PDF/ocr_dataset_1038.jsonl

{"citation_id": "19930085906", "source_url": "https://ntrs.nasa.gov/api/citations/19930085906/downloads/19930085906.pdf", "page_number": 13, "total_pages": 23, "image_filename": "19930085906_p13.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:28:29.125556+00:00"}
{"citation_id": "19930085934", "source_url": "https://ntrs.nasa.gov/api/citations/19930085934/downloads/19930085934.pdf", "page_number": 7, "total_pages": 23, "image_filename": "19930085934_p7.jpg", "text": "6\nNACA RM E9G12\n\nWater injected, lb/sec . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 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. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .", "timestamp": "2026-07-22T05:28:29.867918+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 26, "total_pages": 96, "image_filename": "19930085880_p26.jpg", "text": "24\nNACA RM No. I9C03\n\nResistance, lb\n.30\n.25\n.20\n.15\n.10\n.05\n0\n\nSpeed, fps\n0 10 15 20 25 30 35\n\nWetted area\n(sq ft)\n0\n.05\n.10\n.15\n.20\n.25\n.30\n\nNACA\n\n(c) $\\tau = 12^\\circ$.\n\nFigure 11.- Continued.", "timestamp": "2026-07-22T05:28:30.266228+00:00"}
{"citation_id": "19930085889", "source_url": "https://ntrs.nasa.gov/api/citations/19930085889/downloads/19930085889.pdf", "page_number": 22, "total_pages": 37, "image_filename": "19930085889_p22.jpg", "text": "CONFIDENTIAL\n\nNACA RM L9F14\n\n[Figure: A model aircraft with a sweptback wing and fuselage mounted in a wind tunnel, viewed from below. The model is suspended by a support structure. In the lower right corner of the image, there is a label with the NACA logo and the number L-57655.]\n\nFigure 4.- The $46.7^\\circ$ sweptback wing with fuselage mounted in the rolling-flow section of the stability tunnel.\n\nCONFIDENTIAL\n\n21", "timestamp": "2026-07-22T05:28:30.460991+00:00"}
{"citation_id": "19930085519", "source_url": "https://ntrs.nasa.gov/api/citations/19930085519/downloads/19930085519.pdf", "page_number": 45, "total_pages": 46, "image_filename": "19930085519_p45.jpg", "text": "44\nNACA RM No. I8K19\n\n<!-- Image (129, 105, 830, 732) -->\n\n(a) Flap neutral - flap-slot sealed and faired.\nFigure 22.- Variation of rolling-moment coefficient and yawing-moment coefficient with angle of attack for various deflections of a 49-percent-span by 20-percent-chord plain sealed aileron on the 42° sweptback wing.", "timestamp": "2026-07-22T05:28:30.769200+00:00"}
{"citation_id": "19930085912", "source_url": "https://ntrs.nasa.gov/api/citations/19930085912/downloads/19930085912.pdf", "page_number": 12, "total_pages": 36, "image_filename": "19930085912_p12.jpg", "text": "10\nNACA RM No. E9C16\n\nheat addition can be written as $q_0 \\left( 1 - \\frac{460+T_0}{460+T_{av}} \\right)$. The $\\eta_{calc}$ curve of figure 10 was calculated from the data of figure 5, assuming a tunnel total temperature of $0^\\circ$ F, a plenum-chamber gas temperature of $1000^\\circ$ F, and perfect mixing. The lower curve in figure 10 was computed by subtracting the quantity $\\left( 1 - \\frac{460+T_0}{460+T_{av}} \\right)$ from the upper curve. The experimental data also plotted in figure 10 for a gas temperature of $1000^\\circ$ F show good agreement with the computed curve.\n\nInlet-lip temperature distribution. - The maximum inlet-lip temperatures were obtained at the highest value of bleedback and plenum-chamber gas temperature utilized in the investigation. The maximum lip temperatures encountered are shown in figure 11 and were obtained at a bleedback of 8.65 percent and a plenum-chamber gas temperature of $1000^\\circ$ F. The lip temperatures decreased with decreasing bleedback or decreasing plenum-chamber gas temperature.\n\nDuct-skin temperature. - The highest measured duct-skin temperature was $185^\\circ$ F and was obtained at a point on the skin adjacent to the hot-gas ducts, which were located in the inlet wall. The model-air total temperature was approximately $90^\\circ$ F and the plenum-chamber gas temperature, $1000^\\circ$ F. For an average air-temperature rise of $40^\\circ$ F, the skin temperature did not exceed $100^\\circ$ F. The temperature of the skin, except adjacent to the hot-gas ducting, did not exceed the model-air total temperature.\n\nIcing with Bleedback\n\nIn the analysis of the icing data, the pressure-drop coefficients $\\Delta p/q$ across the screen were computed for each icing run. The screen was considered iced when the value of $\\Delta p/q$ approached 1.5 times the value for the screen at the beginning of each run. The experimental bleedback and plenum-chamber gas temperatures corresponding to this criterion are shown in figure 12 for tunnel velocities of 200, 280, 360, and 410 feet per second at an angle of attack of $0^\\circ$ and for 200 and 280 feet per second at an angle of attack of $8^\\circ$. No ice accretions were observed on the accessory housing nor the nacelle lip when the inlet screen was iced. A very slight ice formation encountered from 4 to 10 inches behind the orifices around the entire periphery of the inlet was apparently caused by the poor mixing obtained immediately behind the orifices.", "timestamp": "2026-07-22T05:28:34.588123+00:00"}
{"citation_id": "19930082617", "source_url": "https://ntrs.nasa.gov/api/citations/19930082617/downloads/19930082617.pdf", "page_number": 44, "total_pages": 58, "image_filename": "19930082617_p44.jpg", "text": "NACA TN 1962\n43\n\n[Figure: Side view of cylinder 72 after buckling. Far side or stringer 11.]\n\nCYL 72\nNACA", "timestamp": "2026-07-22T05:28:43.475235+00:00"}
{"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 46, "total_pages": 78, "image_filename": "19930082618_p46.jpg", "text": ".028\n.024\nSection drag coefficient, $c_d$\n.020\n.016\n.012\n.008\n.004\n0\n-.8 -.4 0 .4 .8 1.2 1.6\nSection lift coefficient, $c_l$\n\nR\nO 0.7 x $10^6$\n□ 1.0\n◇ 1.5\n△ 2.0\nFlagged symbols denote\nstandard roughness\n\n.028\n.024\nSection drag coefficient, $c_d$\n.020\n.016\n.012\n.008\n.004\n0\n-1.2 -.8 -.4 0 .4 .8 1.2 1.6\nSection lift coefficient, $c_l$\n\nR\n▽ 3.0 x $10^6$\n◁ 6.0\n◁ 9.0\nFlagged symbols denote\nstandard roughness\n\nNACA\n\nMoment coefficient, $c_m$\n0\n-.2\n-.4\n-.6\n-.8\n-1.0\n-1.2\n-1.4\n-1.6\n-1.8\n-2.0\n-.8 -.4 0 .4 .8 1.2 1.6\nSection lift coefficient, $c_l$\n\nR\nO 0.7 x $10^6$\n□ 1.0\n◇ 1.5\n△ 2.0\n▽ 3.0\n◁ 6.0\n◁ 9.0\n\na.c. position\nx/c y/c\n.274 .016\n.271 .022\n.273 -.081\n.273 -.069\n.284 -.045\n.284 -.093\n.282 -.056\n\n(c) Section drag characteristics and section pitching-moment characteristics about the aerodynamic center of the plain NACA 63$_2$-415 airfoil section.\n\nFigure 8.- Concluded.\n\n441\nNACA TN 1945", "timestamp": "2026-07-22T05:28:45.376891+00:00"}
{"citation_id": "19930082542", "source_url": "https://ntrs.nasa.gov/api/citations/19930082542/downloads/19930082542.pdf", "page_number": 42, "total_pages": 53, "image_filename": "19930082542_p42.jpg", "text": "```markdown\nNACA TN No. 1867\n47\n\nProperties at room temperature\n\nBrinell hardness\n300\n200\n100\nBrinell hardness\n\nStress, psi\n140,000\n120,000\n100,000\n80,000\n60,000\n40,000\nTensile strength\n0.02-percent-offset yield strength\n\nElongation, percent\n40\n20\n0\nElongation\n\nRupture properties at 1200° F\n\nStress, psi\n50,000\n40,000\n30,000\n20,000\nRupture strength\n100 hr\n1000 hr\n\nElongation, percent\n40\n20\n0\nRupture elongation\n100 hr\n\n| Aging tem- perature, °F | Original | 1400 | 1500 | 1600 | 1700 | 1800 |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| Aging time, hr | | 24 | | | | |\n\nNACA\n\nOriginal treatment\no As-rolled\n□ 2050° F water-quenched 2 hr\n△ 2200° F water-quenched 1 hr\n\nFigure 7.- Effect of aging at various temperatures on bar stock of low-carbon N-155 alloy.\n```", "timestamp": "2026-07-22T05:28:47.393970+00:00"}
{"citation_id": "19930082914", "source_url": "https://ntrs.nasa.gov/api/citations/19930082914/downloads/19930082914.pdf", "page_number": 54, "total_pages": 66, "image_filename": "19930082914_p54.jpg", "text": "NACA TN No. 1857\n53\n\n[Figure: A black and white image showing a pattern of wavy, horizontal lines. The lines are more distorted and irregular in the upper portion of the image and become progressively straighter and more uniform towards the bottom.]\n\n(d) $4\\frac{1}{4}$ to $5\\frac{3}{4}$ inches from nozzle.\nFigure 11.- Continued.", "timestamp": "2026-07-22T05:28:48.494291+00:00"}
{"citation_id": "19930085881", "source_url": "https://ntrs.nasa.gov/api/citations/19930085881/downloads/19930085881.pdf", "page_number": 21, "total_pages": 31, "image_filename": "19930085881_p21.jpg", "text": "NACA RM L9D12\n19\n\nCONFIDENTIAL\n\n.12\n.08\n.04\n0\n$C_D$\n\n.16\n.12\n.08\n.04\n0\n$pb/2V$\n\n.6 .8 1.0 1.2 1.4 1.6 1.8 2.0\nM\n\n| Model | $\\delta_a$ (deg) | $i_w$ (deg) |\n| :--- | :---: | :---: |\n| 50d | 5.2 | 0 |\n| 50e | 5.6 | -0.2 |\n| 50f | 6.0 | 0 |\n\nNACA\n\n(a) NACA 65A009 airfoil section.\nFigure 5.— Experimental results. $\\Lambda = 0^\\circ$.\nCONFIDENTIAL", "timestamp": "2026-07-22T05:28:48.583289+00:00"}
{"citation_id": "19930085542", "source_url": "https://ntrs.nasa.gov/api/citations/19930085542/downloads/19930085542.pdf", "page_number": 28, "total_pages": 46, "image_filename": "19930085542_p28.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:28:55.748720+00:00"}
{"citation_id": "19930085879", "source_url": "https://ntrs.nasa.gov/api/citations/19930085879/downloads/19930085879.pdf", "page_number": 29, "total_pages": 29, "image_filename": "19930085879_p29.jpg", "text": "NACA - Langley Field, Va.\n\nDeceleration, g\n\n100\n50\n40\n30\n20\n15\n10\n\n50\n40\n30\n20\n15\n10\n\nActual deceleration vs time after ejection\nAverage deceleration experienced vs\nlength of time experienced\n\nProbable human tolerance\n\nMach 2, Sea level\n\nMach 2, 30000 ft\n\nMach 2, 45000 ft\n\n[Figure: Graph plotting Deceleration (g) vs Time (sec) with multiple curves representing different Mach numbers and altitudes, including a \"Probable human tolerance\" curve and annotations for actual and average deceleration.]\n\nTime, sec\n\n.05 .01 .2 .5 1.0 2.0 5.0 10.0 20.0 50.0 100.0\n\nNACA\n\nFigure 15.— Calculated decelerations due to drag of a 1500-pound nose jettisoned at various altitudes.\n\nNACA RM L59D11\n\n27", "timestamp": "2026-07-22T05:28:58.179655+00:00"}
{"citation_id": "19930085914", "source_url": "https://ntrs.nasa.gov/api/citations/19930085914/downloads/19930085914.pdf", "page_number": 12, "total_pages": 42, "image_filename": "19930085914_p12.jpg", "text": "NACA RM A9D25\n\nform, but having no camber or twist (reference 9). It must be noted that the wing of reference 9 was tested as a semispan model mounted from the tunnel wall and that the gap at the root chord and the existence of a boundary layer on the tunnel wall would have the effect of reducing the effective aspect ratio. That this effect was small is evidenced by the close agreement between the results of the tests of the semispan model (reference 9) and the results of tests of a complete model of a similar wing (reference 8). Furthermore, the cambered and twisted wing of this investigation had streamwise sections of 5-percent-chord thickness as compared with 6-percent-chord thickness for the sections of the plane wing discussed in reference 9. Wing-alone characteristics for the cambered and twisted wing were calculated by subtracting the data obtained from tests of the fuselage from those obtained from tests of the wing-fuselage combination. No account was taken of wing-fuselage interference.\n\nA comparison is made in figure 16 of the aerodynamic characteristics of the two wings at several Mach numbers for Reynolds numbers of approximately 2 million. The principal effect of camber and twist upon the drag characteristics was a reduction of drag at positive lift coefficients above a lift coefficient of about 0.1, indicating an increase in maximum lift-drag ratio. The lift data (fig. 16(b)) indicate a slightly more pronounced reduction of lift-curve slope due to separation at the tips at a lift coefficient of about 0.2 for the cambered and twisted wing. This reduction of lift-curve slope for the cambered and twisted wing occurred at a slightly higher lift coefficient than for the plane wing of reference 9. This delay was probably the result of the reduced angle of attack of the tips due to wing twist. The angle of attack for zero lift was about $0.5^\\circ$ for the cambered and twisted wing as compared with $0^\\circ$ for the plane wing.\n\nFigure 16(c) shows an increase in static longitudinal stability due to camber and twist. The forward movement of the aerodynamic center at a lift coefficient of approximately 0.2, due to separation at the tips, was, in general, slightly more pronounced for the cambered and twisted wing, and occurred at a higher lift coefficient. The final deterioration of stability of the cambered and twisted wing occurred at a lift coefficient about 0.15 higher (approximately 33 percent) than for the plane wing. The cambered and twisted wing had a moment coefficient at zero lift of approximately -0.01; whereas the plane wing of reference 9 had no pitching moment at zero lift.\n\nCONCLUDING REMARKS\n\nThe results of tests of the cambered and twisted wing with the leading edge swept back $63^\\circ$ in combination with a slender fuselage indicate the following:", "timestamp": "2026-07-22T05:29:01.338428+00:00"}
{"citation_id": "19930085972", "source_url": "https://ntrs.nasa.gov/api/citations/19930085972/downloads/19930085972.pdf", "page_number": 2, "total_pages": 46, "image_filename": "19930085972_p2.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:29:02.293562+00:00"}
{"citation_id": "19930085938", "source_url": "https://ntrs.nasa.gov/api/citations/19930085938/downloads/19930085938.pdf", "page_number": 7, "total_pages": 42, "image_filename": "19930085938_p7.jpg", "text": "```markdown\n6\nNACA RM No. L9B04\n\nwith either configuration. This absence of upper-limit porpoising enabled stable take-offs to be made with full elevator deflection ($-30^\\circ$) as shown in figure 8. The high peak trims and the absence of upper-limit porpoising were probably both due to the high sternpost angles ($15\\frac{1}{2}^\\circ$).\n\nIn figure 10, typical plots of variation in trim of the two configurations at fixed elevator deflections are plotted against speed coefficient for the three locations of the center of gravity investigated. Typical photographs are shown in figures 11 and 12. The static trims of both configurations were high (approx. $10^\\circ$). The trim of the single-boom configuration increased until a speed coefficient of approximately 2.5 was reached. From this speed, until a speed coefficient of approximately 5.0, the trim remained fairly constant at large elevator deflections ($-15^\\circ$ to $-30^\\circ$). This flattening of the trim track was the result of the powerful forebody roach which rose almost vertically and hinged on the boom in this speed range. The twin-boom configuration had higher hump trims since the forebody roach did not strike the booms. With both models, trims obtainable with a wide range of elevator deflection were high enough to permit operation above the lower trim limit of stability and no upper-limit porpoising was encountered.\n\nThe stability and trim characteristics of the two configurations differ chiefly in their range of elevator deflection for stable take-offs and the operating trims for given elevator deflections. These differences in stability and trim characteristics for the two models may be attributed primarily to differences in the tail surfaces, differences in the chine strips, and the change in position of booms relative to the roach behind the forebody. Of these three changes, the last constitutes the only difference that is inherent in the change from single-boom to twin-boom configuration. The significant conclusion appears to be that both the single-boom and twin-boom configurations can be designed to have a large range of fixed elevator deflection for stable take-offs over a wide range of location of the center of gravity.\n\n### Landing Stability\n\nThe maximum amplitudes of oscillation in trim and rise during landing of the twin-boom configuration are shown in figure 13. Landings were stable at all contact trims and positions of the center of gravity.\n\nThe maximum amplitudes of oscillation in trim and rise during landings of the single-boom configuration are shown in figure 14. At forward positions of the center of gravity, violent lower-limit porpoising occurred during the landing runout for all landing trims. At after positions of the center of gravity lower-limit porpoising occurred at landing trims below $7^\\circ$. This instability could not be associated with the boom inasmuch as this portion of the hull was generally clear of the water when porpoising occurred. The presence of the vertical chine strips near the point of the step appeared to introduce an undesirable bow-down hydrodynamic moment which\n```", "timestamp": "2026-07-22T05:29:03.981424+00:00"}
{"citation_id": "19930085869", "source_url": "https://ntrs.nasa.gov/api/citations/19930085869/downloads/19930085869.pdf", "page_number": 31, "total_pages": 36, "image_filename": "19930085869_p31.jpg", "text": "NACA RM L9D15 CONFIDENTIAL 29\n\n[Figure: (a) $C_V = 2.77$; $\\tau = 9.8^\\circ$; $\\delta_\\theta = -15^\\circ$.]\n\n[Figure: (b) $C_V = 5.54$; $\\tau = 11^\\circ$; $\\delta_\\theta = -10^\\circ$.]\n\n[Figure: (c) $C_V = 11.08$; $\\tau = 3.5^\\circ$; $\\delta_\\theta = 0^\\circ$.]\n\nFigure 10.— Strobe-flash pictures of swept hull being tested. Full power; gross load coefficient, 3.87; center-of-gravity location, 0.20$\\bar{c}$.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:29:04.519906+00:00"}
{"citation_id": "19930085906", "source_url": "https://ntrs.nasa.gov/api/citations/19930085906/downloads/19930085906.pdf", "page_number": 14, "total_pages": 23, "image_filename": "19930085906_p14.jpg", "text": "NACA RM E9F20 CONFIDENTIAL 13\n\n[Figure: Three-quarter side view of a cylindrical preheater installation with coiled tubing and external piping. NACA stamp: C-19902, 10-28-47]\n\n(a) Three-quarter side view.\n\n[Figure: View looking upstream from combustion chamber, showing circular arrangement of tubes and internal structure. NACA stamp: C-19901, 10-28-47]\n\n(b) View looking upstream from combustion chamber.\n\nFigure 4. - Installation of preheater in zero position in 20-inch ram jet.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:29:10.226436+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 62, "total_pages": 99, "image_filename": "19930082511_p62.jpg", "text": "60\nNACA TN No. 1826\n\n$$\nP_{00}(\\xi, \\rho) = - \\left[ \\frac{2(b-a)}{\\pi} \\right]^2 - \\frac{\\xi - a}{\\frac{\\pi}{2(b-a)}} + \\frac{J_0 \\left[ i\\rho \\frac{\\pi}{2(b-a)} \\right]}{\\frac{i\\pi}{2(b-a)} J_0' \\left[ i \\frac{\\pi}{2(b-a)} \\right]} \\sin \\frac{\\pi}{2} \\frac{\\xi - a}{b - a}\n$$\n\n$$\n+ \\sum_{s=0}^{\\infty} Q_{00}(s) \\left[ y_{s0} e^{-(b-\\xi)y_{s0}} - \\frac{\\pi}{2(b-a)} e^{-(\\xi-a)y_{s0}} \\right]\n$$\n\nfor $\\xi < a$,\n\n$$\nP_{00}(\\xi, \\rho) = - \\left[ \\frac{2(b-a)}{\\pi} \\right]^2 + \\frac{\\xi - a}{\\frac{\\pi}{2(b-a)}}\n$$\n\n$$\n+ \\sum_{s=0}^{\\infty} Q_{00}(s) \\left[ y_{s0} e^{-(b-\\xi)y_{s0}} - \\frac{\\pi}{2(b-a)} e^{-(a-\\xi)y_{s0}} \\right]\n$$\n\nand, for $\\xi > b$,\n\n$$\nP_{00}(\\xi, \\rho) = - \\left[ \\frac{2(b-a)}{\\pi} \\right]^2 - \\frac{\\xi - a}{\\frac{\\pi}{2(b-a)}}\n$$\n\n$$\n- \\sum_{s=0}^{\\infty} Q_{00}(s) \\left[ y_{s0} e^{-(\\xi-b)y_{s0}} - \\frac{\\pi}{2(b-a)} e^{-(\\xi-a)y_{s0}} \\right]\n$$\n\nFor $n \\neq 0$, the corresponding formulas are, for $a \\le \\xi < b$,\n\n$$\nP_{0n}(\\xi, \\rho) = - \\frac{1}{\\frac{n\\pi}{b-a}} \\left[ (\\xi - a) - (-1)^n (b - \\xi) \\right]\n$$\n\n$$\n+ \\frac{J_0 \\left( i\\rho \\frac{n\\pi}{b-a} \\right)}{\\frac{n\\pi}{b-a} J_0' \\left( i \\frac{n\\pi}{b-a} \\right)} \\sin n\\pi \\frac{\\xi - a}{b - a}\n$$\n\n$$\n- \\frac{n\\pi}{b-a} \\sum_{s=0}^{\\infty} Q_{0n}(s) \\left[ e^{-(\\xi-a)y_{s0}} - (-1)^n e^{-(b-\\xi)y_{s0}} \\right]\n$$", "timestamp": "2026-07-22T05:29:10.908345+00:00"}
{"citation_id": "19930085964", "source_url": "https://ntrs.nasa.gov/api/citations/19930085964/downloads/19930085964.pdf", "page_number": 6, "total_pages": 18, "image_filename": "19930085964_p6.jpg", "text": "NACA RM E9G25\n\nclosed loop is formed; that is, the node line completely encircles the blade. Node lines that are discontinuous may have their ends at either the base or the tip and may travel around the blades several times as typified in figure 3(o).\n\nAn effect termed \"breathing\" occurs predominantly in the mode shown in figure 3(c) at a frequency of 1230 cycles per second. In this vibrational mode, the two sides of the blade move alternately toward and away from each other due to the bending of the blade about the node line. This action is illustrated in figure 9. Severe stress concentrations in the leading and trailing edges are produced by this action.\n\nVibrational modes of blades B₁, B₂, B₃, and B₄. - The nodal patterns determined for the stiffened hollow blades B₁, B₂, B₃, and B₄ are shown in figures 4, 5, 6, and 7, respectively. The lowest frequency modes of 890, 880, 850, and 880 cycles per second in figures 4 to 7, respectively, represent a variation of approximately 4.5 percent. Again, the lowest frequency mode is not a true fundamental bending mode. The first six corresponding modes of blades B₁, B₂, and B₃ are similar both in nodal patterns and in frequencies; the small differences observed are attributed partly to dimension and shape variations inherent in the manufacturing process and partly to slight variations in the clamping of the blades. An additional mode, in the case of blade B₄, was observed at a frequency of 1490 cycles per second (fig. 7(d)). Because the amplitude to which this mode could be excited was very low, it is possible that it existed in the cases of blades B₁, B₂, and B₃, but with an amplitude too small to detect. With the exception of the mode in figure 7(d) of blade B₄, the first six modes of blades B₁, B₂, and B₃ (figs. 4(a) to 4(f), 5(a) to 5(f), and 6(a) to 6(f), respectively) are similar to the modes of blade B₄ (figs. 7(a) to 7(c) and 7(e) to 7(g)). The vibrational modes occurring below 2600 cycles per second have nodal patterns on the concave side similar to those on the convex side. Above 2600 cycles per second, plate vibrations generally exist, the nodal patterns being dissimilar on the two sides and becoming more complex with increasing frequency. Also, above 2600 cycles per second there is little similarity in the nodal patterns of the four blades.\n\nThe breathing effect was found to be prominent in the B-type blades in the same mode (approximately 1230 cycles per second) as determined for the A-type blade. In the B-type blades, the alternate movement of the sides toward and away from each other also produces", "timestamp": "2026-07-22T05:29:11.432397+00:00"}
{"citation_id": "19930085519", "source_url": "https://ntrs.nasa.gov/api/citations/19930085519/downloads/19930085519.pdf", "page_number": 46, "total_pages": 46, "image_filename": "19930085519_p46.jpg", "text": "NACA RM No. L8K19\n45\n\n<!-- Image (153, 81, 885, 782) -->\n\n(b) Partial-span slotted flap at $\\delta_F = 50^\\circ$.\nFigure 22.- Concluded.", "timestamp": "2026-07-22T05:29:13.841390+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 27, "total_pages": 96, "image_filename": "19930085880_p27.jpg", "text": "NACA RM No. L9C03\n25\n\n.30\n.25\n.20\nResistance, lb\n.15\n.10\n.05\n0\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\nNACA\n\n(d) $\\tau = 16^\\circ$.\n\nFigure 11.- Continued.", "timestamp": "2026-07-22T05:29:14.404500+00:00"}
{"citation_id": "19930082617", "source_url": "https://ntrs.nasa.gov/api/citations/19930082617/downloads/19930082617.pdf", "page_number": 45, "total_pages": 58, "image_filename": "19930082617_p45.jpg", "text": "44\n\nPage intentionally left blank\n\nPage intentionally left blank", "timestamp": "2026-07-22T05:29:15.979702+00:00"}
{"citation_id": "19930085962", "source_url": "https://ntrs.nasa.gov/api/citations/19930085962/downloads/19930085962.pdf", "page_number": 6, "total_pages": 51, "image_filename": "19930085962_p6.jpg", "text": "NACA RM A9E05 CONFIDENTIAL 5\n\nTare corrections due to the air forces exerted on the exposed area of the turntable were obtained from force measurements made with the model removed from the tunnel. Possible interference effects between the model and the turntable were not evaluated but they are believed to be small. The magnitude of the measured tare drag coefficient was 0.0063.\n\nTESTS\n\nLift, drag, and pitching-moment data have been obtained for a range of angle of attack at a constant Reynolds number of 2,000,000 and Mach numbers from 0.20 to 0.94. For each angle of attack and Mach number, tests were made with elevator deflections of $0^\\circ$, $2^\\circ$, $4^\\circ$, $6^\\circ$, $10^\\circ$, $20^\\circ$, and $30^\\circ$ except at a Mach number of 0.94 where the maximum elevator deflection was limited to $20^\\circ$. At low speeds, the angle-of-attack range was from $-15^\\circ$ to $15^\\circ$, but at Mach numbers above 0.85 the range was limited by tunnel power and model strength.\n\nRESULTS AND DISCUSSION\n\nLift, drag, and pitching-moment characteristics as a function of angle of attack are presented in figures 3 to 11, inclusive, for elevator deflections of $0^\\circ$, $2^\\circ$, $4^\\circ$, $6^\\circ$, $10^\\circ$, $20^\\circ$, and $30^\\circ$ at Mach numbers of 0.20, 0.50, 0.70, 0.80, 0.85, 0.87, 0.90, 0.92, and 0.94. Since the tail profile is symmetrical, the data presented in these figures for positive elevator deflections can be used to indicate the effect of negative elevator deflections by simply reversing the algebraic signs of the coordinate axes. The variation of lift coefficient with elevator deflection for various Mach numbers is shown in figure 12 and the lift data are plotted in figures 13 and 14 as a function of Mach number.\n\nLift Characteristics\n\nStabilizer effectiveness.— The aerodynamic characteristics of the horizontal tail with the elevator neutral (figs. 3 to 11, inclusive) have been fully reported in reference 1. Despite the symmetry of the profile, the character of the stall with the elevator neutral was dependent on the algebraic sign of the angle of attack. This is especially noticeable at a Mach number of 0.80 (fig. 6) where the tail stalled abruptly with a sizable loss of lift at $8^\\circ$ angle of attack, but had a gentle stall with a relatively small loss of lift at about $-12^\\circ$ angle of attack. This asymmetry may be due in part to\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:29:16.204681+00:00"}
{"citation_id": "19930085912", "source_url": "https://ntrs.nasa.gov/api/citations/19930085912/downloads/19930085912.pdf", "page_number": 13, "total_pages": 36, "image_filename": "19930085912_p13.jpg", "text": "NACA RM No. E9C16\n\nThe theoretical curves A and B shown in figure 12 are based on the assumption that icing occurs when the minimum kinetic temperature (static temperature plus 0.85 times the dynamic temperature) on the screen was $32^\\circ$ F. These curves represent the upper and lower limits of the icing conditions used in the investigation; that is, curve A was calculated for saturated air at $0^\\circ$ F, a tunnel velocity of 410 feet per second, and a liquid-water content of 1.0 gram per cubic meter. Curve B was calculated for saturated air at $0^\\circ$ F, a tunnel velocity of 220 feet per second, and a liquid-water content of 0.7 gram per cubic meter.\n\nIn order to assure a minimum kinetic temperature of $32^\\circ$ F on the screen, an average kinetic temperature of $38^\\circ$ F and a total temperature of $41.3^\\circ$ F are required at a tunnel velocity of 410 feet per second, corresponding to a velocity in the screen of 516 feet per second; a total temperature of $39.4^\\circ$ F is required at a tunnel velocity of 220 feet per second, corresponding to a velocity in the screen of 302 feet per second.\n\nNearly all the data in figure 12 fall within the limits of the two curves. If the conditions of the investigation had been ideal, the lower-speed data would have fallen near curve B and the higher-speed data near curve A, provided that the liquid-water content in all cases was constant. The variation in liquid-water content and the use of other than the optimum amount of bleedback preclude such a correlation.\n\nInlet-lip temperature distribution. - A marked reduction in inlet-lip temperature, particularly near the stagnation region, was observed under icing conditions as compared with nonicing conditions. Typical lip-temperature profiles for both conditions are shown in figure 13 for a bleedback of 4.4 percent and a plenum-chamber gas temperature of $1000^\\circ$ F. The liquid-water content for the icing condition was 0.5 gram per cubic meter.\n\nSUMMARY OF RESULTS\n\nThe following results were obtained from an icing-research-tunnel investigation of a two-thirds-scale model of a turbojet-engine nacelle with a long straight air inlet utilizing a hot-gas bleedback system for ice prevention:\n\n1. Identical temperature distributions were obtained at the simulated engine inlet for a fixed amount of bleedback independent of tunnel velocity.", "timestamp": "2026-07-22T05:29:20.123239+00:00"}
{"citation_id": "19930082914", "source_url": "https://ntrs.nasa.gov/api/citations/19930082914/downloads/19930082914.pdf", "page_number": 55, "total_pages": 66, "image_filename": "19930082914_p55.jpg", "text": "54\n\nPage intentionally left blank\n\nPage intentionally left blank", "timestamp": "2026-07-22T05:29:21.998421+00:00"}
{"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 47, "total_pages": 78, "image_filename": "19930082618_p47.jpg", "text": "NACA TN 1945\n\nSection lift coefficient, $c_l$\nSection angle of attack, $\\alpha_{sec}$, deg\n\nR\n$\\circ$ 0.7 $\\times$ 10$^6$\n$\\square$ 1.0\n$\\triangle$ 1.5\n$\\nabla$ 2.0\n$\\diamond$ 3.0\n$\\blacktriangle$ 6.0\n$\\blacktriangledown$ 9.0\n\nFlagged symbols denote\nstandard roughness\n\nNACA\n\nMoment coefficient, $c_{m_{c/4}}$\nSection angle of attack, $\\alpha_{sec}$, deg\n\n(a) Section lift and pitching-moment characteristics of the plain airfoil section.\nFigure 9.- Aerodynamic characteristics of the NACA 65$_2$-415 airfoil section, 24-inch chord.\n\n45", "timestamp": "2026-07-22T05:29:22.338913+00:00"}
{"citation_id": "19930082542", "source_url": "https://ntrs.nasa.gov/api/citations/19930082542/downloads/19930082542.pdf", "page_number": 43, "total_pages": 53, "image_filename": "19930082542_p43.jpg", "text": "NACA TN No. 1867\n49\n\n<!-- Image (117, 101, 903, 431) -->\n\n100X\n1000X\n(a) Solution-treated 2200° F\n1 hour, water-quenched.\n\n<!-- Image (112, 530, 901, 868) -->\n\n100X\n1000X\n(b) 1400° F 24 hours.\nNACA\n\nFigure 8.- Effect of aging at two temperatures on microstructure of\nlow-carbon N-155 bar stock.", "timestamp": "2026-07-22T05:29:23.860826+00:00"}
{"citation_id": "19930085934", "source_url": "https://ntrs.nasa.gov/api/citations/19930085934/downloads/19930085934.pdf", "page_number": 8, "total_pages": 23, "image_filename": "19930085934_p8.jpg", "text": "NACA RM E9G12\n\n(3) Entropy due to superheated vapor in incoming air. - Although the Mollier diagram for steam does not include values of entropy for a pressure of 0.1110 pound per square inch, the low-pressure region near saturation indicates an entropy increase of 0.018 Btu per pound per °F for each temperature rise of 20° F along an isobaric line. By using the entropy at saturation for a pressure of 0.1110 pound per square inch and correcting for temperature difference from saturation temperature to 77.4° F, an entropy value $s_{s,1}$ of 2.204 Btu per pound per °F was obtained.\n\nEntropy of superheated vapor per pound of dry air at inlet\n\n$$\n\\begin{aligned}\ns_{s,a,1} &= s_{s,1} \\cdot q_1 \\\\\n&= (2.204)(0.01025) \\\\\n&= 0.02259 \\text{ (Btu/(lb)(°F))}\n\\end{aligned}\n$$\n\n(4) Entropy due to dry air. -\n\nPressure of dry air\n\n$$\n\\begin{aligned}\nP_{d,1} &= P_1 - P_{s,1} \\\\\n&= 6.870 - 0.1110 \\\\\n&= 6.759 \\text{ (lb/sq in.)} \\\\\n\\frac{1}{N} &= \\frac{P_0}{P_{d,1}} \\\\\n&= 14.70 / 6.759 \\\\\n&= 2.175 \\\\\nR \\log_e \\frac{1}{N} &= - R \\log_e N = - 0.05326\n\\end{aligned}\n$$\n\nwhere\n\n$$\nR = 53.35 \\text{ (ft-lb/(lb)(°F))}\n$$\n\nThe temperature function at the inlet $\\Phi_{t,1}$ is 0.07067, from table 1 of reference 2.", "timestamp": "2026-07-22T05:29:26.228889+00:00"}
{"citation_id": "19930085881", "source_url": "https://ntrs.nasa.gov/api/citations/19930085881/downloads/19930085881.pdf", "page_number": 22, "total_pages": 31, "image_filename": "19930085881_p22.jpg", "text": "20\nNACA RM L9D12\n\nCONFIDENTIAL\n\n.12\n$C_D$ .08\n.04\n0\n\n.16\n.12\n$pb/2V$ .08\n.04\n0\n\nModel $\\delta_a (deg)$ $i_w (deg)$\n119b 5.0 0\n\n.6 .8 1.0 1.2 1.4 1.6 1.8 2.0\nM\nNACA\n\n(b) 9-percent-thick double-wedge airfoil section.\nFigure 5.- Continued.\nCONFIDENTIAL", "timestamp": "2026-07-22T05:29:26.979970+00:00"}
{"citation_id": "19930085542", "source_url": "https://ntrs.nasa.gov/api/citations/19930085542/downloads/19930085542.pdf", "page_number": 29, "total_pages": 46, "image_filename": "19930085542_p29.jpg", "text": "NACA RM No. L8L29\n27\n\nModel No. Profile\n1 Flat plate\n2 NACA 0012\n3 Biconvex 12%\n\nAngle of attack, $\\alpha$, deg\nLongitudinal-force coefficient, $C_X$\n\nPitching-moment coefficient, $C_m$\nLift coefficient, $C_L$\n\nNACA\n\nFigure 8.— Effect of profile on aerodynamic characteristics of a triangular wing of aspect ratio 2.31. $\\Delta c/\\mu = 52.2^\\circ$.", "timestamp": "2026-07-22T05:29:28.267341+00:00"}
{"citation_id": "19930085889", "source_url": "https://ntrs.nasa.gov/api/citations/19930085889/downloads/19930085889.pdf", "page_number": 23, "total_pages": 37, "image_filename": "19930085889_p23.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:29:39.032770+00:00"}
{"citation_id": "19930085972", "source_url": "https://ntrs.nasa.gov/api/citations/19930085972/downloads/19930085972.pdf", "page_number": 3, "total_pages": 46, "image_filename": "19930085972_p3.jpg", "text": "NACA RM L9B18\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nRESEARCH MEMORANDUM\n\nLOW-SPEED WIND-TUNNEL INVESTIGATION OF THE LONGITUDINAL  \nSTABILITY CHARACTERISTICS OF A MODEL EQUIPPED  \nWITH A VARIABLE-SWEEP WING  \n\nBy Charles J. Donlan and William C. Sleeman, Jr.\n\nSUMMARY\n\nAn investigation has been made to determine the longitudinal stability characteristics of a complete model equipped with a variable-sweep wing at angles of sweepback of $45^\\circ$, $30^\\circ$, $15^\\circ$, and $0^\\circ$. The investigation was directed toward the study of various wing modifications and an external-flap arrangement designed to minimize the shift in neutral point accompanying the change in sweep angle.\n\nThe results indicated that stability at the stall was obtained at a sweep angle of $15^\\circ$ without recourse to stall-control devices. The basic neutral-point movement accompanying the change in sweep angle from $45^\\circ$ to $15^\\circ$ amounted to 56 percent of the mean aerodynamic chord (at zero sweep angle) and the most effective modification investigated only reduced this change to 47 percent of the chord. It appears, therefore, that for designs in which the fuselage is the major load-carrying element some relative movement between the wing and center of gravity will be required to assure satisfactory stability at all sweep angles.\n\nINTRODUCTION\n\nThe use of swept wings on high-speed airplanes has introduced serious longitudinal- and lateral-stability problems at low speeds. Many high-lift and stall-control devices have been investigated in an attempt to improve the low-speed characteristics of highly swept wings but no completely satisfactory solution has been found. One obvious method for avoiding the low-speed problems associated with highly swept wings would be to employ a wing whose sweep angle could be varied in flight. Thus, for maximum high-speed flight and optimum cruising performance, the wing could be adjusted to any desired sweep angle; whereas, for the landing condition, the sweep angle could be decreased to an angle that would assure satisfactory low-speed characteristics without recourse to stall-control devices.\n\nThe present paper presents the results of a wind-tunnel investigation of a complete model equipped with a wing whose sweep angle could be varied for the purpose of studying various wing modifications designed to decrease", "timestamp": "2026-07-22T05:29:43.468602+00:00"}
{"citation_id": "19930085906", "source_url": "https://ntrs.nasa.gov/api/citations/19930085906/downloads/19930085906.pdf", "page_number": 15, "total_pages": 23, "image_filename": "19930085906_p15.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:29:44.004704+00:00"}
{"citation_id": "19930085977", "source_url": "https://ntrs.nasa.gov/api/citations/19930085977/downloads/19930085977.pdf", "page_number": 1, "total_pages": 33, "image_filename": "19930085977_p1.jpg", "text": "FILE COPY\nNO 8\n\nCONFIDENTIAL\n\nCopy\nRM L9H22\n\nNACA RM L9H22\n\nNACA\n\nRESEARCH MEMORANDUM\n\nAERODYNAMIC CHARACTERISTICS OF A WING WITH\nUNSWEEPED QUARTER-CHORD LINE, ASPECT RATIO 4,\nTAPER RATIO 0.6, AND NACA 65A006 AIRFOIL SECTION\n\nTRANSONIC-BUMP METHOD\n\nBy Kenneth W. Goodson and William D. Morrison, Jr.\n\nLangley Aeronautical Laboratory\nLangley Air Force Base, Va.\n\nTHIS DOCUMENT ON LOAN FROM THE NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\nLANGLEY AERONAUTICAL LABORATORY\nLANGLEY FIELD, HAMPTON, VIRGINIA\n\nRETURN TO THE ABOVE ADDRESS\nRE: FOR PUBLICATIONS SHOULD BE ADDRESSED\nAS FOLLOWS:\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n1512 H ST., N. W.\nWASHINGTON 25, D. C.\n\nCLASSIFIED DOCUMENT\nThis document contains classified information affecting the National Defense of the United States within the meaning of the Espionage Act, USC 50:31 and 32. Its transmission or the revelation of its contents in any manner to an unauthorized person is prohibited by law.\nInformation so classified may be imparted only to persons in the military and naval services of the United States, appropriate civilian officers and employees of the Federal Government who have a legitimate interest therein, and to United States citizens of known loyalty and discretion who of necessity must be informed thereof.\n\nCLASSIFICATION CHANGED TO\nUNCLASSIFIED\nDATE 8-18-54\nAUTHORITY J.W. CHOWLEY\nCHANGE # 2460\nF.E.T.\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\nWASHINGTON\nOctober 21, 1949\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:29:45.359490+00:00"}
{"citation_id": "19930085914", "source_url": "https://ntrs.nasa.gov/api/citations/19930085914/downloads/19930085914.pdf", "page_number": 13, "total_pages": 42, "image_filename": "19930085914_p13.jpg", "text": "12\nNACA RM A9D25\n\nEffects of Mach Number\n\nThere was little difference in Mach number effects on the wing-fuselage combination for Reynolds numbers of 0.8 million and 2.0 million. Variation of Mach number from 0.20 to 0.93 at a Reynolds number of 2.0 million affected the aerodynamic characteristics as follows:\n\n1. The abrupt forward movement of the aerodynamic center beginning at a lift coefficient of about 0.5 to 0.6 was reduced in severity.\n\n2. The aerodynamic center at zero lift moved rearward from about 41 percent of the mean aerodynamic chord to about 45 percent.\n\n3. The lift-curve slope increased from about 0.049 to 0.055 per degree.\n\nEffects of Reynolds Number\n\nIncreasing Reynolds number from 0.8 million to 9.0 million at a Mach number of 0.20 affected the aerodynamic characteristics of the wing-fuselage combination as follows:\n\n1. Minor irregularities in static longitudinal stability were virtually eliminated up to a lift coefficient of about 0.55.\n\n2. The drag was reduced at positive lift coefficients above a lift coefficient of about 0.2.\n\n3. The lift-curve slope decreased from 0.051 to 0.046 per degree.\n\n4. The aerodynamic-center position at zero lift was little affected by changes in Reynolds number, moving from 41 percent to 39 percent of the mean aerodynamic chord.\n\n5. The data from these tests indicate that certain important effects of boundary-layer separation which are evident from tests of highly swept-back wings at low Reynolds numbers may not be present under full-scale conditions.\n\nEffects of Camber and Twist\n\nThe following effects of camber and twist were indicated by a comparison of the results for the cambered and twisted wing with those for a wing of identical plan form having no camber or twist:\n\n1. The drag coefficients were reduced at positive lift coefficients above about 0.1.", "timestamp": "2026-07-22T05:29:45.715581+00:00"}
{"citation_id": "19930085938", "source_url": "https://ntrs.nasa.gov/api/citations/19930085938/downloads/19930085938.pdf", "page_number": 8, "total_pages": 42, "image_filename": "19930085938_p8.jpg", "text": "NACA RM No. L9B04\n\nresulted in low trims. Extending the chine strips aft on the twin-boom configuration caused similar landing behavior for this model.\n\nSatisfactory landing stability, therefore, can be attained with either configuration. To avoid instability during the landing runout, however, vertical chine strips near the point of the step should be avoided.\n\nSpray\n\nThe range of speed over which spray entered the propellers is plotted against gross load coefficient in figure 15 for both configurations. At the gross load used for the stability tests (65,000 lb, full size, $C_{\\Delta}$ of 3.87), the propellers of the twin-boom configuration operated in spray over a shorter speed-coefficient range ($C_V = 2.0$ to $2.6$) than did those of the single-boom configuration ($C_V = 1.4$ to $3.8$) as a result of the higher trims and greater nacelle spacing of the twin-boom model. For both models the chine strips produced a confused pattern of light spray which tended to become more intense as load was increased beyond the load at which spray first entered the propellers. At the gross load used for stability tests the propeller spray of both models was satisfactory. Figures 11(b), 11(c), and 12(a) are photographs of the models operating in the spray region at normal gross load of 65,000 pounds, full size.\n\nAt high trims, through a speed range from approximately $C_V = 6$ to take-off, transverse spray from the forebody, aft of the vertical chine strips, wetted the under surface of the wing and the booms of the twin-boom configuration.\n\nResistance\n\nResistance coefficient, load-resistance ratio, trim, and load coefficient at best trim (with $12^\\circ$ considered the maximum usable trim at high speed) are plotted against speed coefficient in figure 16. The hump $\\Delta/R$ values of 3.6 for the single-boom and 2.9 for the twin-boom are considerably less than those obtained in well-designed conventional hulls but are of the same order as those of single-float seaplanes. Actually a lower power loading than was used in the powered model tests would be needed in order to take off without assistance; a high-speed airplane would have such a low power loading. High hump trims and, therefore, high hump resistance were natural results of placing small booms high with respect to the forebody.\n\nThe twin-boom model appeared to have inherently higher resistance than the single-boom model over most of the speed range. At the hump speed the single boom rode on the roach behind the forebody. The resultant decrease in trim tended to lower resistance. At high speed, the differences in the", "timestamp": "2026-07-22T05:29:46.081376+00:00"}
{"citation_id": "19930082617", "source_url": "https://ntrs.nasa.gov/api/citations/19930082617/downloads/19930082617.pdf", "page_number": 46, "total_pages": 58, "image_filename": "19930082617_p46.jpg", "text": "NACA TN 1962\n\n45\n\n[Figure: Side view of a buckled cylindrical structure with riveted seams and visible deformation. Wires are attached to the side.]\n\nFigure 32.- Side view of cylinder 73 after buckling.", "timestamp": "2026-07-22T05:29:48.987907+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 28, "total_pages": 96, "image_filename": "19930085880_p28.jpg", "text": "26\nNACA RM No. L9C03\n\nResistance, lb\nWetted area\n(sq ft)\n0\n.05\n.10\n.15\n.20\n.25\n.30\n\nSpeed, fps\n0 10 15 20 25 30 35\n\n(e) T = 20°.\nFigure 11.- Concluded.", "timestamp": "2026-07-22T05:29:50.609108+00:00"}
{"citation_id": "19930085964", "source_url": "https://ntrs.nasa.gov/api/citations/19930085964/downloads/19930085964.pdf", "page_number": 7, "total_pages": 18, "image_filename": "19930085964_p7.jpg", "text": "6\nNACA RM E9G25\n\na stress concentration at the rivet holes. Failure by fatigue in\nthis mode produced cracks that started at the rivet holes and\nextended to the tip of the blade. Because the stiffening band at\nthe tip of the blade neither shifts the natural frequency nor\nappreciably changes the nodal pattern and because the presence of\nthe rivet holes constitutes a severe stress concentration, the\nadvantages of the stiffening features are doubtful. Use of the\nstiffening band alone would possibly be more beneficial in this\nmode inasmuch as the points of severe stress concentration would\nthen have greater strength.\n\nVibrational modes of blade C. - The nodal patterns determined\nfor the solid blade C are shown in figure 8. This blade had the\nsame dimensions and airfoil contour as blades A and B except for a\nthinner trailing edge. The lowest frequency mode occurred at\n1270 cycles per second and can be considered a true fundamental bend-\ning mode. The first torsional mode occurred at 2050 cycles per sec-\nond. Little similarity exists between the nodal patterns of the\nsolid and the hollow blades, but, as might be expected, the nodal\npatterns on the respective sides of the solid blade were similar in\nall modes. Fewer vibrational modes were detected in the solid blade\nthan in any of the hollow blades. Also, at the higher frequencies,\nmodes of the solid blade did not appear as complex as those of the\nhollow blades. In this investigation, all the modes in which the\nblades are capable of vibrating were not determined. Modes of ampli-\ntudes too small to determine the nodal patterns accurately were\ndetected in all blades. There probably are modes that would occur\nabove the frequency limitation of the equipment, which was approx-\nimately 11,000 cycles per second.\n\nProbability of excitation of blades A and B in an engine. - The\nexcitation orders present in turbojet engines have been shown to vary\nover a wide range. Reference 2 indicates the presence of strong 7th,\n8th, 11th, 14th, and 45th orders in a particular engine. Inasmuch\nas the configuration of this engine is typical of conventional turbo-\njet engines, nearly all the vibrational modes determined for hollow\nblades A and B could be excited during engine operation. Nearly all\nthe vibrational modes determined for blades A and B must therefore\nbe assumed to represent potential sources of fatigue failure. On\nthe basis of the results presented, it is improbable that blades of\nthe types represented by blades A and B could be designed with all\nresonant vibration responses outside the operating range of the\nengine. The amplitudes of those modes excited must therefore be\nlimited to values that do not produce excessive stresses in the\nblades.", "timestamp": "2026-07-22T05:29:56.590636+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 1, "total_pages": 114, "image_filename": "19930086061_p1.jpg", "text": "RESTRICTED\nCopy 277\nRM L9J07\n\nNACA RM L9J07\n\nNACA\n\nRESEARCH MEMORANDUM\n\nCLASSIFICATION CHANGED TO\nUNCLASSIFIED\nAUTHORITY CROWLEY CHANGE #1923\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 $60^\\circ$\nAND BICONVEX SECTIONS\n\nBy Ralph W. May, Jr., and John G. Hawes\n\nLangley Aeronautical Laboratory\nLangley Air Force Base, Va.\n\nCLASSIFIED DOCUMENT\n\nThis document contains classified information\naffecting the National Defense of the United\nStates within the meaning of the Espionage Act,\nUSC 50:31 and 32. Its transmission or the\nrevelation of its contents in any manner to an\nunauthorized person is prohibited by law.\nInformation so classified may be imparted\nonly to persons in the military and naval\nservices of the United States, appropriate\ncivilian officers and employees of the Federal\nGovernment who have a legitimate interest\ntherein, and to United States citizens of known\nloyalty and discretion who of necessity must be\ninformed thereof.\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\n\nWASHINGTON\nNovember 18, 1949\n\nRESTRICTED", "timestamp": "2026-07-22T05:30:00.662533+00:00"}
{"citation_id": "19930082914", "source_url": "https://ntrs.nasa.gov/api/citations/19930082914/downloads/19930082914.pdf", "page_number": 56, "total_pages": 66, "image_filename": "19930082914_p56.jpg", "text": "NACA TN No. 1857\n55\n\n[Figure: A grayscale image showing a pattern of wavy, horizontal lines that appear to be interference fringes or flow visualization. The lines are denser and more regular at the bottom, becoming more distorted and irregular towards the top. Two small, dark, diamond-shaped markers are visible at the bottom left and right corners of the image area.]\n\n(e) $5\\frac{1}{2}$ to 7 inches from nozzle.\nFigure 11.- Continued.\nNACA", "timestamp": "2026-07-22T05:30:02.609047+00:00"}
{"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 48, "total_pages": 78, "image_filename": "19930082618_p48.jpg", "text": "```markdown\n2.8\n2.4\nSection lift coefficient, $c_l$\n2.0\n1.6\n1.2\n.8\n.4\n0\n-.4\n-.8\n-1.2\n-1.6\nMoment coefficient, $c_{m_{c/4}}$\n-16\n-8\n0\n8\n16\nSection angle of attack, $\\alpha_s$, deg\n\nR\n$\\circ$ 0.7 $\\times$ 10$^6$\n$\\square$ 1.0\n$\\diamond$ 1.5\n$\\triangle$ 2.0\n$\\triangledown$ 6.0\nFlagged symbols denote\nstandard roughness\n\nNACA\n\n(b) Section lift and pitching-moment characteristics of the NACA 65$_2$-415 airfoil section with a\n0.20c simulated split flap deflected 60$^\\circ$.\nFigure 9.— Continued.\n\n94\nNACA TN 1945\n```", "timestamp": "2026-07-22T05:30:03.477676+00:00"}
{"citation_id": "19930085962", "source_url": "https://ntrs.nasa.gov/api/citations/19930085962/downloads/19930085962.pdf", "page_number": 7, "total_pages": 51, "image_filename": "19930085962_p7.jpg", "text": "6\nCONFIDENTIAL\nNACA RM A9E05\n\na small inadvertent deflection of the leading-edge flap or to\ndifferences in the surface roughness of the upper and lower surfaces\nas a result of the flap angle brackets. Except at the Mach number\nof 0.80, the variation of lift coefficient with angle of attack with\nthe elevator neutral was nearly symmetrical about the angle of zero\nlift. At Mach numbers of 0.80 and 0.85 with the elevator neutral,\nthe tail stalled abruptly with a sizable loss of lift; whereas at\nMach numbers less than 0.80 the lift curve was rounded at the stall.\nThe Mach number at which the type of stall changed was affected to\nsome extent by the Reynolds number as can be seen from the data of\nreference 1.\n\nThe effect of compressibility on the lift-curve slope with the\nelevator neutral is shown in figure 15. The lift-curve slope increased\nfrom 0.062 to 0.095 as the Mach number increased from 0.20 to 0.94.\n\nElevator effectiveness.- The variation of lift coefficient with\nelevator deflection for various Mach numbers is shown in figure 12.\nAt deflections between 0° and 2°, the elevator effectiveness was\ngenerally lower than for deflections between 2° and 4°. Neither the\nmagnitude nor the extent of this reduced effectiveness at deflections\nnear zero was aggravated by compressibility. At a Mach number of 0.20,\nthe elevator effectiveness was approximately linear from 2° to 15°\nwith a value of slope $\\partial C_L/\\partial \\delta$ of 0.034. The value of $\\partial C_L/\\partial \\delta$\npredicted from thin-airfoil theory (reference 7), assuming the experi-\nmental value of tail lift-curve slope, was 0.035.\n\nThe effects of compressibility on the elevator effectiveness\nparameters are presented in figure 15 where $C_{L\\delta}^*$ and $a_\\delta$ are shown\nas functions of Mach number. Due to the nonlinearity of the variation\nof the lift coefficient with elevator deflection for small elevator\ndeflections, the effectiveness parameter $C_{L\\delta}^*$ was obtained as the\ndifference in the lift coefficient due to 4° of elevator deflection\ndivided by 4. The value of $C_{L\\delta}^*$ was 0.030 and was not affected by\ncompressibility at Mach numbers less than 0.60. As the Mach number\nwas increased to 0.90, $C_{L\\delta}^*$ increased to 0.044, subsequently decreas-\ning to 0.038 at a Mach number of 0.94.\n\nTo demonstrate more clearly the effects of compressibility on the\nelevator effectiveness, the variation of lift coefficient with Mach\nnumber for various angles of attack of the tail with constant elevator\ndeflections is presented in figure 13, and the variation of lift coef-\nficient with Mach number for various elevator deflections at constant\nangles of attack is presented in figure 14. In figure 14, the data\nobtained with a positive elevator deflection and a negative angle of\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:30:03.858005+00:00"}
{"citation_id": "19930082542", "source_url": "https://ntrs.nasa.gov/api/citations/19930082542/downloads/19930082542.pdf", "page_number": 44, "total_pages": 53, "image_filename": "19930082542_p44.jpg", "text": "NACA TN No. 1867\n51\n\n[Figure: Microstructure images showing grain boundaries and precipitates at two magnifications]\n\n100X\n(c) 1600° F 24 hours.\nFigure 8.- Concluded.\n1000X\nNACA", "timestamp": "2026-07-22T05:30:04.905304+00:00"}
{"citation_id": "19930085912", "source_url": "https://ntrs.nasa.gov/api/citations/19930085912/downloads/19930085912.pdf", "page_number": 14, "total_pages": 36, "image_filename": "19930085912_p14.jpg", "text": "12\nNACA RM No. E9C16\n\n2. Optimum temperature distribution was obtained at a bleedback of 4.4 percent. This value of bleedback resulted in an average dry-air-temperature rise of 46° F with a maximum local temperature deviation of 60° F at the engine inlet for a gas temperature of 1000° F.\n\n3. The use of plenum-chamber gas pressure other than the optimum resulted in increased temperature gradients across the inlet. For pressures higher than the optimum, low-temperature regions existed near the duct walls and for pressures lower than optimum the low-temperature regions occurred at the duct center.\n\n4. The introduction of cold gas under pressure through the orifices decreased the ram-pressure recovery linearly with increasing bleedback. The ram-pressure recovery with hot-gas bleedback decreased nearly linearly with increasing average model-air temperature.\n\n5. The decrease in mass flow with hot-gas bleedback was almost entirely attributable to the decrease in air density resulting from the increase in air temperature in the model.\n\n6. Satisfactory agreement was obtained between the calculated heat requirements and the measured heat requirements for icing conditions.\n\nCONCLUDING REMARKS\n\nThe foregoing discussion indicates that it is possible by means of analysis to obtain a satisfactory orifice configuration for the protection of a jet-engine nacelle by hot-gas bleedback and to predict many of its thermodynamic and aerodynamic characteristics. The change in mass flow and the reduction in ram-pressure recovery due to the addition of heat can be accurately computed. It has been established that the temperature distribution at the engine inlet is a function of bleedback alone and that the amount of bleedback required for a given icing condition can be accurately calculated.\n\nLewis Flight Propulsion Laboratory,\nNational Advisory Committee for Aeronautics,\nCleveland, Ohio.", "timestamp": "2026-07-22T05:30:06.556379+00:00"}
{"citation_id": "19930085881", "source_url": "https://ntrs.nasa.gov/api/citations/19930085881/downloads/19930085881.pdf", "page_number": 23, "total_pages": 31, "image_filename": "19930085881_p23.jpg", "text": "NACA RM L9D12\n21\n\nCONFIDENTIAL\n\n.12\n.08\n$C_D$\n.04\n0\n\n.12\n.08\n$pb/2V$\n.04\n0\n-.04\n\nModel $\\delta_a(deg)$ $i_w(deg)$\n115e 4.0 -0.18\n115f 3.6 -0.17\n\n.6 .8 1.0 1.2 1.4 1.6 1.8 2.0\nM\n\n(c) NACA 16-009 airfoil section.\nFigure 5.— Continued.\nCONFIDENTIAL", "timestamp": "2026-07-22T05:30:13.603160+00:00"}
{"citation_id": "19930085869", "source_url": "https://ntrs.nasa.gov/api/citations/19930085869/downloads/19930085869.pdf", "page_number": 32, "total_pages": 36, "image_filename": "19930085869_p32.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:30:14.261200+00:00"}
{"citation_id": "19930085934", "source_url": "https://ntrs.nasa.gov/api/citations/19930085934/downloads/19930085934.pdf", "page_number": 9, "total_pages": 23, "image_filename": "19930085934_p9.jpg", "text": "8\nNACA RM E9G12\n\nEntropy of dry air at inlet\n\n$$\n\\begin{aligned}\ns_{d,1} &= \\varphi_{t,1} - R \\log_e N \\\\\n&= 0.07067 + 0.05326 \\\\\n&= 0.1239 \\text{ (Btu/(lb)(}^\\circ\\text{F))}\n\\end{aligned}\n$$\n\n(5) Entropy of mixture at inlet. - Because ratios of the quantities remain unchanged, the use of tables having different temperature bases (references 1 and 2) is valid. The entropy of mixture per pound of dry air equals the entropy of saturated liquid per pound of dry air plus the entropy of superheated vapor per pound of dry air plus the entropy of dry air.\n\n$$\n\\begin{aligned}\ns_{m,a,1} &= s_{f,a,1} + s_{s,a,1} + s_{d,1} \\\\\n&= 0.002294 + 0.02259 + 0.1239 \\\\\n&= 0.1488 \\text{ (Btu/(lb)(}^\\circ\\text{F))}\n\\end{aligned}\n$$\n\nII. Isentropic Outlet Temperature\n\n(1) Outlet temperature. - An outlet temperature $t_{t,2}$ of $108.1^\\circ$ F is assumed for computing an entropy at the outlet equal to that at the inlet by trial-and-error solution.\n\n(2) Specific humidity at outlet. -\n\n$P_{g,2}$ at $t_{t,2}$ of $108.1^\\circ$F\n\n$$\n= 1.206 \\text{ (lb/sq in.)(from table 1, reference 1)}\n$$\n\n$$\nP_2 = 19.998 \\text{ (lb/sq in.) (original conditions)}\n$$\n\n$$\n\\begin{aligned}\nP_{d,2} &= P_2 - P_{g,2} \\\\\n&= 19.998 - 1.206 \\\\\n&= 18.79 \\text{ (lb/sq in.)}\n\\end{aligned}\n$$\n\n$v_{g,2}$ at $t_{t,2}$ of $108.1^\\circ$F\n\n$$\n= 279.5 \\text{ (cu ft/lb)(from table 1, reference 1)}\n$$", "timestamp": "2026-07-22T05:30:15.955237+00:00"}
{"citation_id": "19930085542", "source_url": "https://ntrs.nasa.gov/api/citations/19930085542/downloads/19930085542.pdf", "page_number": 30, "total_pages": 46, "image_filename": "19930085542_p30.jpg", "text": "28\nNACA RM No. L8L29\n\n.016\n.012\n.008\n.004\n0\n-.004\n$C_{Y\\psi}$\n\nModel No. Profile\n1 Flat plate\n2 NACA 0012\n3 Biconvex (12% Thick)\n\n.002\n0\n-.002\n-.004\n$C_{n\\psi}$\n\n.004\n.002\n0\n-.002\n-.004\n$C_{l\\psi}$\n\n-2 0 .2 .4 .6 .8 1.0 1.2 1.4 1.6\nLift coefficient, $C_L$\n\n[Figure: NACA logo]\n\nFigure 9.- Effect of profile of a triangular wing of aspect ratio 2.31\non $C_{Y\\psi}$, $C_{n\\psi}$, and $C_{l\\psi}$. $\\Lambda_{c/4} = 52.2^\\circ$.", "timestamp": "2026-07-22T05:30:17.057858+00:00"}