PDF/ocr_dataset_1027.jsonl

{"citation_id": "19930082485", "source_url": "https://ntrs.nasa.gov/api/citations/19930082485/downloads/19930082485.pdf", "page_number": 43, "total_pages": 62, "image_filename": "19930082485_p43.jpg", "text": "42\nNACA TN No. 1810\n\n<!-- Image (118, 110, 905, 836) -->\n\nFigure 11.- Radial variation of axial component of velocity $\\frac{V_x}{V_{cr}}$ and weight-flow parameter at 0.1 chord downstream of cascade.", "timestamp": "2026-07-22T05:04:17.213654+00:00"}
{"citation_id": "19930082496", "source_url": "https://ntrs.nasa.gov/api/citations/19930082496/downloads/19930082496.pdf", "page_number": 29, "total_pages": 50, "image_filename": "19930082496_p29.jpg", "text": "28\n\nPage intentionally left blank\n\nPage intentionally left blank", "timestamp": "2026-07-22T05:04:20.353735+00:00"}
{"citation_id": "19930082914", "source_url": "https://ntrs.nasa.gov/api/citations/19930082914/downloads/19930082914.pdf", "page_number": 22, "total_pages": 66, "image_filename": "19930082914_p22.jpg", "text": "NACA TN No. 1857\n21\n\nBut at $z = y = 0$,\n\n$$c_1 = V = \\frac{V_0}{n(y=0)} = \\frac{V_0}{a}$$\n\nTherefore\n\n$$qc_2 = \\log_e 1 = 0$$\n\nand\n\n$$c_2 = 0$$\n\nTherefore\n\n$$\\xi = 1 + \\frac{b}{a} y$$\n\nand\n\n$$\\frac{b}{a} z = \\log_e \\left[ 1 + \\frac{b}{a} y + \\sqrt{\\left(1 + \\frac{b}{a} y\\right)^2 - 1} \\right]$$\n\nor\n\n$$e^{\\frac{b}{a} z} + e^{-\\frac{b}{a} z} = 2 \\left( 1 + \\frac{b}{a} y \\right)$$\n\nThe path of a ray, therefore, through a medium of linear density gradient is a catenary. Because $a$ is very nearly equal to unity, the equation can be simplified to\n\n$$e^{bz} + e^{-bz} = 2(1 + by)$$", "timestamp": "2026-07-22T05:04:20.610689+00:00"}
{"citation_id": "19930082614", "source_url": "https://ntrs.nasa.gov/api/citations/19930082614/downloads/19930082614.pdf", "page_number": 22, "total_pages": 36, "image_filename": "19930082614_p22.jpg", "text": "20\nNACA TN 1939\n\nTABLE I.— SUMMARY OF THE DRAG CHARACTERISTICS OF VARIOUS AERODYNAMIC BRAKES$^a$\n\n| Brake type | Description | Airplane component tested with brake | Chordwise location of brake | | Brake chord | | Brake span | | Brake angle | | $C_{D_B}$ (max area split, $\\frac{1}{2}$) | $C_{D_B}$ (max area split, $\\frac{1}{2}$) | Remarks | Ref. |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| | | | Upper surface | Lower surface | Upper surface | Lower surface | Upper surface | Lower surface | Upper surface | Lower surface | | | | |\n| A | Perforated split flaps | Rectangular wing | 0.80c | 0.80c | 0.800c | 0.800c | 0.600b | 1.000b | 90° | 90° | 0.240 | 0.240 | 1.01 | - - - | 3 |\n| A | Perforated split flaps | Rectangular wing | .80c | .80c | .800c | .800c | 1.000b | 1.000b | | | .400 | .414 | 1.03 | - - - | 3 |\n| B | Perforated split flaps | Tapered wing | .60b | .60b | .800b | .800b | .600b | .600b | 60° | 60° | .2640 | .190 | .79 | - - - | 4 |\n| B | Perforated split flaps | Tapered wing | .80b | .80b | .800b | .800b | 1.000b | 1.000b | 60° | 60° | .4400 | .389 | .88 | - - - | 4 |\n| C | Solid split flaps | Bomber model wing | .64c | .64c | .117c | .117c | .870b | .130b | 45° | 54° | .0861 | .083 | .96 | - - - | - |\n| D | Twelve small split flaps | Rectangular wing | .86c | .86c | .144c | .125c | .300b | .300b | 60° | 60° | .0805 | .070 | .89 | - - - | 5 |\n| D | Twelve small split flaps | Rectangular wing | .63c | .63c | .120c | .120c | .300b | .300b | 60° | 60° | .0612 | .091 | 1.58 | - - - | 5 |\n| E | Solid split flaps | Bomber model wing | .76c | .76c | - - - | .830c | - - - | .343b | - - - | - - - | .1450 | .093 | .64 | - - - | - |\n| F | Solid spoilers | Elliptical wing | .16c | .16c | .043c | .058c | .128b | .128b | 90° | 90° | .0165 | .041 | 2.48 | No gap (See fig. 6) | 6 |\n| G | Solid spoilers | Elliptical wing | .16c | .16c | - - - | .053c | - - - | .244b | - - - | - - - | .0150 | .051 | 3.09 | No gap | 6 |\n| G | Solid spoilers | Elliptical wing | None | .16c | - - - | .071c | - - - | .244b | - - - | - - - | .0200 | .051 | 2.38 | Gap = 0.05c | 6 |\n| G | Solid spoilers | Elliptical wing | None | .16c | - - - | .106c | - - - | .244b | - - - | - - - | .0330 | .040 | 1.21 | Gap = 0.30c | 6 |\n| G | Solid spoilers | Elliptical wing | None | .30c | - - - | .078c | - - - | .190b | - - - | - - - | .0360 | .066 | 2.06 | No gap | 6 |\n| G | Solid spoilers | Elliptical wing | None | .30c | - - - | .066c | - - - | .190b | - - - | - - - | .0630 | .069 | 1.60 | Gap = 0.25c | 6 |\n| G | Solid spoilers | Elliptical wing | None | .30c | - - - | .106c | - - - | .190b | - - - | - - - | .0640 | .069 | .88 | Gap = 0.30c | 6 |\n| H | Solid spoiler | Rectangular wing | None | .20c | - - - | .067c | - - - | 1.000b | - - - | - - - | .0667 | .114 | 1.71 | No gap | 6 |\n| I | Solid spoiler | Fighter model wing | .20c | None | .169c | - - - | .202b | - - - | 90° | - - - | .0340 | .065 | 2.02 | Gap = 0.33c | - |\n| I | Solid spoiler | Fighter model wing | None | .20c | - - - | .169c | - - - | .202b | - - - | - - - | .0340 | .056 | 1.52 | Gap = 0.33c | - |\n| I | Solid spoiler | Fighter model wing | .50c | None | .169c | - - - | .202b | - - - | 90° | - - - | .0340 | .057 | 1.67 | Gap = 0.33c | - |\n| I | Solid spoiler | Fighter model wing | None | .50c | - - - | .169c | - - - | .202b | - - - | - - - | .0340 | .044 | 1.29 | Gap = 0.33c | - |\n| J | Solid spoiler | Tapered wing | .75c | None | .080c | - - - | .400b | - - - | 90° | - - - | .0137 | .020 | 1.46 | No gap | 7 |\n| J | Solid spoiler | Tapered wing | .75c | .75c | .080c | .020c | .400b | .400b | 90° | 90° | .0274 | .038 | 1.39 | No gap | 7 |\n| K | Perforated plate | Fighter model wing | .63c | None | .350c | - - - | .470b | - - - | 90° | - - - | .1180 | .143 | 1.21 | Gap = 0.16c | - |\n| L | Double horn gap and slot | Tapered wing | .50c | None | .092c | - - - | .143b | - - - | 90° | - - - | .0131 | .012 | .92 | Gap = 0.34c | - |\n| L | Double horn gap and slot | Tapered wing | None | .50c | - - - | .092c | - - - | .143b | - - - | - - - | .0125 | .015 | 1.20 | Gap = 0.34c | - |\n| M | Spoiler with vertical slots | Tapered wing | .56c | None | .092c | - - - | .143b | - - - | 90° | - - - | .0131 | .009 | .69 | Gap = 0.34c | - |\n| M | Spoiler with vertical slots | Tapered wing | None | .56c | - - - | .092c | - - - | .143b | - - - | - - - | .0125 | .011 | .88 | Gap = 0.34c | - |\n| N | Frise/Coon-type brake | Bomber model wing | - - - | - - - | - - - | - - - | .165b | .165b | 90° | 90° | .1200 | .041 | .34 | - - - | - |\n| O | Fuselage dive brakes | Bomber fuselage aft of wing | - - - | - - - | - - - | - - - | - - - | - - - | - - - | 90° | .0444 | .044 | 1.04 | - - - | - |\n| P | Fuselage dive-recovery brake | Bottom of fighter fuselage | - - - | - - - | - - - | - - - | - - - | - - - | - - - | 80° | .0336 | .033 | .62 | - - - | - |\n| Q | Fuselage side brakes | Bomber fuselage | - - - | - - - | - - - | - - - | - - - | - - - | - - - | 60° | .0270 | .041 | .72 | - - - | - |\n| R | Fuselage side brakes | Side of fighter fuselage | - - - | - - - | - - - | - - - | - - - | - - - | - - - | 80° | .0516 | .051 | 1.02 | - - - | - |\n\n$^a$ Data are for zero lift; Mach number less than 0.3.\n$^b$ See figure 6.\n$^c$ c = local chord of wing.\n$^d$ $\\frac{1}{2}$ = wing chord at mean spanwise station of brake.\n$^e$ $c_F$ = root chord of wing.\n$^f$ percent: upper surface, 0.50b; lower surface, 0.43b.\n\n[Figure: NACA logo]", "timestamp": "2026-07-22T05:04:21.751557+00:00"}
{"citation_id": "19930082542", "source_url": "https://ntrs.nasa.gov/api/citations/19930082542/downloads/19930082542.pdf", "page_number": 23, "total_pages": 53, "image_filename": "19930082542_p23.jpg", "text": "22\nNACA TN No. 1867\n\nTABLE I.- ROOM-TEMPERATURE PHYSICAL PROPERTIES OF LOW-CARBON M-155 BAR STOCK - Concluded\n\n<!-- Table (100, 120, 888, 935) -->\n\\begin{tabular}{|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|}\n\\hline\n\\multicolumn{3}{|c|}{Heat treatment} & \\multicolumn{2}{c|}{Aging treatment} & \\multicolumn{2}{c|}{Re-cold-rolling} & \\multirow{3}{*}{Brinell hardness} & \\multicolumn{8}{c|}{Room-temperature tensile properties} \\\\\n\\cline{1-7}\\cline{9-16}\n\\multicolumn{2}{|c|}{Solution treatment} & \\multirow{2}{*}{Method of cooling (c)} & \\multirow{2}{*}{Temper-ature ($^\\circ$F)} & \\multirow{2}{*}{Time (hr)} & \\multirow{2}{*}{Temper-ature ($^\\circ$F)} & \\multirow{2}{*}{Percent reduc-tion} & & \\multicolumn{2}{c|}{Tensile strength (psi)} & \\multicolumn{3}{c|}{Offset yield strength (psi)} & \\multirow{2}{*}{Propor-tional limit (psi)} & \\multirow{2}{*}{Elonga-tion in 2 in. (percent)} & \\multirow{2}{*}{Reduction of area (percent)} \\\\\n\\cline{1-2}\\cline{9-11}\nTemper-ature ($^\\circ$F) & Time (hr) & & & & & & & 0.02 percent & 0.1 percent & 0.2 percent & & & & & \\\\\n\\hline\n2150 & 1 & W.Q. & --- & --- & --- & --- & 193 & 116,950 & 30,000 & 47,000 & 53,000 & 10,000 & 45.5 & 62.2 \\\\\n\\hline\n\\multicolumn{17}{|c|}{Solution-treated at 2150$^\\circ$ F} \\\\\n\\hline\n2200 & 1 & W.Q. & --- & --- & --- & --- & 180 & 118,300 & 48,000 & 54,500 & 58,000 & 40,000 & 54 & 64.8 \\\\\n2200 & 1 & W.Q. & --- & --- & --- & --- & 205 & 115,750 & 42,000 & 53,000 & 57,000 & 22,000 & 46.5 & 64.3 \\\\\n2200 & 1 & A.C. & --- & --- & --- & --- & 175 & 118,950 & 42,000 & 48,500 & 55,000 & 30,000 & 54 & 64.6 \\\\\n\\hline\n\\multicolumn{17}{|c|}{Solution-treated at 2200$^\\circ$ F} \\\\\n\\hline\n2200 & 1 & W.Q. & 1400 & 2 & --- & --- & 182 & 117,250 & 47,000 & 57,000 & 61,000 & 40,000 & 42 & 42.4 \\\\\n2200 & 1 & W.Q. & 1400 & 8 & --- & --- & 185 & --- & --- & --- & --- & --- & --- & --- \\\\\n2200 & 1 & W.Q. & 1400 & 24 & --- & --- & 221 & 119,250 & 48,000 & 61,000 & 65,000 & 23,000 & 32 & 36.8 \\\\\n2200 & 1 & W.Q. & 1400 & 24 & --- & --- & 221 & 119,250 & 48,000 & 61,000 & 65,000 & 23,000 & 32 & 36.8 \\\\\n2200 & 1 & W.Q. & 1350 & 2 & --- & --- & 206 & 120,125 & 54,000 & 61,000 & 64,000 & 45,000 & 39.5 & 40.9 \\\\\n2200 & 1 & W.Q. & 1350 & 8 & --- & --- & 213 & 118,000 & 50,000 & 57,000 & 60,000 & 32,000 & 36.5 & 44.9 \\\\\n2200 & 1 & W.Q. & 1350 & 24 & --- & --- & 213 & 118,950 & 42,500 & 51,000 & 57,250 & 32,000 & 31.5 & 43.5 \\\\\n2200 & 1 & W.Q. & 1600 & 2 & --- & --- & 219 & 119,250 & 42,500 & 53,000 & 57,500 & 22,500 & 31.5 & 34.3 \\\\\n2200 & 1 & W.Q. & 1750 & 2 & --- & --- & 197 & 115,400 & 43,000 & 55,000 & 55,500 & 35,000 & 39 & 36.1 \\\\\n\\hline\n\\multicolumn{17}{|c|}{Cold-working amount and temperature:} \\\\\n\\hline\n2200 & 1 & W.Q. & --- & --- & 1200 & 5 & 224 & 122,000 & 69,000 & 76,500 & 80,000 & 40,000 & 40.5 & 29.5 \\\\\n2200 & 1 & W.Q. & --- & --- & 1200 & 10 & 236 & 125,000 & 93,000 & 102,500 & 107,000 & 74,000 & 31.5 & 25.7 \\\\\n2200 & 1 & W.Q. & 1400 & 24 & 1200 & 5 & 289 & 136,000 & 99,000 & 105,000 & 107,000 & 70,000 & 31.5 & 27.7 \\\\\n2200 & 1 & W.Q. & --- & --- & 1200 & 15 & 289 & 142,000 & 97,000 & 112,000 & 117,000 & 75,000 & 28.5 & 30.1 \\\\\n2200 & 1 & W.Q. & 1400 & 24 & 1200 & 15 & 291 & 146,000 & 101,000 & 115,000 & 118,000 & 80,000 & 28.5 & 32.7 \\\\\n2200 & 1 & W.Q. & --- & --- & 1400 & 15 & 289 & 144,000 & 100,000 & 113,000 & 117,000 & 80,000 & 24.5 & 37.8 \\\\\n2200 & 1 & W.Q. & --- & --- & 1400 & 15 & 289 & 140,500 & 94,000 & 111,000 & 116,000 & 61,500 & 26.5 & 46.2 \\\\\n2200 & 1 & W.Q. & --- & --- & 1400 & 15 & 289 & 128,500 & 94,000 & 95,000 & 96,000 & 68,500 & 28 & 40.7 \\\\\n\\hline\n\\multicolumn{17}{|c|}{Solution-treated at 2250$^\\circ$ F} \\\\\n\\hline\n2250 & 1/2 & W.Q. & --- & --- & --- & --- & 178 & 116,000 & 44,500 & 52,000 & 55,000 & 30,000 & 53 & 64.1 \\\\\n\\hline\n\\multicolumn{17}{|c|}{Solution-treated at 2300$^\\circ$ F} \\\\\n\\hline\n2300 & 1/2 & W.Q. & --- & --- & --- & --- & 163 & 116,250 & 48,000 & 53,000 & 57,000 & 40,000 & 56.5 & 60.8 \\\\\n\\hline\n\\end{tabular}\n\n$^1$All solution-treatment periods were 1 hour unless noted otherwise.\n$^2$All hot-rolled material was given a final stress relief at 1200$^\\circ$ F for 1 hr.\n$^3$W.Q., water-quenched; A.C., air-cooled.\n$^4$Aged after rolling.\n\nNACA", "timestamp": "2026-07-22T05:04:22.170140+00:00"}
{"citation_id": "19930085519", "source_url": "https://ntrs.nasa.gov/api/citations/19930085519/downloads/19930085519.pdf", "page_number": 15, "total_pages": 46, "image_filename": "19930085519_p15.jpg", "text": "14\nNACA RM No. L8K19\n\nREFERENCES\n\n1. Schneiter, Leslie E., and Watson, James M.: Low-Speed Wind-Tunnel Investigation of Various Plain-Spoiler Configurations for Lateral Control on a 42° Sweptback Wing. NACA TN No. 1646, 1948.\n\n2. Wenzinger, Carl J., and Rogallo, Francis M.: Wind-Tunnel Investigation of Spoiler, Deflector, and Slot Lateral-Control Devices on Wings with Full-Span Split and Slotted Flaps. NACA Rep. No. 706, 1941.\n\n3. Rogallo, Francis M., and Swenson, Robert S.: Wind-Tunnel Development of a Plug-Type Spoiler-Slot Aileron for a Wing with a Full-Span Slotted Flap and a Discussion of Its Application. NACA ARR, Nov. 1941.\n\n4. Wenzinger, Carl J.: The Effects of Full-Span and Partial-Span Split Flaps on the Aerodynamic Characteristics of a Tapered Wing. NACA TN No. 505, 1934.\n\n5. House, R. O.: The Effects of Partial-Span Plain Flaps on the Aerodynamic Characteristics of a Rectangular and a Tapered Clark Y Wing. NACA TN No. 663, 1938.\n\n6. House, Rufus O.: The Effects of Partial-Span Slotted Flaps on the Aerodynamic Characteristics of a Rectangular and a Tapered N.A.C.A. 23012 Wing. NACA TN No. 719, 1939.\n\n7. Lowry, John G., and Schneiter, Leslie E.: Estimation of Effectiveness of Flap-Type Controls on Sweptback Wings. NACA TN No. 1674, 1948.\n\n8. Pitkin, Marvin, and Maggin, Bernard: Analysis of Factors Affecting Net Lift Increment Attainable with Trailing-Edge Split Flaps on Tailless Airplanes. NACA ARR No. L4I18, 1944.\n\n9. Fischel, Jack, and Schneiter, Leslie E.: High-Speed Wind-Tunnel Investigation of an NACA 65-210 Semispan Wing Equipped with Plug and Retractable Ailerons and a Full-Span Slotted Flap. NACA TN No. 1663, 1948.\n\n10. Rogallo, F. M., and Spano, Bartholomew S.: Wind-Tunnel Investigation of a Spoiler-Slot Aileron on an NACA 23012 Airfoil with a Full-Span Fowler Flap. NACA ARR, Dec. 1941.", "timestamp": "2026-07-22T05:04:28.269252+00:00"}
{"citation_id": "19930085471", "source_url": "https://ntrs.nasa.gov/api/citations/19930085471/downloads/19930085471.pdf", "page_number": 19, "total_pages": 28, "image_filename": "19930085471_p19.jpg", "text": "UNCLASSIFIED\nCONFIDENTIAL\nRESTRICTED\n\nNACA RM No. L8J11\n\n[Figure: A circular test section of a wind tunnel with a flutter model installed inside. The model is mounted on a support structure within the tunnel. The image shows the interior of the tunnel with visible structural elements and mounting hardware.]\n\nNACA\nL-56781\n\nFigure 4.- Flutter model installed in the test section.\n\nRESTRICTED\nCONFIDENTIAL\nUNCLASSIFIED\n\n17", "timestamp": "2026-07-22T05:04:30.424800+00:00"}
{"citation_id": "19930082498", "source_url": "https://ntrs.nasa.gov/api/citations/19930082498/downloads/19930082498.pdf", "page_number": 28, "total_pages": 49, "image_filename": "19930082498_p28.jpg", "text": "NACA TN No. 1838\n\n27\n\n1/16\" thick perforated sheet\n3/16\" on centers\n1/8\" diameter holes\n7/32\" on centers\n\nPerforated sheet steel used in mufflers 18, 19, 40, 43.\n\n48\n\n59\n\n[Figure: NACA logo]\n\nFigure 1.— Concluded.", "timestamp": "2026-07-22T05:04:40.859467+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 35, "total_pages": 99, "image_filename": "19930082511_p35.jpg", "text": "NACA TN No. 1826\n33\n\nTransposing terms in this equation gives\n\n$$\n\\frac{a^2}{l^2 \\sqrt{(1 - l^2)(a^2 - l^2)}} = \\sqrt{\\frac{a^2}{(1 - l^2)(a^2 - l^2)}} - \\sqrt{\\frac{a^2 - l^2}{1 - l^2}} - \\frac{d}{dl} \\frac{\\sqrt{(1 - l^2)(a^2 - l^2)}}{l}\n$$\n\nwhence\n\n$$\nNa^2 \\int_0^1 \\frac{dl}{l^2 \\sqrt{(1 - l^2)(a^2 - l^2)}} = Na^2 \\int_0^1 \\frac{dl}{\\sqrt{(1 - l^2)(a^2 - l^2)}} - N \\int_0^1 \\sqrt{\\frac{a^2 - l^2}{1 - l^2}} \\, dl - N \\frac{\\sqrt{(1 - l^2)(a^2 - l^2)}}{l} \\Bigg|_0^1\n$$\n\n$$\n= NaK - NaE - N \\frac{\\sqrt{(1 - l^2)(a^2 - l^2)}}{l} \\Bigg|_0^1\n$$\n\nThe first two terms on the right are exactly canceled by the two terms previously obtained, so that the final result is:\n\n$$\n\\lim_{z \\to \\infty} R.P.Nw_2(z) = \\lim_{l \\to 0} -N \\frac{\\sqrt{(1 - l^2)(a^2 - l^2)}}{l} \\Bigg|_0^1\n$$\n\n$$\n= \\lim_{z \\to \\infty} Nz \\sqrt{\\left(1 - \\frac{a^2}{z^2}\\right)\\left(1 - \\frac{1}{z^2}\\right)}\n$$\n\n$$\n= \\lim_{z \\to \\infty} Nz - \\frac{N}{z} \\left(\\frac{a^2 + 1}{2}\\right) \\cdot \\cdot \\cdot\n$$", "timestamp": "2026-07-22T05:04:44.663976+00:00"}
{"citation_id": "19930082245", "source_url": "https://ntrs.nasa.gov/api/citations/19930082245/downloads/19930082245.pdf", "page_number": 45, "total_pages": 66, "image_filename": "19930082245_p45.jpg", "text": ".24\n.20\n.16\n.12\n.08\n.04\n0\n-.04\n-.08\n-.12\n-.16\n-.20\n.1 .2 .3 .4 .5 .6 .7 .8 .9\nMach number, M\nSection pitching-moment coefficient, $C_m$\n$\\delta_a$\n(deg)\n12\n6\n4\n2\n0\n-2\n-4\n-6\n12\n18\n30\n-4\n-2\n0\n12\n18\nNACA\n8\n6\n4\n2\n0\n-2\n-4\n-6\n-8\n-10\n-12\n-14\n.1 .2 .3 .4 .5 .6 .7 .8 .9\nMach number, M\nSection angle of attack, $\\alpha$, deg\n$\\delta_a$\n(deg)\n12\n6\n4\n2\n0\n-2\n-4\n-6\n12\n18\n30\n-4\n-2\n0\n2\n4\n(c) $C_h = 0$.\nFigure 9. - Continued.\nNACA TN No. 1596\n44", "timestamp": "2026-07-22T05:04:45.148735+00:00"}
{"citation_id": "19930085544", "source_url": "https://ntrs.nasa.gov/api/citations/19930085544/downloads/19930085544.pdf", "page_number": 6, "total_pages": 33, "image_filename": "19930085544_p6.jpg", "text": "NACA RM No. L8K26\n\n$\\omega$ angular velocity of propeller, radians per second $(2 \\pi n)$\n\nA dot over a quantity denotes the first derivative of the quantity with respect to time; two dots, the second derivative with respect to time.\n\nFORCES ON AN INCLINED PROPELLER\n\nThe Velocity Diagram\n\nFigure 1(a) shows a side view of a propeller disk the thrust axis of which is inclined at an angle $\\alpha_{\\mathrm{T}}$ to the forward velocity $V$. This forward velocity $V$ is shown resolved into a component $V \\cos \\alpha_{\\mathrm{T}}$ perpendicular to the plane of rotation and a component $V \\sin \\alpha_{\\mathrm{T}}$ parallel to the plane of rotation. Figure 1(b), a view perpendicular to the plane of rotation along the thrust axis, shows the velocity component $V \\sin \\alpha_{\\mathrm{T}}$ at a section of a propeller blade which is located at a position $\\omega t$ on the propeller disk. In this paper the time variable $\\omega t$, which defines the position of the blade, is considered to be zero when the blade is initially vertical upward with the propeller axis in positive pitch and is measured in the direction of rotation (these axes may be rotated to comply with propeller attitudes other than pitch). With this convention the vector $V \\sin \\alpha_{\\mathrm{T}}$ may be resolved into a component $V \\sin \\alpha_{\\mathrm{T}} \\sin \\omega t$ in the direction of the tangential velocity $\\omega r$ and a component $V \\sin \\alpha_{\\mathrm{T}} \\cos \\omega t$ in a radial direction along the blade. In the treatment that follows it is assumed that the radial component of the flow ($V \\sin \\alpha_{\\mathrm{T}} \\cos \\omega t$) has a negligible effect on the airfoil characteristics. With this assumption, it remains to determine the effect of the periodic change in the rotational velocity ($\\pi n D x + V \\sin \\alpha_{\\mathrm{T}} \\sin \\omega t$) and the component velocity $V \\cos \\alpha_{\\mathrm{T}}$ on the propeller characteristics.\n\nThe vector diagram for a section of an inclined propeller is shown in figure 1(c). In this figure the induced effects are not included. It should be realized, however, that the aerodynamic helix angles will be somewhat different from the geometric helix angles shown. From figure 1(c) the geometric helix angle for any position of the propeller blade is given by\n\n$$\n\\phi_{\\omega t} = \\tan^{-1} \\frac{V \\cos \\alpha_{\\mathrm{T}}}{\\pi n D x + V \\sin \\alpha_{\\mathrm{T}} \\sin \\omega t}\n$$", "timestamp": "2026-07-22T05:04:45.490242+00:00"}
{"citation_id": "19930082476", "source_url": "https://ntrs.nasa.gov/api/citations/19930082476/downloads/19930082476.pdf", "page_number": 41, "total_pages": 41, "image_filename": "19930082476_p41.jpg", "text": "NACA TN No. 1801\n39\n\n[Figure: Five vertical strips of sequential photographs showing the motion of a model airplane. The strips are labeled with numbers at the bottom: 12, 24, 36, 48, 60. Intermediate frames are labeled 6, 18, 30, 42, 54.]\n\nFigure 7.- Typical motion of the model with elevator deflected to $13^\\circ$ up and wheel set one-half with the spin (loading 2). Pictures taken at 64 frames per second.\nNACA", "timestamp": "2026-07-22T05:04:45.723104+00:00"}
{"citation_id": "19930082585", "source_url": "https://ntrs.nasa.gov/api/citations/19930082585/downloads/19930082585.pdf", "page_number": 28, "total_pages": 30, "image_filename": "19930082585_p28.jpg", "text": "NACA TN 1907\n27\n\nInflow factor, $\\lambda$\n\nBlade stalling should be considered above this boundary\n\nMoment of inertia, $I_1$, slug-ft$^2$\n\nAngle of incidence, $\\theta$, deg\n\nFigure 8.- Effect of blade moment of inertia on $\\lambda$ against $\\theta$ throughout the transition maneuver. Slow exponential pitch change.", "timestamp": "2026-07-22T05:04:49.730324+00:00"}
{"citation_id": "19930085626", "source_url": "https://ntrs.nasa.gov/api/citations/19930085626/downloads/19930085626.pdf", "page_number": 2, "total_pages": 24, "image_filename": "19930085626_p2.jpg", "text": "NACA RM No. L8K23 CONFIDENTIAL\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nRESEARCH MEMORANDUM\n\nFREE-FLIGHT INVESTIGATION AT TRANSONIC AND SUPERSONIC  \nSPEEDS OF THE ROLLING EFFECTIVENESS OF SEVERAL  \nAILERON CONFIGURATIONS ON A TAPERED  \nWING HAVING $42.7^\\circ$ SWEEPBACK  \n\nBy Carl A. Sandahl\n\nSUMMARY\n\nAn investigation was made of several aileron modifications in conjunction with a tapered, sweptback wing having circular-arc airfoil sections of relatively large thickness ratio. The modifications, all of which effectively reduced the aileron trailing-edge angle, included a straight-side extended-chord aileron, a parallel-side aileron of thickness equal to the wing thickness at the hinge line and having blunt trailing edges, and a straight-side aileron having a blunt trailing edge of thickness equal to one-half of the wing thickness at the hinge line. The modified ailerons eliminated the reversal of rolling effectiveness obtained with the true-contour ailerons at small deflections.\n\nINTRODUCTION\n\nIn the early part of 1948, as part of an investigation of wing-aileron rolling effectiveness utilizing rocket-propelled test vehicles in free flight, tests were made of an 0.2-chord, outboard half-semispan, true-contour aileron in conjunction with a wing which had an aspect ratio of 4.00, a taper ratio of 0.5, a sweepback angle of $40^\\circ$ measured at the quarter-chord line, and circular-arc airfoil sections of 10-percent-thickness ratio normal to the quarter-chord line. The above tests, which are reported in reference 1, indicated reversal of the rolling effectiveness in the Mach number range from 0.94 to 1.00 for small aileron deflections. It was believed that the relatively large trailing-edge angle of the circular-arc sections contributed to the observed reversal of effectiveness.\n\nFollowing the previously mentioned tests, an investigation was conducted by means of the \"transonic bump\" technique in the Langley 7-by 10-foot high-speed tunnel for the purpose of developing an aileron for the wing configuration used in the free-flight tests which would not\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:04:56.237363+00:00"}
{"citation_id": "19930085536", "source_url": "https://ntrs.nasa.gov/api/citations/19930085536/downloads/19930085536.pdf", "page_number": 14, "total_pages": 20, "image_filename": "19930085536_p14.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:04:56.499159+00:00"}
{"citation_id": "19930085487", "source_url": "https://ntrs.nasa.gov/api/citations/19930085487/downloads/19930085487.pdf", "page_number": 18, "total_pages": 36, "image_filename": "19930085487_p18.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:04:57.354484+00:00"}
{"citation_id": "19930082450", "source_url": "https://ntrs.nasa.gov/api/citations/19930082450/downloads/19930082450.pdf", "page_number": 33, "total_pages": 37, "image_filename": "19930082450_p33.jpg", "text": "32\nNACA TN No. 1778\n\n$$ \\frac{H}{t_w} = 21 $$\n$$ \\left( \\frac{D_w}{t_w} = 20 \\right) $$\n\n$$ \\sigma_{cr}, ksi $$\n12.7\n8.3\n\n$$ \\frac{P_l}{L\\sqrt{E}}, ksi $$\n\nColors indicate minimum weight\nproportions for $$ \\frac{t_w}{t_s} = 0.79 $$.\nRed means some other,\nblue means no other value\nof $$ \\frac{t_w}{t_s} $$ gives less weight.\n\n$$ \\sigma_f, ksi $$\n\n$$ \\frac{31}{(30)} $$\n\n$$ \\sigma_{cr}, ksi $$\n17.6\n12.7\n8.3\n\n$$ \\frac{P_l}{L\\sqrt{E}}, ksi $$\n\n$$ \\frac{41}{(40)} $$\n\n$$ \\sigma_{cr}, ksi $$\n17.6\n12.7\n8.3\n\n$$ \\frac{P_l}{L\\sqrt{E}}, ksi $$\n\nNACA\n\n$$ \\frac{P_l}{t_s}, ksi $$\n\nFigure 8.—Direct-reading design chart (alternate form) for 24S-T aluminum-alloy Z-stiffened panels. $$ \\frac{t_w}{t_s} = 0.79 $$.", "timestamp": "2026-07-22T05:04:58.608292+00:00"}
{"citation_id": "19930085542", "source_url": "https://ntrs.nasa.gov/api/citations/19930085542/downloads/19930085542.pdf", "page_number": 8, "total_pages": 46, "image_filename": "19930085542_p8.jpg", "text": "6\nNACA RM No. L8L29\n\nThe angle of attack, longitudinal-force coefficient, and rolling-moment coefficient were corrected for jet boundary effects, but corrections were not applied to account for model blocking, which amounts to an error of about 1.5 percent in dynamic pressure.\n\nIn the rolling-flow tests, tares appeared to be negligible up to approximately $\\alpha = 16^\\circ$. However, at high angles of attack, there appeared to be large support interference, and since these effects could not be accurately evaluated, the rolling derivatives are not presented for angles of attack greater than approximately $16^\\circ$.\n\nThe measurements taken are believed to be accurate within the following amounts which are based on the maximum values of the forces and moments for model 6:\n\n| | |\n| :--- | :--- |\n| $\\alpha$, deg | $\\pm 0.1$ |\n| $\\psi$, deg | $\\pm 0.2$ |\n| $C_L$ | $\\pm 0.0029$ |\n| $C_X$ | $\\pm 0.0045$ |\n| $C_m$ | $\\pm 0.0045$ |\n| $C_l$ | $\\pm 0.0004$ |\n| $C_n$ | $\\pm 0.0003$ |\n| $C_Y$ | $\\pm 0.0046$ |\n\nRESULTS AND DISCUSSION\n\nPresentation of Results\n\nThe static and rolling characteristics of the models of the present investigation are presented in the four groups of basic data and the two summary groups shown in the following table:\n\n| | Figure |\n| :--- | :--- |\n| Effect of profile of triangular wings | 8, 9, 10 |\n| Effect of aspect ratio of triangular wings | 11, 12, 13 |\n| Effect of vertical fins | 14, 15, 16 |\n| Effect of aspect ratio of modified triangular wings | 17, 18, 19 |\n| Summary effects of aspect ratio of triangular wings | 20, 21, 22 |\n| Summary effects of aspect ratio of modified triangular wings | 23, 24, 25 |\n\nAll theoretical values obtained from reference 5 have been calculated by use of the equations and have been extrapolated to the appropriate taper ratios.", "timestamp": "2026-07-22T05:04:59.157048+00:00"}
{"citation_id": "19930082485", "source_url": "https://ntrs.nasa.gov/api/citations/19930082485/downloads/19930082485.pdf", "page_number": 44, "total_pages": 62, "image_filename": "19930082485_p44.jpg", "text": "NACA TN No. 1810\n43\n\n1026\n\nRadial pressure-gradient parameter, $\\frac{1}{F_t} \\frac{dP_s}{dr}$\nor centrifugal-force parameter, $\\frac{\\rho_s}{\\rho_t} \\left(\\frac{V_u}{V_{cr}}\\right)^2$\n\n$\\frac{\\rho_s}{\\rho_t} \\left(\\frac{V_u}{V_{cr}}\\right)^2$\n\n$\\frac{r}{V_t} \\frac{dP_s}{dr}$\n\n$\\frac{r}{V_{cr2}} \\frac{\\rho_s}{\\rho_t} \\frac{dV_r}{dt}$\n\nRadial-acceleration parameter, $\\frac{r}{V_{cr}^2} \\frac{\\rho_s}{\\rho_t} \\frac{dV_r}{dt}$\n\nInner shroud\nOuter shroud\n\nExperimental\nDesign\n\nRadius, r, in.\n\nFigure 12.- Radial equilibrium parameters.", "timestamp": "2026-07-22T05:05:01.610719+00:00"}
{"citation_id": "19930082496", "source_url": "https://ntrs.nasa.gov/api/citations/19930082496/downloads/19930082496.pdf", "page_number": 30, "total_pages": 50, "image_filename": "19930082496_p30.jpg", "text": "NACA TN No. 1836\n29\n\n[Figure: A circular mechanical assembly with labeled parts. Labels point to: \"Nozzle-box thermocouple leads\", \"Inlet-gas thermocouple\", \"Turbine blades\", \"Turbine disk\", \"Asbestos sheet\", and \"Fragment shield\". In the bottom right corner of the figure is a NACA logo with text \"C-21895\" and \"7-27-48\".]\n\nFigure 5. - Quasi-service evaluation unit showing shield and asbestos sheet used to preserve blade fragments.", "timestamp": "2026-07-22T05:05:02.224338+00:00"}
{"citation_id": "19930082617", "source_url": "https://ntrs.nasa.gov/api/citations/19930082617/downloads/19930082617.pdf", "page_number": 24, "total_pages": 58, "image_filename": "19930082617_p24.jpg", "text": "NACA TN 1962\n23\n\nStringers\no 1 to 9\nx 10 to 16\n\nMoment\n(in. - lb)\n1 36.0 X $10^3$\n2 72.0 X $10^3$\n3 108.0 X $10^3$\n4 144.0 X $10^3$\n5 216.0 X $10^3$\n6 288.0 X $10^3$\n\n2.57\"\nA\nA\nBond L\n45°\nA-A\n\nDistance from horizontal diameter, in.\n10\n9\n8\n7\n6\n5\n4\n3\n2\n1\n0\n1\n2\n3\n4\n5\n6\n7\n8\n9\n10\n\n20 16 12 8 4 0 -4 -8 -12 -16 -20 X $10^{-4}$\nStrain\n\n1 2 3 4 5 6\n\nNACA\n\nFigure 12.- Strain diagram of cylinder 75. Band L.", "timestamp": "2026-07-22T05:05:07.092068+00:00"}
{"citation_id": "19930085471", "source_url": "https://ntrs.nasa.gov/api/citations/19930085471/downloads/19930085471.pdf", "page_number": 20, "total_pages": 28, "image_filename": "19930085471_p20.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:05:10.960328+00:00"}
{"citation_id": "19930085519", "source_url": "https://ntrs.nasa.gov/api/citations/19930085519/downloads/19930085519.pdf", "page_number": 16, "total_pages": 46, "image_filename": "19930085519_p16.jpg", "text": "```markdown\nNACA RM No. L59K19\n\nPlane of\nsymmetry\n\nOrigin of axes\n\nY-axis\n\n42.00\n37.22\n11.44\n\nX-axis\n\n40°\n42°\n\nM.A.C.\n34.704\n\n28.56\n\n68.25\n\n90°\n\nNACA 64-112 AIRFOIL\n\n0.272 c line (1/4-chord line\nof unswept wing)\n\nNACA\n\n7.15\n\n26.25\n\n61.56\n\nFigure 1.- The 42° sweptback wing. Area, 32.24 square feet; aspect\nratio, 4.01; taper ratio, 0.625. All dimensions are in inches\nunless otherwise noted.\n\n15\n```", "timestamp": "2026-07-22T05:05:19.664347+00:00"}
{"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 22, "total_pages": 78, "image_filename": "19930082618_p22.jpg", "text": "20\nNACA TN 1945\n\nTABLE VIII\nORDINATES OF THE\nNACA 632-415 AIRFOIL SECTION\n[Stations and ordinates given in\npercent of airfoil chord]\n\nTABLE IX\nORDINATES OF THE\nNACA 652-415 AIRFOIL SECTION\n[Stations and ordinates given in\npercent of airfoil chord]\n\n| Upper surface | | Lower surface | |\n| :--- | :--- | :--- | :--- |\n| Station | Ordinate | Station | Ordinate |\n| 0 | 0 | 0 | 0 |\n| .300 | 1.237 | .700 | -1.087 |\n| .525 | 1.595 | .875 | -1.205 |\n| .991 | 2.074 | 1.269 | -1.248 |\n| 2.198 | 2.964 | 2.202 | -2.120 |\n| 4.660 | 4.264 | 3.310 | -2.500 |\n| 7.147 | 5.221 | 4.653 | -2.565 |\n| 9.647 | 6.077 | 10.353 | -4.009 |\n| 14.669 | 7.343 | 15.351 | -4.456 |\n| 19.709 | 8.279 | 20.295 | -5.093 |\n| 24.750 | 8.941 | 25.250 | -5.361 |\n| 29.800 | 9.362 | 30.200 | -5.474 |\n| 34.852 | 9.599 | 35.148 | -5.459 |\n| 39.905 | 9.527 | 40.095 | -5.243 |\n| 44.953 | 9.229 | 45.045 | -4.909 |\n| 50.000 | 8.874 | 50.000 | -4.459 |\n| 55.039 | 8.298 | 54.961 | -3.518 |\n| 60.070 | 7.595 | 59.930 | -3.311 |\n| 65.093 | 6.760 | 64.907 | -2.560 |\n| 70.106 | 5.877 | 69.894 | -1.989 |\n| 75.109 | 4.907 | 74.891 | -1.327 |\n| 80.102 | 3.900 | 79.898 | -.716 |\n| 85.089 | 2.885 | 84.915 | -.153 |\n| 90.059 | 1.834 | 89.941 | .104 |\n| 95.028 | .931 | 94.972 | .333 |\n| 100.000 | 0 | 100.000 | 0 |\n\n| Upper surface | | Lower surface | |\n| :--- | :--- | :--- | :--- |\n| Station | Ordinate | Station | Ordinate |\n| 0 | 0 | 0 | 0 |\n| .313 | 1.208 | .637 | -1.008 |\n| .542 | 1.490 | .958 | -1.200 |\n| 1.016 | 1.900 | 1.404 | -1.472 |\n| 2.231 | 2.690 | 2.769 | -1.936 |\n| 4.697 | 3.863 | 5.305 | -2.599 |\n| 7.184 | 4.796 | 7.816 | -3.098 |\n| 9.682 | 5.578 | 10.318 | -3.510 |\n| 14.697 | 6.842 | 15.305 | -4.150 |\n| 19.726 | 7.809 | 20.274 | -4.625 |\n| 24.764 | 8.550 | 25.236 | -4.970 |\n| 29.807 | 9.093 | 30.193 | -5.205 |\n| 34.854 | 9.455 | 35.146 | -5.323 |\n| 39.903 | 9.639 | 40.097 | -5.333 |\n| 44.953 | 9.617 | 45.047 | -5.237 |\n| 50.000 | 9.374 | 50.000 | -4.962 |\n| 55.043 | 8.910 | 54.957 | -4.530 |\n| 60.079 | 8.260 | 59.921 | -3.976 |\n| 65.106 | 7.462 | 64.894 | -3.342 |\n| 70.124 | 6.542 | 69.876 | -2.654 |\n| 75.131 | 5.532 | 74.869 | -1.952 |\n| 80.126 | 4.447 | 79.874 | -1.263 |\n| 85.109 | 3.320 | 84.891 | -.628 |\n| 90.080 | 2.175 | 89.920 | -.107 |\n| 95.040 | 1.058 | 94.960 | .206 |\n| 100.000 | 0 | 100.000 | 0 |\n\nL.E. radius: 1.594\nSlope of radius through L.E.: 0.168\n\nL.E. radius: 1.505\nSlope of radius through L.E.: 0.168\n\nTABLE X\nORDINATES OF THE\nNACA 662-415 AIRFOIL SECTION\n[Stations and ordinates given in\npercent of airfoil chord]\n\n| Upper surface | | Lower surface | |\n| :--- | :--- | :--- | :--- |\n| Station | Ordinate | Station | Ordinate |\n| 0 | 0 | 0 | 0 |\n| .314 | 1.226 | .686 | -1.006 |\n| .544 | 1.467 | .956 | -1.197 |\n| 1.019 | 1.873 | 1.401 | -1.443 |\n| 2.241 | 2.632 | 2.739 | -1.880 |\n| 4.711 | 3.748 | 5.289 | -2.494 |\n| 7.199 | 4.617 | 7.801 | -2.961 |\n| 9.696 | 5.381 | 10.304 | -3.343 |\n| 14.709 | 6.624 | 15.291 | -3.932 |\n| 19.726 | 7.581 | 20.284 | -4.397 |\n| 24.771 | 8.329 | 25.229 | -4.749 |\n| 29.812 | 8.897 | 30.188 | -5.009 |\n| 34.857 | 9.309 | 35.143 | -5.139 |\n| 39.904 | 9.574 | 40.096 | -5.207 |\n| 44.952 | 9.685 | 45.048 | -5.205 |\n| 50.000 | 9.656 | 50.000 | -5.244 |\n| 55.046 | 9.475 | 54.954 | -5.053 |\n| 60.090 | 9.100 | 59.910 | -4.816 |\n| 65.126 | 8.451 | 64.874 | -4.321 |\n| 70.150 | 7.513 | 69.840 | -3.730 |\n| 75.162 | 6.413 | 74.830 | -2.939 |\n| 80.159 | 5.187 | 79.841 | -2.003 |\n| 85.139 | 3.972 | 84.861 | -1.180 |\n| 90.104 | 2.519 | 89.896 | -.421 |\n| 95.053 | 1.196 | 94.947 | .063 |\n| 100.000 | 0 | 100.000 | 0 |\n\nL.E. radius: 1.435\nSlope of radius through L.E.: 0.168\n\nNACA", "timestamp": "2026-07-22T05:05:22.753316+00:00"}
{"citation_id": "19930082585", "source_url": "https://ntrs.nasa.gov/api/citations/19930082585/downloads/19930082585.pdf", "page_number": 29, "total_pages": 30, "image_filename": "19930082585_p29.jpg", "text": "28\nNACA TN 1907\n\nDescending velocity, V, ft/sec\nAltitude lost, ft\n\nType of pitch change (fig. 2)\nInstantaneous\nExponential (moderate)\nExponential (slow)\nNone\n\n[Figure: Graph showing descending velocity against altitude lost for different types of pitch change.]\n\nFigure 9.- Effect of rate of pitch reduction on descending velocity against altitude lost. $I_1 = 200$ slug-feet$^2$.", "timestamp": "2026-07-22T05:05:23.733742+00:00"}
{"citation_id": "19930082614", "source_url": "https://ntrs.nasa.gov/api/citations/19930082614/downloads/19930082614.pdf", "page_number": 23, "total_pages": 36, "image_filename": "19930082614_p23.jpg", "text": "NACA TN 1939\n21\n\nTABLE II.- CALCULATION OF THE VARIATIONS OF AIRSPEED WITH\nTIME FOR AN AIRPLANE IN A 60° DIVE FROM AN ALTITUDE OF\n25,000 FEET. WING LOADING, 50 POUNDS PER SQUARE FOOT\n\n| t (sec) | $\\Delta$t (sec) | $V_e$ (ft/sec) | $\\overline{V}_e$ (ft/sec) | $\\Delta$h (ft) | h (ft) | $C_{D_n}$ | a/g | a (ft/sec²) | $\\overline{a}$ (ft/sec²) | $\\Delta$V (ft/sec) | V (ft/sec) |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| 0 | 0 | -- | -- | -- | 25,000 | 0.114 | 0.27 | 8.7 | -- | -- | 700 |\n| 1 | 1 | 709 | 704 | -610 | 24,390 | .114 | .24 | 7.7 | 8.0 | 8.0 | 708 |\n| 2 | 1 | 715 | 712 | -620 | 23,770 | .114 | .20 | 6.4 | 7.0 | 7.0 | 715 |\n| 4 | 2 | 725 | 720 | -1240 | 22,530 | .114 | .17 | 5.5 | 6.0 | 12.0 | 727 |\n| 6 | 2 | 737 | 732 | -1270 | 21,260 | .114 | .13 | 4.2 | 4.9 | 9.8 | 737 |\n| 8 | 2 | 744 | 740 | -1280 | 19,980 | .114 | .07 | 2.3 | 3.2 | 6.4 | 743 |\n| 10 | 2 | 746 | 744 | -1290 | 18,690 | .114 | .03 | 1.0 | 1.6 | 3.2 | 746 |\n| 14 | 4 | 745 | 745 | -2580 | 16,110 | .114 | -.04 | -1.3 | -0.1 | -0.4 | 746 |\n| 18 | 4 | 736 | 741 | -2970 | 13,540 | .114 | -.11 | -3.5 | -2.4 | -9.6 | 736 |\n\n[Figure: NACA logo]\n\nExample calculation:\nAn airplane in a 60° dive at a speed of 700 feet per second instantaneously extends air brakes at time t = 0.\nDrag coefficient of airplane without air brakes, $C_D = 0.014$.\nDrag increment due to air brakes, $\\Delta C_D = 0.100$.\n\nAt t = 0\n$$h = 25,000 \\text{ ft}$$\n$$a/g = 0.27 \\text{ (See guide lines, fig. 1)}$$\n$$a_0 = (0.27) (32.2) = 8.7 \\text{ ft/sec}^2$$\n\nt = 1 sec\n$$\\Delta t_1 = 1 \\text{ sec}$$\n$$V_e = 700 + 8.7 (1) = 709 \\text{ ft/sec}$$\n$$\\overline{V}_e = \\frac{1}{2} (700 + 709) = 704 \\text{ ft/sec}$$\n$$\\Delta h = V_e \\Delta t \\sin \\gamma = (704) (1) (-0.866) = -610 \\text{ ft}$$\n$$h = 25,000 - 610 = 24,390 \\text{ ft}$$\n$$a/g = 0.24 \\text{ (fig. 1)}$$\n$$a_1 = (0.24) (32.2) = 7.7 \\text{ ft/sec}^2$$\n$$\\overline{a} = \\frac{1}{2} (8.7 + 7.7) = 8.2 \\text{ ft/sec}^2$$\n$$\\Delta V = \\overline{a} \\Delta t = 8.2 \\text{ ft/sec}$$\n$$V = 700 + 8 = 708 \\text{ ft/sec}$$\n\nt = 2 sec\n$$\\Delta t_2 = 1 \\text{ sec}$$\n$$V_e = V + \\Delta t_2 \\left[ a_1 + \\frac{(a_1 - a_0)}{2} \\frac{\\Delta t_2}{\\Delta t_1} \\right]$$\n$$= 708 + (1) \\left[ 7.7 + \\frac{7.7 - 8.7}{2} \\frac{(1)}{(1)} \\right]$$", "timestamp": "2026-07-22T05:05:29.172139+00:00"}
{"citation_id": "19930082245", "source_url": "https://ntrs.nasa.gov/api/citations/19930082245/downloads/19930082245.pdf", "page_number": 46, "total_pages": 66, "image_filename": "19930082245_p46.jpg", "text": "```markdown\nNACA TN No. 1596\n\n10\n8\n6\n4\n2\n0\n-2\n-4\n-6\n-8\n-10\n-12\n-14\n.1 .2 .3 .4 .5 .6 .7 .8 .9\nMach number, M\nSection angle of attack, $\\alpha$, deg\n$\\delta_a$ (deg)\n-12\n-6\n-4\n-2\n0\n2\n4\n6\n12\n18\n30\n-2\n0\n2\n4\n\n.28\n.24\n.20\n.16\n.12\n.08\n.04\n0\n-.04\n-.08\n-.12\n-.16\n-.20\n.1 .2 .3 .4 .5 .6 .7 .8 .9\nMach number, M\nSection pitching-moment coefficient, $c_m$\n$\\delta_a$ (deg)\n-12\n-6\n-4\n-2\n0\n2\n4\n6\n12\n18\n30\n-2\n0\n2\n4\n12\n18\n[NACA logo]\n\n(d) $c_n = 0.2$.\nFigure 9.- Continued.\n\n45\n```", "timestamp": "2026-07-22T05:05:29.568286+00:00"}
{"citation_id": "19930085536", "source_url": "https://ntrs.nasa.gov/api/citations/19930085536/downloads/19930085536.pdf", "page_number": 15, "total_pages": 20, "image_filename": "19930085536_p15.jpg", "text": "```markdown\nNACA RM No. E5K05\n\n+ Orifice\n\n6.0°\n\n1\"\n\nSection\ncontaining\norifices\n\n1\"\n\n18.5°\n\n3\"\n\n4 3/4\"\n\n90°\n75°\n70°\n60°\n40°\n20°\n0°\n-20°\n-40°\n-60°\n-70°\n-75°\n-90°\n\nNACA\n\nFigure 2. - Sketch of model showing orifice positions and important dimensions.\n\n13\n```", "timestamp": "2026-07-22T05:05:31.877637+00:00"}
{"citation_id": "19930085869", "source_url": "https://ntrs.nasa.gov/api/citations/19930085869/downloads/19930085869.pdf", "page_number": 1, "total_pages": 36, "image_filename": "19930085869_p1.jpg", "text": "NACA RM L9D15\n\nFILE COPY\nNO 7\n\nCONFIDENTIAL\n\nCopy\nRM L9D157\n\nCLASSIFICATION CANCELLED\n\nNACA\n\nRESEARCH MEMORANDUM\n\nHYDRODYNAMIC CHARACTERISTICS OF\nA SWEPT PLANING-TAIL HULL\n\nBy Robert E. McKann, Claude W. Coffee,\nand Donald D. Arabian\n\nLangley Aeronautical Laboratory\nLangley Air Force Base, Va.\n\nCLASSIFICATION CANCELLED\nAUTHORITY H.L. DRYDEN CHANGE #1494\nDATE 6-11-53 T.C. FRASER, JR.\n\nCLASSIFIED DOCUMENT\n\nThis document contains classified information affecting the National Defense of the United States within the meaning of the Espionage Act, USC 50:31 and 32. Its transmission or the revelation of its contents in any manner to an unauthorized person is prohibited by law. Information or material may be disclosed 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\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\nLANGLEY AERONAUTICAL LABORATORY\nLANGLEY FIELD, HAMPTON, VIRGINIA\n\nRETURN TO THE ADDRESSEE\nRECEIPT FOR THIS DOCUMENT SHOULD BE ADDRESSED\nAS FOLLOWS:\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\nWASHINGTON 25, D.C.\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\nWASHINGTON\nSeptember 12, 1949\n\nCONFIDENTIAL\nCLASSIFICATION CANCELLED", "timestamp": "2026-07-22T05:05:32.147026+00:00"}
{"citation_id": "19930085487", "source_url": "https://ntrs.nasa.gov/api/citations/19930085487/downloads/19930085487.pdf", "page_number": 19, "total_pages": 36, "image_filename": "19930085487_p19.jpg", "text": "NACA RM NO. E8J22\n\n[Figure: Sand pattern for second torsional-mode vibration on first-stage blade. A ruler marked \"INCHES\" with scale from 0 to 2 is visible, placed near the blades. The blades show sand accumulation patterns indicating vibration nodes. A NACA logo and document number C-14380 dated 2-27-46 are in the lower right corner of the image.]\n\nFigure 5. - Sand pattern for second torsional-mode vibration on first-stage blade. Frequency of vibration, 6840 cycles per second.\n\n17", "timestamp": "2026-07-22T05:05:32.569640+00:00"}
{"citation_id": "19930085626", "source_url": "https://ntrs.nasa.gov/api/citations/19930085626/downloads/19930085626.pdf", "page_number": 3, "total_pages": 24, "image_filename": "19930085626_p3.jpg", "text": "2\nCONFIDENTIAL\nNACA RM No. L8K23\n\nproduce reversal at transonic speeds. It was desired that the modifications be limited to the aileron only; the remainder of the wing was to be unmodified. Several satisfactory aileron configurations were developed in the \"transonic bump\" tests which are reported in reference 2. Confirmatory free-flight tests of these configurations were subsequently performed at larger scale and are the subject of the present paper.\n\nIn the present investigation, outboard, inboard, and full-span ailerons were tested. The inboard and full-span ailerons had true-contour profiles. The outboard ailerons were tested with several profile modifications including straight-side extended-chord ailerons, parallel-side ailerons of thickness equal to the wing thickness at the hinge line having blunt trailing edges, and straight-side ailerons having a blunt trailing edge of thickness equal to one-half of the wing thickness at the hinge line.\n\nThe present investigation, which was made by means of the technique described in references 3 and 4, permits the evaluation of the wing-aileron rolling effectiveness over the Mach number range from about 0.6 to 1.9 at relatively large scale. The variation of drag coefficient with Mach number was also obtained.\n\nSYMBOLS\n\n| Symbol | Definition |\n| :--- | :--- |\n| $\\frac{pb}{2V}$ | wing-tip helix angle, radians |\n| $p$ | rolling velocity, radians per second |\n| $b$ | diameter of circle swept by wing tips, feet |\n| $V$ | flight-path velocity, feet per second |\n| $C_D$ | total-drag coefficient based on total exposed area of basic wing (1.563 sq ft) |\n| $M$ | Mach number |\n| $R$ | Reynolds number based on average exposed chord of basic wing (0.55 ft) |\n| $c$ | wing chord parallel to model center line |\n| $\\delta_a$ | aileron deflection measured in plane normal to chord plane and to aileron hinge line |\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:05:32.870609+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 36, "total_pages": 99, "image_filename": "19930082511_p36.jpg", "text": "34\nNACA TN No. 1826\n\nDerivation of $w_3$.— In order to cancel the effect of the terms $Mw_1 + Nw_2$ at infinity, the function $w_3$ must approach $-R \\cdot P \\cdot Nw_2(z)$ as $z$ increases without limit. In addition, it must have no singularities in the upper half of the $z$-plane, it must satisfy the condition of continuity at $z = \\pm 1$, it must be a pure real on the fixed boundaries and a pure imaginary on the free boundaries, and it must be zero at $z = 0$. It is readily formulated as\n\n$$\nw_3(z) = -Nz \\sqrt{\\frac{1 - z^2}{a^2 - z^2}}\n$$\n\nThat this function satisfies the first condition is readily shown by writing it in a slightly different form and expanding the radical:\n\n$$\n\\lim_{z \\to \\infty} -Nz \\sqrt{\\frac{1 - z^2}{a^2 - z^2}} = \\lim_{z \\to \\infty} -Nz \\sqrt{\\frac{1 - \\frac{1}{z^2}}{1 - \\frac{a^2}{z^2}}}\n$$\n\n$$\n= \\lim_{z \\to \\infty} -Nz \\left( 1 + \\frac{a^2}{2z^2} - \\frac{1}{2z^2} + \\dots \\right)\n$$\n\n$$\n= \\lim_{z \\to \\infty} -Nz - \\frac{N(a^2 - 1)}{2z} - \\dots\n$$\n\nComparison of this expression with that for $\\lim_{z \\to \\infty} R \\cdot P \\cdot Nw_2(z)$ shows that the difference between the two expressions approaches zero as $z$ approaches infinity. That the function satisfies the remaining conditions is readily verified by inspection.\n\nThe complex velocity function for the closed-open-closed tunnel with unequal pressures is, finally,\n\n$$\nQ(z) = \\frac{2a}{\\pi}(K' - E') \\int_0^z \\frac{dz}{\\sqrt{(1 - z^2)(a^2 - z^2)}}\n$$\n\n$$\n- \\frac{2K'}{a\\pi} \\int_0^z \\sqrt{\\frac{a^2 - z^2}{1 - z^2}} \\, dz + \\frac{2K'}{a\\pi} z \\sqrt{\\frac{1 - z^2}{a^2 - z^2}} \\quad (22)\n$$", "timestamp": "2026-07-22T05:05:35.246964+00:00"}
{"citation_id": "19930082496", "source_url": "https://ntrs.nasa.gov/api/citations/19930082496/downloads/19930082496.pdf", "page_number": 31, "total_pages": 50, "image_filename": "19930082496_p31.jpg", "text": "30\n\nPage intentionally left blank\n\nPage intentionally left blank", "timestamp": "2026-07-22T05:05:36.114738+00:00"}
{"citation_id": "19930082485", "source_url": "https://ntrs.nasa.gov/api/citations/19930082485/downloads/19930082485.pdf", "page_number": 45, "total_pages": 62, "image_filename": "19930082485_p45.jpg", "text": "44\nNACA TN No. 1810\n\n[Figure: Diagram showing the flow of a gas in a channel between curved surfaces. The diagram includes labels for \"Velocity-potential lines\", \"Streamlines\", and variables $r$, $r_1$, $r_2$, $n$, and $dn$.]\n\nFigure 13.- The flow of a gas in the channel between curved surfaces.\n\n1026", "timestamp": "2026-07-22T05:05:36.829060+00:00"}
{"citation_id": "19930082914", "source_url": "https://ntrs.nasa.gov/api/citations/19930082914/downloads/19930082914.pdf", "page_number": 23, "total_pages": 66, "image_filename": "19930082914_p23.jpg", "text": "22\nNACA TN No. 1857\n\nor, by use of the series expansion, to\n$$y \\approx \\frac{1}{2} bz^2$$\n\nThe deviation of a ray on passing through the 3 inches of the mixing region, from $z = 0$ to $z = 3$, is\n$$y = \\frac{1}{2} \\times 0.00042 \\times 9$$\n$$= 0.0019 \\text{ inch}$$\n\nThe index of refraction at $y = 0$ is\n$$n = 1.00028$$\n\nand at $y = 0.0019$ is\n$$n = 1.0002808$$\n\nThe important quantity, though, is $n - 1$. The light ray emerges from the mixing region at a place where $n - 1$ differs by less than 1 percent from its value at the place where the ray entered the mixing region. For the jet under discussion, therefore, the effect of refraction is negligible.\n\nRESULTS\n\nInterferograms\n\nFor obtaining interferograms of the mixing region of the free jet of Mach number 1.6, the interferometer was so adjusted that straight, horizontal interference fringes were produced when there was no air flow, as is shown in figure 10.\n\nFigure 13 shows the portion of the jet of which interferograms were taken for the present investigation. This portion was the bottom part of the horizontal jet for the first 10 inches from the nozzle. The", "timestamp": "2026-07-22T05:05:40.073636+00:00"}
{"citation_id": "19930082498", "source_url": "https://ntrs.nasa.gov/api/citations/19930082498/downloads/19930082498.pdf", "page_number": 29, "total_pages": 49, "image_filename": "19930082498_p29.jpg", "text": "Page intentionally left blank\n\nPage intentionally left blank", "timestamp": "2026-07-22T05:05:44.128170+00:00"}
{"citation_id": "19930082617", "source_url": "https://ntrs.nasa.gov/api/citations/19930082617/downloads/19930082617.pdf", "page_number": 25, "total_pages": 58, "image_filename": "19930082617_p25.jpg", "text": "24\nNACA TN 1962\n\nStringers\nO 1 to 9\nX 10 to 16\n\nMoment\n(in. - lb)\n1 36.0 X 10^3 4 144.0 X 10^3\n2 72.0 X 10^3 5 216.0 X 10^3\n3 108.0 X 10^3 6 288.0 X 10^3\n\n[Figure: Cross-section diagram showing Band V, 2.57\" dimension, and A-A view at 45°]\n\nDistance from horizontal diameter, in.\nStrain\n\nFigure 13.- Strain diagram of cylinder 75. Band V.", "timestamp": "2026-07-22T05:05:46.305289+00:00"}
{"citation_id": "19930082450", "source_url": "https://ntrs.nasa.gov/api/citations/19930082450/downloads/19930082450.pdf", "page_number": 34, "total_pages": 37, "image_filename": "19930082450_p34.jpg", "text": "NACA TN No. 1778\n33\n\n$$\n\\frac{H}{t_W} = 26\n$$\n$$\n\\left(\\frac{t_W}{t_W} = .25\\right)\n$$\n\n$$\n\\frac{S}{t_S} \\text{ or } \\frac{b_S}{t_S}\n$$\n\n$$\n\\sigma_{cr}, \\text{ ksi}\n$$\n12.7\n8.3\n\n$$\n\\frac{P_i}{L\\sqrt{E}}, \\text{ ksi}\n$$\n\n$$\nt_S\n$$\n$$\nb_W\n$$\n$$\nt_W\n$$\n$$\nb_S \\text{ or } S\n$$\n\n$$\n\\bar{\\sigma}_f, \\text{ ksi}\n$$\n\n36\n(35)\n\n$$\n\\sigma_{cr}, \\text{ ksi}\n$$\n19.3\n17.6\n12.7\n8.3\n\n$$\n\\frac{S}{t_S} \\text{ or } \\frac{b_S}{t_S}\n$$\n\n$$\n\\frac{P_i}{L\\sqrt{E}}, \\text{ ksi}\n$$\n\n46\n(45)\n\n$$\n\\sigma_{cr}, \\text{ ksi}\n$$\n17.6\n12.7\n8.3\n\n$$\n\\frac{S}{t_S} \\text{ or } \\frac{b_S}{t_S}\n$$\n\n$$\n\\frac{P_i}{L\\sqrt{E}}, \\text{ ksi}\n$$\n24S-T\n$$\n\\sigma_{cy} = 44 \\text{ ksi}\n$$\n\n$$\n\\frac{P_i}{t_S}, \\text{ ksi}\n$$\n\nNACA\n\nFigure 8.-Concluded. $$ \\frac{t_W}{t_S} = 0.79 $$.", "timestamp": "2026-07-22T05:05:46.520895+00:00"}
{"citation_id": "19930085542", "source_url": "https://ntrs.nasa.gov/api/citations/19930085542/downloads/19930085542.pdf", "page_number": 9, "total_pages": 46, "image_filename": "19930085542_p9.jpg", "text": "NACA RM No. L8L29\n\nEffect of Profile of Triangular Wings\n\nChanges in wing profile appear to have rather large effects on the lift, longitudinal-force, and pitching-moment characteristics at moderate and high lift coefficients as is indicated in figure 8. The highest maximum lift coefficient was obtained with the flat-plate wing (model 1) and the lowest with the biconvex (12-percent-thick) wing (model 3). The NACA 0012 wing (model 2) gave gradually increasing longitudinal stability throughout the lift-coefficient range. Reductions in longitudinal stability were obtained at $C_L = 0.4$ for both the flat-plate wing and the biconvex wing. (See fig. 8.) For all three models the directional stability increased with lift coefficient up to $C_L = 0.9$, after which the directional stability decreased to about zero at maximum lift coefficient. (See fig. 9.)\n\nCertain characteristics appear to have a consistent relation to the shape of the airfoil leading edge. The effective dihedral parameter $C_{l_\\psi}$, for example, varies almost linearly up to $C_L = 0.4$ for the blunt-nose NACA 0012 airfoil. For the sharper nose flat-plate and biconvex airfoils $C_{l_\\psi}$ is linear only to a lift coefficient of 0.25. The decrease in the initial linear range of the effective dihedral parameter as the leading edge was effectively sharpened was noted in tests of untapered swept wing in reference 7. Similar trends are noted, but to a lesser degree, for the lift-curve slope (fig. 8) and for the variations on the rolling derivatives $C_{Y_p}$ and $C_{n_p}$ with lift coefficient (fig. 10). Negative values of $C_{n_p}$ (as predicted by the theories of references 2 and 5) were obtained only for the NACA 0012 profile model and then only to a lift coefficient of 0.5. No consistent effects of airfoil section are noted for the damping in roll over the range of lift coefficients for which the data are presented.\n\nEffect of Aspect Ratio of Triangular Plan Forms\n\nIt should be remembered that in the following discussion of the effect of aspect ratio of triangular plan forms there are also effects of sweep present, since the sweep angle is automatically increased as the aspect ratio is decreased.\n\nAs the aspect ratio is reduced, $C_{L_\\alpha}$ is decreased (at low lift coefficients) and $C_{L_{max}}$ occurs at higher angles of attack. This trend was noted in reference 8 in tests of similar triangular-wing models. At $C_L = 0.3$ a sharp increase in $C_{L_\\alpha}$ occurs for model 4; increased longitudinal stability is noted at the same lift coefficient. An opposite trend is noted for model 7 at $C_L = 0.6$ where a decrease occurs in the", "timestamp": "2026-07-22T05:05:54.993933+00:00"}
{"citation_id": "19930085519", "source_url": "https://ntrs.nasa.gov/api/citations/19930085519/downloads/19930085519.pdf", "page_number": 17, "total_pages": 46, "image_filename": "19930085519_p17.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:05:55.173558+00:00"}
{"citation_id": "19930085536", "source_url": "https://ntrs.nasa.gov/api/citations/19930085536/downloads/19930085536.pdf", "page_number": 16, "total_pages": 20, "image_filename": "19930085536_p16.jpg", "text": "14\nNACA RM No. E8K05\n\n[Figure: Diagram showing concentric circles and an ellipse with numbered points. A radial line is labeled r/x with values .10, .20, .30. A legend indicates 'O Source position'. The NACA logo is present.]\n\nFigure 3. - Cross section of test body showing source configuration for theoretical calculations.", "timestamp": "2026-07-22T05:06:05.437787+00:00"}
{"citation_id": "19930085869", "source_url": "https://ntrs.nasa.gov/api/citations/19930085869/downloads/19930085869.pdf", "page_number": 2, "total_pages": 36, "image_filename": "19930085869_p2.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:06:06.871452+00:00"}
{"citation_id": "19930085487", "source_url": "https://ntrs.nasa.gov/api/citations/19930085487/downloads/19930085487.pdf", "page_number": 20, "total_pages": 36, "image_filename": "19930085487_p20.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:06:07.260105+00:00"}
{"citation_id": "19930085544", "source_url": "https://ntrs.nasa.gov/api/citations/19930085544/downloads/19930085544.pdf", "page_number": 7, "total_pages": 33, "image_filename": "19930085544_p7.jpg", "text": "6\nNACA RM No. L8K26\n\nor\n\n$$\n\\phi_{\\omega t} = \\tan^{-1} \\frac{\\cos \\alpha_T}{\\frac{\\pi x}{J} + \\sin \\alpha_T \\sin \\omega t} \\tag{1}\n$$\n\nThe resultant velocity is given by\n\n$$\nW_{\\omega t} = \\sqrt{V^2 \\cos^2 \\alpha_T + (\\pi n D x + V \\sin \\alpha_T \\sin \\omega t)^2} \\tag{2}\n$$\n\nUsing the relationship in equation (1), the local advance ratio is given by\n\n$$\nJ_{\\omega t} = \\frac{\\pi x \\cos \\alpha_T}{\\frac{\\pi x}{J} + \\sin \\alpha_T \\sin \\omega t} \\tag{3}\n$$\n\nFrom (3) it is seen that the local advance ratio varies depending on the position of the blade.\n\n### Method of Analysis\n\nIn calculating the forces on an inclined propeller it must be realized that not only do the blade sections operate in a variable flow field but that the flow is a compressible one with the possibility of high section Mach numbers along the propeller blades. The method of reference 2 for dealing with the oscillating effects applies to incompressible flow where the slope of the lift curve is approximately $2\\pi$, while in the compressible case, the slope may be considerably higher. Since the wave length in oscillating flow is usually several blade chord lengths (10 or more), it appears logical as a first approximation to consider the oscillating effects to be negligible as compared with the change of slope of the lift curve with change in Mach number. Also, the Goldstein correction factors for a finite number of blades have been found to apply reasonably well when applied to the calculation of forces on nonoptimum propellers (reference 4). Therefore, it appears reasonable to extend their use to the present case.\n\n**Steady state.**- In steady-state calculations of the forces and moments on the blade of a pitched propeller, a change in time (blade position) is treated simply as a change in the operating $V/nD$ of the propeller in accordance with equation (3). The complete propeller is assumed to operate successively at different blade positions under the instantaneous conditions at each particular position. The thrust per", "timestamp": "2026-07-22T05:06:08.380745+00:00"}
{"citation_id": "19930082496", "source_url": "https://ntrs.nasa.gov/api/citations/19930082496/downloads/19930082496.pdf", "page_number": 32, "total_pages": 50, "image_filename": "19930082496_p32.jpg", "text": "NACA TN No. 1836\n31\n\n[Figure: Four images of a ceramic blade]\n\nNACA\nC-19258\n7-31-47\n\nNACA\nC-19257\n7-31-47\n\nRoll\nNeck\nAirfoil\n\nNACA\nC-19259\n7-31-47\n\nNACA\nC-19260\n7-31-47\n\nFigure 6. - Typical ceramal blade.", "timestamp": "2026-07-22T05:06:12.678532+00:00"}
{"citation_id": "19930085471", "source_url": "https://ntrs.nasa.gov/api/citations/19930085471/downloads/19930085471.pdf", "page_number": 21, "total_pages": 28, "image_filename": "19930085471_p21.jpg", "text": "UNCLASSIFIED\nRESTRICTED\nCONFIDENTIAL\n\nNACA RM No. L8G11\n\n60-cycle timer\npressure cell 1 2 3 4\nreference line\nmodel position 1 in.\ntime\ntorsion\nbending\nNACA\n\nFigure 5.- Sample oscillograph record of the flutter of model B-5.\n\nRESTRICTED\nCONFIDENTIAL\n\n19", "timestamp": "2026-07-22T05:06:13.929468+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 37, "total_pages": 99, "image_filename": "19930082511_p37.jpg", "text": "NACA TN No. 1826\n35\n\nor\n\n$$Q(z) = \\frac{2}{\\pi}(K' - E')F(z) - \\frac{2K'}{\\pi}E(z) + \\frac{2K'}{a^2\\pi} z \\sqrt{\\frac{1 - z^2}{1 - \\frac{z^2}{a^2}}} \\quad (23)$$\n\nwhere the modulus of the elliptic integrals is $1/a$.\n\nInduced velocity on the axis.- For the special case in which $z = iy$, the preceding equations for $Q$ reduce to a somewhat simpler form. The procedure will be only outlined here, inasmuch as the manipulative steps are similar to those already described.\n\nReplacing $z$ with $iy$ in the expression for $w_1$ and then substituting $y^2 = \\frac{a^2}{b^2} - a^2$ reduces the first term to\n\n$$\\frac{2i}{\\pi}(K' - E') \\left[ K' - F'\\left( \\frac{1}{\\sqrt{1 + \\frac{y^2}{a^2}}} \\right) \\right]$$\n\nThe same substitutions, together with the previously described technique from reference 12 reduces the second term to\n\n$$- \\frac{2iK'}{\\pi} \\left\\{ \\frac{y}{a^2} \\sqrt{\\frac{1 + y^2}{1 + \\frac{y^2}{a^2}}} + \\left[ K' - F'\\left( \\frac{1}{\\sqrt{1 + \\frac{y^2}{a^2}}} \\right) \\right] - \\left[ E' - E'\\left( \\frac{1}{\\sqrt{1 + \\frac{y^2}{a^2}}} \\right) \\right] \\right\\}$$\n\nThe third term is found directly as\n\n$$\\frac{2iK'}{a^2\\pi} y \\sqrt{\\frac{1 + y^2}{1 + \\frac{y^2}{a^2}}}$$\n\nThe total simplifies to the form\n\n$$Q(iy) = \\frac{2i}{\\pi} \\left[ -K'E'\\left( \\frac{1}{\\sqrt{1 + \\frac{y^2}{a^2}}} \\right) + E'F'\\left( \\frac{1}{\\sqrt{1 + \\frac{y^2}{a^2}}} \\right) \\right] \\quad (24)$$", "timestamp": "2026-07-22T05:06:18.650234+00:00"}
{"citation_id": "19930082542", "source_url": "https://ntrs.nasa.gov/api/citations/19930082542/downloads/19930082542.pdf", "page_number": 24, "total_pages": 53, "image_filename": "19930082542_p24.jpg", "text": "NACA TN No. 1867\n23\n\nTABLE II.—RUPTURE TEST CHARACTERISTICS AT 1200° F OF LOW-CARBON H-15 BAR STOCK\n\n| Heat treatment | | | | | Hot-cold-rolling (b) | | Rupture properties at 1200° F | | | |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| **Solution treatment** | | | **Aging treatment (a)** | | | | | | | |\n| **Temper-ature (°F)** | **Time (hr)** | **Method of cooling (c)** | **Temper-ature (°F)** | **Time (hr)** | **Temper-ature (°F)** | **Percent reduction** | **Stress (psi)** | **Rupture time (hr)** | **Elongation in 1 in. (percent)** | **Reduction of area (percent)** | **Rupture strength (psi)** |\n| | | | | | | | | | | | **100 hr** | **1000 hr** |\n| **Aging and rolling hot-rolled bar stock** | | | | | | | | | | | | |\n| (d) | (a) | (a) | (d) | (d) | (d) | (a) | 25,000 | 8.5 | 89.5 | 8.5 | 49,500 | 37,500 |\n| | | | | | | | 30,000 | 75 | 17 | 23.3 | | |\n| | | | | | | | 35,000 | 252 | 19 | 23.3 | | |\n| | | | | | | | 40,000 | 610 | 34 | 48.3 | | |\n| --- | --- | --- | 1350 | 24 | --- | --- | 50,000 | 123 | 12 | 49.2 | 51,000 | 35,500 |\n| | | | | | | | 55,000 | 241 | 38 | 55.3 | | |\n| | | | | | | | 41,000 | 411 | 33 | 55.3 | | |\n| --- | --- | --- | 1500 | 24 | --- | --- | 50,000 | 54 | 16 | 46.5 | 47,000 | 36,000 |\n| | | | | | | | 45,000 | 166 | 12 | 50.1 | | |\n| | | | | | | | 40,000 | 430 | 40 | 58.6 | | |\n| --- | --- | --- | 1750 | 24 | --- | --- | 50,000 | 40 | 42 | 21.9 | 45,000 | 37,000 |\n| | | | | | | | 45,000 | 101 | 38 | 44.7 | | |\n| | | | | | | | 40,000 | 430 | 36 | 51.0 | | |\n| --- | --- | --- | --- | --- | 75 | 10 (9.8) | 60,000 | 73 | 4 | 4.0 | 58,500 | 49,000 |\n| | | | | | | | 55,000 | 230 | 13 | 32.4 | | |\n| | | | | | | | 50,000 | 809 | 10 | 16.7 | | |\n| --- | --- | --- | --- | --- | 1200 | 5 (4.5) | 60,000 | 48 | 8 | 22.2 | 56,000 | 46,000 |\n| | | | | | | | 55,000 | 216 | 8 | 32.2 | | |\n| | | | | | | | 50,000 | 387 | 12 | 36.0 | | |\n| --- | --- | --- | --- | --- | 1200 | 10 | 65,000 | 56 | 5 | 12.3 | 61,000 | 47,000 |\n| | | | | | | | 60,000 | 115 | 5 | 28.8 | | |\n| | | | | | | | 55,000 | 308 | 16 | 38.8 | | |\n| | | | | | | | 50,000 | 528 | 15 | 40.8 | | |\n| --- | --- | --- | --- | --- | 1200 | 15 (14.1) | 60,000 | 157 | 6 | 20.6 | 63,000 | 49,000 |\n| | | | | | | | 55,000 | 325 | 6 | 24.1 | | |\n| | | | | | | | 50,000 | 845 | 6.5 | 47.0 | | |\n| --- | --- | --- | 81400 | 24 | 1200 | 15 (15.6) | 55,000 | 72 | 16 | 51.9 | 53,000 | 47,000 |\n| | | | | | | | 52,500 | 108 | 27 | 40.8 | | |\n| | | | | | | | 50,000 | 364 | 19 | 38.8 | | |\n| --- | --- | --- | --- | --- | 1200 | 20 | 65,000 | 84 | 10 | 21.8 | 62,000 | 46,000 |\n| | | | | | | | 60,000 | 112 | 10 | 23.7 | | |\n| | | | | | | | 55,000 | 210 | 18 | 44.7 | | |\n| | | | | | | | 50,000 | 536 | 15 | 33.5 | | |\n| (h) | (h) | (h) | (h) | (h) | (h) | (h) | 60,000 | 25 | 30 | 50.4 | 49,000 | 35,000 |\n| | | | | | | | 50,000 | 85 | 26 | 36.6 | | |\n| | | | | | | | 40,000 | 429 | 16 | 24.8 | | |\n| | | | | | | | 35,000 | 785 | 12 | 19.0 | | |\n| (i) | (i) | (i) | (i) | (i) | (i) | (i) | 50,000 | 23 | 36 | 41.1 | 43,500 | 35,000 |\n| | | | | | | | 40,000 | 240 | 34 | 44.6 | | |\n| | | | | | | | 44,000 | 163 | 35 | 41.5 | | |\n| | | | | | | | 35,000 | 1086 | 28 | 35.6 | | |\n| **Solution-treated at 1800° F** | | | | | | | | | | | | |\n| 1800 | 2 | W.Q. | --- | --- | --- | --- | 45,000 | 51 | 40 | 58.6 | 42,000 | 35,000 |\n| | | | | | | | 40,000 | 120 | 32 | 56.1 | | |\n| | | | | | | | 35,000 | 175 | 36 | 55.3 | | |\n| | | | | | | | 35,000 | 1313 | 38 | 52.0 | | |\n| 1800 | 2 | W.Q. | --- | --- | 1200 | 15 | 55,000 | 103 | 13 | 41.5 | 55,000 | 40,000 |\n| | | | | | | | 50,000 | 241 | 9 | 32.8 | | |\n| | | | | | | | 45,000 | 415 | 14 | 46.0 | | |\n| **Solution-treated at 1950° F** | | | | | | | | | | | | |\n| 1950 | 2 | W.Q. | --- | --- | --- | --- | 45,000 | 92 | 27 | 35.0 | 45,000 | 38,000 |\n| | | | | | | | 40,000 | 318 | 25.5 | 40.8 | | |\n| | | | | | | | 37,500 | 1107 | 30 | 40.8 | | |\n| 1950 | 2 | W.Q. | --- | --- | 1200 | 15 | 64,000 | 59 | 4 | 8.1 | 61,000 | 52,000 |\n| | | | | | | | 60,000 | 168 | 4 | 17.2 | | |\n| | | | | | | | 55,000 | 517 | 5 | 16.3 | | |\n\n$^a$All aging treatments preceded hot-cold-rolling except where noted.\n$^b$All hot-cold-rolled material was given a final stress relief at 1200° F for 1 hr.\n$^c$W.Q., water-quenched; A.C., air-cooled.\n$^d$As-hot-rolled.\n$^e$Fractured in gage mark.\n$^f$Estimated.\n$^g$Aged after rolling.\n$^h$100-percent reduction at 1200° F; 65-percent reduction from 1800° to 1400° F.\n$^i$1800° F 2 hr, air-cooled to 1400° F, 50-percent reduction at 1400° F, repeated five more times; then 1800° F 2 hr, air-cooled.\n\n[Figure: NACA logo]", "timestamp": "2026-07-22T05:06:22.605397+00:00"}
{"citation_id": "19930082498", "source_url": "https://ntrs.nasa.gov/api/citations/19930082498/downloads/19930082498.pdf", "page_number": 30, "total_pages": 49, "image_filename": "19930082498_p30.jpg", "text": "NACA TN No. 1838\n\n[Figure: Photograph of various commercial aircraft mufflers and muffler-heaters laid out on a concrete surface. Several components are labeled with numbered callouts: 7, 8, 9, and 10. A ruler is visible for scale. In the bottom right corner of the image, there is a NACA logo with the identifier “I-53827”.]\n\n(a) Commercial aircraft mufflers and muffler-heaters.\n\nFigure 2.— Photographs of typical mufflers shown in table II. Numbers correspond to configurations listed in table II.", "timestamp": "2026-07-22T05:06:23.755349+00:00"}
{"citation_id": "19930082617", "source_url": "https://ntrs.nasa.gov/api/citations/19930082617/downloads/19930082617.pdf", "page_number": 26, "total_pages": 58, "image_filename": "19930082617_p26.jpg", "text": "NACA TN 1962\n25\n\nStringers\nO 1 to 9\nX 10 to 16\n\nMoment\n(in. - lb)\n1 54.0 X $10^3$\n2 108.0 X $10^3$\n3 162.0 X $10^3$\n4 216.0 X $10^3$\n5 270 X $10^3$\n\n3.66\"\nBand A\nA\nA\nA-A\n\n| Distance from horizontal diameter, in. | Strain |\n| :--- | :--- |\n| 10 | |\n| 9 | |\n| 8 | |\n| 7 | |\n| 6 | |\n| 5 | |\n| 4 | |\n| 3 | |\n| 2 | |\n| 1 | |\n| 0 | 16 12 8 4 0 -4 -8 -12 -16 -20 X $10^{-4}$ |\n| 1 | |\n| 2 | |\n| 3 | |\n| 4 | |\n| 5 | |\n| 6 | |\n| 7 | |\n| 8 | 1 2 3 4 5 |\n| 9 | |\n| 10 | |\n\nNACA\n\nFigure 14.- Strain diagram of cylinder 76. Band A.", "timestamp": "2026-07-22T05:06:29.548388+00:00"}