PDF/ocr_dataset_1008.jsonl

{"citation_id": "19930086022", "source_url": "https://ntrs.nasa.gov/api/citations/19930086022/downloads/19930086022.pdf", "page_number": 29, "total_pages": 34, "image_filename": "19930086022_p29.jpg", "text": "NACA RM L9E24\n27\n\n$C_l$\n.02\n.01\n0\n-.01\n-.02\n-.03\n\nPlain wing\n\n$C_l$\n.02\n.01\n0\n-.01\n-.02\n-.03\n\nLeading and trailing-edge flaps\n\n$C_l$\n.02\n.01\n0\n-.01\n-.02\n-.03\n\nLeading and trailing-edge flaps and fences\n\n$\\delta_a$\n-25 -20 -15 -10 -5 0 5 10 15 20 25\n\na\n$\\nabla$ $\\diamondsuit$ $\\square$ $\\triangle$\n0\n8\n12\n16\n\nNACA\n\nFigure 8.— The variation of rolling-moment characteristics with aileron deflection for various model configurations.", "timestamp": "2026-07-22T04:25:35.905970+00:00"}
{"citation_id": "19930085982", "source_url": "https://ntrs.nasa.gov/api/citations/19930085982/downloads/19930085982.pdf", "page_number": 30, "total_pages": 32, "image_filename": "19930085982_p30.jpg", "text": "28\nNACA RM E9E13\n\n$$225 \\times 10^3$$\n\nCorrected\nweight flow\n$W\\sqrt{\\theta}/\\delta$\n(lb/sec)\n\n| | |\n| :--- | :--- |\n| $\\circ$ | 21.36 |\n| $\\square$ | 19.66 |\n| $\\triangle$ | 18.76 |\n| $\\nabla$ | 18.17 |\n| $\\diamond$ | 16.11 |\n| $+$ | 15.48 |\n\nCorrected enthalpy addition, $\\frac{U \\Delta V_{\\theta}}{\\theta}$, ft-lb/slug\n\n[Figure: Graph showing multiple curves plotted against radius ratio. The curves correspond to the symbols in the legend. A horizontal line labeled \"Design\" is present. The x-axis ranges from .5 to 1.0. The y-axis ranges from 0 to 225x10^3. Labels \"Hub\" and \"Tip\" are at the bottom left and right of the plot area respectively.]\n\nRadius ratio, $r/r_t$\n\nFigure 9. - Radial distribution of energy addition.\n\n1131", "timestamp": "2026-07-22T04:25:41.897828+00:00"}
{"citation_id": "19930086015", "source_url": "https://ntrs.nasa.gov/api/citations/19930086015/downloads/19930086015.pdf", "page_number": 30, "total_pages": 54, "image_filename": "19930086015_p30.jpg", "text": "Vertical distance from tunnel center line, z, in.\n\n|  |  |  |\n| :--- | :--- | :--- |\n| 24 | | |\n| 16 | | |\n| 8 | | |\n| 0 | | |\n| -8 | | |\n| -16 | | |\n| -24 | | |\n| -.04 | 0 | .04 |\n\nFlagged symbols denote survey off center line\n\n|  |  |  |\n| :--- | :--- | :--- |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | 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|\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n| | | |\n|", "timestamp": "2026-07-22T04:25:42.943651+00:00"}
{"citation_id": "19930085930", "source_url": "https://ntrs.nasa.gov/api/citations/19930085930/downloads/19930085930.pdf", "page_number": 64, "total_pages": 92, "image_filename": "19930085930_p64.jpg", "text": "NACA RM L9G07\n63\n\n[Figure: Graph showing Average stagnation-pressure recovery, $\\frac{P_{2av}}{P_o}$, versus Area ratio, $\\frac{A_2}{A_1}$. The graph contains three data series represented by circles, squares, and diamonds. A legend in the top right corner indicates:\n- Circle: $\\frac{p_1}{p_A} = 1.28$\n- Square: $\\frac{p_1}{p_A} = 1.00$\n- Diamond: $\\frac{p_1}{p_A} = .85$\nThe graph includes \"UNCLASSIFIED\" and \"CONFIDENTIAL\" stamps. The NACA logo is present in the bottom right corner of the plot area.]\n\n(b) 10.15-percent-span station.\n\nFigure 26.- Concluded.", "timestamp": "2026-07-22T04:25:43.573560+00:00"}
{"citation_id": "19930085842", "source_url": "https://ntrs.nasa.gov/api/citations/19930085842/downloads/19930085842.pdf", "page_number": 62, "total_pages": 104, "image_filename": "19930085842_p62.jpg", "text": "58\n\n$\\alpha = 23.2^\\circ$\n$\\delta_{a_{TA}}, deg$\n-30\n-20\n-10\n10\n20\n\n$\\alpha = 11.3^\\circ$\n$\\delta_{a_{TA}}, deg$\n-30\n-20\n-10\n10\n20\n\n$\\alpha = -0.6^\\circ$\n$\\delta_{a_{TA}}, deg$\n-30\n-20\n-10\n10\n20\n\nLift coefficient, $C_L$\n.8\n.6\n.4\n.2\n0\n-.2\n\nElevator deflection, $\\delta_a$, deg\n-48\n-40\n-32\n-24\n-16\n-8\n0\n8\n16\n\nNATIONAL ADVISORY\nCOMMITTEE FOR AERONAUTICS\n\nFigure 29.- The effect of right-ailavator tab setting on the variation of $C_L$ with $\\delta_a$. Model in basic configuration; $\\delta_{a_{TL}} = 0^\\circ$; $\\delta_r = 0^\\circ$; propellers removed; $\\delta_r = 0^\\circ$.\n\nNACA RM L9G29", "timestamp": "2026-07-22T04:25:46.424589+00:00"}
{"citation_id": "19930085965", "source_url": "https://ntrs.nasa.gov/api/citations/19930085965/downloads/19930085965.pdf", "page_number": 34, "total_pages": 67, "image_filename": "19930085965_p34.jpg", "text": "NACA RM E9E06\n33\n\nAPPENDIX B\n\nCALCULATION OF RATE OF HEAT REQUIRED TO MAINTAIN SPECIFIED\nBLADE-SURFACE TEMPERATURE IN CONDITIONS OF ICING\n\nAn analysis of the heating requirements for thermal-ice prevention of the inlet guide vane of an axial-flow compressor is presented for specified conditions of icing. The feasibility of a design of an ice-prevention system utilizing eddy-current heating under the assumed icing conditions has been presented in the text of this report.\n\nSymbols\n\nThe following symbols are used in this appendix:\n\n| | |\n| :--- | :--- |\n| A | cross-sectional area of flow passage, sq ft |\n| C | blade chord, ft |\n| $c_p$ | specific heat of air at constant pressure, Btu/(lb)($^\\circ$F) |\n| $c_{p,w}$ | specific heat of water, Btu/(lb)($^\\circ$F) |\n| D | diameter of equivalent leading-edge cylinder, ft |\n| d | average distance between adjacent blades, ft |\n| e | vapor pressure, lb/sq in. |\n| $e_s$ | vapor pressure at saturation at temperature of surface, lb/sq in. |\n| $e_x$ | local vapor pressure, determined by equation (B9), lb/sq in. |\n| g | acceleration due to gravity, 32.2 ft/sec$^2$ |\n| H | total heat dissipation, Btu/(hr)(sq ft) |\n| $H_1$ | heat dissipation due to convection, Btu/(hr)(sq ft) |\n| $H_2$ | heat dissipation due to evaporation, Btu/(hr)(sq ft) |\n| $H_3$ | heat dissipation due to heating of impinging water on surface of blade, Btu/(hr)(sq ft) |", "timestamp": "2026-07-22T04:25:47.468515+00:00"}
{"citation_id": "19930086060", "source_url": "https://ntrs.nasa.gov/api/citations/19930086060/downloads/19930086060.pdf", "page_number": 18, "total_pages": 20, "image_filename": "19930086060_p18.jpg", "text": "16\nNACA RM L9F02\n\nCONFIDENTIAL\n\n$C_P$\n.2\n.1\n0\n-.1\n-.2\n-.3\n-.4\n-.5\n\n$K = 80 \\text{ percent}$\n\n[NACA logo]\n\n(d) Maximum diameter at 80-percent station.\nFigure 5.- Concluded.\nCONFIDENTIAL", "timestamp": "2026-07-22T04:25:51.293142+00:00"}
{"citation_id": "19930082090", "source_url": "https://ntrs.nasa.gov/api/citations/19930082090/downloads/19930082090.pdf", "page_number": 114, "total_pages": 122, "image_filename": "19930082090_p114.jpg", "text": "112\nNACA TN No. 1455\n\n<!-- Image (167, 132, 772, 906) -->\n\nFigure 59.- Thermal output and isothermal frictional pressure drops of copper and stainless-steel fin heat exchanger O.", "timestamp": "2026-07-22T04:25:52.623291+00:00"}
{"citation_id": "19930085529", "source_url": "https://ntrs.nasa.gov/api/citations/19930085529/downloads/19930085529.pdf", "page_number": 81, "total_pages": 85, "image_filename": "19930085529_p81.jpg", "text": "80\nNACA RM No. L8A30a\n\nTABLE 74\n$$[A = -45^\\circ, \\delta_{a_0} = -10.0^\\circ, \\alpha = 2^\\circ]$$\nCONFIDENTIAL\n\n| Tube | Per-cent chord | UPPER SURFACE<br>Mach Number | | | | | Tube | Per-cent chord | LOWER SURFACE<br>Mach Number | | | | |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| | | 0.60 | 0.80 | 0.89 | 0.995 | 0.96 | | | 0.60 | 0.80 | 0.89 | 0.995 | 0.96 |\n| A<br>1<br>2<br>3<br>4<br>5<br>6<br>7<br>8<br>9<br>10<br>11 | 2.0<br>6.0<br>15.0<br>27.5<br>40.0<br>50.0<br>59.0<br>67.5<br>77.5<br>87.5<br>96.0 | --<br>--<br>--<br>--<br>--<br>--<br>-0.096<br>-0.067<br>--<br>--<br>-- | --<br>--<br>--<br>--<br>--<br>--<br>-0.091<br>-0.060<br>--<br>--<br>-- | --<br>--<br>--<br>--<br>--<br>--<br>-0.086<br>-0.059<br>--<br>--<br>-- | --<br>--<br>--<br>--<br>--<br>--<br>-0.092<br>-0.069<br>--<br>--<br>-- | --<br>--<br>--<br>--<br>--<br>--<br>-0.066<br>-0.060<br>--<br>--<br>-- | 86<br>87<br>88<br>89<br>90<br>91<br>92<br>93<br>94 | 2.0<br>10.0<br>25.0<br>41.0<br>50.5<br>62.5<br>72.5<br>84.0<br>94.0 | --<br>--<br>--<br>--<br>--<br>-0.031<br>-0.008<br>--<br>-- | --<br>--<br>--<br>--<br>--<br>-0.031<br>-0.009<br>--<br>-- | --<br>--<br>--<br>--<br>--<br>-0.032<br>-0.008<br>--<br>-- | --<br>--<br>--<br>--<br>--<br>-0.029<br>-0.008<br>--<br>-- | --<br>--<br>--<br>--<br>--<br>-0.049<br>-0.012<br>--<br>-- |\n| B12<br>13<br>14<br>15<br>16<br>17<br>18<br>19<br>20<br>21<br>22 | 2.0<br>6.0<br>15.0<br>27.5<br>40.0<br>50.0<br>59.0<br>67.5<br>77.5<br>88.0<br>95.5 | --<br>--<br>--<br>-0.237<br>-0.217<br>-0.192<br>-0.159<br>-0.103<br>-0.060<br>-0.017<br>-- | --<br>--<br>--<br>-0.250<br>-0.214<br>-0.186<br>-0.172<br>-0.116<br>-0.059<br>-0.017<br>-- | --<br>--<br>--<br>-0.243<br>-0.195<br>-0.163<br>-0.142<br>-0.108<br>-0.047<br>-0.020<br>-- | --<br>--<br>--<br>-0.216<br>-0.205<br>-0.168<br>-0.091<br>-0.074<br>-0.010<br>-0.077<br>-- | --<br>--<br>--<br>-0.249<br>-0.237<br>-0.214<br>-0.192<br>-0.151<br>-0.076<br>-0.001<br>-- | 95<br>96<br>97<br>98<br>99<br>100<br>101<br>102<br>103 | 3.0<br>10.0<br>25.0<br>41.0<br>52.5<br>62.5<br>72.5<br>86.5<br>94.5 | --<br>--<br>--<br>-0.098<br>-0.073<br>-0.062<br>-0.063<br>-0.016<br>-0.077 | --<br>--<br>--<br>-0.105<br>-0.088<br>-0.071<br>-0.066<br>-0.016<br>-0.079 | --<br>--<br>--<br>-0.141<br>-0.109<br>-0.081<br>-0.066<br>-0.016<br>-0.080 | --<br>--<br>--<br>-0.175<br>-0.152<br>-0.126<br>-0.069<br>-0.014<br>-0.089 | --<br>--<br>--<br>-0.245<br>-0.261<br>-0.214<br>-0.078<br>-0.013<br>-0.060 |\n| C23<br>24<br>25<br>26<br>27<br>28<br>29<br>30<br>31<br>32<br>33 | 2.0<br>6.0<br>15.0<br>27.8<br>40.0<br>50.0<br>59.0<br>67.5<br>77.5<br>88.0<br>95.5 | -0.793<br>-0.436<br>-0.364<br>-0.310<br>-0.277<br>-0.221<br>-0.191<br>-0.131<br>-0.083<br>-0.012<br>-- | -0.799<br>-0.439<br>-0.296<br>-0.303<br>-0.327<br>-0.221<br>-0.210<br>-0.146<br>-0.090<br>-0.007<br>-- | -0.451<br>-0.444<br>-0.489<br>-0.294<br>-0.261<br>-0.199<br>-0.196<br>-0.138<br>-0.064<br>-0.019<br>-- | -0.378<br>-0.397<br>-0.450<br>-0.468<br>-0.231<br>-0.211<br>-0.200<br>-0.139<br>-0.122<br>-0.019<br>-- | -0.316<br>-0.299<br>-0.421<br>-0.468<br>-0.291<br>-0.241<br>-0.487<br>-0.188<br>-0.298<br>-0.236<br>-- | 104<br>105<br>106<br>107<br>108<br>109<br>110<br>111<br>112 | 3.0<br>10.0<br>25.0<br>41.0<br>52.5<br>62.5<br>72.5<br>85.1<br>94.6 | -0.210<br>-0.068<br>-0.065<br>--<br>--<br>-0.065<br>-0.026<br>-0.090<br>-0.063 | -0.216<br>-0.081<br>-0.086<br>--<br>--<br>-0.081<br>-0.023<br>-0.016<br>-0.058 | -0.264<br>-0.091<br>-0.119<br>--<br>--<br>-0.098<br>-0.037<br>-0.013<br>-0.051 | -0.190<br>-0.097<br>-0.142<br>--<br>--<br>-0.132<br>-0.061<br>-0.014<br>-0.037 | -0.186<br>-0.096<br>-0.157<br>--<br>--<br>-0.198<br>-0.149<br>-0.096<br>-0.066 |\n| D34<br>35<br>36<br>37<br>38<br>39<br>40<br>41<br>42<br>43 | 2.0<br>15.0<br>27.5<br>40.0<br>50.0<br>59.0<br>67.5<br>77.5<br>87.5<br>94.2 | -0.392<br>-0.297<br>-0.279<br>-0.271<br>-0.246<br>-0.186<br>--<br>-0.086<br>-0.024<br>-0.021 | -0.394<br>-0.339<br>-0.377<br>-0.301<br>-0.293<br>-0.251<br>--<br>-0.104<br>-0.033<br>-0.024 | -0.394<br>-0.339<br>-0.396<br>-0.294<br>-0.337<br>-0.246<br>--<br>-0.143<br>-0.074<br>-0.030 | -0.487<br>-0.363<br>-0.353<br>-0.279<br>-0.345<br>-0.246<br>--<br>-0.143<br>-0.087<br>-0.047 | -0.449<br>-0.361<br>-0.360<br>-0.279<br>-0.298<br>-0.246<br>--<br>-0.145<br>-0.025<br>-0.047 | 113<br>114<br>115<br>116<br>117<br>118<br>119<br>120<br>121 | 3.0<br>10.0<br>25.0<br>41.0<br>52.5<br>61.5<br>72.5<br>87.4<br>94.2 | -0.132<br>-0.068<br>-0.058<br>-0.053<br>-0.090<br>-0.090<br>-0.031<br>-0.043<br>-0.041 | -0.129<br>-0.076<br>-0.071<br>-0.066<br>-0.106<br>-0.090<br>-0.042<br>-0.043<br>-0.033 | -0.114<br>-0.096<br>-0.087<br>-0.096<br>-0.128<br>-0.093<br>-0.061<br>-0.043<br>-0.009 | -0.103<br>-0.096<br>-0.096<br>-0.111<br>-0.136<br>-0.093<br>-0.075<br>-0.043<br>-0.009 | -0.087<br>-0.114<br>-0.108<br>-0.114<br>-0.130<br>-0.130<br>-0.072<br>-0.043<br>-0.021 |\n| E44<br>45<br>46<br>47<br>48<br>49<br>50<br>51<br>52<br>53<br>54 | 2.0<br>6.0<br>15.0<br>27.5<br>40.0<br>50.0<br>59.0<br>67.5<br>77.5<br>89.5<br>95.5 | -0.307<br>-0.284<br>-0.271<br>-0.258<br>-0.254<br>-0.229<br>-0.189<br>-0.129<br>-0.070<br>-0.050<br>-0.004 | -0.315<br>-0.309<br>-0.309<br>-0.294<br>-0.254<br>-0.256<br>-0.189<br>-0.163<br>-0.126<br>-0.037<br>-0.009 | -0.309<br>-0.314<br>-0.309<br>-0.330<br>-0.284<br>-0.280<br>-0.229<br>-0.209<br>-0.126<br>-0.037<br>-0.028 | -0.299<br>-0.309<br>-0.314<br>-0.322<br>-0.322<br>-0.274<br>-0.229<br>-0.229<br>-0.173<br>-0.017<br>-0.026 | -0.247<br>-0.292<br>-0.341<br>-0.340<br>-0.322<br>-0.277<br>-0.229<br>-0.196<br>-0.173<br>-0.012<br>-0.031 | 122<br>123<br>124<br>125<br>126<br>127<br>128<br>129<br>130<br>131 | 3.0<br>10.0<br>25.0<br>41.0<br>52.5<br>62.5<br>72.5<br>78.0<br>87.2<br>94.1 | -0.152<br>-0.026<br>-0.026<br>-0.092<br>-0.079<br>-0.062<br>-0.028<br>-0.015<br>-0.077<br>-0.043 | -0.152<br>-0.022<br>-0.026<br>-0.110<br>-0.079<br>-0.076<br>-0.028<br>-0.020<br>-0.077<br>-0.040 | -0.145<br>-0.009<br>-0.096<br>-0.133<br>-0.079<br>-0.092<br>-0.028<br>-0.017<br>-0.077<br>-0.031 | -0.106<br>-0.006<br>-0.096<br>-0.152<br>-0.079<br>-0.092<br>-0.028<br>-0.013<br>-0.077<br>-0.", "timestamp": "2026-07-22T04:25:52.999918+00:00"}
{"citation_id": "19930086081", "source_url": "https://ntrs.nasa.gov/api/citations/19930086081/downloads/19930086081.pdf", "page_number": 14, "total_pages": 44, "image_filename": "19930086081_p14.jpg", "text": "12 CONFIDENTIAL NACA RM L9H05\n\nrange of the moment curves. Both normal-force and bending-moment coefficients varied almost linearly with control deflection for each angle of attack, indicating that the change in hinge-moment characteristics probably was associated with a change in load distribution near the nose of the control root. Such a change in load distribution could well be expected in this region because of the discontinuity in the chord plane accompanying control deflection combined with a peak angle-of-attack loading near the leading edge. Increasing the control-surface thickness increased the chord-force coefficient. This increase, however, would be of little practical significance from design considerations since the maximum value of chord-force coefficient obtained was no greater than 0.04 and was quite small when compared with normal-force loads. The value of $C_{N_{f\\delta}}$ was decreased about 10 percent, and the hinge-moment characteristics remained unchanged.\n\nA comparison of theory with the experimental results for zero angle of attack (fig. 18) indicates that flat-plate linearized theory predicted reasonably well the variation of the coefficients with control deflection though the theoretical normal-force effectiveness was not fully realized.\n\nCONCLUSIONS\n\nFrom an investigation at a Mach number of 1.9 of a delta wing with half-delta control flap in the Langley 9- by 12-inch supersonic blowdown tunnel, the following conclusions may be drawn:\n\n1. The experimental rolling effectiveness of the control amounted to about 85 percent of that calculated by linearized theory.\n\n2. At zero angle of attack of the wing, the normal-force, hinge-moment, and bending-moment characteristics of the control surface were in reasonable agreement with linearized theory. At small angles of attack the control-surface hinge moment exhibited considerable nonlinear variations with control deflection and with angle of attack.\n\n3. Installation of the fence caused no significant changes in either the aerodynamic characteristics of the complete wing or in the loads and moments of the control surface.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T04:25:53.106735+00:00"}
{"citation_id": "19930086151", "source_url": "https://ntrs.nasa.gov/api/citations/19930086151/downloads/19930086151.pdf", "page_number": 8, "total_pages": 34, "image_filename": "19930086151_p8.jpg", "text": "6\nCONFIDENTIAL\nNACA RM L9J28\n\nDISCUSSION\n\nAerodynamic Characteristics in Pitch\n\nThe lift, drag, and pitching-moment coefficients for the plain wing and for the wing with the end plate are presented in figures 3 and 4, respectively.\n\nThe data of figures 3 and 4 show that a change in aileron plan form had little or no effect on the aerodynamic characteristics of the plain wing or the wing with the end plate. For all configurations investigated, the wing aerodynamic center was between about $0.23\\bar{c}$ and $0.25\\bar{c}$ at the low lift coefficients, and stable pitching-moment characteristics were exhibited at the wing stall.\n\nThe effect on the lift characteristics of adding the end plate to the wing was to increase the lift-curve slope from 0.040 to 0.046 and to decrease the maximum lift coefficient by approximately 0.23. (Compare figs. 3 and 4.) Although the effect of an end plate in increasing the wing lift-curve slope has been found previously on unswept wings (reference 10) and results from an increase in the effective aspect ratio of the wing, the unswept wings also showed an increase in maximum lift coefficient when the end plate was added (references 10 and 11). The aforementioned values of lift-curve slope obtained on the wing configurations reported herein (0.040 on the plain wing and 0.046 on the wing with end plate) correspond to effective aspect ratios of about 1.8 and 2.3, according to the charts of reference 12. It is of interest to note that, although the plain wing with the triangular-plan-form aileron had a geometric aspect ratio of 2.31, its effective aspect ratio was less (about 1.8). The reason for this phenomenon is unknown at present.\n\nThe addition of the end plate to the wing also produced an increase in the values of drag coefficient and an appreciable decrease in the values of the lift-drag ratio over the entire lift-coefficient range (figs. 3 and 4). This increase in drag coefficient was fairly small and constant at low values of lift coefficient (up to about 0.6 lift coefficient) and became fairly large at high values of lift coefficient. The break in the curve of pitching-moment coefficient plotted against lift coefficient and the decrease in the slope of the lift curve of the wing with end plate for values of $C_L$ above 0.6 indicate some form of separation or adverse flow effects at the wing end-plate juncture. This in all probability causes the much larger values of drag coefficient for the wing with end plate. Reference 11, however, indicates that the drag coefficient of unswept wings is less at moderate and large lift coefficients with an end plate installed than without one.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T04:25:53.870350+00:00"}
{"citation_id": "19930085936", "source_url": "https://ntrs.nasa.gov/api/citations/19930085936/downloads/19930085936.pdf", "page_number": 51, "total_pages": 59, "image_filename": "19930085936_p51.jpg", "text": "50\nNACA RM No. E9B03\n\nPressure\ncoefficient\n$C_p$\n.4\n.2\n0\n-.2\n-.2\n0\n.2\n\nAngle of\nyaw\n(deg)\n-12\n-6\n0\n\n[Figure: Polar plot showing radial pressure distributions with three curves corresponding to different yaw angles.]\n\nNACA\n\n(b) x/L = 0.898.\nFigure 10. - Concluded. Radial pressure distributions at 0° angle of\nattack for three yaw angles.", "timestamp": "2026-07-22T04:25:58.121342+00:00"}
{"citation_id": "19930086105", "source_url": "https://ntrs.nasa.gov/api/citations/19930086105/downloads/19930086105.pdf", "page_number": 9, "total_pages": 22, "image_filename": "19930086105_p9.jpg", "text": "```markdown\nNACA RM E5H12 CONFIDENTIAL 7\n\nis presented as a function of the total-pressure recovery $P_3/P_0$ without combustion. Over the range investigated, improved diffuser performance in the presence of combustion is shown by the consistently higher values of optimum total-pressure recovery obtained with the regenerative-type burner. Further evidence of improved diffuser performance is indicated by the slight decrease in optimum $P_3/P_0$ with decreasing cold $P_3/P_0$ or with equivalently increasing values of $A_4/A_1$. Experiments using propylene oxide in conjunction with the perforated conical flame holder resulted in approximately the same curve as that obtained with 62-octane gasoline.\n\nSUMMARY OF RESULTS\n\nA preliminary experimental investigation of pressure fluctuations at the combustion-chamber inlet of a 3.6-inch-ram jet employing a shock diffuser and operating at a Mach number of 1.92 gave the following results for two distinct burner configurations:\n\n1. For each outlet-inlet area ratio investigated, the amplitude of the pressure fluctuations built up gradually as the fuel-air ratio was increased to the value corresponding to optimum diffuser mean static pressure (as indicated on manometers by static orifices). With further increase in fuel-air ratio, the pressure pulsations became more intense and the shock oscillated in and out of the diffuser inlet. The fundamental frequency increased from 6 to 35 cycles per second as the fuel-air ratio was increased beyond the value at optimum mean static pressure.\n\n2. With both burner configurations at the conditions of optimum mean combustion-chamber static pressure, the maximum instantaneous pressure coincided within experimental error with the static pressure at optimum cold recovery. With the conical flame-holder configuration, the minimum instantaneous pressure at optimum mean pressure recovery approximated the cold static pressure corresponding to the particular outlet area.\n\n3. A marked improvement in diffuser performance was obtained with the regenerative-type burner as compared with the perforated conical flame-holder configuration.\n\nLewis Flight Propulsion Laboratory,\nNational Advisory Committee for Aeronautics,\nCleveland, Ohio.\n\nCONFIDENTIAL\n```", "timestamp": "2026-07-22T04:25:58.818914+00:00"}
{"citation_id": "19930085952", "source_url": "https://ntrs.nasa.gov/api/citations/19930085952/downloads/19930085952.pdf", "page_number": 48, "total_pages": 50, "image_filename": "19930085952_p48.jpg", "text": "```markdown\nNACA RM L9C24\n47\n\n<!-- Image (293, 136, 790, 755) -->\n\n(b) $\\alpha \\approx 46^\\circ$; $\\frac{Y}{nD} = 0.37$.\nFigure 22.- Continued.\n```", "timestamp": "2026-07-22T04:26:03.426646+00:00"}
{"citation_id": "19930086003", "source_url": "https://ntrs.nasa.gov/api/citations/19930086003/downloads/19930086003.pdf", "page_number": 28, "total_pages": 36, "image_filename": "19930086003_p28.jpg", "text": "CONFIDENTIAL\n\nDownwash angle, $\\epsilon$, deg\n\nM = 0.98\nM = 1.00\nM = 1.03\nM = 1.05\n\n$\\alpha$, deg -2, -1, 0, 1, 2, 3, 4, 6, 8, 10\n$\\square$ $\\diamond$ $\\circ$ $\\triangle$ $\\blacktriangle$ $\\blacktriangledown$ $\\square$ $\\diamond$ $\\nabla$ $\\blacktriangledown$\n\nDownwash angle, $\\epsilon$, deg\n\nM = 1.08\nM = 1.10\nM = 1.15\nM = 1.18\n\nTail-height, $h_t$, percent semispan\nCONFIDENTIAL\n\nFigure 10.— Concluded.\n\nNACA RM L9108\n26", "timestamp": "2026-07-22T04:26:07.069335+00:00"}
{"citation_id": "19930085838", "source_url": "https://ntrs.nasa.gov/api/citations/19930085838/downloads/19930085838.pdf", "page_number": 78, "total_pages": 118, "image_filename": "19930085838_p78.jpg", "text": "```markdown\n76\n\nAileron section hinge-moment coefficient, $c_{h_a}$\n\n$\\delta_f = 0^\\circ$\n\n$\\delta_a$ (deg)\n-15\n-10\n-5\n0\n5\n10\n15\n20\n25\n\n<!-- Image (119, 112, 863, 768) -->\n\nSection angle of attack, $\\alpha_o$, deg\n\n(a) $\\delta_f = 0^\\circ$.\n\nFigure 10.- Hinge-moment characteristics of a straight-sided Frise aileron on the approximately 15.4-percent-chord thick NACA 7-series-type airfoil with double slotted flap and flip. Aileron balance, $0.408c_a$; $R = 6 \\times 10^6$ (approx.)\n\nNACA RM No. L9B23\n```", "timestamp": "2026-07-22T04:26:07.268941+00:00"}
{"citation_id": "19930090382", "source_url": "https://ntrs.nasa.gov/api/citations/19930090382/downloads/19930090382.pdf", "page_number": 7, "total_pages": 37, "image_filename": "19930090382_p7.jpg", "text": "6\nCONFIDENTIAL\nNACA RM L9I07\n\nEffect of advance ratio and forward Mach number on maximum efficiency.- The variation of maximum efficiency with advance ratio for the forward Mach numbers at which the propeller was investigated is shown in figure 7. In general, the results are similar to those presented in reference 1 for the NACA 4-(5)(08)-03 propeller except that the forward Mach number defining the transition from operation at high advance ratio to operation at low advance ratio for highest efficiencies is higher for the NACA 4-(4)(06)-04 propeller (M = 0.85). These results indicate that thinning the blade makes this transition occur at higher forward Mach numbers. Eventually, however, as the forward Mach number is increased the propeller will be operating as a supersonic type of propeller and the highest efficiencies will then be obtained at low values of advance ratio.\n\nComparison of envelope efficiency for NACA 4-(4)(06)-04 and 4-(5)(08)-03 propellers.- The envelope of the maximum-efficiency curves in figure 6 is presented in figure 8. A similar envelope curve for the NACA 4-(5)(08)-03 propeller from results of reference 1 is also presented in figure 8. At low forward Mach numbers there is little difference in envelope efficiency. At supercritical Mach numbers the rate of efficiency loss is less for the thin propeller, and at a forward Mach number of 0.80 the efficiency of the thin propeller is about 10 percent higher than for the NACA 4-(5)(08)-03 propeller. These differences in efficiency are caused by differences in thickness ratio and design pitch. The effect due to differences in design lift coefficient is believed to be secondary, as is shown by the comparison of results (reference 2) for the NACA 4-(3)(08)-03 and 4-(5)(08)-03 propellers.\n\nCONCLUSIONS\n\nInvestigations of an NACA 4-(4)(06)-04 two-blade propeller in the Langley 8-foot high-speed tunnel for blade angles of $20^\\circ$ to $70^\\circ$ and through a forward Mach number range extending up to 0.925 and comparison of results with those for the NACA 4-(5)(08)-03 propeller (NACA RM L9G06a) indicate the following conclusions:\n\n1. In general, the effects of compressibility on maximum efficiency are similar to those for the NACA 4-(5)(08)-03 propeller.\n\n2. At a forward Mach number of 0.8 the envelope efficiency for the thin-blade propeller is 10 percent higher than that for the NACA\n\nCONFIDENTIAL", "timestamp": "2026-07-22T04:26:07.476036+00:00"}
{"citation_id": "19930086078", "source_url": "https://ntrs.nasa.gov/api/citations/19930086078/downloads/19930086078.pdf", "page_number": 17, "total_pages": 42, "image_filename": "19930086078_p17.jpg", "text": "CONFIDENTIAL\nTABLE I\nGEOMETRY OF THE EXTENSIBLE WING-TIP AILERONS\n\n| Aileron plan form | Aileron Extension | | | | Aileron chord | Aileron span $\\frac{b_a}{\\text{Wing semispan, } b/2}$ | | Aileron area, $S_a$ (sq ft) | | Aileron area, $\\frac{S_a}{S}$ Wing area | |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| | Nominal extension | | Actual extension, $b_a$ (ft) | | | $\\Lambda = 0^\\circ$ | $\\Lambda = 45^\\circ$ | $\\Lambda = 0^\\circ$ | $\\Lambda = 45^\\circ$ | $\\Lambda = 0^\\circ$ | $\\Lambda = 45^\\circ$ |\n| | $\\Lambda = 0^\\circ$ | $\\Lambda = 45^\\circ$ | $\\Lambda = 0^\\circ$ | $\\Lambda = 45^\\circ$ | | | | | | | |\n| Large chord | Full | Full | 0.491 | 0.351 | 0.625$^a$ | 0.127 | 0.126 | 0.766 | 0.775 | 0.040 | 0.040 |\n| | 3/4 | 7/10 | .367 | .245 | | .095 | .083 | .573 | .542 | .030 | .028 |\n| | 1/2 | 1/2 | .247 | .178 | | .064 | .064 | .386 | .393 | .020 | .020 |\n| | 1/4 | 1/4 | .121 | .093 | | .031 | .033 | .189 | .206 | .010 | .011 |\n| Triangular | Full | Full | .977 | .695 | .625$^a$ | .252 | .250 | .764 | .767 | .040 | .040 |\n| | 3/4 | 3/4 | .737 | .520 | | .190 | .187 | .576 | .574 | .030 | .030 |\n| | 1/2 | 1/2 | .491 | .348 | | .127 | .125 | .384 | .384 | .020 | .020 |\n| | 1/4 | 1/4 | .244 | .174 | | .063 | .063 | .191 | .192 | .010 | .010 |\n| Short chord | Full | Full | 1.035 | .732 | .156$^a$ | .267 | .263 | .402 | .402 | .021 | .021 |\n| | 3/4 | 3/4 | .776 | .548 | | .200 | .197 | .302 | .302 | .016 | .016 |\n| | 1/2 | 1/2 | .515 | .365 | | .133 | .131 | .200 | .201 | .011 | .011 |\n| | 1/4 | 1/4 | .259 | .186 | | .067 | .067 | .101 | .102 | .005 | .005 |\n\n$^a$At root chord of aileron.\n\nCONFIDENTIAL\nNACA\nNACA RM L9E04\n15", "timestamp": "2026-07-22T04:26:09.213746+00:00"}
{"citation_id": "19930085958", "source_url": "https://ntrs.nasa.gov/api/citations/19930085958/downloads/19930085958.pdf", "page_number": 40, "total_pages": 43, "image_filename": "19930085958_p40.jpg", "text": "```markdown\nNACA RM No. L9B11\n\n1.2\n$q/q$ .8\n.4\nTail height\n(percent $b/2$)\n46.6\n33.9\n21.1\n-1.1\n\n24\n$\\epsilon$, deg\n16\n8\n0\n\n.08\n.04\n0\n$C_m$ -.04\n-.08\n-.12\n-.16\n-.20\n-.24\n-4 0 4 8 12 16 20 24\n$\\alpha$, deg\n\n1.4\n1.2\n1.0\n$C_L$ .8\n.6\n.4\n.2\n0\n-.2\n-4 0 4 8 12 16 20 24\n$\\alpha$, deg\nTail height\n(percent $b/2$)\n$i_t$\noff\n$\\circ$ 46.6 -1.2\n$\\diamond$ 33.9 -1.0\n$\\triangle$ 21.1 -1.3\n$\\nabla$ -1.1 -1.3\n[Figure: NACA logo]\n\nFigure 20.- Characteristics of a 42° sweptback wing-fuselage combination with horizontal tail.\n0.60$\\frac{b}{2}$ drooped-nose flaps; split flaps; upper-surface fences; low wing.\n\n39\n```", "timestamp": "2026-07-22T04:26:10.138718+00:00"}
{"citation_id": "19930086022", "source_url": "https://ntrs.nasa.gov/api/citations/19930086022/downloads/19930086022.pdf", "page_number": 30, "total_pages": 34, "image_filename": "19930086022_p30.jpg", "text": "28\nNACA RM L9E24\n\n<!-- Image (235, 109, 660, 884) -->\n\nFigure 9.- The effects of high-lift and stall-control devices on the aileron hinge-moment and effectiveness parameters $P_{R\\alpha}$, $C_{h\\alpha}$, $P_{R\\delta}$, $C_{h\\delta}$, and $C_{l\\delta}$.", "timestamp": "2026-07-22T04:26:12.026386+00:00"}
{"citation_id": "19930085982", "source_url": "https://ntrs.nasa.gov/api/citations/19930085982/downloads/19930085982.pdf", "page_number": 31, "total_pages": 32, "image_filename": "19930085982_p31.jpg", "text": "NACA-Langley - 7-18-49 - 275\n\nNACA RM E9E13\n\n| | |\n| :--- | :--- |\n| $\\downarrow$ Flow giving maximum efficiency | |\n| $+$ Design point | |\n\nEquivalent turning angle, $\\Delta\\beta_e$, deg\n\nEquivalent turning angle, $\\Delta\\beta_e$, deg\n\nEquivalent turning angle, $\\Delta\\beta_e$, deg\n\nEquivalent turning angle, $\\Delta\\beta_e$, deg\n\nEquivalent angle of attack, $\\alpha_e$, deg\n\n(a) Radial station a near tip, $\\sigma = 0.605$.\n\n(c) Radial station c, $\\sigma = 0.822$.\n\n(b) Radial station b, $\\sigma = 0.696$.\n\n(d) Radial station d near hub, $\\sigma = 1.002$.\n\nFigure 9. - Equivalent turning angles of rotor blades at three-fourths design speed. NACA 65-(12)10 blade section.\n\n29", "timestamp": "2026-07-22T04:26:21.659708+00:00"}
{"citation_id": "19930085930", "source_url": "https://ntrs.nasa.gov/api/citations/19930085930/downloads/19930085930.pdf", "page_number": 65, "total_pages": 92, "image_filename": "19930085930_p65.jpg", "text": "64\nNACA RM L9G07\n\n[Figure: A graph plotting Average Mach number ratio ($\\frac{M_{2av}}{M_1}$) against Area ratio ($\\frac{A_2}{A_1}$). The graph contains three data series represented by circles, squares, and diamonds. There are \"CONFIDENTIAL\" stamps at the top and bottom of the graph area, and a NACA logo at the bottom right.]\n\n$$ \\frac{M_{2av}}{M_1} $$\n\n$$ \\circ \\quad \\frac{p_1}{p_A} = 1.28 $$\n$$ \\square \\quad \\frac{p_1}{p_A} = 1.00 $$\n$$ \\diamond \\quad \\frac{p_1}{p_A} = .85 $$\n\n$$ \\frac{p_2}{p_A} = 1.00 $$\n(Experimental)\n\n$$ \\text{Area ratio, } \\frac{A_2}{A_1} $$\n\nFigure 27.- The variation of the average Mach number ratio with area ratio at three static-pressure ratios for model 2.", "timestamp": "2026-07-22T04:26:24.292101+00:00"}
{"citation_id": "19930085551", "source_url": "https://ntrs.nasa.gov/api/citations/19930085551/downloads/19930085551.pdf", "page_number": 80, "total_pages": 82, "image_filename": "19930085551_p80.jpg", "text": "NACA RM No. L8K30\n79\n\n<!-- Image (249, 129, 757, 880) -->\n\n(b) Flaps and gear down; power for level flight.\nFigure 23.- Concluded.", "timestamp": "2026-07-22T04:26:24.718740+00:00"}
{"citation_id": "19930082090", "source_url": "https://ntrs.nasa.gov/api/citations/19930082090/downloads/19930082090.pdf", "page_number": 115, "total_pages": 122, "image_filename": "19930082090_p115.jpg", "text": "NACA TN No. 1455\n113\n\nVentilating air\n50.5\"\n2.25\"\n15.7\"\nExhaust\ngas\n6.4\"\nUC-1 air shroud\n\n5.1\"\n8.4\"\n9\" I.D.\n3\"\n6.4\"\nExhaust gas\n4\"\n15.7\"\n7\"\nVentilating air\nA-1 air shroud\n\nA\n25\"\n0.75\"\n15.6\"\n0.75\"\n52 rows of fins\n(69 fins per row)\nSection A-A\n0.75\"\n80 rows of fins\n(19 fins per row)\nSlotted-fin heat\nexchanger\nNACA\n\nFigure 60. - Schematic diagram of slotted-fin heat exchanger P and air\nshrouds UC-1 and A-1. Weight of heat exchanger, 32.5 pounds.\n\n| | Air side | | Gas side |\n| :--- | :--- | :--- | :--- |\n| | UC-1 | A-1 | |\n| Cross-sectional area, sq ft | 0.203 | 0.271 | 0.203 |", "timestamp": "2026-07-22T04:26:32.963733+00:00"}
{"citation_id": "19930085842", "source_url": "https://ntrs.nasa.gov/api/citations/19930085842/downloads/19930085842.pdf", "page_number": 63, "total_pages": 104, "image_filename": "19930085842_p63.jpg", "text": "```markdown\nNACA RM L50C29\n\nDrag coefficient, $C_D$\n\n$\\alpha = 23.2^\\circ$\n$\\delta_{a_{TR}}, deg$\n-30\n-20\n-10\n10\n20\n\n$\\alpha = 11.3^\\circ$\n$\\delta_{a_{TR}}, deg$\n-30\n-20\n-10\n10\n20\n\n$\\alpha = -0.6^\\circ$\n$\\delta_{a_{TR}}, deg$\n-30\n-20\n-10\n10\n20\n\nAllavator deflection, $\\delta_a$, deg\n\nNATIONAL ADVISORY\nCOMMITTEE FOR AERONAUTICS\n\nFigure 30.- The effect of right-allavator tab setting on the variation of $C_D$ with $\\delta_a$. Model in basic configuration; $\\delta_{a_{TL}} = 0^\\circ$; $\\delta_r = 0^\\circ$; propellers removed; $\\delta_F = 0^\\circ$.\n\n59\n```", "timestamp": "2026-07-22T04:26:57.937457+00:00"}
{"citation_id": "19930085936", "source_url": "https://ntrs.nasa.gov/api/citations/19930085936/downloads/19930085936.pdf", "page_number": 52, "total_pages": 59, "image_filename": "19930085936_p52.jpg", "text": "NACA RM No. E9B03\n51\n\nAngle of\nyaw\n(deg)\n-12\n-6\n0\n\nPressure\ncoefficient\n$C_p$\n.2\n0\n-.2\n-.2\n0\n.2\n\n[Figure: Polar plot showing radial pressure distributions for three yaw angles (-12, -6, 0 degrees) at x/L = 0.148. The plot includes concentric circles and curves representing pressure coefficient values.]\n\nNACA\n\n(a) x/L = 0.148.\n\nFigure 11. - Radial pressure distributions at $5^\\circ$ angle of attack for three yaw angles.", "timestamp": "2026-07-22T04:26:59.704997+00:00"}
{"citation_id": "19930085952", "source_url": "https://ntrs.nasa.gov/api/citations/19930085952/downloads/19930085952.pdf", "page_number": 49, "total_pages": 50, "image_filename": "19930085952_p49.jpg", "text": "48\nNACA RM L9C24\n\n[Figure: Graph with three plots]\n\n(c) $\\alpha \\approx 58^\\circ$; $\\frac{Y}{nD} = 0.30$.\nFigure 22.- Continued.", "timestamp": "2026-07-22T04:27:00.209015+00:00"}
{"citation_id": "19930086081", "source_url": "https://ntrs.nasa.gov/api/citations/19930086081/downloads/19930086081.pdf", "page_number": 15, "total_pages": 44, "image_filename": "19930086081_p15.jpg", "text": "NACA RM L9H05 CONFIDENTIAL 13\n\n4. Increasing the leading-edge bluntness and airfoil thickness of the control surface decreased the rolling effectiveness about 15 percent and caused no change in the hinge-moment characteristics.\n\nLangley Aeronautical Laboratory\nNational Advisory Committee for Aeronautics\nLangley Air Force Base, Va.\n\nREFERENCES\n\n1. Sandahl, Carl A.: Free-Flight Investigation of the Rolling Effectiveness of Several Delta Wing-Aileron Configurations at Transonic and Supersonic Speeds. NACA RM L8D16, 1948.\n\n2. Conner, D. William: Aerodynamic Characteristics of Two All-Movable Wings Tested in the Presence of a Fuselage at a Mach Number of 1.9. NACA RM L8H04, 1948.\n\n3. Brown, Clinton E., and Heinke, Harry S., Jr.: Preliminary Wind-Tunnel Tests of Triangular and Rectangular Wings in Steady Roll at Mach Numbers of 1.62 and 1.92. NACA RM L8L30, 1948.\n\n4. Brown, Clinton E., and Adams, Mac C.: Damping in Pitch and Roll of Triangular Wings at Supersonic Speeds. NACA Rep. 892, 1948.\n\n5. Lagerstrom, P. A., and Graham, Martha E.: Linearized Theory of Supersonic Control Surfaces. Jour. Aero. Sci., vol. 16, no. 1, Jan. 1949, pp. 31-34.\n\n6. Brown, Clinton E.: Theoretical Lift and Drag of Thin Triangular Wings at Supersonic Speeds. NACA Rep. 839, 1946.\n\n7. Beskin, L.: Determination of Upwash around a Body of Revolution at Supersonic Velocities. CVAC-DEVF Memo BB-6, APL/JHU-CM-251, The Johns Hopkins Univ., Appl. Phys. Lab., May 27, 1946.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T04:27:00.713321+00:00"}
{"citation_id": "19930086151", "source_url": "https://ntrs.nasa.gov/api/citations/19930086151/downloads/19930086151.pdf", "page_number": 9, "total_pages": 34, "image_filename": "19930086151_p9.jpg", "text": "NACA RM L9J28 CONFIDENTIAL 7\n\nThe pitching-moment data obtained on the wing with and without the end plate were about the same, except that the wing with the end plate was slightly less stable than the plain wing.\n\nLateral Control Characteristics\n\nThe rolling-moment and yawing-moment data obtained through the angle-of-attack range from tests of the $45^\\circ$ sweptback wing at positive deflections of the wing-tip ailerons are presented in figures 5 to 8. In order to show the variation of rolling-moment coefficient with aileron deflection, the rolling-moment data of figures 5 to 8 were cross-plotted against aileron deflection as shown in figures 9 and 10. Inasmuch as all wing-aileron configurations investigated were symmetrical and had symmetrical profiles (although the end plate was asymmetrically placed on the wing), the rolling-moment data obtained at positive aileron deflections and negative angles of attack (figs. 5 to 8) were cross-plotted with opposite signs in figures 9 and 10 to provide data at negative aileron deflections and positive angles of attack.\n\nEffect of aileron plan form.— A comparison of the data obtained with the triangular and parallelogram wing-tip ailerons reveals an inconsistent effect of aileron plan form on the rolling moments over the angle-of-attack range (figs. 5 to 8). The rolling-moment data presented in figures 5 to 10 also show that a serious reduction of rolling moment occurred for positive aileron deflections at the higher positive angles of attack, and in some cases, the aileron effectiveness reversed. This loss in effectiveness and the aileron reversal probably result from the stalling of the aileron at large deflections and wing angles of attack. Because wing stall angle generally increases with angle of sweepback, particularly with sharp leading edges, the triangular-plan-form aileron exhibited less tendency toward aileron reversal than the parallelogram-plan-form aileron. Similar effects of a large reduction and reversal of aileron effectiveness at large positive values of $\\alpha$ and $\\delta_a$ were not exhibited by the data of references 1 to 4 because the ailerons of the reference investigations were \"free floating\" — which enabled them to assume low incidences in the neutral condition — and also had conventional airfoil profiles, so that the ailerons did not stall when deflected to moderate deflections.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T04:27:01.424234+00:00"}
{"citation_id": "19930090382", "source_url": "https://ntrs.nasa.gov/api/citations/19930090382/downloads/19930090382.pdf", "page_number": 8, "total_pages": 37, "image_filename": "19930090382_p8.jpg", "text": "**Page intentionally left blank**\n\nmissing\n8, 10, 12, 14\n\n**Page intentionally left blank**", "timestamp": "2026-07-22T04:27:02.655156+00:00"}
{"citation_id": "19930086003", "source_url": "https://ntrs.nasa.gov/api/citations/19930086003/downloads/19930086003.pdf", "page_number": 29, "total_pages": 36, "image_filename": "19930086003_p29.jpg", "text": "CONFIDENTIAL\n\nM = 0.60\nM = 0.70\nM = 0.80\n\nDownwash angle, $\\epsilon$, deg\n\n$\\alpha$, deg -2, -1, 0, 1, 2, 3, 4, 6, 8, 10\n\n□ ◇ ○ △ ▲ ▽ □ ◇ ▽ ▽\n\nM = 0.85\nM = 0.90\nM = 0.93\n\nDownwash angle, $\\epsilon$, deg\n\nTail-height, $h_t$, percent semispan\n\n-80 -40 0 40 80\n\nCONFIDENTIAL\n\nFigure 11.— Effective downwash angles in region of tail plane for a model with 45° sweptback wing, aspect ratio 6, taper ratio 0.6, and NACA 65A006 airfoil section. Wing fuselage.\n\nNACA RM L59L08\n\n27", "timestamp": "2026-07-22T04:27:05.113551+00:00"}
{"citation_id": "19930085838", "source_url": "https://ntrs.nasa.gov/api/citations/19930085838/downloads/19930085838.pdf", "page_number": 79, "total_pages": 118, "image_filename": "19930085838_p79.jpg", "text": "NACA RM No. L9B23\n\nAileron section hinge-moment coefficient, $C_{h_a}$\n\nSection angle of attack, $\\alpha_o$, deg\n\n$\\delta_t = 0^\\circ$\n\n$\\delta_a$ (deg)\n-20\n-15\n-10\n-5\n0\n5\n10\n15\n\n(b) $\\delta_r = 25^\\circ$.\n\nFigure 10.- Continued.\n\n77", "timestamp": "2026-07-22T04:27:06.384203+00:00"}
{"citation_id": "19930086078", "source_url": "https://ntrs.nasa.gov/api/citations/19930086078/downloads/19930086078.pdf", "page_number": 18, "total_pages": 42, "image_filename": "19930086078_p18.jpg", "text": "CONFIDENTIAL\n\nX-axis\n0.50 chord\nline; aileron\npivoting axis\n\nY-axis\n\n7.50\n\n45.00\n$b/2 = 46.50$\n\n9.38\n18.75\n\n30.00\n\n12°\n\nNACA\n64A010 airfoil\n\nCONFIDENTIAL\n\n(a) Large-chord extensible aileron on wing.\n\n(b) Triangular extensible aileron on wing.\n\n(c) Short-chord extensible aileron on wing.\n\nFigure 1.- Schematic drawing of the unswept configuration of the untapered semispan-wing model and the extensible wing-tip ailerons. Wing area = 19.16 square feet; aspect ratio = 3.13. (All dimensions are in inches unless otherwise noted.)\n\n16\n\nNACA RM L9E04", "timestamp": "2026-07-22T04:27:06.756068+00:00"}
{"citation_id": "19930086022", "source_url": "https://ntrs.nasa.gov/api/citations/19930086022/downloads/19930086022.pdf", "page_number": 31, "total_pages": 34, "image_filename": "19930086022_p31.jpg", "text": "NACA RM L9E24\n29\n\n$$C_{h_a}$$\n\n$$\\alpha$$\n(deg)\n0\n8\n12\n16\n\nPlain wing\n\nLeading and trailing-edge flaps\n\nLeading and trailing-edge flaps and fences\n\n$$\\delta_a$$\n-25 -20 -15 -10 -5 0 5 10 15 20 25\n\n[Figure: NACA logo]\n\nFigure 10.— The variation of aileron hinge-moment characteristics with aileron deflection for various model configurations.", "timestamp": "2026-07-22T04:27:08.634948+00:00"}
{"citation_id": "19930085551", "source_url": "https://ntrs.nasa.gov/api/citations/19930085551/downloads/19930085551.pdf", "page_number": 81, "total_pages": 82, "image_filename": "19930085551_p81.jpg", "text": "80\nNACA RM No. L8K30\n\n<!-- Image (83, 297, 882, 675) -->\n\n(a) Clean condition; power for level flight.\nFigure 24.- Variation of maximum $pb/2V$ available without exceeding 80 pounds of wheel force with indicated airspeed. C-54D airplane.", "timestamp": "2026-07-22T04:27:11.403771+00:00"}
{"citation_id": "19930082090", "source_url": "https://ntrs.nasa.gov/api/citations/19930082090/downloads/19930082090.pdf", "page_number": 116, "total_pages": 122, "image_filename": "19930082090_p116.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T04:27:12.821609+00:00"}
{"citation_id": "19930085958", "source_url": "https://ntrs.nasa.gov/api/citations/19930085958/downloads/19930085958.pdf", "page_number": 41, "total_pages": 43, "image_filename": "19930085958_p41.jpg", "text": "1.2\n.8\n$q_t/q$\n.4\n.08\n.04\n0\n-.04\n$C_m$\n-.08\n-.12\n-.16\n-.20\n-.24\n-.28\n-.32\n-4 0 4 8 12 16 20 24\n$\\alpha$, deg\n\nTail height\n(percent $\\frac{b}{2}$)\n46.6\n33.9\n21.1\n-1.1\n\n24\n16\n$i$, deg\n8\n0\n\n1.4\n1.2\n1.0\n.8\n$C_L$\n.6\n.4\n.2\n0\n-.2\n-4 0 4 8 12 16 20 24\n$\\alpha$, deg\n\nTail height\n(percent $\\frac{b}{2}$) $i_t$\noff\n$\\circ$ 46.6 -1.2\n$\\square$ 33.9 -1.2\n$\\diamond$ 21.1 -1.2\n$\\nabla$ -1.1 -2.1\n\nNACA\n\nFigure 21.- Characteristics of a 42° sweptback wing-fuselage combination with a horizontal tail.\n0.55$\\frac{b}{2}$ extensible leading-edge flaps; split flaps; upper-surface fences; low wing.\n\nNACA RM NO. L9B11\n40", "timestamp": "2026-07-22T04:27:13.022227+00:00"}
{"citation_id": "19930085930", "source_url": "https://ntrs.nasa.gov/api/citations/19930085930/downloads/19930085930.pdf", "page_number": 66, "total_pages": 92, "image_filename": "19930085930_p66.jpg", "text": "NACA RM L9G07\n65\n\nCONFIDENTIAL\n\nAverage stagnation-pressure recovery, $\\frac{P_{2av}}{P_0}$\n\n$\\frac{p_2}{p_A} = 1.00$\n(Experimental)\n\n$\\circ \\frac{p_1}{p_A} = 1.28$\n$\\square \\frac{p_1}{p_A} = 1.00$\n$\\diamond \\frac{p_1}{p_A} = .85$\n\nArea ratio, $\\frac{A_2}{A_1}$\n\nFigure 28.- The variation of the average stagnation-pressure recovery with area ratio at three static-pressure ratios for model 2.", "timestamp": "2026-07-22T04:27:13.286353+00:00"}
{"citation_id": "19930085965", "source_url": "https://ntrs.nasa.gov/api/citations/19930085965/downloads/19930085965.pdf", "page_number": 35, "total_pages": 67, "image_filename": "19930085965_p35.jpg", "text": "34\nNACA RM E9E06\n\nh unit thermal convective conductance between vane surface and air, Btu/(hr)(sq ft)(°F)\n\nJ mechanical equivalent of heat, ft-lb/Btu\n\nL latent heat of vaporization of water, Btu/lb\n\nl length of blade, ft\n\nM water impingement per unit area, lb/(hr)(sq ft)\n\nMₑ water impingement per unit area leading edge, lb/(hr)(sq ft)\n\nMₛ water impingement per unit area blade surface, lb/(hr)(sq ft)\n\nm liquid-water content, lb/cu ft\n\nN number of blades\n\nPr Prandtl number\n\np static pressure, lb/sq in. absolute\n\nR gas constant, for air 53.3 ft-lb/(lb)(°R)\n\ns distance along chord from leading edge of blade, ft\n\nTₐᵥ average of inlet-air temperature and blade-surface temperature, °R\n\nt air temperature, °F\n\ntₛ surface temperature with no external heat conditions, °F\n\ntₛ d surface design temperature, °F\n\nV velocity, ft/sec\n\nW air flow, lb/sec\n\nx average distance between adjacent blades, ft\n\nα angle of attack of blade section, degrees\n\nγ ratio of specific heats, for air = 1.4\n\nρ density, slugs/cu ft", "timestamp": "2026-07-22T04:27:33.706034+00:00"}
{"citation_id": "19930086105", "source_url": "https://ntrs.nasa.gov/api/citations/19930086105/downloads/19930086105.pdf", "page_number": 10, "total_pages": 22, "image_filename": "19930086105_p10.jpg", "text": "8\nCONFIDENTIAL\nNACA RM E5H12\n\nREFERENCES\n\n1. Kantrowitz, Arthur, and Donaldson, Coleman duP.: Preliminary Investigation of Supersonic Diffusers. NACA ACR L5D20, 1945.\n\n2. Oswatitsch, and Böhm: Air Forces and Flow Phenomena on Self Propelled Missiles. Rep. No. 1010/2, Kaiser Wilhelm Inst. Aerod. Res., Oct. 1944. (Trans. by Curtiss-Wright Corp., Oct. 1945.)\n\n3. Ferri, Antonio, and Nucci, Louis M.: Preliminary Investigation of a New Type of Supersonic Inlet. NACA RM L6J31, 1946.\n\n4. Evvard, John C., and Blakey, John W.: The Use of Perforated Inlets for Efficient Supersonic Diffusion. NACA RM E7C26, 1947.\n\n5. Connors, J. F., and Schroeder, A. H.: Preliminary Investigation of Effects of Combustion in Ram Jet on Performance of Supersonic Diffusers. I - Shock Diffuser with Triple-Shock Projecting Cone. NACA RM E8F15, 1948.\n\n6. Schroeder, Albert H., and Connors, James F.: Preliminary Investigation of Effects of Combustion in Ram Jet on Performance of Supersonic Diffusers. II - Perforated Supersonic Inlet. NACA RM E8G16, 1948.\n\n7. Schroeder, Albert H., and Connors, James F.: Preliminary Investigation of Effects of Combustion in Ram Jet on Performance of Supersonic Diffusers. III - Normal-Shock Diffuser. NACA RM E8J18, 1948.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T04:27:36.853101+00:00"}
{"citation_id": "19930085936", "source_url": "https://ntrs.nasa.gov/api/citations/19930085936/downloads/19930085936.pdf", "page_number": 53, "total_pages": 59, "image_filename": "19930085936_p53.jpg", "text": "52\nNACA RM No. E9B03\n\nPressure\ncoefficient\n$C_p$\n.4\n.2\n0\n-.2\n-.2\n0\n.2\n.4\n\nAngle of\nyaw\n(deg)\n-12\n-6\n0\n\n[Figure: Polar plot showing radial pressure distributions for three yaw angles]\n\n(b) x/L = 0.898.\nFigure 11. - Concluded. Radial pressure distributions at 5° angle of\nattack for three yaw angles.", "timestamp": "2026-07-22T04:27:37.234195+00:00"}
{"citation_id": "19930086015", "source_url": "https://ntrs.nasa.gov/api/citations/19930086015/downloads/19930086015.pdf", "page_number": 31, "total_pages": 54, "image_filename": "19930086015_p31.jpg", "text": "CONFIDENTIAL\n\nVertical distance from tunnel center line, z, in.\n\n|  |  |  |\n| :--- | :--- | :--- |\n| 24 | | |\n| 16 | | |\n| 8 | | |\n| 0 | | |\n| -8 | | |\n| -16 | | |\n| -24 | | |\n| | -.04 | 0 | .04 |\n\nx=-24\n\nFlagged symbols denote survey off center line\n\n|  |  |  |  |\n| :--- | :--- | :--- | :--- |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | 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|\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| |", "timestamp": "2026-07-22T04:27:37.527185+00:00"}
{"citation_id": "19930086081", "source_url": "https://ntrs.nasa.gov/api/citations/19930086081/downloads/19930086081.pdf", "page_number": 16, "total_pages": 44, "image_filename": "19930086081_p16.jpg", "text": "14\nNACA RM L9H05\n\nCONFIDENTIAL\n5.566\nFuselage &\n($C_2$) axis\n$(+C_n)$\n$(+C_l)$\n4.605\nRoot chord axis of\ncontrol surface ($C_{BM_f}$)\n1.755\nY-axis of wing\nControl-surface\nhinge axis\nControl surface\nchord plane ($C_{C_f}$)\n$(+C_L)$\n$(+C_{N_f})$\n$+\\delta$\n$+\\alpha$\nWind axis\n($C_D$)\n$(+C_m)$\nCONFIDENTIAL\n$(+C_{M_f})$\nNACA\nFigure 1.- Relation between the various reference axes and reference\nplanes used in presenting test data for control surface. All\ndimensions in inches.", "timestamp": "2026-07-22T04:27:38.157327+00:00"}
{"citation_id": "19930085982", "source_url": "https://ntrs.nasa.gov/api/citations/19930085982/downloads/19930085982.pdf", "page_number": 32, "total_pages": 32, "image_filename": "19930085982_p32.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T04:27:39.775251+00:00"}
{"citation_id": "19930085838", "source_url": "https://ntrs.nasa.gov/api/citations/19930085838/downloads/19930085838.pdf", "page_number": 80, "total_pages": 118, "image_filename": "19930085838_p80.jpg", "text": "78\nNACA RM No. L9B23\n\n<!-- Image (117, 109, 842, 874) -->\n\n(a) $\\delta_f = 25^\\circ$.\nFigure 10.- Continued.", "timestamp": "2026-07-22T04:27:42.544187+00:00"}
{"citation_id": "19930085952", "source_url": "https://ntrs.nasa.gov/api/citations/19930085952/downloads/19930085952.pdf", "page_number": 50, "total_pages": 50, "image_filename": "19930085952_p50.jpg", "text": "NACA RM L9C24\n49\n\n[Figure: A graph with three subplots. The x-axis is labeled $\\delta_f, deg$ and ranges from -16 to 24. The top subplot y-axis ranges from 0 to -0.5 and shows curves for $C_{h_a}$ and $C_{h_f}$. The middle subplot y-axis ranges from 9.2 to 9.6 and shows a curve for $C_L$. The bottom subplot y-axis ranges from 0 to -0.3 and shows a curve for $C_m$. A NACA logo is present in the bottom right corner of the graph.]\n\n(d) $\\alpha \\approx 69^\\circ$; $\\frac{V}{nD} = 0.20$.\nFigure 22.— Concluded.\n\nNACA-Langley - 11-7-49 - 275", "timestamp": "2026-07-22T04:27:43.784597+00:00"}
{"citation_id": "19930085551", "source_url": "https://ntrs.nasa.gov/api/citations/19930085551/downloads/19930085551.pdf", "page_number": 82, "total_pages": 82, "image_filename": "19930085551_p82.jpg", "text": "NACA RM No. L8K30\n81\n\n[Figure: Graph plotting Max. $Pb/2V$ against Indicated airspeed, mph. The y-axis ranges from 0 to .08. The x-axis ranges from 100 to 140. The legend indicates circles for Right roll and squares for Left roll. The NACA logo is present in the bottom right corner of the graph.]\n\n(b) Flaps and gear full down; power for level flight.\nFigure 24.— Concluded.", "timestamp": "2026-07-22T04:27:46.932888+00:00"}
{"citation_id": "19930086151", "source_url": "https://ntrs.nasa.gov/api/citations/19930086151/downloads/19930086151.pdf", "page_number": 10, "total_pages": 34, "image_filename": "19930086151_p10.jpg", "text": "8\nCONFIDENTIAL\nNACA RM L9J28\n\nA comparison of the values of the slope of rolling-moment coefficient against aileron deflection $C_{l\\delta_a}$ at $\\alpha = 0^\\circ$ for the four wing-aileron configurations is shown in the following table:\n\n| Aileron plan form | $C_{l\\delta_a}$ | |\n| :--- | :---: | :---: |\n| | Plain wing | Wing with end plate |\n| Triangular | 0.00072 | 0.00061 |\n| Parallelogram | .00072 | .00047 |\n\nAlthough the values of $C_{l\\delta_a}$ for the two aileron plan forms on the wing with end plate differed appreciably at $\\alpha = 0^\\circ$, aileron plan form generally had little effect on the rolling moments of either the plain wing or the wing with end plate over most of the angle-of-attack range. In addition, all aileron configurations exhibited larger values of $C_{l\\delta_a}$ at $\\alpha = 5^\\circ$ and $10^\\circ$ than at $\\alpha = 0^\\circ$ (figs. 9 and 10).\n\nThe yawing-moment data shown in figures 5 to 8 exhibit little or no consistent effect of aileron plan form. Although the yawing-moment data have not been cross-plotted against aileron deflection (as were the rolling-moment data), the values of $C_n$ for positive angles of attack and negative aileron deflections would retain the same signs and values as shown in figures 5 to 8 for negative values of $\\alpha$ and positive values of $\\delta_a$. Analysis of these data in conjunction with the rolling-moment data of figures 9 and 10 shows that the yawing moments were generally adverse and became more adverse with increase in angles of attack and aileron deflection. At the higher angles of attack, the adverse $C_n/C_l$ ratio amounted to as much as 1.5 for all aileron configurations.\n\nEffect of end plate.— The data obtained on the wing with end plate (figs. 7 and 8) generally showed a decrease in the rolling moments obtained through most of the angle-of-attack range and over the aileron-deflection range, compared with the rolling moments produced on the plain wing (figs. 5 and 6). This effect probably results from the fact that the end plate reduces any \"carry-over\" of loading from the aileron to the wing and vice versa, and causes the aileron to act essentially as an independent semispan wing in the presence of the end plate. As an independent wing, the aileron, because of its low aspect ratio and large sweep, produces small increments of lift — hence, small values of rolling moment for given deflections — and is less effective than the\n\nCONFIDENTIAL", "timestamp": "2026-07-22T04:27:52.015289+00:00"}
{"citation_id": "19930085930", "source_url": "https://ntrs.nasa.gov/api/citations/19930085930/downloads/19930085930.pdf", "page_number": 67, "total_pages": 92, "image_filename": "19930085930_p67.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T04:27:52.895102+00:00"}
{"citation_id": "19930085842", "source_url": "https://ntrs.nasa.gov/api/citations/19930085842/downloads/19930085842.pdf", "page_number": 64, "total_pages": 104, "image_filename": "19930085842_p64.jpg", "text": "69\n\n| $\\alpha$, deg | $\\delta_a$, deg | Symbol |\n| :--- | :--- | :--- |\n| 35.2 | -4.8 | $\\triangle$ |\n| 29.3 | -4.8 | $\\triangle$ |\n| 23.2 | -3.7 | $\\circ$ |\n| 11.3 | -1.5 | $\\circ$ |\n| -0.6 | 0 | $\\diamond$ |\n\n$\\left. \\begin{array}{c} \\text{ } \\\\ \\text{ } \\\\ \\text{ } \\\\ \\text{ } \\\\ \\text{ } \\end{array} \\right\\} \\approx \\text{Trim}$\n\nPitching-moment coefficient, $C_m$\n\nFlap deflection, $\\delta_F$, deg\n\nNATIONAL ADVISORY\nCOMMITTEE FOR AERONAUTICS\n\nFigure 31.- Variation of pitching-moment coefficient with flap deflection. Model in basic configuration; $\\delta_{a_{TL}} = \\delta_{a_{TR}} = 0^\\circ$; $\\delta_r = 0^\\circ$; propellers removed.\n\nNACA RM L9C29", "timestamp": "2026-07-22T04:27:53.685633+00:00"}