PDF/ocr_dataset_1060.jsonl

{"citation_id": "19930085859", "source_url": "https://ntrs.nasa.gov/api/citations/19930085859/downloads/19930085859.pdf", "page_number": 5, "total_pages": 31, "image_filename": "19930085859_p5.jpg", "text": "NACA RM No. L9B25\n3\n\n$C_m$\npitching-moment coefficient referred to $0.25\\bar{c}$\n$$ \\left( \\frac{\\text{Twice panel pitching moment}}{qS\\bar{c}} \\right) $$\n\n$C_B$\nbending-moment coefficient at plane of symmetry\n$$ \\left( \\frac{\\text{Root bending moment}}{q \\left(\\frac{S}{2}\\right) \\left(\\frac{b}{2}\\right)} \\right) $$\n\n$q$\neffective dynamic pressure over span of model, pounds per square foot $\\left( \\frac{1}{2}\\rho V^2 \\right)$\n\n$S$\ntwice wing area of semispan model, 0.1250 square foot\n\n$\\bar{c}$\nmean aerodynamic chord of wing, 0.181 foot; based on relationship $\\frac{2}{S} \\int_{0}^{b/2} c^2 dy$ (using theoretical tip)\n\n$c$\nlocal wing chord\n\n$b$\ntwice span of semispan model\n\n$y$\nspanwise distance from plane of symmetry\n\n$\\rho$\nair density, slugs per cubic foot\n\n$V$\nairspeed, feet per second\n\n$M$\neffective Mach number over span of model\n\n$M_a$\naverage chordwise local Mach number\n\n$M_l$\nlocal Mach number\n\n$R$\nReynolds number of wing based on $\\bar{c}$\n\n$\\alpha$\nangle of attack, degrees\n\n$\\epsilon$\neffective downwash angle, degrees\n\n$q_{\\text{wake}}/q$\nratio of point dynamic pressure at the quarter chord of the tail mean aerodynamic chord to free-stream dynamic pressure", "timestamp": "2026-07-22T06:15:19.639472+00:00"}
{"citation_id": "19930082592", "source_url": "https://ntrs.nasa.gov/api/citations/19930082592/downloads/19930082592.pdf", "page_number": 26, "total_pages": 50, "image_filename": "19930082592_p26.jpg", "text": "NACA TN 1914\n25\n\n<!-- Image (151, 267, 812, 666) -->\n\n(d) Composite of figures 6(a), 6(b), and 6(c).\nFigure 6. - Concluded. Effect of time, temperature, and cobalt\ncontent on oxidation penetration of titanium carbide - cobalt\nceramals.", "timestamp": "2026-07-22T06:15:21.224856+00:00"}
{"citation_id": "19930085862", "source_url": "https://ntrs.nasa.gov/api/citations/19930085862/downloads/19930085862.pdf", "page_number": 1, "total_pages": 60, "image_filename": "19930085862_p1.jpg", "text": "Copy No. 158\nRM No. L9A07\n\nFILE COPY\nNO 4\n\nRESTRICTED\n\nCLASSIFICATION CHANGE TO\nUNCLASSIFIED AUTHORITY J.W.CROWLEY\nCHANGE#1895 DATE 12-14-53 W.H.L.\n\nNACA\n\nRESEARCH MEMORANDUM\n\nLOW-SPEED INVESTIGATION OF AILERON AND SPOILER CHARACTERISTICS\nOF A WING HAVING $42^\\circ$ SWEEPBACK OF THE LEADING EDGE AND\nCIRCULAR-ARC AIRFOIL SECTIONS AT REYNOLDS NUMBERS\nOF APPROXIMATELY $6.0 \\times 10^6$\n\nBy\nStanley H. Spooner and Robert L. Woods\n\nLangley Aeronautical Laboratory\nLangley Air Force Base, Va.\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\nWASHINGTON\nMarch 10, 1949\n\nRESTRICTED\n\nTHIS DOCUMENT ON LOAN FROM THE FILES OF\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\nLANGLEY AERONAUTICAL LABORATORY\nLANGLEY FIELD, HAMPTON, VIRGINIA\n\nRETURN TO THE ABOVE ADDRESS.\nRULES FOR PUBLICATIONS SHOULD BE FOLLOWED\nAS FOLLOWS:\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n1512 H STREET, N. W.\nWASHINGTON 25, D. C.\n\nCLASSIFIED DOCUMENT\nThis document contains classified information affecting the National Defense of the United States within the meaning of the Espionage Act, USC 50:31 and 32. Its transmission or the revelation of its contents in any manner to an unauthorized person is prohibited by law. Information so classified may be imparted only to persons in the military and naval services of the United States, appropriate civilian officers and employees of the Federal Government who have a legitimate interest therein, and to United States citizens of known loyalty and discretion who of necessity must be informed thereof.", "timestamp": "2026-07-22T06:15:24.465431+00:00"}
{"citation_id": "19930082546", "source_url": "https://ntrs.nasa.gov/api/citations/19930082546/downloads/19930082546.pdf", "page_number": 25, "total_pages": 65, "image_filename": "19930082546_p25.jpg", "text": "24\nNACA TN No. 1870\n\nAPPENDIX\n\nThe response of an idealized panel to a plane sound wave is given in reference 5, page 220. The panel is assumed to move as an infinite, thin, but rigid piston that can vibrate as a whole under the action of elastic and damping restraints. The equations are reproduced here in somewhat modified form to show the effect of rigidity, mass, and damping on the response of a panel.\n\nThe vibration velocity of the panel is given by the following equation:\n\n$$\n\\dot{\\xi}_{02} e^{i\\omega_1 t} = \\frac{2K\\dot{\\xi}_{01} e^{i\\omega_1 t}}{W} \\tag{4}\n$$\n\nSubstituting $K\\dot{\\xi}_{01} = p$ and $\\dot{\\xi}_{02} = i\\omega_1 \\xi_{02}$ gives\n\n$$\n\\xi_{02} e^{i\\omega_1 t} = \\frac{2p e^{i\\omega_1 t}}{i\\omega_1 W} \\tag{5}\n$$\n\nwhere\n\n$$\nW = (C + 2K) + i\\left(M\\omega_1 - \\frac{s}{\\omega_1}\\right)\n$$\n\nThe absolute value is given by\n\n$$\n\\xi_{02} = \\frac{2p}{\\omega_1 \\sqrt{(C + 2K)^2 + \\left(M\\omega_1 - \\frac{s}{\\omega_1}\\right)^2}} \\tag{6}\n$$\n\nUtilizing the value of the critical damping for single-degree systems gives (p. 50, reference 4)\n\n$$\nC_c = 2M\\omega_n\n$$\n\n$$\n\\xi_{02} = \\frac{2p}{\\omega_1 \\sqrt{\\left(\\frac{C}{C_c} M\\omega_n + 2K\\right)^2 + \\left(M\\omega_1 - \\frac{s}{\\omega_1}\\right)^2}} \\tag{7a}\n$$", "timestamp": "2026-07-22T06:15:26.793179+00:00"}
{"citation_id": "19930082918", "source_url": "https://ntrs.nasa.gov/api/citations/19930082918/downloads/19930082918.pdf", "page_number": 14, "total_pages": 62, "image_filename": "19930082918_p14.jpg", "text": "NACA TN 1940\n13\n\nbeing nucleated, are growing by matrix depletion of the precipitate constituent atoms, or are growing by agglomeration. This usefulness of the parameter measurements is predicated upon the possibility of the average radii of the atoms making up the precipitate being somewhat larger or smaller than the average radii of the matrix atoms. If such is the case, then precipitation by matrix depletion will result in a measurable decrease or increase in the lattice parameter. Nucleation can be ascertained if the lattice parameter remains constant for a period of time at a given temperature and then increases or decreases. Precipitate growth by agglomeration in turn can be noted when the parameter has reached a steady-state value after increasing or decreasing from the initial value and yet the precipitate particles continue to grow as evidenced by metallographic examination.\n\nThe results of parameter measurements on the solution-treated and aged low-carbon N-155 are shown in figure 11. In no case was the aging time sufficient to complete the precipitation reaction by matrix depletion, because the lattice parameters did not reach a steady-state value. For this reason, the relationship between the precipitate and matrix compositions and the temperature of aging was not definitely determined. It appears, however, that the precipitates obtained by aging at $1400^\\circ$ and $1600^\\circ$ F could have been slightly different in composition, since the curves for the matrix parameter approached steady-state values which were probably not quite the same. At the end of 1000 hours, aging at $1200^\\circ$ F had resulted in so little change in lattice parameter that the only conclusion was that the major volume fraction of the material was never out of the nucleation stage. In addition a long nucleation period was shown by the lack of marked parameter change for aging up to approximately 100 hours at $1400^\\circ$ F followed by precipitate growth through matrix depletion. Only matrix depletion was found when aging was done at $1600^\\circ$ F.\n\nLine width measurements.— Since Dehlinger (reference 6) had postulated that long-period lattice distortion (of the type which could be associated with each of the small precipitated particles revealed by the metallographic examination of samples after prolonged aging) would result in line broadening, it was decided to measure the line-broadening effects. Table 2 shows the results of width measurements, expressed as ratios between any given aged sample and unaged material, of the (111) line at $\\theta = 22^\\circ$. These were obtained from the Morelco spectrometer plots. The little effect noted was due to lack of resolving power. Table 3 shows the line widths B, expressed in radians, obtained from photometer plots of film recordings of the (220) line at $\\theta = 65^\\circ$. The increased resolution in the back-reflection region revealed that appreciable broadening appeared only after long-time aging at $1400^\\circ$ F and that a smaller degree of broadening occurred rather quickly when aging was done at $1600^\\circ$ F. Table 3 also summarizes the calculations involved in converting the line broadenings to root-mean-square strains by the method of Haworth (see reference 3).", "timestamp": "2026-07-22T06:15:27.139256+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 46, "total_pages": 98, "image_filename": "19930086073_p46.jpg", "text": "```markdown\n44\n\nLift coefficient, $C_L$\n\n| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | 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| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | |", "timestamp": "2026-07-22T06:15:30.341829+00:00"}
{"citation_id": "19930082613", "source_url": "https://ntrs.nasa.gov/api/citations/19930082613/downloads/19930082613.pdf", "page_number": 24, "total_pages": 46, "image_filename": "19930082613_p24.jpg", "text": "NACA TN 1938\n23\n\n<!-- Image (150, 117, 780, 475) -->\n\nFigure 5. - Cracks at edge of air-intake hole. Type-A liner; etchant, none; condition,\nfailed in service; magnification, X100. Largest crack is approximately 1/4 inch long.\n\n<!-- Image (150, 536, 780, 894) -->\n\nFigure 6. - Cross section showing cracks extending into metal from liner surfaces.\nType-A liner; etchant, none; condition, failed in service; magnification, X100. Cracks\nshown were located near large crack.", "timestamp": "2026-07-22T06:15:30.522589+00:00"}
{"citation_id": "19930083221", "source_url": "https://ntrs.nasa.gov/api/citations/19930083221/downloads/19930083221.pdf", "page_number": 13, "total_pages": 47, "image_filename": "19930083221_p13.jpg", "text": "NACA TN No. 1824\n11\n\nSince\n\n$$C_{I_\\alpha}(t) = \\frac{1}{c_o} \\int_o^{c_o} \\frac{\\Delta p}{q\\alpha} dx$$\n\nthe following results are obtained:\n\nFirst time interval $0 < t < \\frac{c_o}{1+M_o}$\n\n$$C_{I_\\alpha}(t) = \\frac{4}{M_o} \\tag{19a}$$\n\nSecond time interval $\\frac{c_o}{1+M_o} < t < \\frac{c_o}{M_o-1}$\n\n$$C_{I_\\alpha}(t) = \\frac{4}{\\pi} \\left[ \\frac{1}{M_o} \\left( \\frac{\\pi}{2} + \\text{arc sin} \\frac{c_o - M_o t}{t} \\right) + \\frac{1}{\\sqrt{M_o^2-1}} \\text{arc cos} \\frac{t + M_o c_o - t M_o^2}{c_o} \\right.$$\n$$\\left. + \\frac{1}{M_o c_o} \\sqrt{t^2 - (c_o - t M_o)^2} \\right] \\tag{19b}$$\n\nThird time interval $\\frac{c_o}{M_o-1} < t$\n\n$$C_{I_\\alpha}(t) = \\frac{4}{\\sqrt{M_o^2-1}} \\tag{19c}$$\n\nThese results have been discussed in reference 7 for values of $M_o$ greater than one. They still hold, however, for sonic flight speeds and, in fact, can be reduced to the expressions:\n\nFirst time interval $0 < t < \\frac{c_o}{2}$", "timestamp": "2026-07-22T06:15:35.832693+00:00"}
{"citation_id": "19930085548", "source_url": "https://ntrs.nasa.gov/api/citations/19930085548/downloads/19930085548.pdf", "page_number": 11, "total_pages": 46, "image_filename": "19930085548_p11.jpg", "text": "```markdown\n10\nNACA RM No. E8L30\n\n$$P_c = P_e r^{1.35} \\quad (4)$$\n\nwhere\n\n$P_c$ compression pressure, (lb/sq in. abs.)\n$P_e$ exhaust-gas pressure, (lb/sq in. abs.)\n$r$ compression ratio\n\nis shown to result in pressures somewhat lower than the actual compression pressures. In view of the low flow coefficient across the inlet ports of the experimental cylinder, it appears doubtful that compression could proceed from inlet-manifold pressure. Actually, the pressure from which compression begins should be somewhere between inlet- and exhaust-manifold pressures. Indicator-card data show that this pressure is less than the inlet-manifold pressure by about 20 percent of the cylinder pressure drop. The data in figure 14 show that the value of 1.35 for the polytropic exponent very nearly fits the plotted points.\n\n**Maximum cylinder pressures.** - The effect of inlet-manifold pressure and fuel-air ratio on maximum cylinder pressure is shown in figure 15. The curves initially rise quite rapidly and, as the fuel-air ratio is enriched, become more nearly flat. This flatness is a result of increasing the duration of fuel injection at the richer mixtures and of operating with a fixed injection advance angle. If the injection advance had been optimized, increasing with the fuel-air ratio, these curves would be more nearly straight lines. As is to be expected, the maximum cylinder pressure at a given fuel-air ratio is approximately proportional to the inlet-manifold pressure.\n\nRatios of maximum combustion pressures to compression pressure are plotted in figure 16 for both experimental and calculated values of cylinder pressure. The calculated values were obtained by using the method of computation (for constant-volume combustion with certain correction factors applied) described in the analysis (reference 1). The calculated values of combustion pressure are higher than the experimental values and the calculated values of compression pressures $P_e r^{1.35}$ are slightly lower than the experimental values, being nearly $P_m r^{1.35}$. Thus the calculated ratios are accordingly higher than the experimental ratios. The pressure\n```", "timestamp": "2026-07-22T06:15:37.394803+00:00"}
{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 12, "total_pages": 72, "image_filename": "19930085491_p12.jpg", "text": "NACA RM No. A8J04 CONFIDENTIAL 11\n\nnumber of 0.62 million.\n\nThe values of the estimated minimum drag coefficients for the wings must be considered approximate since no solution was available for the determination of the thickness drag of wings with rounded leading edges. The theoretical values used in this analysis were obtained from the curves of reference 11, which consider variations in leading-edge sweep angle, aspect ratio, taper ratio, thickness-chord ratio, and Mach number for wings having a double-wedge profile with maximum thickness at midchord. Although this airfoil section has a sharp leading edge, the experimental results of references 5 and 12 indicate that there is little, if any, change in wing drag associated with the rounding of a sharp leading-edge section on a wing swept within the Mach cone. Probably a more important difference, between the actual wing sections and that used in the theoretical analysis, is the distribution of wing section thickness. However, this deviation will not alter the qualitative variation of wing pressure drag with changes in sweep.\n\nThe thickness drag coefficient for the fuselage alone has been determined by the method of characteristics (reference 13) and in each case has been based on the wing area of the particular configuration. Liquid-film results indicated that behind the point of intersection of the wing leading edge the fuselage boundary layer was turbulent. To account for this, an approximation of 40-percent laminar and 60-percent turbulent flow was used to obtain the total friction drag coefficient using equation (5) of reference 14. This equation assumes that the turbulent boundary layer over the rear of the fuselage is the same as would be obtained if the flow over the entire fuselage were turbulent.\n\nThe components of the drag obtained in the manner discussed previously, have been tabulated below for four of the configurations. For the most highly swept-wing configuration it was not possible to determine the wing-thickness drag coefficient from the curves of reference 11. The effects of wing-fuselage interference have been neglected.\n\n| Configuration | WF-57 | WF-60 | WF-63 | WF-67 |\n|---------------|-------|-------|-------|-------|\n| Wing-thickness drag | 0.0150 | 0.0078 | 0.0047 | 0.0026 |\n| Wing-friction drag | .0034 | .0034 | .0034 | .0034 |\n| Fuselage thickness drag | .0015 | .0015 | .0016 | .0016 |\n| Fuselage friction drag | .0034 | .0034 | .0036 | .0036 |\n| Total | 0.0233 | 0.0161 | 0.0133 | 0.0112 |\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:15:38.009890+00:00"}
{"citation_id": "19930082703", "source_url": "https://ntrs.nasa.gov/api/citations/19930082703/downloads/19930082703.pdf", "page_number": 19, "total_pages": 28, "image_filename": "19930082703_p19.jpg", "text": "NACA TN 1983\n\n[Figure: Tail-surface installation on helicopter B, showing rear fuselage with mounted horizontal stabilizers and vertical fin; NACA logo and identifier L-58392 visible in lower right of image]\n\nFigure 2.- Tail-surface installation on helicopter B.\n\nL-58392\n\n17", "timestamp": "2026-07-22T06:15:38.384916+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 12, "total_pages": 149, "image_filename": "19930083192_p12.jpg", "text": "```markdown\n8\nNACA TN 1976\n\nGUST ALLEVIATION FACTOR\n\nMuch of the data obtained from commercial operations with the NACA V-G recorder and some gust data from special investigations are not amenable to analysis by the extended equations of the previous section because of limited information or insufficiently defined acceleration peaks, and so means to account for the main effects of alleviation had to be established. Since, for much of the work, the level of gust intensity has been set by the airline data and special investigations, the results of the extended equations have been used mainly to adjust measurements of effective gust velocity for the effects of unsteady lift and vertical motion. The gust alleviation or correction factors so obtained were consequently based on a knowledge of the average gust characteristics and the response of the conventional airplanes on which considerable data had been accumulated.\n\nThe gust alleviation factor is defined as the relative response of two airplanes encountering the same gust with the gradient distance defined in chords. If a representative gust size is determined and a particular airplane selected as being representative of the airplanes used for gust measurements, then the effective gust velocities are related by the acceleration ratios. Thus, if all gust measurements were made on airplane 1 and the size of the average gust were known, then effective gust velocities from the accelerations experienced by airplane 2 are related to those for airplane 1 by the expression\n\n$$ \\frac{U_{e_1}}{\\left(\\frac{\\Delta n}{\\Delta n_g}\\right)_1} = \\frac{U_{e_2}}{\\left(\\frac{\\Delta n}{\\Delta n_g}\\right)_2} = U \\sqrt{\\frac{\\rho}{\\rho_o}} $$\n\nThe effective gust velocity representing the conditions is $U_{e_1}$ or, if measured with airplane 2, is\n\n$$ \\frac{\\left(\\frac{\\Delta n}{\\Delta n_g}\\right)_1}{\\left(\\frac{\\Delta n}{\\Delta n_g}\\right)_2} U_{e_2} $$\n```", "timestamp": "2026-07-22T06:15:38.559764+00:00"}
{"citation_id": "19930085572", "source_url": "https://ntrs.nasa.gov/api/citations/19930085572/downloads/19930085572.pdf", "page_number": 12, "total_pages": 17, "image_filename": "19930085572_p12.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:15:40.058021+00:00"}
{"citation_id": "19930082566", "source_url": "https://ntrs.nasa.gov/api/citations/19930082566/downloads/19930082566.pdf", "page_number": 28, "total_pages": 44, "image_filename": "19930082566_p28.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:15:43.759560+00:00"}
{"citation_id": "19930085862", "source_url": "https://ntrs.nasa.gov/api/citations/19930085862/downloads/19930085862.pdf", "page_number": 2, "total_pages": 60, "image_filename": "19930085862_p2.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:15:49.336389+00:00"}
{"citation_id": "19930082592", "source_url": "https://ntrs.nasa.gov/api/citations/19930082592/downloads/19930082592.pdf", "page_number": 27, "total_pages": 50, "image_filename": "19930082592_p27.jpg", "text": "```markdown\n26\nNACA TN 1914\n\nCobalt oxidation-rate constant,\n$\\Delta \\log p / \\Delta \\log t = K_{Co}$\n\nCobalt\n(percent)\nO 5\n$\\square$ 10\n$\\diamond$ 20\n$\\Delta$ 30\n\n<!-- Image (116, 179, 856, 692) -->\n\nReciprocal of absolute temperature in $^\\circ R$, $1/T$\nTemperature, $^\\circ F$\n\nFigure 7. - Effect of temperature and cobalt content on oxidation-rate constant of titanium carbide - cobalt ceramals.\n```", "timestamp": "2026-07-22T06:15:49.887093+00:00"}
{"citation_id": "19930085859", "source_url": "https://ntrs.nasa.gov/api/citations/19930085859/downloads/19930085859.pdf", "page_number": 6, "total_pages": 31, "image_filename": "19930085859_p6.jpg", "text": "4\nNACA RM No. L9B25\n\n$(L/D)_{max}$ maximum ratio of lift to drag\n\n$y_{c.p.}$ lateral center of pressure, percent semispan $(100C_B/C_L)$\n\n$h_t$ tail height relative to wing chord plane extended, percent semispan, positive for tail positions above chord plane extended\n\nTESTS\n\nThe tests were made in the Langley high-speed 7- by 10-foot tunnel utilizing an adaptation of the NACA wing-flow technique for obtaining transonic speeds. The technique used involves placing the model in the high-velocity flow field generated over the curved surface of a bump on the tunnel floor. (See reference 2.)\n\nTypical contours of local Mach number in the vicinity of the model location on the bump obtained from surveys with no model in position are shown in figure 5. It is seen that there is a Mach number gradient of about 0.04 over the model semispan at low Mach numbers and from 0.06 to 0.07 at the highest Mach numbers. The chordwise Mach number gradient is generally less than 0.01. No attempt has been made to evaluate the effects of this chordwise and spanwise Mach number gradient. Note that the long dashed lines shown near the root of the wing (fig. 5) indicate a local Mach number 5 percent below the maximum value and represent a nominal extent of the bump boundary layer. The effective test Mach number was obtained from contour charts similar to those presented in figure 5 using the relationship\n\n$$M = \\frac{2}{b/2} \\int_0^{b/2} cM_a dy$$\n\nThe variation of mean test Reynolds number with Mach number is shown in figure 6. The boundaries on the figure are an indication of the probable range in Reynolds number caused by variations in test conditions in the course of the investigation.\n\nForce and moment data, effective downwash angles, and the ratio of dynamic pressure at 25 percent of the tail mean aerodynamic chord to free-stream dynamic pressure were obtained for various model configurations through a Mach number range of 0.60 to 1.18 and an angle-of-attack range of $-2^\\circ$ to $10^\\circ$.", "timestamp": "2026-07-22T06:15:53.163507+00:00"}
{"citation_id": "19930082613", "source_url": "https://ntrs.nasa.gov/api/citations/19930082613/downloads/19930082613.pdf", "page_number": 25, "total_pages": 46, "image_filename": "19930082613_p25.jpg", "text": "Page intentionally left blank\n\nPage intentionally left blank", "timestamp": "2026-07-22T06:15:58.023521+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 47, "total_pages": 98, "image_filename": "19930086073_p47.jpg", "text": "NACA RM A9H04\n45\n\nLift coefficient, $C_L$\nDrag coefficient, $C_D$\n\nAngle of sideslip, $\\beta$, deg\n$\\circ$ 0.0\n$\\square$ 6.0\n$\\diamond$ 12.0\n$\\triangle$ 15.9\n\nFigure 9- Continued.\n(b) $C_L$ vs $C_D$.\n\n[Figure: A graph plotting Lift coefficient ($C_L$) on the vertical axis against Drag coefficient ($C_D$) on the horizontal axis. The vertical axis ranges from -0.4 to 1.2. The horizontal axis ranges from -0.1 to 0.8. There are four curves plotted with different markers corresponding to angles of sideslip ($\\beta$) of 0.0, 6.0, 12.0, and 15.9 degrees. The NACA logo is visible in the top right corner of the plot area.]", "timestamp": "2026-07-22T06:16:00.185498+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 87, "total_pages": 96, "image_filename": "19930085880_p87.jpg", "text": "NACA RM No. L9C03\n85\n\nTrimming moment, lb-ft\nSpeed (fps)\nWetted area, sq ft\n(a) $\\tau = 40$.\nFigure 24.- Variation of moment with wetted area. Model 250D.", "timestamp": "2026-07-22T06:16:08.779801+00:00"}
{"citation_id": "19930082703", "source_url": "https://ntrs.nasa.gov/api/citations/19930082703/downloads/19930082703.pdf", "page_number": 20, "total_pages": 28, "image_filename": "19930082703_p20.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:16:10.802222+00:00"}
{"citation_id": "19930082918", "source_url": "https://ntrs.nasa.gov/api/citations/19930082918/downloads/19930082918.pdf", "page_number": 15, "total_pages": 62, "image_filename": "19930082918_p15.jpg", "text": "14\nNACA TN 1940\n\nFigure 12 shows the results of such calculations. The results of\ncourse are qualitatively similar to the line widths recorded in\ntable 3.\n\nOne further point is of importance here. The fact that the\nlattice parameter decreased with aging time makes it possible for\na parameter distribution to be present as a result of concentration\ngradients being set up. Broadening of the diffraction lines will\narise from this type of parameter distribution. Broadening could\nalso arise from the elastic strains surrounding the precipitate\nparticles, such elastic strains being due to a difference in atomic\nspacings in the two lattices. Thus the broadening data presented\nhere can be ascribed to two sources and unfortunately enough data\nare not at present on hand to separate the two effects.\n\nHardness Measurements\n\nThe hardness survey was made to provide data for determining\nwhat internal condition gives high hardness. Figure 13 shows the\nresults obtained on the solution-treated stock aged at 1200°, 1400°,\nand 1600° F. From figure 13, it can be seen that conventional aging\nbehavior was apparently followed; that is, the higher the aging\ntemperature, the sooner the approach to a maximum hardness - the\nmaximum hardness, however, increasing in value with decreasing aging\ntemperature. This was certainly true for aging at 1400° and 1600° F\nand probably true for aging at 1200° F.\n\nCreep Properties\n\nThe purpose of the creep and rupture testing carried out on\nsolution-treated and aged low-carbon N-155 alloy was to measure the\nmechanical behavior of samples used in the physical measurements\ndiscussed previously. Figures 14 and 15 and table 4 show the results\nof creep testing.\n\nAt the stress level of 30,000 psi (fig. 14), aging at 1600° F\nresulted in a rather uniform increase in the secondary creep rate.\nAging at 1400° F for time periods up to approximately 10 hours resulted\nin little change in creep rate over the unaged material. For aging\ntime periods over 10 hours the creep rate increased rather rapidly\nto approach that for the material aged at 1600° F. From the two tests\nrun after aging 100 and 1000 hours at 1200° F (100 hr being somewhat\nlonger than the total period of creep testing) it appeared that aging\nat 1200° F had little effect on creep rate up to the longer aging\nperiod considered, 1000 hours, and then served only to reduce the creep\nresistance slightly.", "timestamp": "2026-07-22T06:16:11.271173+00:00"}
{"citation_id": "19930082646", "source_url": "https://ntrs.nasa.gov/api/citations/19930082646/downloads/19930082646.pdf", "page_number": 12, "total_pages": 37, "image_filename": "19930082646_p12.jpg", "text": "```markdown\nNACA TN 1980\n11\n\n[Figure: Lines of model]\n\n| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | |", "timestamp": "2026-07-22T06:16:13.279795+00:00"}
{"citation_id": "19930085548", "source_url": "https://ntrs.nasa.gov/api/citations/19930085548/downloads/19930085548.pdf", "page_number": 12, "total_pages": 46, "image_filename": "19930085548_p12.jpg", "text": "```markdown\nNACA RM No. E8L30\n\nratio increases as the compression ratio is decreased. For example, the value of the experimental pressure ratio for a compression ratio of 7 is 1.42 whereas for a compression ratio of 4 it is 1.72 (at a fuel-air ratio of 0.055).\n\nAt high compression ratios, the temperature is, of course, higher at the end of compression than at low compression ratios. This increase in temperature caused the decrease in the ratio of combustion pressure to compression pressure with increasing compression ratio (fig. 16) for the calculated data and is a contributing factor in accounting for this variation in the case of the experimental data. For the experimental data, however, a second effect is present, because combustion is incomplete at crank top dead center and because the rate of change of cylinder volume with respect to time for constant clearance volume is greater for high than for low compression ratios.\n\nThe combustion pressure rise was about one-half of that to be expected with constant-volume combustion. The combustion pressure rise apparently had little bearing on the performance of the cylinder when the limitations were maximum cylinder pressure and exhaust-gas temperature.\n\n### Thermal Efficiency\n\nThe effects of pertinent engine operating variables on indicated specific fuel consumption and thermal efficiency of the experimental cylinder are shown in figure 17, which shows that the efficiency decreases with increasing fuel-air ratio. This change is caused by the greater divergence of the properties of the working fluid from the properties of a perfect gas at the rich mixtures, the occurrence of more of the combustion process at constant pressure, and the burning of some fuel at bottom center, as previously discussed. In addition, these values include a combustion efficiency, which was not determined but was estimated to be approximately 90 to 100 percent.\n\nThe variation of thermal efficiency with compression ratio is compared to that of an air standard cycle in figure 18. The trend of the experimental data is shown to be similar to that expected from theoretical considerations.\n\nThe experimental data are compared with the efficiencies calculated by the methods of the reference analysis (reference 1) in figure 19. The figure shows that the experimental values are somewhat less than the theoretical results. In the course of the investigation,\n\n```", "timestamp": "2026-07-22T06:16:14.588965+00:00"}
{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 13, "total_pages": 72, "image_filename": "19930085491_p13.jpg", "text": "12 CONFIDENTIAL NACA RM No. A8J04\n\nDrag due to lift.— The theoretical drag-rise factor as given in reference 5 is\n\n$$\n\\frac{\\Delta C_D}{(\\Delta C_L)^2} = \\frac{k_a}{dC_L/d\\alpha}\n$$\n\nwhere $k_a$ defines the rearward inclination of the resultant force on the flat lifting surface as a fraction of the angle of attack. As was discussed in references 1 and 5, the theoretical value of $k_a$ equals one when the lifting wing has a supersonic leading edge. However, for a lifting wing with a subsonic leading edge, suction pressures develop near the leading edge (see fig. 5, which is the qualitative lifting pressure distribution determined from the results of Stewart, reference 15) reducing the rearward inclination of the resultant force and the theoretical value of $k_a$ to less than one.\n\nThe amount of theoretical leading-edge suction for a wing of the present investigation is the same as that for a swept-back triangle having the same leading edges. Based on this consideration, the following expression obtained from the results of reference 1 may be used to determine $k_a$.\n\n$$\nk_a = 1 - \\frac{S_T}{S} \\frac{\\pi m \\sqrt{1-m^2}}{\\sqrt{M_0^2-1} \\left(dC_L/d\\alpha\\right) E^2}\n$$\n\nIn this equation, $S_T$ is the area of the equivalent triangle and $E$ is the complete elliptic integral of the second kind with the modulus $\\sqrt{1-m^2}$. Values of $\\Delta C_D / (\\Delta C_L)^2$ were obtained by substituting equation (2) into equation (1).\n\nMaximum lift-drag ratio and optimum lift coefficient.— The theoretical maximum lift-drag ratio of the configurations in the present study may be determined from the minimum drag coefficient and the drag-rise factor. As was shown in reference 5\n\n$$\n(L/D)_{\\text{max}} = \\frac{1}{2} \\sqrt{\\frac{1}{C_{D_{\\text{min}}} \\left[ \\frac{\\Delta C_D}{(\\Delta C_L)^2} \\right]}}\n$$\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:16:18.044914+00:00"}
{"citation_id": "19930083221", "source_url": "https://ntrs.nasa.gov/api/citations/19930083221/downloads/19930083221.pdf", "page_number": 14, "total_pages": 47, "image_filename": "19930083221_p14.jpg", "text": "12\nNACA TN No. 1824\n\n$$C_{L_\\alpha}(t) = 4 \\tag{20a}$$\n\nSecond time interval $\\frac{c_o}{2} < t$\n\n$$C_{L_\\alpha}(t) = \\frac{4}{\\pi} \\left( \\frac{\\pi}{2} + \\text{arc sin } \\frac{c_o - t}{t} + 2 \\sqrt{\\frac{2t - c_o}{c_o}} \\right) \\tag{20b}$$\n\nThe indicial lift coefficient is seen to be constant and equal to $4\\alpha$ up to the time $t' = \\frac{c_o}{2a_o}$ or up to the time required to travel one-half chord. Following this first time interval the indicial function rises monotonically, reaching an infinitely large value as time increases. The growth of $C_{L_\\alpha}(t)$ is, of course, in agreement with the fact that the steady-state load coefficient becomes infinitely large in linear theory for $M_o=1$. This means that the theory cannot be used to predict the complete extent of the $C_{L_\\alpha}(t)$ variation with time but that during the earlier part of the motion the assumptions remain valid.\n\nIn figure 3, curves of $C_{L_\\alpha}(t)$ are plotted as functions of $\\frac{2M_o t}{c_o}$ for values of $M_o=1$ as given by equations (20a), (20b), and $M_o=1.2$, $M_o=1.4$ as given by equations (19a), (19b), and (19c). Also included in the figure are variations of $C_{L_\\alpha}(t)$ for $M_o=0$ as calculated from Wagner's results (reference 8) by R. T. Jones (reference 9) and also for $M_o=0.8$. The derivation of results leading to the $M_o=0.8$ curve will be given subsequently in this paper. The value of $C_{L_\\alpha}(t)$ at $M_o=0.4$ for a short interval of time is also drawn. The dashed portions of the curves were not calculated but were drawn to agree with the known asymptotic value of the lift function.\n\n[Figure: A 3D plot showing curves of $C_{L_\\alpha}(t)$ vs $\\frac{2M_o t}{c_o}$ for various $M_o$ values. The vertical axis is labeled $C_{L_\\alpha}(t)/\\alpha$ with ticks at 0, 4, 8. The horizontal axis is labeled $\\frac{2M_o t}{c_o}$ with ticks at 2, 4, 8, 12, 16. The depth axis is labeled $M_o$ with ticks at 0, .8, 1.2, 1.6, 2.2. A label $\\frac{2M_o t}{c_o}$ points to the horizontal axis.]\n\nFigure 3.— Indicial-lift-curve slope for Mach numbers between 0 and 1.4 shown to time required to travel 12 half-chord lengths.", "timestamp": "2026-07-22T06:16:21.740646+00:00"}
{"citation_id": "19930085572", "source_url": "https://ntrs.nasa.gov/api/citations/19930085572/downloads/19930085572.pdf", "page_number": 13, "total_pages": 17, "image_filename": "19930085572_p13.jpg", "text": "NACA RM No. FRL02\n\n11\n\nPressure\naltitude\n(ft)\nO 5,000\n□ 10,000\n◇ 20,000\n△ 30,000\n\nFuel\n—○— AN-F-58\n--●-- AN-F-32\n\nCorrected net thrust, lb\n6000\n5000\n4000\n3000\n2000\n1000\n0\n3000 4000 5000 6000 7000 8000 9000\nCorrected engine speed, rpm\n\n[Figure: Graph showing corrected net thrust vs. corrected engine speed for two fuels at various pressure altitudes, with data points and trend lines. NACA logo in bottom right corner of plot area.]\n\nFigure 2. - Comparison of corrected net thrust with AN-F-58 and\nAN-F-32 fuels. Mach number, 0.37.", "timestamp": "2026-07-22T06:16:21.995867+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 13, "total_pages": 149, "image_filename": "19930083192_p13.jpg", "text": "NACA TN 1976\n\nThe ratio of acceleration ratios is the gust alleviation factor and may be determined by calculation or by experiment.\n\nFor load calculations, a specific alleviation factor K was derived and is shown in figure 5, as reproduced from reference 1. The factor K is calculated with the Boeing B-247 airplane as a reference and the assumptions were made that wing loading is proportional to mass parameter and that the effect of pitch on the gust load increment is the same for all airplanes. The curve is calculated as the acceleration ratio for any airplane to the acceleration ratio for the Boeing B-247 airplane $\\left(\\frac{W}{C} = 16,\\right.$ C = 11) when both airplanes traverse a gust with a gradient distance of 10 chords. The resulting curve is based on the assumption that all airplanes are similar to the Boeing B-247 airplane but have different wing loadings.\n\nTHE STRUCTURE OF ATMOSPHERIC GUSTS\n\nThe determination of the characteristics of atmospheric gusts pertinent to the airplane, their variations, and their frequency of occurrence in the atmosphere is of fundamental importance. Considerable research has been undertaken, therefore, to cover various aspects of the gust-structure problem. The research performed includes flight tests in the neighborhood of Langley Air Force Base over a period of about 10 years and special investigations made in conjunction with commercial airlines, the military services, and the Civil Aeronautics Administration.\n\nMETHODS OF GUST-STRUCTURE MEASUREMENTS\n\nThe general method followed in the investigation of gust structure has been to fly airplanes in rough air and to deduce the gust characteristics from the reactions of the airplane. Whenever possible, the airplane which was the primary instrument, as indicated in reference 7, has been calibrated by testing dynamically similar scaled models in the Langley gust tunnel in order to obtain the actual relation between acceleration and gust intensity for various gust shapes. Most of the data on over-all gust structure have been obtained from records of acceleration and airspeed as a function of time. These records have provided information on the gust velocities, the gradient distance, and the gust spacing. In the evaluation of the records, the pitching motion of the airplane is neglected except in those cases where it can be included through calibrations made in the gust tunnel.", "timestamp": "2026-07-22T06:16:23.715277+00:00"}
{"citation_id": "19930082592", "source_url": "https://ntrs.nasa.gov/api/citations/19930082592/downloads/19930082592.pdf", "page_number": 28, "total_pages": 50, "image_filename": "19930082592_p28.jpg", "text": "NACA TN 1914\n27\n\n[Figure: Micrograph showing a circular cross-section with labeled regions: Bakelite (top), Oxide layer, Oxidation interface, and Unoxidized ceramal (bottom). A NACA stamp is present in the lower right corner of the image.]\n\nNACA\nC-22907\n2-7-49\n\nFigure 8. - Oxidation of 30-percent-molybdenum - titanium carbide ceramal. Temperature, 1785° F; time at temperature, 7 hours; unetched; magnification X50.", "timestamp": "2026-07-22T06:16:25.769129+00:00"}
{"citation_id": "19930082566", "source_url": "https://ntrs.nasa.gov/api/citations/19930082566/downloads/19930082566.pdf", "page_number": 29, "total_pages": 44, "image_filename": "19930082566_p29.jpg", "text": "NACA TN NO. 1899\n\n[Figure: Panel for measurement of pressures, showing various gauges and valves labeled M, P, H, I, A, F, G, C, D, R]\n\nFigure 10.- Panel for measurement of pressures.\n\n27", "timestamp": "2026-07-22T06:16:26.503313+00:00"}
{"citation_id": "19930085847", "source_url": "https://ntrs.nasa.gov/api/citations/19930085847/downloads/19930085847.pdf", "page_number": 9, "total_pages": 32, "image_filename": "19930085847_p9.jpg", "text": "NACA RM A9D04 CONFIDENTIAL 7\n\nnumber can be traced in large part to the formation of shock and subsequent rearward shock movement on the lower surface, although some of the increase (roughly 30 percent) is due to the upper-surface separation then present. It would appear from these results that the major portion of the drag rise on the test airfoil comes from the formation of supersonic regions on the upper and lower surfaces (as has also been noted for other airfoils in references 4 and 5). Since the beneficial effect of suction is anticipated to derive solely from the elimination of separation, these results indicate the limited possibilities of boundary-layer suction on the test airfoil even if fully effective in causing reattachment of the separated wake.\n\nCONCLUDING REMARKS\n\nThe results of a limited flight investigation to study the effect of boundary-layer suction behind the shock wave on airfoil drag at supercritical Mach numbers showed no measurable effect on the test airfoil for the suction coefficient attainable. An inspection of the pressure distributions revealed that the major portion of the drag increase with Mach number was due to pressure changes directly associated with supersonic flow on upper and lower surfaces, and a minor portion (never exceeding 30 percent) was attributable to upper-surface separation behind the shock wave. Thus, the possible drag reduction due to elimination of flow separation behind the upper-surface shock wave by means of suction was limited. The extent to which this condition applies to other airfoils or other angles of attack cannot be inferred from the results of the subject tests.\n\nAmes Aeronautical Laboratory,  \nNational Advisory Committee for Aeronautics,  \nMoffett Field, Calif.\n\nREFERENCES\n\n1. Regenscheit, B.: Drag Reduction by Suction of the Boundary Layer Separated Behind Shock Wave Formation at High Mach Numbers. NACA TM 1168, 1947.\n\n2. Fage, A., and Sargent, R. F.: Effect on Aerofoil Drag of Boundary-Layer Suction Behind a Shock Wave. R. & M. No. 1913, Oct. 26, 1943.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:16:32.415204+00:00"}
{"citation_id": "19930082613", "source_url": "https://ntrs.nasa.gov/api/citations/19930082613/downloads/19930082613.pdf", "page_number": 26, "total_pages": 46, "image_filename": "19930082613_p26.jpg", "text": "NACA TN 1938\n25\n\n[Figure: Micrograph showing transcrystalline cracking in a liner]\n\nNACA\nC-22627\n12-9-48\n\nFigure 7. - Transcrystalline cracking in liner run for 66 hours and 57 minutes. Type-B liner; etchant, 5-percent aqua regia in water, electrolytic; magnification, X350.", "timestamp": "2026-07-22T06:16:33.257839+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 48, "total_pages": 98, "image_filename": "19930086073_p48.jpg", "text": "46\nNACA RM A9H04\n\n[Figure: A graph plotting Lift coefficient, $C_L$ versus Pitching-moment coefficient, $C_m$. The vertical axis ranges from -0.4 to 1.4. The horizontal axis ranges from 0.12 to -0.08. There are four curves plotted with different symbols: circles, squares, diamonds, and triangles. The NACA logo is visible in the bottom right corner of the graph area.]\n\nLift coefficient, $C_L$\n\nPitching-moment coefficient, $C_m$\n\n$\\bigcirc$ $\\square$ $\\diamond$ $\\triangle$\n0.0 6.0 12.0 15.9\nAngle of sideslip, $\\beta$, deg\n\n(c) $C_L$ vs $C_m$.\n\nFigure 9.— Continued.", "timestamp": "2026-07-22T06:16:33.587465+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 88, "total_pages": 96, "image_filename": "19930085880_p88.jpg", "text": "86\nNACA RM No. L9C03\n\n<!-- Image (109, 133, 881, 935) -->\n\n(b) $\\tau = 8^\\circ$.\nFigure 24.-- Continued.", "timestamp": "2026-07-22T06:16:38.628927+00:00"}
{"citation_id": "19930085862", "source_url": "https://ntrs.nasa.gov/api/citations/19930085862/downloads/19930085862.pdf", "page_number": 3, "total_pages": 60, "image_filename": "19930085862_p3.jpg", "text": "NACA RM No. L9A07\nRESTRICTED\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nRESEARCH MEMORANDUM\n\nLOW-SPEED INVESTIGATION OF AILERON AND SPOILER CHARACTERISTICS\nOF A WING HAVING $42^\\circ$ SWEEPBACK OF THE LEADING EDGE AND\nCIRCULAR-ARC AIRFOIL SECTIONS AT REYNOLDS NUMBERS\nOF APPROXIMATELY $6.0 \\times 10^6$\n\nBy Stanley H. Spooner and Robert L. Woods\n\nSUMMARY\n\nA low-speed investigation has been conducted in the Langley 19-foot\npressure tunnel at Reynolds numbers from $5.3 \\times 10^6$ to $6.9 \\times 10^6$ to deter-\nmine the effectiveness of a conventional aileron and of various spanwise\nspoiler arrangements on a $42^\\circ$ sweptback wing. The wing had an aspect\nratio of 3.94, a taper ratio of 0.625, and thin, symmetrical, circular-\narc airfoil sections. The rolling-moment characteristics of the aileron\nand the spoilers, together with the aileron hinge-moment, normal-force,\nand balance-chamber pressure characteristics were determined for both\nthe plain wing and the wing equipped with various high-lift and stall-\ncontrol devices.\n\nThe results of the investigation indicate that the effectiveness of\nthe aileron $C_{l\\delta}$ on the plain wing decreased slightly at high angles of\nattack. At low angles of attack, the effectiveness of the aileron was\napproximately the same regardless of the flap configuration. As the\nangle of attack was increased, however, deflection of inboard-located,\nhalf-span, split flaps resulted in a loss of aileron effectiveness.\nThe combination of leading-edge flaps and stall-control fences almost\nentirely offset the detrimental effects which resulted when the split\nflaps were deflected. For the plain wing configuration, the aileron\nhinge-moment characteristics were such that a conventional, sealed,\ninternal aerodynamic balance of approximately 30 percent of the aileron\nchord would be required to completely balance the aileron at low angles\nof attack. With this amount of balance, the aileron probably would be\nunderbalanced at high angles of attack of the plain wing and at all\nangles of attack of the flapped configurations. When stalling occurred\non the outboard portions of the wing, as it did without the stall-\ncontrol devices, an inboard spoiler location was more effective than an\noutboard location, and when inboard stalling occurred the outboard\nspoiler location proved more effective. The spoilers on the plain wing\nbecame ineffective in the maximum lift range. The maximum rolling\neffectiveness of the 10-percent-chord step spoilers on the wing equipped\nwith the high-lift and stall-control devices was equivalent to that\nproduced by a total aileron deflection of approximately $35^\\circ$.\n\nRESTRICTED", "timestamp": "2026-07-22T06:16:38.820249+00:00"}
{"citation_id": "19930082546", "source_url": "https://ntrs.nasa.gov/api/citations/19930082546/downloads/19930082546.pdf", "page_number": 26, "total_pages": 65, "image_filename": "19930082546_p26.jpg", "text": "NACA TN No. 1870\n\nWhen $s = M\\omega_n^2$ is substituted equation (7a) may be written as\n\n$$\n\\xi_{02} = \\frac{2p}{\\omega_1 \\sqrt{\\left(2\\frac{C}{C_c}M\\omega_n + 2K\\right)^2 + M^2\\omega_1^2 \\left[1 - \\left(\\frac{\\omega_n}{\\omega_1}\\right)^2\\right]^2}}\n\\tag{7b}\n$$\n\nFor the case of zero damping, radiation resistance, and stiffness, equation (7a) reduces to\n\n$$\n\\xi_{02} = \\frac{2p}{M\\omega_1^2}\n\\tag{8}\n$$\n\nThis is the same equation as equation (3) in text with the exception of the factor 2. The pressure used in the text is the pressure at the panel surface which for a large plane panel is double the free-space pressure because of reflection. The above equations are based on the free-space pressure of the incident wave.\n\nThe resonant condition of the panel is given by $\\omega_1 = \\omega_n$. For this condition the amplitude of vibration is given by\n\n$$\n\\xi_{02} = \\frac{p}{\\frac{C}{C_c}M\\omega_n^2 + K\\omega_n}\n\\tag{9}\n$$\n\nThe relation of the panel vibration amplitude to air amplitude at resonance may be written as\n\n$$\n\\frac{\\xi_{02}}{\\xi_{01}} = \\frac{1}{1 + \\frac{C}{C_c} \\frac{M\\omega_n}{K}}\n\\tag{10}\n$$\n\nEquation (10) shows that if the structural damping $\\frac{C}{C_c}$ is zero, the panel amplitude at resonance is equal to the amplitude of the impinging sound wave. The term $\\frac{C}{C_c} \\frac{M\\omega_n}{K}$ must be greater than unity for the damping to make an appreciable difference in the amplitude. The value of this", "timestamp": "2026-07-22T06:16:44.902146+00:00"}
{"citation_id": "19930082703", "source_url": "https://ntrs.nasa.gov/api/citations/19930082703/downloads/19930082703.pdf", "page_number": 21, "total_pages": 28, "image_filename": "19930082703_p21.jpg", "text": "Top view\n\n103 in.\n\nChord, 12 in.\n\n$25\\frac{1}{2}$ in.\n\nR = 24 ft\n\nPlane normal to rotor shaft axis (through flapping hinges)\n\n0.61R\n\n0.24R\n\nSide view\n\nRotor shaft axis\n\nGap, $11\\frac{1}{2}$ in.\n\n$\\frac{1}{4}$ M.A.C. tail\n\nNACA\n\nFigure 3.- Location and principal dimensions of biplane tail surface of helicopter B. Angle of incidence of tail surface is $0^\\circ$ relative to plane normal to rotor shaft axis.\n\nNACA TN 1983\n\n19", "timestamp": "2026-07-22T06:16:45.137846+00:00"}
{"citation_id": "19930085859", "source_url": "https://ntrs.nasa.gov/api/citations/19930085859/downloads/19930085859.pdf", "page_number": 7, "total_pages": 31, "image_filename": "19930085859_p7.jpg", "text": "NACA RM No. L9B25\n\nNo tares have been applied to the data to account for the presence of the end plates on the models. Jet-boundary corrections have not been evaluated because the boundary conditions to be satisfied are not rigorously defined. However, inasmuch as the effective flow field is large compared with the span and chord of the model the corrections are believed to be small.\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.) For the wing-fuselage tests the effective downwash angles at the chord plane extended were determined by mounting a free-floating tail on the center line of the fuselage. 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 the angle 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 $0.2^\\circ$ 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 with no model in position.\n\nRESULTS AND DISCUSSION\n\nA table of the figures presenting the results is given as follows:\n\n| Wing-alone force data | Figure |\n| --- | --- |\n| Wing-fuselage force data | 7 |\n| Effective downwash angles (wing alone) | 8 |\n| Effective downwash angles (wing fuselage) | 9 |\n| Downwash gradients | 10 |\n| Dynamic-pressure surveys | 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-coefficient range of $\\pm 0.1$ of the nominal lift coefficient.", "timestamp": "2026-07-22T06:16:47.939747+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 47, "total_pages": 114, "image_filename": "19930086061_p47.jpg", "text": "```markdown\nNACA RM L9J07\n\n$\\alpha = 4.1^\\circ$\n$C_L = 0.13$\n\n$\\alpha = 8.1^\\circ$\n$C_L = 0.28$\n\n$\\alpha = 14.1^\\circ$\n$C_L = 0.50$\n\n$\\alpha = 24.1^\\circ$\n$C_L = 0.83$\n\n| P | | | |\n| :--- | :--- | :--- | :--- |\n| -2 | | | |\n| -1 | | | |\n| 0 | | | |\n| 1 | | | |\n\n| | Station 4 | | |\n| :--- | :--- | :--- | :--- |\n| | $\\frac{y}{b/2}, 0.500$ | | |\n\n| | Station 5 | | |\n| :--- | :--- | :--- | :--- |\n| | $\\frac{y}{b/2}, 0.667$ | | |\n\n| | Station 6 | | |\n| :--- | :--- | :--- | :--- |\n| | $\\frac{y}{b/2}, 0.833$ | | |\n\n| | Station 7 | | |\n| :--- | :--- | :--- | :--- |\n| | $\\frac{y}{b/2}, 0.916$ | | |\n\n| | | | |\n| :--- | :--- | :--- | :--- |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n\n| | | | |\n| :--- | :--- | :--- | :--- |\n| | | | 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| | | |\n| | | | |\n\n| | | | |\n| :--- | :--- | :--- | :--- |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n\n| | | | |\n| :--- | :--- | :--- | :--- |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n\n| | | | |\n| :--- | :--- | :--- | :--- |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n\n| | | | |\n| :--- | :--- | :--- | :--- |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n\n| | | | |\n| :--- | :--- | :--- | :--- |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n\n| | | | |\n| :--- | :--- | :--- | :--- |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n\n| | | | |\n| :--- | :--- | :--- | :--- |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n\n| | | | |\n| :--- | :--- | :--- | :--- |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n\n| | | | |\n| :--- | :--- | :--- | :--- |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n\n| | | | |\n| :--- | :--- | :--- | :--- |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n\n| | | | |\n| :--- | :--- | :--- | :--- |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n\n| | | | |\n| :--- | :--- | :--- | :--- |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n\n| | | | |\n| :--- | :--- | :--- | :--- |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n\n| | | | |\n| :--- | :--- | :--- | :--- |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n\n| | | | |\n| :--- | :--- | :--- | :--- |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n\n| | | | |\n| :--- | :--- | :--- | :--- |\n|", "timestamp": "2026-07-22T06:16:48.120399+00:00"}
{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 14, "total_pages": 72, "image_filename": "19930085491_p14.jpg", "text": "NACA RM No. A8J04 CONFIDENTIAL 13\n\nand the lift coefficient for maximum lift-drag ratio is\n\n$$\nC_{Lopt} = \\sqrt{\\frac{C_{Dmin}}{\\Delta C_D / (\\Delta C_L)^2}}\n$$\n\nThus the maximum lift-drag ratio depends equally upon the minimum drag coefficient and the drag-rise factor.\n\nTheoretical Location of Line of Laminar\n\nSeparation on Basic Configuration at Zero Lift\n\nThe pressure distribution shown in figure 6 has been used to determine the theoretical line of laminar boundary-layer separation on the wing of WF-63 at zero angle of attack. To obtain this pressure distribution, a graphical modification of the method of Jones (reference 2) was used. In applying the method, the streamwise airfoil sections were approximated as closely as possible with symmetrical 16-sided polygons; the lengths of the sides were shortest where the section curvature was the greatest. It is believed that this pressure distribution is sufficiently accurate for the prediction of the line of laminar separation although the pressures near the leading edge, because of the large wedge angle required to fit the nose radius, are uncertain. (Because of the uncertainty at the leading edge, this pressure field was not used in the previously discussed determination of the wing pressure drag.)\n\nThe heavy solid line shown on the half-wing plan form in figure 6 is the theoretical line of laminar boundary-layer separation which was determined from the theoretical pressure distributions on sections normal to the wing leading edge. In reference 16, von Kármán and Millikan have shown that the point of laminar boundary-layer separation depends only upon the location of the section minimum-pressure coefficient and the rate of pressure recovery behind the minimum-pressure point. The criterion for separation developed in that reference has been used with the pressure distribution of figure 6 by applying the results of Jones (reference 17) which indicate that the characteristics of a laminar boundary layer on an oblique cylinder are determined entirely by the normal-flow Mach and Reynolds numbers. When the nomenclature of reference 16 is used, the pressure distributions normal to the wing leading edge (fig. 6) are the double-roof-profile type since the minimum-pressure location occurs\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:16:51.835378+00:00"}
{"citation_id": "19930082592", "source_url": "https://ntrs.nasa.gov/api/citations/19930082592/downloads/19930082592.pdf", "page_number": 29, "total_pages": 50, "image_filename": "19930082592_p29.jpg", "text": "Page intentionally left blank\n\nPage intentionally left blank", "timestamp": "2026-07-22T06:16:52.473160+00:00"}
{"citation_id": "19930082566", "source_url": "https://ntrs.nasa.gov/api/citations/19930082566/downloads/19930082566.pdf", "page_number": 30, "total_pages": 44, "image_filename": "19930082566_p30.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:16:54.606724+00:00"}
{"citation_id": "19930083221", "source_url": "https://ntrs.nasa.gov/api/citations/19930083221/downloads/19930083221.pdf", "page_number": 15, "total_pages": 47, "image_filename": "19930083221_p15.jpg", "text": "NACA TN No. 1824\n13\n\nApplication of the indicial lift function at $M_o=1$.— Once the\nindicial lift function is known, it is possible to determine the\nlift corresponding to a given variable motion. Consider, as an\nidealized example, the case\nwhere the airfoil experiences\nan abrupt rising and sinking\nmotion at regular time inter-\nvals. Such a motion involves\nabrupt plus and minus angles\nof attack without rotation or\npitching so that $\\alpha(t')$ is\ngiven by the meander or square-\nwave function shown in figure\n4(a). In this example the\nvariation of $\\alpha$ is such that\nthe curve for $C_L(t')$ can be\ncalculated easily. In figure\n4(b) $C_L(t')$ is shown for the\ncase in which the discontin-\nuities occur at intervals of\ntime equal to $c_o/V_o$, that is,\nafter each chord length of\ntravel. The principal point\nof interest in this example\nis the fact that such a\nmotion yields no excessive\nvalue of lift or perturbation\nvelocities and the entire\nanalysis is within the frame-\nwork of linear methods.\n\n[Figure: (a) Impressed angle of attack.]\n[Figure: (b) Resulting variation of lift.]\nFigure 4.— Lift resulting from\nsquare-wave angle-of-attack\nvariation.\n\nWhen the variable motion is more complex in character the lift\ncoefficient can be expressed by means of Duhamel's integral. Corre-\nsponding to the angle-of-attack variation $\\alpha(t')$ as a function of\ntime, the lift coefficient $C_L(t')$ is given by the expression\n\n$$C_L(t') = \\frac{d}{dt'} \\int_0^{t'} C_{L\\alpha}(t'-\\tau') \\alpha(\\tau') d\\tau' \\quad (21)$$\n\nIn analysis related to equation (21) it is convenient to employ\ntechniques associated with the use of the Laplace transformation. (See\nreference 10.) Thus, if the Laplace transform $\\bar{F}(s)$ of the function\n$f(t)$ is defined by the relation", "timestamp": "2026-07-22T06:16:55.143398+00:00"}
{"citation_id": "19930085572", "source_url": "https://ntrs.nasa.gov/api/citations/19930085572/downloads/19930085572.pdf", "page_number": 14, "total_pages": 17, "image_filename": "19930085572_p14.jpg", "text": "12\nNACA RM No. E8L02\n\n[Figure: A line graph plotting Corrected jet-fuel consumption against Corrected engine speed. The graph includes a legend for Pressure altitude (5,000, 10,000, 20,000, 30,000 ft) and Fuel (AN-F-58, AN-F-32). The NACA logo is visible in the bottom right corner of the plot area.]\n\nCorrected jet-fuel consumption, lb/hr\nCorrected engine speed, rpm\n\nFigure 3. - Comparison of corrected jet-fuel consumption with AN-F-58 and AN-F-32 fuels. Mach number, 0.37.", "timestamp": "2026-07-22T06:16:55.413150+00:00"}
{"citation_id": "19930082613", "source_url": "https://ntrs.nasa.gov/api/citations/19930082613/downloads/19930082613.pdf", "page_number": 27, "total_pages": 46, "image_filename": "19930082613_p27.jpg", "text": "Page intentionally left blank\n\nPage intentionally left blank", "timestamp": "2026-07-22T06:16:55.778465+00:00"}
{"citation_id": "19930082918", "source_url": "https://ntrs.nasa.gov/api/citations/19930082918/downloads/19930082918.pdf", "page_number": 16, "total_pages": 62, "image_filename": "19930082918_p16.jpg", "text": "NACA TN 1940\n\nSome ambiguity was present in determining the creep rate and aging relationship for aging at $1400^\\circ$ F. However, in view of the magnitude of experimental errors involved, it is felt that the curve shown is a reasonable compromise. In any event the trend that long-time aging at $1400^\\circ$ or $1500^\\circ$ F tends to reduce markedly the creep resistance was clearly shown.\n\nAt the stress level of 60,000 psi the effect of aging at either $1400^\\circ$ or $1600^\\circ$ F seemed qualitatively the same as at 30,000 psi except for the lower relative creep strength of the unaged material and the specimens aged 1 hour at $1400^\\circ$ F. It will be further noted that the results at 60,000 psi with the two types of extensometer agreed only within an average factor of 1.5. Previous experience with creep testing leads to the conclusion that creep rates are generally reproducible only within a factor of about the same magnitude.\n\nRupture Characteristics\n\nFigures 16, 17, and 18 and table 5 show the results of the rupture testing on material aged at $1400^\\circ$ F. When aging was carried out at $1400^\\circ$ F, a gradual approach to a flat maximum in very short-time rupture strength occurred with increased aging time. Subject to additional investigation, for rupture times up to 10 hours, it appeared that no appreciable alteration of the initial structure occurred by virtue of reactions at the test temperature, and that the preceding result was thus due to the initial structure of the material. With increased time for rupture (greater than 10 hr), the main alteration in the relationship of rupture time and aging time appeared to be the marked improvement of the unaged and short-time-aged material in comparison with the long-time-aged material. (See fig. 16.) This improvement most probably was due to reactions occurring in the material during testing and will be considered under DISCUSSION OF RESULTS. Inspection of figure 17 also shows that the aging periods at $1400^\\circ$ F for maximum short-time rupture strength were associated with the highest values of deformation prior to fracture. The initial deformation which occurred upon loading also increased slightly with increased aging time. Some alteration of these deformation characteristics was evident with increasing test time, because the curves of figure 17 in general slope downward. Again the experimental error involved in determining rupture times and deformation obscured to a certain extent the exact shape of the curves in figures 18 and 17. However, trends were clearly established. From figure 18, it is evident that, upon aging, the mode of failure changed from being primarily intergranular to being both intergranular and transgranular.\n\nFigures 19, 20, and 21 and table 5 show the results of rupture testing on material aged at $1600^\\circ$ F. At this temperature a very broad maximum in the short-time rupture strength occurred with aging times longer than 0.5 hour (see fig. 19). A very slight maximum possibly", "timestamp": "2026-07-22T06:16:56.143550+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 49, "total_pages": 98, "image_filename": "19930086073_p49.jpg", "text": "```markdown\nNACA RM A9E04\n\nLift coefficient, $C_L$\n\n$\\beta$, deg\n$\\circ$ 0.0\n$\\square$ 6.0\n$\\diamond$ 12.0\n$\\triangle$ 15.9\n\nRolling-moment coefficient, $C_l$\n\nYawing-moment coefficient, $C_n$\n\nSide-force coefficient, $C_Y$\n\n(d) $C_L$ vs $C_l$, $C_n$ and $C_Y$.\n\nFigure 9.— Concluded.\n\n47\n```", "timestamp": "2026-07-22T06:16:59.113506+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 14, "total_pages": 149, "image_filename": "19930083192_p14.jpg", "text": "10\nNACA TN 1976\n\nIn addition to measurements of over-all gust structure, some\nmeasurements have been made of the distribution of gust velocity along\nthe airplane span by recording the local wing pressures at various\nstations along the span and the airspeed. From thin-airfoil theory, the\nchange in pressure at any location on an airfoil divided by the dynamic\npressure q is a function of the angle-of-attack change and that, in\nturn, is equal to the effective gust velocity divided by the equivalent\nairspeed of the airplane. Local values of true gust velocity cannot be\nmeasured because the motions of the airplane are somewhat involved and\ncannot be taken into account.\n\nThe intensity of horizontal gusts has been obtained from airspeed\nrecords on the assumption that the absolute velocity of the airplane\nremains constant in spite of rapid changes in wind speed which the air-\nplane may experience. The apparent changes in airspeed are taken as the\nhorizontal gust velocities when not associated with large normal\naccelerations.\n\nAPPARATUS AND TESTS\n\nFive airplanes having the characteristics shown in table I and\nfigure 6 have been used to obtain gust-structure data. As may be noted\nfrom figure 6, all airplanes except the XBM-1 were monoplanes, and the\nF-61C airplane had twin tail booms instead of the usual fuselage. Table I\nshows that the weight of the airplanes varied from about 700 pounds\nto 55,000 pounds, and the wing mean geometric chord varied from about\n4 feet for the Aeronca C-2 airplane to about 19 feet for the XB-15 air-\nplane. The \"sea-level\" mass parameters of the various airplanes varied\nfrom about 6 to 23. In the case of the Aeronca C-2 and the XBM-1 air-\nplanes, the response to known gusts had been obtained by tests in the\nLangley gust tunnel (references 7 and 8, respectively).\n\nFor the determination of the gust intensity, gradient distance, and\ngust spacing, all airplanes carried a standard set of instruments\nconsisting of:\n\n(1) Recording accelerometer located at the center of gravity\n\n(2) Airspeed altitude recorder\n\n(3) Synchronizing timer", "timestamp": "2026-07-22T06:16:59.282524+00:00"}
{"citation_id": "19930085847", "source_url": "https://ntrs.nasa.gov/api/citations/19930085847/downloads/19930085847.pdf", "page_number": 10, "total_pages": 32, "image_filename": "19930085847_p10.jpg", "text": "8\nCONFIDENTIAL\nNACA RM A9D04\n\n3. Beals, Donald D., and Mourhess, Mary J.: Numerical Evaluation of\nthe Wake-Survey Equations for Subsonic Flow Including the Effect\nof Energy Addition. NACA ARR, Nov. 1945.\n\n4. Oswatitsch, K.: The Drag Increase at High Subsonic Speeds. R.A.E.\nTN No. Aero 1919, Oct. 1947.\n\n5. Davies, H.: Flight Research at High Subsonic Speeds. Jour. of the\nRoyal Aero. Society, Aug. 1948, vol. 52, no. 452, pp. 483-512.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:16:59.413771+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 89, "total_pages": 96, "image_filename": "19930085880_p89.jpg", "text": "NACA RM No. L9C03\n87\n\nTrimming moment, lb-ft\nSpeed (fps)\nWetted area, sq ft\n(c) $\\tau = 12^\\circ$.\nFigure 24.- Continued.\nNACA", "timestamp": "2026-07-22T06:17:00.218087+00:00"}