delta n/\\Delta n_g$ as a function of the mass parameter $\\mu_g$. The calculation for aspect ratio 6 is based on the unsteady-lift function given in figure 26 and that for infinite aspect ratio is based on the unsteady-lift functions given by Rhode in reference 4. The test points shown in figure 28 represent experimental values of the acceleration ratio determined from tests in both the old and the new gust tunnels. The results are for all types of airplanes from tailless high-speed configurations with wings swept back 35° to transports. As can be seen from figure 28, the experimental data scatter about the curve based on the unsteady-lift functions for aspect ratio $\\infty$ and below the curve for aspect ratio 6. The results indicate that, for airplanes that include a fuselage, the use of the unsteady functions for infinite aspect ratio have given adequate and in most cases more accurate predictions of the acceleration ratio than the computations based on the finite-aspect-ratio theory of Jones.\n\nThe discrepancy between experiment and theory was considered as to the effect of interference on the experimental results. If the fuselage is considered as an elongated wing having a chord three times that", "timestamp": "2026-07-22T06:30:51.629013+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 50, "total_pages": 78, "image_filename": "19930082483_p50.jpg", "text": "48\nNACA TN No. 1807\n\n8. Theoretical jet velocity $V_j'$ based on ratio of inlet to discharge stagnation pressures from equation (49)\n\n$$V_j' = \\sqrt{2gJc_p T_1'} \\left[ 1 - \\frac{1}{\\left(\\frac{p_1'}{p_e'}\\right)^{\\frac{\\gamma-1}{\\gamma}}} \\right]$$\n\nwhere\nJ 778 ft-lb/Btu\n$c_p$ 0.240 Btu/(lb)($^\\circ$F)\n\ntherefore\n\n$$V_j' = 1092 \\text{ ft/sec}$$\n\n9. Theoretical jet velocity $V_j$ based on ratio of inlet total pressure to discharge static pressure by equation (50)\n\n$$V_j = \\sqrt{2gJc_p T_1'} \\left[ 1 - \\frac{1}{\\left(\\frac{p_1'}{p_e}\\right)^{\\frac{\\gamma-1}{\\gamma}}} \\right]$$\n\nTherefore\n\n$$V_j = 1255 \\text{ ft/sec}$$\n\n10. Velocity ratio $v'$ based on ratio of inlet total pressure to discharge total pressure by equation (51)\n\n$$v' = \\frac{u_m}{V_j'}$$\n\n$$= \\frac{528}{1092}$$\n\n$$= 0.4835$$", "timestamp": "2026-07-22T06:30:54.126359+00:00"}
{"citation_id": "19930082646", "source_url": "https://ntrs.nasa.gov/api/citations/19930082646/downloads/19930082646.pdf", "page_number": 34, "total_pages": 37, "image_filename": "19930082646_p34.jpg", "text": "NACA TN 1980\n33\n\nWarped forebody and extended afterbody ———\nBasic forebody and basic afterbody - - - - - - -\n\n[Figure: Graph showing Gross load, lb (y-axis, 40 to 90 x 10^3) vs Speed, mph (x-axis, 0 to 50). Two curves are plotted: a solid line and a dashed line. Regions are labeled \"Spray in propellers\", \"Propellers clear\", and \"Clear\". A NACA logo is present in the bottom right corner of the graph area.]\n\nFigure 16.- Variation of range of speed with gross load for spray in propellers during taxying in waves.", "timestamp": "2026-07-22T06:30:54.357765+00:00"}
{"citation_id": "19930082918", "source_url": "https://ntrs.nasa.gov/api/citations/19930082918/downloads/19930082918.pdf", "page_number": 32, "total_pages": 62, "image_filename": "19930082918_p32.jpg", "text": "NACA TN 1940\n31\n\nTABLE 3\nINTERNAL ROOT-MEAN-SQUARE STRAINS FOR LOW-CARBON M-155 ALLOY\n\n| Aging temperature (°F) | Aging time (hr) | Lattice parameter (Å) | d (Å) | cot θ | K | B (radians) | $\\frac{\\overline{\\Delta d}}{d}$ (a) |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| 1400 | 0.5 | 3.5862 | $b_{1.27}$ | $b_{4.4843}$ | $b_{6.15} \\times 10^{-4}$ | $b_0$ | 0 |\n| | 1.0 | 3.5874 | $b_{1.27}$ | $b_{4.4843}$ | $b_{6.15}$ | $b_{1.47} \\times 10^{-3}$ | $1.40 \\times 10^{-3}$ |\n| | 3.0 | 3.5876 | $b_{1.270}$ | $b_{4.4843}$ | $b_{6.15}$ | $b_0$ | 0 |\n| | 10.0 | 3.5898 | $b_{1.269}$ | $b_{4.4843}$ | $b_{6.15}$ | $b_{2.10}$ | 2.0 |\n| | 30.0 | 3.5840 | $b_{1.267}$ | $b_{4.4830}$ | $b_{6.14}$ | $b_{4.53}$ | 4.31 |\n| | 100.0 | 3.5814 | $b_{1.266}$ | $b_{4.4820}$ | $b_{6.12}$ | $b_{5.18}$ | 4.92 |\n| | 1000.0 | 3.5770 | $c_{1.000}$ | $c_{.6724}$ | $c_{7.26}$ | $c_{5.95}$ | 6.25 |\n| 1600 | 3.0 | 3.583 | $c_{1.082}$ | $c_{.6851}$ | $c_{7.41}$ | $c_{4.10}$ | 4.70 |\n| | 100.0 | 3.581 | $c_{1.081}$ | $c_{.6847}$ | $c_{7.40}$ | $c_{3.66}$ | 4.20 |\n\nSolution-treated 10 hours at 2200° F, water-quenched, and aged as indicated\n\n| | | | | | | | |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| | | 3.587 | $b_{1.27}$ | $b_{0.4843}$ | $b_{6.15} \\times 10^{-4}$ | $b_{11.5} \\times 10^{-3}$ | $10.90 \\times 10^{-3}$ |\n\nSolution-treated 1 hour at 2200° F, water-quenched, and rolled to 15 percent reduction in cross section at 2400° F\n\n$$\n\\frac{\\overline{\\Delta d}}{d} = 1.75 |K| B \\times 10^3\n$$\n\nwhere $|K| = |d \\cot \\theta \\times 10^{-3}|$\n$B = \\sqrt{B_0^2 - B_g^2}$ radians $= 0.00245 \\sqrt{W_0^2 - W_g^2}$ radians\n$W_0$ half width of indicated line from microphotometer plot, cm\n$W_g$ half width of standard line from microphotometer plot, cm\n\n[Figure: NACA logo]\n\nFrom (220) line and chromium $K_{\\alpha 1}$ radiation.\nFrom (311) line and cobalt $K_{\\alpha 1}$ radiation.", "timestamp": "2026-07-22T06:30:56.146883+00:00"}
{"citation_id": "19930085900", "source_url": "https://ntrs.nasa.gov/api/citations/19930085900/downloads/19930085900.pdf", "page_number": 9, "total_pages": 33, "image_filename": "19930085900_p9.jpg", "text": "8\nCONFIDENTIAL\nNACA RM L9D20\n\n5. With similar strips arranged as V-steps pointed forward, the\nresistance and trim were very high. For V-steps pointed aft, the\nresistance and trim were considerably lower than for either strips or\njets in the chine configuration.\n\nLangley Aeronautical Laboratory\nNational Advisory Committee for Aeronautics\nLangley Air Force Base, Va.\n\nREFERENCE\n\n1. Weinflash, Bernard: The Effect of Air Jets Simulating Chines or\nMultiple Steps on the Hydrodynamic Characteristics of a Stream-\nline Fuselage. NACA RM L8J21, 1948.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:30:57.058012+00:00"}
{"citation_id": "19930083221", "source_url": "https://ntrs.nasa.gov/api/citations/19930083221/downloads/19930083221.pdf", "page_number": 31, "total_pages": 47, "image_filename": "19930083221_p31.jpg", "text": "NACA TN No. 1824\n29\n\nIntegrating by parts and using the fact that $g(0,\\mu)=0$\n\n$$\nu = \\frac{\\partial \\varphi}{\\partial x} = \\frac{\\sin \\theta}{2\\pi r} \\int_{0}^{x-\\beta r} \\frac{\\beta^2 r^2 g'(x_1,\\mu) dx_1}{[(x-x_1)^2 - \\beta^2 r^2]^{3/2}} \\tag{47}\n$$\n\nand\n\n$$\nv_r = \\frac{\\partial \\varphi}{\\partial r} = \\frac{\\sin \\theta}{2\\pi r} \\left\\{ \\frac{1}{r} \\int_{0}^{x-\\beta r} \\frac{(x-x_1) g'(x_1,\\mu) dx_1}{\\sqrt{(x-x_1)^2 - \\beta^2 r^2}} - \\int_{0}^{x-\\beta r} \\frac{(x-x_1) \\beta^2 r g'(x_1,\\mu) dx_1}{[(x-x_1)^2 - \\beta^2 r^2]^{3/2}} \\right\\} \\tag{48}\n$$\n\nwhere $g'(x_1,\\mu)$ indicates $\\frac{\\partial}{\\partial x_1} g(x_1,\\mu)$. Setting $x = x_0 + \\beta r$ and letting $r$ approach infinity, the asymptotic expressions for equations (47) and (48) become, if $g'(0,\\mu) = 0$,\n\n$$\nu = \\frac{-\\sin \\theta}{2\\pi} \\sqrt{\\frac{\\beta}{2r}} \\int_{0}^{x_0} \\frac{g''(x_1,\\mu) dx_1}{\\sqrt{x_0 - x_1}} \\tag{49}\n$$\n\nand\n\n$$\nv_r = \\frac{\\sin \\theta}{2\\pi} \\beta \\sqrt{\\frac{\\beta}{2r}} \\int_{0}^{x_0} \\frac{g''(x_1,\\mu) dx_1}{\\sqrt{x_0 - x_1}} \\tag{50}\n$$\n\nThe relations just derived may be used in conjunction with equation (38) to give the wave drag of a lifting surface. This result takes the form\n\n$$\nD = \\frac{\\beta^2 \\rho_0}{8\\pi^2} \\int_{0}^{2\\pi} \\sin^2 \\theta \\int_{0}^{\\infty} dx_0 \\int_{0}^{x_0} \\frac{g''(x_1,\\mu) dx_1}{\\sqrt{x_0 - x_1}} \\int_{0}^{x_0} \\frac{g''(x_2,\\mu) dx_2}{\\sqrt{x_0 - x_2}} \\tag{51}\n$$", "timestamp": "2026-07-22T06:30:59.451195+00:00"}
{"citation_id": "19930085899", "source_url": "https://ntrs.nasa.gov/api/citations/19930085899/downloads/19930085899.pdf", "page_number": 10, "total_pages": 29, "image_filename": "19930085899_p10.jpg", "text": "NACA RM No. E9J21\n\n0.25-chord line\nCenter of gravity (0.25 M.A.C.)\n(MAC)\n2.166\n0.541\n90°\n2.652\n2.607\n7.07\n11.8\n1.18 D Max\n2.25\n-Wing-fuselage end plate\n-Wing-alone end plate\n1.05\n0.56\nBump surface\nCenterline of balance\nnormal to bump surface\n1.591\n4.242\nReference centerline\n0.52\n\nTABULATED DATA\nWing\nTwice semispan area 0.125 sq ft\nAspect ratio 4.0\nTaper ratio 0.60\nMean aerodynamic chord 0.180 ft\nIncidence 0°\nDihedral 0°\nAirfoil section parallel to\nfree airstream NACA 65A006\n\n0 1 2\nScale, inches\nNACA\n\nFigure 1.- General arrangement of model with 45° sweptback wing, aspect ratio 4, taper ratio 0.6,\nand NACA 65A006 airfoil.", "timestamp": "2026-07-22T06:30:59.627482+00:00"}
{"citation_id": "19930085859", "source_url": "https://ntrs.nasa.gov/api/citations/19930085859/downloads/19930085859.pdf", "page_number": 25, "total_pages": 31, "image_filename": "19930085859_p25.jpg", "text": "NACA RM NO. L9B25\n\nDownwash angle, $\\epsilon$, deg\n\nM = 0.70\n\nM = 0.80\n\nM = 0.85\n\n$\\alpha$, deg -2, -1, 0, 1, 2, 3, 4, 6, 8, 10\n\nDownwash angle, $\\epsilon$, deg\n\nM = 0.90\n\nM = 0.93\n\nM = 0.95\n\nTail height, $h_t$, percent semispan\n\nFigure 10.- Effective downwash angles in region of tail plane for a model with $35^\\circ$ sweptback wing, aspect ratio 4, taper ratio 0.6, and NACA 65A006 airfoil. Wing-fuselage.\n\n23", "timestamp": "2026-07-22T06:31:01.753948+00:00"}
{"citation_id": "19930085862", "source_url": "https://ntrs.nasa.gov/api/citations/19930085862/downloads/19930085862.pdf", "page_number": 24, "total_pages": 60, "image_filename": "19930085862_p24.jpg", "text": "```markdown\n22\nNACA RM No. L9A07\n\n<!-- Image (172, 116, 775, 877) -->\n\n(b) $C_{h_a}$ and $P_R$ against $\\alpha$.\nFigure 5.— Continued.\n```", "timestamp": "2026-07-22T06:31:03.843404+00:00"}
{"citation_id": "19930085890", "source_url": "https://ntrs.nasa.gov/api/citations/19930085890/downloads/19930085890.pdf", "page_number": 12, "total_pages": 26, "image_filename": "19930085890_p12.jpg", "text": "NACA RM No. E9C11\n\nAPPENDIX - SENSITIVITY OF DIBORANE TO TEMPERATURE AND SHOCK\n\nFor the purpose of establishing a safe handling procedure for diborane, the sensitivity of diborane to temperature and to the shock produced by the detonation of a number 6 detonator cap was determined.\n\nA schematic diagram of the equipment for determining the temperature sensitivity of diborane is shown in figure 9(a). A stainless-steel bomb, 1 inch in diameter and $4\\frac{1}{2}$ inches in length, fitted with a Teflon-packed needle valve and stainless-steel safety disk tested to a bursting pressure of 3700 pounds per square inch, contained the diborane as a gas. Asbestos-covered resistance wire was wrapped around the container and the temperature was measured by means of two thermocouples placed on the cylinder. In addition, the pressure increase during the heating was observed by means of a steel Bourdon-tube pressure gage. The temperature-sensitivity apparatus was filled with gaseous diborane and heated to approximately $1000^\\circ$ F over a period of about 1/2 hour.\n\nFor the three temperature-sensitivity determinations made, the only reaction observed was a gradual pressure increase from 60 pounds per square inch gage at dry-ice temperature to 280 pounds per square inch gage at approximately $1000^\\circ$ F, with the pressure remaining constant at any intermediate temperature for a few minutes. These experiments indicate no explosions resulted from gradual application of heat.\n\nThe sketch of the detonation experimental apparatus is shown in figure 9(b). Two brass containers having volumes of 50 and 60 cubic centimeters were used to hold the diborane. A thin brass separating tube for containing the detonator cap was centrally located in the diborane container.\n\nPrior to the experiments with diborane, preliminary detonation work was performed with the detonator cap alone, and with the cap and water, methyl alcohol, and a mixture of 80-percent tetranitromethane and 20-percent nitrobenzene. For the investigation of diborane, two experiments were made with gaseous diborane in the 30-cubic-centimeter container; one experiment was made with 13.1 grams of liquid diborane (at a temperature of approximately $-72^\\circ$ C) in the 30-cubic-centimeter container and one experiment was made with 24.6 grams of liquid diborane (at a temperature of approximately $-72^\\circ$ C) in a 60-cubic-centimeter container. In the preliminary detonation work, the results obtained with the detonator cap alone and with water and methyl alcohol were all similar", "timestamp": "2026-07-22T06:31:06.083119+00:00"}
{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 33, "total_pages": 72, "image_filename": "19930085491_p33.jpg", "text": "32 CONFIDENTIAL NACA RM No. ABJ04\n\nthe Reynolds number. Inspection of figures 17(a) and (b) and the corresponding liquid-film results, figures 13(a) and (b), indicates the effect of increasing the Reynolds number on the trailing shock wave of WF-63 at zero lift. This comparison shows that as the Reynolds number was increased the trailing shock wave moved closer to the trailing edge of the wing. This result is similar to that obtained with tests of bodies of revolution, reference 14, where increasing the Reynolds number moved the point of laminar separation to the rear and also moved the trailing shock wave closer to the body base. Although no appreciable rearward movement of the line of laminar separation was apparent in the present tests on the outboard wing sections, there was a decrease in the separated-flow area near the wing root which moved the inboard origin of the compression forward.\n\nFor the lifting wings, the point at which the compression wave joins the trailing edge is associated with the area of separated flow on the upper wing surface since the compression is coincident with the trailing edge on those sections with reattached turbulent boundary layer. This result is also similar to that observed in reference 14 with turbulent flow over bodies of revolution. In this case the compression wave is attached to the body base. Figures 16(b), (d), (f), (h), and (j) show that the point of intersection of the compression line and the trailing edge moves toward the tip as the sweep angle increases, this progression being the same as that shown in figure 14(b) of the extent of the turbulent flow at the trailing edges.\n\nCONCLUSIONS\n\nWind-tunnel tests have been made at a Mach number of 1.53 to determine the longitudinal characteristics of a wing-fuselage combination which linear theory indicates should be capable of efficient flight (maximum lift-drag ratio of approximately 10) up to this Mach number.\n\n1. The following conclusions were obtained from tests with the basic configuration ($63^\\circ$ sweep of leading edge) at a Reynolds number of 0.62 million:\n\n(a) The experimental lift-drag ratio was 6.7 as compared to 10.1 predicted by theory.\n\n(b) The experimental total center-of-pressure travel with lift coefficient was approximately 20 percent of the mean aerodynamic chord as compared to zero travel predicted by theory.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:31:12.475281+00:00"}
{"citation_id": "19930085870", "source_url": "https://ntrs.nasa.gov/api/citations/19930085870/downloads/19930085870.pdf", "page_number": 15, "total_pages": 92, "image_filename": "19930085870_p15.jpg", "text": "14 CONFIDENTIAL NACA RM No. L9D07\n\nsuperimposed on the schlieren photographs to show that the inboard shock arises from the upper surface and the outboard shock from the lower surface. Considering the effect of change in surface Mach number with increase in angle of attack, the Mach number of the flow over the lower wing surface aft of the ridge line would decrease while that of the corresponding upper wing surface would increase. The Mach lines from a fixed point of origin would change their inclination with angle of attack in a direction which is in agreement with the observed changes of the shock inclinations. However, the curvature of the shocks and the shift in the apparent point at center line are not so simply accounted for.\n\nThe profile schlieren photographs of wings 1, 5, and 11 (fig. 18) apparently show no separation of the boundary layer. The shocks emanating from the rear portion of the model may be traced to the trailing edge only. In some instances a very weak shock may be traced to the sting tip on the wing surface; however, this is confined to the profile view and its over-all effect is probably negligible.\n\nPressure Distributions\n\nPressure distributions were made in an effort to show that the location of the steep adverse pressure gradient and the line of transition were practically coincident. Pressure-distribution tests of wedge-leading-edge wing 5 were made at a Mach number of 1.62 at the wing center line, 25.5 percent semispan, and 60.3 percent semispan. The results are presented in figure 21. Similar tests were made of wing 11 for both the elliptical- and wedge-leading-edge configurations at 22.5 percent and 64.1 percent semispan. These results are presented in figures 22 and 23. Except for the elliptical-leading-edge wing, for which a smooth pressure-distribution curve void of sharp peaks has been assumed to exist, no attempt has been made to fair the curves ahead of the ridge line because of insufficient test points in this vicinity.\n\nFor the wedge-leading-edge wings the theoretical pressure distribution at the test stations has been computed for zero angle of attack by the method given in reference 14. (See appendix B.) In all cases the theory gives a fair prediction of the actual results, the greatest discrepancies appearing in the curve for wing 11 at 64.1 percent semispan. Most of the discrepancies are undoubtedly a result of the presence of the shocks on the wing surfaces not accounted for in the theoretical solution.\n\nAt the center-line station of wing 5, test results indicate that no effect is transmitted forward through the boundary layer from the presence of the sting tip. At the 25.5-percent-semispan station the difference in the abruptness of the pressure rise aft of the ridge\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:31:14.643661+00:00"}
{"citation_id": "19930085847", "source_url": "https://ntrs.nasa.gov/api/citations/19930085847/downloads/19930085847.pdf", "page_number": 23, "total_pages": 32, "image_filename": "19930085847_p23.jpg", "text": "NACA RM A9D04\nCONFIDENTIAL\n21\n\nDrag-producing area\nThrust-producing area\n\n<!-- Image (286, 100, 768, 856) -->\n\n(d) M = 0.78 ($C_L = 0.15$).\n\nFigure 7.- Continued.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:31:16.205073+00:00"}
{"citation_id": "19930085911", "source_url": "https://ntrs.nasa.gov/api/citations/19930085911/downloads/19930085911.pdf", "page_number": 1, "total_pages": 52, "image_filename": "19930085911_p1.jpg", "text": "NACA RM E9F22\n\nCopy 467\nRM E9F22\n\nCONFIDENTIAL\n\nNACA\n\nRESEARCH MEMORANDUM\n\nFREE-FLIGHT PERFORMANCE OF 16-INCH-DIAMETER SUPERSONIC\nRAM-JET UNITS\n\nI - FOUR UNITS DESIGNED FOR COMBUSTION-CHAMBER-INLET\nMACH NUMBER OF 0.12 AT FREE-STREAM MACH NUMBER\nOF 1.6 (UNITS A-2, A-3, A-4, AND A-5)\n\nBy William W. Carlton and Wesley E. Messing\n\nLewis Flight Propulsion Laboratory\nCleveland, Ohio\n\nCLASSIFIED DOCUMENT\n\nThis document contains classified information\naffecting the National Defense of the United\nStates within the meaning of the Espionage Act,\nUSC 50:31 and 32. Its transmission or the\nrevelation of its contents in any manner to an\nunauthorized person is prohibited by law.\nInformation so classified may be imparted\nonly to persons in the military and naval\nservices of the United States, appropriate\ncivilian officers and employees of the Federal\nGovernment who have a legitimate interest\ntherein, and to United States citizens of proven\nloyalty and discretion who of necessity must be\ninformed thereof.\n\nCLASSIFICATION CHANGED TO UNCLASSIFIED\nAUTHORITY: RESEARCH ABSTRACT NO. 101\nDATE: MAY 25, 1956\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\nWASHINGTON\nSeptember 22, 1949\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:31:17.427021+00:00"}
{"citation_id": "19930082646", "source_url": "https://ntrs.nasa.gov/api/citations/19930082646/downloads/19930082646.pdf", "page_number": 35, "total_pages": 37, "image_filename": "19930082646_p35.jpg", "text": "34\nNACA TN 1980\n\nMaximum angular acceleration, radians/sec$^2$\nO Maximum positive\n$\\Delta$ Maximum negative\nWarped forebody and extended afterbody ———\nBasic forebody and basic afterbody - - - - - - -\n\nMaximum vertical acceleration, g\nNACA\n\nFigure 17.- Variation of maximum positive and negative angular and maximum vertical accelerations with wave length, for landings in waves 4 feet high:", "timestamp": "2026-07-22T06:31:20.264046+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 51, "total_pages": 78, "image_filename": "19930082483_p51.jpg", "text": "11. Velocity ratio $v$ based on ratio of inlet total pressure to discharge static pressure by equation (52)\n\n$$\n\\begin{aligned}\nv &= \\frac{u_m}{V_j} \\\\\n&= \\frac{528}{1255} \\\\\n&= 0.421\n\\end{aligned}\n$$\n\nEstimation of Power at $180^\\circ$ Admission\n\nAt $180^\\circ$ admission, equation (30) or (53) becomes\n\n$$\n\\begin{aligned}\n(\\text{net power estimated})_{180^\\circ} &= \\frac{1}{2} \\left[ (\\text{power observed})_{360^\\circ} + (\\text{shaft losses})_{360^\\circ} \\right] \\\\\n&- (\\text{shaft losses})_{180^\\circ} \\\\\n&- (\\text{driving-fluid losses})_{180^\\circ}\n\\end{aligned}\n$$", "timestamp": "2026-07-22T06:31:21.079815+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 65, "total_pages": 98, "image_filename": "19930086073_p65.jpg", "text": "```markdown\nNACA RM A59H04\n\nLift coefficient, $C_L$\n\n| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | 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| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | |", "timestamp": "2026-07-22T06:31:22.117319+00:00"}
{"citation_id": "19930085900", "source_url": "https://ntrs.nasa.gov/api/citations/19930085900/downloads/19930085900.pdf", "page_number": 10, "total_pages": 33, "image_filename": "19930085900_p10.jpg", "text": "NACA RM L9D20\n\nCONFIDENTIAL\n\nNormal jets\nSlanted jets\n\nJet outlets\n\n.026 D. tubes\n\nCenterline\n\nSta. 38.50\n45°\n\nSta. 41.00\n\n45°\n\nNACA\n\nNote:\nAll dimensions in inches model size, feet full size\n\nFigure 1.— Exploded view showing the location of jets normal to the center line and jets slanted aft at an angle of $45^\\circ$.\n\nCONFIDENTIAL\n\n9", "timestamp": "2026-07-22T06:31:23.952057+00:00"}
{"citation_id": "19930085899", "source_url": "https://ntrs.nasa.gov/api/citations/19930085899/downloads/19930085899.pdf", "page_number": 11, "total_pages": 29, "image_filename": "19930085899_p11.jpg", "text": "10\n\n<!-- Image (33, 224, 909, 632) -->\n\nFigure 2.- Details of wing fences tested on model with 45° sweptback wing, aspect ratio 4, taper ratio 0.6, and NACA 65A006 airfoil.\n\nNACA RM No. L9K21", "timestamp": "2026-07-22T06:31:26.088205+00:00"}
{"citation_id": "19930085859", "source_url": "https://ntrs.nasa.gov/api/citations/19930085859/downloads/19930085859.pdf", "page_number": 26, "total_pages": 31, "image_filename": "19930085859_p26.jpg", "text": "```markdown\n24\n\nDownwash angle, $\\epsilon$, deg\n\nM = 0.98\n\nM = 1.00\n\nM = 1.03\n\nM = 1.05\n\nM = 1.10\n\n$\\alpha$, deg\n-2 $\\circ$\n-1 $\\square$\n0 $\\diamond$\n1 $\\triangle$\n2 $\\triangledown$\n3 $\\square$\n4 $\\circ$\n6 $\\diamond$\n8 $\\triangledown$\n\nTail height, $h_t$, percent semispan\n\nFigure 10.- Concluded.\n\nNACA RM No. 19B25\n```", "timestamp": "2026-07-22T06:31:30.469407+00:00"}
{"citation_id": "19930082918", "source_url": "https://ntrs.nasa.gov/api/citations/19930082918/downloads/19930082918.pdf", "page_number": 33, "total_pages": 62, "image_filename": "19930082918_p33.jpg", "text": "32\nNACA TN 1940\n\nTABLE 4\nEFFECT OF AGING ON CREEP STRENGTH OF LOW-CARBON\nN-155 ALLOY SOLUTION-TREATED 10 HOURS\nAT 2200° F AND WATER-QUENCHED\n\n| Aging temperature (°F) | Aging time (hr) | Test temperature (°F) | Stress (psi) | Secondary creep rate (in./in./hr) (1) |\n| :--- | :--- | :--- | :--- | :--- |\n| 1200 | 100.0 | 1200 | 30,000 | 0.000007 |\n| | 1000.0 | | 30,000 | .000012 |\n| 1400 | 1.0 | 1200 | 30,000 | .000011 |\n| | | | 60,000 | .009 |\n| | | | | .004 |\n| | 3.0 | | 60,000 | .002 |\n| | 10.0 | | 30,000 | .000007 |\n| | | | 60,000 | .0026 |\n| | | | | .0025 |\n| | 30.0 | | 60,000 | .005 |\n| | 100.0 | | 30,000 | .000063 |\n| | | | | .000070 |\n| | | | 60,000 | .013 |\n| | | | | .004 |\n| | 1000.0 | | 30,000 | .000095 |\n| | | | 60,000 | .019 |\n| | | | | .018 |\n| 1600 | 1.0 | 1200 | 30,000 | .000017 |\n| | | | 60,000 | .0038 |\n| | | | | .004 |\n| | 10.0 | | 30,000 | .000047 |\n| | | | 60,000 | .006 |\n| | | | | .008 |\n| | 100.0 | | 30,000 | .000112 |\n| | | | 60,000 | .013 |\n| | | | | .013 |\n| | 1000.0 | | 30,000 | .00025 |\n| | | | 60,000 | .021 |\n| | | | | .017 |\n| Unaged | | 1200 | 30,000 | .000005 |\n| | | | | .000008 |\n| | | | 60,000 | .015 |\n| | | | | .007 |\n\n1The minimum observed rate at 60,000 psi or the rate between 40 and 50 hr after start of testing at 30,000 psi.\nNACA", "timestamp": "2026-07-22T06:31:32.367220+00:00"}