Buckets:
| {"citation_id": "19930082485", "source_url": "https://ntrs.nasa.gov/api/citations/19930082485/downloads/19930082485.pdf", "page_number": 3, "total_pages": 62, "image_filename": "19930082485_p3.jpg", "text": "2\nNACA TN No. 1810\n\nThese methods, however, do not consider compressibility and are mathematically complex. Stream-filament theories that consider compressibility may be used to design blades of high solidity. These methods involve graphical integration of the flow equations (reference 5). Satisfactory experimental verification of either the potential-flow theories or of stream-filament theories have previously been lacking.\n\nThe design performance of turbine stator blades based on free-vortex flow as measured in a sector of an annular cascade tunnel was determined at the NACA Lewis laboratory. The blade-surface velocities were calculated by an approximation, which was developed at the NACA Lewis laboratory, of the stream-filament method of reference 5 and are compared with experimental values.\n\nSYMBOLS\n\nThe following symbols are used in this report:\n\n| | |\n| :--- | :--- |\n| C | curvature, (1/ft, unless otherwise noted) |\n| $c_p$ | specific heat at constant pressure, (ft-lb/(slug)($^\\circ$R)) |\n| $c_v$ | specific heat at constant volume, (ft-lb/(slug)($^\\circ$R)) |\n| exp | base of Napierian logarithmic system e raised to power in parentheses following exp |\n| n | distance from blade suction surface to any point on velocity-potential line, (ft, unless otherwise noted) |\n| $n_o$ | length of velocity potential line from suction surface to pressure surface, (ft, unless otherwise noted) |\n| P | pressure, (lb/sq ft absolute) |\n| r | radius, (ft, unless otherwise noted) |\n| S | blade-surface length, (ft, unless otherwise noted) |\n| T | temperature, ($^\\circ$R) |\n| V | velocity, (ft/sec) |\n| $V_{cr}$ | critical velocity, $\\sqrt{\\frac{\\gamma-1}{\\gamma+1} 2c_p T_t}$, (ft/sec) |", "timestamp": "2026-07-22T04:33:34.894617+00:00"} | |
| {"citation_id": "19930082496", "source_url": "https://ntrs.nasa.gov/api/citations/19930082496/downloads/19930082496.pdf", "page_number": 2, "total_pages": 50, "image_filename": "19930082496_p2.jpg", "text": "NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nTECHNICAL NOTE NO. 1836\n\nINITIAL INVESTIGATION OF CARBIDE-TYPE CERAMAL OF\n\n80-PERCENT TITANIUM CARBIDE PLUS 20-PERCENT\n\nCOBALT FOR USE AS GAS-TURBINE-BLADE MATERIAL\n\nBy Charles A. Hoffman, G. Mervin Ault, and James J. Gangler\n\nSUMMARY\n\nAn investigation was conducted to determine the material problems arising in the use of a carbide-type ceramal for gas-turbine blades. Specimens of an 80-percent titanium carbide plus 20-percent cobalt (by weight) ceramal were investigated for short-time tensile-strength characteristics at $1800^\\circ$ and $2200^\\circ$ F and for thermal-shock characteristics at $1800^\\circ$, $2000^\\circ$, $2200^\\circ$, and $2400^\\circ$ F. Gas-turbine blades of this material were operated in a quasi-service-evaluation unit. The conditions of operation were indicated inlet-gas temperatures between $1700^\\circ$ and $2200^\\circ$ F and turbine tip speeds between 478 and 835 feet per second. A commercial tensile machine, a specially designed furnace and air-quench-chamber apparatus, and a small gas turbine were used for these investigations. X-ray-diffraction studies were made of the material before and after operation in the form of turbine blades. For comparison purposes, zircon and titanium carbide ceramics and an alloy were investigated for thermal-shock characteristics and turbine-blade performance, respectively.\n\nThe ceramal had a short-time tensile strength of 33,200 pounds per square inch at $1800^\\circ$ F and as high as 13,200 pounds per square inch at $2200^\\circ$ F. On a strength-to-weight basis, this ceramal was, in general, superior to alloys and ceramics. The ceramal survived 25 thermal-shock cycles at a temperature of $1800^\\circ$ F, 25 cycles at $2000^\\circ$ F, 25 cycles at $2200^\\circ$ F, and 25 cycles at $2400^\\circ$ F, whereas the zircon ceramic survived 1 cycle at $1800^\\circ$ F and the titanium carbide ceramic survived through 21 cycles at $2400^\\circ$ F.\n\nThree ceramal blades survived 9 hours and 42 minutes of quasi-service evaluation, whereas six of the twelve original metal blades remained; at this time, one ceramal blade was accidentally broken. At 12 hours and 13 minutes, two of the three ceramal blades remained, whereas four metal blades remained. At this time, a second ceramal blade had been destroyed because of wheel-dovetail failure and the third ceramal blade failed as a result of service operation.", "timestamp": "2026-07-22T04:33:35.261048+00:00"} | |
| {"citation_id": "19930085838", "source_url": "https://ntrs.nasa.gov/api/citations/19930085838/downloads/19930085838.pdf", "page_number": 89, "total_pages": 118, "image_filename": "19930085838_p89.jpg", "text": "NACA RM No. L9B23\n\n87\n\nFlap section hinge-moment coefficient, $c_{h_1}$\n\nFlap section hinge-moment coefficient, $c_{h_1}$\n\nSection angle of attack, $\\alpha_o$, deg\n\n(d) $\\delta_f = 25^\\circ$.\n\nFigure 11.- Continued.", "timestamp": "2026-07-22T04:33:37.328786+00:00"} | |
| {"citation_id": "19930086151", "source_url": "https://ntrs.nasa.gov/api/citations/19930086151/downloads/19930086151.pdf", "page_number": 17, "total_pages": 34, "image_filename": "19930086151_p17.jpg", "text": "```markdown\nNACA RM L9J28\n\nCONFIDENTIAL\n\nX-axis\nY-axis\n0.50 chord line\n40.98 (M.A.C)\n45°\nAileron pivot axis\n10.25\n18.59\n37.65\n13.26\n26.52\nWing with parallelogram aileron.\n\nX-axis\nY-axis\n42.43\n0.50 chord line\n40.32 (M.A.C)\n45°\nAileron pivot axis\n10.08\n19.31\n31.82\n33.32\n41.82\nEnd plate (removable)\n13.26\n26.52\nWing with triangular aileron.\n\n48\n4 radius\nWing upper surface\n30\n20\nEnd plate\nView A-A\n\n1/2\nrounded corners\nCONFIDENTIAL\n\nAreas\nComplete wing-\nwith parallelogram aileron 21.02 sq.ft.\nwith triangular aileron 21.02 sq.ft.\nEnd plate (on both wing semispans) 20.00 sq.ft.\n\nAspect ratio\nWing:-\nwith parallelogram aileron 1.87\nwith triangular aileron 2.31\n\nNACA\n\nFigure 1.- Geometric characteristics of the 45° sweptback wing, wing-tip ailerons, and end plate. (All dimensions in inches unless otherwise noted.)\n\n15\n```", "timestamp": "2026-07-22T04:33:38.107015+00:00"} | |
| {"citation_id": "19930086015", "source_url": "https://ntrs.nasa.gov/api/citations/19930086015/downloads/19930086015.pdf", "page_number": 38, "total_pages": 54, "image_filename": "19930086015_p38.jpg", "text": "NACA RM A9E24 CONFIDENTIAL 37\n\n| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | |", "timestamp": "2026-07-22T04:33:40.304708+00:00"} | |
| {"citation_id": "19930085529", "source_url": "https://ntrs.nasa.gov/api/citations/19930085529/downloads/19930085529.pdf", "page_number": 85, "total_pages": 85, "image_filename": "19930085529_p85.jpg", "text": "84\nNACA RM No. L8A30a\n\nTABLE 75\n$$[\\Lambda = -45^\\circ, \\delta_{te} = 9.8^\\circ, \\alpha = 7^\\circ]$$\nCONFIDENTIAL\n\n| | | UPPER SURFACE | | | | | | | LOWER SURFACE | | | | |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| **Tube** | **Per-cent chord** | **Mach Number** | | | | | **Tube** | **Per-cent chord** | **Mach Number** | | | | |\n| | | **0.60** | **0.80** | **0.89** | **0.995** | **0.96** | | | **0.60** | **0.80** | **0.89** | **0.995** | **0.96** |\n| A 1 | 2.0 | -- | -- | -- | -- | -- | 86 | 3.0 | -- | -- | -- | -- | -- |\n| 2 | 6.0 | -- | -- | -- | -- | -- | 87 | 10.0 | -- | -- | -- | -- | -- |\n| 3 | 15.0 | -- | -- | -- | -- | -- | 88 | 25.0 | -- | -- | -- | -- | -- |\n| 4 | 27.5 | -- | -- | -- | -- | -- | 89 | 41.0 | -- | -- | -- | -- | -- |\n| 5 | 40.0 | -- | -- | -- | -- | -- | 90 | 52.5 | -- | -- | -- | -- | -- |\n| 6 | 50.0 | -- | -- | -- | -- | -- | 91 | 62.5 | -0.001 | -0.008 | -0.079 | -0.136 | -0.159 |\n| 7 | 29.0 | -0.165 | -0.282 | -0.349 | -0.436 | -0.517 | 92 | 72.5 | -.009 | -.046 | -.071 | -.129 | -.159 |\n| 8 | 67.5 | -.136 | -.284 | -.363 | -.448 | -.532 | 93 | 84.0 | -- | -- | -- | -- | -- |\n| 9 | 77.5 | -- | -- | -- | -- | -- | 94 | 94.0 | -- | -- | -- | -- | -- |\n| 10 | 86.0 | -- | -- | -- | -- | -- | | | | | | | |\n| 11 | 96.0 | -- | -- | -- | -- | -- | 95 | 3.0 | -- | -- | -- | -- | -- |\n| | | | | | | | 96 | 10.0 | -- | -- | -- | -- | -- |\n| B12 | 2.0 | -- | -- | -- | -- | -- | 97 | 25.0 | .040 | .009 | -.006 | -.038 | -.050 |\n| 13 | 6.0 | -- | -- | -- | -- | -- | 98 | 41.0 | -.019 | -.056 | -.075 | -.113 | -.104 |\n| 14 | 15.0 | -- | -- | -- | -- | -- | 99 | 52.5 | -.021 | -.064 | -.090 | -.137 | -.137 |\n| 15 | 27.5 | -.468 | -.646 | -.696 | -.907 | -1.112 | 100 | 62.5 | -.028 | -.059 | -.090 | -.135 | -.123 |\n| 16 | 40.0 | -.611 | -.648 | -.640 | -.667 | -.795 | 101 | 72.5 | .004 | -.038 | -.074 | -.099 | -.095 |\n| 17 | 50.0 | -.503 | -.604 | -.621 | -.953 | -.756 | 102 | 84.0 | .005 | -.016 | -.048 | -.106 | -.110 |\n| 18 | 19.0 | -.392 | -.521 | -.572 | -.628 | -.694 | 103 | 94.5 | .011 | -.037 | -.051 | -.104 | -.117 |\n| 19 | 67.5 | -.298 | -.447 | -.518 | -.580 | -.658 | | | | | | | |\n| 20 | 77.5 | -.208 | -.361 | -.444 | -.511 | -.589 | 104 | 3.0 | .528 | .695 | .722 | .708 | .707 |\n| 21 | 86.0 | -.150 | -.235 | -.303 | -.383 | -.409 | 105 | 10.0 | .326 | .453 | .447 | .448 | .436 |\n| 22 | 95.5 | -- | -- | -- | -- | -- | 106 | 25.0 | .085 | .161 | .177 | .156 | .167 |\n| | | | | | | | 107 | 41.0 | -- | -- | -- | -- | -- |\n| C23 | 2.0 | .886 | .991 | .958 | .797 | .793 | 108 | 52.5 | -- | -- | -- | -- | -- |\n| 24 | 6.0 | .085 | .718 | .808 | .768 | .816 | 109 | 62.5 | -.010 | -.042 | -.054 | -.082 | -.073 |\n| 25 | 15.0 | -.765 | -.825 | -.845 | -.723 | -.763 | 110 | 72.5 | .017 | -.016 | -.024 | -.053 | -.040 |\n| 26 | 27.5 | -.071 | -.605 | -.743 | -.713 | -.740 | 111 | 82.5 | .013 | -.008 | -.028 | -.063 | -.056 |\n| 27 | 40.0 | .468 | -.777 | -.697 | -.695 | -.736 | 112 | 94.6 | .020 | -.020 | -.026 | -.061 | -.053 |\n| 28 | 50.0 | .439 | -.567 | -.594 | -.594 | -.686 | | | | | | | |\n| 29 | 59.0 | .350 | -.480 | -.549 | -.574 | -.581 | 113 | 3.0 | .591 | .561 | .597 | .574 | .562 |\n| 30 | 67.5 | .297 | -.388 | -.460 | -.548 | -.547 | 114 | 10.0 | .381 | .336 | .353 | .339 | .348 |\n| 31 | 77.5 | .238 | -.287 | -.343 | -.417 | -.387 | 115 | 25.0 | .191 | .201 | .221 | .202 | .211 |\n| 32 | 88.0 | .140 | -.206 | -.188 | -.202 | -.258 | 116 | 41.0 | .058 | .098 | .108 | .098 | .099 |\n| 33 | 95.0 | -- | -- | -- | -- | -- | 117 | 52.5 | .046 | .042 | .065 | .038 | .046 |\n| | | | | | | | 118 | 62.5 | -- | -- | -- | -- | -- |\n| D34 | 2.0 | -1.477 | -1.184 | -1.135 | -1.184 | -1.077 | 119 | 72.5 | .046 | .032 | .042 | .024 | .033 |\n| 35 | 15.0 | -.655 | -1.012 | -.957 | -1.076 | -.998 | 120 | 82.5 | .059 | .048 | .048 | .047 | .050 |\n| 36 | 27.5 | -.513 | -.808 | -.818 | -.772 | -.772 | 121 | 94.2 | .053 | .053 | .004 | .000 | .000 |\n| 37 | 40.0 | -.481 | -.593 | -.775 | -.832 | -.839 | | | | | | | |\n| 38 | 50.0 | -.401 | -.501 | -.551 | -.759 | -.866 | 122 | 3.0 | .581 | .586 | .586 | .585 | .481 |\n| 39 | 59.0 | -.285 | -.395 | -.390 | -.323 | -.413 | 123 | 10.0 | .382 | .393 | .391 | .391 | .388 |\n| 40 | 67.5 | -- | -- | -- | -- | -- | 124 | 25.0 | .184 | .213 | .213 | .213 | .213 |\n| 41 | 77.5 | -.167 | -.204 | -.176 | -.135 | -.176 | 125 | 41.0 | .108 | .115 | .117 | .116 | .116 |\n| 42 | 87.5 | -.108 | -.117 | -.114 | -.077 | -.077 | 126 | 52.5 | .083 | .098 | .087 | .085 | .084 |\n| 43 | 94.2 | -.078 | -.114 | -.113 | -.053 | -.063 | 127 | 62.5 | .058 | .051 | .047 | .051 | .051 |\n| | | | | | | | 128 | 72.5 | .050 | .032 | .032 | .022 | .005 |\n| E44 | 2.0 | -1.410 | -1.433 | -1.336 | -1.195 | -1.119 | 129 | 78.0 | .060 | .049 | .049 | .026 | .026 |\n| 45 | 6.0 | -1.438 | -1.472 | -1.398 | -1.224 | -1.087 | 130 | 82.5 | .070 | .058 | .042 | .050 | .051 |\n| 46 | 15.0 | -1.121 | -1.207 | -1.207 | -1.088 | -.907 | 131 | 94.0 | .099 | .039 | .032 | .037 | .034 |\n| 47 | 27.5 | -.483 | -.675 | -1.129 | -.955 | -.770 | | | | | | | |\n| 48 | 40.0 | -.433 | -.470 | -.644 | -.865 | -.812 | 132 | 3.0 | .581 | .586 | .586 | .585 | .481 |\n| 49 | 50.0 | -.313 | -.440 | -.474 | -.644 | -.489 | 133 | 10.0 | .382 | .393 | .391 | .391 | .388 |\n| 50 | 59.0 | -.284 | -.355 | -.277 | -.297 | -.590 | 134 | 25.0 | .218 | .225 | .225 | .226 | .227 |\n| 51 | 67.5 | -.221 | -.281 | -.261 | -.297 | -.234 | 135 | 41.0 | .135 | .135 | .137 | .137 | .138 |\n| 52 | 77.5 | -.139 | -.172 | -.131 | -.180 | -.209 | 136 | 52.5 | .094 | .102 | .103 | .104 | .106 |\n| 53 | 88.5 | -.060 | -.098 | -.105 | -.110 | -.142 | 137 | 62.5 | .068 | .068 | .067 | .067 | .067 |\n| 54 | 95.5 | -.062 | -.101 | -.096 | -.112 | -.135 | 138 | 72.5 | .134 | .145 | .151 | .153 | .157 |\n| | | | | | | | 139 | 83.4 | .171 | .180 | .181 | .182 | .186 |\n| F55 | 2.0 | -- | -- | -- | -- | -- | 140 | 94.0 | .057 | .050 | .040 | .032 | .031 |\n| 56 | 6.0 | -1.336 | -1.472 | -1.279 | -1.247 | -1.041 | | | | | | | |\n| 57 | 15.0 | -.777 | -1.178 | -1.181 | -1.067 | -.870 | 141 | 3.0 | .575 | .592 | .601 | .605 | .606 |\n| 58 | 27.5 | -.595 | -.73", "timestamp": "2026-07-22T04:33:40.592235+00:00"} | |
| {"citation_id": "19930085930", "source_url": "https://ntrs.nasa.gov/api/citations/19930085930/downloads/19930085930.pdf", "page_number": 77, "total_pages": 92, "image_filename": "19930085930_p77.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T04:33:42.931189+00:00"} | |
| {"citation_id": "19930082487", "source_url": "https://ntrs.nasa.gov/api/citations/19930082487/downloads/19930082487.pdf", "page_number": 3, "total_pages": 33, "image_filename": "19930082487_p3.jpg", "text": "NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nTECHNICAL NOTE NO. 1813\n\nA STUDY OF FLOW CHANGES ASSOCIATED WITH AIRFOIL \nSECTION DRAG RISE AT SUPERCRITICAL SPEEDS\n\nBy Gerald E. Nitzberg and Stewart Crandall\n\nSUMMARY\n\nA study of experimental pressure distributions and section characteristics for several moderately thick airfoil sections was made. A correlation appears to exist between the drag-divergence Mach number and the free-stream Mach number for which sonic velocity occurs at the airfoil crest, the chordwise station at which the airfoil surface is tangent to the free-stream direction. It was found that, since the Mach number for which sonic velocity occurs at the airfoil crest can be estimated satisfactorily by means of the Prandtl-Glauert rule, a method is provided whereby the drag-divergence Mach number of an airfoil section at a given angle of attack can be estimated from the low-speed pressure distribution and the airfoil profile. This method was used to predict with a reasonable degree of accuracy the drag-divergence Mach number of a considerable number of airfoil sections having diverse shapes and a wide range of thickness-chord ratios.\n\nThe pressure distributions and section force characteristics of several moderately thick airfoil sections at Mach numbers above the drag-divergence Mach number were analyzed. Some of the characteristics of the flow over these airfoils at supercritical Mach numbers are discussed.\n\nINTRODUCTION\n\nThe first intensive efforts to develop airfoil sections suitable for high-subsonic-speed applications were directed toward obtaining sections with the highest possible critical speeds. It was reasoned that, if the first occurrence of local sonic velocity on the airfoil surface could be delayed to higher free-stream Mach numbers, then, the", "timestamp": "2026-07-22T04:33:43.820596+00:00"} | |
| {"citation_id": "19930082245", "source_url": "https://ntrs.nasa.gov/api/citations/19930082245/downloads/19930082245.pdf", "page_number": 8, "total_pages": 66, "image_filename": "19930082245_p8.jpg", "text": "NACA TN No. 1596\n\nRESULTS\n\nThe aerodynamic force and moment data presented herein were determined from the mechanical integration of diagrams of pressure coefficient P plotted against chord for pressures over the upper and lower surfaces of the main portion of the airfoil and of the aileron. These data may be considered to be section data. In the preparation of the figures, the various aerodynamic coefficients first were plotted against angle of attack at a given test Mach number and with aileron deflection as a parameter. From these basic plots, the variation of the coefficients with Mach number at a constant value of airfoil section normal-force coefficient and with aileron deflection as a parameter were determined. Most of the data included in this paper have been presented in this manner and thus show the variation of the coefficients with Mach number at a constant value of airfoil section normal-force coefficient.\n\nTypical pressure-distribution plots at several Mach numbers for an angle of attack of $1^\\circ$ and an aileron deflection of $0^\\circ$ are given in figure 5 for the airfoil with the true-contour aileron. In figure 6 is shown the variation of section airfoil angle of attack and section pitching-moment coefficient with Mach number for the airfoil with the true-contour aileron at constant values of airfoil section normal-force coefficient. Plots of aileron section normal-force coefficient and section hinge-moment coefficient for the true-contour aileron against Mach number are to be found in figure 7. Aileron section loads may be determined from these data.\n\nRepresentative pressure distributions for the airfoil with the beveled-trailing-edge aileron are given in figure 8. These data are for an angle of attack of $1^\\circ$ and an aileron deflection of $0^\\circ$. Figure 9 shows the variation with Mach number of the section airfoil angle of attack and section pitching-moment coefficient of the airfoil with the beveled-trailing-edge aileron at constant values of airfoil section normal-force coefficient. Figure 10 presents the aileron section normal-force and section hinge-moment characteristics of the beveled-trailing-edge aileron.\n\nThe effects of the true-contour aileron and the beveled-trailing-edge aileron on the airfoil section-normal-force-coefficient-curve slopes $\\left(\\frac{\\Delta c_n}{\\Delta \\alpha}\\right)_{\\delta_a=0^\\circ}$ and $\\left(\\frac{\\Delta c_n}{\\Delta \\delta_a}\\right)_{\\alpha=0^\\circ}$ are compared in figure 11. The slopes shown are the average values for angles of attack from $-1^\\circ$ to $1^\\circ$ and for aileron deflections from $-1^\\circ$ to $1^\\circ$. The variation of aileron section effectiveness $-\\left(\\frac{\\Delta \\alpha}{\\Delta \\delta_a}\\right)_{c_n}$ with Mach number for the airfoil with the two ailerons at various values of airfoil section normal-force", "timestamp": "2026-07-22T04:33:45.984340+00:00"} | |
| {"citation_id": "19930086081", "source_url": "https://ntrs.nasa.gov/api/citations/19930086081/downloads/19930086081.pdf", "page_number": 26, "total_pages": 44, "image_filename": "19930086081_p26.jpg", "text": "```markdown\nCONFIDENTIAL\n24\n\n<!-- Image (94, 72, 848, 755) -->\n\nFigure 8.- Aerodynamic characteristics of a semispan delta wing with a half-delta tip control surface\ntested in the presence of a large fuselage. Large fence on. Three-percent-thick control;\nR = $4.0 \\times 10^6$; M = 1.90. Flagged symbols denote repeat tests.\n\nNACA RM L9H05\n```", "timestamp": "2026-07-22T04:33:46.431622+00:00"} | |
| {"citation_id": "19930085842", "source_url": "https://ntrs.nasa.gov/api/citations/19930085842/downloads/19930085842.pdf", "page_number": 74, "total_pages": 104, "image_filename": "19930085842_p74.jpg", "text": "70\nNACA RM L9C29\n\n$\\beta, deg$ $\\alpha, deg$\n20 11.4\n100 20 3.4\n20 0.5\n\n$\\alpha, deg$ $\\beta, deg$\n11.3 30\n5.3 30\n-0.6 30\n\nPropulsive efficiency, $\\eta$\n0 20 40 60 80\n\nPropeller advance-diameter ratio, V/nD\n0 2 4 6 8 10 12\n\nNATIONAL ADVISORY\nCOMMITTEE FOR AERONAUTICS\n\n(c) Variation of $\\eta$ with V/nD.\nFigure 39.— Concluded.", "timestamp": "2026-07-22T04:33:49.501889+00:00"} | |
| {"citation_id": "19930093773", "source_url": "https://ntrs.nasa.gov/api/citations/19930093773/downloads/19930093773.pdf", "page_number": 8, "total_pages": 47, "image_filename": "19930093773_p8.jpg", "text": "NACA RM E9G09\n7\n\nconsumption (fig. 6(d)) formed a single curve for data obtained\nat altitudes below 35,000 feet; however, higher corrected specific\nfuel consumptions were obtained at altitudes of 45,000 and\n50,000 feet. At low engine speeds the trend of the data with\nincreasing altitude was inconsistent. The corrected engine fuel-\nair ratio (fig. 6(e)) and the corrected exhaust-gas temperature\n(fig. 6(f)) increased with altitude at all corrected engine speeds;\nhowever, the increase in corrected engine fuel-air ratio was\ninsignificant at a corrected engine speed of 7900 rpm and altitudes\nup to 35,000 feet.\n\nGeneralization in terms of pumping characteristics. - If a\nturbojet engine is considered as a pump that increases the energy\nlevel of the working fluid as it passes through the engine, the\nthrust may be determined by an evaluation of the energy change.\nThis change in available energy is determined by the change in\ntotal pressure and total temperature of the air flowing through\nthe engine. In this method of generalization, as in the method\npreviously discussed, changes in component efficiencies including\nthe effects of Reynolds number lessen the possibility of generaliz-\ning the data obtained at various altitudes to a single curve.\n\nThe variation of engine total-temperature ratio with engine\ntotal-pressure ratio is shown in figure 7(a) for altitudes from\n5000 to 50,000 feet at a flight Mach number of 0.21 and in\nfigure 7(b) for flight Mach numbers from 0.21 to 0.97 at an\naltitude of 25,000 feet. As the altitude was increased, the engine-\ntotal-temperature ratio increased at all values of engine-total-\npressure ratio. The data for the range of flight Mach numbers\ninvestigated at an altitude of 25,000 feet plotted as a single\ncurve at all engine-pressure ratios above approximately 1.4.\nSimilar data obtained over a range of flight Mach numbers at\nother altitudes also formed a single curve for each altitude at\nengine total-pressure ratios above approximately 1.4. From\nthe data presented in figure 7, the total pressure at the exhaust-\nnozzle outlet can be determined for any flight Mach number and\nexhaust-gas temperature at altitudes between 5000 and 50,000 feet\nand engine-total-pressure ratios above approximately 1.4. The\njet thrust can then be calculated by use of equation (8) or (9)\npresented in the appendix.\n\nEngine Windmilling Characteristics\n\nThe engine windmilling speed is shown in figure 8 as a function\nof true airspeed for altitudes from 5000 to 45,000 feet. The engine\nwindmilling speed was unaffected by changes in altitude in the range\nof airspeeds investigated.", "timestamp": "2026-07-22T04:33:51.444210+00:00"} | |
| {"citation_id": "19930082450", "source_url": "https://ntrs.nasa.gov/api/citations/19930082450/downloads/19930082450.pdf", "page_number": 6, "total_pages": 37, "image_filename": "19930082450_p6.jpg", "text": "NACA TN No. 1778\n\nFirst the values of $P_1/t_S$ and $\\frac{P_1}{L/\\sqrt{c}}$ are calculated\n\n$\\frac{P_1}{t_S} = \\frac{3.0}{0.064}$\n\n$= 46.9 \\text{ ksi}$\n\n$\\frac{P_1}{L/\\sqrt{c}} = \\frac{3.0}{20/\\sqrt{1}}$\n\n$= 0.15 \\text{ ksi}$\n\nThen a trial value of $t_W/t_S$ is assumed (for the example $\\frac{t_W}{t_S} = 0.79$ will be used). In the chart for this value of $t_W/t_S$ (fig. 4) the points corresponding to the design values of $P_1/t_S$ and $\\frac{P_1}{L/\\sqrt{c}}$ lie on the red line at $\\frac{H}{t_W} = 26$ (or $\\frac{b_W}{t_W} = 25$). Accordingly, the value of $H/t_W$ for minimum weight for $\\frac{t_W}{t_S} = 0.79$ is 26, and because the value is established by a red line, not a blue line, some value of $t_W/t_S$ other than 0.79 will give less weight. Inspection of the charts for other values of $t_W/t_S$ reveals that at the given design values of $P_1/t_S$ and $\\frac{P_1}{L/\\sqrt{c}}$ the blue region lies between $\\frac{H}{t_W} = 26$ and $\\frac{H}{t_W} = 31$ on the chart for $\\frac{t_W}{t_S} = 0.63$.\n\nBy interpolation, the panel proportions corresponding to this blue region are found to be $\\frac{H}{b_W} \\approx 29.5$ ($\\frac{b_W}{t_W} \\approx 28.5$) and $\\frac{S}{t_S} = \\frac{b_S}{t_S} \\approx 35$, and for these proportions $\\bar{\\sigma}_F \\approx 30.5 \\text{ ksi}$ and $\\sigma_{cr} \\approx 30.5 \\text{ ksi}$, which are the values for minimum weight. The actual panel dimensions can be calculated from these proportions as\n\n$t_W = \\frac{t_W}{t_S} t_S$\n\n$= 0.63(0.064)$\n\n$= 0.0403 \\text{ inch}$", "timestamp": "2026-07-22T04:33:53.287105+00:00"} | |
| {"citation_id": "19930090382", "source_url": "https://ntrs.nasa.gov/api/citations/19930090382/downloads/19930090382.pdf", "page_number": 18, "total_pages": 37, "image_filename": "19930090382_p18.jpg", "text": "20\nNACA RM L9I07\n\nCONFIDENTIAL\n\n<!-- Image (70, 109, 922, 999) -->\n\nCONFIDENTIAL\nAdvance ratio, J\n(d) M=0.43. Concluded.\nFigure 5 - Continued.", "timestamp": "2026-07-22T04:33:59.011546+00:00"} | |
| {"citation_id": "19930082447", "source_url": "https://ntrs.nasa.gov/api/citations/19930082447/downloads/19930082447.pdf", "page_number": 7, "total_pages": 24, "image_filename": "19930082447_p7.jpg", "text": "NACA TN No. 1775\n\nRESULTS AND DISCUSSION\n\nCorrelation of Experimental Data with Theoretical Solutions\n\nA theoretical investigation of the motions and hydrodynamic impact loads experienced by V-bottom seaplanes during step-landing impacts is presented in reference 1. The entire immersion process, including the conditions at the instants of maximum acceleration, maximum draft, and rebound, was analyzed from water contact until rebound. This analysis showed that the motion and time characteristics of an impact may be represented by means of generalized variables designated the load-factor coefficient, the draft coefficient, the time coefficient, and the vertical-velocity ratio. The variation of these variables during an impact was shown to be governed solely by the magnitude of the approach parameter $\\kappa$ which depends only on the trim and the flight-path angle at the instant of initial contact with the water and which may be considered a criterion of impact similarity. A single variation with $\\kappa$ consequently exists for each of the generalized variables representing the state of motion and the time corresponding to any given stage of the impact.\n\nThe basic data obtained in the present investigation are shown in table I. The experimental data corresponding to the instants of maximum acceleration, maximum draft, and rebound are compared in figures 3 to 6 with the theoretical variations of the generalized variables with the approach parameter, as presented in reference 1. The solid-line curves show the theoretical relationships and the symbols represent the experimental data. Reduction of the experimental data to the form of generalized variables was accomplished by use of the dead-rise function $f(\\beta) = \\frac{\\pi}{2\\beta} - 1$ and the aspect-ratio factor $\\varphi(A) = 1 - \\frac{\\tan \\tau}{2 \\tan \\beta}$. These relations were presented in reference 5 and correspond to the theoretical and experimental relations obtained by reference 6 and reference 7, respectively.\n\nThe variation of load-factor coefficient\n\n$$\nC_{l} = \\frac{n_{1} w g}{j_{o}^{2}} \\left( \\frac{w}{g} \\left[ \\frac{6 \\sin \\tau + \\cos^{2} \\tau}{[f(\\beta)]^{2} \\varphi(A) \\rho m} \\right] \\right)^{1/3}\n$$\n\nwith approach parameter $\\kappa$ is shown in figure 3. The upper curve shows the maximum load-factor coefficient, whereas the lower curve shows the load-factor coefficient at the instant of maximum draft. The experimental values agree well with the theoretical variation of maximum load-factor coefficient with approach parameter. At the time of maximum draft, however, the experimental values show greater scatter as a result of the inaccuracies in measuring the time of maximum draft and the acceleration at that time.", "timestamp": "2026-07-22T04:33:59.135674+00:00"} | |
| {"citation_id": "19930085965", "source_url": "https://ntrs.nasa.gov/api/citations/19930085965/downloads/19930085965.pdf", "page_number": 42, "total_pages": 67, "image_filename": "19930085965_p42.jpg", "text": "```markdown\n1160\n\nNACA RM E9E06\n\nTABLE I - SUMMARY OF PERTINENT VALUES\n\n| 1 | 2 | 3 | 4 | 5 | 6 | 7 | 8 | 9 | 10 | 11 | 12 | 13 | 14 | 15 |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| $H_{max}$ 2<br>(oer-steds) | $H_{max}$ 1<br>(oer-steds) | $B_2$<br>(deg) | $\\beta_2 - \\beta_2'$<br>(deg) | $P_T$<br>(cm) | $l_T$<br>(cm) | $N_a I_{max}$ a | $N_b I_{max}$ b | NI | $E_2$<br>(volts/cm) | Measured<br>$E_2$<br>(volts/cm) | $P_e'$<br>(watts/sq in.) | Measured<br>$P_e'$<br>(watts/sq in.) | $\\mu$ | c<br>(1/cm) |\n\nSAE 1020 steel, measured $\\gamma = 0.594 \\times 10^5$ mho/cm\n\n| 6 | 4.31 | 0.60 | 1.30 | 8.60 | 0.0049 | 16.8 | 56.1 | 51.6 | 0.0140 | 0.0127 | 0.22 | 0.25 | 2665 | 195.5 |\n| 30 | 5.63 | 4.25 | 8.82 | 9.01 | .0123 | 37.2 | 280.5 | 224.7 | .0379 | .0400 | 3.33 | 3.50 | 2665 | 195.5 |\n| 50 | 5.93 | 5.26 | 9.82 | 9.24 | .0163 | 46.7 | 468.0 | 364.0 | .0502 | .0535 | 7.44 | 7.60 | 2665 | 195.5 |\n\nArmco Magnetic Ingot Iron, measured $\\gamma = 0.873 \\times 10^5$ mho/cm\n\n| 6 | 2.83 | 1.65 | 3.90 | 8.56 | 0.0043 | 14.7 | 56.1 | 50.0 | 0.0128 | 0.0115 | 0.21 | 0.21 | 4590 | 311 |\n| 30 | 3.49 | 5.30 | 9.88 | 8.90 | .0104 | 31.1 | 280.5 | 220.2 | .0323 | .0350 | 2.87 | 2.80 | 4590 | 311 |\n| 50 | 3.60 | 6.05 | 10.10 | 9.10 | .0138 | 40.0 | 468.0 | 359.0 | .0426 | .0444 | 6.34 | 6.70 | 4590 | 311 |\n\nHipernik, measured $\\gamma = 0.248 \\times 10^5$ mho/cm\n\n| 6 | 0.256 | 6.68 | 10.15 | 8.66 | 0.0104 | 23.7 | 56.1 | 56.4 | 0.0241 | 0.0240 | 0.43 | 0.40 | 50,000 | 546 |\n| 20 | .280 | 7.55 | 10.25 | 9.13 | .0190 | 42.6 | 187.0 | 162.3 | .0458 | .0510 | 2.74 | 2.50 | 50,000 | 546 |\n| 40 | .300 | 7.88 | 10.30 | 9.54 | .0264 | 59.3 | 374.0 | 306.5 | .0667 | .0780 | 7.96 | 8.40 | 50,000 | 546 |\n\n[Figure: NACA logo]\n\n41\n```", "timestamp": "2026-07-22T04:34:05.648311+00:00"} | |
| {"citation_id": "19930086151", "source_url": "https://ntrs.nasa.gov/api/citations/19930086151/downloads/19930086151.pdf", "page_number": 18, "total_pages": 34, "image_filename": "19930086151_p18.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T04:34:06.376066+00:00"} | |
| {"citation_id": "19930092013", "source_url": "https://ntrs.nasa.gov/api/citations/19930092013/downloads/19930092013.pdf", "page_number": 6, "total_pages": 21, "image_filename": "19930092013_p6.jpg", "text": "2\nREPORT 948—NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\n$F$ lateral stick force, pounds\n$\\beta$ sideslip angle, degrees\n$\\varphi$ bank angle, degrees\n$p$ rolling velocity, radians per second\n$P$ period of oscillation, seconds\n$T_{\\frac{1}{2}}$ time for oscillation to damp to half amplitude, seconds\n$C_{\\frac{1}{2}}$ number of cycles for oscillation to damp to half amplitude\n$T_2$ time for oscillation to double amplitude, seconds\n$C_l$ rolling-moment coefficient $\\left( \\frac{\\text{rolling moment}}{qSb} \\right)$\n$C_{l\\beta}$ rate of change of rolling-moment coefficient with respect to sideslip angle $\\left( \\frac{\\partial C_l}{\\partial \\beta} \\right)$, per degree\n$C_{l\\delta_a}$ rate of change of rolling-moment coefficient with wing-tip helix angle $\\left[ \\frac{\\partial C_l}{\\partial (pb/2V)} \\right]$\n$C_{l\\delta}$ rate of change of rolling-moment coefficient with aileron deflection $\\left( \\frac{\\partial C_l}{\\partial \\delta_a} \\right)$, per degree\n$\\left( \\frac{\\partial \\delta_a}{\\partial \\beta} \\right)_s$ aileron servo-gearing ratio\n$\\left( \\frac{\\partial \\delta_t}{\\partial \\delta_a} \\right)_s$ aileron tab servo-gearing ratio\n$(\\Delta C_{l\\beta})_s$ change in $C_{l\\beta}$ due to servo action, per degree\n$\\left| \\frac{p}{\\beta} \\right|$ ratio of amplitude of rolling velocity to amplitude of sideslip angle of the oscillatory mode, per second\n$\\left| \\frac{\\varphi}{\\beta} \\right|$ ratio of amplitude of angle of bank to amplitude of sideslip angle of the oscillatory mode\n\n[Figure: Top-view diagram of an airplane with labels: \"Vane for sideslip recorder\", \"Vane for dihedral apparatus\", \"Airspeed head\"]\n\n[Figure: Front-view diagram of an airplane with dimensions: \"42'-10\"\" and \"7.5'\"]\n\n[Figure: Side-view diagram of an airplane with dimension: \"33'-10\"\"]\n\nFIGURE 1.—Three-view drawing of the test airplane. Wing area, 334 sq ft; aspect ratio, 5.5; taper ratio, 0.5.\n\nDESCRIPTION OF APPARATUS\n\nTEST AIRPLANE\n\nThe airplane used in the investigation was a conventional propeller-driven, low-midwing, single-place fighter airplane. A three-view drawing of the airplane as instrumented for flight tests is given in figure 1.\n\nEFFECTIVE-DIHEDRAL CONTROL APPARATUS\n\nTheory and design conditions.—Dihedral effect can be expressed quantitatively by the stability coefficient $C_{l\\beta}$, the rate of change of rolling-moment coefficient with angle of sideslip. The design of the present apparatus is based on the fact that a change in apparent $C_{l\\beta}$ can be obtained from actuation, by a servomechanism with an output motion $\\delta$ proportional to sideslip angle $\\beta$, of a control surface which produces rolling moment. Then\n\n$$(\\Delta C_{l\\beta})_s = C_{l\\delta} \\left( \\frac{\\partial \\delta}{\\partial \\beta} \\right)_s \\quad (1)$$\n\nPreliminary investigation showed that the most practicable method of obtaining large servo-actuated rolling-moment coefficients proportional to sideslip angle on the test airplane was by use of the normal ailerons. In order to simulate changes in $C_{l\\beta}$, the servo motion of the ailerons should not be accompanied by any resultant movement of the stick or increment in lateral stick force. This condition arises from the fact that the aileron stick-deflection and stick-force gradients $d\\theta/d\\beta$ and $dF/d\\beta$ required for balance in steady straight sideslips are, to a pilot, measures of the stick-fixed and stick-free dihedral effect. In order to obtain changes in $d\\theta/d\\beta$ and apparent stick-fixed $C_{l\\beta}$, a differential linkage is required in the control system, with aileron deflection as the output and pilot-applied stick motion and an independent servo motion as inputs. The maximum value of servo-gear ratio $(\\partial \\delta_a/\\partial \\beta)$, which then can be utilized is restricted in two ways: First, the maximum servo-actuated aileron deflection must be limited to allow the pilot sufficient aileron deflection for normal maneuvering and emergency control; and, second, the sideslip-angle range over which the apparatus is operative for any servo-gear ratio must be greater than that encountered during the desired maneuvers. These restrictions", "timestamp": "2026-07-22T04:34:07.478669+00:00"} | |
| {"citation_id": "19930082496", "source_url": "https://ntrs.nasa.gov/api/citations/19930082496/downloads/19930082496.pdf", "page_number": 3, "total_pages": 50, "image_filename": "19930082496_p3.jpg", "text": "2\nNACA TN No. 1836\n\nExamination by X-ray diffraction revealed that during operation a scale consisting of two layers formed on the ceramal blades. The outer layer was mainly titanium dioxide $TiO_2$ (rutile). The inner layer consisted mainly of cobalt titanate $CoTiO_3$. No significant change in the base material was indicated.\n\nThe investigation further yielded data, which indicated that:\n\n(a) More care should be exercised in the handling of blades made of carbide ceramals than is customary with metal blades.\n\n(b) Blades of carbide-type ceramals, having high thermal conductivities, cause the turbine-disk rim to run hotter than do metal blades.\n\n(c) For long-time operation, coatings will be required on carbide-type ceramal blades for protection against oxidation if the naturally formed oxide coating is not protective.\n\nINTRODUCTION\n\nThe current service life of metal turbine blades is relatively short at operating gas temperatures above 1800° F, and the ultimate operating temperature of such blades is limited by the melting point of their lowest melting constituent, which generally lies between 2300° and 2500° F. Research on materials for gas-turbine blades is guided by the following immediate goals: longer life at the relatively low gas temperatures of 1500° to 1800° F, a practical life at temperatures of 1800° to 2600° F, and a short but useful life at temperatures above 2600° F. These objectives necessitate research on materials having better high-temperature characteristics than those of metal alloys currently available. Research has led to a consideration of ceramic materials as a possible substitute; however, ceramic materials are at present poor in thermal-shock properties (the ability to withstand severe and abrupt temperature fluctuations of the surrounding gas). Some aspects of research in the field of ceramics are discussed in reference 1. The possibility of supplementing the desirable characteristics of a ceramic with the thermal-shock resistance of a metal suggests a composite material consisting of ceramic and metallic constituents that would result in a material of long life at high temperatures. Such materials have been termed \"ceramals\" (references 2 and 3). If the combinations and proportions of constituents in such materials are varied, some control over the properties of the material should be possible.\n\nCeramal bodies are not new, but have been suggested, in part, by the development of cemented carbides, which have been widely used", "timestamp": "2026-07-22T04:34:12.859198+00:00"} | |
| {"citation_id": "19930085930", "source_url": "https://ntrs.nasa.gov/api/citations/19930085930/downloads/19930085930.pdf", "page_number": 78, "total_pages": 92, "image_filename": "19930085930_p78.jpg", "text": "NACA RM L59G07\n\nCONFIDENTIAL\nUNCLASSIFIED\n\nAverage exit Mach number, $M_{2,av}$\n\n| | | |\n| :--- | :--- | :--- |\n| $\\bigcirc$ | 50-percent-span station | |\n| $\\square$ | 25-percent-span station | |\n| $\\diamondsuit$ | 10.15-percent-span station | |\n\n1.8\n1.6\n1.4\n1.2\n1.0\n.8\n\n0 .2 .4 .6 .8 1.0 .2 1.4 1.6\nDistance from convex surface\n\nConcave surface\n\nCONFIDENTIAL\nUNCLASSIFIED\n\nNACA\n\nFigure 36.- The variation of the average exit Mach number with distance from convex surface for model 4.\n\n77", "timestamp": "2026-07-22T04:34:14.122532+00:00"} | |
| {"citation_id": "19930086078", "source_url": "https://ntrs.nasa.gov/api/citations/19930086078/downloads/19930086078.pdf", "page_number": 29, "total_pages": 42, "image_filename": "19930086078_p29.jpg", "text": "NACA RM L9H04\n27\n\nCONFIDENTIAL\n\n$$\n\\begin{array}{c|c}\n\\frac{S_a}{S} & \\text{Nominal} \\\\\n& \\text{extension} \\\\\n\\hline\n\\square & .010 & \\frac{1}{4} \\\\\n\\diamond & .020 & \\frac{1}{2} \\\\\n\\triangle & .030 & \\frac{3}{4} \\\\\n\\circ & .040 & \\text{Full}\n\\end{array}\n$$\n\n$$\n\\begin{array}{c|c|c|c|c|c|c|c|c|c|c|c}\n\\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{.02} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c}{\\text{Rolling-moment coefficient, } C_l} \\\\\n\\cline{2-2} \\cline{12-12}\n\\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{.01} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c}{} \\\\\n\\cline{2-2}\n\\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{0} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c}{} \\\\\n\\cline{2-2}\n\\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{-.01} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c}{} \\\\\n\\cline{2-2}\n\\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{-.02} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c}{} \\\\\n\\cline{2-2}\n\\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{-.03} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c}{} \\\\\n\\cline{2-2}\n\\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{-.04} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c}{} \\\\\n\\hline\n\\text{Yawing-moment} & \\multicolumn{1}{c|}{.02} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c}{} \\\\\n\\text{coefficient, } C_n & \\multicolumn{1}{c|}{.01} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c}{} \\\\\n\\cline{2-2}\n\\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{0} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c}{} \\\\\n\\cline{2-2}\n\\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{-.01} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c}{} \\\\\n\\hline\n\\multicolumn{1}{c|}{} & -8 & -4 & 0 & 4 & 8 & 12 & 16 & 20 & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c|}{} & \\multicolumn{1}{c}{} \\\\\n\\multicolumn{1}{c|}{} & \\multicolumn{11}{c}{\\text{Angle of attack, } \\alpha, \\text{ deg}} \\\\\n\\end{array}\n$$\n\nCONFIDENTIAL\nNACA\n\n(a) $\\delta_a = 4^\\circ$.\n\nFigure 9.- Lateral control characteristics of unswept wing with triangular wing-tip aileron at various extensions.", "timestamp": "2026-07-22T04:34:14.535236+00:00"} | |
| {"citation_id": "19930086105", "source_url": "https://ntrs.nasa.gov/api/citations/19930086105/downloads/19930086105.pdf", "page_number": 19, "total_pages": 22, "image_filename": "19930086105_p19.jpg", "text": "1179\n\nNACA RM E9H12\n\nCONFIDENTIAL\n\n(a) Fuel flow, 76.5 pounds per hour; optimum total-pressure recovery, $P_2/P_0$, 0.80.\n\n(b) Fuel flow, 78.0 pounds per hour; total-pressure recovery, $P_3/P_0$, 0.66; frequency, 28 cycles per second.\n\nFigure 5. - High-speed schlieren photographs of shock pattern at diffuser inlet (approximately 2500 frames/sec). Outlet-inlet area ratio, $A_4/A_1$, 0.8.\n\nNACA\nC-23871\n8-10-49\n\nCONFIDENTIAL\n\n17", "timestamp": "2026-07-22T04:34:16.874052+00:00"} | |
| {"citation_id": "19930085838", "source_url": "https://ntrs.nasa.gov/api/citations/19930085838/downloads/19930085838.pdf", "page_number": 90, "total_pages": 118, "image_filename": "19930085838_p90.jpg", "text": "```markdown\n89\n\nFlap section hinge-moment coefficient, $C_{h_f}$\n\n$\\delta_a = 0^\\circ$\n\n$\\delta_t$ (deg)\n-5\n-4\n-3\n-2\n0\n\nSection angle of attack, $\\alpha_o$, deg\n\n(e) $\\delta_F = 40^\\circ$.\nFigure 11.- Continued.\n\nNACA\n\nNACA RM No. 19B23\n```", "timestamp": "2026-07-22T04:34:18.299750+00:00"} | |
| {"citation_id": "19930085842", "source_url": "https://ntrs.nasa.gov/api/citations/19930085842/downloads/19930085842.pdf", "page_number": 75, "total_pages": 104, "image_filename": "19930085842_p75.jpg", "text": "NACA RM 19C29\n71\n\n<!-- Image (168, 109, 905, 840) -->\n\n(a) $\\beta = 20^\\circ$.\n(b) $\\beta = 30^\\circ$.\n\nFigure 40.- Variation of $C_p$, $C_{T_e}$, and $\\eta$ with V/nD. Basic model configuration; all control surfaces neutral; data for $\\beta = 20^\\circ$ obtained with wing-tip support.", "timestamp": "2026-07-22T04:34:18.931896+00:00"} | |
| {"citation_id": "19930082245", "source_url": "https://ntrs.nasa.gov/api/citations/19930082245/downloads/19930082245.pdf", "page_number": 9, "total_pages": 66, "image_filename": "19930082245_p9.jpg", "text": "8\nNACA TN No. 1596\n\ncoefficient is given in figure 12. The values of $-\\left(\\frac{\\Delta c_l}{\\Delta \\delta_a}\\right)_{c_n}$ given are the average values for aileron deflections from $-6^\\circ$ to $4^\\circ$ for the airfoil with the true-contour aileron and from $-4^\\circ$ to $6^\\circ$ for the airfoil with the beveled-trailing-edge aileron.\n\nThe variation of the section critical Mach number of the airfoil with aileron deflection for the airfoil with the two ailerons is given in figure 13 for values of airfoil section normal-force coefficient from 0 to 0.6. The section critical Mach number was determined from the intersection of curves of minimum airfoil pressure coefficient plotted against Mach number with the curve of critical pressure coefficient plotted against Mach number. In a few cases, where the test Mach numbers were below the critical Mach number, the test data have been extrapolated a moderate amount to higher Mach numbers to obtain the critical Mach number values.\n\nOne of the problems of high-speed flight is the wing twist during rolling caused by the pitching moments developed by the lateral-control device. The rate of change of airfoil section pitching-moment coefficient with angle of attack $\\left(\\frac{\\Delta c_m}{\\Delta \\alpha}\\right)_{c_n}$ at a constant value of airfoil section normal-force coefficient is an index of the tendency of an aileron to twist a wing (as a result of the pitching moment developed by the aileron) in terms of the section effectiveness developed by the aileron, and therefore affords a proper comparison of the two ailerons as regards wing twisting. The ratio $\\left(\\frac{\\Delta c_m}{\\Delta \\alpha}\\right)_{c_n}$ was obtained by dividing values of $\\left(\\frac{\\Delta c_m}{\\Delta \\delta_a}\\right)_{c_n}$ by the corresponding values of aileron section effectiveness $\\left(\\frac{\\Delta c_l}{\\Delta \\delta_a}\\right)_{c_n}$ given in figure 12 and applies for the same deflection range as the data of figure 12. The variation with Mach number of the ratio $\\left(\\frac{\\Delta c_m}{\\Delta \\alpha}\\right)_{c_n}$ is given in figure 14.\n\nThe section hinge-moment-coefficient derivatives $\\left(\\frac{\\Delta c_h}{\\Delta \\delta_a}\\right)_{\\alpha=0^\\circ}$ and $\\left(\\frac{\\Delta c_h}{\\Delta \\alpha}\\right)_{\\delta_a=0^\\circ}$ for the two configurations are presented in figure 15.\n\nThese slopes are the average values for angles of attack from $-1^\\circ$ to $1^\\circ$ and for aileron deflections from $-1^\\circ$ to $1^\\circ$. The action of the thickened trailing edge in relieving hinge moments is shown in figure 16 by representative pressure distributions over the aileron for deflections of $-4^\\circ$ and $4^\\circ$ at an angle of attack of $1^\\circ$.", "timestamp": "2026-07-22T04:34:22.810013+00:00"} | |
| {"citation_id": "19930082485", "source_url": "https://ntrs.nasa.gov/api/citations/19930082485/downloads/19930082485.pdf", "page_number": 4, "total_pages": 62, "image_filename": "19930082485_p4.jpg", "text": "NACA TN No. 1810\n\nW weight of gas flowing across velocity-potential line per unit depth, (slug/sec)\n\n$\\sqrt{Z}$ $\\frac{V}{\\sqrt{2c_{p}T_{t}}} = \\sqrt{\\frac{\\gamma-1}{\\gamma+1}} \\frac{V}{V_{cr}}$\n\n$\\alpha$ angle between normal to cascade axis and flow direction of gas, (deg)\n\n$\\Gamma$ circulation, (sq ft/sec)\n\n$\\gamma$ ratio of specific heats, $c_{p}/c_{v}$\n\n$\\nu$ local deflection angle, (deg)\n\n$\\rho$ density, (slugs/cu ft)\n\n$\\tau$ blade pitch, (ft, unless otherwise noted)\n\n$\\phi$ velocity potential, (sq ft/sec)\n\n$\\psi$ stream function (slugs/sec)\n\nSubscripts:\n\n0 ambient static conditions\n\n1 suction (convex) surface\n\n2 pressure (concave) surface\n\ne cascade exit\n\ni cascade entrance\n\nm value at point of average streamline curvature\n\nr component in radial direction\n\ns static conditions\n\nt stagnation conditions\n\nu component in tangential direction\n\nx component in axial direction", "timestamp": "2026-07-22T04:34:27.053592+00:00"} | |
| {"citation_id": "19930082450", "source_url": "https://ntrs.nasa.gov/api/citations/19930082450/downloads/19930082450.pdf", "page_number": 7, "total_pages": 37, "image_filename": "19930082450_p7.jpg", "text": "6\nNACA TN No. 1778\n\n$$H = \\frac{\\bar{H}}{t_W} t_W$$\n\n$$= 29.5 (0.040)$$\n\n$$= 1.18 \\text{ inches}$$\n\n$$S = \\frac{\\bar{S}}{t_S} t_S$$\n\n$$= 35(0.064)$$\n\n$$= 2.24 \\text{ inches}$$\n\nand the section properties can be determined from table 3 as\n\n$$\\bar{h} = \\frac{\\bar{h}}{t_S} t_S$$\n\n$$= 3.92(0.064)$$\n\n$$= 0.251 \\text{ inch}$$\n\n$$\\rho = \\frac{\\bar{\\rho}}{t_S} t_S$$\n\n$$= 6.02(0.064)$$\n\n$$= 0.385 \\text{ inch}$$\n\nIn order to illustrate the use of the direct-reading design charts when more accuracy than that corresponding to interpolation by inspection is desired, a plot has been made (fig. 10) of the values of $\\bar{\\sigma}_f$, $\\sigma_{cr}$, and $H/t_W$ given by the charts at the design values of $P_i/t_S$ and $\\frac{P_i}{L/\\sqrt{C}}$.\n\nThe proportions which give the highest value of $\\bar{\\sigma}_f$ can be readily selected from a plot of this kind. (For the example these proportions are so nearly the same as were obtained by inspection that the values will not be repeated.)", "timestamp": "2026-07-22T04:34:28.463478+00:00"} | |
| {"citation_id": "19930082487", "source_url": "https://ntrs.nasa.gov/api/citations/19930082487/downloads/19930082487.pdf", "page_number": 4, "total_pages": 33, "image_filename": "19930082487_p4.jpg", "text": "```markdown\n2\nNACA TN No. 1813\n\nadverse lift and drag changes associated with supercritical speeds\ncould be delayed. However, experimental measurements show that the\nonset of these adverse effects is not directly related to the\ncritical Mach number. For some types of airfoil sections the drag\nrises rapidly as soon as the free-stream Mach number exceeds the\ncritical; whereas for other types no appreciable drag rise occurs\nuntil the Mach number of the free stream is considerably above the\ncritical.\n\nNo adequate method has been presented for predicting the free-\nstream Mach numbers at which the various supercritical flow changes\noccur. The purpose of the present report is to investigate these\nflow changes for two-dimensional transonic flow past conventional and\nlow-drag airfoil sections. The investigation is based primarily on\na systematic analysis of experimental data for airfoil sections of\n15-percent-chord thickness. Although this thickness is greater than\nis generally desirable for use at high speeds, it has the advantage\nthat airfoil-shape effects are less likely to be masked by boundary-\nlayer effects.\n\nSYMBOLS\n\n| Symbol | Definition |\n| :--- | :--- |\n| $\\alpha$ | airfoil section angle of attack |\n| $c$ | airfoil chord |\n| $c_d$ | airfoil section drag coefficient |\n| $c_l$ | airfoil section lift coefficient |\n| $M$ | ratio of local velocity to local velocity of sound |\n| $M_p$ | free-stream Mach number at which sonic velocity is first reached at airfoil crest |\n| $M_{cr}$ | critical Mach number of airfoil section (free-stream Mach number at which local sonic velocity is first reached on airfoil surface) |\n| $M_d$ | drag-divergence Mach number (Mach number at which slope of curve of drag coefficient versus Mach number attains a value of 0.10) |\n| $M_0$ | ratio of free-stream velocity to free-stream velocity of sound |\n```", "timestamp": "2026-07-22T04:34:30.064331+00:00"} | |
| {"citation_id": "19930086081", "source_url": "https://ntrs.nasa.gov/api/citations/19930086081/downloads/19930086081.pdf", "page_number": 27, "total_pages": 44, "image_filename": "19930086081_p27.jpg", "text": "NACA RM L9H05\nCONFIDENTIAL\n25\n\n<!-- Image (214, 78, 819, 809) -->\n\n(a) Variation of $C_L$, $C_m$, $C_D$, and $C_n$ with $\\alpha$.\n\nFigure 9.- Aerodynamic characteristics of a semispan delta wing with a half-delta tip control surface tested in the presence of a large fuselage. Fence off. Seven-percent-thick control; $R = 4.0 \\times 10^6$; $M = 1.90$. Flagged symbols denote repeat tests.", "timestamp": "2026-07-22T04:34:30.300236+00:00"} | |
| {"citation_id": "19930085965", "source_url": "https://ntrs.nasa.gov/api/citations/19930085965/downloads/19930085965.pdf", "page_number": 43, "total_pages": 67, "image_filename": "19930085965_p43.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T04:34:33.782963+00:00"} | |
| {"citation_id": "19930090382", "source_url": "https://ntrs.nasa.gov/api/citations/19930090382/downloads/19930090382.pdf", "page_number": 19, "total_pages": 37, "image_filename": "19930090382_p19.jpg", "text": "NACA RM L6I07\n21\n\nCONFIDENTIAL\n\nTip Mach number, $M_t$\nEfficiency, $\\eta$\n\n<!-- Image (89, 110, 868, 997) -->\n\nThrust coefficient, $C_T$\nPower coefficient, $C_P$\nAdvance ratio, J\n($\\beta_{0.7R}$=40°)\nFigure 5 - Continued.\n(e) M=0.53.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T04:34:36.753408+00:00"} | |
| {"citation_id": "19930086151", "source_url": "https://ntrs.nasa.gov/api/citations/19930086151/downloads/19930086151.pdf", "page_number": 19, "total_pages": 34, "image_filename": "19930086151_p19.jpg", "text": "CONFIDENTIAL\n\nNACA RM L9D28\n\n[Figure: A sweptback semispan wing mounted in a wind tunnel, with visible support structure and tunnel walls. A label on the lower right of the image reads “NACA L-58099”.]\n\nFigure 2.— The $45^\\circ$ sweptback semispan wing mounted in the Langley 300 MPH 7- by 10-foot tunnel. Plain wing with triangular wing-tip aileron.\n\nCONFIDENTIAL\n\n17", "timestamp": "2026-07-22T04:34:38.728637+00:00"} | |
| {"citation_id": "19930085930", "source_url": "https://ntrs.nasa.gov/api/citations/19930085930/downloads/19930085930.pdf", "page_number": 79, "total_pages": 92, "image_filename": "19930085930_p79.jpg", "text": "```markdown\nCONFIDENTIAL\nUNCLASSIFIED\n\n78\n\n$\\frac{P_{22}}{P_0}$\nStagnation-pressure recovery,\n\n1.0\n.8\n.6\n.4\n.2\n0\n\n0 .2 .4 .6 .8 1.0 1.2 1.4 1.6\nDistance from convex surface\n\n$\\square$ 50-percent-span station\n$\\square$ 25-percent-span station\n$\\diamond$ 10.15-percent-span station\n\nConcave surface\n\nUNCLASSIFIED\nCONFIDENTIAL\n\nNACA\n\nFigure 37.- The variation of the stagnation-pressure recovery with distance from\nconvex surface for model 4.\n\nNACA RM L9G07\n```", "timestamp": "2026-07-22T04:34:41.367523+00:00"} | |
| {"citation_id": "19930086078", "source_url": "https://ntrs.nasa.gov/api/citations/19930086078/downloads/19930086078.pdf", "page_number": 30, "total_pages": 42, "image_filename": "19930086078_p30.jpg", "text": "28\nNACA RM L9H04\n\nCONFIDENTIAL\n\n<!-- Image (89, 110, 888, 901) -->\n\n(b) $\\delta_a = 6^\\circ$.\nFigure 9.- Concluded.", "timestamp": "2026-07-22T04:34:42.410468+00:00"} | |
| {"citation_id": "19930086105", "source_url": "https://ntrs.nasa.gov/api/citations/19930086105/downloads/19930086105.pdf", "page_number": 20, "total_pages": 22, "image_filename": "19930086105_p20.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T04:34:43.350991+00:00"} | |
| {"citation_id": "19930092013", "source_url": "https://ntrs.nasa.gov/api/citations/19930092013/downloads/19930092013.pdf", "page_number": 7, "total_pages": 21, "image_filename": "19930092013_p7.jpg", "text": "APPARATUS FOR VARYING EFFECTIVE DIHEDRAL IN FLIGHT\n3\n\nbecome more severe as airspeed is decreased, since both the pilot-required aileron control and the sideslip-angle range then generally increase.\n\nIn order to obtain changes in $dF/d\\beta$ and apparent stick-free $C_{l\\beta}$, a means of canceling the hinge moment due to servo-actuated deflection of the ailerons is required, since with common differential linkages the entire hinge moment is transmitted back to the stick. It was desired for the first tests that the ratio of the stick-free value of $C_{l\\beta}$ to the stick-fixed value remain constant as the stick-fixed value was changed. This leads to the requirement that the stick-free value be zero when the stick-fixed value is zero, which is equivalent to assuming that the change in aileron hinge moment with sideslip is zero. The desired effect is approximated on the present installation by servo actuation of the aileron trim tab to furnish an aileron hinge moment equal and opposite to that arising from the servo-actuated aileron motion. As was the case for the aileron system, a differential gearing with tab angle as the output motion and the tab servo and pilot-actuated trim-tab motions as inputs is required.\n\nAlthough the discussion thus far has been confined to the static flight condition of steady straight sideslips, a similar explanation which yields similar requirements can be developed for maneuvers in which sideslip angle varies rapidly. The ideal servomechanism for producing a change in $C_{l\\beta}$ which is constant under any dynamic condition would be one with an output motion always in phase with, and a constant proportion of, the input quantity. Deviations of actual servomechanisms from this ideal cause undesired variations in $C_{l\\beta}$.\n\n**Aileron drive system.**—There are a number of mechanisms which will give the desired differential aileron motion, and the choice between them depends on the particular control system under consideration. The linkage which was used for the test airplane is illustrated schematically in figure 2. In the original aileron circuit, lateral stick motion imparted a corresponding angular motion to a control horn attached to the forward end of a torque tube which was supported by two fixed bearings. This rotation was transmitted as a linear motion by push rods attached by self-aligning bearings to the horn. In the modified installation an additional torque-tube bearing was attached to the fuselage structure just forward of the stick. The torque tube was cut immediately forward of this bearing and a universal joint installed. The original forward fixed bearing was replaced by two bearings. The one nearest the torque-tube horn restrains the tube vertically by means of roller guides but permits the tube to rotate in a horizontal plane about the vertical axis of the universal joint. The second bearing is bolted to a plate which is also free to rotate in a horizontal plane about the vertical axis of the universal joint. This plate is attached by cables to a drum on the servo motor and rotates in a horizontal plane when the drum rotates. Thus, the forward portion of the torque tube swings about the universal joint when the servo responds to a sideslip signal. The torque-tube horn and the aileron push rods then move laterally if the stick is held fixed, and this motion results in an aileron deflection proportional to drum rotation. The\n\n<!-- Image (503, 73, 921, 537) -->\n\nFIGURE 2.—Sketch of original and modified aileron-control system.\n\n<!-- Image (503, 556, 921, 913) -->\n\nFIGURE 3.—Kinematics of aileron-control system for various servo positions.", "timestamp": "2026-07-22T04:34:43.710564+00:00"} | |
| {"citation_id": "19930082447", "source_url": "https://ntrs.nasa.gov/api/citations/19930082447/downloads/19930082447.pdf", "page_number": 8, "total_pages": 24, "image_filename": "19930082447_p8.jpg", "text": "```markdown\n6\nNACA TN No. 1775\n\nAt high values of $\\kappa$ the trend of the experimental variation at maximum draft is below the theoretical curve and indicates somewhat lower accelerations. These low values of acceleration are believed to result from the time lag in the displacement measurements which results in recorded values of the time of maximum draft that are slightly greater than the actual time of maximum draft. Since the time of maximum draft occurs after the time of maximum acceleration, the greater the time lag of maximum draft, the smaller the acceleration at the indicated time of maximum draft. At low values of $\\kappa$ (high flight-path angles) the trend of the experimental data at maximum draft is somewhat above the theoretical curve. This result is explained by the presence of buoyant forces, which were neglected in the theoretical solutions. These buoyant forces become of significance only at high flight-path angles beyond the range for conventional seaplanes.\n\nThe variation of draft coefficient\n\n$$\nC_d = y \\left( \\frac{g}{W} \\left[ \\frac{f(\\beta)^2 g(A) \\rho \\kappa}{6 \\sin \\tau \\cos^2 \\tau} \\right] \\right)^{1/3} \\quad (2)\n$$\n\nwith approach parameter $\\kappa$ is presented in figure 4. The upper curve shows the maximum draft coefficient and the lower curve shows the draft coefficient at time of maximum acceleration. The experimental data are in good agreement with the theoretical curves.\n\nThe variation of time coefficient\n\n$$\nC_t = t \\dot{y}_o \\left( \\frac{g}{W} \\left[ \\frac{f(\\beta)^2 g(A) \\rho \\kappa}{6 \\sin \\tau \\cos^2 \\tau} \\right] \\right)^{1/3} \\quad (3)\n$$\n\nwith the approach parameter $\\kappa$ is shown in figure 5. The upper curve shows values for the time coefficient at the instant that the model leaves the water on the rebound. The middle curve shows the time coefficient at the instant of maximum draft. The lower curve shows the time coefficient at the instant of maximum acceleration. The test points show good agreement with the theoretical curves; the buoyant forces again account for the lower values of the experimental data at low values of $\\kappa$ (high flight-path angles).\n\nIn figure 6, the ratio of vertical velocity to initial vertical velocity $\\dot{y}/\\dot{y}_o$ is plotted against the approach parameter $\\kappa$. The upper curve shows this ratio at the instant of maximum acceleration and the lower curve shows the ratio at the instant of rebound. The experimental data show general agreement with the theoretical curves despite the low measured values of velocity which result in greater scatter of the points because of measurement error.\n```", "timestamp": "2026-07-22T04:34:44.188586+00:00"} | |
| {"citation_id": "19930093773", "source_url": "https://ntrs.nasa.gov/api/citations/19930093773/downloads/19930093773.pdf", "page_number": 9, "total_pages": 47, "image_filename": "19930093773_p9.jpg", "text": "8\nNACA RM E9G09\n\nThe internal drag of a windmilling turbojet engine is of interest, particularly on multiengine airplanes when it may be desirable to cruise with one or more engines inoperative. The ratio of windmilling drag to net thrust at maximum permissible engine speed is shown in figure 9 as a function of true airspeed for an altitude of 25,000 feet. The internal drag of a windmilling engine varied from 2 percent of the available net thrust at a true airspeed of 200 miles per hour to 15 percent at a true airspeed of 650 miles per hour. The desirability of blocking the inlet of an inoperative engine is apparent.\n\nSUMMARY OF RESULTS\n\nThe following results were obtained from an investigation of a J47 turbojet engine in the NACA Lewis altitude wind tunnel at simulated altitudes from 5000 to 50,000 feet and simulated flight Mach numbers from 0.21 to 0.97:\n\n1. The correction factors commonly used to generalize turbojet-engine performance can be used to predict performance for only a limited range of altitudes and corrected engine speeds.\n\n2. From the engine pumping characteristics, jet thrust could be predicted for any desired flight Mach number and exhaust-gas temperature at altitudes from 5000 to 50,000 feet and engine-pressure ratios above approximately 1.4.\n\n3. The temperature-limited engine speed decreased with increasing altitude, which indicated the need for a variable-area exhaust nozzle.\n\n4. In general, the exhaust-gas temperature was reduced at all engine speeds by an increase in flight Mach number.\n\n5. The specific fuel consumption at temperature-limited engine speed and a flight Mach number of 0.21 varied from 1.20 to 1.30 pounds per hour per pound of net thrust over the range of altitudes investigated. Minimum specific fuel consumption of 1.05 pounds per hour per pound of net thrust was obtained at an engine speed of approximately 6400 rpm at altitudes from 15,000 to 45,000 feet.\n\n6. As the flight Mach number was increased from 0.21 to 0.97 at temperature-limited engine speed, the specific fuel consumption increased from 1.21 to 1.43 pounds per hour per pound of net thrust. At low engine speeds the increase was much larger.", "timestamp": "2026-07-22T04:34:47.024193+00:00"} | |
| {"citation_id": "19930085842", "source_url": "https://ntrs.nasa.gov/api/citations/19930085842/downloads/19930085842.pdf", "page_number": 76, "total_pages": 104, "image_filename": "19930085842_p76.jpg", "text": "72\nNACA RM L9C29\n\n<!-- Image (102, 100, 836, 853) -->\n\nFigure 41.- Variation of $C_P$, $C_{T_e}$, and $C_{T_e}/C_Q$ with $\\beta$ of the model propellers. Basic model configuration; $\\frac{V}{nD} = 0$; data for $\\beta = 20^\\circ$ obtained with wing-tip support.", "timestamp": "2026-07-22T04:34:49.078721+00:00"} | |
| {"citation_id": "19930082450", "source_url": "https://ntrs.nasa.gov/api/citations/19930082450/downloads/19930082450.pdf", "page_number": 8, "total_pages": 37, "image_filename": "19930082450_p8.jpg", "text": "NACA TN No. 1778\n\nAs a check on the accuracy of interpolation, the magnitude of $\\overline{t}/t_S$ for these proportions can be determined from table 3 and multiplied by the values of $t_S$ and $\\overline{\\sigma}_r$ for the design. This product should be equal to the design value of $P_1$. For the example\n\n$\\overline{\\sigma}_r = 30.5 \\text{ ksi}$\n\n$\\frac{\\overline{t}}{t_S} = 1.538$\n\nand\n\n$P_1 = \\overline{\\sigma}_r \\overline{t}$\n\n$= \\overline{\\sigma}_r \\frac{\\overline{t}}{t_S} t_S$\n\n$= 30.5(1.538)(0.064)$\n\n$= 3.0 \\text{ kips per inch}$\n\nwhich agrees with the design value of $P_1$ originally assumed.\n\nLangley Aeronautical Laboratory \nNational Advisory Committee for Aeronautics \nLangley Field, Va., August 2, 1948", "timestamp": "2026-07-22T04:34:53.242672+00:00"} | |
| {"citation_id": "19930082485", "source_url": "https://ntrs.nasa.gov/api/citations/19930082485/downloads/19930082485.pdf", "page_number": 5, "total_pages": 62, "image_filename": "19930082485_p5.jpg", "text": "```markdown\n4\nNACA TN No. 1810\n\nParameters:\n\n| | |\n| :--- | :--- |\n| a | $\\frac{n_o C_1}{2}$ |\n| b | $-\\frac{C_1}{\\Delta C}$ |\n| f | $\\sqrt{z_m} (1 - z_m)^{\\frac{2-\\gamma}{\\gamma-1}} \\left[ 1 + z_m \\left( \\frac{2-\\gamma}{\\gamma-1} \\right) \\right]$ |\n| $F(t_1)$ | $\\int_0^{t_1} \\exp(t^2) dt$ |\n| $F(t_2)$ | $\\int_0^{t_2} \\exp(t^2) dt$ |\n| g | $\\frac{z_m^{3/2}}{\\gamma-1} (1-z_m)^{\\frac{2-\\gamma}{\\gamma-1}}$ |\n| J | $\\int_{C_1}^{C_2} \\exp \\left[ -\\frac{n_o}{2\\Delta C} (C^2 - C_m^2) \\right] \\frac{dC}{\\Delta C}$ |\n| K | $\\int_{C_1}^{C_2} \\exp \\left[ -\\frac{3n_o}{2\\Delta C} (C^2 - C_m^2) \\right] \\frac{dC}{\\Delta C}$ |\n| t | $C \\sqrt{-\\frac{n_o}{2\\Delta C}}$ |\n| y | $\\sqrt{-\\frac{n_o \\Delta C}{2}}$ |\n| $\\mu$ | $\\frac{W}{\\rho_t n_o \\sqrt{2c_p T_t}}$ |\n\n3024\n```", "timestamp": "2026-07-22T04:34:58.047837+00:00"} | |
| {"citation_id": "19930082487", "source_url": "https://ntrs.nasa.gov/api/citations/19930082487/downloads/19930082487.pdf", "page_number": 5, "total_pages": 33, "image_filename": "19930082487_p5.jpg", "text": "NACA TN No. 1813\n\nMₛ shock-stall Mach number (free-stream Mach number at which boundary-layer separation first occurs at rear of airfoil)\n\nP pressure coefficient $\\left( \\frac{P-P_0}{q_0} \\right)$\n\np local static pressure on the airfoil surface\n\nP₀ static pressure of the free stream\n\nq₀ dynamic pressure of the free stream\n\nx chordwise distance from airfoil leading edge\n\n(x/c)β airfoil crest (chordwise station at which the airfoil surface is tangent to the direction of the free stream)\n\nDETERMINATION OF DRAG-DIVERGENCE MACH NUMBER\n\nAs the free-stream Mach number is increased through the critical value, the first abrupt change in airfoil section characteristics, for an airfoil section at a moderate angle of attack, is the rapid increase in drag coefficient. For some airfoil sections at certain angles of attack this drag rise begins at the critical Mach number; whereas for other cases it does not begin until the free-stream Mach number is considerably greater than the critical. Some high-speed pressure distributions and section characteristics (reference 1) for the NACA 65₂-215 (a = 0.5), 66,2-215 (a = 0.6), 0015, 23015, and 4415 airfoil sections have been studied in an attempt to determine the flow changes which are associated with the appearance of this more or less abrupt drag rise. In order to supplement these data, simultaneous pressure distributions and schlieren pictures of the flow around the NACA 23015 airfoil section were obtained in the Ames 1- by 3-1/2-foot high-speed wind tunnel. Before considering the specific problem of the flow changes associated with the appearance of the supercritical drag rise, the data for this airfoil section will be discussed in detail.", "timestamp": "2026-07-22T04:34:58.308470+00:00"} | |
| {"citation_id": "19930085965", "source_url": "https://ntrs.nasa.gov/api/citations/19930085965/downloads/19930085965.pdf", "page_number": 44, "total_pages": 67, "image_filename": "19930085965_p44.jpg", "text": "NACA RM E9E06\n43\n\n1125\n\n74-1412\n\nSection A-A\n\n[Figure: Axial-flow compressor with front view of inlet guide vanes.]\n\nNACA\nC-21056\n4-1-48\n\nFigure 1. - Axial-flow compressor with front view of inlet guide vanes.", "timestamp": "2026-07-22T04:35:03.079343+00:00"} | |
| {"citation_id": "19930086081", "source_url": "https://ntrs.nasa.gov/api/citations/19930086081/downloads/19930086081.pdf", "page_number": 28, "total_pages": 44, "image_filename": "19930086081_p28.jpg", "text": "26\nNACA RM L9H05\n\nCONFIDENTIAL\n\n| | |\n| :--- | :--- |\n| $\\delta$ | |\n| *(deg)* | |\n| $\\circ$ 0 | |\n| $\\square$ 2.1 | |\n| $\\diamond$ 3.9 | |\n| $\\triangleright$ 6.4 | |\n| $\\triangleleft$ 8.2 | |\n| $\\nabla$ 10.4 | |\n| $\\triangledown$ 12.4 | |\n\n$C_l$\n.036\n.032\n.028\n.024\n.020\n.016\n.012\n.008\n.004\n0\n-.004\n-.008\n-.012\n-.016\n-.020\n-.024\n-.028\n-.032\n-.036\n\n-8 -7 -6 -5 -4 -3 -2 -1 0 1 2 3 4 5 6\n$\\alpha$, deg\n\nCONFIDENTIAL\nNACA\n\n(b) Variation of $C_l$ with $\\alpha$.\nFigure 9.- Concluded.", "timestamp": "2026-07-22T04:35:03.242255+00:00"} | |
| {"citation_id": "19930082245", "source_url": "https://ntrs.nasa.gov/api/citations/19930082245/downloads/19930082245.pdf", "page_number": 10, "total_pages": 66, "image_filename": "19930082245_p10.jpg", "text": "NACA TN No. 1596\n\nThe wing-drag coefficient $C_D$ at an aileron deflection of $0^\\circ$ for the two aileron configurations is shown plotted against Mach number in figure 17 for various values of airfoil section normal-force coefficient. The drag coefficients were determined from force-test measurements of the drag of the complete wing and are based on the effective area of the complete wing in the tunnel. The drag coefficients shown are not two-dimensional data. The data are useful, however, in showing the changes in drag coefficient with changes in Mach number, and the relative effect of the two ailerons on the drag coefficient.\n\nDISCUSSION\n\nAileron Section Effectiveness\n\nThe section effectiveness $-\\left(\\frac{\\Delta\\alpha}{\\Delta\\beta_a}\\right)_{C_n}$ of the airfoil with either aileron at moderate deflections decreased markedly with an increase in Mach number (fig. 12). This decrease between the Mach numbers of 0.25 and 0.75 amounted to about one-half the low-speed values, at the lower values of airfoil section normal-force coefficient.\n\nAt a Mach number of 0.25, thickening the trailing edge had no effect on the aileron effectiveness at moderate deflections at low values of airfoil section normal-force coefficient and reduced this effectiveness at the higher values of airfoil section normal-force coefficient (fig. 12). Low-speed two-dimensional tests (references 4 and 7) have shown losses in aileron section effectiveness when the aileron trailing edge was beveled, and low-speed three-dimensional tests (references 8 and 9) have shown similar losses in rolling effectiveness for beveled ailerons. At test Mach numbers greater than 0.25, the present tests (fig. 12) showed appreciable losses in section aileron effectiveness when the aileron trailing edge was beveled.\n\nThe airfoil with the beveled-trailing-edge aileron showed a rather abrupt loss in effectiveness at deflections greater than $12^\\circ$ for airfoil section normal-force coefficients of 0.4 and less, as indicated by the data of figure 9. This abrupt loss in effectiveness was the result of stalling of the air flow at deflections greater than $12^\\circ$ as indicated by the pressure diagrams (not shown). At airfoil section normal-force coefficients greater than 0.4, the air flow was stalling at a deflection of $12^\\circ$ and at lower deflections. The stalling at the higher values of airfoil section normal-force coefficient occurred more gradually with increase in deflection than at the lower values of airfoil section normal-force coefficient, with a corresponding more uniform change in airfoil characteristics with change in deflection.\n\nThe characteristics of the airfoil with the true-contour aileron (fig. 6) at large deflections are more uniform with change in deflection", "timestamp": "2026-07-22T04:35:04.498640+00:00"} | |
| {"citation_id": "19930086151", "source_url": "https://ntrs.nasa.gov/api/citations/19930086151/downloads/19930086151.pdf", "page_number": 20, "total_pages": 34, "image_filename": "19930086151_p20.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T04:35:09.714224+00:00"} | |
| {"citation_id": "19930086078", "source_url": "https://ntrs.nasa.gov/api/citations/19930086078/downloads/19930086078.pdf", "page_number": 31, "total_pages": 42, "image_filename": "19930086078_p31.jpg", "text": "NACA RM L9H04\n29\n\nCONFIDENTIAL\n\n<!-- Image (118, 192, 907, 760) -->\n\n(a) Aileron pivoted at 0.50-chord station.\nFigure 10.- Lateral control characteristics of unswept wing with small-\nchord wing-tip aileron at various deflections, fully extended.", "timestamp": "2026-07-22T04:35:13.299243+00:00"} | |
| {"citation_id": "19930085930", "source_url": "https://ntrs.nasa.gov/api/citations/19930085930/downloads/19930085930.pdf", "page_number": 80, "total_pages": 92, "image_filename": "19930085930_p80.jpg", "text": "NACA RM L9G07\n\nCONFIDENTIAL\nUNCLASSIFIED\n\nLocal static-pressure ratio, $\\frac{p_2}{p_0}$\n\n○ 50-percent-span station\n□ 25-percent-span station\n◇ 10.15-percent-span station\n\nConcave surface\n\nUNCLASSIFIED\nCONFIDENTIAL\n\nNACA\n\nDistance from convex surface\n\nFigure 38.- The variation of the local static-pressure ratio with distance from convex surface for model 4.\n\n79", "timestamp": "2026-07-22T04:35:16.286108+00:00"} | |
| {"citation_id": "19930086105", "source_url": "https://ntrs.nasa.gov/api/citations/19930086105/downloads/19930086105.pdf", "page_number": 21, "total_pages": 22, "image_filename": "19930086105_p21.jpg", "text": "NACA RM E9H12 CONFIDENTIAL 19\n\n$$\n\\frac{P_3}{P_0} \\text{ at optimum cold recovery}\n$$\n\no Average of 40 static orifices\n□ Maximum and minimum pressures\nindicated by balanced\ndiaphragm gage\n\n(a) Outlet-inlet area ratio, $A_4/A_1$, 0.9.\n\nCombustion-chamber-inlet static pressure, $\\frac{P_3}{P_0}$\nFree-stream total pressure\n\n(b) Outlet-inlet area ratio, $A_4/A_1$, 1.0.\n\n(c) Outlet-inlet area ratio, $A_4/A_1$, 1.1.\n\nFigure 6. - Effect of fuel-air ratio on combustion-chamber-inlet static\npressure. Regenerative-type burner.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T04:35:17.655831+00:00"} | |
| {"citation_id": "19930093773", "source_url": "https://ntrs.nasa.gov/api/citations/19930093773/downloads/19930093773.pdf", "page_number": 10, "total_pages": 47, "image_filename": "19930093773_p10.jpg", "text": "NACA RM E9G09\n9\n\n7. At an altitude of 25,000 feet, the internal drag of a windmilling engine varied from 2 percent of the available net thrust at a true airspeed of 200 miles per hour to 15 percent at a true airspeed of 650 miles per hour.\n\nLewis Flight Propulsion Laboratory,\nNational Advisory Committee for Aeronautics,\nCleveland, Ohio.", "timestamp": "2026-07-22T04:35:21.110523+00:00"} | |
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