PDF/ocr_dataset_1085.jsonl

{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 108, "total_pages": 114, "image_filename": "19930086061_p108.jpg", "text": "104\nNACA RM L9J07\n\n<!-- Image (92, 177, 822, 838) -->\n\nFigure 53.- Variation of $C_L$ with $\\alpha$ for the three wings investigated.", "timestamp": "2026-07-22T07:02:13.870766+00:00"}
{"citation_id": "19930085997", "source_url": "https://ntrs.nasa.gov/api/citations/19930085997/downloads/19930085997.pdf", "page_number": 19, "total_pages": 40, "image_filename": "19930085997_p19.jpg", "text": "```markdown\nNACA RM A9I29\n\nCONFIDENTIAL\n\nA\n8° RAMP\n.154\n.028\n8°\n5.5°\n.292\nSta.\n3.206\nSta.\n3.500\nIntersection radial \"A\" & plane of symmetry\nSta.\n3.780\nSta.\n3.934\nSta.\n4.172\nSta.\n4.424\n\nB\n13° RAMP\n.134\n.032\n8°\n5.5°\n.286\nSta.\n3.206\nSta.\n3.500\nSta.\n3.780\nSta.\n3.934\nSta.\n4.172\nSta.\n4.424\n\nC\n21° RAMP\n.100\n.040\n8°\n5.5°\n.284\n.032\n\nD\n21° RAMP\n.134\n.015 O.D.\n.137\n6°\n5°\n.271\n.032\n\nAll dimension are in inches.\n\nFigure 3.— Inlet configurations.\n\nNACA\n\nCONFIDENTIAL\n\n17\n```", "timestamp": "2026-07-22T07:02:16.787756+00:00"}
{"citation_id": "19930085992", "source_url": "https://ntrs.nasa.gov/api/citations/19930085992/downloads/19930085992.pdf", "page_number": 14, "total_pages": 32, "image_filename": "19930085992_p14.jpg", "text": "```markdown\nTABLE I.- EXPERIMENTAL DATA\n\n| Run | Weight | Spanwise position (in. from root) | q (lbs/sq ft) | $V_1$ (fps) | Mach number | Reynolds number (based on wing chord) | v (fps) | (slugs/cu ft) | $f_{B_1}$ (cps) | $f_{B_2}$ (cps) | $f_e$ (cps) | $f_a$ (cps) | $\\theta_{B_1}$ | $\\theta_{B_2}$ | $\\theta_a$ | Remarks |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| 1 | None | ------- | 0 | 0 | 0 | 0 | 443.1 | 0.0002135 | 4.94 | 30.9 | 58.3 | ---- | 0.0099 | 0.0048 | 0.0030 | Fluttered in second bending mode |\n| | None | ------- | 208.5 | 419.9 | .3929 | $1.6694 \\times 10^6$ | 0 | --------- | ---- | ---- | ---- | 34.5 | ---- | ---- | ---- | with node about 8 inches from |\n| | None | ------- | 0 | 0 | 0 | 0 | 0 | --------- | 5.02 | 31.1 | 58.4 | ---- | .0130 | .0022 | .0011 | tip; clear, sustained response |\n| 2 | 1a | 40 (tip) | 0 | 0 | 0 | 0 | 0 | --------- | 4.45 | 27.3 | 42.2 | ---- | .0283 | .0069 | .0064 | Fluttered in second bending mode |\n| | 1a | 40 (tip) | 148.1 | 355.0 | .3278 | 1.4326 | 369.6 | .002196 | ---- | ---- | ---- | 29.2 | ---- | ---- | ---- | with $1\\frac{1}{2}$-inch tip amplitude |\n| | 1a | 40 (tip) | 0 | 0 | 0 | 0 | 0 | --------- | 4.42 | 27.2 | 41.7 | ---- | .0236 | .0054 | .0075 | |\n| 3 | 2a | 40 (tip) | 0 | 0 | 0 | 0 | 0 | --------- | 4.51 | 28.2 | 43.8 | ---- | .0110 | .0047 | .0012 | Fluttered with 3-inch tip amplitude |\n| | 2a | 40 (tip) | 198.8 | 366.9 | .3393 | 1.4756 | 382.6 | .002185 | ---- | ---- | ---- | 26.0 | ---- | ---- | ---- | |\n| | 2a | 40 (tip) | 0 | 0 | 0 | 0 | 0 | --------- | 4.29 | 27.5 | 42.9 | ---- | .0166 | .0200 | .0220 | |\n| 4 | None | ------- | 204.5 | 416.0 | .3882 | 1.6596 | 437.5 | .002148 | ---- | ---- | ---- | 35.7 | ---- | ---- | ---- | Check on run 1; flutter response |\n| | None | ------- | 0 | 0 | 0 | 0 | 0 | --------- | 4.92 | 30.5 | 57.1 | ---- | .0095 | .0037 | .0013 | kept smaller to prevent damage |\n| | | | | | | | | | | | | | | | | to model |\n| 5 | 1b | 40 (tip) | 0 | 0 | 0 | 0 | 0 | --------- | 4.29 | 26.2 | 38.9 | ---- | .0070 | .0139 | .0153 | Fluttered with 4 to 5 inch |\n| | 1b | 40 (tip) | 121.1 | 320.3 | .2960 | 1.3269 | 330.0 | .002240 | ---- | ---- | ---- | 28.6 | ---- | ---- | ---- | amplitude on forward end of |\n| | 1b | 40 (tip) | 0 | 0 | 0 | 0 | 0 | --------- | 4.29 | 26.9 | 32.2 | ---- | .0143 | Not clear | .0137 | weight; weight bent at angle to |\n| | | | | | | | | | | | | | | | | airstream because of flutter |\n| 6 | 2b | 40 (tip) | 0 | 0 | 0 | 0 | 0 | --------- | 4.34 | 27.3 | 40.3 | ---- | .0124 | .0107 | .0085 | Fluttered in second bending mode |\n| | 2b | 40 (tip) | 130.5 | 332.5 | .3059 | 1.4314 | 336.5 | .002320 | ---- | ---- | ---- | 31.3 | ---- | ---- | ---- | with strong torsion response; |\n| | 2b | 40 (tip) | 0 | 0 | 0 | 0 | 0 | --------- | 4.32 | 27.7 | 40.4 | ---- | .0110 | .0152 | .0299 | small amplitude |\n| 7 | None | ------- | 207.0 | 419.5 | .3927 | 1.7583 | 433.5 | .0002225 | ---- | ---- | ---- | 34.7 | ---- | ---- | ---- | Check on run 1; good flutter response |\n| | None | ------- | 0 | 0 | 0 | 0 | 0 | --------- | 5.00 | 30.7 | 57.0 | ---- | .0107 | .0060 | .0019 | |\n| 8 | 2c | 40 (tip) | 0 | 0 | 0 | 0 | 0 | --------- | 4.39 | 28.0 | 46.8 | ---- | .0126 | .0051 | .0140 | Fluttered violently in second |\n| | 2c | 40 (tip) | 178.3 | 388.9 | .3618 | 1.6272 | 400.8 | .0002238 | ---- | ---- | ---- | 28.5 | ---- | ---- | ---- | bending mode with 2-inch tip |\n| | 2c | 40 (tip) | 0 | 0 | 0 | 0 | 0 | --------- | 4.34 | 27.7 | 47.1 | ---- | .0115 | .0090 | .0136 | amplitude |\n| 9 | 2d | 40 (tip) | 0 | 0 | 0 | 0 | 0 | --------- | 4.34 | 27.7 | 53.3 | ---- | .0083 | .0131 | .0091 | Fluttered violently in second |\n| | 2d | 40 (tip) | 197.1 | 408.9 | .3836 | 1.6687 | 427.7 | .0002173 | ---- | ---- | ---- | 29.5 | ---- | ---- | ---- | bending mode with large |\n| | 2d | 40 (tip) | 0 | 0 | 0 | 0 | 0 | --------- | 4.29 | 27.4 | 52.2 | ---- | .0142 | .0039 | .0038 | amplitude and node 8 to 10 inches |\n| | | | | | | | | | | | | | | | | from tip |\n| 10 | 3 | $30\\frac{1}{4}$ | 0 | 0 | 0 | 0 | 0 | --------- | 4.74 | 30.8 | 46.1 | ---- | .0089 | .0073 | .0059 | Good flutter response with small |\n| | 3 | $30\\frac{1}{4}$ | 148.0 | 354.5 | .3273 | 1.4573 | 366.1 | .002229 | ---- | ---- | ---- | 33.5 | ---- | ---- | ---- | amplitude |\n| | 3 | $30\\frac{1}{4}$ | 0 | 0 | 0 | 0 | 0 | --------- | 4.71 | 30.7 | 45.9 | ---- | .0113 | .0054 | .0019 | |\n| 11 | 3 | 24 | 0 | 0 | 0 | 0 | 0 | --------- | 4.83 | 29.7 | 48.0 | ---- | .0123 | .0039 | .0075 | Fluttered with about $1\\frac{1}{2}$-inch ampli- |\n| | 3 | 24 | 169.3 | 379.0 | .3514 | 1.5452 | 393.6 | .002203 | ---- | ---- | ---- | 32.8 | ---- | ---- | ---- | tude on trailing edge of weight |\n| | 3 | 24 | 0 | 0 | 0 | 0 | 0 | --------- | 4.83 | 30.0 | 47.8 | ---- | .0142 | .0045 | .0047 | |\n\nNACA\nMACH RM 19E17\n21\n```", "timestamp": "2026-07-22T07:02:17.312160+00:00"}
{"citation_id": "19930086083", "source_url": "https://ntrs.nasa.gov/api/citations/19930086083/downloads/19930086083.pdf", "page_number": 5, "total_pages": 48, "image_filename": "19930086083_p5.jpg", "text": "NACA RM L9F10\n3\n\n$C_m$ pitching-moment coefficient (M/qcS)\n$C_n$ yawing-moment coefficient (N/qbS)\nL lift, pounds (-Z)\nD drag (-X when $\\Psi = 0$), pounds\nX force along X-axis, pounds\nY force along Y-axis, pounds\nZ force along Z-axis, pounds\nL' rolling moment, foot-pounds\nM pitching moment, foot-pounds\nN yawing moment, foot-pounds\nq free-stream dynamic pressure, pounds per square foot ($\\frac{1}{2}\\rho V^2$)\nS wing area (3.67 sq ft)\n$\\bar{c}$ wing mean aerodynamic chord (M.A.C.) (1.68 ft) ($\\frac{2}{S} \\int_0^{b/2} c^2 dy$)\nb wing span (2.91 ft)\nV free-stream velocity\nR Reynolds number\n$\\delta_n$ leading-edge flap deflection measured perpendicular to hinge line, degrees\n$\\alpha$ angle of attack of wing with respect to chord plane at root of model, degrees\nc local wing chord, foot\ny lateral distance from plane of symmetry, measured parallel to Y-axis, feet", "timestamp": "2026-07-22T07:02:19.556743+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 71, "total_pages": 149, "image_filename": "19930083192_p71.jpg", "text": "NACA TN 1976\n67\n\n| Symbol | Definition |\n| :--- | :--- |\n| $u$ | vertical gust velocity at any point, feet per second |\n| $V$ | true airspeed, feet per second |\n| $V_e$ | equivalent airspeed, feet per second |\n| $V_p$ | probable speed, miles per hour |\n| $V_L$ | design level-flight speed, miles per hour |\n| $V_{max}$ | maximum indicated airspeed, miles per hour |\n| $V_{nev}$ | placard do-not-exceed speed, miles per hour |\n| $V_o$ | airspeed at maximum acceleration increment, miles per hour |\n| $W$ | weight of airplane, pounds |\n| $z$ | vertical displacement of airplane, feet |\n| $dp/dt$ | angular acceleration in roll, radians per second per second |\n| $dq/dt$ | angular acceleration in pitch, radians per second per second |\n| $d\\epsilon/d\\alpha$ | downwash factor |\n| $\\alpha$ | angle of attack, degrees |\n| $\\beta$ | angle of gust to direction of flight of airplane, degrees |\n| $\\Delta$ | incremental value |\n| $\\delta_f$ | absolute displacement of equivalent fuselage, feet |\n| $\\delta_w$ | absolute displacement of equivalent wing, feet |\n| $\\delta_{st}$ | deflection of equivalent wing under conditions corresponding to normal static design procedure, feet |\n| $\\delta_{d_{max}}$ | maximum value of $\\delta_w - \\delta_f$, feet |\n| $\\theta$ | pitch angle, degrees |", "timestamp": "2026-07-22T07:02:20.989372+00:00"}
{"citation_id": "19930086110", "source_url": "https://ntrs.nasa.gov/api/citations/19930086110/downloads/19930086110.pdf", "page_number": 2, "total_pages": 42, "image_filename": "19930086110_p2.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T07:02:21.249814+00:00"}
{"citation_id": "19930085911", "source_url": "https://ntrs.nasa.gov/api/citations/19930085911/downloads/19930085911.pdf", "page_number": 48, "total_pages": 52, "image_filename": "19930085911_p48.jpg", "text": "NACA RM E9F22 CONFIDENTIAL 47\n\nOutlet static pressure,\n$P_7$, lb/sq ft\n4000\n2000\n\nNet-thrust coefficient, $C_F$\n.6\n.4\n.2\n0\n-.2\n\nGas total-temperature\nratio, $T_7/T_0$\n7.0\n5.0\n3.0\n1.0\n\nCombustion efficiency,\n$\\eta_b$, percent\n100\n80\n60\n40\n20\n\n0 10 20 30 40 50\nTime after release, $\\tau$, sec\n\n(e) Performance variables.\n\nFigure 10. - Concluded. Time history of flight data and performance of ram-\njet unit 16-A-5.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T07:02:23.903623+00:00"}
{"citation_id": "19930085999", "source_url": "https://ntrs.nasa.gov/api/citations/19930085999/downloads/19930085999.pdf", "page_number": 13, "total_pages": 20, "image_filename": "19930085999_p13.jpg", "text": "NACA RM E9107\n\nA Strain-gage location  \nB Hardened-steel sphere  \nC Lead-wire passages  \nD Positioning guide  \nE Adjustment screw  \n\nFigure 1. – Method of blade-root fastening used in investigation of vibration in loosely mounted turbine blades.\n\n[Figure: Photograph showing a turbine blade with labeled components: “Lead-wire conduit”, “Ceramic-coated strain gage”, “Precoat”, and a scale marked “INCHES” with 0 to 4 inch marks.]\n\nNACA  \nC-22431  \n10-18-48  \n\nFigure 2. – High-temperature strain-gage instrumentation of turbine blade showing location of strain gage and lead-wire conduit.", "timestamp": "2026-07-22T07:02:26.255729+00:00"}
{"citation_id": "19930085975", "source_url": "https://ntrs.nasa.gov/api/citations/19930085975/downloads/19930085975.pdf", "page_number": 25, "total_pages": 30, "image_filename": "19930085975_p25.jpg", "text": "NACA RM L9E10\n23\n\nCONFIDENTIAL\n\n| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | 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| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | |", "timestamp": "2026-07-22T07:02:26.307265+00:00"}
{"citation_id": "19930085983", "source_url": "https://ntrs.nasa.gov/api/citations/19930085983/downloads/19930085983.pdf", "page_number": 24, "total_pages": 46, "image_filename": "19930085983_p24.jpg", "text": "22\nCONFIDENTIAL\nNACA RM A9I27\n\n<!-- Image (229, 94, 720, 851) -->\n\n(b) $C_L$ vs $C_D$, $C_h$ vs $\\alpha$.\nFigure 6.- Concluded.\nCONFIDENTIAL", "timestamp": "2026-07-22T07:02:27.602546+00:00"}
{"citation_id": "19930085991", "source_url": "https://ntrs.nasa.gov/api/citations/19930085991/downloads/19930085991.pdf", "page_number": 16, "total_pages": 24, "image_filename": "19930085991_p16.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T07:02:27.904817+00:00"}
{"citation_id": "19930085919", "source_url": "https://ntrs.nasa.gov/api/citations/19930085919/downloads/19930085919.pdf", "page_number": 42, "total_pages": 47, "image_filename": "19930085919_p42.jpg", "text": "NACA RM No. A9C21\nCONFIDENTIAL\n\nLift coefficient, $C_L$\nAngle of attack, $\\alpha$, deg\nPitching-moment coefficient, $C_m$\n\n$\\delta_f$, deg $\\delta_{f,ext}$, deg Extended-nose flap\n$\\square$ -40 40\n$\\circ$ 0 off\n---constant-chord slotted\n\n$\\delta_f$, deg $\\delta_{f,ext}$, deg Extended-nose flap\n$\\diamond$ 0 40\n$\\circ$ 0 off\n---Constant-chord slotted\n0.25c split flap\n$\\delta_f=45^\\circ$\n\n0.15c\n0.100c\n\nNACA\n\nFigure 16.- Effect of the extended-nose flap of full span on the lift and pitching-moment characteristics of the model with short fuselage. $R, 4.2 \\times 10^6$\n\n41", "timestamp": "2026-07-22T07:02:29.608637+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 98, "total_pages": 98, "image_filename": "19930086073_p98.jpg", "text": "```markdown\n.0008\n$\\Delta C_{l_{\\delta a}}$\n.0004\n0\nAngle of sideslip,\n$\\beta$, deg\n0.0\n12.1\n0 .2 .4 .6 .8 1.0 1.2\n\n.0004\n$\\Delta C_{n_{\\delta a}}$\n0\n-.0004\n-.0008\n0 .2 .4 .6 .8 1.0 1.2\n\n.0004\n$\\Delta C_{Y_{\\delta a}}$\n0\n-.0004\n0 .2 .4 .6 .8 1.0 1.2\n\nLift coefficient, $C_L$\n(a) Wing alone ($\\delta_{a_L}=+11.7^\\circ$ & $\\delta_{a_R}=-11.3^\\circ$).\n\n.0012\n$\\Delta C_{l_{\\delta a}}$\n.0008\n.0004\n0\nAngle of sideslip,\n$\\beta$, deg\n0.0\n12.0\n0 .2 .4 .6 .8 1.0 1.2\n\n$\\Delta C_{n_{\\delta a}}$\n0\n-.0004\n0 .2 .4 .6 .8 1.0 1.2\n\n.0004\n$\\Delta C_{Y_{\\delta a}}$\n0\n-.0004\n-.0008\n0 .2 .4 .6 .8 1.0 1.2\n\nLift coefficient, $C_L$\n(b) Wing + body ($\\delta_{a_L}=+10.8^\\circ$ & $\\delta_{a_R}=-10.8^\\circ$).\n\nFigure 28. — Increments of rolling-moment, yawing-moment, and side-force coefficients per degree of total aileron deflection.\n\nNACA-Langley\nNACA RM A57E04\n```", "timestamp": "2026-07-22T07:02:32.656467+00:00"}
{"citation_id": "19930085870", "source_url": "https://ntrs.nasa.gov/api/citations/19930085870/downloads/19930085870.pdf", "page_number": 61, "total_pages": 92, "image_filename": "19930085870_p61.jpg", "text": "62\nNACA RM No. L9D07\n\nCONFIDENTIAL\n\nElliptical L.E. $\\circledcirc$ $C_L$ $\\square$ $C_m$\nWedge L.E. $\\triangle$ $C_L$ $\\diamond$ $C_m$\n\n.24\n.16\n.08\n$C_L$\n0\n-.08\n-.16\n-.24\n\n.01\n$C_m$\n0\n-.01\n\nElliptical L.E. $\\circledcirc$ $C_D$ $\\square$ $L/D$\nWedge L.E. $\\triangle$ $C_D$ $\\diamond$ $L/D$\n\n.06\n.04\n$C_D$\n.02\n0\n\n6\n4\n$L/D$\n2\n0\n\n-8 -6 -4 -2 0 2 4 6 8\n$\\alpha$, deg\n\n(j) Wing 10. $w=1.826$; $R=530,000$.\nFigure 7. - Continued\nCONFIDENTIAL", "timestamp": "2026-07-22T07:02:33.084055+00:00"}
{"citation_id": "19930085970", "source_url": "https://ntrs.nasa.gov/api/citations/19930085970/downloads/19930085970.pdf", "page_number": 28, "total_pages": 30, "image_filename": "19930085970_p28.jpg", "text": "CONFIDENTIAL\n\nMaximum lift-drag ratio, $(\\frac{L}{D})_{max}$\n\n| Configuration | Reynolds number | Source |\n| :--- | :--- | :--- |\n| Wing and body | $.35-.52 \\times 10^6$ | 1- by $3\\frac{1}{2}$-ft wind tunnel |\n| Wing and body | $.35-.52 \\times 10^6$ | Calculated |\n| Wing alone | $2.35 \\times 10^6$ | Reference 3 |\n| Wing and body | $.69 \\times 10^6$ | Reference 2 |\n\nlaminar boundary layer\nturbulent boundary layer\n\nMach number, $M$\n\nFigure 10.— Effect of Mach number on the maximum lift-drag ratio.\n\nCONFIDENTIAL\n26\nNACA RM A9D09", "timestamp": "2026-07-22T07:02:33.634256+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 109, "total_pages": 114, "image_filename": "19930086061_p109.jpg", "text": "```markdown\nNACA RM L9J07\n\n1.4\n1.2\n1.0\n.8\n.6\n.4\n.2\n0\n0 4 8 12 16 20 24 28 32 36 40 44\n$\\alpha$, deg\n\n$C_L$\n\n[Figure: Graph showing $C_L$ vs $\\alpha$ for different wings and Reynolds numbers]\n\n*   Wing 2, $R \\approx 0.85 \\times 10^6$\n*   Wing 2, $R \\approx 1.42 \\times 10^6$\n*   Full-scale wing of ref. 1, $R \\approx 2.91 \\times 10^6$\n\nNACA\n\nFigure 54.- Effect of Reynolds number on the variation of $C_L$ with $\\alpha$ for wing and a comparison of the lift data of wing 2 with that of the comparable full-scale wing of reference 1.\n\n105\n```", "timestamp": "2026-07-22T07:02:38.607001+00:00"}
{"citation_id": "19930085997", "source_url": "https://ntrs.nasa.gov/api/citations/19930085997/downloads/19930085997.pdf", "page_number": 20, "total_pages": 40, "image_filename": "19930085997_p20.jpg", "text": "18\nCONFIDENTIAL\nNACA RM A9I29\n\n[Figure: Radial \"a\" (see fig. 1)]\n\nStation | $A/A_i$\n--- | ---\n3.934 | 1.000\n4.172 | 1.058\n4.424 | 1.304\n4.704 | 1.597\n4.900 | 1.825\n5.400 | 2.639\n5.900 | 4.115\n6.400 | 7.806\n6.900 | 11.760\n7.500 | 15.130\n\n0.10\n\nPlane of symmetry\n\nAll dimensions are in inches.\n\n[Figure: Area divisions used in weighted average $\\bar{P}_t/P_0$]\n\n0.02\n\n0.02\n\n0.02\n\n0.015 O.D. tubing\n\nNACA\n\nFigure 4. - Internal shape of main ducts and total-pressure-measurement locations at position 2.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T07:02:39.105996+00:00"}
{"citation_id": "19930085911", "source_url": "https://ntrs.nasa.gov/api/citations/19930085911/downloads/19930085911.pdf", "page_number": 49, "total_pages": 52, "image_filename": "19930085911_p49.jpg", "text": "```markdown\n48\nCONFIDENTIAL\nNACA RM E9F22\n\n<!-- Image (148, 110, 868, 880) -->\n\n(a) Free-stream Mach numbers as function of time.\n\n(b) Fuel flow as function of free-stream total pressure.\nFigure 11. - Comparison of free-stream Mach numbers and fuel flows for ram-\njet units 16-A-2, 16-A-3, 16-A-4, and 16-A-5.\n\nCONFIDENTIAL\n```", "timestamp": "2026-07-22T07:02:42.586290+00:00"}
{"citation_id": "19930085990", "source_url": "https://ntrs.nasa.gov/api/citations/19930085990/downloads/19930085990.pdf", "page_number": 23, "total_pages": 132, "image_filename": "19930085990_p23.jpg", "text": "```markdown\nCONFIDENTIAL\n\nAll dimensions given in\ninches unless otherwise\nspecified.\n\nNote: Leading- and trailing-\nedge radii are 0.005 inch\n\nR=1.686c\n.015 gap\n.042c .045c\n.50c\n.15c\n5.1°\nWing and tail section\n\nSym Q\n\nMoment axis\n12.00\n2.70\n14.09\n36.00\n16.00\nc=18.67\n3.60\n9.33\n6.00\n18.00\n70.50\n24.00\n63.17\n12.00\n8.00\n146.00\n\nFuselage Coordinates\n(Percent length)\n| x | r | x | r |\n| :--- | :--- | :--- | :--- |\n| 0 | 0 | 61.60 | 4.398 |\n| 3.42 | 0.904 | 68.50 | 4.252 |\n| 6.85 | 1.480 | 75.30 | 3.992 |\n| 10.27 | 1.958 | 83.20 | 3.575 |\n| 13.70 | 2.370 | 84.95 | 3.445 |\n| 20.55 | 3.034 | 86.30 | 3.310 |\n| 27.40 | 3.553 | 89.00 | 2.938 |\n| 34.23 | 3.939 | 91.80 | 2.460 |\n| 41.10 | 4.211 | 94.50 | 1.835 |\n| 47.92 | 4.375 | 97.30 | 1.034 |\n| 54.80 | 4.430 | 99.30 | 0.293 |\n| 55.50 | 4.439 | 100.00 | 0 |\n\nNACA\n\n(a) Horizontal tail mounted in extended wing-chord plane.\nFigure 1.- Semispan model of an airplane with a wing and all- movable horizontal tail of aspect ratio 4.\n\nNACA RM A9J01\nCONFIDENTIAL\n21\n```", "timestamp": "2026-07-22T07:02:43.641584+00:00"}
{"citation_id": "19930085999", "source_url": "https://ntrs.nasa.gov/api/citations/19930085999/downloads/19930085999.pdf", "page_number": 14, "total_pages": 20, "image_filename": "19930085999_p14.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T07:02:44.914375+00:00"}
{"citation_id": "19930086020", "source_url": "https://ntrs.nasa.gov/api/citations/19930086020/downloads/19930086020.pdf", "page_number": 8, "total_pages": 22, "image_filename": "19930086020_p8.jpg", "text": "6 CONFIDENTIAL NACA RM A9J06\n\nwing-flow data and data at higher Reynolds numbers from the Ames 12-foot pressure wind tunnel has been obtained in the case of a wing with an unswept plan form (reference 6) and in unreported tests of a triangular wing, it appears that the discrepancy on the test wings might be attributed to the extremely high sweep and/or the lower than ordinary Reynolds number of the tests which the limitations on the model size made necessary. One of the more likely factors was thought to be a boundary-layer drain spanwise along the wing which would be likely to be present on the test wings in view of their high sweepback and which would be aggravated if the low Reynolds number of the tests caused separation (and a resulting \"tunnel\" along which the boundary layer from the wing-flow test station could drain). Another possible source of error could have been the spanwise velocity gradients which existed on the wing-flow test station which, if they caused a change in spanwise loading, would, on a wing of such high sweep, show up as an appreciable longitudinal shift of the aerodynamic center. In an attempt to determine which of the foregoing factors might contribute to the anomalous results, the supplementary tests outlined below were performed on the symmetrical untwisted wing.\n\nTo determine the effect of the spanwise velocity gradient which existed at the wing-flow station, the test setup was duplicated on the side wall of the Ames 1- by 3-1/2-foot high-speed wind tunnel. The entire wing-flow balance was mounted on the outside of the tunnel, with the turntable flush with the inside of the tunnel wall and the semispan wing model projecting into the tunnel air stream. This gave a test configuration which duplicated in all essential respects the wing-flow setup with the exceptions that the spanwise velocity gradient was negligible and the ratio of boundary-layer-displacement thickness to model span was considerably larger. The results summarized in figure 9 show negligible change for pitching-moment-curve slope, checking the wing-flow data within the measurement accuracy limitations. The discrepancy therefore does not appear to be caused by spanwise velocity gradient.\n\nWhile the model and balance were mounted in the tunnel, the effect of oscillation of the model on the test data was also determined. Tests were conducted at constant Mach number both by continuous recording of forces and moments while oscillating the model over the angle-of-attack range and by recording at various fixed angles of attack. There was no observable difference between the results of these two techniques.\n\nTo either eliminate or change any possible spanwise boundary-layer drain along the test wing, two model modifications were tested by the wing-flow technique. The first was the addition of a fuselage, which it was reasoned would place the model wing root well out of the test station boundary layer and thus reduce the tendency for spanwise drain. (See fig. 3(b).) The other modification tested was a boundary-layer fence placed 0.4 inch above the test-station surface where it would obstruct the spanwise drain of the boundary layer along the span of the model\n\nCONFIDENTIAL", "timestamp": "2026-07-22T07:02:48.006111+00:00"}
{"citation_id": "19930085983", "source_url": "https://ntrs.nasa.gov/api/citations/19930085983/downloads/19930085983.pdf", "page_number": 25, "total_pages": 46, "image_filename": "19930085983_p25.jpg", "text": "```markdown\nNACA RM A9I27\nCONFIDENTIAL\n\nLift coefficient, $C_L$\n$\\delta_u$, deg\no 0\n□ -5\n◇ -10\n△ -15\n▽ -20\n▼ -25\n\n.8\n.6\n.4\n.2\n0\n-.2\n-.4\n-.6\n\n-8 -4 0 4 8 12 16\nAngle of attack, $\\alpha$, deg\n\nfor $\\delta_u = 0^\\circ$\n.16 .12 .08 .04 0 -.04 -.08\nPitching-moment coefficient, $C_m$\n\n[Figure: Graph showing $C_L$ vs $\\alpha$ and $C_L$ vs $C_m$]\n\n(a) $C_L$ vs $\\alpha$, $C_L$ vs $C_m$.\n\nFigure 7.- The effect of elevon deflection on the aerodynamic characteristics of the wing-fuselage combination and on the elevon hinge-moment coefficients at a Mach number of 0.89.\n\nCONFIDENTIAL\n23\n```", "timestamp": "2026-07-22T07:02:51.732610+00:00"}
{"citation_id": "19930085870", "source_url": "https://ntrs.nasa.gov/api/citations/19930085870/downloads/19930085870.pdf", "page_number": 62, "total_pages": 92, "image_filename": "19930085870_p62.jpg", "text": "NACA RM No. L9D07\n63\n\nCONFIDENTIAL\n\nElliptical L.E. {○ $C_L$\n                {□ $C_m$\nWedge L.E. {△ $C_L$\n           {◇ $C_m$\n\n$C_L$\n.24\n.16\n.08\n0\n-.08\n-.16\n-.24\n\n$C_m$\n.01\n0\n-.01\n\nElliptical L.E. {○ $C_D$\n                {□ $L/D$\nWedge L.E. {△ $C_D$\n           {◇ $L/D$\n\n$C_D$\n.06\n.04\n.02\n0\n\n$L/D$\n6\n4\n2\n0\n\n$\\alpha$, deg\n-8 -6 -4 -2 0 2 4 6 8\n\n[NACA logo]\n\n(k) Wing II. w=2.193; R = 460,000.\nFigure 7. - Concluded.\nCONFIDENTIAL", "timestamp": "2026-07-22T07:02:59.893550+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 72, "total_pages": 149, "image_filename": "19930083192_p72.jpg", "text": "68\nNACA TN 1976\n\n$\\lambda$\ngust spacing (distance between gust peaks), chords\n\n$\\mu_g$\nmass parameter ($\\mu_g$ is twice the relative density (reference 24) divided by the lift-curve slope)\n$$ \\left( \\frac{2W}{\\rho g S c \\frac{dC_L}{d\\alpha}} \\right) $$\n\n$\\rho$\ndensity of air, slugs per cubic foot\n\n$\\rho_o$\ndensity of air at sea level, slugs per cubic foot\n\nSubscripts:\n\nav\naverage\n\np\nprobable; denotes the most probable value of a quantity as determined from experimental data\n\no\ndenotes value computed from local measurements\n\ncg\ncenter of gravity\n\nw\nwing\n\nl\nfor extended equations denotes point at which solution applies\n\nmax\nmaximum\n\nSubscripts used with $\\Delta n$:\n\no\nacceleration increment due to a gust\n\nm\nacceleration increment due to vertical motion of airplane\n\n$\\theta$\nacceleration increment due to pitch of airplane\n\nq\nacceleration increment due to angular velocity of airplane\n\n$\\epsilon$\nacceleration increment due to downwash\n\nw\nacceleration increment on wing", "timestamp": "2026-07-22T07:03:05.348818+00:00"}
{"citation_id": "19930085975", "source_url": "https://ntrs.nasa.gov/api/citations/19930085975/downloads/19930085975.pdf", "page_number": 26, "total_pages": 30, "image_filename": "19930085975_p26.jpg", "text": "24\nNACA RM L9E10\n\nCONFIDENTIAL\n\n$\\alpha$\n(deg)\n0.30\n3.45\n6.50\n0 Theory, ref. 2\n0 Theory, ref. 3\n\n$C_{l_p}$\n0\n-.1\n-.2\n-.3\n-.4\n-.5\n\n$\\left(\\frac{pb}{2V}\\right)_\\delta$\n.006\n.004\n.002\n0\n\n$C_{l_\\delta}$\n.002\n.001\n0\n\nMach number, M\n4 5 6 7 8 9 10\n\nCONFIDENTIAL\nNACA\n\nFigure 13.— The variation with Mach number of the parameters $C_{l_p}$, $C_{l_\\delta}$, and $\\left(\\frac{pb}{2V}\\right)_\\delta$ at several angles of attack for the 32.6° sweptback wing.", "timestamp": "2026-07-22T07:03:06.800798+00:00"}
{"citation_id": "19930085970", "source_url": "https://ntrs.nasa.gov/api/citations/19930085970/downloads/19930085970.pdf", "page_number": 29, "total_pages": 30, "image_filename": "19930085970_p29.jpg", "text": "```markdown\nNACA RM A5E09\nCONFIDENTIAL\n\n| M | Configuration | Reynolds number | Source |\n| :--- | :--- | :--- | :--- |\n| $\\circ$ 0.50 | Wing and body | $.35-.52 \\times 10^6$ | 1- by $3\\frac{1}{2}$-ft wind tunnel |\n| $\\square$ 0.70 | Wing alone | $2.35 \\times 10^6$ | Reference 3 |\n| $\\diamond$ 0.80 | Wing and body | $.69 \\times 10^6$ | Reference 2 |\n| $\\nabla$ 0.90 | | | |\n| $\\Delta$ 0.95 | | | |\n| $\\nabla$ 1.09 | | | |\n| $\\Delta$ 1.14 | | | |\n| $\\nabla$ 1.24 | | | |\n| $\\Delta$ 1.51 | | | |\n\nLift coefficient, $C_L$\nPitching-moment coefficient, $C_m$ (for $M=.50$)\n\n[Graph plotting Lift coefficient vs Pitching-moment coefficient with multiple data series corresponding to the legend above]\n\nNACA\n\nFigure 11.- Variations of pitching-moment coefficient with lift coefficient at the various test Mach numbers.\n\nCONFIDENTIAL\n27\n```", "timestamp": "2026-07-22T07:03:06.979546+00:00"}
{"citation_id": "19930085997", "source_url": "https://ntrs.nasa.gov/api/citations/19930085997/downloads/19930085997.pdf", "page_number": 21, "total_pages": 40, "image_filename": "19930085997_p21.jpg", "text": "NACA RM A9I29 CONFIDENTIAL 19\n\nMaximum total-pressure ratio, $(H_2/H_0)_{max}$, $(H_3/H_0)_{max}$\n\n1.0\nNormal shock wave\nNose inlet models\n(reference 4)\n\n.9\n\n.8\n\n.7\n\n.6\n□ Configuration A, $(H_3/H_0)_{max}$\n◇ Configuration B, \"\n△ Configuration C, \"\n○ Configuration D, \"\n◊ Configuration D, $(\\bar{H}_2/H_0)_{max}$\n\n.5\n\n.4\n\n.3\n\n.2\n\n1.0 1.2 1.4 1.6 1.8 2.0 2.2\nMach number, $M_o$\n\nFigure 5.—Variation of maximum total-pressure ratio with free-stream Mach number; $\\alpha$, $0^\\circ$; $m_4/m_0$, max.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T07:03:07.958680+00:00"}
{"citation_id": "19930085999", "source_url": "https://ntrs.nasa.gov/api/citations/19930085999/downloads/19930085999.pdf", "page_number": 15, "total_pages": 20, "image_filename": "19930085999_p15.jpg", "text": "NACA RM E9107\n13\n\n[Figure: Oscillograph records showing four traces labeled \"Frequency standard\", \"Vibration signal\", \"Interference level\", and \"Tachometer\"]\n\n(a) Engine speed, 10,000 rpm.\n\n[Figure: Oscillograph records showing four traces labeled \"Frequency standard\", \"Vibration signal\", \"No signal\", and \"Tachometer\"]\n\n(b) Engine speed, 11,500 rpm.\n\nNACA\nC-23987\n9-6-49\n\nFigure 3. - Oscillograph records of data used for computing turbine speed, vibration frequency, and vibratory stress.", "timestamp": "2026-07-22T07:03:11.179034+00:00"}
{"citation_id": "19930086110", "source_url": "https://ntrs.nasa.gov/api/citations/19930086110/downloads/19930086110.pdf", "page_number": 3, "total_pages": 42, "image_filename": "19930086110_p3.jpg", "text": "NACA RM E9H15 CONFIDENTIAL\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nRESEARCH MEMORANDUM\n\nDOWNWASH IN VORTEX REGION BEHIND TRAPEZOIDAL-WING TIP AT MACH NUMBER 1.91\n\nBy J. L. Cummings, H. Mirels, and L. E. Baughman\n\nSUMMARY\n\nThe results of an experimental investigation to determine the downwash in the region of the trailing vortex system behind a trapezoidal-wing tip in a supersonic stream are presented. The wing was cut along the inner Mach line from the tip and had a 5-percent-thick symmetrical diamond cross section. The investigation was made at a Mach number of 1.91 and a Reynolds number of $1.56 \\times 10^6$ based on the wing chord. A wake survey was also conducted.\n\nFor small angles of attack, the experimental spanwise variation of $-d\\epsilon/d\\alpha$ (where $\\epsilon$ is the downwash angle and $\\alpha$ is the angle of attack) at each chordwise station was generally similar to the variation predicted by linearized theory. At higher angles of attack, the spanwise location of the maximum value of $-d\\epsilon/d\\alpha$ at each chordwise station occurred nearer the center of the theoretical vortex sheet than linearized theory would indicate. This result was mainly attributed to differences between the theoretical and actual spanwise distribution of vorticity and to the distortion of the vortex sheet, but the individual effects could not be isolated. A survey of the friction wake indicated several characteristics similar to those observed in subsonic flow.\n\nINTRODUCTION\n\nA knowledge of the downwash behind supersonic lifting surfaces is necessary for the rational design of supersonic airplane or missile configurations. Methods for obtaining linearized solutions for the downwash behind supersonic wings are now available (for example, references 1 to 3). Linearized theory, however, assumes thin wings at small angles of attack and neglects viscous effects and the displacement and the distortion of the trailing vortex system. An experimental program is required to check the indications of linearized theory and to determine the necessary\n\nCONFIDENTIAL", "timestamp": "2026-07-22T07:03:12.144265+00:00"}
{"citation_id": "19930086083", "source_url": "https://ntrs.nasa.gov/api/citations/19930086083/downloads/19930086083.pdf", "page_number": 6, "total_pages": 48, "image_filename": "19930086083_p6.jpg", "text": "```markdown\n4\nNACA RM L9F10\n\n$$C_{l\\psi} = \\left(\\frac{\\partial C_l}{\\partial \\psi}\\right)_\\alpha$$\n\n$$C_{n\\psi} = \\left(\\frac{\\partial C_n}{\\partial \\psi}\\right)_\\alpha$$\n\n$$C_{Y\\psi} = \\left(\\frac{\\partial C_Y}{\\partial \\psi}\\right)_\\alpha$$\n\n$$C_{L\\alpha} = \\left(\\frac{\\partial C_L}{\\partial \\alpha}\\right)_\\psi$$\n\nThe subscripts $\\alpha$ and $\\psi$ indicate the factor held constant.\n\nAPPARATUS AND MODEL\n\nThe model was tested in the Langley 300 MPH 7- by 10-foot tunnel on a single strut as shown in figure 3.\n\nThe general arrangement of the 60° delta-wing model with round and beveled leading edges is shown in figure 1. The wing had a 60° apex angle and the aspect ratio was 2.31. The model was made from a flat steel plate 1/2 inch thick and had a fixed beveled trailing edge with an included angle of 10°. The flat airfoil was used because of simple construction, and, because of the preliminary aspect of the investigation, it was felt that a flat plate would suffice. The model was interchangeably equipped with either a 0.25-inch radius or a beveled leading-edge flap with an included angle of 25°. Each leading-edge flap was attached to the main body of the model by means of a brass strip in the lower surface of the wing which also served as a hinge. For the flap-deflected condition, the upper-surface gap between the flap and the wing was filled with wax to a circular contour which was tangent to the surface of the flap and the wing.\n```", "timestamp": "2026-07-22T07:03:12.525625+00:00"}
{"citation_id": "19930085991", "source_url": "https://ntrs.nasa.gov/api/citations/19930085991/downloads/19930085991.pdf", "page_number": 17, "total_pages": 24, "image_filename": "19930085991_p17.jpg", "text": "NACA RM L9T28\n15\n\n[Figure: Photograph of the test section of the Langley 20-foot free-spinning tunnel with an airplane model spinning in the tunnel. The image shows a large, octagonal chamber with observation windows. Inside, a model airplane is suspended in the center. Three people are visible looking through the windows from an upper level. A label in the bottom right corner of the photo reads \"NACA L-49005\".]\n\nFigure 2.- Photograph of the test section of the Langley 20-foot free-spinning tunnel with an airplane model spinning in the tunnel.", "timestamp": "2026-07-22T07:03:13.878701+00:00"}
{"citation_id": "19930085983", "source_url": "https://ntrs.nasa.gov/api/citations/19930085983/downloads/19930085983.pdf", "page_number": 26, "total_pages": 46, "image_filename": "19930085983_p26.jpg", "text": "24\nCONFIDENTIAL\nNACA RM A9I27\n\n<!-- Image (209, 109, 702, 850) -->\n\n(b) $C_L$ vs $C_D$, $C_h$ vs $\\alpha$.\nFigure 7.- Concluded.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T07:03:14.538750+00:00"}
{"citation_id": "19930085990", "source_url": "https://ntrs.nasa.gov/api/citations/19930085990/downloads/19930085990.pdf", "page_number": 24, "total_pages": 132, "image_filename": "19930085990_p24.jpg", "text": "22\nCONFIDENTIAL\nNACA RM A9I0L\n\nAll dimensions given in inches\nunless otherwise specified\n\nTail fairing body for mounting\nhorizontal tail above the fuselage\n\n$\\mathcal{L}$ of horizontal tail\n\nOriginal fuselage\nModified fuselage\n\n| Tail fairing body coordinates (percent length) | | Modified fuselage coordinates (percent length of original fuselage) | |\n| :--- | :--- | :--- | :--- |\n| $x_1$ | $r_1$ | $x_2$ | $r_2$ |\n| 0 | 0 | 68.50 | 4.252 |\n| 7.41 | 4.96 | 75.90 | 3.972 |\n| 14.81 | 14.58 | 82.20 | 3.960 |\n| 22.22 | 18.48 | 89.09 | 3.890 |\n| 29.63 | 20.78 | 95.90 | 3.822 |\n| 37.04 | 21.85 | 102.80 | 2.897 |\n| 50.00 | 22.22 | 108.25 | 2.040 |\n| 62.96 | 21.85 | 109.60 | 1.746 |\n| 70.37 | 20.78 | 111.00 | 1.438 |\n| 77.78 | 18.48 | 112.30 | 1.103 |\n| 85.19 | 14.58 | 113.70 | 0.758 |\n| 92.59 | 4.96 | | |\n| 100.00 | 0 | | |\n\n(b) Horizontal tail mounted above the original and the modified fuselage.\n\nFigure 1.- Concluded.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T07:03:15.747253+00:00"}
{"citation_id": "19930085870", "source_url": "https://ntrs.nasa.gov/api/citations/19930085870/downloads/19930085870.pdf", "page_number": 63, "total_pages": 92, "image_filename": "19930085870_p63.jpg", "text": "64\nNACA RM No. L9D07\n\nCONFIDENTIAL\n\n.24\nRound L.E. $\\begin{cases} \\bigcirc C_L \\\\ \\square C_m \\end{cases}$\nSharp L.E. $\\begin{cases} \\triangle C_L \\\\ \\diamond C_m \\end{cases}$\n\n.16\n\n.08\n$C_L$\n0\n\n-.08\n\n-.16\n\n-.24\n\n.01\n$C_m$\n0\n\n-.01\n\n.06\nRound L.E. $\\begin{cases} \\bigcirc C_D \\\\ \\square L/D \\end{cases}$\nSharp L.E. $\\begin{cases} \\triangle C_D \\\\ \\diamond L/D \\end{cases}$\n\n.04\n$C_D$\n\n.02\n\n0\n-8 -6 -4 -2 0 2 4 6 8\n$\\alpha$, deg\n\n12\n$L/D$\n\n8\n\n4\n\n0\n\n[NACA logo]\n\n(a) Wing 1. R (round L.E.) = 1,060,000.\nFigure 8.- Aerodynamic characteristics of flat-\nplate triangular wings at M = 1.92.\nCONFIDENTIAL", "timestamp": "2026-07-22T07:03:18.175921+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 73, "total_pages": 149, "image_filename": "19930083192_p73.jpg", "text": "NACA TN 1976\n69\n\ns\nacceleration increment on stabilizing surface\n\nT\ntotal acceleration increment\n\nExamples:\n\n$\\Delta n_{O_W}$\nacceleration increment on wing due to gust\n\n$\\Delta n_{T_S}$\ntotal acceleration increment due to tail", "timestamp": "2026-07-22T07:03:18.571589+00:00"}
{"citation_id": "19930085919", "source_url": "https://ntrs.nasa.gov/api/citations/19930085919/downloads/19930085919.pdf", "page_number": 43, "total_pages": 47, "image_filename": "19930085919_p43.jpg", "text": "42\nCONFIDENTIAL\nNACA RM No. A9C21\n\n<!-- Image (109, 121, 868, 817) -->\n\nFigure 17- Effect of the drooped-nose flap and the extended-nose flap of full span on the drag characteristics of the model with short fuselage. R, 4.2x10⁵\n\nCONFIDENTIAL", "timestamp": "2026-07-22T07:03:20.121940+00:00"}
{"citation_id": "19930086020", "source_url": "https://ntrs.nasa.gov/api/citations/19930086020/downloads/19930086020.pdf", "page_number": 9, "total_pages": 22, "image_filename": "19930086020_p9.jpg", "text": "NACA RM A9J06 CONFIDENTIAL 7\n\nwing. (See fig. 3(a).) Neither of these modifications resulted in significant changes in the pitching-moment data (fig. 9) so that no confirmation of the hypothesis as to boundary-layer drain was obtained.\n\nThe fact that low Reynolds number alone is not sufficient to account for the doubtful pitching-moment results is deducible from the fact that results of Ames 1- by 3-1/2-foot high-speed wind-tunnel tests (reference 3) of a full-span model of the symmetrical untwisted wing gave an extreme aft position of the aerodynamic center rather than an extreme forward position as in the wing-flow tests. The comparison of these various tests is presented in the following table:\n\n| Mach number | Wing-flow method | | 12-foot pressure wind tunnel | | 1- by 3-1/2-foot high-speed wind tunnel | |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| | Aerodynamic center ($\\% \\bar{c}$) | Reynolds number | Aerodynamic center ($\\% \\bar{c}$) | Reynolds number | Aerodynamic center ($\\% \\bar{c}$) | Reynolds number |\n| 0.6 | 25 | $0.55 \\times 10^6$ | 42 | $2.35 \\times 10^6$ | 54 | $0.42 \\times 10^6$ |\n| .9 | 25.5 | $.73 \\times 10^6$ | 44 | $2.35 \\times 10^6$ | 60 | $.51 \\times 10^6$ |\n| 1.1 | 26 | $.81 \\times 10^6$ | - | - - - - | 74 | $.53 \\times 10^6$ |\n\nThat the discrepancies cannot be attributed to the semispan mounting alone is deducible from the fact that the 12-foot pressure wind tunnel has obtained good correlation on results of semispan and full-span $63^\\circ$ swept wings at a Reynolds number of the order of 2 million. Further verification of the semispan testing technique (at high Reynolds number) is contained in reference 7, where a comparison is presented of the data obtained from both semispan and full-span models of a $40^\\circ$ swept-back wing.\n\nIn view of the foregoing discussion no substantiated explanation can be presented of the cause of the discrepancy between the wing-flow pitching-moment characteristics and those presented in references 1 and 4. Therefore it can only be concluded that the wing-flow data on a wing of this plan form cannot be relied upon even qualitatively as an indication of trends.\n\nCONCLUDING REMARKS\n\nThe data presented in this report indicate considerable discrepancy in the pitching-moment characteristics for a highly swept and tapered plan form as measured by the wing-flow method and by the larger-scale Ames 12-foot pressure wind tunnel. Attempts to account for the differences by modifying the wing-flow model configuration and technique were inconclusive.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T07:03:21.458057+00:00"}
{"citation_id": "19930085970", "source_url": "https://ntrs.nasa.gov/api/citations/19930085970/downloads/19930085970.pdf", "page_number": 30, "total_pages": 30, "image_filename": "19930085970_p30.jpg", "text": "```markdown\n28\n\nCONFIDENTIAL\n\n80\n60\n40\n20\n0\n0 .2 .4 .6 .8 1.0 1.2 1.4 1.6\nMach number, M\nLocation of the aerodynamic center, percent m.a.c.\n\nConfiguration Reynolds number Source\nWing and body .35-.52 x 10^6 1- by 3 1/2-ft wind tunnel\nWing alone 2.35 x 10^6 Reference 3\nWing alone Calculated\n\nNACA\n\nFigure 12.- Effect of Mach number on the location of the aerodynamic center.\n\nCONFIDENTIAL\nNACA RM A9E09\nNACA-Langley-7-7-49 - 350\n```", "timestamp": "2026-07-22T07:03:23.542606+00:00"}
{"citation_id": "19930085991", "source_url": "https://ntrs.nasa.gov/api/citations/19930085991/downloads/19930085991.pdf", "page_number": 18, "total_pages": 24, "image_filename": "19930085991_p18.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T07:03:25.016470+00:00"}
{"citation_id": "19930086110", "source_url": "https://ntrs.nasa.gov/api/citations/19930086110/downloads/19930086110.pdf", "page_number": 4, "total_pages": 42, "image_filename": "19930086110_p4.jpg", "text": "2 CONFIDENTIAL NACA RM E9H15\n\nmodifications that will result in agreement between theory and experiment. A comprehensive program of this type for subsonic wings is presented in reference 4.\n\nAn investigation of the downwash behind a rectangular wing at a Mach number of 1.53 is reported in reference 5. Experimental trends of the variation of downwash angle with angle of attack at zero lift were similar to those predicted by linearized theory. The comparison between theory and experiment also indicated that the displacement of the vortex sheet and the resulting influence on the downwash distribution should be considered in calculating downwash angles at finite angles of attack. These results are analogous to those presented in reference 4 for the subsonic case and illustrate the interdependence of the downwash and the trailing vortex system.\n\nAn investigation was undertaken at the NACA Lewis laboratory to obtain downwash data in the immediate vicinity of the trailing vortex sheet, rather than in a probable tail-surface location, in order to provide data relating the shed vorticity and the downwash behind supersonic wings. The results for a trapezoidal wing are presented.\n\nSYMBOLS\n\nThe following symbols are used in this report:\n\n| Symbol | Description |\n| :--- | :--- |\n| $C_{p,B}$ | pressure coefficient, bottom surface of wing |\n| $C_{p,T}$ | pressure coefficient, top surface of wing |\n| c | wing chord (0.5 ft on model) |\n| H | pitot pressure in free stream |\n| $H_w$ | pitot pressure in wake |\n| $\\Delta H$ | $H - H_w$ |\n| M | Mach number |\n| m | slope of inclined vortex sheet (fig. 6) |\n| R | $$ \\frac{x - \\sqrt{x^2 - \\beta^2(y^2+z^2)}}{\\beta \\sqrt{y^2+z^2}} $$ |\n\nCONFIDENTIAL", "timestamp": "2026-07-22T07:03:28.753153+00:00"}
{"citation_id": "19930085983", "source_url": "https://ntrs.nasa.gov/api/citations/19930085983/downloads/19930085983.pdf", "page_number": 27, "total_pages": 46, "image_filename": "19930085983_p27.jpg", "text": "NACA RM A9I27\n\nCONFIDENTIAL\n\nLift coefficient, $C_L$\n\n$\\delta_u$, deg\n- 0\n- -5\n- -10\n- -15\n- -20\n- -25\n\nAngle of attack, $\\alpha$, deg\n\nfor $\\delta_u = 0^\\circ$\n\nPitching-moment coefficient, $C_m$\n\n(a) $C_L$ vs $\\alpha$, $C_L$ vs $C_m$.\n\nFigure 8.- The effect of elevon deflection on the aerodynamic characteristics of the wing-fuselage combination and on the elevon hinge-moment coefficients at a Mach number of 0.93.\n\nCONFIDENTIAL\n\n25", "timestamp": "2026-07-22T07:03:29.545967+00:00"}
{"citation_id": "19930085911", "source_url": "https://ntrs.nasa.gov/api/citations/19930085911/downloads/19930085911.pdf", "page_number": 50, "total_pages": 52, "image_filename": "19930085911_p50.jpg", "text": "```markdown\nCONFIDENTIAL\n\nNACA RM E9F22\n\nCONFIDENTIAL\n\n49\n\n[Figure: Diffuser total-pressure recovery as function of free-stream Mach number at various gas total-temperature ratios for ram-jet units 16-A-2, 16-A-3, 16-A-4, and 16-A-5.]\n\n| Diffuser total-pressure recovery, $P_4/P_0$ | Gas total-temperature ratio, $T_7/T_0$ |\n| :--- | :--- |\n| 1.0 | 4.4 4.9 5.7 6.6 6.5 5.7 |\n| .9 | 3.8 5.8 5.4 5.6 6.1 5.6 5.1 5.0 |\n| .8 | 1.6 1.2 2.3 2.7 3.9 4.2 4.5 3.5 4.5 3.6 4.0 4.3 4.5 4.6 4.0 |\n| .7 | 1.7 1.5 1.2 1.6 2.5 3.7 2.8 3.6 4.0 |\n| .6 | 1.2 1.9 2.5 3.0 |\n| .5 | 1.2 1.2 1.0 1.4 2.0 |\n| .4 | 1.2 1.0 |\n\n| Ram-jet unit | |\n| :--- | :--- |\n| $\\circ$ | A-2 |\n| $\\square$ | A-3 |\n| $\\diamond$ | A-4 |\n| $\\triangle$ | A-5 |\n\nFree-stream Mach number, $M_0$\n\nFigure 12. - Diffuser total-pressure recovery as function of free-stream Mach number at various gas total-temperature ratios for ram-jet units 16-A-2, 16-A-3, 16-A-4, and 16-A-5.\n\nNACA\n```", "timestamp": "2026-07-22T07:03:30.647876+00:00"}
{"citation_id": "19930085990", "source_url": "https://ntrs.nasa.gov/api/citations/19930085990/downloads/19930085990.pdf", "page_number": 25, "total_pages": 132, "image_filename": "19930085990_p25.jpg", "text": "NACA RM A9I01\n\nCONFIDENTIAL\n\n[Figure: A black and white photograph showing a semispan model of an airplane mounted in a wind tunnel. The model is positioned diagonally, with its nose pointing towards the lower right corner. The horizontal tail is visible, mounted in the extended wing-chord plane. The background shows the interior of the wind tunnel.]\n\n(a) Horizontal tail mounted in the extended wing-chord plane.\n\nFigure 2.- Semispan model of the airplane mounted in the Ames 12-foot pressure wind tunnel.\n\nCONFIDENTIAL\n\nA-12380\n\n23", "timestamp": "2026-07-22T07:03:32.576110+00:00"}
{"citation_id": "19930086083", "source_url": "https://ntrs.nasa.gov/api/citations/19930086083/downloads/19930086083.pdf", "page_number": 7, "total_pages": 48, "image_filename": "19930086083_p7.jpg", "text": "```markdown\nNACA RM L9F10\n5\n\n# TESTS\n\n## Test Conditions\n\nThe tests were made in the Langley 300 MPH 7- by 10-foot tunnel at dynamic pressures of approximately 25 and 100 pounds per square foot, corresponding to airspeeds of about 100 and 200 miles per hour. Reynolds numbers for these airspeeds, based on the mean aerodynamic chord (1.68 ft) of the $60^\\circ$ delta wing, were approximately $1.5 \\times 10^6$ and $3.0 \\times 10^6$, respectively. Corresponding Mach numbers were 0.13 and 0.27. The tests were run throughout a range of angles of attack of $0^\\circ$ to $40^\\circ$, and through a $-5^\\circ$ to $20^\\circ$ yaw range.\n\n## Corrections\n\nBlocking, jet-boundary, and air-stream-inclination corrections have been applied to the data. The jet-boundary corrections were obtained from methods outlined in reference 6. These corrections are strictly applicable only to wings of larger span and aspect ratio. However, it is believed that the error in using these corrections for the present setup is negligible. The effects of the support strut, which were determined by using an image system, have been subtracted from the data; no tares for the effects of yaw have been applied to the yaw data.\n\n## RESULTS AND DISCUSSION\n\nThe aerodynamic characteristics in pitch of the $60^\\circ$ delta-wing model with round and beveled leading-edge flaps are presented in figures 4 and 5, respectively; lift-drag ratios are presented in figures 6 and 7. The original data for $\\delta_n = 0^\\circ$ (fig. 4(a)) were rejected because of extremely large scatter of the test data resulting from temporary malfunction of the wind-tunnel scale system. The test was later rerun for only part of the angle-of-attack range as shown. Aerodynamic characteristics in yaw are given in figures 8 and 9, and lateral-stability parameter variations with lift coefficient are given in figures 10 and 11. Longitudinal and lateral-stability parameters determined at a Reynolds number of $3.0 \\times 10^6$ are presented in table I.\n```", "timestamp": "2026-07-22T07:03:33.140578+00:00"}
{"citation_id": "19930085951", "source_url": "https://ntrs.nasa.gov/api/citations/19930085951/downloads/19930085951.pdf", "page_number": 41, "total_pages": 92, "image_filename": "19930085951_p41.jpg", "text": "NACA RM L9D29\n39\n\n[Figure: Graph plotting Efficiency ($\\eta$) and Mach number ($M$) against Advance ratio ($J$). The graph contains four curves representing different $\\beta_{0.75R}$ values ($20^\\circ$, $25^\\circ$, $30^\\circ$, $35^\\circ$). Two additional lines indicate Helical-tip Mach number and Air-stream Mach number. A \"CONFIDENTIAL\" stamp is visible at the top.]\n\n(c) Efficiency.\nFigure 12.— Concluded. Rotational speed, 2000 rpm.\nCONFIDENTIAL", "timestamp": "2026-07-22T07:03:33.531777+00:00"}
{"citation_id": "19930085919", "source_url": "https://ntrs.nasa.gov/api/citations/19930085919/downloads/19930085919.pdf", "page_number": 44, "total_pages": 47, "image_filename": "19930085919_p44.jpg", "text": "NACA RM No. A9C21\nCONFIDENTIAL\n\nCONFIDENTIAL\n\nLift coefficient, $C_L$\n1.6\n1.2\n.8\n.4\n0\n-.4\n\nAngle of attack, $\\alpha$, deg\n-8\n0\n8\n16\n24\n32\n40\n\nPitching-moment coefficient, $C_m$\n.08\n0\n-.08\n\nSharp leading edge of\n50-percent span\n\nFull-span sharp leading edge\n\n$\\circ$ Plain wing + short fuselage\n\nNACA\n\n(a) $C_L$ vs $\\alpha$ and $C_m$.\n\nFigure 18- Effect of sharp leading edges on the lift, drag, and pitching-moment characteristics of\nthe model. $R, 4.2 \\times 10^6$.\n\n43", "timestamp": "2026-07-22T07:03:39.157083+00:00"}
{"citation_id": "19930085870", "source_url": "https://ntrs.nasa.gov/api/citations/19930085870/downloads/19930085870.pdf", "page_number": 64, "total_pages": 92, "image_filename": "19930085870_p64.jpg", "text": "NACA RM No. L9D07\n65\n\nCONFIDENTIAL\n\n.24\nRound L.E. {Circle: $C_l$, Square: $C_m$}\n.16\nSharp L.E. {Triangle: $C_l$, Diamond: $C_m$}\n.08\n$C_l$\n0\n-.08\n-.16\n-.24\n\n.01\n$C_m$\n0\n-.01\n\nRound L.E. {Circle: $C_D$, Square: $L/D$}\nSharp L.E. {Triangle: $C_D$, Diamond: $L/D$}\n\n.06\n.04\n$C_D$\n.02\n0\n\n12\n8\n$L/D$\n4\n0\n\n-8 -6 -4 -2 0 2 4 6 8\n$\\alpha$, deg\n\n[NACA logo]\n\n(b) Wing 2. $R(\\text{round L.E.}) = 850,000$.\nFigure 8 - Continued.\nCONFIDENTIAL", "timestamp": "2026-07-22T07:03:40.755013+00:00"}
{"citation_id": "19930085991", "source_url": "https://ntrs.nasa.gov/api/citations/19930085991/downloads/19930085991.pdf", "page_number": 19, "total_pages": 24, "image_filename": "19930085991_p19.jpg", "text": "NACA RM L9I28\n17\n\nArrangement a\n(triangular or\ntrapezoidal\nfins)\n\nArrangement b\n(triangular or\ntrapezoidal\nfins)\n120°\n\nArrangement\nc\n\n[Figure: Sketches illustrating various fin arrangements tested on the models in the Langley 20-foot free-spinning tunnel (models tested are not shown).]\n\nFigure 3.- Sketches illustrating various fin arrangements tested on the models in the Langley 20-foot free-spinning tunnel (models tested are not shown).", "timestamp": "2026-07-22T07:03:45.139973+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 110, "total_pages": 114, "image_filename": "19930086061_p110.jpg", "text": "```markdown\n106\n\n.1\n\n0\n\n-.1\n\nCm -.2\n\n-.3\n\n-.4\n\n0 .2 .4 .6 .8 1.0 1.2 1.4\n$C_L$\n\n| Planform | Wing | $R/10^6$ | A |\n| :--- | :--- | :--- | :--- |\n| [Symbol: Circle with horizontal line] | 1 | 0.57 | 3.46 |\n| [Symbol: Square] | 2 | 0.85 | 2.31 |\n| [Symbol: Diamond] | 3 | 1.14 | 1.73 |\n| [Symbol: Triangle] | | | |\n\nFull-scale wing of ref.3, $R \\approx 6.00 \\times 10^6$\n(comparable to wing 2).\n\nNACA\n\nFigure 55.- Variation of $C_m$ with $C_L$ for the three wings investigated and a comparison of the variation of $C_m$ with $C_L$ of wing 2 with that of the comparable full-scale wing of reference 1.\n\nNACA RM L9J07\n```", "timestamp": "2026-07-22T07:03:45.417286+00:00"}
{"citation_id": "19930085992", "source_url": "https://ntrs.nasa.gov/api/citations/19930085992/downloads/19930085992.pdf", "page_number": 15, "total_pages": 32, "image_filename": "19930085992_p15.jpg", "text": "```markdown\nTABLE I.- Concluded\n\n| Run | Weight | Spanwise position (in. from root) | q (lbs/sq ft) | $V_1$ (fps) | Mach number | Reynolds number (based on wing chord) | v (fps) | $\\rho$ (slugs/cu ft) | $f_{B1}$ (cps) | $f_{B2}$ (cps) | $f_t$ (cps) | $f_\\theta$ (cps) | $\\phi_{B1}$ | $\\phi_{B2}$ | $\\phi_\\alpha$ | Remarks |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| 12 | 3<br>3<br>3 | 16<br>16<br>16 | 195.1 | 407.0 | .3793 | $1.6469 \\times 10^6$ | 425.5 | 0.002176 | 4.90<br>---<br>4.90 | 30.6<br>---<br>29.6 | 51.6<br>---<br>51.2 | 38.6 | 0.0103<br>---<br>.0111 | 0.0036<br>---<br>.0045 | 0.0032<br>---<br>.0041 | Fluttered very violently with large amplitude |\n| 13 | 3<br>3<br>3 | 10<br>10<br>10 | 205.0 | 417.7<br>0 | .3893 | 1.6808 | 437.5<br>0 | --- | 5.07<br>.002199<br>4.98 | 30.0<br>---<br>30.0 | 54.7<br>---<br>54.7 | 38.3 | .0088<br>---<br>.0146 | .0036<br>---<br>.0040 | .0031<br>---<br>.0041 | Fluttered in second bending mode with large amplitude and node 8 to 10 inches from tip |\n| 14 | 3<br>3<br>3 | 38$\\frac{1}{2}$ (tip)<br>38$\\frac{1}{2}$ (tip)<br>38$\\frac{1}{2}$ (tip) | 146.8 | 349.5 | .3233 | 1.4170 | 363.4 | .002198 | 4.53<br>---<br>4.55 | 28.9<br>---<br>28.9 | 44.3<br>---<br>44.5 | 28.7 | .0097<br>---<br>.0084 | .0047<br>---<br>.0055 | .0024<br>---<br>.0051 | Violent flutter with large amplitude |\n| 15 | 3 | 35 | 144.7 | 350.1 | .3238 | 1.4172 | 364.1 | .002137 | --- | --- | --- | (a) | --- | --- | --- | Violent flutter with large amplitude |\n| 16 | None<br>None<br>None | ---<br>---<br>--- | 210.0 | 422.0<br>0 | .3970 | 1.6768<br>0 | 446.8<br>0 | .002120<br>4.88 | 4.93<br>---<br>4.88 | 30.6<br>---<br>30.7 | 56.9<br>---<br>57.2 | 34.7 | .0094<br>---<br>.0076 | .0052<br>---<br>.0043 | .0037<br>---<br>.0044 | Check on run 1; excellent, sustained flutter response in second bending mode with node 8 inches from tip |\n| 17 | 4<br>4<br>4 | 35$\\frac{1}{2}$ (tip)<br>35$\\frac{1}{2}$ (tip)<br>35$\\frac{1}{2}$ (tip) | 0<br>157.6<br>0 | 0<br>365.5<br>0 | 0<br>.3399<br>0 | 0<br>1.4650<br>0 | 0<br>383.0<br>0 | ---<br>.002166<br>--- | 4.53<br>---<br>4.49 | 28.5<br>---<br>28.5 | 44.4<br>---<br>44.1 | 30.4 | .0032<br>---<br>.0084 | .0066<br>---<br>.0053 | .0124<br>---<br>.0127 | Fluttered in second bending mode with 1-inch amplitude and node 8 inches from tip |\n| 18 | 4<br>4<br>4 | 35<br>35<br>35 | 0<br>150.2<br>0 | 0<br>357.5<br>0 | 0<br>.3305<br>0 | 0<br>1.4438<br>0 | 0<br>372.3<br>0 | ---<br>.002193<br>4.62 | 4.62<br>---<br>4.62 | 30.0<br>---<br>30.0 | 44.5<br>---<br>44.5 | 30.5 | .0089<br>---<br>.0091 | .0058<br>---<br>.0013 | .0179<br>---<br>.0165 | Good flutter response with 3-inch amplitude |\n| 19 | 4<br>4<br>4 | 31<br>31<br>31 | 0<br>154.7<br>0 | 0<br>362.0<br>0 | 0<br>.3360<br>0 | 0<br>1.4629<br>0 | 0<br>378.5<br>0 | ---<br>.002173<br>4.71 | 4.71<br>---<br>4.71 | 30.6<br>---<br>30.0 | 45.5<br>---<br>45.3 | 33.4 | .0114<br>---<br>.0097 | .0138<br>---<br>.0051 | .0146<br>---<br>.0120 | Good flutter response with 3-inch amplitude |\n| 20 | 4<br>4<br>4 | 24<br>24<br>24 | 0<br>186.5<br>0 | 0<br>390.2<br>0 | 0<br>.3630<br>0 | 0<br>1.5547<br>0 | 0<br>410.0<br>0 | ---<br>.002134<br>--- | 4.71<br>---<br>4.71 | 29.8<br>---<br>29.4 | 47.7<br>---<br>47.1 | 38.0 | .0089<br>---<br>.0113 | .0084<br>---<br>.0041 | .0094<br>---<br>.0232 | Fluttered in second bending mode with 3-inch amplitude with node 8 inches from tip |\n| 21 | 4<br>4<br>4 | 16<br>16<br>16 | 0<br>198.6<br>0 | 0<br>410.0<br>0 | 0<br>.3813<br>0 | 0<br>1.6286<br>0 | 0<br>430.9<br>0 | ---<br>---<br>--- | 4.87<br>---<br>4.87 | 29.2<br>---<br>29.0 | 50.8<br>---<br>50.5 | 38.9 | .0074<br>---<br>.0121 | .0116<br>---<br>.0072 | .0115<br>---<br>.0084 | Fluttered violently with large amplitude; strong divergence tendencies; nose section of weight loosened during flutter |\n| 22 | 4<br>4<br>4 | 10<br>10<br>10 | 0<br>204.6<br>0 | 0<br>416.0<br>0 | 0<br>.3875<br>0 | 0<br>1.6525<br>0 | 0<br>437.7<br>0 | ---<br>.002147<br>--- | 4.87<br>---<br>4.80 | 29.9<br>---<br>29.8 | 54.6<br>---<br>54.3 | Not clear | .0119<br>---<br>.0051 | .0047<br>---<br>.0045 | .0083<br>---<br>.0044 | Fluttered violently with large amplitude; strong divergence tendencies |\n\n$^a$No vibration records.\n\nNACA\n13\n```", "timestamp": "2026-07-22T07:03:46.056624+00:00"}