Buckets:
| {"citation_id": "19930085487", "source_url": "https://ntrs.nasa.gov/api/citations/19930085487/downloads/19930085487.pdf", "page_number": 10, "total_pages": 36, "image_filename": "19930085487_p10.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T04:59:45.339294+00:00"} | |
| {"citation_id": "19930085536", "source_url": "https://ntrs.nasa.gov/api/citations/19930085536/downloads/19930085536.pdf", "page_number": 7, "total_pages": 20, "image_filename": "19930085536_p7.jpg", "text": "$$U _ { \\theta } = \\mathrm { K m } \\beta ^ { 2 } \\cos \\left( \\theta - \\delta \\right) \\left( \\frac { 1 } { 1 - \\left( \\mathrm { m } \\beta \\right) ^ { 2 } } \\cosh ^ { - 1 } \\left\\{ \\frac { 1 - \\mathrm { m } \\displaystyle \\frac { \\mathrm { r } } { \\mathrm { x } } \\beta ^ { 2 } \\sin \\left( \\theta - \\delta \\right) } { \\sqrt { \\left[ 1 - \\mathrm { m } \\displaystyle \\frac { \\mathrm { r } } { \\mathrm { x } } \\beta ^ { 2 } \\sin \\left( \\theta - \\delta \\right) \\right] ^ { 2 } - \\left[ 1 - \\left( \\displaystyle \\frac { \\mathrm { r } \\beta } { \\mathrm { x } } \\right) ^ { 2 } \\right] \\left[ 1 - \\left( \\mathrm { m } \\beta \\right) ^ { 2 } \\right] } } \\right\\} \\right.$$\n\n$$- \\frac { 1 - \\mathrm { m } \\displaystyle \\frac { \\mathrm { r } } { \\mathrm { x } } \\beta ^ { 2 } \\sin \\left( \\theta - \\delta \\right) } { \\left[ 1 - \\mathrm { m } \\displaystyle \\frac { \\mathrm { r } } { \\mathrm { x } } \\beta ^ { 2 } \\sin \\left( \\theta - \\delta \\right) \\right] ^ { 2 } - \\left[ 1 - \\left( \\displaystyle \\frac { \\mathrm { r } \\beta } { \\mathrm { x } } \\right) ^ { 2 } \\right] \\left[ 1 - \\left( \\mathrm { m } \\beta \\right) ^ { 2 } \\right] } \\sqrt { \\frac { 1 - \\left( \\displaystyle \\frac { \\mathrm { r } \\beta } { \\mathrm { x } } \\right) ^ { 2 } } { 1 - \\left( \\mathrm { m } \\beta \\right) ^ { 2 } } } \\Bigg )$$\n\nThe boundary condition of the flow is\n\n$$\\frac { \\mathrm { U } _ { \\mathrm { r } } } { \\mathrm { U } } - \\left( \\frac { 1 } { \\mathrm { r } } \\frac { \\partial \\mathrm { r } } { \\partial \\theta } \\right) \\frac { \\mathrm { U } _ { \\theta } } { \\mathrm { U } } - \\frac { \\mathrm { r } } { \\mathrm { x } } \\frac { \\mathrm { U } _ { \\mathrm { x } } } { \\mathrm { U } } = \\frac { \\mathrm { r } } { \\mathrm { x } }$$\n\nFrom integration of the Bernoulli equation for isentropic flow, the pressure coefficient becomes\n\n$$C _ { \\mathrm { P } } = { \\frac { 2 } { \\gamma \\mathrm { M } ^ { 2 } } } \\left( \\left\\{ 1 - { \\frac { \\gamma - 1 } { \\gamma } } \\mathrm { M } ^ { 2 } \\left[ { \\frac { 2 \\mathrm { U } _ { \\mathrm { x } } } { \\mathrm { U } } } + \\left( { \\frac { \\mathrm { U } _ { \\mathrm { x } } } { \\mathrm { U } } } \\right) ^ { 2 } + \\left( { \\frac { \\mathrm { U } _ { \\mathrm { r } } } { \\mathrm { U } } } \\right) ^ { 2 } + \\left( { \\frac { \\mathrm { U } _ { \\theta } } { \\mathrm { U } } } \\right) ^ { 2 } \\right] \\right\\} ^ { \\frac { \\gamma } { \\gamma - 1 } } - 1 \\right)$$", "timestamp": "2026-07-22T04:59:51.270333+00:00"} | |
| {"citation_id": "19930082485", "source_url": "https://ntrs.nasa.gov/api/citations/19930082485/downloads/19930082485.pdf", "page_number": 38, "total_pages": 62, "image_filename": "19930082485_p38.jpg", "text": "NACA TN NO. 1810\n37\n\n1026\n\nStagnation pressure\n$P_t$, lb/sq ft\n\nStatic pressure\n$P_s$, lb/sq ft\n\nRadius, r, in.\n\nO Pitot tube\n□ Wall taps\n\nInner shroud\nOuter shroud\n\nNACA\n\nFigure 7. - Measured stagnation and static pressure at 0.1 chord downstream of cascade.", "timestamp": "2026-07-22T04:59:51.839092+00:00"} | |
| {"citation_id": "19930082914", "source_url": "https://ntrs.nasa.gov/api/citations/19930082914/downloads/19930082914.pdf", "page_number": 16, "total_pages": 66, "image_filename": "19930082914_p16.jpg", "text": "NACA TN No. 1857\n15\n\nor\n\n$$X = \\frac{L}{\\lambda^T} (\\lambda - \\lambda^T)$$\n\nBut, since\n\n$$n = \\frac{V_O}{V} = \\frac{\\lambda_O}{\\lambda}$$\n\nand\n\n$$n^T = \\frac{\\lambda_O}{\\lambda^T}$$\n\ntherefore\n\n$$X = \\frac{L}{n} (n^T - n)$$\n$$= L \\frac{\\lambda}{\\lambda_O} (n^T - n)$$\n\nBy similar triangles, it follows that\n\n$$\\frac{Y}{b} = \\frac{X}{\\lambda} = \\frac{L}{\\lambda_O} (n^T - n)$$", "timestamp": "2026-07-22T04:59:55.445795+00:00"} | |
| {"citation_id": "19930093773", "source_url": "https://ntrs.nasa.gov/api/citations/19930093773/downloads/19930093773.pdf", "page_number": 43, "total_pages": 47, "image_filename": "19930093773_p43.jpg", "text": "42\nNACA RM E9G09\n\n[Figure: A graph plotting Engine total-temperature ratio, $T_7/T_1$ against Engine total-pressure ratio, $P_7/P_1$. The graph contains multiple data series represented by different symbols (circle, square, diamond, triangle, inverted triangle) corresponding to different Flight Mach numbers. A legend is present in the upper center of the plot area. The NACA logo is in the bottom right corner of the plot area.]\n\n| Flight Mach number |\n| :--- |\n| $\\circ$ 0.21 |\n| $\\square$ .53 |\n| $\\diamond$ .72 |\n| $\\triangle$ .85 |\n| $\\nabla$ .97 |\n\nEngine total-temperature ratio, $T_7/T_1$\nEngine total-pressure ratio, $P_7/P_1$\n\n(b) Flight Mach number, 0.21 to 0.97; altitude, 25,000 feet.\nFigure 7. - Concluded. Variation of engine total-temperature ratio\nwith engine total-pressure ratio.\n\n6511", "timestamp": "2026-07-22T05:00:00.173410+00:00"} | |
| {"citation_id": "19930082498", "source_url": "https://ntrs.nasa.gov/api/citations/19930082498/downloads/19930082498.pdf", "page_number": 22, "total_pages": 49, "image_filename": "19930082498_p22.jpg", "text": "```markdown\nTABLE II.- RESULTS OF MUFFLER INVESTIGATION MADE WITH PROPELLER REMOVED - Continued\n\n| Muffler configuration | Engine speed (rpm) | Fundamental firing frequency, F (cps) | Sound-pressure level (db) | | | | | | | | | | | | | | Other sounds | | | Back pressure |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| (a) | | | Over-all | 0.5F | 1.0F | 1.5F | 2.0F | 2.5F | 3.0F | 3.5F | 4.0F | 4.5F | 5.0F | 5.5F | 6.0F | 7.0F | cps | db | db | |\n| 57 (See fig. 2(b)) | 2000<br>2790 | 100.0<br>139.5 | 91.0<br>102.0 | 60<br>(b) | 68<br>94 | 67<br>(b) | 84<br>85 | 70<br>75 | 84<br>(b) | 74<br>(b) | 75<br>(b) | 70<br>(b) | 65<br>(b) | 72<br>(b) | 65<br>(b) | (b)<br>(b) | 233<br>233 | 77<br>73 | 533<br>45 | 76<br>-- |\n| 58 (See fig. 2(c)) | 1650<br>2000<br>2790 | 82.5<br>100.0<br>139.5 | 89.5<br>90.5<br>96.0 | 65<br>--<br>-- | 77<br>--<br>-- | 55<br>--<br>-- | 65<br>--<br>-- | 56<br>--<br>-- | 70<br>--<br>-- | (b)<br>--<br>-- | 76<br>--<br>-- | (b)<br>--<br>-- | (b)<br>--<br>-- | (b)<br>--<br>-- | (b)<br>--<br>-- | (b)<br>--<br>-- | 360<br>--<br>-- | 75<br>--<br>-- | 80<br>--<br>-- | 55<br>--<br>-- |\n| 59 [Figure: Diagram of muffler configuration with dimensions 2 1/4, 2 1/4, 2 1/4, 2, 11, 12, 11, 34. (see fig. 1)] | 1650<br>2000 | 82.5<br>100.0 | 84.0<br>87.0 | 58<br>55 | 84<br>85 | 62<br>58 | 68<br>64 | (b)<br>53 | (b)<br>(b) | (b)<br>77 | (b)<br>86 | (b)<br>64 | 72<br>57 | (b)<br>52 | (b)<br>58 | (b)<br>(b) | 56<br>-- | 46<br>-- | 225<br>-- | 66<br>-- | Low |\n| 60 [Figure: Diagram of muffler configuration with dimensions 2, 2 1/4, 2 1/4, 2, 3 1/4, 34, 24.] | 1650<br>2000 | 82.5<br>100.0 | 85.0<br>85.0 | 66<br>63 | 85<br>88 | 59<br>67 | 68<br>72 | 56<br>58 | 57<br>60 | 60<br>72 | (b)<br>88 | 69<br>66 | 72<br>60 | (b)<br>60 | 66<br>60 | (b)<br>(b) | 700<br>720 | 66<br>74 | 80<br>-- | 65<br>-- | Low |\n| 61 Muffler 60 reversed | 1650<br>2000<br>2790 | 82.5<br>100.0<br>139.5 | 94.0<br>88.0<br>94.5 | 55<br>--<br>-- | 88<br>--<br>-- | 62<br>--<br>-- | 82<br>--<br>-- | 50<br>--<br>-- | 66<br>--<br>-- | 50<br>--<br>-- | 67<br>--<br>-- | 50<br>--<br>-- | 55<br>--<br>-- | (b)<br>--<br>-- | (b)<br>--<br>-- | (b)<br>--<br>-- | --<br>--<br>-- | --<br>--<br>-- | --<br>--<br>-- | --<br>--<br>-- | Medium |\n| 62 [Figure: Diagram of muffler configuration with dimensions 12 1/2, 2 3/4, 24, 33 1/2. (see fig. 2(e))] | 1650<br>2000<br>2790 | 82.5<br>100.0<br>139.5 | 79.0<br>81.5<br>88.0 | (b)<br>(b)<br>(b) | 62<br>57<br>68 | (b)<br>(b)<br>50 | (b)<br>(b)<br>60 | (b)<br>50<br>65 | (b)<br>65<br>66 | 52<br>66<br>64 | 68<br>52<br>68 | 59<br>55<br>62 | 50<br>52<br>(b) | 50<br>(b)<br>(b) | 50<br>(b)<br>(b) | (b)<br>(b)<br>(b) | 235<br>470<br>2300 | 65<br>65<br>71 | 117<br>115<br>130 | 67<br>55<br>70 | Medium |\n\n$^a$In the sketches of the configurations, all dimensions are in inches and all cross sections are circular, except where otherwise indicated.\n$^b$The sound-pressure level was below the range of the analyzer.\n\nNACA\n\nNACA TN No. 1838\n21\n```", "timestamp": "2026-07-22T05:00:03.953733+00:00"} | |
| {"citation_id": "19930082476", "source_url": "https://ntrs.nasa.gov/api/citations/19930082476/downloads/19930082476.pdf", "page_number": 34, "total_pages": 41, "image_filename": "19930082476_p34.jpg", "text": "32\nNACA TN No. 1801\n\n<!-- Image (38, 175, 853, 801) -->\n\nFigure 5.- Variation of rudder and aileron deflection with wheel position for the model as tested with linked rudder and aileron controls.", "timestamp": "2026-07-22T05:00:04.265676+00:00"} | |
| {"citation_id": "19930082585", "source_url": "https://ntrs.nasa.gov/api/citations/19930082585/downloads/19930082585.pdf", "page_number": 22, "total_pages": 30, "image_filename": "19930082585_p22.jpg", "text": "NACA TN 1907\n21\n\n[Figure: Graph showing Angle of incidence, $\\theta$, deg vs Time after power failure, sec. The graph contains four curves labeled: \"No pitch change\", \"Exponential (slow)\", \"Exponential (moderate)\", and \"Instantaneous\". The NACA logo is present in the bottom right corner of the graph.]\n\nFigure 2.- Types of pitch change considered.", "timestamp": "2026-07-22T05:00:07.921612+00:00"} | |
| {"citation_id": "19930082450", "source_url": "https://ntrs.nasa.gov/api/citations/19930082450/downloads/19930082450.pdf", "page_number": 27, "total_pages": 37, "image_filename": "19930082450_p27.jpg", "text": "26\nNACA TN No. 1778\n\n110\n100\n90\n80\n70\n60\n55\n50\n45\n40\n35\n30\n25\n20\n\n$H/t_w = 21$\n$(b_w/t_w = 20)$\n\n$\\sigma_{cr}, ksi$\n13.8\n6.9\n\n$\\sigma_F, ksi - 40$\n35\n30\n25\n20\n15\n\n$S/t_S$ or $b_S/t_S$\n25\n30\n35\n40\n50\n60\n75\n\nColors indicate minimum-weight proportions for $t_w/t_S = 1.00$.\nRed means some other, blue means no other value of $t_w/t_S$ gives less weight.\n\n$P_t/t_S, ksi$\n110\n100\n90\n80\n70\n60\n55\n50\n45\n40\n35\n30\n\n31\n(30)\n24\n28\n32\n36\n20\n\n$\\sigma_{cr}, ksi$\n18.4\n13.8\n6.9\n\n25\n30\n35\n40\n50\n60\n75\n\n$P_t/t_S, ksi$\n110\n100\n90\n80\n70\n60\n55\n50\n45\n40\n35\n30\n\n41\n(40)\n23\n26\n29\n20\n\n$\\sigma_{cr}, ksi$\n18.4\n13.8\n6.9\n\nNACA\n\n25\n30\n35\n40\n50\n60\n75\n\n.05 .06 .07 .08 .09 .10 .15 .20 .25 .30 .40 .50 .60\n$P_t/L\\sqrt{E}, ksi$\n\nFigure 5: Direct-reading design chart for 24S-T aluminum-alloy Z-stiffened panels. $t_w/t_S = 1.00$.", "timestamp": "2026-07-22T05:00:10.207377+00:00"} | |
| {"citation_id": "19930082712", "source_url": "https://ntrs.nasa.gov/api/citations/19930082712/downloads/19930082712.pdf", "page_number": 14, "total_pages": 14, "image_filename": "19930082712_p14.jpg", "text": "```markdown\n006 - 04-81-51 - Airport-V3VN\n\n1.8\nMaximum section lift\ncoefficient, $c_{l,max}$\n1.6\n1.4\n1.2\n1.0\n\n0\nSection angle\nof zero lift,\n$\\alpha_{l_0}$, deg\n-1.0\n-2.0\n\n.12\nSection lift-\ncurve slope\nper\ndegree, $d c_l/d\\alpha$\n.10\n.08\n\n.008\nSection drag\ncoefficient at\nmeasured $c_{l_1}$\n.006\n.004\n.002\n\n1.0 2.0 3.0 4.0 5.0 10.0 15.0 x $10^6$\nReynolds number\n\n(a) Airfoils with smooth surfaces.\n\n1.4\nMaximum section lift\ncoefficient, $c_{l,max}$\n1.2\n1.0\n0.8\n0.6\n\n0\nSection angle\nof zero lift,\n$\\alpha_{l_0}$, deg\n-1.0\n-2.0\n\n.12\nSection lift-\ncurve slope\nper\ndegree, $d c_l/d\\alpha$\n.10\n.08\n\n.014\nSection drag\ncoefficient at\nmeasured $c_{l_1}$\n.012\n.010\n.008\n\n1.0 2.0 3.0 4.0 5.0 10.0 15.0 x $10^6$\nReynolds number\n\n(b) Airfoils with standard leading-\nedge roughness.\n\n$\\circ$ NACA 8-H-12\n$\\square$ NACA 0012\n$\\diamond$ NACA 23012\n\nFigure 2.- Variation with Reynolds number of section maximum lift coefficient, angle of zero lift,\nlift-curve slope, and drag coefficient at design lift coefficient for the NACA 8-H-12 airfoil\nsection, in comparison with the NACA 0012 and NACA 23012 sections.\n\n12\n\nNACA TW 1998\n```", "timestamp": "2026-07-22T05:00:12.702965+00:00"} | |
| {"citation_id": "19930085487", "source_url": "https://ntrs.nasa.gov/api/citations/19930085487/downloads/19930085487.pdf", "page_number": 11, "total_pages": 36, "image_filename": "19930085487_p11.jpg", "text": "NACA RM No. E8J22\n9\n\n[Figure: Electromagnetic coil, amplifier, and oscillator used to determine natural frequencies of compressor blades.]\n\nFigure 1. - Electromagnetic coil, amplifier, and oscillator used to determine natural frequencies of compressor blades.\n\nNACA\nC-14374\n2-27-46", "timestamp": "2026-07-22T05:00:15.129546+00:00"} | |
| {"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 30, "total_pages": 99, "image_filename": "19930082511_p30.jpg", "text": "28\nNACA TN No. 1826\n\nsince, as is clear from figure 15, $a \\neq 1$. The interference velocity at a point on the axis ($z = iy$) due to a vortex on the axis is thus\n\n$$\nq_1\\left(\\xi+\\frac{i}{2}, \\xi_1+\\frac{i}{2}\\right) = - \\frac{i\\Gamma'y_1y}{y^2 - y_1^2} \\sqrt{\\frac{1+y^2}{a^2+y^2} \\frac{a^2+y_1^2}{1+y_1^2}} + \\frac{i\\Gamma'}{2\\pi(\\xi - \\xi_1)} \\quad (19)\n$$\n\nwhich is normal to the axis.\n\nThe interference velocity at the vortex itself is the limit of expression (17) as $z$ approaches $z_1$. Proceeding as before gives\n\n$$\nq_1(\\xi_1) = i \\left[ \\frac{\\Gamma'}{4} + \\frac{\\Gamma'z_1^2}{2} \\frac{a^2 - 1}{(a^2 - z_1^2)(1 - z_1^2)} + (B + 2Cz_1)\\sqrt{\\frac{1 - z_1^2}{a^2 - z_1^2}} + \\frac{\\Gamma'z_1}{4iy_1} \\right] \\quad (20)\n$$\n\nFor the case in which the vortex is on the axis ($z_1 = iy_1$) the normal velocity at the free boundary ($z = x$, where $1 < x < a$) is given by\n\n$$\nQ(x) = i \\left( \\frac{1}{x - iy_1} - \\frac{1}{x + iy_1} \\right) Bx \\sqrt{\\frac{1 - x^2}{a^2 - x^2}}\n$$\n\nor\n\n$$\nQ(x) = \\frac{2iy_1}{x^2 + y_1^2} Bx \\sqrt{\\frac{x^2 - 1}{a^2 - x^2}} \\quad (21)\n$$\n\nCASE 4 - CLOSED-OPEN-CLOSED TUNNEL WITH UNEQUAL PRESSURES ON THE FREE SURFACES\n\nBoundary conditions.- As indicated in part I, the two-dimensional closed-open-closed tunnel may develop unequal pressures on the two free surfaces if a closed space exists below the lower free surface. Within the limits of the present linear theory, this pressure difference corresponds to superposing on the flow discussed in the preceding section an additional perturbation velocity field $Q(z)$ that\n\n(1) Has no singularities within the tunnel\n\n(2) Satisfies the condition of continuity at the inlet lips", "timestamp": "2026-07-22T05:00:18.150525+00:00"} | |
| {"citation_id": "19930085519", "source_url": "https://ntrs.nasa.gov/api/citations/19930085519/downloads/19930085519.pdf", "page_number": 10, "total_pages": 46, "image_filename": "19930085519_p10.jpg", "text": "```markdown\nNACA RM No. L8K19\n9\n\nof the wing with the full-span slotted flap at $\\delta_f = 50^\\circ$ and vanes B\non and off indicated that better gliding characteristics would be obtained\nfor the vanes-off condition. Such may not be the case, however, at other\nflap deflections.\n\nLateral Control Characteristics\n\nPlug ailerons (flap retracted).— Figure 17(a) shows the variation\nwith angle of attack of the rolling-moment and yawing-moment coefficients\nproduced by various projections of the plug aileron with the sharp plug-\nslot lower lip. The rolling-moment coefficient increased with plug-\naileron projection and with angle of attack to an angle of attack from\n$14^\\circ$ to $16^\\circ$, at which point the rolling-moment-coefficient curve showed a\nsharp decrease. This decrease is caused by the abrupt tip stalling of\nthis particular wing, as has been mentioned previously. The maximum\nrolling moment produced by this plug-aileron configuration was obtained\nat $\\delta_p = -7$ percent and was about the same value as that produced by\nspoiler 18 of reference 1 at the same spoiler projection. In the lower\nangle-of-attack range, however, the rolling moments were lower for the\nplug aileron with $\\delta_p = -5$ percent and $-7$ percent than for spoiler 18 of\nreference 1 at comparable projections and angles of attack. At the lower\nplug-aileron projections, the rolling moments were higher over the angle-\nof-attack range than those produced by spoiler 18, and the reversal of\nrolling moment noted for spoiler 18 did not occur. However, the plug\naileron with the sharp plug-slot lower lip was ineffective in producing\nfavorable rolling moment at low plug-aileron projections.\n\nIn an attempt to remedy the plug-aileron ineffectiveness at low pro-\njections, the plug-slot lower lip was faired to offer a better air inlet\n(as shown in fig. 9). The faired plug-slot lower lip improved the\neffectiveness of the plug aileron at all projections and increased the\nmaximum rolling moments approximately 20 percent. (See fig. 17(b).) How-\never, there was still a large reduction in rolling-moment coefficient at\nall projections at angles of attack above the wing-tip stall angle.\n\nIn the low and moderate angle-of-attack range, the yawing-moment coef-\nficients produced by the plug ailerons with either the sharp or the faired\nplug-slot lower lip were of the same sign (positive) as the rolling-moment\ncoefficients (a condition usually referred to as favorable yaw) and were\nequal to about 30 to 40 percent of the rolling-moment coefficient at the\nmaximum values of rolling-moment coefficient. The yawing moments usually\nbecame negative above an angle of attack of about $11^\\circ$ to $13^\\circ$, which is in\nproximity to the angle of attack at which the wing tip stalled and the\npitching moments became unstable.\n```", "timestamp": "2026-07-22T05:00:19.787310+00:00"} | |
| {"citation_id": "19930085471", "source_url": "https://ntrs.nasa.gov/api/citations/19930085471/downloads/19930085471.pdf", "page_number": 12, "total_pages": 28, "image_filename": "19930085471_p12.jpg", "text": "```markdown\nTABLE II\n\nFLUTTER DATA ON A CANTILEVER MODEL WITH TIP WEIGHTS\n\nWing chord = $\\frac{1}{3}$ ft; elastic axis = 47 percent from L. E.; $r_\\alpha^2 = 0.23$ (for bare wing with no tip weights); weight of bare wing = 0.0806 lb; length of wing = $6\\frac{1}{4}$ in.; tip-weight center of gravity coincides with the wing center of gravity\n\n| Frequency (cps) | | | Center of gravity of wing and tip weights (percent chord) | $c_w \\omega_\\alpha$ | Tip weights (lb) | | Moment of inertia of tip weights about c. g. (in.-lb-sec$^2$) |\n| :---: | :---: | :---: | :---: | :---: | :---: | :---: | :---: |\n| Torsion | Bending | Flutter | | | L. E. | T. E. | |\n| 244 | 134 | 153 | 50 | 510 | 0 | 0 | 0 |\n| 133 | 102 | 105 | 50 | 278 | .00949 | .00949 | .000162 |\n| 103 | 81 | 86 | 49 | 216 | .01766 | .01754 | .000342 |\n| 80 | 68 | 74 | 53 | 167 | .02747 | .03044 | .000686 |\n\nNACA\n\n10\nUNCLASSIFIED\nRECONFIDENTIAL\nNACA RM No. L8J11\n```", "timestamp": "2026-07-22T05:00:22.726640+00:00"} | |
| {"citation_id": "19930085536", "source_url": "https://ntrs.nasa.gov/api/citations/19930085536/downloads/19930085536.pdf", "page_number": 8, "total_pages": 20, "image_filename": "19930085536_p8.jpg", "text": "6\nNACA RM No. E8K05\n\nIn connection with the linearized theory, this relation is usually approximated as\n\n$$C_P = - 2 \\frac{U_x}{U} \\quad (6)$$\n\nTheoretical calculations for several bodies showed that the pressure distributions predicted by the two relations were enough different that the approximate equation omits more terms from the exact relation than is justified. Therefore, although the use of the exact relation for the pressure coefficient may be mathematically inconsistent with the approximations of the linearized theory, it has been used in the theoretical calculations presented herein, except where otherwise noted.\n\nFor systematic calculations of flows at angles of attack or yaw, the procedure outlined in reference 6 is too tedious. A simpler means is to consider the flow slightly inclined with respect to the x-axis rather than to move the body relative to this axis. This procedure of turning the flow rather than the body in obtaining the angle-of-attack solution means that the Mach cones are assumed to follow the body rather than the flow. Although in the actual case the Mach cones would follow the flow more closely than the body, this assumption was made to facilitate numerical calculations. If the angle of attack or yaw is kept small, such an assumption should introduce little error. The boundary condition was obtained in the same manner as in reference 6.\n\n$$U_r - \\left(\\frac{1}{r} \\frac{\\partial r}{\\partial \\theta}\\right) U_\\theta - \\frac{r}{x} U_x = U \\left[ \\frac{\\frac{r}{x}}{\\sqrt{\\tan^2 \\alpha + \\tan^2 \\Psi + 1}} - \\frac{1}{\\sqrt{1 + \\cot^2 \\Psi \\sec^2 \\alpha}} \\left( \\cos \\theta + \\frac{1}{r} \\frac{\\partial r}{\\partial \\theta} \\sin \\theta \\right) - \\frac{1}{\\sqrt{1 + \\cot^2 \\alpha \\sec^2 \\Psi}} \\left( \\sin \\theta - \\frac{1}{r} \\frac{\\partial r}{\\partial \\theta} \\cos \\theta \\right) \\right] \\quad (7)$$", "timestamp": "2026-07-22T05:00:28.710204+00:00"} | |
| {"citation_id": "19930082476", "source_url": "https://ntrs.nasa.gov/api/citations/19930082476/downloads/19930082476.pdf", "page_number": 35, "total_pages": 41, "image_filename": "19930082476_p35.jpg", "text": "NACA TN No. 1801\n33\n\n<!-- Image (142, 137, 921, 824) -->\n\nFigure 6.- Typical motion of the model with elevator deflected to 13° up\nand wheel set full with the spin (loading 3). Pictures taken at\n64 frames per second.\nNACA", "timestamp": "2026-07-22T05:00:29.603737+00:00"} | |
| {"citation_id": "19930082542", "source_url": "https://ntrs.nasa.gov/api/citations/19930082542/downloads/19930082542.pdf", "page_number": 21, "total_pages": 53, "image_filename": "19930082542_p21.jpg", "text": "20\nNACA TN No. 1867\n\nTABLE I.- ROOM-TEMPERATURE PHYSICAL PROPERTIES OF LOW-CARBON M-125 BAR STOCK\n\n<!-- Table (117, 109, 875, 937) -->\n\\begin{tabular}{|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|}\n\\hline\n\\multicolumn{3}{|c|}{\\textbf{Heat treatment:}} & \\multicolumn{3}{|c|}{\\textbf{Aging treatment (a)}} & \\multicolumn{2}{|c|}{\\textbf{Hot-cold-rolling (b)}} & \\textbf{Brinell hardness} & \\multicolumn{4}{|c|}{\\textbf{Room-temperature tensile properties}} & \\textbf{Elongation in 2 in. (percent)} & \\textbf{Reduction of area (percent)} \\\\\n\\cline{1-16}\n\\multicolumn{3}{|c|}{\\textbf{Solution treatment}} & \\multicolumn{3}{|c|}{} & \\multicolumn{2}{|c|}{} & & \\multicolumn{4}{|c|}{\\textbf{Offset yield strength (psi)}} & & \\\\\n\\cline{1-16}\n\\textbf{Temper-} & \\textbf{Time} & \\textbf{Method} & \\textbf{Temper-} & \\textbf{Time} & \\textbf{Temper-} & \\textbf{Percent} & & & \\textbf{0.02 percent} & \\textbf{0.1 percent} & \\textbf{0.2 percent} & \\textbf{Propor-} & & \\\\\n\\textbf{ature} & \\textbf{(hr)} & \\textbf{of} & \\textbf{ature} & \\textbf{(hr)} & \\textbf{ature} & \\textbf{reduc-} & & & & & & \\textbf{tional} & & \\\\\n\\textbf{($^\\circ$F)} & & \\textbf{cooling} & \\textbf{($^\\circ$F)} & & \\textbf{($^\\circ$F)} & \\textbf{tion} & & & & & & \\textbf{limit} & & \\\\\n& & \\textbf{(c)} & & & & & & & & & & \\textbf{(psi)} & & \\\\\n\\hline\n\\multicolumn{16}{|c|}{\\textbf{Aging and rolling hot-rolled bar stock}} \\\\\n\\hline\n(a) & --- & (a) & 1350 & 34 & --- & --- & 233 & 128,500 & 72,500 & 76,500 & 78,500 & 57,500 & 40.5 & 52.7 \\\\\n--- & --- & --- & 1350 & 34 & --- & --- & 214 & 124,500 & 52,700 & 62,000 & 66,500 & 317,500 & 36 & 38.2 \\\\\n--- & --- & --- & 1500 & 34 & --- & --- & 212 & 135,000 & 55,000 & 65,000 & 69,500 & 39,000 & 36 & 40.8 \\\\\n--- & --- & --- & --- & --- & --- & --- & 212 & 135,000 & 54,500 & 59,500 & 62,500 & 39,000 & 36 & 40.8 \\\\\n--- & --- & --- & --- & --- & 75 & 10 & 307 & 145,750 & 92,500 & 123,000 & 128,700 & 57,500 & 31 & 53.0 \\\\\n--- & --- & --- & --- & --- & 1200 & 10 & 312 & 147,000 & 97,500 & 127,000 & 132,000 & 57,500 & 31 & 53.0 \\\\\n--- & --- & --- & --- & --- & 1200 & 15 & 312 & 147,750 & 96,500 & 128,000 & 133,000 & 57,500 & 31 & 53.0 \\\\\n--- & --- & --- & --- & --- & 1200 & 15 & 328 & 146,700 & 101,000 & 132,000 & 137,000 & 70,000 & 21.5 & 43.0 \\\\\n--- & --- & --- & --- & --- & 1200 & 20 & 328 & 145,700 & 101,500 & 135,500 & 140,000 & 82,500 & 22 & 43.5 \\\\\n--- & --- & --- & --- & --- & 1200 & 20 & 321 & 146,500 & 102,000 & 136,000 & 140,500 & 82,500 & 22 & 43.5 \\\\\n--- & --- & --- & 9100 & 924 & --- & --- & 321 & 146,500 & 102,000 & 136,000 & 140,500 & 82,500 & 22 & 43.5 \\\\\n--- & --- & --- & --- & --- & (f) & (f) & 325 & 146,500 & 102,000 & 136,000 & 140,500 & 82,500 & 22 & 43.5 \\\\\n--- & --- & --- & --- & --- & (f) & (f) & 325 & 146,500 & 102,000 & 136,000 & 140,500 & 82,500 & 22 & 43.5 \\\\\n--- & --- & --- & --- & --- & (f) & (f) & 189 & --- & --- & --- & --- & --- & --- & --- \\\\\n\\hline\n\\multicolumn{16}{|c|}{\\textbf{Solution-treated at 1800$^\\circ$ F}} \\\\\n\\hline\n1800 & 2 & W.Q. & --- & --- & --- & --- & 207 & 125,000 & 91,000 & 97,500 & 62,000 & 30,000 & 38 & 46.8 \\\\\n1800 & 2 & W.Q. & --- & --- & 1200 & 15 & 207 & 135,000 & 104,000 & 129,500 & 127,000 & 85,000 & 28.5 & 42.1 \\\\\n\\hline\n\\multicolumn{16}{|c|}{\\textbf{Solution-treated at 1950$^\\circ$ F}} \\\\\n\\hline\n1950 & 2 & W.Q. & --- & --- & --- & --- & 198 & 120,250 & 90,000 & 97,000 & 61,000 & 49,000 & 49 & 64.8 \\\\\n1950 & 2 & W.Q. & --- & --- & 1200 & 15 & 274 & 148,250 & 109,000 & 125,000 & 129,000 & 70,000 & 26 & 52.5 \\\\\n\\hline\n\\end{tabular}\n\nAll aging treatments preceded hot-cold-rolling except where noted.\nAll hot-cold-rolled material was given a final stress relief at 1200$^\\circ$ F for 1 hr.\nW.Q., water-quenched; A.C., air-cooled.\nAged after rolling.\n(a) Solution-treated at 1950$^\\circ$ F (2-percent reduction from 1800$^\\circ$ to 1450$^\\circ$ F.\n(b) 20-percent reduction at 1400$^\\circ$ F; repeated five more times; then 1800$^\\circ$ F 2 hr, air-cooled.\n(c) 63000$^\\circ$ F 2 hr, air-cooled to 1400$^\\circ$ F; 20-percent reduction at 1400$^\\circ$ F; repeated five more times; then 1800$^\\circ$ F 2 hr, air-cooled.\n\n[Figure: NACA logo]", "timestamp": "2026-07-22T05:00:31.037207+00:00"} | |
| {"citation_id": "19930082914", "source_url": "https://ntrs.nasa.gov/api/citations/19930082914/downloads/19930082914.pdf", "page_number": 17, "total_pages": 66, "image_filename": "19930082914_p17.jpg", "text": "```markdown\n16\nNACA TN No. 1857\n\nThen, by use of equation (2),\n\n$$\n\\frac{Y}{b} = \\frac{L}{\\lambda_0} (n - 1) \\left( \\frac{\\rho'}{\\rho} - 1 \\right)\n$$\n\nor\n\n$$\n\\frac{\\rho'}{\\rho} = \\frac{Y}{b} \\left( \\frac{\\lambda_0}{L} \\frac{1}{n - 1} \\right) + 1\n$$\n\nIf $\\frac{Y}{b}$, the fringe shift in terms of fringe width, is designated by $S(y)$, and the quantity in parentheses is designated by $C$, then\n\n$$\n\\frac{\\rho'}{\\rho} = CS(y) + 1 \\tag{3}\n$$\n\nIn this equation $\\rho'/\\rho$ is the density ratio between some position in the disturbance and the undisturbed air.\n\nThe technique of obtaining the values of the fringe shifts for a given cross section of the flow field is simplified by the plotting of graphs. An interferogram is taken of the undisturbed fringes. (See, for example, fig. 10.) Then interferograms of the flow field are taken. (See fig. 11.) Then enlargements to about 7 diameters are made. A position along the horizontal axis is chosen at which the density variation is to be determined. At that position on the enlargement of the undisturbed fringe pattern, a vertical line is drawn, and the positions of the fringes are measured with a scale. (A piece of cross-section paper pasted on the picture is useful.) Then the fringe positions are plotted as a function of the vertical coordinate $y$. Such a plot is shown in figure 12. (A convenient position, such as the lower edge of the nozzle opening, is chosen as the zero of the y-coordinate. This position was located on the picture in the following manner: Two small pieces of drill rod were inserted into the lower nozzle block. From their actual diameter and their measured diameter on the picture, the magnification over actual size was determined. From the known distances of these rods to the inside edge of the nozzle, the location of the edge could be found on the picture. Two small pointers placed in the camera, the shadows of which can be seen in\n```", "timestamp": "2026-07-22T05:00:36.903480+00:00"} | |
| {"citation_id": "19930085487", "source_url": "https://ntrs.nasa.gov/api/citations/19930085487/downloads/19930085487.pdf", "page_number": 12, "total_pages": 36, "image_filename": "19930085487_p12.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:00:37.464156+00:00"} | |
| {"citation_id": "19930082450", "source_url": "https://ntrs.nasa.gov/api/citations/19930082450/downloads/19930082450.pdf", "page_number": 28, "total_pages": 37, "image_filename": "19930082450_p28.jpg", "text": "NACA TN No. 1778\n\n27\n\n$\\frac{b_s}{t_s}$ or $\\frac{b_s}{t_s}$\n\n$\\bar{\\sigma}_f$, ksi\n\n$H/t_w = 26$\n\n$(b_w/t_w = 25)$\n\n$\\sigma_{cr}$, ksi\n\n13.8\n\n8.9\n\n$H$\n\n$b_w$\n\n$t_w$\n\n$b_s$ or $S$\n\n$t_s$\n\n$P_l/t_s$, ksi\n\n(35)\n\n$\\sigma_{cr}$, ksi\n\n19.2\n\n13.8\n\n8.9\n\n(45)\n\n$\\sigma_{cr}$, ksi\n\n18.7\n\n19.4\n\n$\\sigma_{cy} = 44$ ksi\n\n24S-T\n\n$P_l/LA\\sqrt{S}$, ksi\n\n$t_w/t_s = 1.00$\n\nFigure 5-Concluded.\n\nNACA", "timestamp": "2026-07-22T05:00:42.023921+00:00"} | |
| {"citation_id": "19930085544", "source_url": "https://ntrs.nasa.gov/api/citations/19930085544/downloads/19930085544.pdf", "page_number": 1, "total_pages": 33, "image_filename": "19930085544_p1.jpg", "text": "Copy No. 128\nRM No. L8K26\n\nNACA RM No. L8K26\nC1. c. 128\n\n[Figure: NACA logo]\n\nRESEARCH MEMORANDUM\n\nENGINE RING DEPT. LIBRARY\nCHANCE-VOUGHT AIRCRAFT\nDALLAS, TEXAS\n\nCALCULATION OF AERODYNAMIC FORCES ON A\nPROPELLER IN PITCH OR YAW\n\nBy\nJohn L. Crigler and Jean Gilman, Jr.\n\nLangley Aeronautical Laboratory\nLangley Air Force Base, Va.\n\nCLASSIFICATION CHANGED TO Unclassified\nBY AUTHORITY OF NASA Bull. # 6\nON 1/29/60 OF G&R\n\nCLASSIFIED DOCUMENT\nThis document contains information affecting the National Defense of the United States within the meaning of the Espionage Act, Title 18, U.S.C., Sections 793 and 794, the transmission or the revelation of its contents in any manner to an unauthorized person is prohibited by law.\nInformation concerning military and naval services of the United States, appropriate civilian defense employment of the Federal Government, or a regulated defense thereof, or United States citizens of known loyalty to persons who, if necessary must be informed.\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\nWASHINGTON\nFebruary 15, 1949\n\nRM L8H26", "timestamp": "2026-07-22T05:00:43.759955+00:00"} | |
| {"citation_id": "19930085471", "source_url": "https://ntrs.nasa.gov/api/citations/19930085471/downloads/19930085471.pdf", "page_number": 13, "total_pages": 28, "image_filename": "19930085471_p13.jpg", "text": "UNCLASSIFIED\nRESTRICTED\nCONFIDENTIAL\n\nNACA RM No. L8J11\n\nHot room\nEntrance cone\nTest section\nQuick operating valve\nTo vacuum sphere\n\nCONFIDENTIAL\nNACA\n\nFigure 1.- Diagram of the supersonic flutter apparatus.\n\n11", "timestamp": "2026-07-22T05:00:52.780432+00:00"} | |
| {"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 31, "total_pages": 99, "image_filename": "19930082511_p31.jpg", "text": "NACA TN No. 1826\n\n(3) Has a horizontal component equal to, say, +1 on the upper free boundary and -1 on the lower free boundary\n\n(4) Has no vertical component on the closed boundaries\n\n(5) Is zero at infinity upstream and downstream\n\nRewritten as conditions on the complex velocity Q(z) in the z-plane (fig. 15), these conditions become\n\n(1) Q(z) has no singularities in the upper half of the z-plane\n\n(2) Q(z) = 0 at z = ±1\n\n(3) On the real axis, Q(z) = -1 for 1 < z < a, and Q(z) = +1 for -a < z < -1\n\n(4) On the real axis, I.P.Q(z) = 0 for |z| < 1 and for |z| > a\n\n(5) Q(z) = 0 for z = 0 and for z = ∞\n\nOutline of method.- Consider the following two functions of z:\n\n$$\nw_1 = \\int_0^z \\frac{dz}{\\sqrt{(1 - z^2)(a^2 - z^2)}}\n$$\n\n$$\nw_2 = \\int_0^z \\sqrt{\\frac{a^2 - z^2}{1 - z^2}} \\, dz\n$$\n\nThey can be considered as complex velocities having the following properties along the real axis (compare reference 10):\n\n$w_1$ is real between 1 and -1; between 1 and a, or between -1 and -a, its real part is constant but an imaginary part is introduced; beyond a or -a, the imaginary part is constant while the real part approaches zero; R.P.$w_1$(z) = -R.P.$w_1$(-z); I.P.$w_1$(z) = I.P.$w_1$(-z)\n\n$w_2$ has the same properties as $w_1$ except that beyond a and -a its real part approaches ∞ and -∞, respectively", "timestamp": "2026-07-22T05:00:53.723683+00:00"} | |
| {"citation_id": "19930085519", "source_url": "https://ntrs.nasa.gov/api/citations/19930085519/downloads/19930085519.pdf", "page_number": 11, "total_pages": 46, "image_filename": "19930085519_p11.jpg", "text": "10\nNACA RM No. L5K19\n\nPresented for comparison in figure 18 are the rolling-moment coefficients against plug-aileron projection at various angles of attack for the plug-aileron configurations with a sharp and a faired plug-slot lower lip and for spoiler 18 (from reference 1). These data show the reversal of rolling effectiveness of the spoiler at low projections and the elimination of the reversal by use of the plug aileron. Also shown is the increase in rolling effectiveness (noted previously) obtained with the faired plug-slot lower lip compared to the effectiveness obtained with the sharp plug-slot lower lip.\n\nPlug aileron (flap deflected).— Figure 19 shows the rolling-moment and yawing-moment coefficients produced by the plug aileron with the faired plug-slot lower lip and with the full-span flap deflected 30° at the optimum nose position. In general, the rolling-moment coefficient increased with increasing plug-aileron projection and increased slightly with increasing angle of attack to the angle of attack for the tip stall (approximately 10°). Comparison of the plug-aileron data of figures 17(b) and 19 shows that deflection of the full-span slotted flap resulted in an increase in the maximum rolling-moment coefficient produced by the plug aileron of about 130 percent over the rolling-moment coefficient produced by the plug aileron on the unflapped wing.\n\nAt low angles of attack, the yawing-moment coefficients produced by the plug ailerons with the full-span slotted flap deflected were generally of the same sign as the rolling-moment coefficients, except at projections of -1/2 percent and -1 percent where the sign was the opposite of the rolling-moment coefficient. The yawing-moment coefficients were about 10 percent to 15 percent of the rolling-moment coefficient at the maximum value of rolling-moment coefficient. The yawing moments became negative above an angle of attack of about 10° which, for the flap-deflected condition, is the angle of attack at which the wing tip stalled.\n\nEffect of plug-aileron actuating-arm configuration.— The plug-aileron actuating arms were normally open as shown in figure 8. In order to determine the effects of this opening, the actuating arms were filled-in to the wing surface in such a manner as to form a solid actuating arm.\n\nThe data of figures 20(a) ($\\delta_F = 0^\\circ$) and 20(b) ($\\delta_F = 50^\\circ$) indicate that with the flap neutral, the filled-in actuating arms had little effect on the rolling-moment coefficients produced by the plug aileron. With the flap deflected, however, the rolling-moment coefficients produced by the plug aileron with the filled-in actuating arms were generally lower by as much as 13 percent than the rolling moments produced by the plug aileron with the open actuating arms.\n\nThe yawing-moment coefficients produced by the plug aileron with the filled-in actuating arms were slightly higher at the maximum plug projections ($\\delta_p = -0.07c$) than those produced by the plug with the open", "timestamp": "2026-07-22T05:00:56.221753+00:00"} | |
| {"citation_id": "19930093773", "source_url": "https://ntrs.nasa.gov/api/citations/19930093773/downloads/19930093773.pdf", "page_number": 44, "total_pages": 47, "image_filename": "19930093773_p44.jpg", "text": "1159\n\nNACA RM E9G09\n\nAltitude\n(ft)\n○ 5,000\n□ 15,000\n◇ 25,000\n△ 35,000\n▽ 45,000\n\nEngine windmilling speed, N, rpm\n5000\n4000\n3000\n2000\n1000\n0\n100 200 300 400 500 600 700 800\nTrue airspeed, $V_0$, mph\n\n[Figure: A line graph plotting Engine windmilling speed against True airspeed. Data points are marked with symbols corresponding to different altitudes as defined in the legend. A NACA logo is present in the bottom right corner of the plot area.]\n\nFigure 8. - Variation of engine windmilling speed with true airspeed at altitudes from 5000 to 45,000 feet.\n\n43", "timestamp": "2026-07-22T05:00:58.365664+00:00"} | |
| {"citation_id": "19930082585", "source_url": "https://ntrs.nasa.gov/api/citations/19930082585/downloads/19930082585.pdf", "page_number": 23, "total_pages": 30, "image_filename": "19930082585_p23.jpg", "text": "22\nNACA TN 1907\n\nInduced velocity, v, ft/sec\n30\nk = 1.0; no pitch change\n20\nPitch change to 4°\nk = 1.0\nk = ∞\n10\n0\n2\n4\n6\n8\n10\nTime after power failure, sec\nNACA\n\nFigure 3.- Assumed variations of induced velocity with time after power failure.", "timestamp": "2026-07-22T05:01:03.342055+00:00"} | |
| {"citation_id": "19930082476", "source_url": "https://ntrs.nasa.gov/api/citations/19930082476/downloads/19930082476.pdf", "page_number": 36, "total_pages": 41, "image_filename": "19930082476_p36.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:01:03.638577+00:00"} | |
| {"citation_id": "19930085487", "source_url": "https://ntrs.nasa.gov/api/citations/19930085487/downloads/19930085487.pdf", "page_number": 13, "total_pages": 36, "image_filename": "19930085487_p13.jpg", "text": "NACA RM No. E8J22\n\n[Figure: Sand pattern on turbine blades with measurement scale and NACA logo]\n\nC-14377 \n2-27-46\n\nFigure 2. - Sand pattern for second bending-mode vibration on first-stage blade. Frequency of vibration, 1940 cycles per second.\n\n11", "timestamp": "2026-07-22T05:01:14.418669+00:00"} | |
| {"citation_id": "19930085536", "source_url": "https://ntrs.nasa.gov/api/citations/19930085536/downloads/19930085536.pdf", "page_number": 9, "total_pages": 20, "image_filename": "19930085536_p9.jpg", "text": "NACA RM No. E5K05\n\nThe advantage of using this method is that one source distribution may be used for all angles of attack or yaw and the only variable with the angle is therefore the strength of each source. Furthermore, for small angles the strength of each source will be linearly proportional to the angle of attack or yaw.\n\nInasmuch as the free stream is no longer in the axial direction relative to the body, the pressure-coefficient relations (equations (5) and (6)) must be revised. These relations become\n\n$$\nC_{P} = \\frac{2}{\\gamma M^{2}} \\left( \\left\\{ 1 - \\frac{\\gamma-1}{2} M^{2} \\left[ 2 \\frac{U_{x}'}{U} + \\left( \\frac{U_{x}}{U} \\right)^{2} + \\left( \\frac{U_{y}}{U} \\right)^{2} + \\left( \\frac{U_{\\theta}}{U} \\right)^{2} \\right] \\right\\}^{\\frac{\\gamma}{\\gamma-1}} - 1 \\right)\n$$\n\nand\n\n$$\nC_{P} = -2 \\frac{U_{x}'}{U}\n$$\n\nwhere\n\n$$\nU_{x}' = \\frac{U_{x}}{\\sqrt{\\tan^{2} \\alpha + \\tan^{2} \\psi + 1}} + \\frac{U_{r} \\cos \\theta - U_{\\theta} \\sin \\theta}{\\sqrt{1 + \\cot^{2} \\psi \\sec^{2} \\alpha}} + \\frac{U_{r} \\sin \\theta + U_{\\theta} \\cos \\theta}{\\sqrt{1 + \\cot^{2} \\alpha \\sec^{2} \\psi}}\n$$\n\nThe source configurations used to calculate the pressure distribution over the test body are shown in figure 3. For angles of attack and yaw of $0^\\circ$, sources 1 to 7 were used and the strengths of sources 1 and 7, 2 and 6, and 3 and 5 were respectively equal. These positions were found with the aid of the rules given in reference 6. Instead of putting the source nearest to a peak at the center of curvature of the peak, a better approximation is to place this source at the focus of the peak. This procedure is similar to that sometimes employed in subsonic-flow problems solved by source distributions. For angle of attack, sources 1 to 7 were used with different strengths. For angle of yaw, all the sources were used. In this case, the strengths of sources 1 and 7, 2 and 6, 3 and 5, 8 and 14, 9 and 13, and 10 and 12 were respectively equal. The positions and the number of sources added for yaw were arbitrarily chosen, except that the sources could not be close to the surface. With the exception of this limitation, the accuracy of the solution is insensitive to small changes in the position of sources 8 to 14.", "timestamp": "2026-07-22T05:01:15.441708+00:00"} | |
| {"citation_id": "19930085542", "source_url": "https://ntrs.nasa.gov/api/citations/19930085542/downloads/19930085542.pdf", "page_number": 4, "total_pages": 46, "image_filename": "19930085542_p4.jpg", "text": "2\nNACA RM No. L8L29\n\nINTRODUCTION\n\nThe advantages of low-aspect-ratio pointed wings for high-speed flight have been outlined by Jones in reference 1. Extensive theoretical investigations of the stability characteristics of triangular wings have been made in the supersonic speed range. Theoretical investigations in the subsonic speed range, however, are very limited. The theory of reference 2 is considered to be applicable only to thin triangular wings of aspect ratios less than 0.5 and the theory of references 3 and 4 presents only a few of the stability characteristics. The approximate theory of swept wings presented in reference 5 might be expected to be unreliable as the taper ratio decreases from unity.\n\nThe present investigation was conducted to determine experimentally the effects of changes in profile and aspect ratio on the low-speed static-stability and rolling characteristics of triangular wings. The investigation was extended to determine the effects of adding fins to the upper surface and of cutting portions from the tips of a triangular wing to form low-aspect-ratio tapered wings. Rolling characteristics were determined by means of the rolling-flow equipment of the Langley stability tunnel. (See reference 6.) The experimental data are compared with available theory.\n\nSYMBOLS\n\nThe data presented herein are in the form of standard NACA coefficients of forces and moments which are referred to the stability system of axes with the origin at the projection of the quarter-chord point of the mean aerodynamic chord on the plane of symmetry. The positive directions of the forces, moments, and angular displacements are shown in figure 1. The symbols and coefficients used herein are defined as follows:\n\n| | |\n| :--- | :--- |\n| $C_L$ | lift coefficient ($L/qS$) |\n| $C_{L_{max}}$ | maximum lift coefficient |\n| $C_X$ | longitudinal-force coefficient ($X/qS$) |\n| $C_Y$ | lateral-force coefficient ($Y/qS$) |\n| $C_l$ | rolling-moment coefficient ($L'/qSb$) |\n| $C_m$ | pitching-moment coefficient ($M/qS\\bar{c}$) |", "timestamp": "2026-07-22T05:01:22.400933+00:00"} | |
| {"citation_id": "19930082914", "source_url": "https://ntrs.nasa.gov/api/citations/19930082914/downloads/19930082914.pdf", "page_number": 18, "total_pages": 66, "image_filename": "19930082914_p18.jpg", "text": "NACA TN No. 1857\n\nthe lower part of the interferograms, served to locate, in the pictures that did not include the nozzle end, the horizontal line that would extend to the nozzle edge.) (The two pieces of drill rod were also used for checking the alinement of the light beam with the nozzle edge. They are placed on opposite sides of the center of the nozzle. If the alinement is correct, then when, on an interferogram, the proper distance is measured from each rod to give the location of the edge of the nozzle opening, the same location is obtained for both measurements.) Then, on the enlargement of the interferogram of the flow field, at the same position horizontally, the positions of the fringes were measured on a vertical scale. Then obtaining the fringe shift as function of $y$ was simply a matter of going to the first plot, figure 12, with a value of $y$ and the corresponding fringe number, reading there the undisplaced fringe number for the same value of $y$, and subtracting. For example, suppose that fringe number 20 lay at $y = 17$ in the flow interferogram, and fringe number 8.4 lay at $y = 17$ in the no-flow interferogram; then the fringe shift at $y = 17$ is 11.6 fringe widths. The density ratio between the place where $y = 17$ and the undisturbed air is obtained from equation (3). This procedure was carried out for each cross section through the flow field where it was decided to obtain the density distribution.\n\nThe \"constant\" $C = \\frac{\\lambda_0}{L} \\frac{1}{n - 1}$ was determined for each cross section through the jet. The value of $\\lambda_0$ was 5170 angstroms. The value of $n - 1$ was determined from equation (1). The value of $\\rho$, the density of room air, was obtained from pressure and temperature measurements. The best critical-table value of $k$ is 0.1167 cubic foot per slug for 5170 angstroms. The width of the nozzle opening was 2.999 inches, and this value would be the value of $L$ if there were no end-effects. A corrected value of $L$ was arbitrarily obtained, as follows: The fringe shift was plotted against $y$ from the undisturbed region, through the boundary layer, into the free-stream region of the jet. The area under the curve was obtained and divided by the fringe shift in the free stream. This gave a position at which the same fringe shift would occur if it took place abruptly, rather than gradually through the boundary layer. The difference between this position and the position of the edge of the nozzle was subtracted twice from the actual nozzle width to obtain the value of $L$ used in the \"constant\" $C$. This method is sufficiently accurate for cross sections close to the nozzle.\n\nEffect of Refraction\n\nWhen light passes through a medium in which the index of refraction varies in a direction that is perpendicular to the initial direction of propagation of the light, the direction of propagation is changed, or the light is refracted. In the mixing region of the jet considered herein, the density of the air varies, and consequently also the index of refraction varies. The purpose of the present section is to determine", "timestamp": "2026-07-22T05:01:25.535971+00:00"} | |
| {"citation_id": "19930082450", "source_url": "https://ntrs.nasa.gov/api/citations/19930082450/downloads/19930082450.pdf", "page_number": 29, "total_pages": 37, "image_filename": "19930082450_p29.jpg", "text": "28\nNACA TN No. 1778\n\n$$\n\\frac{H}{t_w} = 21\n$$\n$$\n\\left( \\frac{b_w}{t_w} = 20 \\right)\n$$\n\n$$\n\\bar{\\sigma}_f, \\text{ksi}\n$$\n\n$$\n\\sigma_{cr}, \\text{ksi}\n$$\n7.3\n\n$$\n\\frac{S_c}{t_s} \\text{ or } \\frac{b_s}{t_s}\n$$\n\n$$\n\\frac{P_l}{L\\sqrt{c}}, \\text{ksi}\n$$\n\nColors indicate minimum weight proportions for $\\frac{t_w}{t_s} = 0.51$.\nRed means some other, blue means no other value of $\\frac{t_w}{t_s}$ gives less weight.\n\n$$\n\\frac{31}{(30)}\n$$\n\n$$\n\\sigma_{cr}, \\text{ksi}\n$$\n11.4\n7.3\n\n$$\n\\frac{S_c}{t_s} \\text{ or } \\frac{b_s}{t_s}\n$$\n\n$$\n\\frac{P_l}{L\\sqrt{c}}, \\text{ksi}\n$$\n\n$$\n\\frac{41}{(40)}\n$$\n\n$$\n\\sigma_{cr}, \\text{ksi}\n$$\n18.2\n11.4\n7.3\n\n$$\n\\frac{S_c}{t_s} \\text{ or } \\frac{b_s}{t_s}\n$$\n\n$$\n\\frac{P_l}{L\\sqrt{c}}, \\text{ksi}\n$$\n\nNACA\n\n$$\n\\frac{P_l}{t_s}, \\text{ksi}\n$$\n\nFigure 6.-Direct-reading design chart (alternate form) for 24S-T aluminum-alloy Z-stiffened panels, $\\frac{t_w}{t_s} = 0.51$.", "timestamp": "2026-07-22T05:01:26.078326+00:00"} | |
| {"citation_id": "19930082485", "source_url": "https://ntrs.nasa.gov/api/citations/19930082485/downloads/19930082485.pdf", "page_number": 39, "total_pages": 62, "image_filename": "19930082485_p39.jpg", "text": "38\nNACA TN No. 1810\n\n1.65\nNACA\n1.60\n$\\frac{P_t}{P_s}$ 1.55\nRatio of stagnation-to-static pressure,\n1.50\n1.45\n1.40\n1.35\n1.30\n1.25\n1.20\n9.0 9.4 9.8 10.2 10.6 11.0 11.4 11.8\nRadius, r, in.\n\nExperimental\nDesign\n\nInner shroud\nOuter shroud\n\nFigure 8. - Comparison of ratios of experimental and design\nstagnation-to-static pressure at 0.1 chord downstream of\nblades.\n\n1026", "timestamp": "2026-07-22T05:01:26.670945+00:00"} | |
| {"citation_id": "19930085471", "source_url": "https://ntrs.nasa.gov/api/citations/19930085471/downloads/19930085471.pdf", "page_number": 14, "total_pages": 28, "image_filename": "19930085471_p14.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:01:28.135040+00:00"} | |
| {"citation_id": "19930085544", "source_url": "https://ntrs.nasa.gov/api/citations/19930085544/downloads/19930085544.pdf", "page_number": 2, "total_pages": 33, "image_filename": "19930085544_p2.jpg", "text": "NACA RM No. L8K26\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nRESEARCH MEMORANDUM\n\nCALCULATION OF AERODYNAMIC FORCES ON A\nPROPELLER IN PITCH OR YAW\n\nBy John L. Crigler and Jean Gilman, Jr.\n\nSUMMARY\n\nAn analysis has been made to determine the applicability of existing propeller theory and the theory of oscillating airfoils to the problem of determining the magnitude of the forces on propellers in pitch or yaw. Strip calculations including the Goldstein correction factors and using compressible airfoil characteristics were first made as though steady-state conditions existed successively at several blade positions of the propeller blades during one revolution. A theory of oscillating airfoils in pulsating incompressible linearized potential flow was then considered from which it was possible to determine factors which would modify the forces as calculated under the assumption of steady-state compressible flow.\n\nComparisons of the steady-state calculations with experimental results show that the magnitude of the force changes experienced by the blades can be predicted with satisfactory accuracy. Results of calculations made by the oscillating theory indicate that the actual forces on the blade may be somewhat lower than the values calculated by the steady-state method. It was not possible to establish this conclusion definitely because of the lack of sufficient experimental data for comparison.\n\nThe turning moment on the shaft of a two-blade propeller fluctuates between approximately zero and its maximum value twice per revolution. For the operating condition investigated the turning moment on the shaft of a three-blade propeller remains nearly constant at about 75 percent of the maximum value attained with the two-blade propeller.\n\nINTRODUCTION\n\nLarge-diameter propellers incorporating thin blade sections are becoming a necessity for certain aircraft installations using large unit power plants at high altitude and high speed. On such propeller installations, the oscillating air forces due to yaw or pitch of the propeller axis may cause dangerous vibratory stresses with a frequency of once per", "timestamp": "2026-07-22T05:01:28.558463+00:00"} | |
| {"citation_id": "19930082614", "source_url": "https://ntrs.nasa.gov/api/citations/19930082614/downloads/19930082614.pdf", "page_number": 20, "total_pages": 36, "image_filename": "19930082614_p20.jpg", "text": "18\nNACA TN 1939\n\nThe relations for speed as a function of time when the airplane is in a $60^\\circ$ dive are shown in figure 9.\n\nA more accurate solution results from a step-by-step calculation. The velocity relations have been calculated by this method for comparison with the relations given by the equations. In order for the results of the step-by-step calculations to be comparable to the formulas, the assumption is again made that there is no variation in $C_{Dn}$. However, the effect of the variation in density is included, instead of assuming an average value. The step-by-step calculations for an initial altitude of 25,000 feet are presented in table II and the results of the calculations are shown in figure 9.\n\nCalculations were made also to indicate the effect of a lag in the time for the drag due to the air brakes to reach its full value. It was assumed that the increase in drag caused by extending the brakes takes place during the interval between 1 and 2 seconds after the brake actuation is started. The amount by which the curve is displaced (fig. 9) indicates the gain in braking effect that can be realized by designing the brakes for minimum delay in opening.\n\nIt is seen from the slopes of the velocity curves in figure 9 that a change in flight-path angle from level flight to a dive of $60^\\circ$ results in a change from an initial deceleration of 18 feet per second squared to an initial acceleration of 6.8 feet per second squared for the assumed airplane with aerodynamic brakes at an altitude of 25,000 feet.\n\nEntry Into a Dive\n\nAn example in which the flight-path angle is variable is provided in the calculations of the speed during a dive entry. The same initial speed, altitudes, and coefficients as in the preceding example are used.\n\nIt is assumed that the airplane is flown so that from level flight the indicated normal acceleration factor decreases to -1.5 within the first second and is then held constant until the airplane is in a $60^\\circ$ dive. The detailed calculations are presented in table III for an initial altitude of 25,000 feet. The results, plotted in figure 10, show the variation of longitudinal acceleration, dive angle, and speed with time for initial altitudes of 25,000 and 10,000 feet.\n\nCONCLUDING REMARKS\n\nAerodynamic brakes afford a means of avoiding undesired increases in speed during the operation of an airplane, make possible rapid decelerations in flight, and allow a considerable increase in the angle of descent at constant speed. A measure of the utility of aerodynamic", "timestamp": "2026-07-22T05:01:29.135133+00:00"} | |
| {"citation_id": "19930082617", "source_url": "https://ntrs.nasa.gov/api/citations/19930082617/downloads/19930082617.pdf", "page_number": 21, "total_pages": 58, "image_filename": "19930082617_p21.jpg", "text": "20\nNACA TN 1962\n\nStringers\n$\\circ$ 1 to 9\n$\\times$ 10 to 16\n\nMoment\n(in. - lb)\n1 $36.0 \\times 10^3$\n2 $108.0 \\times 10^3$\n3 $180.0 \\times 10^3$\n4 $252.0 \\times 10^3$\n\n[Figure: Cross-section diagram showing Band L, dimension 2.57\", and section A-A with 45° angle]\n\nDistance from horizontal diameter, in.\nStrain\n\n| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | |", "timestamp": "2026-07-22T05:01:40.415213+00:00"} | |
| {"citation_id": "19930082498", "source_url": "https://ntrs.nasa.gov/api/citations/19930082498/downloads/19930082498.pdf", "page_number": 23, "total_pages": 49, "image_filename": "19930082498_p23.jpg", "text": "TABLE II.- RESULTS OF MUFFLER INVESTIGATION MADE WITH PROPELLER REMOVED - Concluded\n\n| Muffler configuration | Engine speed (rpm) | Fundamental firing frequency, F (cps) | Sound-pressure level (db) | | | | | | | | | | | | | | | Other sounds | | | Back pressure |\n|---|---|---|---|---|---|---|---|---|---|---|---|---|---|---|---|---|---|---|---|---|---|\n| | | | Over-all | 0.5F | 1.0F | 1.5F | 2.0F | 2.5F | 3.0F | 3.5F | 4.0F | 4.5F | 5.0F | 5.5F | 6.0F | 6.5F | 7.0F | cps | db | cps | db | |\n| [Figure: Diagram (a) showing a rectangular box with dimensions 24, 36, 16, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 24, 2", "timestamp": "2026-07-22T05:01:40.925048+00:00"} | |
| {"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 32, "total_pages": 99, "image_filename": "19930082511_p32.jpg", "text": "30\nNACA TN No. 1826\n\nMaps of the two functions are shown in figure 16. It should\nobviously be possible to find a linear combination of these two func-\ntions, $Mw_1 + Nw_2$, such that for $1 < z < a$\n$$R.P.(Mw_1 + Nw_2) = -1$$\nfor $-a < z < -1$\n$$R.P.(Mw_1 + Nw_2) = +1$$\nand beyond $a$ or $-a$\n$$I.P.(Mw_1 + Nw_2) = 0$$\nA simple additional function $w_3$ to be discussed subsequently is needed\nto satisfy the condition at infinity. The desired velocity function for\nthe closed-open-closed tunnel with unequal pressures is thus of the form:\n$$Q(z) = Mw_1(z) + Nw_2(z) + w_3(z)$$\nThe constants M and N are derived in the two following sections.\n\nEvaluation of integrals.- In the following development, the modulus\nof all the elliptic integrals is $1/a$; the modulus will therefore not be\nindicated in the symbols. In the designations for the incomplete\nelliptic integrals, E, E', F, and F', the terms in parentheses are the\nupper limits of integration. Then\n$$w_1(1) = R.P.w_1(a) = \\int_0^1 \\frac{dz}{\\sqrt{(1-z^2)(a^2-z^2)}} = \\frac{1}{a}K$$\n$$w_2(1) = R.P.w_2(a) = \\int_0^1 \\sqrt{\\frac{a^2-z^2}{1-z^2}} dz = aE$$\n$$I.P.w_1(a) = \\int_1^a \\frac{dz}{\\sqrt{(1-z^2)(a^2-z^2)}}$$\nwhich by the substitution $z^2 = a^2 - (a^2-1)t^2$ reduces to\n$$\\frac{1}{a} \\int_0^1 \\frac{dt}{\\sqrt{(1-t^2)\\left(1 - \\frac{a^2-1}{a^2}t^2\\right)}} = \\frac{1}{a}K'$$", "timestamp": "2026-07-22T05:01:41.164118+00:00"} | |
| {"citation_id": "19930082476", "source_url": "https://ntrs.nasa.gov/api/citations/19930082476/downloads/19930082476.pdf", "page_number": 37, "total_pages": 41, "image_filename": "19930082476_p37.jpg", "text": "NACA TN No. 1801\n35\n\n[Figure: A series of sequential photographs of an airplane in flight, arranged in vertical strips. The strips are labeled with numbers at the bottom and top, indicating time or frame sequence. The numbers visible are 66, 72, 78, 84, 90, 96, 102, 108, 114, and 120. The airplane is shown from various angles as it maneuvers.]\n\nFigure 6.- Continued.\nNACA", "timestamp": "2026-07-22T05:01:47.431537+00:00"} | |
| {"citation_id": "19930085519", "source_url": "https://ntrs.nasa.gov/api/citations/19930085519/downloads/19930085519.pdf", "page_number": 12, "total_pages": 46, "image_filename": "19930085519_p12.jpg", "text": "NACA RM No. L9K19\n\nactuating arms. The yawing-moment coefficients produced by the plug at smaller projections were only slightly affected by variation of the actuating-arm configuration.\n\nEffect of gap between wing upper surface and plug-aileron lower edge.— At plug-aileron projections of $\\delta_{\\mathrm{p}} = -0.03c$ or greater, the lower edge of each plug segment emerged from the upper surface of the wing at the inboard end of the plug segment in such a manner that a wedge-shaped gap existed between the wing upper surface and the plug-segment lower edge. Figure 21 shows that filling in this gap resulted in an appreciable increase in rolling-moment coefficient over that produced in the gap-open condition with the plug at $\\delta_{\\mathrm{p}} = -0.07c$ but had little effect on the rolling-moment coefficient at $\\delta_{\\mathrm{p}} = -0.05c$. This effect of gap between the plug lower edge and the wing upper surface has been obtained previously for plug ailerons on unswept wings (references 9 and 10).\n\nFilling-in the gap between the plug-aileron lower edge and the wing upper surface increased slightly the yawing-moment coefficient produced by the plug aileron at both $\\delta_{\\mathrm{p}} = -0.05c$ and $-0.07c$.\n\nHalf-span plain aileron.— The rolling-moment and yawing-moment characteristics of the wing with the $0.20c$ by $0.49 \\frac{b}{2}$ ailerons are shown in figure 22(a) with the flap neutral and in figure 22(b) with the half-span slotted flap deflected $50^\\circ$. For both the flap-neutral and flap-deflected conditions, the rolling-moment coefficient increased with increasing aileron deflections and decreased as the wing angle of attack was increased either positively or negatively from $\\alpha = 0^\\circ$. At angles of attack below the wing-stall angle, the values of total rolling-moment coefficient, for any combination of equal up-aileron and down-aileron deflections, are equal to or slightly higher for the slotted flap-deflected condition than with the flap neutral.\n\nFor both flap conditions, the total yawing-moment coefficient resulting from an equal up and down deflection of the aileron was generally small at angles of attack below the wing stall and was adverse (sign of yawing moment opposite to sign of rolling moment). At angles of attack below the wing-stall angle the total adverse yawing-moment coefficient produced by the aileron on the wing with the flap deflected, although small, was somewhat greater than that produced by the aileron on the wing with the flap neutral. The total yawing-moment coefficients produced by the plain aileron at angles of attack greater than the wing-stall angle were higher than those at low angles of attack for both flap conditions.", "timestamp": "2026-07-22T05:01:48.536639+00:00"} | |
| {"citation_id": "19930085487", "source_url": "https://ntrs.nasa.gov/api/citations/19930085487/downloads/19930085487.pdf", "page_number": 14, "total_pages": 36, "image_filename": "19930085487_p14.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:01:49.112308+00:00"} | |
| {"citation_id": "19930082245", "source_url": "https://ntrs.nasa.gov/api/citations/19930082245/downloads/19930082245.pdf", "page_number": 41, "total_pages": 66, "image_filename": "19930082245_p41.jpg", "text": "```markdown\n1.6\n1.4\n1.2\n1.0\n.8\n.6\n.4\n.2\n0\n-.2\n-.4\n-.6\n-.8\n.1 .2 .3 .4 .5 .6 .7 .8 .9\nMach number, M\nAileron section normal-force coefficient, $C_{n\\alpha}$\n\n$\\delta a$\n(deg)\n30\n18\n12\n4\n2\n0\n-2\n-4\n-6\n-12\n\n.16\n.12\n.08\n.04\n0\n-.04\n-.08\n-.12\n-.16\n-.20\n-.24\n-.28\n-.58\n-.60\n.1 .2 .3 .4 .5 .6 .7 .8 .9\nMach number, M\nAileron section hinge-moment coefficient, $C_h$\n\n$\\delta a$\n(deg)\n-12\n+6\n+4\n-2\n0\n2\n4\n12\n18\n30\n\n(h) $c_n = 0.8$.\nFigure 7 — Concluded.\n\n40\nNACA TN NO. 1596\n```", "timestamp": "2026-07-22T05:01:57.261303+00:00"} | |
| {"citation_id": "19930085536", "source_url": "https://ntrs.nasa.gov/api/citations/19930085536/downloads/19930085536.pdf", "page_number": 10, "total_pages": 20, "image_filename": "19930085536_p10.jpg", "text": "8\nNACA RM No. E5K05\n\nRESULTS AND DISCUSSION\n\nExperimental data for several angles of yaw and attack are presented in figures 4 and 5. Theoretical calculations based on the linearized theory using equations (8) and (9) for the pressure coefficient are also shown for comparison. The experimental points represent the average of the pressures at corresponding stations on the body. Data were obtained for angles of yaw ranging from -16° to 16° and angles of attack from -10° to 10°. Schlieren observation indicated no shock separation on the cone or interference from the shock caused by the strut over the range of angles of the investigation.\n\nThe linearized theory using equation (8) agrees well with the experimental results for moderate angles of yaw (fig. 4). As the angle was increased, the deviation between theory and experiment slightly increased on the compressive side of the cone. On the expansive side the agreement remained good, which is to be expected because an angle of yaw of 6° corresponds to zero flow deflection on the midpoint of this side. The increasing variation between theory and experiment with increasing flow angle is also illustrated by the fact that the agreement is best over the slenderest parts of the body, that is, the parts of the body that least disturb the flow.\n\nComparison of the effects of using the complete equation for the pressure coefficient (equation (8)) with the use of the linearized one (equation (9)) shows that the values predicted by the use of the linearized relation are consistently high, especially when the flow deflection is large.\n\nThe linearized theory using equation (8) shows close agreement with experiment throughout the range of angles of attack over which the experiments were conducted (fig. 5). The excellent agreement between theory and experiment at an angle of attack of 10°, especially at the station $\\theta = -90^\\circ$ where the flow angle was 28.5°, indicates that the effect of assuming that the Mach cones follow the body rather than the flow is negligible. The Mach angle corresponding to the experimental Mach number is about 32°.\n\nThe variation of pressure coefficient with angle of attack at a station is predicted very closely by the linearized theory using equation (8) (fig. 6). The linearized relation for the pressure coefficient (equation (9)) did not show nearly as good agreement with experiment, nor did it correctly predict the rate of change", "timestamp": "2026-07-22T05:01:59.536019+00:00"} | |
| {"citation_id": "19930085542", "source_url": "https://ntrs.nasa.gov/api/citations/19930085542/downloads/19930085542.pdf", "page_number": 5, "total_pages": 46, "image_filename": "19930085542_p5.jpg", "text": "NACA RM No. L8L29\n\n$C_n$ yawing-moment coefficient (N/qSb)\n\nL lift, pounds\n\nX longitudinal force, pounds\n\nY lateral force, pounds\n\nL' rolling moment about X-axis, foot-pounds\n\nM pitching moment about Y-axis, foot-pounds\n\nN yawing moment about Z-axis, foot-pounds\n\nA aspect ratio $\\left(\\frac{b^2}{S}\\right)$\n\nb span, feet\n\nS area, square feet\n\nc local chord measured parallel to plane of symmetry, feet\n\n$\\overline{c}$ mean aerodynamic chord, feet $\\left(\\frac{2}{S} \\int_0^{b/2} c^2 \\, dy\\right)$\n\n$c_r$ root chord, feet\n\n$\\lambda$ taper ratio\n\nx longitudinal distance from apex of triangle to quarter-chord point of any chordwise section, feet\n\n$\\overline{x}$ longitudinal distance from apex of triangle to quarter-chord point of mean aerodynamic chord, feet $\\left(\\frac{2}{S} \\int_0^{b/2} cx \\, dy\\right)$\n\nR Reynolds number\n\n$\\rho$ density of air, slugs per cubic foot\n\nV free-stream velocity, feet per second", "timestamp": "2026-07-22T05:02:04.723382+00:00"} | |
| {"citation_id": "19930082914", "source_url": "https://ntrs.nasa.gov/api/citations/19930082914/downloads/19930082914.pdf", "page_number": 19, "total_pages": 66, "image_filename": "19930082914_p19.jpg", "text": "18\nNACA TN No. 1857\n\nwhether the refraction, or bending, of the light has a significant effect on the mixing-region density distribution obtained from interferograms.\n\nLet $y$ be the vertical coordinate through the mixing zone and $z$ the horizontal coordinate in the direction of the light beam. Because the air density in the mixing region increases as $y$ increases, the light will be refracted in the manner indicated qualitatively in the following figure:\n\n[Figure: A graph with a vertical axis labeled $y$ and a horizontal axis labeled $z$. A curved line labeled \"Light ray\" starts near the origin and curves upward. A tangent line is drawn to the curve, forming an angle $\\phi$ with the horizontal.]\n\nIf $\\phi$ represents the angle of incidence, then Snell's law states that, at any point in the medium,\n\n$$ \\sin \\phi = \\frac{V}{c_1} $$\n\nwhere $V$ is the velocity of the light and $c_1$ is a constant. Accordingly, therefore,\n\n$$ \\tan \\phi = \\frac{\\sin \\phi}{\\sqrt{1 - \\sin^2 \\phi}} $$\n\n$$ = \\sqrt{\\frac{V^2}{c_1^2 - V^2}} $$\n\n$$ = \\frac{dz}{dy} $$\n\n$$ \\frac{dy}{dz} = \\sqrt{\\frac{c_1^2}{V^2} - 1} $$", "timestamp": "2026-07-22T05:02:07.055408+00:00"} | |
| {"citation_id": "19930093773", "source_url": "https://ntrs.nasa.gov/api/citations/19930093773/downloads/19930093773.pdf", "page_number": 45, "total_pages": 47, "image_filename": "19930093773_p45.jpg", "text": "44\n\n$$\n\\frac{D_m}{P_n}\n$$\n\nWindmilling drag \nNet thrust at maximum permissible engine speed,\n\n.16 \n.12 \n.08 \n.04 \n0 \n\n100 200 300 400 500 600 700 \nTrue airspeed, $V_0$, mph\n\n[Figure: Graph showing variation of ratio of windmilling drag to net thrust at maximum permissible engine speed with true airspeed at altitude of 25,000 feet. Curve starts near (100, 0.02) and rises to approximately (650, 0.16). Grid lines are present. NACA logo in bottom right corner of plot area.]\n\nFigure 9. - Variation of ratio of windmilling drag to net thrust at maximum permissible engine speed with true airspeed at altitude of 25,000 feet.\n\nNACA RM E5G09\n\nNACA-Langley - 11-11-49 - 600\n\n1159", "timestamp": "2026-07-22T05:02:09.800386+00:00"} | |
| {"citation_id": "19930082585", "source_url": "https://ntrs.nasa.gov/api/citations/19930082585/downloads/19930082585.pdf", "page_number": 24, "total_pages": 30, "image_filename": "19930082585_p24.jpg", "text": "NACA TN 1907\n23\n\n<!-- Image (124, 104, 857, 761) -->\n\nFigure 4.- Effect of two different assumptions for induced velocity against time on the variations of descending velocity and rotor angular velocity with time after power failure. $v = v_f + (v_0 - v_f)e^{-kt}$. (See fig. 3.)", "timestamp": "2026-07-22T05:02:10.087812+00:00"} | |
| {"citation_id": "19930085471", "source_url": "https://ntrs.nasa.gov/api/citations/19930085471/downloads/19930085471.pdf", "page_number": 15, "total_pages": 28, "image_filename": "19930085471_p15.jpg", "text": "UNCLASSIFIED\nRESTRICTED\nCONFIDENTIAL\n\nNACA RM No. L8J11\n\nHot room\nRecording instruments\nPneumatic cylinder\nQuick operating valve\nVacuum tank\n\nNACA\nL-56783\n\nFigure 2.- General view of the Langley supersonic flutter apparatus.\n\nRESTRICTED\nCONFIDENTIAL\nUNCLASSIFIED\n\n13", "timestamp": "2026-07-22T05:02:10.192867+00:00"} | |
| {"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 20, "total_pages": 78, "image_filename": "19930082618_p20.jpg", "text": "18\nNACA TN 1945\n\nTABLE I\nORDINATES OF THE\nNACA 641-409 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| .377 | -.829 | .623 | -.629 |\n| .613 | 1.021 | .897 | -.711 |\n| 1.095 | 1.331 | 1.405 | -.903 |\n| 2.222 | 1.895 | 2.476 | -1.151 |\n| 4.803 | 2.732 | 5.137 | -1.168 |\n| 7.297 | 3.553 | 7.705 | -1.687 |\n| 9.798 | 3.925 | 10.202 | -1.857 |\n| 14.810 | 4.796 | 15.180 | -2.104 |\n| 19.830 | 5.452 | 20.170 | -2.272 |\n| 24.854 | 5.957 | 25.146 | -2.377 |\n| 29.882 | 6.315 | 30.118 | -2.427 |\n| 34.912 | 6.536 | 35.088 | -2.416 |\n| 39.942 | 6.632 | 40.058 | -2.348 |\n| 44.972 | 6.534 | 45.028 | -2.174 |\n| 50.000 | 6.242 | 50.000 | -1.930 |\n| 55.024 | 6.016 | 54.976 | -1.636 |\n| 60.045 | 5.594 | 59.955 | -1.310 |\n| 65.060 | 5.069 | 64.940 | -.965 |\n| 70.069 | 4.508 | 69.931 | -.616 |\n| 75.072 | 3.858 | 74.928 | -.278 |\n| 80.069 | 3.154 | 79.931 | -.090 |\n| 85.059 | 2.413 | 84.941 | .279 |\n| 90.043 | 1.644 | 89.957 | 1.44 |\n| 95.021 | .858 | 94.979 | 1.06 |\n| 100.000 | 0 | 100.000 | 0 |\n\nL.E. radius: 0.579\nSlope of radius through L.E.: 0.168\n\nTABLE II\nORDINATES OF THE\nNACA 641-412 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| .330 | 1.064 | .662 | -.864 |\n| .569 | 1.305 | .931 | -1.025 |\n| 1.045 | 1.690 | 1.455 | -1.282 |\n| 2.204 | 2.393 | 2.736 | -1.649 |\n| 4.738 | 3.430 | 5.262 | -2.166 |\n| 7.229 | 4.231 | 7.771 | -2.335 |\n| 9.730 | 4.896 | 10.270 | -2.520 |\n| 14.742 | 5.959 | 15.222 | -2.847 |\n| 19.772 | 6.760 | 20.220 | -3.076 |\n| 24.805 | 7.363 | 25.195 | -3.203 |\n| 29.842 | 7.706 | 30.158 | -3.248 |\n| 34.882 | 8.037 | 35.118 | -3.217 |\n| 39.923 | 8.154 | 40.077 | -3.091 |\n| 44.963 | 7.988 | 45.037 | -2.868 |\n| 50.000 | 7.686 | 50.000 | -2.574 |\n| 55.032 | 7.246 | 54.968 | -2.246 |\n| 60.059 | 6.690 | 59.941 | -1.906 |\n| 65.078 | 6.039 | 64.922 | -1.513 |\n| 70.090 | 5.293 | 69.910 | -1.105 |\n| 75.094 | 4.483 | 74.906 | -.693 |\n| 80.089 | 3.619 | 79.911 | -.435 |\n| 85.076 | 2.722 | 84.924 | -.088 |\n| 90.055 | 1.818 | 89.945 | .250 |\n| 95.027 | .919 | 94.973 | .345 |\n| 100.000 | 0 | 100.000 | 0 |\n\nL.E. radius: 1.040\nSlope of radius through L.E.: 0.168\n\nTABLE III\nORDINATES OF THE\nNACA 642-415 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| .299 | 1.291 | .701 | -1.091 |\n| .526 | 1.579 | .974 | -1.259 |\n| .956 | 2.036 | 1.504 | -1.610 |\n| 2.207 | 2.883 | 2.723 | -2.132 |\n| 4.673 | 4.121 | 5.257 | -2.857 |\n| 7.162 | 5.075 | 7.838 | -3.379 |\n| 9.662 | 5.864 | 10.338 | -3.736 |\n| 14.681 | 7.122 | 15.319 | -4.136 |\n| 19.714 | 8.066 | 20.286 | -4.382 |\n| 24.756 | 8.771 | 25.244 | -4.470 |\n| 29.803 | 9.260 | 30.197 | -4.372 |\n| 34.853 | 9.541 | 35.147 | -4.121 |\n| 39.904 | 9.614 | 40.096 | -3.730 |\n| 44.954 | 9.414 | 45.046 | -3.294 |\n| 50.000 | 9.016 | 50.000 | -2.868 |\n| 55.040 | 8.456 | 54.960 | -2.476 |\n| 60.072 | 7.762 | 59.928 | -2.149 |\n| 65.096 | 6.954 | 64.904 | -1.834 |\n| 70.111 | 6.055 | 69.889 | -1.467 |\n| 75.115 | 5.084 | 74.885 | -1.344 |\n| 80.109 | 4.062 | 79.891 | -.878 |\n| 85.092 | 3.020 | 84.899 | -.328 |\n| 90.066 | 1.982 | 89.934 | .006 |\n| 95.032 | .976 | 94.968 | .288 |\n| 100.000 | 0 | 100.000 | 0 |\n\nL.E. radius: 1.590\nSlope of radius through L.E.: 0.168\n\nTABLE IV\nORDINATES OF THE\nNACA 643-418 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| .263 | 1.508 | .737 | -1.308 |\n| .486 | 1.840 | 1.014 | -1.560 |\n| .850 | 2.370 | 1.530 | -1.842 |\n| 2.152 | 3.357 | 2.848 | -2.413 |\n| 4.609 | 4.800 | 5.391 | -3.336 |\n| 7.095 | 5.908 | 7.905 | -4.212 |\n| 9.595 | 6.823 | 10.405 | -4.715 |\n| 14.617 | 8.277 | 15.393 | -5.183 |\n| 19.657 | 9.366 | 20.343 | -5.182 |\n| 24.707 | 10.176 | 25.283 | -4.996 |\n| 29.763 | 10.730 | 30.227 | -4.642 |\n| 34.823 | 11.037 | 35.177 | -4.217 |\n| 39.883 | 11.093 | 40.115 | -3.809 |\n| 44.945 | 10.820 | 45.055 | -3.440 |\n| 50.000 | 10.220 | 50.000 | -3.068 |\n| 55.047 | 9.635 | 54.953 | -2.555 |\n| 60.086 | 8.799 | 59.907 | -2.274 |\n| 65.114 | 7.861 | 64.886 | -1.721 |\n| 70.131 | 6.784 | 69.869 | -1.296 |\n| 75.135 | 5.635 | 74.855 | -.874 |\n| 80.127 | 4.477 | 79.873 | -1.293 |\n| 85.108 | 3.294 | 84.889 | -.602 |\n| 90.077 | 2.132 | 89.923 | -.061 |\n| 95.037 | 1.030 | 94.963 | .234 |\n| 100.000 | 0 | 100.000 | 0 |\n\nL.E. radius: 1.590\nSlope of radius through L.E.: 0.168\n\nNACA", "timestamp": "2026-07-22T05:02:10.359869+00:00"} | |
Xet Storage Details
- Size:
- 86.8 kB
- Xet hash:
- c4e93d8b949e2f44e3428d902186b1a7402018f62cfd29c33e6858ab4ae9d8f9
·
Xet efficiently stores files, intelligently splitting them into unique chunks and accelerating uploads and downloads. More info.