PDF/ocr_dataset_1059.jsonl

{"citation_id": "19930082566", "source_url": "https://ntrs.nasa.gov/api/citations/19930082566/downloads/19930082566.pdf", "page_number": 23, "total_pages": 44, "image_filename": "19930082566_p23.jpg", "text": "NACA TN No. 1889\n21\n\n[Figure: Drive unit and eccentric. Labels visible on the machinery include G1, G2, and E1.]\n\nFigure 7.- Drive unit and eccentric.", "timestamp": "2026-07-22T06:13:12.209991+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 33, "total_pages": 78, "image_filename": "19930082483_p33.jpg", "text": "NACA TN No. 1807\n31\n\ncontrol curve in figure 12 serves an additional purpose in that it\norients the areas of turbine operation at any degree of admission\nwith respect to corresponding areas at other degrees of admission.\nThis orientation permits rapid visual estimates to be made regarding\nthe characteristics for all mutual relations of these control\nparameters.\n\nIn figure 13, the parameter of corrected rotor speed is intro-\nduced. The construction of the plot is similar to that of figure 12\nwith the exception that corrected rotor speed replaces inlet pressure\nas a variable. For the plot shown in figure 13, the inlet total\npressure is maintained constant at 45 inches of mercury absolute.\n\nBecause the points of peak efficiency occur at lower speeds\nwith decreasing pressure ratio, the planes formed by the operational\nareas appear to fold back upon themselves, with the effect magnified\nat the lower admissions.\n\nACCURACY\n\nThe calculated probable error in the over-all efficiency over\nmost of the range of turbine operation is within $\\pm 1$ point. Power-\noutput calculations over this range are accurate within $\\pm 1/2$ percent\nof the true net power output. The greatest error occurs at the\nlowest power levels with reduced admission, reduced inlet pressure,\nand high rotor speed. At these conditions, efficiency is accurate\nwithin $\\pm 2$ points and power output is accurate within $\\pm 4.6$ percent.\n\nTemperature readings are considered accurate to $\\pm 5^\\circ$ R, turbine\nspeed measurements to $\\pm 5$ rpm, turbine-torque estimates to $\\pm 15.09$ inch-\npounds, and manometer readings to $\\pm 0.05$ inch of mercury.\n\nThe accuracy of the power estimates based on the more refined\nmethods, using uncorrected data, as were used in the construction of\nfigure 11, is within $\\pm 1$ percent of the observed data over most of\nthe speed range and within $\\pm 3$ percent at the highest speeds. Turbine-\nover-all-efficiency estimates by this method are within $\\pm 1$ point\nover most of the speed range and within $\\pm 1 \\frac{1}{2}$ points at the lower and\nhigher ends of the speed range.\n\nEstimates of power output for $180^\\circ$ admission by the method\nbased on use of corrected data are approximately 1 percent lower\nthan the estimations that use the more refined method based on\nobserved data. The efficiency estimated for $180^\\circ$ by this rapid\nmethod is approximately 1 point lower than the estimation obtained", "timestamp": "2026-07-22T06:13:18.844805+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 44, "total_pages": 114, "image_filename": "19930086061_p44.jpg", "text": "```markdown\n$\\alpha=34.1^\\circ$\n$C_L=1.11$\n\n$\\alpha=39.1^\\circ$\n$C_L=1.04$\n\n$\\alpha=44.1^\\circ$\n$C_L=0.75$\n\n$\\alpha=48.1^\\circ$\n$C_L=0.67$\n\n40\n\nP\n-4\n-3\n-2\n-1\n0\n1\n\nStation 1\n$\\frac{y}{b/2}, 0$\n\n[Figure: Triangle symbol]\n\nP\n-4\n-3\n-2\n-1\n0\n1\n\nUpper\nLower\n\nStation 2\n$\\frac{y}{b/2}, .0167$\n\nP\n-3\n-2\n-1\n0\n1\n\nStation 3\n$\\frac{y}{b/2}, .0333$\n\n0 2 4 6 8 10\nx/c\n\n0 2 4 6 8 10\nx/c\n\n0 2 4 6 8 10\nx/c\n\n0 .2 .4 .6 .8 1.0\nx/c\n\nNACA\n\n(a) Stations 1, 2, 3.\n\nFigure 9.- Chordwise pressure distribution about wing 2 at angles of attack of $34.1^\\circ$, $39.1^\\circ$, $44.1^\\circ$, and $48.1^\\circ$.\n\nNACA RM L9J07\n```", "timestamp": "2026-07-22T06:13:21.815727+00:00"}
{"citation_id": "19930085572", "source_url": "https://ntrs.nasa.gov/api/citations/19930085572/downloads/19930085572.pdf", "page_number": 8, "total_pages": 17, "image_filename": "19930085572_p8.jpg", "text": "```markdown\n6\nNACA RM No. E5L02\n\nTABLE I - SPECIFICATIONS AND ANALYSIS OF FUELS USED\n\n<!-- Table (104, 168, 870, 803) -->\n\\begin{tabular}{|l|c|c|c|c|}\n\\hline\n\\multicolumn{1}{|c|}{} & \\multicolumn{2}{c|}{Specification} & \\multicolumn{2}{c|}{Analysis} \\\\\n\\cline{2-5}\n\\multicolumn{1}{|c|}{NACA fuel} & AN-F-58 & AN-F-32 & AN-F-58 & AN-F-32 \\\\\n\\cline{4-5}\n\\multicolumn{1}{|c|}{} & & & 48-210 & 48-306 \\\\\n\\hline\nA.S.T.M. distillation D 86-46, $^\\circ$F & & & & \\\\\n\\quad Initial boiling point & --------- & --------- & 102 & 336 \\\\\n\\quad 5 (evaporated) & --------- & --------- & 149 & 349 \\\\\n\\quad 10 & --------- & 410(max.) & 174 & 355 \\\\\n\\quad 20 & --------- & --------- & 234 & 360 \\\\\n\\quad 30 & --------- & --------- & 286 & 365 \\\\\n\\quad 40 & --------- & --------- & 322 & 370 \\\\\n\\quad 50 & --------- & --------- & 360 & 375 \\\\\n\\quad 60 & --------- & --------- & 390 & 381 \\\\\n\\quad 70 & --------- & --------- & 412 & 387 \\\\\n\\quad 80 & --------- & --------- & 444 & 394 \\\\\n\\quad 90 & 425(min.) & 490(max.) & 480 & 405 \\\\\n\\quad Final boiling point & 600(max.) & 572(max.) & 545 & 446 \\\\\n\\quad Residue, (percent) & 1.5(max.) & 1.5(max.) & 0.8 & 1.0 \\\\\n\\quad Loss, (percent) & 1.5(max.) & 1.5(max.) & 0.2 & 1.0 \\\\\n\\hline\nFreezing point, $^\\circ$F & -76(max.) & -76(max.) & $<$-76 & --------- \\\\\nAromatics, (percent by volume) & & & & \\\\\n\\quad A.S.T.M. D-875-46T & 30(max.) & 20(max.) & 23 & --------- \\\\\n\\quad Silica gel$^a$ & --------- & --------- & 29 & 15 \\\\\nAccelerated gum (mg/100 ml) & 20(max.) & 8.0(max.) & --------- & 0 \\\\\nAir jet residue (mg/100 ml) & 10(max.) & 5.0(max.) & 20 & 1 \\\\\nSulfur, (percent by weight) & 0.5(max.) & 0.20(max.) & 0.09 & 0.02 \\\\\nViscosity, (centistokes at & & & & \\\\\n\\quad -40$^\\circ$ F) & 10.0(max.) & 10.0(max.) & 4.26 & --------- \\\\\nBromine number & 14.0(max.) & 3.0(max.) & 12.0 & --------- \\\\\nReid vapor pressure & & & & \\\\\n\\quad (lb/sq in.) & 5-7(max.) & --------- & 5.7 & --------- \\\\\nHydrogen-carbon ratio & --------- & --------- & 0.153 & 0.154 \\\\\nHeat of combustion (Btu/lb) & 18,200 & --------- & 18,475 & 18,530 \\\\\n\\quad & (min.) & & & \\\\\nSpecific gravity & --------- & 0.950(max.) & 0.794 & 0.831 \\\\\nFlash point, $^\\circ$F & --------- & 110(min.) & --------- & --------- \\\\\n\\hline\n\\end{tabular}\n\n$^a$Reference 2.\n\n1061\nNACA\n```", "timestamp": "2026-07-22T06:13:29.939596+00:00"}
{"citation_id": "19930085485", "source_url": "https://ntrs.nasa.gov/api/citations/19930085485/downloads/19930085485.pdf", "page_number": 16, "total_pages": 26, "image_filename": "19930085485_p16.jpg", "text": "```markdown\n14\n\nLocal Mach number, $M_l$\n1.4\n1.2\n1.0\n.8\n.6\n.4\n\n$M_T$\n.785\n.75\n.725\n.70\n.675\n.65\n.6\n.5\n.4\n\nCONFIDENTIAL\n\n$M_T$\n.4\n.5\n.6\n.65\n.70\n.725\n.75\n.775\n\nVertical distance, inches\n8\n6\n4\n2\n0\n\nLocal Mach number, $M_l$, at\n12-inch chordwise position\n.4\n.6\n.8\n1.0\n1.2\n\nChordwise distance, inches, from L.E. of bump\n10\n12\n14\n16\n18\n20\n\nBump surface\n90°\n\nNACA RM No. L5K02\n\nFigure 5.- Spanwise and chordwise distribution of Mach number over transonic bump. Model off.\nCONFIDENTIAL\n```", "timestamp": "2026-07-22T06:13:35.692245+00:00"}
{"citation_id": "19930082613", "source_url": "https://ntrs.nasa.gov/api/citations/19930082613/downloads/19930082613.pdf", "page_number": 20, "total_pages": 46, "image_filename": "19930082613_p20.jpg", "text": "NACA TN 1938\n19\n\n[Figure: (a) Most common type of crack. Labels: 1, 2, 3, Upper bend in louver flap, Lower bend in louver flap, Stress-relieving hole, Louver flap]\n\n(a) Most common type of crack.\n\n[Figure: (b) Second most common type of crack.]\n\n(b) Second most common type of crack.\nNACA\nC-22310\n9-29-48\n\n[Figure: (c) Large buckle and typical cracks enlarged from figure 4. Labels: Buckle, Uncommon crack]\n\n(c) Large buckle and typical cracks enlarged from figure 4.\nFigure 3. - Typical cracks and buckle at louver.\nNACA\nC-20562\n1-29-48", "timestamp": "2026-07-22T06:13:43.171250+00:00"}
{"citation_id": "19930083221", "source_url": "https://ntrs.nasa.gov/api/citations/19930083221/downloads/19930083221.pdf", "page_number": 10, "total_pages": 47, "image_filename": "19930083221_p10.jpg", "text": "```markdown\n8\nNACA TN No. 1824\n\nThe change in load distribution brought about by this maneuver is the so-called indicial load distribution and a knowledge of such indicial functions is important for applications using operational calculus.\n\n--- characteristic traces\n\n<!-- Image (173, 186, 466, 516) -->\n\n(a) Supersonic wing.\n\n(b) Subsonic wing.\n\nFigure 1.- Boundary conditions for two-dimensional unsteady-lift problem.\n\nFigures 1(a) and 1(b) furnish an insight into the nature of the boundary conditions for airfoils traveling, respectively, at supersonic and subsonic speeds. The chord is initially on the x axis with leading edge at the origin and trailing edge at $x = c_0$. With increasing time, the wing section travels in the negative x direction and sweeps out a portion of the xt plane as shown in the figures (indicated by shaded areas). Throughout this part of the plane the boundary conditions require that the induced vertical velocity is $-V_0\\alpha$, while elsewhere on the plane the induced velocities are continuous functions of z. In figure 2, sketches of the airfoil in the supersonic and subsonic cases are shown together with indications of the manner in which the disturbance field spreads. These wings are presented in xyz space with time as a parameter so that their coordinate system is not to be confused with that of figure 1. The airfoils are traveling from right to left at Mach numbers of 0.8 in the subsonic case and 1.2 in the supersonic case, and for a time corresponding to that required for the wing to travel a distance of one-third chord length. At $t = 0$ cylindrical waves are induced at each disturbance point, that is, at each point of the chord. These waves expand radially at the center of the expanding waves moves relative to the initial disturbance point. At a given instant in time the entire disturbance region\n```", "timestamp": "2026-07-22T06:13:43.761501+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 9, "total_pages": 149, "image_filename": "19930083192_p9.jpg", "text": "NACA TN 1976\n\nare to be compared. For load-evaluation studies of operating conditions, the gross weight is used because it yields conservative values of the effective gust velocity. In all cases, the acceleration increment is defined as the acceleration minus one. The other quantities have been standardized as the result of instrumental limitations so that sea-level density is used with indicated airspeed (assumed to be equal to $V_e$) and gross wing area. The slope of the lift curve is generally obtained from the formula\n\n$$\n\\left(\\frac{dC_L}{d\\alpha}\\right)_{\\text{airplane}} = \\left(\\frac{dC_L}{d\\alpha}\\right)_{\\text{wing}} = \\frac{6A}{A + 2}\n$$\n\nFor preliminary evaluation of acceleration data, sometimes a lift-curve slope must be assumed since the wing configuration is not known. In these cases, a value of 4.5 per radian has usually been selected.\n\nIn the case where loads are to be calculated by utilizing effective gust velocities in the sharp-edge-gust formula, the quantities used must be consistent with those described. If different procedures are followed the answers obtained obviously are incorrect in proportion to the deviation. The simple formula with its simplifying assumptions limits the operations that can be performed on derived data.\n\nEXTENDED GUST EQUATIONS\n\nThe restrictive assumptions of the sharp-edge-gust formula caused little concern until it was to be used for design calculations of gliders and airplanes whose wing loading and other characteristics differed widely from those of the airplanes that were initially used to establish the effective design gust velocities. The assumptions of infinite gust gradient, steady lift, and no vertical motion had to be eliminated in the derivation of the equations for airplane response to a gust. The basis for the final solutions was the work of Küssner reported in reference 2, which resulted in the analytical study presented in reference 4.\n\nFor the application of Küssner's work to the problem of gust loads, the assumptions already cited in the derivation of the sharp-edge-gust formula have been continued with the following exceptions: (1) The gust velocity has been assumed to be uniform across the span of the airplane at any instant and to increase linearly with distance in the direction of flight until the maximum gust velocity $U$ is attained; (2) the airplane can rise but does not pitch under the action of the gust and is in steady level flight; (3) the unsteady-lift functions of Küssner and", "timestamp": "2026-07-22T06:13:44.562130+00:00"}
{"citation_id": "19930082566", "source_url": "https://ntrs.nasa.gov/api/citations/19930082566/downloads/19930082566.pdf", "page_number": 24, "total_pages": 44, "image_filename": "19930082566_p24.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:13:45.068273+00:00"}
{"citation_id": "19930085548", "source_url": "https://ntrs.nasa.gov/api/citations/19930085548/downloads/19930085548.pdf", "page_number": 8, "total_pages": 46, "image_filename": "19930085548_p8.jpg", "text": "NACA RM No. E8L30\n\nwhere\n\n$R_{S}$ scavenging ratio\n\n$p_{e}$ exhaust-gas pressure, (lb/sq in. abs.)\n\n$p_{m}$ inlet-manifold pressure, (lb/sq in. abs.)\n\n$T_{m}$ inlet-manifold temperature, ($^{\\circ}$R)\n\nThis equation was intended to apply to cruising engine speeds corresponding to a mean piston speed of approximately 1800 feet per minute. The experimental cylinder, however, was operated at a mean piston speed of 1350 feet per minute. If the cylinder and the ports are assumed to have the characteristics of an orifice and the weight flow is assumed directly proportional to the average piston speed, the cylinder pressure drop will vary as the square of the average piston speed. Consequently, equation (1) becomes\n\n$$\np_{m} - p_{e} = \\left(\\frac{R_{S}}{0.0910}\\right)^{2} \\left(\\frac{1350}{1800}\\right)^{2} \\frac{p_{m}}{T_{m}}\n$$\n\nPressure drops calculated from equation (2) are compared with the experimental data for a fuel-air ratio of 0.03 in figure 9. The differences between the calculated and experimental data are in a large part attributed to the low inlet-port flow coefficients and the inadequate exhaust lead of the cylinder under investigation. Because the analytical expression is independent of fuel-air ratio, the pronounced effect of fuel-air ratio on the pressure drop across the experimental cylinder (as illustrated in fig. 8) is further evidence of inadequate exhaust lead or time-area for the exhaust blowdown process. The required exhaust lead, determined by a method presented in reference 6, was about six times that of the experimental cylinder.\n\nCharging efficiency. - In reference 1 it is assumed that perfect mixing accompanies the charging process (reference 6). For this scavenging process, the charging efficiency $\\eta_{S}$ is given by reference 7 as\n\n$$\n\\eta_{S} = 1 - e^{-R_{S}}\n$$", "timestamp": "2026-07-22T06:13:46.257689+00:00"}
{"citation_id": "19930082592", "source_url": "https://ntrs.nasa.gov/api/citations/19930082592/downloads/19930082592.pdf", "page_number": 23, "total_pages": 50, "image_filename": "19930082592_p23.jpg", "text": "22\nNACA TN 1914\n\nTemperature\n($^\\circ$F)\n1625\n1755\n2000\n\nTungsten\n(percent)\n20\n5\n10\n30\n30\n30\n20\n10\n5\n20\n5\n10\n\nOxidation penetration, P, in.\n.014\n.012\n.010\n.008\n.006\n.004\n.002\n0\n\nTime, t, hr\n1\n2\n3\n4\n6\n8\n10\n20\n30\n40\n60\n80\n100\n200\n\n[Figure: Graph showing oxidation penetration vs. time for various temperatures and tungsten percentages]\n\n(d) Composite of figures 4(a), 4(b), and 4(c).\nFigure 4. - Concluded. Effect of time, temperature, and tungsten content on oxidation penetration of titanium carbide - tungsten ceramals.", "timestamp": "2026-07-22T06:13:47.737293+00:00"}
{"citation_id": "19930082918", "source_url": "https://ntrs.nasa.gov/api/citations/19930082918/downloads/19930082918.pdf", "page_number": 12, "total_pages": 62, "image_filename": "19930082918_p12.jpg", "text": "NACA TN 1940\n\nin diameter, terminated at the ends in 1/4-inch radii to 3/8-inch threads, were machined in these quarter sections.\n\nTests at 60,000 psi were simply the creep tests carried to rupture. Tests at stresses less than 60,000 psi were run in conventional beam-loaded rupture units with one-piece electric furnaces and with temperature control and uniformity comparable with those of the other rupture and creep tests. Specimen form was the same as that used for the rupture tests in the hydraulic tensile machine.\n\nRESULTS\n\nMetallographic Examination\n\nFigures 5 to 9 show the micrographs taken of the aged samples and the following description summarizes the results:\n\n(1) Aging at $1200^\\circ$ F resulted in little but the progressive development of a distinct grain boundary constituent which resisted etching. At aging periods up to 10 hours, the boundary constituent was incomplete in that it did not surround all the individual grains. At aging periods of 1000 hours, the boundary constituents completely surrounded the grains and had become an approximately 0.5-micron-wide band. After aging 1000 hours slight precipitation was observable in the matrix near the grain boundaries. At 10,000 diameters, this precipitate did not appear to be a distinct phase with an interface but was surrounded by a concentration gradient as revealed by a sloping surface from the center of the precipitate particles to the matrix proper as a result of etching. As postulated by Nabarro (see reference 5) this could indicate that appreciable strains would exist around each such particle because of the probable difference in equilibrium lattice spacing between matrix and precipitate.\n\n(2) Aging at $1400^\\circ$ F resulted first in the progressive development of a grain boundary phase. Initially only a concentration gradient was present at the boundary, as revealed by a sloping surface toward the grain boundary after etching. At 10 hours the first separate boundary-phase particles appeared. At 100 hours the boundary band was approximately 0.8 micron wide and changed little in character with further aging. In addition to the grain-boundary reaction, general matrix precipitation appeared after aging about 10 hours, first along the grain boundaries, and increased rapidly in number and size up to the longest aging period used, 1000 hours. The particles at first had concentration gradients surrounding them; however, for aging times of approximately over 100 hours, no appreciable concentration gradient appeared around the precipitate particles but rather a definite interface was present. Average size of the particles at 1000 hours was estimated visually to be 0.2", "timestamp": "2026-07-22T06:13:49.597307+00:00"}
{"citation_id": "19930082646", "source_url": "https://ntrs.nasa.gov/api/citations/19930082646/downloads/19930082646.pdf", "page_number": 10, "total_pages": 37, "image_filename": "19930082646_p10.jpg", "text": "NACA TN 1980\n9\n\nTABLE I\n\nDATA OBTAINED DURING LANDINGS IN WAVES OF LANGLEY TANK MODEL 20-4\n\n[All values are model size;\nwave height = 0.4 foot for all landings]\n\n| Landing | Wave length (ft) | Initial impact | | | | | | Maximum acceleration | | | | | |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| | | $\\tau_L$ (deg) | $V_v$ (fps) | V (fps) | $\\gamma$ (deg) | $n_v$ (g) | $\\tau$ (radians/sec) | Impact | $\\tau$ (deg) | $V_v$ (fps) | V (fps) | $\\gamma$ (deg) | $n_v$ (g) | $\\tau$ (radians/sec$^2$) |\n| 1 | 13.4 | 7.4 | 1.08 | 35.6 | 1.7 | 2.6 | 9 | 4 | 5.0 | 2.10 | 30.6 | 3.9 | 2.9 | 21 |\n| 2 | 13.4 | 7.4 | 1.04 | 35.1 | 1.7 | 1.5 | 5 | 4 | 6.5 | 1.52 | 28.8 | 3.2 | 3.0 | 18 |\n| 3 | 13.7 | 7.4 | 1.13 | 36.0 | 2.0 | 1.8 | 10 | 7 | 5.0 | 2.10 | 31.0 | 3.8 | 2.8 | 22 |\n| 4 | 14.0 | 7.4 | 1.05 | 35.0 | 1.7 | 1.7 | 10 | 4 | 4.8 | 2.01 | 31.2 | 4.6 | 4.0 | 28 |\n| 5 | 14.2 | 7.4 | 1.04 | 35.0 | 1.7 | 2.2 | 11 | 4 | 4.3 | 2.01 | 27.4 | 4.1 | 3.7 | 31 |\n| 6 | 14.7 | 7.4 | 1.30 | 36.0 | 2.2 | 2.1 | 16 | 2 | 4.8 | 2.10 | 31.8 | 3.8 | 3.4 | 32 |\n| 7 | 14.5 | 7.4 | 1.17 | 36.1 | 1.9 | 1.5 | 0 | 4 | 4.3 | 2.06 | 28.7 | 4.1 | 3.7 | 31 |\n| 8 | 15.6 | 7.3 | 1.43 | 34.9 | 2.1 | .9 | 0 | 4 | 3.9 | 2.52 | 28.5 | 3.3 | 2.9 | 40 |\n| 9 | 15.8 | 7.4 | 1.30 | 36.0 | 2.3 | .11 | 3 | 7 | 4.4 | 2.08 | 31.3 | 4.1 | 3.4 | 32 |\n| 10 | 16.0 | 7.4 | 1.53 | 36.1 | 2.4 | .9 | 3 | 4 | 4.4 | 2.77 | 29.7 | 5.3 | 3.3 | 32 |\n| 11 | 16.1 | 7.7 | 1.10 | 34.0 | 2.3 | 2.9 | 22 | 4 | 4.6 | 1.71 | 33.6 | 4.7 | 2.9 | 22 |\n| | | | | | | | | a3 | 2.8 | 3.49 | 28.6 | 7.0 | 2.0 | 41 |\n| 12 | 16.4 | 7.6 | .81 | 37.5 | 1.2 | .9 | 0 | 4 | 3.2 | 2.72 | 30.5 | 4.7 | 2.1 | 31 |\n| | | | | | | | | a4 | 2.9 | 3.65 | 29.2 | 7.1 | 2.0 | 44 |\n| 13 | 16.4 | 7.4 | 1.30 | 36.1 | 2.1 | 1.9 | 9 | 7 | 5.0 | 2.04 | 29.4 | 4.0 | 2.8 | 20 |\n| | | | | | | | | a6 | 2.9 | 2.99 | 27.8 | 4.5 | 2.3 | 25 |\n| 14 | 16.5 | 7.3 | .79 | 37.4 | 1.2 | 1.8 | 10 | 4 | 6.5 | 2.56 | 28.8 | 4.5 | 1.9 | 12 |\n| | | | | | | | | 7 | 6.7 | 2.16 | 27.2 | 4.5 | 1.9 | 16 |\n| 15 | 16.5 | 7.6 | 1.41 | 35.9 | 2.2 | .7 | -10 | (1.4) | 4.7 | 2.23 | 29.0 | 4.5 | 2.6 | 31 |\n| 16 | 16.6 | 7.3 | 1.49 | 36.0 | 2.4 | 2.1 | 16 | 7 | 1.3 | 2.56 | 31.5 | 4.3 | 2.8 | 28 |\n| | | | | | | | | a4 | 4.2 | 2.48 | 30.0 | 5.0 | 2.9 | 37 |\n| 17 | 16.9 | 7.4 | .82 | 37.1 | 1.3 | 1.2 | 0 | a4 | 2.7 | 2.63 | 30.5 | 4.9 | 2.4 | 30 |\n| | | | | | | | | a5 | 2.7 | 2.48 | 29.5 | 4.9 | 2.1 | 41 |\n| 18 | 17.1 | 7.6 | 1.03 | 36.3 | 1.6 | 0 | 0 | a4 | 2.0 | 2.78 | 27.3 | 4.8 | 1.9 | 25 |\n| | | | | | | | | a4 | 2.0 | 2.78 | 27.3 | 4.8 | 1.9 | 25 |\n| 19 | 17.4 | 7.3 | .74 | 38.2 | 1.1 | 0 | 0 | a4 | 1.9 | 3.25 | 32.9 | 5.6 | 2.8 | 30 |\n| 20 | 18.2 | 7.4 | 1.29 | 36.0 | 2.0 | 2.0 | 20 | 3 | 3.1 | 4.21 | 29.0 | 6.2 | 1.9 | 40 |\n| 21 | 18.7 | 7.4 | .89 | 35.5 | 1.4 | 0 | 0 | a4 | --- | 2.47 | 31.7 | 4.5 | 3.2 | 30 |\n| | | | | | | | | a4 | --- | 3.58 | 31.4 | 4.9 | 2.0 | 32 |\n| 22 | 18.9 | 7.7 | 1.09 | 35.2 | 1.8 | .8 | 0 | 7 | 4.6 | 2.61 | 28.7 | 4.2 | 2.3 | 36 |\n| 23 | 19.2 | 7.7 | 1.28 | 35.4 | 2.0 | 1.2 | 0 | 7 | 4.2 | 2.68 | 31.4 | 4.9 | 2.3 | 30 |\n| 24 | 19.4 | 7.4 | 1.06 | 36.0 | 1.7 | 1.0 | 0 | a4 | --- | 2.76 | 24.7 | 4.5 | 2.0 | 9 |\n| | | | | | | | | a4 | --- | 3.29 | 31.2 | 4.9 | 1.6 | 32 |\n| 25 | 19.7 | 7.4 | .81 | 37.0 | 1.2 | 1.4 | 7 | a4 | --- | 1.93 | 30.0 | 4.5 | 2.1 | 12 |\n| | | | | | | | | 7 | 2.9 | 2.48 | 31.4 | 4.9 | 1.1 | 27 |\n| 26 | 20.1 | 7.8 | .99 | 36.1 | 1.5 | 1.6 | 15 | 7 | 2.6 | 2.55 | 31.4 | 4.7 | 3.0 | 55 |\n| 27 | 20.3 | 7.5 | 1.41 | 35.5 | 2.0 | 1.7 | -10 | a4 | 4.9 | 3.39 | 27.5 | 4.7 | 2.6 | 45 |\n| 28 | 21.7 | 7.5 | 1.34 | 36.5 | 2.1 | .8 | 0 | 7 | 5.3 | 1.09 | 36.0 | 1.7 | 2.6 | 19 |\n| | | | | | | | | a3 | 4.9 | 4.39 | 24.7 | 5.7 | 2.6 | 40 |\n| 29 | 22.7 | 8.0 | 1.06 | 36.0 | 1.4 | 1.2 | -10 | 7 | 8.5 | 1.20 | 38.0 | 1.8 | 2.3 | -10 |\n| | | | | | | | | a4 | 4.5 | 2.68 | 28.4 | 4.5 | 2.1 | 29 |\n| 30 | 23.0 | 7.5 | --- | 36.8 | --- | .9 | 0 | a4 | 9.5 | --- | 27.2 | --- | 2.1 | 12 |\n| | | | | | | | | a4 | 9.5 | --- | 26.1 | --- | 1.2 | 20 |\n| 31 | 23.0 | 7.7 | 1.00 | 35.8 | 1.8 | 1.2 | 9 | a4 | 10.0 | 1.30 | 39.0 | 2.3 | 2.2 | 44 |\n| | | | | | | | | 7 | 3.1 | 4.15 | 35.4 | 5.3 | 1.6 | 35 |\n| 32 | 23.1 | 8.0 | 1.40 | 36.0 | 2.2 | 1.6 | 17 | a4 | 11.0 | 3.75 | 29.0 | 4.3 | 2.2 | 19 |\n| | | | | | | | | a4 | 10.8 | 4.15 | 31.4 | 5.3 | 1.7 | 21 |\n| 33 | 23.2 | 7.7 | 1.06 | 36.3 | 1.7 | 0 | -10 | 7 | 8.8 | 4.47 | 29.1 | 5.7 | 2.9 | 21 |\n| 34 | 23.3 | 7.5 | 1.32 | 36.6 | 2.0 | 1.9 | 19 | 7 | 8.6 | 4.47 | 30.1 | 5.7 | 2.9 | 21 |\n| | | | | | | | | a3 | 12.9 | 3.29 | 32.9 | 5.8 | 1.9 | 31 |\n| 35 | 23.4 | 7.3 | 1.48 | 35.8 | 2.4 | 1.3 | 10 | 3 | 12.9 | 4.12 | 30.3 | 5.3 | 2.1 | 29 |\n| 36 | 23.4 | 7.6 | 1.01 | 36.1 | 1.6 | 1.4 | 8 | 3 | 13.1 | 3.30 | 30.0 | 4.1 | 2.0 | 10 |\n| | | | | | | | | a4 | 12.9 | 3.68 | 28.8 | 5.3 | 1.1 | 17 |\n| 37 | 23.6 | 8.5 | .87 | 35.3 | 1.4 | 0 | 0 | 7 | 8.3 | 3.02 | 31.4 | 4.4 | 2.1 | 4 |\n| | | | | | | | | a4 | 8.3 | 3.18 | 31.8 | 4.4 | 2.0 | 30 |\n| 38 | 23.6 | 7.5 | 1.22 | 36.9 | 1.9 | 1.0 | 0 | 7 | 6.6 | 3.84 | 29.0 | 4.5 | 2.1 | 31 |\n| 39 | 24.1 | 7.5 | 1.27 | 35.5 | 1.5 | 1.1 | 0 | 7 | 6.6 | 3.79 | 30.4 | 4.5 | 2.1 | 31 |\n| 40 | 24.8 | 8.4 | 1.38 | 36.8 | 1.3 | 1.8 | 18 | 7 | 12.1 | 3.59 | 29.0 | 5.7 | 2.1 | 29 |\n| 41 | 25.9 | 7.8 | 1.01 | 36.1 | 1.4 | 1.0 | 0 | a4 | 8.0 | 4.54 | 27.2 | 5.7 | 1.9 | 46 |\n| | | | | | | | | a4 | 5.5 | 4", "timestamp": "2026-07-22T06:13:50.128308+00:00"}
{"citation_id": "19930085847", "source_url": "https://ntrs.nasa.gov/api/citations/19930085847/downloads/19930085847.pdf", "page_number": 6, "total_pages": 32, "image_filename": "19930085847_p6.jpg", "text": "4\nCONFIDENTIAL\nNACA RM A9D04\n\ntest panel developed progressively above a Mach number of 0.76.\n\n4. The measurements gave no indication of boundary-layer separation until about 0.79 Mach number, although the Mach number for drag divergence is about 0.73.\n\nThe suction coefficient$^1$ attained with the full-length slot was constant over the Mach number range and equal to about 0.00175. This value of suction coefficient corresponds to removal of about one-quarter of the air in the boundary layer at the slot location at the lower Mach numbers and is about one-third of the suction coefficient it had been hoped to attain at the beginning of the tests. More suction could not be attained due to the limited pressure drop available and the high duct losses present. In an effort to increase the percentage of boundary-layer air removed, the slot length was reduced from each end by 25 percent (reducing the total length to 1 ft). This change resulted, approximately, in a doubling of the suction coefficient due to the 50-percent reduction in affected area, since the volume rate of flow remained about the same. No further attempt to increase the flow rate by additional mechanical means was made in view of the conclusions drawn from a study of the pressure-distribution results presented hereinafter.\n\nAlthough the wake-survey measurements were made to 0.78 Mach number, the results were questionable above a Mach number of 0.76 due to the probable invalidating influence of the boundary-layer cross flow on the boundary layer and wake measurements above that Mach number. Accordingly, since the wake measurements were considered unreliable and since the frictional drag becomes an increasingly smaller percentage of the total drag at supercritical Mach numbers, it was decided to study the effect of suction on the pressure drag as determined from the pressure-distribution measurements. It was realized that a rigorous study of pressure drag would involve the measurement of angle of attack $\\alpha$ and the determination of the drag coefficient $C_d$ by the equation\n\n$$C_d = C_c \\cos \\alpha + C_n \\sin \\alpha$$\n\nwhere $C_c$ and $C_n$ are the chordwise- and normal-force coefficients. In view of the difficulties of measuring angle of attack in flight, however, the chordwise-force coefficient was used to study drag changes due to suction, since at the low angles of attack which existed for these tests the magnitude of the $C_n \\sin \\alpha$ term was small. This\n\n$^1$Suction coefficient is defined as $Q/VA$ where $Q$ is the volume of air removed under the test conditions of temperature and pressure, $V$ is the true airspeed of the airplane, and $A$ is the total wing area ahead and behind the slot.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:13:59.167229+00:00"}
{"citation_id": "19930085859", "source_url": "https://ntrs.nasa.gov/api/citations/19930085859/downloads/19930085859.pdf", "page_number": 3, "total_pages": 31, "image_filename": "19930085859_p3.jpg", "text": "NACA RM No. L9B25\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nRESEARCH MEMORANDUM\n\nAERODYNAMIC CHARACTERISTICS OF A WING WITH QUARTER-CHORD LINE  \nSWEPT BACK $35^\\circ$, ASPECT RATIO 4, TAPER RATIO 0.6,  \nAND NACA 65A006 AIRFOIL SECTION\n\nTRANSONIC-BUMP METHOD\n\nBy William C. Sleeman, Jr. and Robert E. Becht\n\nSUMMARY\n\nAs part of an NACA transonic research program, a series of wing-body combinations are being investigated in the Langley high-speed 7- by 10-foot tunnel over a Mach number range of 0.60 to 1.18 utilizing the transonic bump.\n\nThis paper presents the results of the investigation of a wing-alone and a wing-fuselage combination employing a wing with the quarter-chord line swept back $35^\\circ$, aspect ratio 4, taper ratio 0.6, and an NACA 65A006 airfoil section. Lift, drag, pitching moment, and root bending moment were obtained for the wing-alone and wing-body configurations. Effective downwash angles and dynamic-pressure characteristics in the region of a probable tail location were also obtained for these configurations and are presented for a range of tail heights at one tail length. In order to expedite publishing of these data, only a brief analysis is included.\n\nINTRODUCTION\n\nThe urgent need for aerodynamic design data in the transonic speed range has led to the establishment of a special NACA committee for transonic research. As part of the NACA transonic research program recommended by this committee a series of wing-body configurations having wing plan form as the chief variable are being investigated in the Langley high-speed 7- by 10-foot tunnel utilizing the transonic-bump test technique. For each wing-fuselage combination investigated the lift, drag, pitching-moment, and root bending-moment characteristics are determined over a Mach number range of 0.60 to 1.18. In addition, effective downwash angles and dynamic-pressure characteristics are obtained for a range of tail heights at one tail length.", "timestamp": "2026-07-22T06:14:07.124423+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 45, "total_pages": 114, "image_filename": "19930086061_p45.jpg", "text": "NACA RM L9J07\n\n$\\alpha=34.1^\\circ$\n$C_L=1.11$\n\n$\\alpha=39.1^\\circ$\n$C_L=1.04$\nStation 4\n$\\frac{y}{b/2}, 0.500$\n\n$\\alpha=44.1^\\circ$\n$C_L=0.75$\n\n$\\alpha=48.1^\\circ$\n$C_L=0.67$\n\n-2\n-1\nP\n0\n1\n\n-2\n-1\nP\n0\n1\n\n-2\n-1\nP\n0\n1\n\n-2\n-1\nP\n0\n1\n\nUpper\nLower\n\nStation 5\n$\\frac{y}{b/2}, 0.667$\n\nStation 6\n$\\frac{y}{b/2}, 0.833$\n\nStation 7\n$\\frac{y}{b/2}, 0.916$\n\n0 2 4 6 8 10\n$x/c$\n\n0 2 4 6 8 10\n$x/c$ (b) Stations 4,5,6,7.\n\n0 2 4 6 8 10\n$x/c$\n\n0 2 4 6 8 10\n$x/c$\n\nNACA\n\nFigure 9.- Concluded.\n\n41", "timestamp": "2026-07-22T06:14:08.351024+00:00"}
{"citation_id": "19930082613", "source_url": "https://ntrs.nasa.gov/api/citations/19930082613/downloads/19930082613.pdf", "page_number": 21, "total_pages": 46, "image_filename": "19930082613_p21.jpg", "text": "Page intentionally left blank\n\nPage intentionally left blank", "timestamp": "2026-07-22T06:14:09.364344+00:00"}
{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 10, "total_pages": 72, "image_filename": "19930085491_p10.jpg", "text": "NACA RM No. A8J04 CONFIDENTIAL 9\n\nat all sweepback settings for zero lift and at selected lift coefficients. The technique, adapted for supersonic flow from a method developed by Grey (reference 8) depends primarily upon the difference in the rates of evaporation within laminar and turbulent flow areas. In addition to indicating areas of laminar and turbulent flow as in reference 5, where the patterns obtained were photographed outside of the tunnel after the test was completed, the photographs obtained in the present tests while the tunnel was operating indicate the location of the line of laminar separation and in some instances the direction of the boundary-layer flow.\n\nSchlieren plan-form photographs were taken during all tests in which liquid-film patterns were recorded. It was determined during a specific test that the presence of the liquid film and fluid ridges therein did not alter the shock-wave pattern or aerodynamic forces on the model.\n\nCorrections and Precision\n\nThe effect of support interference was taken into account in the manner described in reference 5. Liquid-film studies showed that the boundary layer was turbulent over the rear of the fuselage and remained unseparated up to the fuselage base. Reference 9 indicates that the effect of the support system on the pressures on the model will then be confined to the base pressure. The base pressure was measured in each test and the drag force was corrected for the difference between the test and the static pressure of the free stream at the fuselage base. Drag corrections for the longitudinal pressure gradient of the stream were calculated for the wings and fuselage and were found to be negligible.\n\nAll experimental lift curves have been plotted against the angle of attack of the root chord and no attempt has been made to determine the average angle of attack due to wing twist under load. The variation of the wing twist with angle of attack was found to be approximately linear for the basic configuration WF-63 and corresponded to $0.3^\\circ$ washout at the optimum lift coefficient.\n\nThe accuracy of the experimental data is the same as that determined in reference 5 since the experimental technique and equipment were essentially the same. However, the sting moment gage was changed prior to the present investigation and the improvement in construction eliminated the discrepancies noted near zero lift in the previous tests. More experimental points were also obtained in the present investigation to permit more accurate fairing of the data.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:14:09.547916+00:00"}
{"citation_id": "19930085485", "source_url": "https://ntrs.nasa.gov/api/citations/19930085485/downloads/19930085485.pdf", "page_number": 17, "total_pages": 26, "image_filename": "19930085485_p17.jpg", "text": "NACA RM No. L8K02\n15\n\nCONFIDENTIAL\n\n20.1\nTunnel wall\nTo strain gage balance\n30\n1/6 1/8\nAll dimensions\nin inches\n\n16.0\n14.0\n10.0\nFairing\nReflection plane\nNACA\n\nFigure 6.- Drawing of reflection-plane mounting of model on vertical\nwall of the Langley high-speed 7- by 10-foot tunnel.\nCONFIDENTIAL", "timestamp": "2026-07-22T06:14:10.152929+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 34, "total_pages": 78, "image_filename": "19930082483_p34.jpg", "text": "32\nNACA TN No. 1807\n\nfrom data uncorrected to sea level. Efficiency estimations performed this way will always be lower than those shown in figure 11, becoming 3 points lower at 90° admission.\n\nRESULTS\n\nLosses\n\nWith the exception of disk-windage loss, a quantitative statement of the types of turbine loss that have been considered is presented in figure 9. These quantities are based on data obtained at an inlet pressure of 45 inches of mercury absolute and a total-pressure ratio of 2.0. Because the losses are expressed in corrected form, they may be considered representative of the losses at other inlet pressures, although all the losses except bearing losses would require adjustment for changes in pressure ratio.\n\nPresented in figure 9 are the losses as functions of corrected rotor speed for a range of admissions from 90° to 360°, where applicable. Use of logarithmic scales for abscissas and ordinates results in the slopes of these loss curves being equal to the exponential rate of rotor speed at which the losses vary.\n\nRotor-tip leakage loss. - Estimates of rotor-tip leakage losses for a range of gas admissions are plotted in figure 9(a). Because the tip leakage is invariant with rotor speed, the curves appear as horizontal lines over the range of corrected rotor speeds.\n\nDisk-windage loss. - In general, disk-windage loss forms a small fraction of the dissipated power output, which is exemplified by the results of these studies where the maximum value of the corrected disk loss amounted to less than 1 horsepower at the highest corrected operating speed of 14,000 rpm. Because of the minor nature of this loss, it does not appear in figure 9. However, were the disk vaned, as is the case for some commercial gas turbines, the additional pumping action would act to increase greatly the disk power loss.\n\nBearing loss. - Bearing loss is plotted against corrected rotor speed in figure 9(b). The single bearing-loss curve is considered applicable for any degree of admission. The temperature rise of the oil across the journal-thrust bearing did not decrease for the runs with partial admission indicating little change in the thrust-bearing losses. At the higher speeds, the bearing-loss curve approaches a 1.4-power variation with the speed. That this loss is", "timestamp": "2026-07-22T06:14:12.837514+00:00"}
{"citation_id": "19930082703", "source_url": "https://ntrs.nasa.gov/api/citations/19930082703/downloads/19930082703.pdf", "page_number": 16, "total_pages": 28, "image_filename": "19930082703_p16.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:14:14.213070+00:00"}
{"citation_id": "19930085572", "source_url": "https://ntrs.nasa.gov/api/citations/19930085572/downloads/19930085572.pdf", "page_number": 9, "total_pages": 17, "image_filename": "19930085572_p9.jpg", "text": "NACA RM No. E5L02\n7\n\nTABLE II - WINDMILLING STARTING AND ACCELERATION DATA\n[Flight Mach number, 0.37]\n\n| Alti- tude (ft) | Wind- milling speed (rpm) | Free-air tempera- ture ($^\\circ$F) | Success- ful start and accel- eration | Maximum tail-pipe tempera- ture ($^\\circ$F) | Acceleration | |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| | | | | | Time (sec) | Engine speed (rpm) |\n| **AN-F-32 Fuel** | | | | | | |\n| 5,000 | 1320 | 50 | Yes | 1150 | 36 | 6930 |\n| 10,000 | 1280 | 47 | ---do.--- | 1140 | 61 | 6970 |\n| 20,000 | 1300 | 15 | ---do.--- | 1300 | 68 | 7000 |\n| 28,000 | 1130 | -22 | ---do.--- | 1310 | 106 | 6930 |\n| **AN-F-58 Fuel** | | | | | | |\n| 6,000 | 1300 | 44 | ---do.--- | 950 | 91 | 6780 |\n| 10,000 | 1270 | 33 | ---do.--- | 810 | 74 | $^a$4940 |\n| 20,000 | 1160 | 2 | ---do.--- | 1400 | 80 | 6990 |\n| 30,000 | 1040 | -38 | ---do.--- | 1380 | 153 | 6680 |\n\n$^a$No data obtained above 4940 rpm.\n\n[Figure: NACA logo]", "timestamp": "2026-07-22T06:14:17.265887+00:00"}
{"citation_id": "19930083221", "source_url": "https://ntrs.nasa.gov/api/citations/19930083221/downloads/19930083221.pdf", "page_number": 11, "total_pages": 47, "image_filename": "19930083221_p11.jpg", "text": "NACA TN No. 1824\n9\n\nof the airfoil is contained within the closed surfaces shown in the\nfigures, the outer surfaces corresponding to the largest values of\n\n[Figure: Sketch showing two wing sections with disturbance fields. The top section is labeled \"Subsonic\" and the bottom section is labeled \"Supersonic\".]\n\nFigure 2.- Sketch showing extent of disturbance fields after\ntravel of one-third chord length.\n\ntime. In the supersonic case, the pressure distribution over the\nwing reaches a steady-state value as soon as the wing moves ahead of\nthe expanding cylindrical wave produced at $t = 0$ by the leading\nedge. In the subsonic case, the wing never leaves the disturbance\nfield of the cylinders and, as will be seen later, the steady-state\npressure distribution is approached asymptotically.\n\nIt is apparent from equation (17) that the characteristic cones\nhave semivertex angles equal to $45^\\circ$ and that the cones with vertices\non the $xt$ plane have traces with slopes equal to $\\pm 1$. These cones\ndetermine the upstream boundary of the field of influence of the\nvertex point and their cross sections in the plane $t = \\text{constant}$\nare the disturbance regions of the cylindrical waves arising at the\nvertex. Thus, perturbations in pressure produced initially at the\nleading edge of the wing section are confined at later time, in the\nmodified coordinate system, to the cone with vertex at the origin\nand traces $x = \\pm t$.", "timestamp": "2026-07-22T06:14:18.290563+00:00"}
{"citation_id": "19930082566", "source_url": "https://ntrs.nasa.gov/api/citations/19930082566/downloads/19930082566.pdf", "page_number": 25, "total_pages": 44, "image_filename": "19930082566_p25.jpg", "text": "NACA TN No. 1889\n23\n\n[Figure: A black and white photograph of a Bosch pump unit for fluctuating internal pressures. The unit is a complex mechanical assembly with various components labeled with letters such as I, C, T, E2, and U.]\n\nFigure 8.- Bosch pump unit for fluctuating internal pressures.", "timestamp": "2026-07-22T06:14:21.119924+00:00"}
{"citation_id": "19930093769", "source_url": "https://ntrs.nasa.gov/api/citations/19930093769/downloads/19930093769.pdf", "page_number": 38, "total_pages": 39, "image_filename": "19930093769_p38.jpg", "text": "1070\n\nCONFIDENTIAL\n\nNACA RM No. E8L10a\n\nCONFIDENTIAL\n\n[Figure: Combustor basket looking upstream from turbine showing carbon deposits on two rings.]\n\nNACA\nC-21970\n8-4-48\n\nFigure 11. - Combustor basket looking upstream from turbine showing carbon deposits on two rings.\n\n37", "timestamp": "2026-07-22T06:14:23.759967+00:00"}
{"citation_id": "19930082592", "source_url": "https://ntrs.nasa.gov/api/citations/19930082592/downloads/19930082592.pdf", "page_number": 24, "total_pages": 50, "image_filename": "19930082592_p24.jpg", "text": "NACA TN 1914\n23\n\nTungsten oxidation-rate constant, $\\Delta p/\\Delta \\log t = K_W$\n\nTungsten\n(percent)\n$\\circ$ 5\n$\\square$ 10\n$\\diamond$ 20\n$\\triangle$ 30\n\nReciprocal of absolute temperature in $^\\circ R$, $1/T$\nTemperature, $^\\circ F$\n\n[Figure: Graph showing the effect of temperature and tungsten content on oxidation-rate constant of titanium carbide - tungsten cermals.]\n\nFigure 5. - Effect of temperature and tungsten content on oxidation-rate constant of titanium carbide - tungsten cermals.", "timestamp": "2026-07-22T06:14:26.785010+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 10, "total_pages": 149, "image_filename": "19930083192_p10.jpg", "text": "6\nNACA TN 1976\n\nWagner (reference 2), which were derived for the two-dimensional wing, are assumed to be made applicable to the finite wing by substituting the slope of the lift curve for the finite wing in place of the slope of the lift curve for the infinite-aspect-ratio or two-dimensional wing; and (4) except for the shorter gust-gradient distances, the acceleration peak was assumed to coincide in space with the position of the point of first attainment of maximum gust velocity.\n\nThe symbols used in the equation for the gradient gust are shown in figure 2, and the unsteady-lift functions used are shown in figure 3, taken from reference 5. It is indicated that, while a sudden angle-of-attack change, such as is obtained by rotating or pitching a wing, yields one variation of lift $C_{L_\\alpha}$, the gradual immersion of the wing into a gust yields a different curve $C_{L_g}$.\n\nThe equation of vertical motion can be written as\n\n$$\n\\frac{W}{g} \\left( \\frac{d^2z}{dt^2} \\right)_1 = \\int_0^{t_1} \\frac{\\rho S V^2}{2} \\frac{dC_L}{d\\alpha} C_{L_g}(t_1 - t) \\frac{u}{V} dt - \\int_0^{t_1} \\frac{\\rho S V^2}{2} \\frac{dC_L}{d\\alpha} C_{L_\\alpha}(t_1 - t) \\frac{d^2z}{dt^2} \\frac{dt}{V}\n$$\n\nwhere the first integral is the force due to the gust and the second integral results from the vertical motion of the airplane. Since the unsteady-lift functions $C_{L_\\alpha}$ and $C_{L_g}$ depend on chord lengths from the start of a disturbance, the equation is generally transformed so that the distance traveled $s = \\frac{V}{c} t$ rather than $t$ is the independent variable. The equation becomes\n\n$$\n\\left( \\frac{\\Delta n}{\\Delta n_s} \\right)_1 = \\frac{\\bar{c}}{H} \\int_0^{s_1} C_{L_g}(s_1 - s) ds - \\frac{1}{\\mu_g} \\int_0^{s_1} C_{L_\\alpha}(s_1 - s) \\frac{\\Delta n(s)}{\\Delta n_s} ds\n$$", "timestamp": "2026-07-22T06:14:28.048101+00:00"}
{"citation_id": "19930082646", "source_url": "https://ntrs.nasa.gov/api/citations/19930082646/downloads/19930082646.pdf", "page_number": 11, "total_pages": 37, "image_filename": "19930082646_p11.jpg", "text": "```markdown\n10\nNACA TN 1980\n\n<!-- Image (108, 109, 812, 819) -->\n\nFigure 1.- General arrangement.\n```", "timestamp": "2026-07-22T06:14:28.144661+00:00"}
{"citation_id": "19930085548", "source_url": "https://ntrs.nasa.gov/api/citations/19930085548/downloads/19930085548.pdf", "page_number": 9, "total_pages": 46, "image_filename": "19930085548_p9.jpg", "text": "8\nNACA RM No. E8L30\n\nAt the selected operating condition, that is, a scavenging ratio of unity, the charging efficiency from equation (3) is 63.2 percent. Samples of gas from the experimental cylinder, however, indicated that the charging efficiency was less than 55 percent, which represents a loss of 13 percent in power output at rich mixtures.\n\nFor operation at over-all fuel-air ratios in excess of 0.035, experimental data indicate that as the fuel-air ratio is enriched the charging efficiency decreases and ultimately approaches zero. This relation is caused by the cylinder fuel-air ratio being stoichiometric or richer so that all the fuel cannot burn. When the scavenging process begins, the unburned fuel remaining from the previous cycle apparently undergoes combustion, using part of the scavenge air. As a result, the cylinder is in part being scavenged with products of combustion. On succeeding cycles, the cylinder fuel-air ratio becomes increasingly richer until equilibrium is reached, at which time the concentration of products of combustion in the cylinder is very high at inlet-port closure.\n\nThe effect of high concentration of the products of combustion is demonstrated by the data of figure 10. The indicated mean effective pressure decreases rather rapidly as the cylinder fuel-air ratio goes beyond stoichiometric, corresponding to an over-all fuel-air ratio of about 0.035, which is indicative of the poor scavenging under these operating conditions. The burning of practically all the fuel during some part of the cycle, whether it is during the power stroke or during the initial stages of the scavenging process, is clearly illustrated by the curve of exhaust-gas temperature, which continues to increase with increasing over-all fuel-air ratio.\n\nOperation on a four-stroke cycle (with fuel injection every other cycle) should demonstrate the effect of the unburned fuel that exists in the cylinder at the beginning of scavenging. In this case, the unburned fuel and products of combustion resulting from burning during the scavenging period are carried out of the cylinder on the nonfiring cycle. The power curve as a function of fuel-air ratio therefore should not peak but should become substantially flat as the cylinder fuel-air ratio becomes richer than stoichiometric. The four-stroke-cycle data, shown only in figure 10, confirm this conclusion.\n\nPower Output\n\nEffect of engine operating conditions. - The effect of inlet-manifold temperature, inlet-manifold pressure, and fuel-air ratio on the indicated mean effective pressure of the experimental cylinder is shown in figures 11 and 12.", "timestamp": "2026-07-22T06:14:29.353115+00:00"}
{"citation_id": "19930082613", "source_url": "https://ntrs.nasa.gov/api/citations/19930082613/downloads/19930082613.pdf", "page_number": 22, "total_pages": 46, "image_filename": "19930082613_p22.jpg", "text": "NACA TN 1938\n21\n\n[Figure: A photograph of a cylindrical metal combustion-chamber liner with various holes and louvers. Annotations point to specific features on the liner.]\n\nEvidence of carbonaceous deposit present on outer surfaces of upper portion of liner\n\nHot metal upstream of louver\n\nBuckling between stress-relieving hole of louver and air intake\n\nCooled metal downstream of louver\n\n[Scale bar labeled INCHES with markings for 0, 1, and 2]\n\n[NACA logo]\nC-22311\n9-28-48\n\nFigure 4. - Combustion-chamber liner showing cracks extending from louver holes.", "timestamp": "2026-07-22T06:14:36.224461+00:00"}
{"citation_id": "19930085847", "source_url": "https://ntrs.nasa.gov/api/citations/19930085847/downloads/19930085847.pdf", "page_number": 7, "total_pages": 32, "image_filename": "19930085847_p7.jpg", "text": "NACA RM A9D04 CONFIDENTIAL 5\n\nterm was estimated to be roughly 6 percent of the $C_D \\cos \\alpha$ term at a Mach number of 0.78 for the conditions of these tests. Actually, even if the absolute magnitude of $C_d$ were not closely represented by $C_D$ for these tests, the change in $C_d$ with Mach number would still be given very closely by changes in $C_D$, since the tests were conducted at an essentially constant angle of attack near zero $\\alpha$. (The angle-of-attack deviation from the mean value over the test range was only $\\pm 0.6^\\circ$ as estimated from the measured values of lift coefficient.) The estimated change in $C_n \\sin \\alpha$ due to Mach number (applying Glauert factor to average test $C_n$) over the test range is of insignificant magnitude compared to the change in $C_D \\cos \\alpha$.\n\nCurves of chordwise force coefficient and profile drag plotted against Mach number are presented in figure 6 for the original slot to show the suction effect. In addition, data for the reduced length slot is shown in figure 6(a). The chordwise force coefficients were obtained from integration of thickness-wise pressure distributions. The profile drag coefficients were obtained from the wake-survey data using the method of reference 3. The momentum loss of the removed boundary-layer air (determined from a boundary-layer survey at the slot entrance) was added to that measured by the survey rake$^2$ in order to obtain the total section profile drag$^3$. As can be seen from the figure, any reduction in drag or chordwise force due to boundary-layer removal is within experimental accuracy and would appear to be negligible.\n\nDISCUSSION\n\nA qualitative idea of the causes for the drag rise (and the failure of the suction to modify this rise) can be obtained from examination of figures 7 and 8 which present representative chordwise and thickness-wise pressure distributions for selected test Mach numbers throughout the test range with and without suction. The chordwise distributions show that the critical pressure coefficient (for local Mach number equal 1.0) was exceeded on the upper surface over the entire range of these tests. The lower-surface pressures, however, did not exceed the critical until a Mach number of about 0.76 was attained. The shock apparently first formed on the upper surface near the 50-percent-chord station,\n\n$^2$Such an evaluation of the drag with suction applied neglects the drag equivalent of the suction power (assumes 100-percent duct efficiency).\n\n$^3$Comparison should be made only to a Mach number of 0.76 because of boundary-layer cross flow. Points affected by cross flow are indicated by the use of flags in figure 6(b).\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:14:37.572853+00:00"}
{"citation_id": "19930085859", "source_url": "https://ntrs.nasa.gov/api/citations/19930085859/downloads/19930085859.pdf", "page_number": 4, "total_pages": 31, "image_filename": "19930085859_p4.jpg", "text": "2\nNACA RM No. L9B25\n\nThis paper presents the results of the investigation of the wing-alone and wing-fuselage combinations employing a wing with the quarter-chord line swept back 35°, aspect ratio 4, taper ratio 0.6, and an NACA 65A006 airfoil section.\n\nMODEL AND APPARATUS\n\nThe wing of the semispan model had 35° of sweepback referred to the quarter-chord line, a taper ratio of 0.60, aspect ratio of 4, and an NACA 65A006 airfoil section parallel to the free stream. The wing was made of beryllium copper and the fuselage of brass. A two-view drawing of the model is presented in figure 1 while ordinates of the fuselage of fineness ratio 10 can be found in table I.\n\nThe model was mounted on an electrical strain-gage balance, which was enclosed in the bump, and the lift, drag, pitching moment, and bending moment about the model plane of symmetry were measured with calibrated galvanometers. The angle of attack was changed with a small electric motor and the value of the angle was determined with a calibrated slide-wire potentiometer.\n\nEffective downwash angles were determined for a range of tail heights by measuring the floating angles of five free-floating tails with the aid of calibrated slide-wire potentiometers. Details of the floating tails are shown in figures 2 and 3, while a photograph of the test setup on the bump, showing the floating tail mounted in the fuselage, is given in figure 4. The tails used in this investigation were the same as those used in the investigation reported in reference 1.\n\nA total-head comb was used to determine dynamic-pressure ratios for a range of tail heights in a plane which contained the 25-percent mean-aerodynamic-chord point of the free-floating tails. The total-head tubes were spaced 0.25 inch apart.\n\nSYMBOLS\n\n| | | |\n| :--- | :--- | :--- |\n| $C_L$ | lift coefficient | $\\left(\\frac{\\text{Twice panel lift}}{qS}\\right)$ |\n| $C_D$ | drag coefficient | $\\left(\\frac{\\text{Twice panel drag}}{qS}\\right)$ |", "timestamp": "2026-07-22T06:14:38.107602+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 46, "total_pages": 114, "image_filename": "19930086061_p46.jpg", "text": "```markdown\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\nP\n-4\n-3\n-2\n-1\n0\n1\n\nStation 1\n$\\frac{y}{b/2}, 0$\n\nP\n-4\n-3\n-2\n-1\n0\n1\n\nStation 2\n$\\frac{y}{b/2}, .0167$\n\nP\n-3\n-2\n-1\n0\n1\n\nStation 3\n$\\frac{y}{b/2}, .0333$\n\nUpper\nLower\nTwo dimensional\n(calculated at\nequal $c_l$)\n\n0 2 4 6 8 10\nx/c\n\n0 2 4 6 8 10\nx/c\n\n0 .2 .4 .6 .8 1.0\nx/c\n\n0 2 4 6 8 10\nx/c\n\n(a) Stations 1,2,3.\n\nNACA\n\nFigure 10.- Chordwise pressure distribution about wing 3 at angles of attack of $4.1^\\circ$, $8.1^\\circ$, $14.1^\\circ$, and $24.1^\\circ$; $\\Psi = 0^\\circ$.\n\nNACA RM L9J07\n\n24\n```", "timestamp": "2026-07-22T06:14:38.279437+00:00"}
{"citation_id": "19930085572", "source_url": "https://ntrs.nasa.gov/api/citations/19930085572/downloads/19930085572.pdf", "page_number": 10, "total_pages": 17, "image_filename": "19930085572_p10.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:14:38.969109+00:00"}
{"citation_id": "19930082703", "source_url": "https://ntrs.nasa.gov/api/citations/19930082703/downloads/19930082703.pdf", "page_number": 17, "total_pages": 28, "image_filename": "19930082703_p17.jpg", "text": "NACA TN 1983\n\n15\n\n[Figure: A black-and-white photograph of a helicopter on the ground, viewed from the side. The helicopter has a long tail boom with a star insignia and a two-bladed tail rotor. The main rotor blades are visible, and there is a person standing near the front landing gear. A label on the image reads “NACA L-55866”.]\n\nFigure 1.- Helicopter A.", "timestamp": "2026-07-22T06:14:41.302559+00:00"}
{"citation_id": "19930082566", "source_url": "https://ntrs.nasa.gov/api/citations/19930082566/downloads/19930082566.pdf", "page_number": 26, "total_pages": 44, "image_filename": "19930082566_p26.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:14:43.047513+00:00"}
{"citation_id": "19930082918", "source_url": "https://ntrs.nasa.gov/api/citations/19930082918/downloads/19930082918.pdf", "page_number": 13, "total_pages": 62, "image_filename": "19930082918_p13.jpg", "text": "```markdown\n12\nNACA TN 1940\n\nby 0.2 micron in the plane of polish and the particles were estimated\nto be spaced an average of 1 micron apart.\n\n(3) Aging at 1600° F resulted in almost the same type of\nreactions as at 1400° F with the exception that they were accelerated.\nThe boundary phase, for example, appeared after aging for 1 hour.\nThe precipitate size at the end of 1000-hour aging was estimated to\naverage 0.7 by 0.7 micron in the plane of polish and the particles\nwere estimated to be spaced an average of 4 microns apart. Definite\ninterfaces were present for each precipitate particle after aging\ntimes as short as 10 hours.\n\nX-Ray Studies\n\nLine intensity studies.- Figure 10 and table 1 show the results\nof line intensity studies on the 10-hour solution-treated material\nwhen aged at 1200°, 1400°, and 1600° F. Dehlinger (reference 6) and\nsubsequent authors have considered the effects that short- and long-\nperiod lattice distortions have on diffraction lines. In essence the\nconclusions are that short-period disturbances ($10^{-8}$ to $10^{-6}$ cm)\nresult in reduction of line peak intensities without appreciable\nbroadening and that long-period disturbances ($10^{-5}$ to $10^{-4}$ cm) cause\nline broadening with, however, the integrated intensity remaining\nconstant. Accordingly, the results of line peak-intensity studies\nshown in table 1 could be interpreted, in the absence of line-\nbroadening data, as either short-period disturbances or long-period\ndisturbances. If the latter were present, however, broadening should\nbe the predominant effect.\n\nThe two minimums for the material aged at 1200° F in figure 10\nlead to the possibility that two separate processes were observed\nat 1200° F. Aging at 1400° or 1600° F apparently resulted in only\none process occurring at each temperature which decreased line peak\nintensity.\n\nIt will be noted that the scatter for the measurements on material\naged at 1400° F was approximately the same as for the measurements on\nmaterial aged at 1200° or 1600° F. Considering the two methods of\nmeasurement that were used, it is felt that the scatter was still due\nto unrandom grain distribution despite rotation of the samples. Before\nquantitative calculations can be made with such line intensity measure-\nments, additional methods for alleviating unrandom grain distribution\nwill have to be developed. Only qualitative conclusions are thus drawn\nfrom the intensity data in this report.\n\nMatrix lattice-parameter measurements.- The measurement of lattice\nparameter as a function of aging time at a particular temperature can\ngive direct evidence of whether the precipitate particles are still\n```", "timestamp": "2026-07-22T06:14:44.531131+00:00"}
{"citation_id": "19930093769", "source_url": "https://ntrs.nasa.gov/api/citations/19930093769/downloads/19930093769.pdf", "page_number": 39, "total_pages": 39, "image_filename": "19930093769_p39.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:14:47.256066+00:00"}
{"citation_id": "19930082592", "source_url": "https://ntrs.nasa.gov/api/citations/19930082592/downloads/19930082592.pdf", "page_number": 25, "total_pages": 50, "image_filename": "19930082592_p25.jpg", "text": "24\nNACA TN 1914\n\n<!-- Image (104, 120, 939, 874) -->\n\nFigure 6. - Effect of time, temperature, and cobalt content on oxidation penetration of titanium carbide - cobalt cermets.", "timestamp": "2026-07-22T06:14:49.230303+00:00"}
{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 11, "total_pages": 72, "image_filename": "19930085491_p11.jpg", "text": "10 CONFIDENTIAL NACA RM No. A8J04\n\nTHEORETICAL CONSIDERATIONS\n\nAerodynamic Characteristics\n\nThe theoretical characteristics of the wings in this investigation have been determined on the basis of linear theory or approximate linear theory insofar as practicable. Existing theory permits the determination, exclusive of the effects of viscosity, of the lift and pitching moment and the drag due to lift for four of the wings. For the most highly swept plan form (WF-70) the Mach lines from the root trailing edge intersect the wing leading edge and the solution for this case was not attempted.\n\nSince by symmetry, the values of lift and moment at zero angle of attack are zero, and the drag at zero angle of attack is a minimum, the longitudinal aerodynamic characteristics can be defined by the lift-curve slope $dC_L/da$, the moment-curve slope $dC_m/dC_L$, the minimum drag coefficient $C_{D\\text{min}}$, and the drag-rise factor $(\\Delta C_D/\\Delta C_L)^2$.\n\nSlopes of lift and pitching-moment curves.— The linear theory as applied by Cohen (reference 10) has been used to determine the slopes of the lift and pitching-moment curves. The moment curves obtained by the linear theory are linear; whereas the experimental curves were found in all cases to be nonlinear. Hence, the lengthy calculations to obtain the moment-curve slopes were carried out only for WF-63, the basic configuration. The theory of reference 10 is exact to the order of the linear theory for the wings of WF-57 and WF-60. The values obtained for the wings of WF-63 and WF-67 must be considered approximate since the solution for the pressures in the area between the root trailing-edge Mach line and the trailing edge is obtained by a method which involves certain minor violations of the boundary conditions. It is believed, however, that these calculated slopes are close to the values that would be obtained from the exact linear theory solution.\n\nMinimum drag coefficient.— For convenience in the analysis, the minimum drag coefficients have been treated in terms of their components, the thickness and friction drag of the wings, and the thickness and friction drag of the fuselage. It should be noted that in determining the skin-friction coefficients the low-speed skin-friction coefficients have been used. Because it was not possible to determine quantitatively the skin-friction coefficient within observed separated flow regions, this component of wing drag was obtained by assuming completely laminar flow at a Reynolds\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:14:50.089105+00:00"}
{"citation_id": "19930082546", "source_url": "https://ntrs.nasa.gov/api/citations/19930082546/downloads/19930082546.pdf", "page_number": 24, "total_pages": 65, "image_filename": "19930082546_p24.jpg", "text": "NACA TN No. 1870\n23\n\n3. Blade plan form and solidity do not seem to be significant parameters. Tip clearance divided by propeller diameter is shown to be significant.\n\n4. At all tip Mach numbers the four-blade propeller produced smaller pressures than the two-blade propeller for the same power coefficient. At low tip Mach numbers these differences are large, whereas at tip Mach number 1.00, where a large amount of energy appears in the higher harmonics, they are relatively small.\n\n5. A flat vertical wall in the pressure field approximately doubles the free-space pressures in the plane of the wall; a circular wall also increases the pressures but by a lesser amount.\n\n6. Pressures of the fundamental frequency which impinge on the fuselage wall in front of the propeller plane tend to be out of phase with those behind the propeller plane.\n\n7. Oscillating pressures and their phase relations at any point in space may be predicted satisfactorily by the theory in this paper. This analysis is primarily for use in the region near the propeller tips where the Gutin simplified solution is not valid.\n\nLangley Aeronautical Laboratory\nNational Advisory Committee for Aeronautics\nLangley Air Force Base, Va., February 18, 1949", "timestamp": "2026-07-22T06:14:50.742305+00:00"}
{"citation_id": "19930083221", "source_url": "https://ntrs.nasa.gov/api/citations/19930083221/downloads/19930083221.pdf", "page_number": 12, "total_pages": 47, "image_filename": "19930083221_p12.jpg", "text": "10\nNACA TN No. 1824\n\nThe solution of equation (17) for boundary values of the type\nunder discussion has been indicated in reference 7 through considera-\ntion of an analogue problem in supersonic lifting-surface theory.\nThus, the shaded areas in figures 1(a) and 1(b) are thought of as\nswept-forward lifting surfaces situated in a stream directed along\nthe positive t axis at a Mach number $M_0 = \\sqrt{2}$. The boundary values\nremain the same; that is, $\\Phi_z = w = -V_0\\alpha$ on the wing and $\\Phi_t, \\Phi_y, \\Phi_z$\nare continuous functions of z elsewhere in the xt plane. In\nlifting-surface terminology, the unsteady case for supersonic speed\nbecomes a wing with supersonic leading edge, while the case indicated\nin figure 1(b) involves a subsonic leading edge.\n\nThe solution for the wing traveling at supersonic speed has been\ngiven in reference 7 in a form valid for all Mach numbers greater\nthan or equal to one. The expressions for load coefficient $\\Delta p/q$,\nwhere\n$$ \\frac{\\Delta p}{q} = \\frac{p_l - p_u}{\\frac{1}{2}\\rho_0 V_0^2} $$\ndiffer analytically in various regions of the xt plane. These\nexpressions are:\n\nRegion A (between lines $x = -M_0 t$, $x = -t$)\n$$ \\frac{\\Delta p}{q} = \\frac{4\\alpha}{\\sqrt{M_0^2 - 1}} \\quad (18a) $$\n\nRegion B (between lines $x = -t$, $x = t$, and $x = c_0 - M_0 t$)\n$$ \\frac{\\Delta p}{q} = \\frac{4\\alpha}{\\sqrt{M_0^2 - 1}} \\left[ \\frac{1}{\\pi} \\text{arc cos} \\frac{M_0 x + t}{x + M_0 t} + \\frac{\\sqrt{M_0^2 - 1}}{\\pi M_0} \\left( \\frac{\\pi}{2} + \\text{arc sin} \\frac{x}{t} \\right) \\right] \\quad (18b) $$\n\nRegion C (between lines $x = t$, $t = 0$, and $x = c_0 - M_0 t$)\n$$ \\frac{\\Delta p}{q} = \\frac{4\\alpha}{M_0} \\quad (18c) $$\n\nFrom the pressure distributions it is possible to calculate the\nindicial lift coefficient $C_{L\\alpha}(t)$ as a function of $M_0$ and t.", "timestamp": "2026-07-22T06:15:00.246952+00:00"}
{"citation_id": "19930082613", "source_url": "https://ntrs.nasa.gov/api/citations/19930082613/downloads/19930082613.pdf", "page_number": 23, "total_pages": 46, "image_filename": "19930082613_p23.jpg", "text": "Page intentionally left blank\n\nPage intentionally left blank", "timestamp": "2026-07-22T06:15:01.607127+00:00"}
{"citation_id": "19930085548", "source_url": "https://ntrs.nasa.gov/api/citations/19930085548/downloads/19930085548.pdf", "page_number": 10, "total_pages": 46, "image_filename": "19930085548_p10.jpg", "text": "NACA RM No. E8L30\n\nAt very lean fuel-air ratios (below 0.025), figure 11 indicates that the indicated mean effective pressure is roughly proportional to the fuel flow. As the mixture is enriched, the slope of the curve decreases and ultimately reaches a value of zero. This point corresponds to approximately stoichiometric mixture in the cylinder, which is approximately twice the over-all fuel-air ratio for the chosen operating conditions and resultant charging efficiency.\n\nFor changes in inlet-manifold pressure or inlet-manifold temperature, the data of figures 11 and 12 show upon analysis that the indicated mean effective pressure is almost directly proportional to the density of the air in the inlet manifold. This proportionality is a natural result of the manner in which the operating conditions were changed inasmuch as the air flow was held constant at 1 cylinder volume per cycle. Small changes in charging and thermal efficiencies occurring through changes in manifold conditions caused slight variations from this relation.\n\nComparison with calculated results. - The power output from the experimental cylinder cannot be directly compared with that calculated in the analysis (reference 1) because the combustion pressure rise used in the analysis was considerably greater than that experimentally found. A valid comparison between the calculated and experimental power results may be made, however, if the inlet-manifold pressure is kept constant and the results are compared at the same values of maximum cylinder pressure and exhaust-gas temperature, the two limiting factors in the gas-generator engine. Such a comparison is shown in figure 13 for values of mean effective pressure calculated according to thermodynamic relations shown in reference 1 for equivalent values of experimental maximum cylinder pressure and corresponding exhaust-gas temperatures. The results are in fair agreement at conditions of low exhaust temperature, which correspond to lean fuel-air ratios (less than 0.03). Optimizing the injection advance angle may improve the agreement here by increasing the power output and also decreasing the cylinder pressure and the exhaust temperature. At conditions corresponding to rich fuel-air ratios, wide differences are noted between the calculated and experimental values. Most of the difference is attributed to the poor charging efficiency of the experimental cylinder, although injection advance, duration, and pattern also exert some influence on the power output.\n\nCylinder Pressures\n\nCompression pressures. - A comparison of calculated and experimental compression pressures is shown in figure 14. The expression used in the reference analysis (reference 1),", "timestamp": "2026-07-22T06:15:02.187073+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 11, "total_pages": 149, "image_filename": "19930083192_p11.jpg", "text": "```markdown\nNACA TN 1976\n7\n\nfrom which the acceleration ratio $\\left(\\frac{\\Delta n}{\\Delta n_g}\\right)_1$ is seen to be a function of the rate of development of transient lift, the mass parameter, and the shape of the acceleration curve.\n\nThe preceding equation has been solved by Fredholm's method (reference 6) and the solutions obtained are of the form\n\n$$\n\\left(\\frac{\\Delta n}{\\Delta n_g}\\right)_{max} = C - \\frac{2B}{2\\mu_g + H}\n$$\n\nwhere C and B are constants representing solution of the integrals in the equation and whose values depend only on the distance penetrated into the gust. Figure 4 gives the solution for a range of gradient distance and mass parameter. Since the unsteady-lift functions $C_{L_g}$ and $C_{L_\\alpha}$ given in figure 3 were approximated for distances above about 8 chords, the curves for the larger gradient distances in figure 4 are also approximate. The dash line in the figure represents a rough limit of applicability of the gradient-gust solution obtained by this method. Above the dash line, the solutions for finite gradients are not valid, and the curve for H = 0 is to be used. The dash line was determined on the assumption that no gradient gust has a peak acceleration occurring earlier than a gust with zero gradient distance.\n\nFigure 4, together with the sharp-edge-gust formula, can be used to evaluate flight records to obtain the \"true\" gust intensity. The gradient distance $H = Vt_1$ and airplane mass parameter $\\mu_g$ permit the determination of the acceleration ratio from figure 4. The substitution of pertinent values in the sharp-edge-gust formula yields an effective gust velocity which, when divided by the acceleration ratio, gives the true gust velocity. In this case, true airspeed and the actual air density are to be used in the sharp-edge-gust equation.\n\nIn the evaluation of time-history data of airspeed, altitude, and acceleration to obtain gust-gradient distance and true gust intensity, the following methods have been used: The acceleration increment is taken as the acceleration increment at the peak minus the steady value prior to the start of the gust. In the determination of the gradient distance, the time from zero to peak acceleration and the true airspeed is used. In expressing the gradient distance in chords, the mean geometric chord is generally used.\n```", "timestamp": "2026-07-22T06:15:02.776689+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 86, "total_pages": 96, "image_filename": "19930085880_p86.jpg", "text": "84\nNACA RM No. L9C03\n\n| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | |", "timestamp": "2026-07-22T06:15:07.796061+00:00"}
{"citation_id": "19930082703", "source_url": "https://ntrs.nasa.gov/api/citations/19930082703/downloads/19930082703.pdf", "page_number": 18, "total_pages": 28, "image_filename": "19930082703_p18.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:15:10.045379+00:00"}
{"citation_id": "19930085572", "source_url": "https://ntrs.nasa.gov/api/citations/19930085572/downloads/19930085572.pdf", "page_number": 11, "total_pages": 17, "image_filename": "19930085572_p11.jpg", "text": "1061\n\nNACA RM No. E8L02\n\n521808\nBF-808\n\nFigure 1. - Turbojet engine extended from medium-bomber-type airplane in flight.\n\nNACA\nC-21990\n6-8-48\n\n6", "timestamp": "2026-07-22T06:15:13.990456+00:00"}
{"citation_id": "19930085847", "source_url": "https://ntrs.nasa.gov/api/citations/19930085847/downloads/19930085847.pdf", "page_number": 8, "total_pages": 32, "image_filename": "19930085847_p8.jpg", "text": "6 CONFIDENTIAL NACA RM A9D04\n\nwhich is 10 percent behind the point of maximum thickness. The upper-surface shock then moved back with increase in Mach number reaching about 59-percent chord at 0.76 Mach number and about 63-percent chord at 0.83 Mach number. The lower-surface shock apparently forms near the maximum thickness point initially and then moves back much more abruptly than the upper-surface shock, arriving at about 71-percent chord at 0.83 Mach number. As can be seen from the figure, the wing-surface pressure rise became more abrupt as the shock moved rearward. Examination of the trailing-edge pressures for the various distributions presented shows separation to have begun between a Mach number of 0.78 and a Mach number of 0.80 and to be present at higher Mach numbers, separation being indicated by the degree of trailing-edge pressure recovery. The meaning of these changes in terms of changes in airfoil pressure drag can be seen by referring to the thickness-wise pressure distributions also presented in the figures.\n\nThe thickness-wise pressure distributions show the contribution of the various regions along the airfoil contour to the chordwise pressure force.⁴ The area between the forebody and afterbody pressure distributions has been crosshatched to show whether the resultant differential force at any given vertical ordinate is a thrust or a drag force. Integration of the areas so enclosed yields the contribution to chordwise force coefficient. The large drag area roughly centered about the line of zero thickness and with its centroid to the left of the line of zero pressure coefficient is due largely to the basic pressure distribution of the airfoil (unmodified by any boundary-layer or separation effects). It should be noted that most of this area would exist at low Mach numbers as well as the Mach numbers of these tests. The primary effect of flow separation at the Mach number of 0.83 is to increase the area above the low-speed size by making the trailing-edge pressures on the upper surface more negative. The drag areas near the points of maximum thickness and with centroids to the right of the line of critical pressure coefficient are due to the occurrence of supersonic flow and attendant rearward shock movement. These drag areas are due to the combined effect of pressure decrease ahead of the shock associated with the change from subsonic to supersonic flow and the decreasing vertical ordinates back of the point of maximum thickness.\n\nAn over-all inspection of the curves just discussed shows that the drag rise to 0.76 Mach number can be traced to the rearward movement of shock on the upper surface with no increase due to separation since separation has not yet formed. The additional drag rise to 0.83 Mach\n\n⁴ Chordwise pressure force is used for the reasons given in the text of the section on Tests and Results.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:15:16.710929+00:00"}
{"citation_id": "19930082566", "source_url": "https://ntrs.nasa.gov/api/citations/19930082566/downloads/19930082566.pdf", "page_number": 27, "total_pages": 44, "image_filename": "19930082566_p27.jpg", "text": "NACA TN No. 1889\n25\n\n[Figure: Photograph of a measurement apparatus mounted on a wall. The apparatus includes various electrical boxes, cables, and a large circular gauge with a needle. Labels on the apparatus include 'N', 'B', 'J', and 'K'.]\n\nNACA\nFigure 9.- Measurement of fluctuating axial load.", "timestamp": "2026-07-22T06:15:18.746479+00:00"}