PDF/ocr_dataset_1040.jsonl

{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 31, "total_pages": 96, "image_filename": "19930085880_p31.jpg", "text": "NACA RM No. L9C03\n29\n\n[Figure: A graph plotting Lift (lb) against Speed (fps). The y-axis ranges from 0 to .35. The x-axis ranges from 0 to 35. There are multiple lines originating from (0,0) representing different Wetted areas (sq ft) from 0 to .30. An inset shows two rectangular shapes labeled 250A and 250B.]\n\n(c) $\\tau = 16^\\circ$.\nFigure 12.- Continued.", "timestamp": "2026-07-22T05:32:28.155194+00:00"}
{"citation_id": "19930085964", "source_url": "https://ntrs.nasa.gov/api/citations/19930085964/downloads/19930085964.pdf", "page_number": 11, "total_pages": 18, "image_filename": "19930085964_p11.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:32:32.062548+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 5, "total_pages": 114, "image_filename": "19930086061_p5.jpg", "text": "NACA RM L9J07\nRESTRICTED\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nRESEARCH MEMORANDUM\n\nLOW-SPEED PRESSURE-DISTRIBUTION AND FLOW INVESTIGATION FOR A\nLARGE PITCH AND YAW RANGE OF THREE LOW-ASPECT-RATIO\nPOINTED WINGS HAVING LEADING EDGE SWEPT BACK 60°\nAND BICONVEX SECTIONS\n\nBy Ralph W. May, Jr., and John G. Hawes\n\nS U M M A R Y\n\nPressure distributions and flow characteristics were investigated\nat low speed through a yaw range from 0° to 35° and an angle-of-attack\nrange through the stall for three small-scale low-aspect-ratio pointed\nwings having 10-percent-thick biconvex sections, 60° sweptback leading-\nedge, and 0°, 30°, and -30° trailing-edge sweep.\n\nAn effort was made to correlate the pressure distributions with the\nstrong conical vortex flow observed. At zero yaw, separation vortices,\nemanating in the region of the wing apexes, increased in size and were\nswept back farther from the leading edge along the span as the angle of\nattack was increased. Flow observations showed that the center of\nvortex rotation coincided with the maximum depth of a region of turbulent\nseparated flow and with a negative pressure peak. Behind the center of\nvortex rotation a negative-pressure dip occurred as the depth of the\nturbulent region diminished rather rapidly. With increasing angle of\nyaw the separation vortices along the leading and trailing semispans\nbecame more clearly defined as bound and trailing vortices, respectively.\n\nSection lift coefficients and local centers of pressure at zero yaw\nand spanwise load distributions throughout the yaw range are presented\nand discussed with reference to the flow analysis. Force and moment\ncharacteristics of the three wings are compared throughout the large\nyaw range.\n\nI N T R O D U C T I O N\n\nIn the Langley full-scale-tunnel investigations of the German\ndelta-wing DM-1 glider (reference 1), a remarkable effect of a sharp\n\nRESTRICTED", "timestamp": "2026-07-22T05:32:43.637466+00:00"}
{"citation_id": "19930085934", "source_url": "https://ntrs.nasa.gov/api/citations/19930085934/downloads/19930085934.pdf", "page_number": 12, "total_pages": 23, "image_filename": "19930085934_p12.jpg", "text": "NACA RM E9G12\n\nEntropy of dry air at outlet\n\n$$\n\\begin{aligned}\ns_{d,2} &= \\Phi_{t,2} - R \\log_e N \\\\\n&= 0.08403 - 0.01681 \\\\\n&= 0.06722 \\text{ (Btu/(lb)(}^\\circ\\text{F))}\n\\end{aligned}\n$$\n\n(6) Entropy of mixture at outlet. - The entropy of mixture per pound of dry air at the outlet equals the entropy of saturated liquid per pound of dry air plus the entropy of saturated vapor per pound of dry air plus the entropy of dry air.\n\n$$\n\\begin{aligned}\ns_{m,a,2} &= s_{f,a,2} + s_{g,a,2} + s_{d,2} \\\\\n&= 0.002900 + 0.07859 + 0.06722 \\\\\n&= 0.1487 \\text{ (Btu/(lb)(}^\\circ\\text{F))}\n\\end{aligned}\n$$\n\nWhen $s_{m,a,1}$ and $s_{m,a,2}$ are equal, the assumed temperature is that for isentropic compression. If they are unequal, a new temperature $t_{t,2}$ must be selected until equality is attained. The entropy value as computed is sufficiently close to that at the inlet to assume an isentropic process.\n\nIII. Enthalpy Change Through Compressor\n\n(1) Enthalpy at inlet. -\n\nEnthalpy of saturated liquid at inlet\n\n$$\nh_{f,1} \\text{ at } t_{f,1} \\text{ of } 55^\\circ \\text{ F} = 23.07 \\text{ (Btu/lb)}\n$$\n\nEnthalpy of saturated liquid per pound of dry air at inlet\n\n$$\n\\begin{aligned}\nh_{f,a,1} &= h_{f,1} \\text{ w/a} \\\\\n&= (23.07)(0.04997) \\\\\n&= 1.153 \\text{ (Btu/lb)}\n\\end{aligned}\n$$\n\nEnthalpy of saturated vapor at inlet\n\n$$\nh_{g,1} = 1096 \\text{ (Btu/lb)} \\text{ at } p_{s,1} \\text{ of } 0.1110 \\text{ and } t_{t,1} \\text{ of } 77.4^\\circ \\text{ F}\n$$", "timestamp": "2026-07-22T05:32:44.835152+00:00"}
{"citation_id": "19930085906", "source_url": "https://ntrs.nasa.gov/api/citations/19930085906/downloads/19930085906.pdf", "page_number": 18, "total_pages": 23, "image_filename": "19930085906_p18.jpg", "text": "NACA RM E9F20 CONFIDENTIAL 17\n\n1153\n\nFuel temperature, °F\n\n| Distance between preheater and flame holder (in.) |\n|---|\n| 0 |\n| 4 |\n| 12 |\n\nTime, min\n\n(b) Fuel flow, 2600 pounds per hour; fuel-air ratio, 0.054 ±0.001.\n\nFigure 6. - Continued. Effect of preheater position on fuel-heating rate.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:32:45.507786+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 65, "total_pages": 99, "image_filename": "19930082511_p65.jpg", "text": "NACA TN No. 1826\n63\n\nAPPENDIX C\n\nDERIVATION OF EQUATION (32)\n\nEquation (32) was derived for use in calculating some of the results given in reference 13, but it was not explicitly stated and discussed in that paper. Because certain steps in its development are not obvious, the present outline of its derivation is given. Familiarity with reference 13 will be assumed.\n\nCertain difficulties arise in the treatment of the doublet line directly; so the result is found by considering a horseshoe vortex of finite span and letting the span approach zero. Equation (6) of reference 13 gives the formula for $\\left. \\frac{\\partial \\Phi_2}{\\partial \\rho} \\right|_{\\rho=1}$ (where $\\Phi_2$ is defined in reference 13) corresponding to a horseshoe vortex of strength $\\Gamma$ and span $\\sigma$ having one trailing vortex along the tunnel axis and the other to the right of the axis. The procedure for the doublet consists of letting the yaw angle $\\Psi$ be zero, expanding the radicals in ascending powers of $\\sigma$, and proceeding to the next step in the analysis, where $\\sigma$ will eventually be made to approach zero. In the expansions, powers of $\\sigma$ higher than the first may be neglected except where $\\sigma$ occurs in the product $\\xi \\sigma$, since $\\xi$ takes on infinite values; furthermore, since for the doublet the field should be symmetrical about the vertical plane of symmetry ($\\theta = \\frac{\\pi}{2}$), unsymmetrical factors, as $\\sigma \\cos \\theta$, may be immediately eliminated. The formula for $\\left. \\frac{\\partial \\Phi_2}{\\partial \\rho} \\right|_{\\rho=1}$ is thus\n\n$$\n\\left. \\frac{\\partial \\Phi_2}{\\partial \\rho} \\right|_{\\rho=1} = -\\frac{\\Gamma \\sigma}{4\\pi} \\lim_{\\sigma \\to 0} \\left\\{ \\sin \\theta \\left( \\frac{\\xi \\sigma}{\\sqrt{1 + \\xi^2 \\sigma^2}} - \\frac{\\xi}{\\sqrt{1 + \\xi^2}} \\right) \\right.\n$$\n$$\n\\left. - \\frac{\\xi \\sin \\theta}{\\sigma(\\xi^2 + \\sin^2 \\theta)} \\left( 1 - \\frac{1}{\\sqrt{1 + \\xi^2 \\sigma^2}} \\right) \\right.\n$$\n$$\n\\left. - \\frac{\\xi \\sin \\theta}{\\xi^2 + \\sin^2 \\theta} \\left[ \\frac{1}{\\sqrt{1 + \\xi^2}} - \\frac{\\cos^2 \\theta}{(1 + \\xi^2)^{3/2}} \\right] \\right\\} \\quad \\text{(C1)}\n$$\n\nAccording to the procedure of reference 13, it is necessary to make a Fourier analysis of the three terms in the braces and then to insert the Fourier coefficients in equation (8) of reference 13.", "timestamp": "2026-07-22T05:32:46.164269+00:00"}
{"citation_id": "19930085914", "source_url": "https://ntrs.nasa.gov/api/citations/19930085914/downloads/19930085914.pdf", "page_number": 16, "total_pages": 42, "image_filename": "19930085914_p16.jpg", "text": "NACA RM A9D25\n15\n\n[Figure: A man in a white shirt and tie stands next to a model aircraft wing mounted on a stand inside a wind tunnel. The wing is swept back. The NACA logo and identifier A-13253 are visible in the bottom right corner.]\n\n(a) Rear view.\n\n[Figure: A man in a white shirt and tie stands behind a model aircraft wing mounted on a stand inside a wind tunnel, viewed from above. The wing has a distinct V-shape planform. The NACA logo and identifier A-13254 are visible in the bottom right corner.]\n\n(b) Plan view.\n\nFigure 1.- Model of the cambered and twisted wing with the leading edge swept back $63^\\circ$ in combination with a fuselage.", "timestamp": "2026-07-22T05:32:48.305212+00:00"}
{"citation_id": "19930085869", "source_url": "https://ntrs.nasa.gov/api/citations/19930085869/downloads/19930085869.pdf", "page_number": 36, "total_pages": 36, "image_filename": "19930085869_p36.jpg", "text": "```markdown\nCONFIDENTIAL\n\nLoad coefficient, $C_A$\nLoad-resistance ratio, $\\Delta/R$\nResistance coefficient, $C_R$\nTrim, deg\n\nSpeed coefficient, $C_V$\n\nLoad-resistance ratio\nTrim\nTail boom cleared water\nLoad coefficient\nResistance coefficient\n\nNACA\n\nFigure 14.- Resistance coefficient, load-resistance ratio, trim and load coefficient for minimum stable resistance.\n\nCONFIDENTIAL\n\nNACA-Langley - 9-12-49 - 300\nNACA RM L9D15\n34\n```", "timestamp": "2026-07-22T05:32:48.942565+00:00"}
{"citation_id": "19930085962", "source_url": "https://ntrs.nasa.gov/api/citations/19930085962/downloads/19930085962.pdf", "page_number": 11, "total_pages": 51, "image_filename": "19930085962_p11.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:32:50.560450+00:00"}
{"citation_id": "19930085977", "source_url": "https://ntrs.nasa.gov/api/citations/19930085977/downloads/19930085977.pdf", "page_number": 5, "total_pages": 33, "image_filename": "19930085977_p5.jpg", "text": "4\nCONFIDENTIAL\nNACA RM L9H22\n\n$M_1$\nlocal Mach number\n\n$M_a$\naverage chordwise local Mach number\n\nR\nReynolds number of wing based on $\\bar{c}$\n\n$\\alpha$\nangle of attack, degrees\n\n$\\epsilon$\neffective downwash angle, degrees\n\n$q_{wake}/q$\nratio of point dynamic pressure to free-stream dynamic pressure\n\n$y_{cp}$\nlateral center of pressure, percent semispan\n\n$$\n\\left( \\frac{100 C_B}{C_L} \\right)\n$$\n\n$h_t$\ntail 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 4.)\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 6. It is seen that there is a Mach number variation of about 0.05 over the model semispan at low Mach numbers and from 0.07 to 0.08 at the higher Mach numbers. The chordwise Mach number variation is generally less than 0.01. No attempt has been made to evaluate the effects of this chordwise and spanwise Mach number variation. Note that the long-dashed lines shown near the root of the wing (fig. 6) represent a local Mach number 5 percent below the maximum value and indicate the extent of the bump boundary layer. The effective test Mach number was obtained from contour charts similar to those presented in figure 6 using the relationship\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:32:51.370726+00:00"}
{"citation_id": "19930085542", "source_url": "https://ntrs.nasa.gov/api/citations/19930085542/downloads/19930085542.pdf", "page_number": 34, "total_pages": 46, "image_filename": "19930085542_p34.jpg", "text": "```markdown\n32\nNACA RM No. L8L29\n\n$C_{Yp}$\n.4\n0\n-.4\n\nModel No. $\\Lambda_{c/4}$ A\n4 $70.4^\\circ$ 1.07\n2 $52.2^\\circ$ 2.31\n7 $36.9^\\circ$ 4.00\n\n$C_{np}$\n.2\n0\n-.2\n\n$C_{lp}$\n0\n-.2\n-.4\n\n-2 0 .2 .4 .6 .8 1.0\nLift coefficient, $C_L$\n\nFigure 13.— Effect of aspect ratio of triangular wings of NACA 0012 profile on $C_{Yp}$, $C_{np}$, and $C_{lp}$.\n```", "timestamp": "2026-07-22T05:32:53.489322+00:00"}
{"citation_id": "19930085889", "source_url": "https://ntrs.nasa.gov/api/citations/19930085889/downloads/19930085889.pdf", "page_number": 28, "total_pages": 37, "image_filename": "19930085889_p28.jpg", "text": "NACA RM L9F14\n27\n\nCONFIDENTIAL\n\n$R \\times 10^{-6}$\n.280\n.395\n.558\n.718\n.880\n1.116\n\nPitching-moment coefficient, $C_m$\n.04\n0\n-.04\n\nLift coefficient, $C_L$\n-2 0 .2 .4 .6 .8 1.0 1.2\n\no Transition strips off\n□ Transition strips on\n\nCONFIDENTIAL\nNACA\n\nFigure 9.— Variation of the pitching-moment coefficient with lift coefficient of the 46.7° sweptback wing alone for various values of Reynolds number with and without transition strips on wing leading edge.", "timestamp": "2026-07-22T05:32:59.987023+00:00"}
{"citation_id": "19930085912", "source_url": "https://ntrs.nasa.gov/api/citations/19930085912/downloads/19930085912.pdf", "page_number": 17, "total_pages": 36, "image_filename": "19930085912_p17.jpg", "text": "```markdown\nNACA RM No. E9C16\n15\n\nWhen these computations were checked with a Mollier chart for saturated air, the same result was obtained.\n\nDetermination of total orifice area $A_j$. - The flow coefficient $C$ of a choked jet is known to be approximately 0.87 at a plenum-chamber pressure of 80 inches of mercury and a stream static pressure of 25 inches of mercury from unpublished data. From the mass-flow equation,\n\n$$A_j = \\frac{W_g/g}{C V_j \\rho_j}$$\n\nThe jet velocity and density were calculated as in reference 2, where\n\n$$\\rho_j = 0.0261 \\frac{P_j}{T_j} = 0.0261 \\times \\frac{80}{1460} = 0.00143 \\text{ slug/cu ft}$$\n\n$$V_j = 44.8 \\sqrt{T_j} = 44.8 \\sqrt{1460} = 1710 \\text{ ft/sec}$$\n\nTherefore\n\n$$A_j = \\frac{1.3/32.2}{0.87 \\times 0.00143 \\times 1710} = 0.01890 \\text{ sq ft} = 2.72 \\text{ sq in.}$$\n\nDetermination of orifice configuration. - The orifice configuration was determined in accordance with the design criterions:\n\n(a) The total jet area must be approximately 2.72 square inches.\n(b) The layout of orifices must be symmetrical because the inlet and the ducting are symmetrical.\n\nThe area of the duct protected by a jet at a section corresponding to the position of the tip of the accessory housing was constructed for the maximum pressure and temperature available in the plenum chamber and for design free-stream velocity. The maximum available plenum-chamber temperature and pressure were $1000^\\circ$ F and 80 inches of mercury absolute, respectively. By use of the results of reference 2, the jet penetration for a given orifice diameter was calculated from the jet-penetration equation\n\n$$\\left(\\frac{l}{D_j}\\right)^{1.65} = 2.91 \\frac{\\rho_j V_j}{\\rho_1 V_1} \\sqrt{\\frac{s}{D_j}}$$\n\nwhere\n\n$l$ depth of jet penetration into air stream at distance $s$ downstream of orifice center line measured from duct wall, in.\n\n$s$ distance from orifices to accessory-housing tip, 47.25 in.\n```", "timestamp": "2026-07-22T05:33:01.964422+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 32, "total_pages": 96, "image_filename": "19930085880_p32.jpg", "text": "30\nNACA RM No. L9C03\n\n.35\n.30\n.25\n.20\n.15\n.10\n.05\n0\nLift, lb\n\n0 10 15 20 25 30 35\nSpeed, fps\n\nWetted area\n(sq ft)\n0\n.05\n.10\n.15\n.20\n.25\n.30\n\n[Figure: Two vertical rectangular shapes labeled 250A and 250B]\n\n(d) T = 20°.\nFigure 12.- Concluded.", "timestamp": "2026-07-22T05:33:03.162968+00:00"}
{"citation_id": "19930085964", "source_url": "https://ntrs.nasa.gov/api/citations/19930085964/downloads/19930085964.pdf", "page_number": 12, "total_pages": 18, "image_filename": "19930085964_p12.jpg", "text": "```markdown\nNACA RM E9G25\n\n11\n\nTrailing\nedge\n\nExciting frequency, cps\n\nLeading\nedge\n\n<!-- Image (139, 126, 877, 877) -->\n\n(a) 845\n(b) 950\n(c) 1230\n(d) 1870\n(e) 2590\n\n(f) 3180\n(g) 3400\n(h) 4360\n(i) 4490\n(j) 4830\n\n(k) 5230\n(l) 5470\n(m) 6000\n(n) 6540\n(o) 6800\n\n(p) 7000\n(q) 7700\n(r) 8530\n(s) 8730\n(t) 9750\n\nNACA\n\nFigure 3. - Vibrational modes of plain hollow blade A. (Solid lines represent node lines\non concave side of blade; dashed lines represent node lines on convex side. Exciting\nfrequency in cycles per second is shown below each nodal pattern.)\n```", "timestamp": "2026-07-22T05:33:07.276530+00:00"}
{"citation_id": "19930085914", "source_url": "https://ntrs.nasa.gov/api/citations/19930085914/downloads/19930085914.pdf", "page_number": 17, "total_pages": 42, "image_filename": "19930085914_p17.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:33:16.952192+00:00"}
{"citation_id": "19930085906", "source_url": "https://ntrs.nasa.gov/api/citations/19930085906/downloads/19930085906.pdf", "page_number": 19, "total_pages": 23, "image_filename": "19930085906_p19.jpg", "text": "```markdown\n18\nCONFIDENTIAL\nNACA RM E9F20\n\nDistance between\npreheater and\nflame holder\n(in.)\n$\\square$ 4\n$\\diamond$ 8\n$\\triangle$ 12\n\nFuel temperature, $^\\circ$F\nTime, min\n\n[Figure: A graph plotting Fuel temperature ($^\\circ$F) against Time (min). The y-axis ranges from 40 to 320. The x-axis ranges from 0 to 6. There are three data series represented by square, diamond, and triangle markers.]\n\nNACA\n\n(c) Fuel flow, 2600 pounds per hour; fuel-air ratio, 0.061 $\\pm$0.001.\nFigure 6. - Concluded. Effect of preheater position on fuel-\nheating rate.\n\nCONFIDENTIAL\n1153\n```", "timestamp": "2026-07-22T05:33:17.313187+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 6, "total_pages": 114, "image_filename": "19930086061_p6.jpg", "text": "```markdown\n2\nNACA RM L9J07\n\nleading edge was observed. Whereas the flow over the original glider\nwith a round leading edge was essentially as expected (characterized by\nturbulent separation from the trailing edge with the separated region\nincreasing with angle of attack), the flow over the modified glider\nwith a sharp leading edge was characterized by a large vortex on the\nupper surface just behind the leading edge. The vortex remained\nattached up to high angles of attack and provided a considerably higher\nmaximum lift coefficient than that of the original configuration.\nAlthough such upper-surface attached vortices had been reported previously\nfor low-aspect-ratio airfoils, only relatively meager information was\navailable as to their causes and effects. In view of the likelihood that\nsuch flows would be encountered frequently on highly swept wings with\nsharp or small-radius leading edges, further efforts to define the flow\nand its effects on the wing characteristics were considered desirable.\n\nThe project herein reported represents one of the first steps in this\ndirection. Three related small-scale low-aspect-ratio pointed wings,\nliberally equipped with pressure orifices, were constructed and studied\nat low speeds in the entrance cone of the Langley full-scale tunnel.\nThe wings had 10-percent-thick biconvex sections parallel to the air\nstream, 60° sweptback leading edge, and 0°, 30°, and -30° sweep of the\ntrailing edge. Pressure distributions were obtained for a range of\nangles of attack through the stall and for yaw angles up to 35°. Extensive tuft and smoke studies were made to help clarify the flow and to\ncorrelate its characteristics with the measured pressure distributions.\n\nA number of independent but related studies, all for unyawed wings,\nexist: References 2 and 3 describe force and limited flow studies of\ndelta wings with sharp leading edges; and references 4 to 6 give pressure-\ndistribution and flow studies of delta wings with sharp and round\nleading edges of different relative radii of curvature. Reference 7\ndescribes a pressure-distribution and flow study through a yaw range\nof a wing with 47.5° of leading-edge sweep and with a sharp leading\nedge. Pressure distributions on a two-dimensional 6-percent-thick\nbiconvex airfoil are given in reference 8.\n\nS Y M B O L S\n\nConventional NACA coefficients, reduced from pressure-distribution\ndata neglecting chord force, are referred to the standard stability axes.\nThe Z-axis is in the plane of symmetry and perpendicular to the relative\nwind, the X-axis is in the plane of symmetry and perpendicular to the\nZ-axis, and the Y-axis is perpendicular to the plane of symmetry at the\nquarter chord of the mean aerodynamic chord.\n```", "timestamp": "2026-07-22T05:33:18.574206+00:00"}
{"citation_id": "19930085962", "source_url": "https://ntrs.nasa.gov/api/citations/19930085962/downloads/19930085962.pdf", "page_number": 12, "total_pages": 51, "image_filename": "19930085962_p12.jpg", "text": "NACA RM A9E05 CONFIDENTIAL 11\n\n[Figure: Semispan model of a horizontal tail mounted in a wind tunnel, with a ruler marked in inches and an arrow indicating airflow direction. NACA logo and identifier A-11999 visible.]\n\nFigure 2.— Semispan model of a horizontal tail of aspect ratio 4 mounted in the Ames 12-foot pressure wind tunnel.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:33:21.369391+00:00"}
{"citation_id": "19930085934", "source_url": "https://ntrs.nasa.gov/api/citations/19930085934/downloads/19930085934.pdf", "page_number": 13, "total_pages": 23, "image_filename": "19930085934_p13.jpg", "text": "```markdown\n12\nNACA RM E9G12\n\nPressure effect on enthalpy is negligible in this region of pressure.\n\nEnthalpy of saturated vapor per pound of dry air at inlet\n$$h_{g,a,1} = h_{g,1} q_1$$\n$$= (1096)(0.01025)$$\n$$= 11.23 \\text{ (Btu/lb)}$$\n\nEnthalpy of dry air at inlet\n$$h_{d,1} \\text{ at } T_{t,1} \\text{ of } 537.1^\\circ \\text{ R} = 32.87 \\text{ (Btu/lb)}$$\n\nThe enthalpy of mixture per pound of dry air at the inlet equals the enthalpy of saturated liquid per pound of dry air plus the enthalpy of saturated vapor per pound of dry air plus the enthalpy of dry air\n$$h_{m,a,1} = h_{f,a,1} + h_{g,a,1} + h_{d,1}$$\n$$= 1.153 + 11.23 + 32.87$$\n$$= 45.25 \\text{ (Btu/lb)}$$\n\n(2) Enthalpy at outlet. -\n\nEnthalpy of saturated liquid at outlet\n$$h_{f,2} \\text{ at } t_{t,2} \\text{ of } 108.1^\\circ \\text{ F} = 76.05 \\text{ (Btu/lb)}$$\n\nEnthalpy of saturated liquid per pound of dry air at outlet\n$$h_{f,a,2} = h_{f,2} \\left( \\frac{\\text{lb liquid water}}{\\text{lb dry air}} \\right)$$\n$$= (76.05)(0.02017)$$\n$$= 1.534 \\text{ (Btu/lb)}$$\n\nEnthalpy of saturated vapor at outlet\n$$h_{g,2} \\text{ at } t_{t,2} \\text{ of } 108.1^\\circ \\text{ F} = 1108.6 \\text{ (Btu/lb)}$$\n\n1164\n```", "timestamp": "2026-07-22T05:33:22.966752+00:00"}
{"citation_id": "19930085889", "source_url": "https://ntrs.nasa.gov/api/citations/19930085889/downloads/19930085889.pdf", "page_number": 29, "total_pages": 37, "image_filename": "19930085889_p29.jpg", "text": "28\nNACA RM L9F14\n\n<!-- Image (79, 226, 855, 792) -->\n\nFigure 10.- Effect of transition strips and Reynolds number on the lift-curve slope and location of the aerodynamic center for the 46.7° sweptback wing alone.", "timestamp": "2026-07-22T05:33:24.865396+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 33, "total_pages": 96, "image_filename": "19930085880_p33.jpg", "text": "NACA RM No. L9C03\n31\n\n[Figure: A graph plotting Lift (lb) against Speed (fps). The y-axis ranges from 0 to .35. The x-axis ranges from 0 to 35. The graph contains multiple lines originating from (0,0) representing different Wetted areas (sq ft) from 0 to .30. There is a schematic of an inverted triangle in the upper left corner of the plot area.]\n\n(a) $\\tau = 8^\\circ$.\n\nFigure 13.- Aerodynamic lift of model 250D.", "timestamp": "2026-07-22T05:33:26.508379+00:00"}
{"citation_id": "19930085977", "source_url": "https://ntrs.nasa.gov/api/citations/19930085977/downloads/19930085977.pdf", "page_number": 6, "total_pages": 33, "image_filename": "19930085977_p6.jpg", "text": "```markdown\nNACA RM L9H22 CONFIDENTIAL 5\n\n$$\nM = \\frac{2}{S} \\int_{0}^{b/2} c M_a \\, dy\n$$\n\nThe variation of mean test Reynolds number with Mach number is shown in figure 7. The boundaries in the figure indicate the range in Reynolds number caused by variations in atmospheric 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 mean aerodynamic chord of the free-floating tails to free-stream dynamic pressure were obtained for the model wing-alone and wing-fuselage configurations tested through a Mach number range of 0.60 to 1.18 and an angle-of-attack range of $-2^\\circ$ to $12^\\circ$. A few surveys were also made to determine the spanwise variation of wake dynamic pressure at a Mach number of 1.10.\n\nThe end-plate tare corrections to the drag and to the downwash data were obtained through the test Mach number range at $0^\\circ$ angle of attack by testing the model configurations without end plates. A gap of about 1/16 inch was maintained between the wing root chord and the bump surface, and a sponge wiper seal was fastened to the wing butt beneath the surface of the bump to minimize leakage. The end-plate tares were assumed to be constant with angle of attack, and the tares obtained at zero angle of attack were applied to all drag and downwash data. 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. No base-pressure correction has been applied to the wing-fuselage drag data.\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\n\nCONFIDENTIAL\n```", "timestamp": "2026-07-22T05:33:26.983079+00:00"}
{"citation_id": "19930085542", "source_url": "https://ntrs.nasa.gov/api/citations/19930085542/downloads/19930085542.pdf", "page_number": 35, "total_pages": 46, "image_filename": "19930085542_p35.jpg", "text": "5E\nNACA RM No. L8L29\n33\n\n| Model No. | $A_{WING}$ | $A_{FIN}$ |\n| :--- | :--- | :--- |\n| 2 | 2.31 | - |\n| 5 | 2.31 | .77 |\n| 6 | 2.31 | 1.15 |\n\nAngle of attack, $\\alpha$, deg\nLongitudinal-force coefficient, $C_X$\nPitching-moment coefficient, $C_m$\nLift coefficient, $C_L$\n\n[Figure: Graph showing aerodynamic characteristics with data points for Model Nos. 2, 5, and 6 plotted against Lift coefficient ($C_L$). The graph includes curves for Angle of attack ($\\alpha$), Longitudinal-force coefficient ($C_X$), and Pitching-moment coefficient ($C_m$).]\n\nNACA\n\nFigure 14.- Effect of fins on aerodynamic characteristics of a triangular wing of NACA 0012 profile. $\\Lambda_{0/4} = 52.2^\\circ$.", "timestamp": "2026-07-22T05:33:28.240519+00:00"}
{"citation_id": "19930085972", "source_url": "https://ntrs.nasa.gov/api/citations/19930085972/downloads/19930085972.pdf", "page_number": 6, "total_pages": 46, "image_filename": "19930085972_p6.jpg", "text": "4\nNACA RM L9B18\n\ntunnel is not known, but is believed to be small because of the large contraction ratio (14 to 1).\n\nThe aerodynamic coefficients for all configurations were based on the wing area without cutout. All pitching-moment data were based on a chord of 1.181 feet ($c'$ at $\\Lambda = 0^\\circ$).\n\nCorrections\n\nThe data have not been corrected for tares caused by the model-support system inasmuch as, with the arrangement used, the tares are believed to be small. Jet-boundary corrections have been applied to the angles of attack, the drag coefficients, and the tail-on pitching-moment coefficients. The corrections were computed by use of reference 2, which unpublished calculations have indicated to be satisfactory for sweep angles up to 45°.\n\nAll forces and moments were corrected for blocking by the method given in reference 3. An increment of longitudinal-force coefficient has been applied to account for the horizontal buoyancy.\n\nRESULTS AND DISCUSSION\n\nThe basic aerodynamic characteristics of the model are presented in figures 6 to 9. The longitudinal-stability parameters are presented in figures 10 to 15. For convenient reference, an outline of the summary figures presenting the results is given as follows:\n\nFigure\n\nVariable sweep:\n(a) Effect of sweep . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 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. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 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. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .", "timestamp": "2026-07-22T05:33:31.682059+00:00"}
{"citation_id": "19930085914", "source_url": "https://ntrs.nasa.gov/api/citations/19930085914/downloads/19930085914.pdf", "page_number": 18, "total_pages": 42, "image_filename": "19930085914_p18.jpg", "text": "Equation for fuselage ordinates:\n\n$$\n\\frac{r}{r_o} = \\left[ 1 - \\left( 1 - \\frac{2x}{l} \\right)^2 \\right]^{\\frac{3}{4}}\n$$\n\nNote: All dimensions given in feet unless otherwise specified.\n\nFineness ratio; $\\frac{l}{2r_o} = 12.5$\n\n<!-- Image (133, 129, 864, 803) -->\n\nFigure 2.— Dimensions of wing and fuselage.\n\nNACA RM A9D25\n17", "timestamp": "2026-07-22T05:33:36.171586+00:00"}
{"citation_id": "19930085912", "source_url": "https://ntrs.nasa.gov/api/citations/19930085912/downloads/19930085912.pdf", "page_number": 18, "total_pages": 36, "image_filename": "19930085912_p18.jpg", "text": "16\nNACA RM No. E9C16\n\nThe inlet-stream velocity at the orifices $V_i$ can be determined from the mass-flow equation if it is assumed that the mass density of the inlet air was approximately equal to that of the free stream:\n\n$$V_i = \\frac{W_a - W_g}{\\rho_0 A_i \\times 32.2} = \\frac{29.5 - 1.30}{32.2 \\times 0.0024 \\times 0.835} = 437 \\text{ ft/sec}$$\n\nThe assumption of a value for the penetration $l$ involves several considerations. First, the penetration as used in the equation of reference 2 was defined as the point at which the temperature has returned to $1^\\circ$ F above the free-stream temperature and, consequently, an allowance must be made for overlap of the jets in order that the temperatures at the simulated engine inlet be uniform. Secondly, the results of reference 2 were obtained in a straight rectangular duct, whereas the inlet ducting used in this investigation was a diffuser of $2^\\circ$ half-angle. Because the jets were introduced at the minimum area of the inlet, the jets would tend to penetrate the air stream more rapidly, not only because the wall is inclined away from the jet but also because of the decreasing air-stream velocity downstream of the jets.\n\nThe requirement for penetrations greater than those predicted by the jet-penetration equation will be partly compensated for by the effect of the diverging duct. Hence, the depth of penetration was assumed equal to that calculated from the equation of reference 2.\n\nIn order to validate this procedure for calculating penetration, the temperature profiles obtained from the three thermocouple rakes mounted in the duct were plotted and the data were analyzed in the manner of reference 2. Typical temperature profiles at the three rakes are shown in figure 14 for a tunnel velocity of 374 feet per second, a plenum-chamber gas temperature of $837^\\circ$ F, and a plenum-chamber gas pressure of 2900 pounds per square foot absolute.\n\nIt has been noted that the results of reference 2 were obtained in a rectangular duct and the penetrations were measured from a horizontal plane, whereas the profile data of figure 14 were obtained in a circular diffuser of $2^\\circ$ half-angle. The penetrations obtained from the profiles of figure 14 were based on the distance measured from the duct wall. The corresponding penetration coefficients along with the predicted values are shown in figure 15 as a function of the mixing-distance - diameter ratio. The experimental values at the simulated engine inlet are about 9 percent higher than those calculated. This increase is caused by the diverging duct.\n\nIn laying out the jet coverage, the assumption was made that the jet divergence angle was approximately $22^\\circ$. A series of similar coverages for various sized orifices was used to facilitate laying out various orifice configurations.", "timestamp": "2026-07-22T05:33:36.424198+00:00"}
{"citation_id": "19930085964", "source_url": "https://ntrs.nasa.gov/api/citations/19930085964/downloads/19930085964.pdf", "page_number": 13, "total_pages": 18, "image_filename": "19930085964_p13.jpg", "text": "12\nNACA RM E9G25\n\nTrailing\nedge\nExciting frequency, cps\nLeading\nedge\n\n(a) 890\n(b) 950\n(c) 1260\n(d) 1820\n(e) 2370\n(f) 2580\n\n(g) 3100\n(h) 3230\n(i) 3330\n(j) 3450\n(k) 4420\n(l) 4650\n\n(m) 4780\n(n) 5020\n(o) 5200\n(p) 5940\n(q) 7010\n(r) 7580\n\n(s) 8240\n(t) 8590\n(u) 9130\n(v) 9740\n\nNACA\n\nFigure 4. - Vibrational modes of stiffened hollow blade B₁. (Solid lines represent node lines on concave side of blade; dashed lines represent node lines on convex side. Exciting frequency in cycles per second is shown below each nodal pattern.)", "timestamp": "2026-07-22T05:33:37.648603+00:00"}
{"citation_id": "19930085906", "source_url": "https://ntrs.nasa.gov/api/citations/19930085906/downloads/19930085906.pdf", "page_number": 20, "total_pages": 23, "image_filename": "19930085906_p20.jpg", "text": "NACA RM E9F20 CONFIDENTIAL 19\n\n[Figure: Graph showing \"Rise in fuel temperature / Maximum rise in fuel temperature • percent\" on the y-axis (0 to 100) versus \"Time, min\" on the x-axis (0 to 6). A curve starts at (0,0), rises steeply, and asymptotically approaches 100%. The NACA logo is in the bottom right corner of the graph area.]\n\nFigure 7. - Variation of ratio between rise in fuel temperature to maximum rise in fuel temperature as function of heating time. (All data are within ± 6 percent of curve.)\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:33:40.255357+00:00"}
{"citation_id": "19930085962", "source_url": "https://ntrs.nasa.gov/api/citations/19930085962/downloads/19930085962.pdf", "page_number": 13, "total_pages": 51, "image_filename": "19930085962_p13.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:33:41.659180+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 7, "total_pages": 114, "image_filename": "19930086061_p7.jpg", "text": "NACA RM L9J07\n\n$c_{n}$ section normal-force coefficient $\\left(\\int_{0}^{1.0} P \\, d\\left(\\frac{x}{c}\\right)\\right)$\n\n$c_{l}$ section lift coefficient $\\left(c_{n} \\cos \\alpha\\right)$\n\n$c_{m}$ section pitching-moment coefficient about Y-axis $\\left(c_{n} \\cdot \\bar{x}_{c} / 4\\right)$\n\n$c_{l_{\\alpha}}$ section lift-curve slope at $c_{l}=0$, per degree\n\n$P$ pressure coefficient $\\left(\\frac{p-p_{o}}{q}\\right)$\n\n$C_{N}$ wing normal-force coefficient $\\left(\\frac{1}{2} \\int_{-1.0}^{1.0} \\frac{c_{n} c}{c_{av}} \\, d\\left(\\frac{y}{b / 2}\\right)\\right)$\n\n$C_{L}$ wing lift coefficient $\\left(C_{N} \\cos \\alpha\\right)$\n\n$C_{L_{\\max }}$ maximum wing lift coefficient\n\n$C_{L_{\\alpha}}$ wing lift-curve slope, per degree\n\n$\\alpha_{C_{L_{\\max }}}$ angle of attack for $C_{L_{\\max }}$, degrees\n\n$C_{m}$ wing pitching-moment coefficient $\\left(\\frac{c_{av}}{2 \\bar{c}} \\int_{-1.0}^{1.0} c_{m}\\left(\\frac{c}{c_{av}}\\right)^{2} \\, d\\left(\\frac{y}{b / 2}\\right)\\right)$\n\n$C_{l}$ wing rolling-moment coefficient $\\left(\\frac{1}{4} \\int_{-1.0}^{1.0} \\frac{c_{n} c}{c_{av}} \\frac{y}{b / 2} \\, d\\left(\\frac{y}{b / 2}\\right)\\right)$\n\n$p$ local static pressure, pounds per square foot\n\n$p_{o}$ free-stream static pressure, pounds per square foot", "timestamp": "2026-07-22T05:33:45.803686+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 34, "total_pages": 96, "image_filename": "19930085880_p34.jpg", "text": "32\nNACA RM No. L9C03\n\n[Figure: A graph plotting Lift against Speed. The y-axis is labeled \"Lift, lb\" and ranges from 0 to .35. The x-axis is labeled \"Speed, fps\" and ranges from 0 to 35. Inside the graph area, there is a small diagram of an inverted triangle above a horizontal line. Several lines originate from the origin (0,0) and extend upwards to the right, each labeled with a \"Wetted area (sq ft)\" value on the right side: 0, .05, .10, .15, .20, .25, .30. The NACA logo is present near the bottom right of the graph area.]\n\n(b) $\\tau = 12^\\circ$.\n\nFigure 13.- Continued.", "timestamp": "2026-07-22T05:33:49.990870+00:00"}
{"citation_id": "19930085889", "source_url": "https://ntrs.nasa.gov/api/citations/19930085889/downloads/19930085889.pdf", "page_number": 30, "total_pages": 37, "image_filename": "19930085889_p30.jpg", "text": "NACA RM L9F14\n29\n\nCONFIDENTIAL\n\n$$\n\\begin{array}{c|c}\n\\text{Sweepback, } \\Lambda, \\text{deg} & \\\\\n\\hline\n\\circ & 3.6 \\\\\n\\square & 32.6 \\\\\n\\diamond & 46.7 \\\\\n\\end{array}\n$$\n\n[Figure: Three plots of aerodynamic coefficients ($C_{Y\\psi}$, $C_{n\\psi}$, $C_{l\\psi}$) versus Lift coefficient ($C_L$). The plots show data points for three different sweepback angles (3.6, 32.6, and 46.7 degrees) represented by circles, squares, and diamonds respectively. The top plot shows $C_{Y\\psi}$ ranging from -0.004 to 0.008. The middle plot shows $C_{n\\psi}$ ranging from -0.004 to 0.004. The bottom plot shows $C_{l\\psi}$ ranging from -0.002 to 0.006. The x-axis for all plots is Lift coefficient, $C_L$, ranging from -4 to 12. The plots contain \"CONFIDENTIAL\" and \"NACA\" watermarks.]\n\nCONFIDENTIAL\nNACA\n\nFigure 11.- Variation of $C_{Y\\psi}$, $C_{n\\psi}$, and $C_{l\\psi}$ with lift coefficient for the wings tested with a fuselage.", "timestamp": "2026-07-22T05:33:50.564605+00:00"}
{"citation_id": "19930085542", "source_url": "https://ntrs.nasa.gov/api/citations/19930085542/downloads/19930085542.pdf", "page_number": 36, "total_pages": 46, "image_filename": "19930085542_p36.jpg", "text": "```markdown\n34\nNACA RM No. L8L29\n\n<!-- Image (122, 109, 831, 874) -->\n\nFigure 15.- Effect of fins on a triangular wing of NACA 0012 profile on $C_{Y\\psi}$, $C_{n\\psi}$, and $C_{l\\psi}$. $A_c/4 = 52.2^\\circ$.\n```", "timestamp": "2026-07-22T05:33:51.032010+00:00"}
{"citation_id": "19930085934", "source_url": "https://ntrs.nasa.gov/api/citations/19930085934/downloads/19930085934.pdf", "page_number": 14, "total_pages": 23, "image_filename": "19930085934_p14.jpg", "text": "NACA RM E9G12\n\nEnthalpy of saturated vapor per pound of dry air at outlet\n\n$h_{g,a,2} = h_{g,2} \\cdot q_2$\n\n$= (1108.6)(0.04005)$\n\n$= 44.40 \\text{ (Btu/lb)}$\n\nEnthalpy of dry air at outlet\n\n$h_{d,2} \\text{ at } T_{t,2} = 40.25 \\text{ (Btu/lb)}$\n\nThe enthalpy of mixture per pound of dry air at the outlet equals the enthalpy of saturated liquid per pound of dry air plus the enthalpy of saturated vapor per pound of dry air plus the enthalpy of dry air.\n\n$h_{m,a,2} = h_{f,a,2} + h_{g,a,2} + h_{d,2}$\n\n$= 1.534 + 44.40 + 40.25$\n\n$= 86.18 \\text{ (Btu/lb)}$\n\n(3) Enthalpy change from inlet to outlet.\n\n$\\Delta h_{m,a} = h_{m,a,2} - h_{m,a,1}$\n\n$= 86.18 - 45.25$\n\n$= 40.93 \\text{ (Btu/lb)}$\n\nIf the vapor at the outlet is in the superheat region, the entropy and the enthalpy are computed for the assumed temperature in the same manner as the enthalpy or the actual work in section IV.\n\nIV. Actual Work Determined from Total Temperature at Outlet\n\nTemperature at outlet = $264^\\circ \\text{ F}$\n\nFor this temperature, the vapor contained in the air is in the superheated condition. The pressure exerted by the vapor in the mixture will therefore be equal to that exerted at saturation. The following method is used to determine the temperature at which saturation occurs. By assuming values of saturation temperature $t$", "timestamp": "2026-07-22T05:33:52.179881+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 1, "total_pages": 98, "image_filename": "19930086073_p1.jpg", "text": "NACA RM A9H04\n\nNACA\n\nRESEARCH MEMORANDUM\n\nAN INVESTIGATION AT LOW SPEED OF A LARGE-SCALE TRIANGULAR\nWING OF ASPECT RATIO TWO.- III. CHARACTERISTICS OF\nWING WITH BODY AND VERTICAL TAIL\n\nBy Adrien E. Anderson\n\nAmes Aeronautical Laboratory\nMoffett Field, Calif.\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\nWASHINGTON\n\nOctober 14, 1949\nDeclassified June 11, 1953", "timestamp": "2026-07-22T05:34:00.548045+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 66, "total_pages": 99, "image_filename": "19930082511_p66.jpg", "text": "64\nNACA TN No. 1826\n\nIn the first term in the braces, the expression in the brackets is the first and only Fourier coefficient. Changing $\\xi$ to $\\beta$ and inserting the expression into the inner integral of equation (8) of reference 13 gives\n\n$$\n\\lim_{\\sigma \\to 0} \\frac{1}{q} \\int_{-\\infty}^{\\infty} \\left( \\frac{\\beta \\sigma}{\\sqrt{1 + \\beta^2 \\sigma^2}} - \\frac{\\beta}{\\sqrt{1 + \\beta^2}} \\right) \\cos q(\\beta - \\xi) \\, d\\beta\n$$\n\nwhich, if integrated by parts, reduces to\n\n$$\n\\lim_{\\sigma \\to 0} \\frac{1}{q} \\int_{-\\infty}^{\\infty} \\left[ \\frac{1}{(1 + \\beta^2)^{3/2}} - \\frac{\\sigma}{(1 + \\beta^2 \\sigma^2)^{3/2}} \\right] \\sin q(\\beta - \\xi) \\, d\\beta\n$$\n\nThe contribution of the first term in the brackets is\n\n$$\n\\frac{1}{q} \\int_{-\\infty}^{\\infty} \\frac{\\sin q(\\beta - \\xi) \\, d\\beta}{(1 + \\beta^2)^{3/2}} = -2K_1(q) \\sin q\\xi\n$$\n\n(See reference 15, p. 52.) The fact that the contribution of the second term in the brackets is zero follows immediately, upon performing the change of variable $p = \\beta \\sigma$, from the Riemann-Lebesgue lemma (reference 11, p. 172).\n\nThe third term in the braces of equation (C1) is converted as follows:\n\n$$\n- \\frac{\\xi \\sin \\theta}{\\xi^2 + \\sin^2 \\theta} \\left[ \\frac{1}{\\sqrt{1 + \\xi^2}} - \\frac{\\cos^2 \\theta}{(1 + \\xi^2)^{3/2}} \\right] = - \\frac{\\xi \\sin \\theta}{(1 + \\xi^2)^{3/2}}\n$$\n\nso that again the first and only Fourier coefficient is given directly. Inserting it into the inner integral of equation (8) of reference 13 and integrating by parts gives\n\n$$\n- \\int_{-\\infty}^{\\infty} \\frac{\\beta}{(1 + \\beta^2)^{3/2}} \\cos q(\\beta - \\xi) \\, d\\beta = \\int_{-\\infty}^{\\infty} \\frac{q \\sin q(\\beta - \\xi)}{\\sqrt{1 + \\beta^2}} \\, d\\beta\n$$\n\n$$\n= -2qK_0(q) \\sin q\\xi\n$$", "timestamp": "2026-07-22T05:34:00.749949+00:00"}
{"citation_id": "19930085972", "source_url": "https://ntrs.nasa.gov/api/citations/19930085972/downloads/19930085972.pdf", "page_number": 7, "total_pages": 46, "image_filename": "19930085972_p7.jpg", "text": "```markdown\nNACA RM L9B18\n5\n\nBasic Configurations\n\nEffect of sweep angle, wing without cutouts.- The destabilizing\nmovement of the neutral-point location as the wing sweep angle was\ndecreased amounted to about 2 percent of the chord ($c^l$ at $\\Lambda = 0^\\circ$) per\ndegree change in sweep angle (fig. 10). Inasmuch as the parameters\naffecting the tail contribution to stability show relatively minor\nvariations with sweep angle, it would appear that the shift in neutral\npoint is primarily associated with the geometric movement of the wing-\naerodynamic-center position as the wing is rotated. Up to about $35^\\circ$ of\nsweep angle the experimental rate of variation of neutral point with\nsweep angle is in good agreement with that estimated from the simple\ngeometric consideration that the centroid of lift on each wing panel\nacts at 25 percent of the mean aerodynamic chord of the wing at each\nsweep angle.\n\nFor the $30^\\circ$ and $45^\\circ$ sweep configurations, an unstable variation of\npitching-moment coefficient with lift coefficient at the high lift\ncoefficients is indicated (figs. 8 and 9); whereas, for $0^\\circ$ and $15^\\circ$ configu-\nrations, stable pitching-moment characteristics were obtained in the\nvicinity of the maximum lift coefficient.\n\nEffect of sweep angle, wing with faired cutout.- If a cutout is\nallowed to develop at the juncture of the wing-root trailing edge and\nthe fuselage as the wing sweep angle is decreased (figs. 1 and 4),\nsignificant changes in the stability characteristics exhibited by the\nmodel can occur. It was anticipated that the cutout would move the wing-\naerodynamic-center position forward somewhat but that the downwash field\nin the vicinity of the horizontal tail would be changed in such a manner\nthat the over-all stability of the model would be increased. The extent\nto which these effects were manifested at various sweep angles is indi-\ncated in figures 11 to 13.\n\nA study of these data indicates that the faired cutout afforded\nsomewhat greater stability than the configurations having no cutout but\nthe over-all effect on stability is small compared to the large changes\nproduced by the geometric movement of the wing aerodynamic center. The\neffects on the stability parameters were greatest at $0^\\circ$ sweep and decreased\nas the sweep angle was increased. The tail contribution to stability at\nsweep angles of $0^\\circ$ and $15^\\circ$ was increased considerably because of favorable\nflow changes at the tail but this beneficial effect was partially canceled\nby the forward movement of the wing aerodynamic center caused by the\ncutouts.\n\nThe effect of the faired cutout on the neutral-point position for all\nsweep angles is summarized in figure 15.\n```", "timestamp": "2026-07-22T05:34:02.806259+00:00"}
{"citation_id": "19930085914", "source_url": "https://ntrs.nasa.gov/api/citations/19930085914/downloads/19930085914.pdf", "page_number": 19, "total_pages": 42, "image_filename": "19930085914_p19.jpg", "text": "Note: All dimensions given in feet unless otherwise specified.\n\nTypical section parallel to plane of symmetry\n\nAll sections have NACA a=1.0 mean-camber lines and 64A005 thickness distributions. (See Table I for section coordinates.)\n\nSpanwise camber distribution\n\n| Station | Percent semi-span | Camber $y_c/c$ |\n| :--- | :--- | :--- |\n| $c_0$ | 0 | 0 |\n| $c_1$ | 20 | .0082 |\n| $c_2$ | 40 | .0108 |\n| $c_3$ | 60 | .0117 |\n| $c_4$ | 80 | .0115 |\n| $c_5$ | 100 | .0114 |\n\n[Graph: Plot of Spanwise station, percent semi-span vs Angle of twist, $a_1$, deg. Curves labeled \"Model static twist\" and \"Theoretical twist ($C_L=.25, M=1.5$)\"]\n\nFigure 3.— Plan form of right half of wing showing spanwise variation of camber and twist and location of sections for which coordinates have been calculated.\n\n18\nNACA RM A9D25", "timestamp": "2026-07-22T05:34:06.690799+00:00"}
{"citation_id": "19930085906", "source_url": "https://ntrs.nasa.gov/api/citations/19930085906/downloads/19930085906.pdf", "page_number": 21, "total_pages": 23, "image_filename": "19930085906_p21.jpg", "text": "20\nCONFIDENTIAL\nNACA RM E9F20\n\n[Figure: A line graph with a shaded region between two lines. The y-axis is labeled \"Time required to reach fuel temperature of 200° F, min\" with values 0, 1, 2, 3. The x-axis is labeled \"Distance between preheater and flame holder, in.\" with values 0, 4, 8, 12. The NACA logo is in the bottom right corner of the graph.]\n\nFigure 8. - Effect of preheater position on time required to attain fuel temperature of 200° F. Initial temperature, 65° ±7° F.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:34:09.413547+00:00"}
{"citation_id": "19930085964", "source_url": "https://ntrs.nasa.gov/api/citations/19930085964/downloads/19930085964.pdf", "page_number": 14, "total_pages": 18, "image_filename": "19930085964_p14.jpg", "text": "NACA RM E9G25\n13\n\n1171\n\nTrailing\nedge\n\nExciting frequency, cps\n\nLeading\nedge\n\n<!-- Image (143, 131, 864, 853) -->\n\n(a) 880\n(b) 955\n(c) 1220\n(d) 1840\n(e) 2330\n(f) 2580\n\n(g) 3110\n(h) 3310\n(i) 3360\n(j) 4300\n(k) 4620\n(l) 4690\n\n(m) 4780\n(n) 5030\n(o) 5210\n(p) 5780\n(q) 6440\n(r) 6720\n\n(s) 7030\n(t) 7290\n(u) 8310\n(v) 8670\n(w) 9810\n\n40-1574\n\nFigure 5. - Vibrational modes of stiffened hollow blade B2. (Solid lines represent node\nlines on concave side of blade; dashed lines represent node lines on convex side.\nExciting frequency in cycles per second is shown below each nodal pattern.)\n\nNACA", "timestamp": "2026-07-22T05:34:10.243395+00:00"}
{"citation_id": "19930085962", "source_url": "https://ntrs.nasa.gov/api/citations/19930085962/downloads/19930085962.pdf", "page_number": 14, "total_pages": 51, "image_filename": "19930085962_p14.jpg", "text": "NACA RM A52D05\nCONFIDENTIAL\n\n12\n10\n8\n6\n4\n2\n0\n-2\n-4\n-6\n-8\n-20 -16 -12 -8 -4 0 4 8 12 16\nAngle of attack, $\\alpha$, deg\n\nLift coefficient, $C_L$\n\n$\\delta_e$\n(deg)\n$\\circ$ 0\n$\\square$ 2\n$\\diamond$ 4\n$\\triangle$ 6\n$\\nabla$ 10\n$\\blacktriangledown$ 20\n$\\blacktriangleleft$ 30\n\n.16 .12 .08 .04 0 -.04 -.08 -.12 -.16 -.20 -.24 -.28\nPitching-moment coefficient, $C_m$\n\nNACA\n\n(a) $C_L$ vs $\\alpha$, $C_L$ vs $C_m$.\n\nFigure 3. — The effect of elevator deflection on the aerodynamic characteristics of the tail at a Mach number of 0.20.\n\nCONFIDENTIAL\n13", "timestamp": "2026-07-22T05:34:12.354564+00:00"}
{"citation_id": "19930085912", "source_url": "https://ntrs.nasa.gov/api/citations/19930085912/downloads/19930085912.pdf", "page_number": 19, "total_pages": 36, "image_filename": "19930085912_p19.jpg", "text": "NACA RM No. E9C16\n\nNumerous orifice configurations were laid out in accordance with the design criterions. The configuration selected for trial provided the best over-all coverage with a minimum of jet overlap and represented a compromise between the two design conditions. The configuration chosen and the area protected by each jet are shown in figure 3 and the excellent results obtained with this configuration have been discussed in the text.\n\nIn order to check the design criterions and to be certain that the chosen configuration was the best, several other orifice configurations were experimentally investigated. A configuration consisting of six 3/4-inch-diameter jets equally spaced was investigated. Because of the small number of holes, very poor mixing was obtained. The screen and the accessory housing could be protected only at excessive values of bleedback. Another configuration consisting of nine holes with three 3/4-inch-diameter holes spaced at $120^\\circ$ intervals and two 33/64-inch-diameter holes equally spaced between the larger holes was also investigated. As in the previous configuration, poor mixing was obtained. For both of these experimental configurations, the jet layouts also showed that poor coverage would be obtained.\n\n---\n\n**REFERENCES**\n\n1. Callaghan, Edmund E., Ruggeri, Robert S., and Krebs, Richard P.: Experimental Investigation of Hot-Gas Bleedback for Ice Protection of Turbojet Engines. I- Nacelle with Offset Air Inlet. NACA RM No. E8D13, 1948.\n\n2. Callaghan, Edmund E., and Ruggeri, Robert S.: Investigation of the Penetration of an Air Jet Directed Perpendicularly to an Air Stream. NACA TN No. 1615, 1948.", "timestamp": "2026-07-22T05:34:14.426238+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 8, "total_pages": 114, "image_filename": "19930086061_p8.jpg", "text": "4\nNACA RM L9J07\n\nq\nreference dynamic pressure at pitot-tube location (fig. 1.),\npounds per square foot $\\left(\\frac{\\rho V^2}{2}\\right)$\n\n$q_l$\nlocal dynamic pressure, pounds per square foot $\\left(\\frac{\\rho V_l^2}{2}\\right)$\n\nV\nvelocity at pitot-tube location (fig. 1), feet per second\n\n$V_l$\nlocal velocity, feet per second\n\n$\\rho$\nmass density of air, slugs per cubic foot\n\n$\\nu$\nkinematic viscosity, square feet per second\n\nS\nwing area, square feet\n\nb\nwing span, feet\n\nc\nlocal wing chord, feet\n\n$c_{av}$\naverage wing chord, feet (S/b)\n\n$\\bar{c}$\nmean aerodynamic chord, M.A.C., feet $\\left(\\frac{2}{S} \\int_0^{b/2} c^2 dy\\right)$\n\nA\naspect ratio, $\\left(b^2/S\\right)$\n\n$\\alpha$\nangle of attack, degrees\n\n$\\psi$\nangle of yaw, degrees\n\nR\nReynolds number $\\left(\\frac{V_l \\bar{c}}{\\nu}\\right)$\n\nx\ndistance along chord from leading edge, feet", "timestamp": "2026-07-22T05:34:17.923136+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 35, "total_pages": 96, "image_filename": "19930085880_p35.jpg", "text": "NACA RM No. L9C03\n33\n\nLift, lb\nWetted area\n(sq ft)\n.35\n.30\n.25\n.20\n.15\n.10\n.05\n0\n0 10 15 20 25 30 35\nSpeed, fps\nNACA\n(c) $\\tau = 16^\\circ$.\nFigure 13.- Continued.", "timestamp": "2026-07-22T05:34:20.660003+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 2, "total_pages": 98, "image_filename": "19930086073_p2.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:34:22.794415+00:00"}
{"citation_id": "19930085542", "source_url": "https://ntrs.nasa.gov/api/citations/19930085542/downloads/19930085542.pdf", "page_number": 37, "total_pages": 46, "image_filename": "19930085542_p37.jpg", "text": "NACA RM No. L8L29\n35\n\n$C_{Y_p}$\n.4\n0\n-.4\n\nModel No. $A_{WING}$ $A_{FIN}$\n2 2.31 -\n5 2.31 .77\n6 2.31 1.15\n\n.2\n$C_{n_p}$ 0\n-.2\n\n0\n$C_{l_p}$ -.2\n-.4\n\n-.2 0 .2 .4 .6 .8 1.0\nLift coefficient, $C_L$\n\nFigure 16.- Effect of fins on a triangular wing of NACA 0012 profile\non $C_{Y_p}$, $C_{n_p}$, and $C_{l_p}$. $\\Lambda_{c/4} = 52.2^\\circ$.", "timestamp": "2026-07-22T05:34:22.931938+00:00"}
{"citation_id": "19930085889", "source_url": "https://ntrs.nasa.gov/api/citations/19930085889/downloads/19930085889.pdf", "page_number": 31, "total_pages": 37, "image_filename": "19930085889_p31.jpg", "text": "30\nNACA RM L9F14\n\nCONFIDENTIAL\n\n.004\n$C_{Y\\psi}$ 0\n-.004\n\n.004\n.002\n$C_{n\\psi}$ 0\n-.002\n\n.002\n0\n$C_{l\\psi}$\n-.002\n-.004\n\n$\\circ$ $46.7^\\circ \\Lambda$ wing alone\n$\\square$ $46.7^\\circ \\Lambda$ wing + fuselage\n\n-.2 0 .2 .4 .6 .8 1.0 1.2\nLift coefficient, $C_L$\n\nCONFIDENTIAL\nNACA\n\nFigure 12.- Variation of $C_{Y\\psi}$, $C_{n\\psi}$, and $C_{l\\psi}$ with lift coefficient for the\n$46.7^\\circ$ sweptback wing tested alone and in combination with a fuselage.", "timestamp": "2026-07-22T05:34:23.222480+00:00"}
{"citation_id": "19930085962", "source_url": "https://ntrs.nasa.gov/api/citations/19930085962/downloads/19930085962.pdf", "page_number": 15, "total_pages": 51, "image_filename": "19930085962_p15.jpg", "text": "14\nCONFIDENTIAL\nNACA RM A9E05\n\n<!-- Image (103, 163, 863, 821) -->\n\nFigure 3. — Concluded.\n(b) $C_L$ vs $C_D$.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:34:30.891860+00:00"}
{"citation_id": "19930085934", "source_url": "https://ntrs.nasa.gov/api/citations/19930085934/downloads/19930085934.pdf", "page_number": 15, "total_pages": 23, "image_filename": "19930085934_p15.jpg", "text": "14\nNACA RM E9G12\n\nand using the corresponding values of specific volume $v_g$ found\nin table 1 of reference 1, a value of specific humidity equal\nto the total water-air ratio at the outlet is obtained.\n\nFrom section II, item (3) the total water-air ratio is 0.06022.\nAssume the temperature $t$ of saturation to be $121.3^\\circ$ F. The pres-\nsure $P_{g,2}$ corresponding to $121.3^\\circ$ F is 1.755 pounds per square inch\nin the steam tables (table 1, reference 1).\n\n(1) Pressure of superheated vapor. -\n$$P_{d,2} = P_2 - P_{g,2}$$\n$$= 19.998 - 1.755$$\n$$= 18.24 \\text{ (lb/sq in.)}$$\n\nThe volume occupied by 1 pound of vapor at saturation $v_g$ for\n$121.3^\\circ$ F is 196.5 cubic feet per pound from table 1, reference 1.\n\nThe weight of air that would occupy the same volume as $v_g$ at\n$P_{d,2}$ is\n$$W = \\frac{P_{d,2} v_g}{R(t + 459.7)}$$\n$$= \\frac{(144)(18.24)(196.5)}{(53.35)(581.0)}$$\n$$= 16.65 \\text{ (lb)}$$\n\nWater-air ratio\n$$\\frac{1}{W} = \\frac{1}{16.65} = 0.0601$$\n\nInasmuch as the actual $q_{m,2}$ is 0.06022 pound water per pound\nair, the assumed temperature is acceptable; the pressure exerted by\nthe vapor in the superheat region $P_{s,2}$ is therefore 1.755 pounds\nper square inch.\n\n(2) Enthalpy of mixture. -\nEnthalpy of vapor $h_g$ at $264^\\circ$ F = 1179 (Btu/lb) (reference 1, fig. 10)", "timestamp": "2026-07-22T05:34:31.095394+00:00"}