PDF/ocr_dataset_1066.jsonl

{"citation_id": "19930085899", "source_url": "https://ntrs.nasa.gov/api/citations/19930085899/downloads/19930085899.pdf", "page_number": 6, "total_pages": 29, "image_filename": "19930085899_p6.jpg", "text": "NACA RM No. L9A21\n\nBy measuring tail floating angles without a model installed it was determined that a tail spacing of 2 inches would produce negligible interference effects of reflected shock waves on the tail floating angles. Downwash angles for the wing-alone configuration were therefore obtained simultaneously for the middle, highest, and lowest tail positions in one series of tests, and simultaneously for the two intermediate positions in succeeding runs. (See fig. 3.) The downwash angles presented are increments from the tail floating angles without a model in position. It should be noted that the floating angles measured are in reality a measure of zero pitching moment about the tail pivot axis rather than the angle of zero lift. It has been estimated, however, that for the tail arrangement used a downwash gradient of $2^\\circ$ across the span of the tail will result in an error of less than $0.1^\\circ$ in the measured downwash angle.\n\nTotal-head readings obtained from the tail survey comb have been corrected for bow wave loss. The static-pressure values used in computing the dynamic-pressure ratios were obtained by use of a static probe.\n\nRESULTS AND DISCUSSION\n\nA table of the figures presenting the results is given below:\n\n| Basic wing-alone data | Figure |\n| --- | --- |\n| Basic wing-fuselage data | 7 |\n| Effect of wing fences on wing-fuselage characteristics | 8 |\n| Dynamic-pressure surveys | 9 |\n| Effective downwash angles (wing alone) | 10 |\n| Effective downwash angles (wing fuselage) | 11 |\n| Summary of aerodynamic characteristics | 12 |\n|  | 13 |\n\nThe discussion is based on the summarized values given in figure 13 unless otherwise noted. Note that the slopes summarized in figure 13 have been averaged over a lift range of $\\pm 0.1$ of the nominal lift coefficient.\n\nLift and Drag Characteristics\n\nThe wing-alone lift-curve slope measured near zero lift was about 0.059 at a Mach number of 0.60. This compared with $\\partial C_L / \\partial \\alpha$ of 0.057 obtained from the unpublished low-speed semispan data of a wing with identical geometry which was tested in the Langley two-dimensional tunnel at Reynolds numbers up to $12 \\times 10^6$. The lift curves below M = 1.05 were nonlinear; the slope at higher lift coefficients was somewhat less than that near zero lift. Above M = 1.05 the lift curves were essentially", "timestamp": "2026-07-22T06:28:25.686752+00:00"}
{"citation_id": "19930082566", "source_url": "https://ntrs.nasa.gov/api/citations/19930082566/downloads/19930082566.pdf", "page_number": 41, "total_pages": 44, "image_filename": "19930082566_p41.jpg", "text": "```markdown\nNACA TN NO. 1889\n\n[Graph: S-N diagram comparing fatigue stress results]\n\nY-axis: Stress, psi\nX-axis: N, cycles\n\nData series:\n- Solid specimen; diam., 0.2\"\n- Tubing; inside diam., 2\"; wall thickness, 0.05\"\n\n[Logo: NACA]\n\nFigure 17.- Comparison between S-N diagram based on present longitudinal fatigue-stress results and S-N diagram based on data given in reference 3 for 0.2-inch-diameter specimens.\n\n39\n```", "timestamp": "2026-07-22T06:28:28.327292+00:00"}
{"citation_id": "19930082646", "source_url": "https://ntrs.nasa.gov/api/citations/19930082646/downloads/19930082646.pdf", "page_number": 30, "total_pages": 37, "image_filename": "19930082646_p30.jpg", "text": "NACA TN 1980\n29\n\nV = 23.7 mph; $\\tau = 3.1^\\circ$\nV = 38.8 mph; $\\tau = 8.7^\\circ$\n\nV = 28.0 mph; $\\tau = 3.2^\\circ$\nV = 41.0 mph; $\\tau = 9.4^\\circ$\n\nV = 32.2 mph; $\\tau = 3.3^\\circ$\nV = 43.1 mph; $\\tau = 9.9^\\circ$\n\nV = 36.6 mph; $\\tau = 4.0^\\circ$\n[annotation: NACA L-59840]\nV = 45.3 mph; $\\tau = 10.5^\\circ$\n\n(a) Warped forebody and extended afterbody.\n(b) Basic forebody and basic afterbody.\n\nFigure 14.- Spray on flaps during take-off at design gross load.\n$\\delta_e = -10^\\circ$.", "timestamp": "2026-07-22T06:28:28.867585+00:00"}
{"citation_id": "19930082592", "source_url": "https://ntrs.nasa.gov/api/citations/19930082592/downloads/19930082592.pdf", "page_number": 46, "total_pages": 50, "image_filename": "19930082592_p46.jpg", "text": "NACA TN 1914\n45\n\n[Figure: Micrograph showing a cross-section of a material with three distinct regions labeled: \"Oxide\" at the top, \"Oxide-penetration interface\" in the middle, and \"Unoxidized ceramal\" at the bottom. The image includes a NACA logo with identifiers \"C-22916\" and \"2-7-49\".]\n\nFigure 17. - Oxidation interface of 30-percent-tungsten - titanium carbide ceramal. Oxidation progresses as linear front and grain-boundary oxide penetration is inappreciable. Temperature, $2000^\\circ$ F; time at temperature, 4 hours; unetched; magnification, X750.", "timestamp": "2026-07-22T06:28:29.022205+00:00"}
{"citation_id": "19930082613", "source_url": "https://ntrs.nasa.gov/api/citations/19930082613/downloads/19930082613.pdf", "page_number": 37, "total_pages": 46, "image_filename": "19930082613_p37.jpg", "text": "Page intentionally left blank\n\nPage intentionally left blank", "timestamp": "2026-07-22T06:28:34.016569+00:00"}
{"citation_id": "19930083221", "source_url": "https://ntrs.nasa.gov/api/citations/19930083221/downloads/19930083221.pdf", "page_number": 28, "total_pages": 47, "image_filename": "19930083221_p28.jpg", "text": "26\nNACA TN No. 1824\n\n$$\n\\left.\n\\begin{aligned}\n\\xi &= -y_1 \\tan \\mu + x_1 \\\\\n\\eta &= y_1 \\sec \\mu\n\\end{aligned}\n\\right\\}\n\\tag{39}\n$$\n\nwhere\n\n$$\n\\tan \\mu = \\beta \\cos \\theta\n$$\n\nequation (29) becomes\n\n$$\n\\varphi(x,y,z) = -\\frac{1}{2\\pi} \\iint_{\\tau} \\frac{\\cos \\mu \\Delta w_0(\\xi,\\eta) d\\xi d\\eta}{\\sqrt{(x-x_1)^2 - \\beta^2(y-y_1)^2 - \\beta^2 z^2}}\n$$\n\nSince, however, it has been shown that $\\varphi$ evaluated infinitely far away from the wing does not change if a source is moved along the line $\\xi$=constant, it follows that the source strengths can be integrated along these lines. The second integration is then along $\\eta$=0 where, from equation (39), $\\xi$=$x_1$ and the value of the potential at an infinite distance is\n\n$$\n\\varphi(x,y,z) = -\\frac{1}{2\\pi} \\int \\frac{\\cos \\mu \\ dx_1}{\\sqrt{(x-x_1)^2 - \\beta^2 y^2 - \\beta^2 z^2}} \\int \\Delta w_0(x_1,\\eta) d\\eta\n$$\n\nSetting\n\n$$\nf(x_1,\\mu) = -\\cos \\mu \\int \\Delta w_0(x_1,\\eta) d\\eta\n$$\n\nit follows that\n\n$$\n\\varphi(x,r,\\theta) = \\frac{1}{2\\pi} \\int_0^{x-\\beta r} \\frac{f(x_1,\\mu) dx_1}{\\sqrt{(x-x_1)^2 - \\beta^2 r^2}}\n\\tag{40}\n$$\n\nand this is the same as the potential for a body of revolution with source strength per unit length given by $f(x_1)$. The induced velocities corresponding to the potential in equation (40) are found to be, after first integrating by parts and using the notation $\\partial/\\partial x_1 \\ f(x_1,\\mu) = f'(x_1,\\mu)$ together with the relation $f(0,\\mu) = 0$,", "timestamp": "2026-07-22T06:28:34.786296+00:00"}
{"citation_id": "19930082546", "source_url": "https://ntrs.nasa.gov/api/citations/19930082546/downloads/19930082546.pdf", "page_number": 44, "total_pages": 65, "image_filename": "19930082546_p44.jpg", "text": "NACA TN No. 1870\n\n43\n\n[Figure: Graph showing Pressure, p, dynes/cm² vs. x/D for various a/D values (.042, .083, .167, .333), with handwritten calculations in top right corner including:  \n$ \\frac{dynes}{cm^2} = .00102 \\cdot \\rho_{air} $  \n$ \\rho_{air} = .0822 \\ lb/ft^3 $  \n$ V_{max} = 3937 \\ in/sec $  \n$ (1 \\frac{lb}{in^2}) (.00102 \\frac{dynes}{cm^2}) / (\\frac{lb}{ft^3}) $  \n$ (1 \\ cm^2) (39.37 \\ in)^2 $  \n$ = .000145 \\ \\#/in^2 $  \nand NACA logo at bottom right of graph]\n\nFigure 6.— Effect of tip clearance on the free-space pressures for NACA 4-(5)(08)-03 propeller. B = 4; β₀.₇₅ = 10°; Mₜ = 0.60.\n\n[Figure: Graph showing Cₚ vs. x/D for various β₀.₇₅ (deg) values (8, 15, 20) with corresponding Cₚ values (.0195, .0454, .0815), and NACA logo at bottom right of graph]\n\nFigure 7.— Effect of blade loading on the free-space pressure distribution for NACA 4-(5)(08)-03 propeller. B = 2; Mₜ = 0.60; $\\frac{a}{D}$ = 0.083.", "timestamp": "2026-07-22T06:28:35.186294+00:00"}
{"citation_id": "19930085890", "source_url": "https://ntrs.nasa.gov/api/citations/19930085890/downloads/19930085890.pdf", "page_number": 9, "total_pages": 26, "image_filename": "19930085890_p9.jpg", "text": "```markdown\n8\nNACA RM No. E9C11\n\nmethod of injection are also shown. The experimental performance data cover a range of ratios of fuel weight to total propellant weight of 0.22 to 0.58 (stoichiometric, 0.224). The maximum specific impulse, as indicated by the faired curve, is 249 pound-seconds per pound and occurred at a ratio of fuel weight to total propellant weight of approximately 0.37.\n\nA curve obtained by correcting the experimental data for heat losses and for small combustion-chamber-pressure differences from 300 pounds per square inch is also shown in figure 5. The corrections are predominantly heat corrections. The curve with these corrections increased to a maximum of 274 pound-seconds per pound at a ratio of fuel weight to total propellant weight of about 0.36.\n\nFor comparison, the theoretical performance for the fuel and the expansion nozzle used is shown in figure 5; chemical equilibrium during expansion is assumed. Figure 5 also shows the theoretical performance curve of 100-percent liquid diborane and liquid oxygen taken from reference 1.\n\nThe shapes of the experimental and theoretical curves are similar with the peak occurring in nearly the same position near a ratio of fuel weight to total propellant weight of 0.36. The maximum uncorrected experimental specific impulse is approximately 83 percent of the theoretical value for the fuel and nozzle used. The peak value of the corrected experimental curve is 92 percent of the peak (299 lb-sec/lb) for the theoretical curve for equilibrium expansion and the fuel and the nozzle used.\n\nChanging the $I^*$ of the engine from 325 to 159 inches produced a small decrease in performance (fig. 5). Changing the method of injection from the eight-hole, solid-jet system to the four-hole system in an engine having an $I^*$ of 325 inches appeared to have no definite effect on the average performance of the propellant.\n\nThe injection systems and the combustion volumes used were not sufficient in number and variation to determine optimum values of these design variables.\n\nCurves of theoretical and experimental volume specific impulse plotted against ratio of fuel weight to total propellant weight are presented in figure 6. The values of theoretical volume specific impulse were obtained by multiplying values of theoretical specific impulse by the density of the propellant combination at corresponding propellant mixture ratios.\n```", "timestamp": "2026-07-22T06:28:35.525128+00:00"}
{"citation_id": "19930085900", "source_url": "https://ntrs.nasa.gov/api/citations/19930085900/downloads/19930085900.pdf", "page_number": 5, "total_pages": 33, "image_filename": "19930085900_p5.jpg", "text": "4 CONFIDENTIAL NACA RM L9D20\n\nno indication of any reduction in the rate of increase. The trim quickly rose to $18.6^\\circ$ at 17 feet per second and then remained against the trim stops (set at $20^\\circ$) from 25 feet per second up. The effective hydrodynamic lift was very low. The low lift and the high trim are an indication of the strong suction forces acting on the unmodified fuselage.\n\nEffect of Jet Spacing\n\nThe effect of jet spacing on resistance, trim, and lift are shown in figure 6. The resistance and trim decreased as the jet spacing was decreased but were always considerably less than for the basic model. The lift was practically the same for all spacings except at 15 and 22 feet per second.\n\nThe lower trims obtained for the more closely spaced jets indicate that they reduced the suction forces more than the jets spaced further apart. The presence of some suction force for all jet spacings was shown by the model maintaining a trim of at least $6^\\circ$ at the higher speeds even though the center of gravity was forward of the wetted area.\n\nThe photographs in figure 7 show the spray characteristics of the 2-inch spacing and the $\\frac{1}{4}$-inch spacing at 35 feet per second. At this speed the trim for the 2-inch spacing was about $2^\\circ$ higher than for the $\\frac{1}{4}$-inch spacing. The spray height was about the same for both, but the density of the spray was less for the $\\frac{1}{4}$-inch spacing. The direction of the spray at the side of the model was more nearly vertical for the 2-inch spacing than for the $\\frac{1}{4}$-inch spacing. The jets caused the spray to separate from the fuselage and the $\\frac{1}{4}$-inch-spaced jets were apparently more effective in this respect than the 2-inch-spaced jets.\n\nEffect of Length of Jet Rows\n\nResults of tests to determine the effect of varying the length of the rows of jets simulating chines are given in figure 8. The resistance was decreased with an increase in length of the jet rows. Because the curves for the 24- and 32-inch lengths are practically the same, no further reduction in resistance could be expected if the rows of jets were extended to the nose of the fuselage. The omission of the forward portion of the simulated chines, however, permitted the water to run up over that part of the fuselage at low speeds.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:28:36.359992+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 63, "total_pages": 98, "image_filename": "19930086073_p63.jpg", "text": "```markdown\nNACA RM A50H04\n\n1.4\n1.2\n1.0\n.8\n.6\n.4\n.2\n0\n-.2\n0 .1 .2 .3 .4 .5 .6 .7 .8 .9\nLift coefficient, $C_L$\nDrag coefficient, $C_D$\n\n$\\circ$ $\\delta_{a_L}=+10.8, \\delta_{a_R}=-10.8$\nAileron deflection, $\\delta_a$, deg\n(b) $C_L$ vs $C_D$.\n\nFigure 13.— Continued.\n\n61\n```", "timestamp": "2026-07-22T06:28:41.507956+00:00"}
{"citation_id": "19930085548", "source_url": "https://ntrs.nasa.gov/api/citations/19930085548/downloads/19930085548.pdf", "page_number": 27, "total_pages": 46, "image_filename": "19930085548_p27.jpg", "text": "26\nNACA RM No. E8L30\n\nAlternate locations for fuel-injector\nmaximum cylinder-pressure indicator\nvalve and exhaust-gas sampling valve\n\nFuel injector\n\nCompression ratio\n7.0\n5.25\n4.5\n4.0\n\nInlet-air\nside of cylinder\n\nFuel-nozzle-\norifice diameters\n(in.)\nA 0.019\nB .014\nC .008\n\nSection through plane of sprays\n\nFigure 5. - Outlines of combustion chamber (for\ncompression ratios of 4, 4.5, 5.25, and 7),\nlocation of fuel-injection valve, and pattern of\nfuel sprays.", "timestamp": "2026-07-22T06:28:42.095515+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 45, "total_pages": 78, "image_filename": "19930082483_p45.jpg", "text": "NACA TN No. 1807\n43\n\n$\\eta'$ turbine over-all efficiency based on ratio of inlet stagnation conditions to discharge stagnation conditions\n\n$\\theta$ temperature correction ratio, $T/T_0$\n\n$\\lambda$ empirical constant, depending upon inactive blade shrouding\n\n$\\mu$ absolute viscosity, (lb)(sec)/sq ft\n\n$\\nu$ velocity ratio based on ratio of inlet total pressure to discharge static pressure\n\n$\\nu'$ velocity ratio based on ratio of inlet total pressure to discharge total pressure\n\n$\\rho$ mass density, slugs/cu ft\n\nSubscripts:\n\nd fluid surrounding turbine-rotor disk\n\ne turbine discharge (measuring station)\n\nF degree of admission corresponding to fraction of active nozzle arc\n\nG degree of admission corresponding to fraction of active nozzle arc other than F\n\nh hub or inner radius position (blade root)\n\ni inlet (measuring station)\n\nJ jet\n\nm pitch line\n\ns isentropic\n\nT tip or outer-radius position\n\nu tangential\n\nx axial\n\n0 NACA sea-level air", "timestamp": "2026-07-22T06:28:43.136376+00:00"}
{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 29, "total_pages": 72, "image_filename": "19930085491_p29.jpg", "text": "28 CONFIDENTIAL NACA RM No. A8J04\n\nunless otherwise stated, the effect of sweep also includes the effects of the wing thickness and aspect ratio changes.\n\nEffect of sweep on lift-curve slope.— Data presented in figures 7(a) through (e) show that the break in the lift curve which has been previously discussed with WF-63 is evident at all sweep angles. For this reason two values of the slope have been listed in table II which indicate the difference in slope between zero lift and the optimum lift coefficient. Also shown in table II and on the experimental plots are the theoretical values for all wings except that of WF-70. Figure 10(b) is a cross plot of $(dC_L/d\\alpha)_{opt}$ and $(dC_L/d\\alpha)_{theo}$ against the factor $m$ which shows that both values decrease with increasing angles of sweep when the Mach number remains constant. The liquid-film test results give some insight as to the boundary-layer-flow changes associated with the differences between experiment and theory. The changes in boundary-layer flow on the upper surfaces of the wings near the optimum lift coefficients are shown in figure 14(b). The differences in flow pattern due to differences in lift coefficient from the optimum are relatively small and can be neglected. (See figs. 13(c) and (d).) These liquid-film patterns indicate that at the lower angles of sweep where the greatest difference between experiment and theory exists, the area of separated flow at the trailing edge is also the greatest.\n\nEffect of sweep on minimum drag coefficient.— The theoretical and experimental variations of minimum drag coefficient with sweep are shown in figure 10(a) where it will be noted that there is a marked reduction in minimum drag coefficient with increasing sweep. The more rapid rise of $C_{Dmin}$ obtained theoretically as $m$ approaches a value of one (decreasing sweep) is a result of the use of a double-wedge airfoil section in the theoretical determination of the wing pressure drag; that is, at the lower angles of sweep where the ridge line of the theoretical wing is nearly sonic, the theoretical pressure drag is somewhat higher than would be obtained with the test wing section which has no abrupt change in slope at the maximum-thickness position. The variation of the theoretical pressure-drag increment for the wings as tabulated under the section Theoretical Considerations does, however, indicate that exclusive of the effect of thickness distribution just discussed the variation in wing-pressure drag almost completely accounts for the measured reduction in total minimum drag coefficient with increased sweep. The results of reference 11 show that the primary factors in reducing the wing-thickness drag are\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:28:45.769332+00:00"}
{"citation_id": "19930085862", "source_url": "https://ntrs.nasa.gov/api/citations/19930085862/downloads/19930085862.pdf", "page_number": 19, "total_pages": 60, "image_filename": "19930085862_p19.jpg", "text": "NACA RM No. L9A07\n17\n\n[Figure: Front view of bottom of wing.]\n\n(a) Front view of bottom of wing.\n\n[Figure: Rear view of top of wing.]\n\n(b) Rear view of top of wing.\n\nFigure 2.— Wing mounted in 19-foot pressure tunnel.", "timestamp": "2026-07-22T06:28:46.172108+00:00"}
{"citation_id": "19930085870", "source_url": "https://ntrs.nasa.gov/api/citations/19930085870/downloads/19930085870.pdf", "page_number": 12, "total_pages": 92, "image_filename": "19930085870_p12.jpg", "text": "NACA RM No. L9D07 CONFIDENTIAL 11\n\nfor sharp-leading-edge wings with and without the effect of leading-edge suction. The theoretical $(L/D)_{\\max}$ for uncambered wings is\n\n$$(L/D)_{\\max} = \\frac{1}{2} \\sqrt{\\frac{1}{C_{D_{\\min}} \\left( \\Delta C_D / C_L^2 \\right)}} \\tag{4}$$\n\nIn the theoretical calculations it was assumed that turbulent flow existed over the greater portion of the wing aft of the ridge line. Accordingly, a friction-drag coefficient based on turbulent flow and a mean value of the test Reynolds numbers was assumed to be 0.0093. This value was added to the previously calculated wave-drag values in determining the theoretical $(L/D)_{\\max}$. No points are indicated on the test curves as it was often necessary to extrapolate the $L/D$ curves of the individual wings to obtain the value of $(L/D)_{\\max}$, a result of the low angle-of-attack range of the tests. The extrapolated values are given in table 2. As expected, the highest values of $(L/D)_{\\max}$ were obtained at low values of $\\tan \\epsilon / \\tan m$, the region of low values of minimum drag. In the vicinity of $\\frac{\\tan \\epsilon}{\\tan m} = 1$, the test values are greater than the theoretical because of the abnormally large drag values predicted by theory. At the higher values of $\\tan \\epsilon / \\tan m$, the test results are less than theory primarily because the experimental lift-curve slopes are less than theory and the experimental drag is greater than theory. The higher $(L/D)_{\\max}$ of the elliptical-leading-edge wings at low values of $\\tan \\epsilon / \\tan m$ may be traced to the smaller minimum drag of these wings rather than any large realization of leading-edge-suction force. In general, the linear theory gives a fair approximation of maximum $L/D$ for wings of this thickness ratio. It is interesting to note that values of $(L/D)_{\\max}$ as high as 8.1 were obtained for the thin-plate wings (see table 3) as compared with a value of 5.8 for the thick-wing series.\n\nCenter of Pressure and Pitching Moment\n\nPitching-moment-curve slopes $\\frac{dC_M}{d\\alpha}$ at zero lift are presented in figure 14 as a function of $\\tan \\epsilon / \\tan m$ and show that the center of area is a good approximation of the center of pressure. Figure 15 gives the actual center-of-pressure location. For both the elliptical- and wedge-leading-edge series, the center of pressure shifts forward with increase in $\\tan \\epsilon / \\tan m$, the over-all travel being approximately 10 percent. The location of the center of pressure appears relatively independent of Mach number for the wings of a given leading-edge shape. However, the center of pressure of the elliptical-leading-edge wings\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:28:47.184048+00:00"}
{"citation_id": "19930082918", "source_url": "https://ntrs.nasa.gov/api/citations/19930082918/downloads/19930082918.pdf", "page_number": 29, "total_pages": 62, "image_filename": "19930082918_p29.jpg", "text": "28\nNACA TN 1940\n\nREFERENCES\n\n1. Freeman, J. W., Reynolds, E. E., Frey, D. N., and White, A. E.: A Study of Effects of Heat Treatment and Hot-Cold-Work on Properties of Low-Carbon N-155 Alloy. NACA TN 1867, 1949.\n\n2. Thomassen, Lars, Williams, Robley C., and Wyckoff, Ralph W. G.: Surface Replicas for Electron Microscopy. Rev. Sci. Instr., vol. 16, no. 6, June 1945, pp. 155-156.\n\n3. Haworth, F. E.: Energy of Lattice Distortion in Cold Worked Permalloy. Phys. Rev., vol. 52, Sept. 15, 1937, pp. 613-620.\n\n4. Warren, B. E., and Biscoe, J.: The Structure of Silica Glass by X-Ray Diffraction Studies. Jour. Am. Ceramic Soc., vol. 21, 1938, pp. 49-54.\n\n5. Nabarro, F. R. N.: The Strains Produced by Precipitation in Alloys. Proc. Roy. Soc. (London), ser. A, vol. 175, no. 963, July 18, 1940, pp. 519-538.\n\n6. Dehlinger, U.: Theory of Distorted Lattices. Z. Kristallogr. Bd. 65, 1927, p. 615.\n\n7. Mehl, R. F., and Jetter, L. K.: Symposium on the Age-Hardening of Metals. Am. Soc. Metals (Cleveland, Ohio), 1940.\n\n8. Seitz, Frederick: Modern Theory of Solids. McGraw-Hill Book Co.; Inc., 1943, ch. 10.\n\n9. Hume-Rothery, William: The Structure of Metals and Alloys. Third ed., Inst. of Metals (London), 1948.\n\n10. Hume-Rothery, William: Atomic Theory for Students of Metallurgy. Inst. of Metals (London), 1948, p. 261.\n\n11. Bridgman, P. W.: The Stress Distribution at the Neck of a Tension Specimen. Trans. Am. Soc. Metals, vol. 32, 1944, pp. 553-574.\n\n12. Sachs, George: Effect of Strain on Fracture. Symposium on Fracturing of Metals. Am. Soc. Metals, 1948.\n\n13. Lyman, Taylor (ed.): Metals Handbook. Am. Soc. Metals, 1948, p. 21.\n\n14. Barrett, Charles S.: Structure of Metals. First ed., McGraw-Hill Book Co., Inc., 1943, p. 553.", "timestamp": "2026-07-22T06:28:48.442863+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 57, "total_pages": 114, "image_filename": "19930086061_p57.jpg", "text": "NACA RM L9J07\n53\n\nLeft semispan\nUpper\nLower\nRight semispan\n-3\n-2\n-1 P\n0\n1\n-3\n-2\n-1\n0\n1\n20°\n(c) $\\psi = 20^\\circ$\n\nLeft semispan\nUpper\nLower\nRight semispan\n-3\n-2\n-1 P\n0\n1\n-3\n-2\n-1\n0\n1\n35°\n(d) $\\psi = 35^\\circ$\nNACA\n\nFigure 15.- Concluded.", "timestamp": "2026-07-22T06:28:51.098072+00:00"}
{"citation_id": "19930082566", "source_url": "https://ntrs.nasa.gov/api/citations/19930082566/downloads/19930082566.pdf", "page_number": 42, "total_pages": 44, "image_filename": "19930082566_p42.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:28:51.912833+00:00"}
{"citation_id": "19930085847", "source_url": "https://ntrs.nasa.gov/api/citations/19930085847/downloads/19930085847.pdf", "page_number": 18, "total_pages": 32, "image_filename": "19930085847_p18.jpg", "text": "16\nCONFIDENTIAL\nNACA RM A9D04\n\n<!-- Image (96, 220, 874, 666) -->\n\n(a) Chordwise force coefficient.\n\nFigure 6.- Variation of chordwise force coefficient and profile drag coefficient with Mach number with and without suction on test panel.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:28:54.618218+00:00"}
{"citation_id": "19930085899", "source_url": "https://ntrs.nasa.gov/api/citations/19930085899/downloads/19930085899.pdf", "page_number": 7, "total_pages": 29, "image_filename": "19930085899_p7.jpg", "text": "6\nNACA RM No. L9A21\n\nlinear over the test lift range. (See figs. 7 and 13.) The basic lift-curve slope was increased by an average of 20 percent by the addition of the fuselage.\n\nThe drag rise at zero lift occurred at a Mach number of about 0.93 for both wing and wing-fuselage conditions. It should be remembered that the absolute drag coefficients are probably high because of the presence of end-plate tare and also because of the low Reynolds numbers at which the tests were run. At a Mach number of 0.6 the minimum drag coefficient for the wing alone was 0.010. This value compares with a minimum drag value of approximately 0.0050 obtained from the previously mentioned low-speed investigation made at a Reynolds number of $12 \\times 10^6$. The values of $(L/D)_{max}$ shown in figure 13 are somewhat low because of high absolute drag and are presented primarily for comparison with the other wings to be investigated in this series.\n\nThe lateral center of pressure for the wing alone ($C_L = 0.4$) increased from about 48 percent semispan at $M = 0.60$ to about 50 percent semispan at $M = 0.98$. Between $M = 0.98$ and 1.05 there is a fairly abrupt outboard movement of $y_{cp}$ to about 55 percent of the semispan. The addition of the fuselage moved the lateral center of pressure inboard an average of about 3 percent semispan throughout the Mach number range.\n\nPitching-Moment Characteristics\n\nNear zero lift coefficient, the wing aerodynamic center was about 32 percent mean aerodynamic chord and was almost constant throughout the Mach number range. This value compares with a low-speed aerodynamic-center location of 25 percent mean aerodynamic chord near zero lift obtained from the unpublished Langley two-dimensional-tunnel data. The addition of the fuselage moved the aerodynamic center forward about 7 percent mean aerodynamic chord below $M = 1.00$ and forward to a lesser extent at the higher Mach numbers. At $C_L = 0.4$ the wing-alone aerodynamic center varied from about 22 percent mean aerodynamic chord at low Mach numbers to 45 percent mean aerodynamic chord at $M = 1.18$. The unstable shift in aerodynamic center at low Mach numbers in the higher $C_L$ range is even more pronounced for the wing-fuselage condition.\n\nEffect of Wing Fences\n\nIn an attempt to alleviate the unstable aerodynamic-center shift, which was probably caused by premature flow separation at the tip in the higher lift range at low Mach numbers, it was decided to investigate two upper-surface wing fences on the wing-fuselage combination.", "timestamp": "2026-07-22T06:29:00.832800+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 64, "total_pages": 98, "image_filename": "19930086073_p64.jpg", "text": "62\nNACA RM A9H04\n\n<!-- Image (115, 91, 853, 819) -->\n\n(c) $C_L$ vs $C_m$.\n\nFigure 13.—Continued.", "timestamp": "2026-07-22T06:29:01.018793+00:00"}
{"citation_id": "19930085548", "source_url": "https://ntrs.nasa.gov/api/citations/19930085548/downloads/19930085548.pdf", "page_number": 28, "total_pages": 46, "image_filename": "19930085548_p28.jpg", "text": "```markdown\nNACA RM No. E8L30\n\nPressure tap\nThermocouples\nInsulation\nInlet\nOutlet\nNACA\n\nFigure 6. - Longitudinal section through exhaust surge tank instrumented for obtaining exhaust-gas pressures and temperatures.\n\n27\n```", "timestamp": "2026-07-22T06:29:02.581705+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 46, "total_pages": 78, "image_filename": "19930082483_p46.jpg", "text": "44\nNACA TN No. 1807\n\n1 rotor inlet\n2 rotor discharge\n360° full admission\n120° 120° admission\n180° 180° admission\n\nSuperscripts:\n- (bar) average value\n' (prime) stagnation state", "timestamp": "2026-07-22T06:29:05.780132+00:00"}
{"citation_id": "19930082613", "source_url": "https://ntrs.nasa.gov/api/citations/19930082613/downloads/19930082613.pdf", "page_number": 38, "total_pages": 46, "image_filename": "19930082613_p38.jpg", "text": "NACA TN 1938\n37\n\n<!-- Image (100, 191, 899, 801) -->\n\n(a) View of crack, which\nextends from punched\nhole.\n\n(b) Cutaway view of crack.\n\n(c) Top view of crack and\ncutaway.\n\n(d) Top view of two-dimensional\ncrack after surface has been\npolished to section A-A.\nCompare with photomicrographs\n(figs. 5, 8, and 9).\n\n(e) Section through two-\ndimensional crack.\n\nFigure 12. - Pictorial explanation of tubular formation.", "timestamp": "2026-07-22T06:29:06.687807+00:00"}
{"citation_id": "19930085862", "source_url": "https://ntrs.nasa.gov/api/citations/19930085862/downloads/19930085862.pdf", "page_number": 20, "total_pages": 60, "image_filename": "19930085862_p20.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:29:09.302820+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 28, "total_pages": 149, "image_filename": "19930083192_p28.jpg", "text": "24\nNACA TN 1976\n\ninformation. The old NACA gust tunnel described in reference 19 was capable of handling 3-foot-span models of airplanes at speeds up to about 50 miles per hour but in 1945 was replaced by a tunnel that is able to handle 6-foot-span models at speeds up to 100 miles per hour.\n\nThe accuracy of measurement at present is about 0.05g for acceleration, 0.3 foot per second for speed, and 0.01 inch for deflection of the wing or flap. Angular motions are determined within $0.1^{\\circ}$ of pitch-angle increment and about $0.2^{\\circ}$ for the position of any flap. All measurements except that of speed are incremental values from steady conditions and, as such, depend upon the steadiness of flight prior to entry into the gust. At peak acceleration for sharp-edge gusts, the acceleration increment is considered accurate to within 0.1g and the pitch-angle increment to about $0.1^{\\circ}$, but the accuracy is questionable for the longest gust-gradient distance of about 16 chords.\n\nIn fundamental studies such as the determination of unsteady-lift functions, the solution obtained from gust-tunnel tests is an indirect one since the acceleration increment due to a gust is dependent on the difference between two terms. Thus, it is not possible to determine whether the correct values of the unsteady-lift functions are used, but only whether the use of unsteady-lift functions for a given shape is adequate. If sufficient data are collected utilizing different models of many sizes and shapes in different gusts, the various unsteady-lift functions can be evaluated to a limited degree.\n\nFlight Investigations\n\nTwo approaches have been utilized for flight investigations of airplane reactions, one consisting simply of detailed analysis of flight records to obtain the relations desired, for example, between the pitch of an airplane and the acceleration imposed on the airplane, and the other consisting of flying an airplane in rough air to obtain statistical data for different conditions and then statistically comparing the reactions of the airplane. The first approach was used in a rather crude manner in reference 3, in which a determination was made of the effect of airplane wing loading and speed on the gust load factor. The statistical approach may permit checks to be made of both theoretical and gust-tunnel results, but the procedures and techniques have not been developed to the point where precise results can be obtained.\n\nTRANSIENT AERODYNAMICS\n\nThe material presented in this section covers the development of the lift force and downwash but does not cover the variation in moment", "timestamp": "2026-07-22T06:29:10.492990+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 58, "total_pages": 114, "image_filename": "19930086061_p58.jpg", "text": "54\nNACA RM L9J07\n\nLeft semispan\nRight semispan\nUpper\nLower\n\n(a) $\\psi = 0^\\circ$\n\nLeft semispan\nRight semispan\nUpper\nLower\n\n(b) $\\psi = 10^\\circ$\n\nFigure 16.- Pressure distribution about wing 1 at various angles of yaw;\n$\\alpha = 44.1^\\circ$.", "timestamp": "2026-07-22T06:29:12.955879+00:00"}
{"citation_id": "19930085900", "source_url": "https://ntrs.nasa.gov/api/citations/19930085900/downloads/19930085900.pdf", "page_number": 6, "total_pages": 33, "image_filename": "19930085900_p6.jpg", "text": "NACA RM L9D20 CONFIDENTIAL 5\n\nThe trim and lift for the 16-, 24-, and 32-inch lengths were practically the same. The resistance for the 16-inch length, however, was greater than for the other two.\n\nThe difference in spray characteristics between the 8-inch length and 32-inch length is shown in figure 9. The spray for the 8-inch length was heavier than that for the 32-inch length. The arrow in the photograph of the 8-inch length points to station 34 at which a spume of spray comes off the forward jet. The water forward of this station can be seen running up the side of the model with some of the water going over the top; the spray aft of station 34 slants back in a more nearly horizontal direction. For the 32-inch length, the spray broke from the model along the entire wetted length and the top of the model was free of water.\n\nEffect of Slanting Jets Aft\n\nThe effect on resistance, trim, and lift of substituting chine jets slanted back at an angle of $45^\\circ$ to the center line for jets normal to the center line is shown in figure 10. The substitution of slanted jets for jets perpendicular to the center line had little effect on resistance or lift but increased the trim over most of the speed range. The horizontal thrust component of the slanted jets, measured at rest with a load on the water of 7.6 pounds, was about 0.1 pound.\n\nThe photographs in figure 11 compare the spray pattern of the slanted jets with that of the normal jets in the chine configuration. The upper surface and the stern of the model with the slanted jets was completely free of water as shown in figure 11(c). The spray characteristics of the model with the jets perpendicular to the center line were similar.\n\nEffect of Air Flow\n\nThe curves in figure 12 show the effect of air flow on resistance at three representative speeds for each of three different configurations. The jet spacing for all three configurations was $1/4$ inch and the number of jets in each was approximately the same.\n\nThe variation in resistance with air flow at each speed was approximately the same for all three configurations. The very high resistance at extremely low air flows shows that merely venting the bottom of the fuselage through the jets would have had little effect on the resistance. 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\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:29:13.896720+00:00"}
{"citation_id": "19930085847", "source_url": "https://ntrs.nasa.gov/api/citations/19930085847/downloads/19930085847.pdf", "page_number": 19, "total_pages": 32, "image_filename": "19930085847_p19.jpg", "text": "NACA RM A9D04 CONFIDENTIAL 17\n\nNote: Cross flow present on all flagged points.\n\nProfile drag coefficient, $C_d$\n\n| | |\n| :--- | :--- |\n| $\\circ$ Slot closed | |\n| $\\triangle$ Slot open ($C_q \\approx 0.00175$) | |\n\nMach number\n\n(b) Profile drag coefficient.\n\nFigure 6.- Concluded.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:29:14.808549+00:00"}
{"citation_id": "19930082546", "source_url": "https://ntrs.nasa.gov/api/citations/19930082546/downloads/19930082546.pdf", "page_number": 45, "total_pages": 65, "image_filename": "19930082546_p45.jpg", "text": "44\nNACA TN No. 1870\n\n.040\n2-blade NACA 4-(5)(08)-03 propeller\n2-blade NACA 4-(5)(08)-03 propeller\n2-blade NACA 4-(5)(06.5)-06 propeller\n2-blade Clark Y propeller\n\n.036\n\n.032\n$M_t = 1.00$\n\n.028\n\n.024\n\n.020\n$C_p$\n\n.016\n$M_t = 0.60$\n\n.012\n\n.008\n\n.004\n\n0\n.02 .04 .06 .08 .10\n$C_p$\n\n[Figure: NACA logo]\n\nFigure 8.- Effect of power coefficient and tip Mach number on the oscillating-pressure coefficients of two- and four-blade propellers in the plane of rotation. $\\frac{d}{D} = 0.042$.", "timestamp": "2026-07-22T06:29:15.987181+00:00"}
{"citation_id": "19930082566", "source_url": "https://ntrs.nasa.gov/api/citations/19930082566/downloads/19930082566.pdf", "page_number": 43, "total_pages": 44, "image_filename": "19930082566_p43.jpg", "text": "NACA TN No. 1889\n41\n\n[Figure: Four cylindrical metal specimens, each with a different fracture pattern, arranged horizontally. Below each specimen is a label indicating a stress ratio.]\n\n$\\sigma_2/\\sigma_1 = 0$\n$\\sigma_2/\\sigma_1 = 2$\n$\\sigma_2/\\sigma_1 = 1$\n$\\sigma_2/\\sigma_1 = 0.5$\n\n[annotation: NACA logo]\n\nFigure 18.- Typical fractured specimens.", "timestamp": "2026-07-22T06:29:17.679173+00:00"}
{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 30, "total_pages": 72, "image_filename": "19930085491_p30.jpg", "text": "NACA RM No. A8J04 CONFIDENTIAL 29\n\nsmall thickness-chord ratios, high angles of sweep, and large aspect ratios. For the present wings, where the thickness-chord ratio and aspect ratio both decrease with increased sweep (fig. 10(a)), it is probable that both the theoretical and experimental smaller rates of decrease in $C_{D_{\\min}}$ at the highest sweep angles are largely due to the adverse reduction in aspect ratio.\n\nIn addition to changes in wing-fuselage interference, other factors which may influence the minimum drag-coefficient variation with sweep are changes in skin friction and separation pressure drag. Figure 14(a) which presents liquid-film results at zero lift for the test configurations shows that there are small changes in the area of separated flow, particularly at the three highest angles of sweep. These results give further indication that the large variation in minimum drag coefficient with sweep is due primarily to changes in wing-thickness pressure drag rather than to changes in separation or friction drag.\n\nEffect of sweep on drag due to lift.— As was discussed in the preceding sections, the drag curves obtained with all configurations are composed essentially of two parabolic segments which join slightly below the optimum lift coefficient. Thus the values of the drag-rise factors at the optimum lift coefficients are slightly greater than those in the lower range of lift coefficients. Both experimental values are higher than indicated by theory for the reasons previously discussed with the results of the tests of WF-63.\n\nThe experimental variation of drag-rise factor with sweep can be studied by considering the factors which determine $\\Delta C_D$ at a given value of $\\Delta C_L$ by use of equation (6). For a constant lift coefficient as the sweep angle is increased, the increase in $\\Delta C_D$ can be attributed either to an increase in angle of attack or to a change in $k_a$ or to changes in both. As was considered in the preceding section the lift-curve slope decreases with increasing sweep. Consequently the angle of attack for a given lift coefficient increases and contributes to an increase in $\\Delta C_D$. The variations with sweep of $k_a$ near zero lift and at the optimum lift coefficient are shown in table II and the latter values are plotted in figure 10(c). Since the values from $57.0^\\circ$ to $67.0^\\circ$ sweep are nearly constant, this variation of $k_a$ has little influence on the noted increase in drag-rise factor. However, above $67.0^\\circ$ sweep there is an abrupt increase in the value of $k_a$ that, coupled with the decreased lift-curve slope, results in a rapid increase in the rate of drag-rise.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:29:19.914812+00:00"}
{"citation_id": "19930082613", "source_url": "https://ntrs.nasa.gov/api/citations/19930082613/downloads/19930082613.pdf", "page_number": 39, "total_pages": 46, "image_filename": "19930082613_p39.jpg", "text": "Page intentionally left blank\n\nPage intentionally left blank", "timestamp": "2026-07-22T06:29:20.710240+00:00"}
{"citation_id": "19930085548", "source_url": "https://ntrs.nasa.gov/api/citations/19930085548/downloads/19930085548.pdf", "page_number": 29, "total_pages": 46, "image_filename": "19930085548_p29.jpg", "text": "```markdown\n28\n\nNACA\n\nDischarge coefficient\n.8\n.6\n.4\n.2\n0\n\nExhaust port opens\nInlet port opens\n\nInlet\nExhaust\n\nInlet port closes\nExhaust port closes\n\n100 120 140 160 180 200 220 240 260\n\nCrank angle, deg A.T.C.\n\n(a) Inlet- and exhaust-flow coefficients based on maximum port area.\nAverage discharge coefficient: inlet, 0.238; exhaust, 0.372.\n\nFigure 7. - Flow coefficients of $3\\frac{1}{4}$- by $4\\frac{1}{2}$-inch ported cylinder.\n\nNACA RM No. E8L30\n```", "timestamp": "2026-07-22T06:29:22.577890+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 47, "total_pages": 78, "image_filename": "19930082483_p47.jpg", "text": "NACA TN No. 1807\n45\n\nAPPENDIX B\n\nCALCULATIONS\n\nNumerical examples illustrating the calculation methods described\nin the text are presented.\n\nCalculations from Figure 8\n\nThe operating characteristics of this turbine are computed\nfrom the information presented in figure 8 at the design operating\npoint D. For $360^\\circ$ admission, a rotor speed of 8650 rpm, and a\ntotal-pressure ratio of 2.1, power output corrected to sea level\nP equals 388 horsepower; efficiency based on total-pressure ratio\n$\\eta'$ is 0.755; inlet total or stagnation temperature $T_1'$ is 518.6$^\\circ$ R;\nand inlet total or stagnation pressure $p_1'$ equals 29.92 inches of\nmercury absolute.\n\nThe following quantities may be calculated:\n\n1. Theoretical power input based on stagnation conditions from\nequation (42)\n\n$$\n\\frac{W(\\Delta_g h')}{0.707} = \\frac{P}{\\eta'}\n$$\n\n$$\n= \\frac{388}{0.755}\n$$\n\n$$\n= 514 \\text{ hp}\n$$\n\n2. Weight flow by equation (43)\n\n$$\nW = \\frac{\\left[ \\frac{W(\\Delta_g h')}{0.707} \\right]}{\\left[ \\frac{\\Delta_g h'}{0.707} \\right]}\n$$\n\nFrom tables in reference 8,\n\n$$\n\\Delta_g h' = 23.76 \\text{ Btu/lb}\n$$", "timestamp": "2026-07-22T06:29:23.566048+00:00"}
{"citation_id": "19930082592", "source_url": "https://ntrs.nasa.gov/api/citations/19930082592/downloads/19930082592.pdf", "page_number": 47, "total_pages": 50, "image_filename": "19930082592_p47.jpg", "text": "Page intentionally left blank\n\nPage intentionally left blank", "timestamp": "2026-07-22T06:29:23.748413+00:00"}
{"citation_id": "19930082646", "source_url": "https://ntrs.nasa.gov/api/citations/19930082646/downloads/19930082646.pdf", "page_number": 31, "total_pages": 37, "image_filename": "19930082646_p31.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:29:24.939058+00:00"}
{"citation_id": "19930082918", "source_url": "https://ntrs.nasa.gov/api/citations/19930082918/downloads/19930082918.pdf", "page_number": 30, "total_pages": 62, "image_filename": "19930082918_p30.jpg", "text": "NACA TN 1940\n29\n\nTABLE 1\nEFFECT OF AGING ON INTENSITY OF (111) LINE OF LOW-CARBON\nB-175 ALLOY SOLUTION-TREATED 10 HOURS AT 2000° F\nAND WATER-QUENCHED\n\n| Aging temperature (°F) | Aging time (hr) | Intensity, $I/I_0$ (1) | Average intensity | Mean deviation from average |\n| :--- | :--- | :--- | :--- | :--- |\n| 1200 | 0.5 | 1.100<br>1.049<br>1.075<br>1.169 | 1.099 | $\\pm$.037 |\n| | 1.0 | .865<br>.935<br>.943<br>.802<br>.857 | .885 | $\\pm$.044 |\n| | 3.0 | 1.112<br>1.075<br>.975<br>.905<br>.955 | 1.016 | $\\pm$.06 |\n| | 10.0 | 1.019<br>.987<br>1.041<br>1.000 | 1.007 | $\\pm$.025 |\n| | 30.0 | 1.037<br>1.073<br>1.021<br>1.024 | 1.039 | $\\pm$.02 |\n| | 1000.0 | .654<br>.900<br>.872<br>.927 | .888 | $\\pm$.023 |\n| 1400 | .5 | .89<br>.82<br>.85 | .85 | $\\pm$.02 |\n| | 1.0 | .83<br>.87<br>.80 | .83 | $\\pm$.02 |\n| | 3.0 | .94<br>.94<br>.98 | .95 | $\\pm$.02 |\n| | 10.0 | .99<br>.99 | .99 | .00 |\n| | 30.0 | .97<br>1.05<br>1.12 | 1.05 | $\\pm$.05 |\n| | 100.0 | 1.02<br>1.18<br>1.17 | 1.12 | $\\pm$.07 |\n| | 1000.0 | 1.14<br>1.16 | 1.15 | $\\pm$.01 |\n| 1600 | .5 | .91<br>1.01 | .96 | $\\pm$.05 |\n| | 1.0 | .84<br>.88 | .86 | $\\pm$.02 |\n| | 3.0 | .88<br>.72<br>.87 | .82 | $\\pm$.07 |\n| | 10.0 | .92<br>.89<br>.93 | .91 | $\\pm$.02 |\n| | 100.0 | 1.01<br>1.05<br>1.06 | 1.04 | $\\pm$.02 |\n| | 1000.0 | 1.21<br>1.02<br>1.14 | 1.12 | $\\pm$.07 |\n\n$^1$ peak intensity of sample aged as indicated.\n$I_0$ peak intensity of unaged sample.\n\nNACA", "timestamp": "2026-07-22T06:29:25.140808+00:00"}
{"citation_id": "19930085899", "source_url": "https://ntrs.nasa.gov/api/citations/19930085899/downloads/19930085899.pdf", "page_number": 8, "total_pages": 29, "image_filename": "19930085899_p8.jpg", "text": "NACA RM No. L9A21\n\nIt would appear that the use of either fence would reduce the severity of the unstable aerodynamic-center shift considerably at low Mach numbers. (See fig. 9.) Note also that with a wing fence the stabilizing aerodynamic-center shift at M = 1.10 is reduced considerably, and $y_{cp}$ does not vary appreciably with Mach number. The fences produced only minor effects on the lift and drag characteristics. (See fig. 9.)\n\nDownwash and Dynamic-Pressure Surveys in Region of Tail Plane\n\nThe variation of isolated wing effective downwash angle with tail height and angle of attack for various Mach numbers are presented in figure 11. There is a fairly small effect of tail height and Mach number on $\\partial \\epsilon / \\partial \\alpha$ near zero lift for a tail location between 30 percent semispan above or below the wing chord line extended. (See fig. 13). At the higher lift coefficients the variations of $\\partial \\epsilon / \\partial \\alpha$ appear more erratic, but there appears to be a marked decrease in $\\partial \\epsilon / \\partial \\alpha$ for all tail locations at the highest test Mach numbers. (See fig. 11).\n\nThe results of limited downwash data obtained for the wing-fuselage condition are presented in figure 12. For angles of attack greater than $4^\\circ$ it was not possible to obtain data for the two innermost vane positions because of fuselage interference. The dashed curves in the region of the chord line extended have been estimated from unpublished results obtained for other models of this series with a free-floating tail mounted on the fuselage center line.\n\nThe results of point dynamic-pressure surveys, made in a plane perpendicular to the chord plane extended at $\\alpha = 0^\\circ$ and containing the 25-percent mean-aerodynamic-chord point of the free-floating tails used in the downwash surveys, indicate that the loss in free-stream dynamic pressure was almost always less than 10 percent up to M = 1.15 regardless of tail height. (See figs. 10 and 13.)\n\nLangley Aeronautical Laboratory\nNational Advisory Committee for Aeronautics\nLangley Air Force Base, Va.\n\nREFERENCE\n\n1. Schneiter, Leslie E., and Ziff, Howard L.: Preliminary Investigation of Spoiler Lateral Control on a $42^\\circ$ Sweptback Wing at Transonic Speeds. NACA RM No. L7F19, 1947.", "timestamp": "2026-07-22T06:29:25.383270+00:00"}
{"citation_id": "19930085890", "source_url": "https://ntrs.nasa.gov/api/citations/19930085890/downloads/19930085890.pdf", "page_number": 10, "total_pages": 26, "image_filename": "19930085890_p10.jpg", "text": "```markdown\nNACA RM No. E9C11\n\nIn the same manner, the values of experimental volume specific impulse were obtained from the values of experimental specific impulse. The maximum value for the experimental volume specific impulse as indicated by the faired curve is approximately $182 \\times 62.4$ pound-seconds per cubic foot and occurs at a ratio of fuel weight to total propellant weight of about 0.25. The maximum corrected value of volume specific impulse of $199 \\times 62.4$ pound-seconds per cubic foot is approximately 92 percent of the maximum theoretical value of $217 \\times 62.4$ pound-seconds per cubic foot, which occurs at a mixture ratio of 0.22 and is calculated for the fuel and the expansion nozzle used.\n\nA summary of the experimental results is presented in table I, which includes specific impulse, specific impulse corrected, volume specific impulse, characteristic velocity, and thrust coefficient. These results show a considerable variation of characteristic velocity and thrust coefficient under comparable conditions for these experiments.\n\nAn investigation of the accuracy of the data showed that there are three possible sources of error: in the calibration of the weighing or thrust systems, in the recording instruments used for recording the data, and in the interpretation of the data records. The instrument errors have been previously described. For data interpretation, a variation of 0.5 percent could be made in reading the thrust records and 1.0 percent could be made in reading the flow records. The maximum instrument variation and the variations interpreting data permit a maximum possible variation of 4.0 percent in specific impulse.\n\nDuring operation of the engine, a deposit accumulated over the entire inner surface of the engine. Photographs of typical deposits are shown in figure 7. Generally this deposit was approximately 1/32 inch thick from the rear of the chamber (including the surfaces of the injector nozzles) to the convergent section of the exhaust nozzle where it increased slightly. Along the convergent section to the throat of the exhaust nozzle, the deposit generally thinned to about 1/64 inch. The radial distribution of the deposit at the throat was nonuniform. The throat deposit is probably a result of afterburning of diborane during the purging operation rather than of combustion during operation. The deposit was crusty in nature, usually with a smooth, glazed, dark-gray to dark-brown surface. Some difficulties were experienced in the investigation because the high heat release caused several severe burnouts of the exhaust nozzle, the injector nozzles, and the injector head. Typical injector-plate and exhaust-nozzle failures are shown in figure 8.\n```", "timestamp": "2026-07-22T06:29:25.553173+00:00"}
{"citation_id": "19930083221", "source_url": "https://ntrs.nasa.gov/api/citations/19930083221/downloads/19930083221.pdf", "page_number": 29, "total_pages": 47, "image_filename": "19930083221_p29.jpg", "text": "NACA TN No. 1824\n\n$$\nu = \\frac{\\partial \\varphi}{\\partial x} = \\frac{1}{2\\pi} \\int_{0}^{x-\\beta r} \\frac{f'(x_1, \\mu) dx_1}{\\sqrt{(x-x_1)^2 - \\beta^2 r^2}} \\tag{41}\n$$\n\nand\n\n$$\nv_r = \\frac{\\partial \\varphi}{\\partial r} = \\frac{-1}{2\\pi r} \\int_{0}^{x-\\beta r} \\frac{(x-x_1) f'(x_1, \\mu) dx_1}{\\sqrt{(x-x_1)^2 - \\beta^2 r^2}} \\tag{42}\n$$\n\nAsymptotic values of the velocity components for large values of $r$ are readily seen to be, after first setting $x = x_0 + \\beta r$,\n\n$$\nu = \\frac{1}{2\\pi \\sqrt{2\\beta r}} \\int_{0}^{x_0} \\frac{f'(x_1, \\mu) dx_1}{\\sqrt{x_0 - x_1}} \\tag{43}\n$$\n\nand\n\n$$\nv_r = - \\frac{1}{2\\pi} \\sqrt{\\frac{\\beta}{2r}} \\int_{0}^{x_0} \\frac{f'(x_1, \\mu) dx_1}{\\sqrt{x_0 - x_1}} \\tag{44}\n$$\n\nEquations (43) and (44) may be used together with equation (38) to give for the value of drag the expression\n\n$$\nD = \\frac{\\rho_0}{8\\pi^2} \\int_{0}^{2\\pi} d\\theta \\int_{0}^{\\infty} dx_0 \\int_{0}^{x_0} \\frac{f'(x_1, \\mu) dx_1}{\\sqrt{x_0 - x_1}} \\int_{0}^{x_0} \\frac{f'(x_2, \\mu) dx_2}{\\sqrt{x_0 - x_2}} \\tag{45}\n$$\n\nAssuming that the body is of finite length so that $f'(x) = 0$ for $x > l$ reversal of the order of integration yields the relation\n\n$$\nD = - \\frac{\\rho_0}{8\\pi^2} \\int_{0}^{2\\pi} d\\theta \\int_{0}^{l} \\int_{0}^{l} f'(x_1, \\mu) f'(x_2, \\mu) \\ln |x_1 - x_2| dx_1 dx_2 \\tag{46}\n$$", "timestamp": "2026-07-22T06:29:27.076073+00:00"}
{"citation_id": "19930085859", "source_url": "https://ntrs.nasa.gov/api/citations/19930085859/downloads/19930085859.pdf", "page_number": 22, "total_pages": 31, "image_filename": "19930085859_p22.jpg", "text": "20\nNACA RM No. L9B25\n\nBending-moment coefficient, $C_B$\n\n| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | |", "timestamp": "2026-07-22T06:29:28.233550+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 59, "total_pages": 114, "image_filename": "19930086061_p59.jpg", "text": "NACA RM L9J07\n55\n\nUpper\nLower\n\nLeft semispan\nRight semispan\n\n(c) $\\psi = 20^\\circ$\n\nUpper\nLower\n\nLeft semispan\nRight semispan\n\n(d) $\\psi = 35^\\circ$\n\nFigure 16.- Concluded.", "timestamp": "2026-07-22T06:29:30.917356+00:00"}
{"citation_id": "19930085870", "source_url": "https://ntrs.nasa.gov/api/citations/19930085870/downloads/19930085870.pdf", "page_number": 13, "total_pages": 92, "image_filename": "19930085870_p13.jpg", "text": "12 CONFIDENTIAL NACA RM No. L9D07\n\nlies 3 to 4 percent ahead of its location for the wedge-leading-edge wings, probably as a result of the difference in profile and associated differences in shock locations.\n\nLiquid-Film and Schlieren Photographs\n\nSchlieren photographs were taken of wings 1, 5, and 11. Wing 1 represents the highly sweptback wing near the center of the Mach cone; wing 5, the condition of the leading edge near the Mach cone; and wing 11, the condition of supersonic leading edge for all test Mach numbers.\n\nIn figure 16(a) plan-form schlieren photographs of wedge-leading-edge wing 1 are shown for $0^\\circ$ and $4^\\circ$ angle of attack at a Mach number of 1.62. The corresponding liquid-film patterns are shown in figure 17(c), the upper surface being shown for the $4^\\circ$ angle-of-attack condition. In the schlieren photographs a distinct wake or trailing vortex may be seen leaving the trailing edge near the tips at zero angle of attack. At an angle of attack of $4^\\circ$ the vortices are much more intense and exhibit a tendency to form two distinct line vortices from either tip. The liquid-film photographs show similar patterns to exist on the wing surface. The dry regions obviously are due to the large shear intensity through momentum transfer along the lines of vorticity. It appears that the location of the outer line of vorticity approaches coincidence with the position just aft of the ridge line, at which point the adverse pressure gradient is steepest. The attendant thickening of the boundary layer favors transition, and it has been shown in the past by numerous high-speed boundary-layer investigations that the transition point coincides rather accurately with the beginning of the steep pressure rise. It is believed that the inboard lines of vorticity are the result of an overlapping effect or rolling up of the shed vortices along the transition line directly associated with the high sweep of the transition line and leading edge. The outer lines of vorticity are probably due in part to a realization of the Kutta-Joukowski condition calling for strong parallel vortices extending downstream from the point of maximum width of the airfoil. With sufficient drying time allowed, the entire area enclosed by the vorticity lines in the liquid-film tests became dry, indicating a complete turbulent region in this area. However, in order to associate the phenomenon better with that shown by the schlieren photographs, the drying time was shortened for the figures presented herein. No separation is apparent from the profile schlieren photographs of figure 18.\n\nThe plan-form schlieren photographs of wings 5 and 11 show a somewhat different phenomenon than that exhibited by wing 1. (See figs. 16(b) and 16(c).) Similar photographs of wing 5 at a Mach number of 1.92 are shown in figure 19. At zero angle of attack, shocks are seen leaving the trailing edge of each wing well inboard of the tips and are apparently composed of two or more shocks arising from points on the wing.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:29:31.486040+00:00"}
{"citation_id": "19930085862", "source_url": "https://ntrs.nasa.gov/api/citations/19930085862/downloads/19930085862.pdf", "page_number": 21, "total_pages": 60, "image_filename": "19930085862_p21.jpg", "text": "NACA RM No. L9A07\n19\n\n0.60 $\\frac{b}{2}$\n4.82\"\nDrooped-nose flap\nC\nA\nA\nB\nB\nSplit flap\n0.50 $\\frac{b}{2}$\n7.71\"\n8.40\"\n30°\nSection A-A\n(enlarged)\nSection B-B\n(enlarged)\n90°\n30°\nFence\n1.99\"\n30°\nSection C-C\n(enlarged)\n\n0.025 $\\frac{b}{2}$\n0.55 $\\frac{b}{2}$\nExtensible\nleading-edge\nflap\nE\n0.13c\n0.18c\nD\nD\nB\nB\nSplit flap\nFence\n1.99\"\n3.80\"\n37°\n$\\frac{1}{2}$\" D\nSection D-D\n(enlarged)\nSection E-E\n(enlarged)\n\nNACA\n\nFigure 3.- Details of high-lift and stall-control devices.", "timestamp": "2026-07-22T06:29:34.319298+00:00"}
{"citation_id": "19930082566", "source_url": "https://ntrs.nasa.gov/api/citations/19930082566/downloads/19930082566.pdf", "page_number": 44, "total_pages": 44, "image_filename": "19930082566_p44.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:29:35.730658+00:00"}
{"citation_id": "19930082546", "source_url": "https://ntrs.nasa.gov/api/citations/19930082546/downloads/19930082546.pdf", "page_number": 46, "total_pages": 65, "image_filename": "19930082546_p46.jpg", "text": "NACA TN No. 1870\n45\n\n<!-- Image (172, 125, 818, 481) -->\n\nFigure 9.- Effect of propeller tip shape on the free-space pressures.\n$B = 2; \\beta_{0.75} = 15^\\circ; M_t = 0.75; \\frac{d}{D} = 0.083.$\n\n<!-- Image (172, 555, 818, 868) -->\n\nFigure 10.- Effect of reflecting surfaces in the pressure field of the\nNACA 4-(5)(08)-03 propeller. $B = 2; \\beta_{0.75} = 20^\\circ; M_t = 0.60;$\n$\\frac{d}{D} = 0.083.$", "timestamp": "2026-07-22T06:29:36.688049+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 29, "total_pages": 149, "image_filename": "19930083192_p29.jpg", "text": "NACA TN 1976\n25\n\ncoefficient. The development of lift following a sudden change in angle of attack is assumed to be independent of the slope of the lift curve and can be treated as a separate subject. The neglect of possible variations in moment coefficient appears justified at this time on the basis that theory indicates that the zero-lift moment coefficient is unaffected.\n\nUnsteady-Lift Functions\n\nIn the present paper, the variation of lift coefficient for a unit change in angle of attack has been referred to as $C_{L_\\alpha}$ or as the Wagner function, and the corresponding variation in lift coefficient following a change in angle of attack due to penetrating a unit gust has been called $C_{L_g}$ or the Küssner function. The Küssner function is related to the Wagner function in that it has been derived from it by considering the airfoil penetrating a gust to be replaced by a deforming airfoil where the deformations correspond to the chordwise angle-of-attack distribution due to gust penetration. Unless otherwise stated, the lift at any time after the start of the angle-of-attack change is expressed as a fraction of the final lift coefficient.\n\nInfinite aspect ratio.- Both the Wagner function and the Küssner function for unsteady lift have been derived a number of times. The former was originally given by Wagner in reference 20 and has been checked experimentally by Walker in reference 21 by measuring the circulation about a wing following a sudden change in forward speed. Figure 24 summarizes the computed and observed circulations obtained by Walker. Although the amount of direct experimental verification is limited, several analytical studies agree and indicate that the Wagner function should be close to correct.\n\nThe Küssner function has also been computed a number of times and different results have been obtained. Figure 25, which is based on data taken from references 2, 5, and 22 to 25, presents the results of six separate computations of the lift ratio $C_{L_g}$ as a function of the distance penetrated into a sharp-edge gust and shows considerable scatter. The derivation given by Küssner in reference 5 is probably the best. No direct experimental verification of this factor is available, but Küssner has made several attempts at verification by dropping airfoils equipped with end plates through the boundary of a wind-tunnel jet. The tests made by Küssner have been described in reference 5 and the comparisons of the computed and experimental flight-path curvatures indicate good agreement, at least for penetrations up to 3 to 4 chords.", "timestamp": "2026-07-22T06:29:36.847147+00:00"}
{"citation_id": "19930082613", "source_url": "https://ntrs.nasa.gov/api/citations/19930082613/downloads/19930082613.pdf", "page_number": 40, "total_pages": 46, "image_filename": "19930082613_p40.jpg", "text": "NACA TN 1938\n39\n\n[Figure: Micrograph showing two distinct regions labeled \"Elongated grains\" and \"Equiaxed grains\"]\n\nFigure 13. - Edge of stress-relieving hole of louver. Type-A liner; etchant, 10-percent sodium cyanide, electrolytic; condition, as fabricated; magnification X750. Note fissure and elongated grains. Fissure is 0.0019 inch deep.\n\n[Figure: Micrograph showing intercrystalline cracks in a material structure. A label in the bottom right corner reads: \"NACA C-22632 12-9-48\"]\n\nFigure 14. - Intercrystalline cracks in liner run for 16 hours and 40 minutes. Type-B liner; etchant, aqua regia and glycerine; condition, cracked during accelerated-life runs; magnification, X250. Incipient intergranular cracks formed at punched edge. Note small grains at edge. Cracks are lined with oxide scale, which does not show.", "timestamp": "2026-07-22T06:29:42.687230+00:00"}
{"citation_id": "19930085847", "source_url": "https://ntrs.nasa.gov/api/citations/19930085847/downloads/19930085847.pdf", "page_number": 20, "total_pages": 32, "image_filename": "19930085847_p20.jpg", "text": "18\nCONFIDENTIAL\nNACA RM A9D04\n\nDrag-producing area\nThrust-producing area\n\n<!-- Image (193, 98, 642, 816) -->\n\n(a) M = 0.70 ($C_L = 0.21$).\n\nFigure 7.- Curves of measured chordwise and thickness-wise pressure distributions over the test panel at selected Mach numbers in the test range. No suction.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:29:43.632359+00:00"}