PDF/ocr_dataset_1031.jsonl

{"citation_id": "19930085544", "source_url": "https://ntrs.nasa.gov/api/citations/19930085544/downloads/19930085544.pdf", "page_number": 14, "total_pages": 33, "image_filename": "19930085544_p14.jpg", "text": "NACA RM No. L8K26\n\nbe in exact agreement. With this consideration in mind the agreement between experiment and theory is good.\n\nThe shift in the position of the maximum force between the experimental and calculated data noted particularly on figure 5 is due to several factors. First, the survey rake for the experimental data was 18 inches (0.375 propeller diameter) behind the center line of the propeller. Calculations showed that, in this distance, the twist of the propeller slipstream accounted for approximately $12^\\circ$ shift of the maximum force for the operating condition of $\\frac{V}{nD} = 1.2$ and $\\beta = 26^\\circ$ but was negligible for the operating condition of $\\frac{V}{nD} = 3.1$ and $\\beta = 53^\\circ$. Second, the unsteady flow on the blade sections causes a lag in the forces (about $5^\\circ$) which means that the maximum force does not occur on the horizontal as would be indicated from steady-flow calculations. This lag in the position of the maximum force is a function of the frequency of the oscillation and decreases with the propeller rotational speed, which means that it would decrease as the $V/nD$ is increased for constant forward speed. Third, the inclusion of the flapping of the blade section in the calculations causes an additional lag in the position of the maximum force (about $16^\\circ$).\n\nFigure 8 shows calculated differential thrust-coefficient curves of a two-blade propeller for three blade positions for the $26^\\circ$ blade-angle setting operating at a $V/nD$ of 1.2 and with the thrust axis inclined $4^\\circ$ to the air stream. The experimental curves for two positions for a $30^\\circ$ lag in phase angle are shown for comparison. The calculated distribution of thrust coefficient with radius is in good agreement with the experimental distribution.\n\nForces computed by oscillating-airfoil theory.- Figures 9 to 12, inclusive, show the results of some of the calculations made for an oscillating airfoil in a pulsating flow field in incompressible flow. The results for all the curves in these figures apply to any particular blade section. The total force on the propeller blade for any blade position is found by summing up the forces along the blade.\n\nThe variation of $L_c$ with $\\omega t$ is shown in figure 9 at $\\epsilon = 0.10$ for several values of $k$ with the steady part of the angle of attack $\\alpha$ equal to zero. In any particular case, the quantity $k$ is fixed by the operating conditions of the propeller and by the blade chord. Interpreted physically, the quantity $1/k$ is a measure of the wave length between successive waves in the vortex wake in terms of the half-chord length; in the steady-state calculations this wave length is arbitrarily assumed to be very large with respect to the chord. Thus, in figure 9 the curve of", "timestamp": "2026-07-22T05:12:44.055106+00:00"}
{"citation_id": "19930082498", "source_url": "https://ntrs.nasa.gov/api/citations/19930082498/downloads/19930082498.pdf", "page_number": 40, "total_pages": 49, "image_filename": "19930082498_p40.jpg", "text": "NACA TN No. 1838\n39\n\n[Figure: Top photograph showing a close-up of an aircraft engine cowling or nacelle, with visible exhaust pipes and mounting structure.]\n\nNACA\nL-52532\n\n[Figure: Bottom photograph showing an exposed radial aircraft engine mounted on a test stand outdoors, with exhaust pipes and wiring visible; background shows open land and water.]\n\nFigure 3.— Original engine-exhaust installation.\n\nNACA\nL-53877", "timestamp": "2026-07-22T05:12:44.581098+00:00"}
{"citation_id": "19930082496", "source_url": "https://ntrs.nasa.gov/api/citations/19930082496/downloads/19930082496.pdf", "page_number": 44, "total_pages": 50, "image_filename": "19930082496_p44.jpg", "text": "NACA TN No. 1836\n43\n\n[Figure: A close-up photograph of a damaged mechanical component, likely a turbine disk with dovetail slots. A ruler marked \"INCHES\" is placed across the top for scale. In the bottom right corner of the image, there is a NACA logo with the text \"C-21269\" and \"4-22-48\".]\n\n(a) Close-up of failure after 12 hours and 13 minutes of operation.\nFigure 12. - Disk-dovetail failure.", "timestamp": "2026-07-22T05:12:44.773106+00:00"}
{"citation_id": "19930085879", "source_url": "https://ntrs.nasa.gov/api/citations/19930085879/downloads/19930085879.pdf", "page_number": 9, "total_pages": 29, "image_filename": "19930085879_p9.jpg", "text": "NACA RM L9D11\n7\n\nREFERENCES\n\n1. Marzella, J., and Hess, J. R.: Ground and Flight Tests of Martin-Baker Aircraft Company Pilot's Ejection Seat from Model JD-1 Airplane. Rep. ASL NAM 256005.1, Part I, Naval Air Exp. Sta., Bur. Aero., Feb. 27, 1947.\n\n2. Armstrong, Harry G., and Heim, J. W.: The Effect of Acceleration on the Living Organism. ACTR 4362, Materiel Div., Army Air Corps. Dec. 1, 1937.\n\n3. Scher, Stanley H.: Wind-Tunnel Investigation of the Stability of Jettisoned Nose Sections of the D-558 Airplane - Phases I and II. NACA RM L7K10, 1948.\n\n4. Angle, Ellwyn E.: Initial Flight Test of the NACA FR-1-A, a Low-Acceleration Rocket-Propelled Vehicle for Transonic Flutter Research. NACA RM L7J08, 1948.", "timestamp": "2026-07-22T05:12:47.188890+00:00"}
{"citation_id": "19930082542", "source_url": "https://ntrs.nasa.gov/api/citations/19930082542/downloads/19930082542.pdf", "page_number": 25, "total_pages": 53, "image_filename": "19930082542_p25.jpg", "text": "24\nNACA TN No. 1867\n\nTABLE II.--RUPTURE TEST CHARACTERISTICS AT 1200° F OF LOW-CARBON M-155 BAR STOCK - Continued\n\n<!-- Table (88, 118, 893, 905) -->\n\\begin{tabular}{|c|c|c|c|c|c|c|c|c|c|c|c|c|}\n\\hline\n\\multicolumn{5}{|c|}{Heat treatment} & \\multicolumn{2}{c|}{Hot-cold-rolling} & \\multicolumn{6}{c|}{Rupture properties at 1200° F} \\\\\n\\cline{1-13}\n\\multicolumn{3}{|c|}{Solution treatment} & \\multicolumn{2}{c|}{Aging treatment (a)} & \\multicolumn{2}{c|}{(b)} & \\multicolumn{6}{c|}{} \\\\\n\\cline{1-13}\nTemper- & Time & Method & Temper- & Time & Temper- & Percent & Stress & Rupture & Elongation & Reduction & \\multicolumn{2}{c|}{Rupture strength} \\\\\nature & (hr) & of & ature & (hr) & ature & reduction & (psi) & time & in 1 in. & of area & \\multicolumn{2}{c|}{(psi)} \\\\\n($^\\circ$F) & & cooling & ($^\\circ$F) & & ($^\\circ$F) & & & (hr) & (percent) & (percent) & 100 hr & 1000 hr \\\\\n& & (c) & & & & & & & & & & \\\\\n\\hline\n\\multicolumn{13}{|c|}{Solution-treated at 2050° F} \\\\\n\\hline\n\\multicolumn{5}{|c|}{Time and cooling rate} & \\multicolumn{8}{c|}{} \\\\\n\\hline\n2050 & 1 & W.Q. & ---- & -- & ---- & ---- & 45,000 & 25 & $^e$1 & 17.8 & 45,000 & 40,000 \\\\\n& & & & & & & 44,000 & 420 & 18 & 22.6 & & \\\\\n& & & & & & & 43,000 & 619 & 24 & 26.7 & & \\\\\n& & & & & & & 40,000 & 1138 & 17 & 15.6 & & \\\\\n\\hline\n2050 & 2 & W.Q. & ---- & -- & ---- & ---- & 48,000 & 33 & $^e$3 & 17.3 & 45,500 & 39,000 \\\\\n& & & & & & & 45,000 & 180 & 10 & 18.9 & & \\\\\n& & & & & & & 40,000 & 687 & 19 & 23.3 & & \\\\\n\\hline\n2050 & 2 & A.C. & ---- & -- & ---- & ---- & 45,000 & 35 & $^e$1.5 & 14.4 & 43,000 & 39,000 \\\\\n& & & & & & & 42,000 & 119 & 5 & 7.3 & & \\\\\n& & & & & & & 40,000 & 614 & $^e$8.5 & 17.8 & & \\\\\n\\hline\n2050 & 5 & W.Q. & ---- & -- & ---- & ---- & 51,000 & 62 & (a) & 8.0 & 49,000 & 40,500 \\\\\n& & & & & & & 48,000 & 181 & 12 & 12.5 & & \\\\\n& & & & & & & 45,000 & 141 & 10 & 10.9 & & \\\\\n& & & & & & & 40,000 & 1138 & 21 & 25.6 & & \\\\\n\\hline\n\\multicolumn{5}{|c|}{Aging time and temperature} & \\multicolumn{8}{c|}{} \\\\\n\\hline\n2050 & 2 & W.Q. & 1400 & 2 & ---- & ---- & 54,000 & 35 & 21 & 14.4 & 49,000 & 39,000 \\\\\n& & & & & & & 50,000 & 116 & 11 & 18.9 & & \\\\\n& & & & & & & 45,000 & 163 & 12 & 10.9 & & \\\\\n& & & & & & & 40,000 & 885 & 17 & 20.4 & & \\\\\n\\hline\n2050 & 2 & W.Q. & 1400 & 8 & ---- & ---- & 50,000 & 118 & 28 & 24.5 & 51,000 & 40,000 \\\\\n& & & & & & & 45,000 & 310 & 24 & 23.3 & & \\\\\n& & & & & & & 40,000 & 954 & 29 & 35.8 & & \\\\\n\\hline\n2050 & 2 & W.Q. & 1400 & 16 & ---- & ---- & 50,000 & 97 & 29 & 34.0 & 50,000 & 39,500 \\\\\n& & & & & & & 45,000 & 300 & 37 & 40.8 & & \\\\\n& & & & & & & 40,000 & 860 & 33 & 41.2 & & \\\\\n\\hline\n2050 & 2 & W.Q. & 1400 & 24 & ---- & ---- & 55,000 & 31 & 24 & 21.2 & 50,500 & 35,000 \\\\\n& & & & & & & 50,000 & 123 & 38 & 33.0 & & \\\\\n& & & & & & & 45,000 & 205 & 37 & 36.9 & & \\\\\n& & & & & & & 40,000 & 436 & 44 & 44.7 & & \\\\\n\\hline\n2050 & 2 & W.Q. & 1400 & 50 & ---- & ---- & 50,000 & 75 & 9 & 38.8 & 48,000 & 36,500 \\\\\n& & & & & & & 45,000 & 300 & 33 & 37.9 & & \\\\\n& & & & & & & 40,000 & 885 & 37 & 40.8 & & \\\\\n& & & & & & & 40,000 & 423 & 5 & 25.6 & & \\\\\n\\hline\n2050 & 2 & W.Q. & 1350 & 24 & ---- & ---- & 50,000 & 145 & 13 & 12.5 & 52,000 & 40,000 \\\\\n& & & & & & & 45,000 & 343 & 12 & 21.0 & & \\\\\n& & & & & & & 40,000 & 1078 & 20 & 29.7 & & \\\\\n\\hline\n2050 & 2 & W.Q. & 1500 & 24 & ---- & ---- & 50,000 & 87 & 40 & 36.9 & 48,000 & 36,000 \\\\\n& & & & & & & 45,000 & 122 & 24 & 35 & & \\\\\n& & & & & & & 40,000 & 462 & 37 & 48.3 & & \\\\\n\\hline\n2050 & 2 & W.Q. & 1600 & 24 & ---- & ---- & 50,000 & 74 & 40 & 41.8 & 47,500 & $^f$35,000 \\\\\n& & & & & & & 45,000 & 132 & 31 & 43.7 & & \\\\\n& & & & & & & 40,000 & 361 & 38.5 & 45.8 & & \\\\\n\\hline\n2050 & 2 & W.Q. & 1750 & 24 & ---- & ---- & 50,000 & 121 & 33 & 29.8 & 51,500 & 38,000 \\\\\n& & & & & & & 45,000 & 306 & 32 & 52.8 & & \\\\\n& & & & & & & 39,000 & 659 & 36 & 53 & & \\\\\n\\hline\n\\end{tabular}\n\n$^a$All aging treatments preceded hot-cold-rolling except where noted.\n$^b$All hot-cold-rolled material was given a final stress relief at 1200° F for 1 hr.\n$^c$W.Q., water-quenched; A.C., air-cooled.\n$^d$Fractured in gage mark.\n$^e$Estimated.\n\nNACA", "timestamp": "2026-07-22T05:12:51.392624+00:00"}
{"citation_id": "19930082245", "source_url": "https://ntrs.nasa.gov/api/citations/19930082245/downloads/19930082245.pdf", "page_number": 56, "total_pages": 66, "image_filename": "19930082245_p56.jpg", "text": "```markdown\nNACA TN No. 1596\n\n1.6\n1.4\n1.2\n1.0\n.8\n.6\n.4\n.2\n0\n-.2\n-.4\n-.6\n.1 .2 .3 .4 .5 .6 .7 .8 .9\nMach number, M\nAileron section normal-force coefficient, $C_{n\\delta}$\n$\\delta_a$ (deg)\n30\n18\n12\n6\n4\n2\n0\n-2\n-4\n-6\n-12\n\n.16\n.12\n.08\n.04\n0\n-.04\n-.08\n-.12\n-.16\n-.20\n-.24\n-.28\n.1 .2 .3 .4 .5 .6 .7 .8 .9\nMach number, M\nAileron section hinge-moment coefficient, $C_h$\n$\\delta_a$ (deg)\n12\n4\n2\n0\n-2\n-4\n-6\n-12\n-18\n-30\n\n(f) $C_n = 0.6$.\nFigure 10.—Continued.\n\n[Figure: NACA logo]\n\n55\n```", "timestamp": "2026-07-22T05:12:54.746078+00:00"}
{"citation_id": "19930085626", "source_url": "https://ntrs.nasa.gov/api/citations/19930085626/downloads/19930085626.pdf", "page_number": 12, "total_pages": 24, "image_filename": "19930085626_p12.jpg", "text": "NACA RM No. L8K23\n11\n\nCONFIDENTIAL\n\n[Figure: A photograph of a rocket model with extended chord ailerons. A scale in inches is visible at the bottom. A label on the image reads \"NACA L-56017\".]\n\n(b) Extended chord aileron configuration.\nFigure 2.— Continued.\nCONFIDENTIAL", "timestamp": "2026-07-22T05:12:54.940460+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 6, "total_pages": 96, "image_filename": "19930085880_p6.jpg", "text": "```markdown\n4\nNACA RM No. L9C03\n\ngiven in figure 8(a). The area involved for model 250D varied appreciably only with wetted length (or wetted area) and is given in figure 8(b) plotted against total wetted area.\n\n### Aerodynamic Tares\n\nThe aerodynamic tares for resistance, moment, and lift were determined for all trims. The aerodynamic drag and moment were determined with the model attached to the towing gear, with the model removed, and with the model and strut structure removed (towing staff alone). When the model was attached it was positioned approximately one-half inch above the water. When the model was removed the position of the staff and struts was the same as when the model was attached. With the strut structure also removed, the drag was measured at positions of the staff to cover the range for the drafts obtained in the tests. The aerodynamic moment tares were found to be negligible for all models.\n\nThe aerodynamic drag of the gear one-half inch above the zero draft position but with the model removed was the same for all models and did not vary appreciably with trim. This drag, plus the increments due to change in draft (as determined from the runs made with the staff alone) is given in figure 9.\n\nThe aerodynamic drag of the gear one-half inch above the zero draft position and with the model attached was found to be the same for models 250A and 250B. Its variation with trim was negligible for these models but not for model 250D. The difference in the drags with and without the model attached was considered to be the aerodynamic drag of the model alone. The aerodynamic drag correction for the model alone was assumed to be equal to the ratio of the unwetted plan-form area to the total plan-form area multiplied by the total aerodynamic drag of the model alone. This correction is given in figure 10 for models 250A and 250B and in figure 11 for model 250D. These values, in addition to those for the gear alone, were subtracted as tare corrections from all the resistance data before plotting.\n\nThe aerodynamic lift was determined by counterbalancing the model in the air at zero speed, then running at the desired speeds and trims and adding weight until the model moved downward, then removing the weights until the model moved upward. The average of these two weight limits was considered to be the lift. This lift, which varied appreciably with trim, was the same for models 250A and 250B but different for model 250D. The aerodynamic lift correction was also determined as a function of unwetted area in the same manner as for the aerodynamic drag; it is given in figure 12 for models 250A and 250B and in figure 13 for model 250D. The values given were subtracted as tare corrections from the values of load applied to the models.\n```", "timestamp": "2026-07-22T05:12:55.594924+00:00"}
{"citation_id": "19930085542", "source_url": "https://ntrs.nasa.gov/api/citations/19930085542/downloads/19930085542.pdf", "page_number": 17, "total_pages": 46, "image_filename": "19930085542_p17.jpg", "text": "NACA RM No. L8L29\n15\n\n<!-- Image (157, 102, 852, 868) -->\n\nFigure 1.- System of stability axes. Positive forces, moments, and angles are indicated.", "timestamp": "2026-07-22T05:12:56.935070+00:00"}
{"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 30, "total_pages": 78, "image_filename": "19930082618_p30.jpg", "text": "```markdown\n28\n\nNACA TN 1945\n\n<!-- Image (85, 103, 893, 791) -->\n\n(a) Section lift and pitching-moment characteristics of the plain airfoil section.\nFigure 3.- Aerodynamic characteristics of the NACA 64$_2$-415 airfoil section, 24-inch chord.\n```", "timestamp": "2026-07-22T05:13:08.512459+00:00"}
{"citation_id": "19930085519", "source_url": "https://ntrs.nasa.gov/api/citations/19930085519/downloads/19930085519.pdf", "page_number": 28, "total_pages": 46, "image_filename": "19930085519_p28.jpg", "text": "NACA RM No. L8K19\n\nPlane of\nSymmetry\n\n.50 b/2\n\n90°\n\nA\nA\n\nchord\nline\n\n60°\n\n0.20c'\n\nSection A-A\n\nNACA\n\nFigure 7.- Details of the half-span Zap flap tested on the 42° sweptback wing.\n\n27", "timestamp": "2026-07-22T05:13:15.927298+00:00"}
{"citation_id": "19930082914", "source_url": "https://ntrs.nasa.gov/api/citations/19930082914/downloads/19930082914.pdf", "page_number": 33, "total_pages": 66, "image_filename": "19930082914_p33.jpg", "text": "32\n\nPage intentionally left blank\n\nPage intentionally left blank", "timestamp": "2026-07-22T05:13:16.117064+00:00"}
{"citation_id": "19930085487", "source_url": "https://ntrs.nasa.gov/api/citations/19930085487/downloads/19930085487.pdf", "page_number": 31, "total_pages": 36, "image_filename": "19930085487_p31.jpg", "text": "NACA RM No. EBJ22\n29\n\nFrequency, cps\n\n100,000\n\n10,000\n\nRear stator row, 62 blades\nFront stator row, 40 blades\n\nFirst torsional mode\nSecond bending mode\n\n1,000\n\nFirst bending mode\n\nSecond order, split compressor case\n\nFirst order, rotor speed\n\n100\n\n4 6 8 10 12 14 16x10³\nRotor speed, rpm\n\n(e) Fifth stage.\n\nFigure 9. - Continued. Critical-speed diagrams for 10 stages of\ncompressor rotor.", "timestamp": "2026-07-22T05:13:18.395655+00:00"}
{"citation_id": "19930085889", "source_url": "https://ntrs.nasa.gov/api/citations/19930085889/downloads/19930085889.pdf", "page_number": 3, "total_pages": 37, "image_filename": "19930085889_p3.jpg", "text": "2\nCONFIDENTIAL\nNACA RM L9F14\n\nThe derivative of yawing moment caused by rolling was either zero\nor positive through most of the lift-coefficient range for each of the\nwings tested. At zero lift coefficient there was a decrease of the\ndamping in roll with an increase of sweepback. The values of damping in\nroll obtained in the rolling-flow test section of the Langley stability\ntunnel show good agreement with those obtained by free rotation of the\nmodels in the Langley 7- by 10-foot tunnel and with the values\ncalculated by Weissinger's theory.\n\nAn increase in sweepback caused large reductions in the rolling-\nmoment coefficient and in the wing-tip helix angle resulting from a unit\nangular deflection of the ailerons about their hinge axis.\n\nINTRODUCTION\n\nThe influence of a number of different geometric parameters on the\nrolling stability derivatives of wings have been investigated in the\nLangley stability tunnel by means of the rolling-flow technique. (See\nreference 1.) The investigations have included the effects of aspect\nratio and sweep (reference 2), taper ratio (reference 3), dihedral\n(reference 4), and airfoil section (reference 5). All of the investi-\ngations were performed at low Mach numbers and with moderately thick\nwings. In order to obtain an indication of the rolling characteristics\nof sweptback wings at higher subsonic speeds, a series of thin wings\n(NACA 65A006 airfoil section) were tested in the Langley high-speed\n7- by 10-foot wind tunnel at Mach numbers from about 0.4 to about 0.9.\n(See reference 6.) Results were obtained over an angle-of-attack range\nfrom $0.3^{\\circ}$ to $6.5^{\\circ}$ for the damping-in-roll derivative $C_{l_p}$ and for the\naileron effectiveness.\n\nThe results of the investigation reported herein were obtained in\nthe rolling-flow test section of the Langley stability tunnel, and the\nmodels were those used for the investigation reported in reference 6.\nThe purpose of the present tests was to obtain more complete information,\nat least at low speeds, on the static and rolling characteristics of\nthe wings and also to obtain a correlation between techniques of the\nLangley stability tunnel (rolling flow) and the Langley 7- by 10-foot\ntunnel (free rotation) for determining the damping in roll.\n\nThe wings tested were sweptback $3.6^{\\circ}$, $32.6^{\\circ}$, and $46.7^{\\circ}$ and had an\naspect ratio of 4 and taper ratio of 0.6.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:13:21.024387+00:00"}
{"citation_id": "19930085881", "source_url": "https://ntrs.nasa.gov/api/citations/19930085881/downloads/19930085881.pdf", "page_number": 5, "total_pages": 31, "image_filename": "19930085881_p5.jpg", "text": "NACA RM L9D12 CONFIDENTIAL 3\n\ncharacteristics, the interference effects between the wings were negligible. Several test vehicles, differing nominally only in the aileron deflection, of each configuration were tested. The models were finished with clear lacquer and were smooth and fair.\n\nThe wings were constructed of laminated wood and were stiffened by means of steel plates cycle-welded into the upper and lower wing surfaces as shown in figure 1. The outer surfaces of the steel plates were covered with a wood veneer to facilitate construction. In order to establish that this construction possessed sufficient stiffness to minimize the effects of wing twisting, several test vehicles of configurations 50 ($\\Lambda = 0^\\circ$) and 53 ($\\Lambda = 45^\\circ$) (see table II) having reduced wing stiffness were constructed. The reduction in wing stiffness was obtained by reducing the thickness of the steel stiffening plates and by using plates of duralumin of reduced thickness. In this way, the wing stiffness was progressively reduced to about one-third of that of the wings employed in the present tests. The results of the flight tests showed no measurable variation of rolling effectiveness with wing stiffness for the range of stiffness values investigated indicating that the effects of wing twisting on the present experimental results are negligible.\n\nThe wings were constructed with the aft 20 percent of the airfoil sections deflected to the desired settings. This method of construction simulated plain, sealed, full-span ailerons. The measured values of aileron deflection and wing incidence are estimated to be within $\\pm 0.1^\\circ$ and $\\pm 0.05^\\circ$, respectively, of the actual values.\n\nThe test vehicles were propelled by a two-stage rocket-propulsion system to a Mach number of about 1.9. During coasting flight following burnout of the rocket motor, time histories of the rolling velocity produced by the ailerons (obtained by means of a small radio transmitter designated \"spinsonde\" contained in the nose of the test vehicles) and the flight-path velocity (obtained by Doppler radar) were recorded. These data, in conjunction with atmospheric data obtained by means of radiosonde, permitted the evaluation of the wing-aileron rolling-effectiveness parameter $\\frac{pb}{2V}$ as a function of Mach number. The drag coefficient was also obtained by a process involving a graphical differentiation of the flight-path velocity-time relation. The scale of the tests is indicated by the curves of Reynolds number versus Mach number in figure 3. A more complete description of the technique is given in references 1 and 2.\n\nACCURACY\n\nThe accuracy of the test results is estimated to be within the following limits:\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:13:23.489178+00:00"}
{"citation_id": "19930082498", "source_url": "https://ntrs.nasa.gov/api/citations/19930082498/downloads/19930082498.pdf", "page_number": 41, "total_pages": 49, "image_filename": "19930082498_p41.jpg", "text": "Page intentionally left blank\n\nPage intentionally left blank", "timestamp": "2026-07-22T05:13:26.959276+00:00"}
{"citation_id": "19930082496", "source_url": "https://ntrs.nasa.gov/api/citations/19930082496/downloads/19930082496.pdf", "page_number": 45, "total_pages": 50, "image_filename": "19930082496_p45.jpg", "text": "44\n\nPage intentionally left blank\n\nPage intentionally left blank", "timestamp": "2026-07-22T05:13:29.258992+00:00"}
{"citation_id": "19930082614", "source_url": "https://ntrs.nasa.gov/api/citations/19930082614/downloads/19930082614.pdf", "page_number": 31, "total_pages": 36, "image_filename": "19930082614_p31.jpg", "text": "NACA TN 1939\n29\n\n<!-- Image (222, 110, 827, 877) -->\n\n<!-- Table (227, 643, 578, 869) -->\n\\begin{tabular}{|l|l|l|l|l|l|l|l|}\n\\hline\nBrake & \\multicolumn{2}{l|}{Chordwise} & Brake chord & \\multicolumn{2}{l|}{Gap width} & \\multicolumn{2}{l|}{Brake span} \\\\\n\\cline{2-3}\\cline{5-6}\\cline{7-8}\n& \\multicolumn{2}{l|}{location} & & \\multicolumn{2}{l|}{} & \\multicolumn{2}{l|}{semispan} \\\\\n\\cline{2-3}\\cline{5-6}\\cline{7-8}\n& $d_u$ & $d_l$ & $c_B$ & $c_{B_l}$ & $e_u$ & $e_l$ & upper & lower \\\\\n\\hline\nF & 0.16c & 0.16c & 0.043$c_B$ & 0.058$c_B$ & 0 & 0 & 0.128 & 0.128 \\\\\n\\hline\nG & None & .16c & None & .033$c_B$ & 0 & & & .244 \\\\\n& \" & .16c & \" & .071$c_B$ & 0.25$c_B$ & & & .244 \\\\\n& \" & .16c & \" & .106$c_B$ & .50$c_B$ & & & .244 \\\\\n& \" & .30c & \" & .072$c_B$ & 0 & & & .350 \\\\\n& \" & .30c & \" & .096$c_B$ & .25$c_B$ & & & .350 \\\\\n& \" & .30c & \" & .145$c_B$ & .50$c_B$ & & & .350 \\\\\n\\hline\nH & None & .20c & None & .067$c_B$ & 0 & & & 1.000 \\\\\n\\hline\nI & 0.20c & None & 0.165c & & 0.33$c_B$ & & 0.202 & \\\\\n& None & 0.20c & & 0.165c & & 0.33$c_B$ & & 0.202 \\\\\n& 0.50c & None & 0.165c & & 0.33$c_B$ & & 0.202 & \\\\\n& None & 0.50c & & 0.165c & & 0.33$c_B$ & & 0.202 \\\\\n\\hline\nJ & 0.75c & None & 0.040c & & 0 & & 0.409 & \\\\\n\\hline\nK & .63c & None & .250c & & 0.16$c_B$ & & .479 & \\\\\n\\hline\nL & .36c & None & .092c & & .34$c_B$ & & .143 & \\\\\n& None & 0.56c & & 0.092c & & 0.32$c_B$ & & 0.143 \\\\\n\\hline\nM & 0.36c & None & 0.092c & & 0.34$c_B$ & & 0.143 & \\\\\n& None & 0.56c & & 0.092c & & 0.32$c_B$ & & 0.143 \\\\\n\\hline\nN & --- & --- & --- & --- & --- & --- & 0.163 & .163 \\\\\n\\hline\n\\end{tabular}\n\n(b) Spoiler- and picket-fence-type air brakes.\nFigure 6. — Continued.", "timestamp": "2026-07-22T05:13:32.005946+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 46, "total_pages": 99, "image_filename": "19930082511_p46.jpg", "text": "44\nNACA TN No. 1826\n\nSince $g_m^{(J)}(a)$ thereby equals zero, condition (4) is automatically satisfied. Values of $g_m^{(J)}(b)$ are here assumed to be finite, instead of infinite (see part I); the corresponding inaccuracy, however, is considered to exist mainly in the immediate region of the exit lip ($\\xi=b, \\rho=1$).\n\nSubstituting this series in equation (26) gives\n\n$$\n\\phi_A = \\sum_{m=1}^{\\infty} \\sum_{n=0}^{\\infty} \\left[ h_{mn}^{(1)} P_{mn}(\\xi, \\rho) \\sin m\\theta + h_{mn}^{(2)} P_{mn}(\\xi, \\rho) \\cos m\\theta \\right] \\quad (28)\n$$\n\nwhere\n\n$$\nP_{m0}(\\xi, \\rho) = \\frac{1}{\\pi} \\int_0^{\\infty} \\frac{J_m(i\\rho q)}{iq J_m'(iq)} dq \\int_a^b \\sin \\frac{\\pi}{2} \\frac{\\beta - a}{b - a} \\cos q(\\beta - \\xi) d\\beta\n$$\n\nand, for $n \\neq 0$,\n\n$$\nP_{mn}(\\xi, \\rho) = \\frac{1}{\\pi} \\int_0^{\\infty} \\frac{J_m(i\\rho q)}{iq J_m'(iq)} dq \\int_a^b \\sin n\\pi \\frac{\\beta - a}{b - a} \\cos q(\\beta - \\xi) d\\beta\n$$\n\nIn the evaluation of these two expressions, the inner integrals may be found analytically and the outer integrals, which converge rapidly, may be found numerically. It is possible, however, by means of contour integration similar to that discussed in reference 14, to transform the infinite integrals into infinite series that are more convenient for the present purpose. The contour integration and the resulting infinite series are given in appendix A.\n\nDifferentiating equation (28) with respect to $\\xi$ and taking $\\rho = 1$ gives\n\n$$\n\\left. \\frac{\\partial \\phi_A}{\\partial \\xi} \\right|_{\\rho=1} = \\sum_{m=1}^{\\infty} \\sum_{n=0}^{\\infty} \\left[ h_{mn}^{(1)} \\sin m\\theta \\frac{\\partial P_{mn}(\\xi, 1)}{\\partial \\xi} + h_{mn}^{(2)} \\cos m\\theta \\frac{\\partial P_{mn}(\\xi, 1)}{\\partial \\xi} \\right] \\quad (29)\n$$", "timestamp": "2026-07-22T05:13:33.717756+00:00"}
{"citation_id": "19930085626", "source_url": "https://ntrs.nasa.gov/api/citations/19930085626/downloads/19930085626.pdf", "page_number": 13, "total_pages": 24, "image_filename": "19930085626_p13.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:13:37.880012+00:00"}
{"citation_id": "19930085542", "source_url": "https://ntrs.nasa.gov/api/citations/19930085542/downloads/19930085542.pdf", "page_number": 18, "total_pages": 46, "image_filename": "19930085542_p18.jpg", "text": "```markdown\n16\n\n$\\frac{5}{8}\" R.$\n$\\frac{3}{4}\"$\n$5.0\"$\n$\\frac{1}{16}$\n$31.23\"$\n$7.8^\\circ$\nFLAT PLATE\n\n$31.60\"$\nBICONVEX (12% THICK)\n\n$31.60\"$\nNACA 0012\n\n[Figure: NACA logo]\n\nFigure 2.- Root-chord profiles of $60^\\circ$ triangular wings.\n\nNACA RM No. L5E29\n```", "timestamp": "2026-07-22T05:13:41.503971+00:00"}
{"citation_id": "19930082485", "source_url": "https://ntrs.nasa.gov/api/citations/19930082485/downloads/19930082485.pdf", "page_number": 52, "total_pages": 62, "image_filename": "19930082485_p52.jpg", "text": "NACA TN No. 1810\n51\n\n1026\n\n1134-1447\n\n[Figure: Diagram showing three airfoil-related shapes. The top shape is a curved airfoil profile with a series of inscribed circles along its length. The middle shape is a simple curved line representing a camber line. The bottom shape is a symmetrical airfoil profile, also with a series of inscribed circles. A vertical axis labeled 'x' passes through the maximum thickness point of the profiles. At the trailing edge of the bottom airfoil, an angle is indicated as $12^\\circ$-$30'$ included angle.]\n\nNACA\n\nFigure 20.- Distribution of symmetrical airfoil on assumed camber lines.", "timestamp": "2026-07-22T05:13:51.588990+00:00"}
{"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 31, "total_pages": 78, "image_filename": "19930082618_p31.jpg", "text": "NACA TN 1945\n\nSection lift coefficient, $c_l$\nMoment coefficient, $c_{m_{c/4}}$\nSection angle of attack, $\\alpha_0$, deg\n\nR\n$\\circ$ 0.7 x $10^6$\n$\\square$ 1.0\n$\\diamond$ 1.5\n$\\triangle$ 2.0\n$\\nabla$ 6.0\nFlagged symbols denote\nstandard roughness\n\nNACA\n\n(b) Section lift and pitching-moment characteristics of the NACA 64$_2$-415 airfoil section with a\n0.20c simulated split flap deflected 60°.\nFigure 3.— Continued.\n\n29", "timestamp": "2026-07-22T05:13:55.594828+00:00"}
{"citation_id": "19930085869", "source_url": "https://ntrs.nasa.gov/api/citations/19930085869/downloads/19930085869.pdf", "page_number": 11, "total_pages": 36, "image_filename": "19930085869_p11.jpg", "text": "CONFIDENTIAL\nTABLE I\nOFFSETS OF LANGLEY TANK MODEL 237-GB\n[All dimensions are in inches]\n\nOffsets model scale 1/16\n\n| Station | Distance to F.P. | Keel above base line | Chine above base line | Half breadth of chine | Height of hull at center line | Radius of boom | Line of centers above base line | 1-inch buttock | | 2-inch buttock | | 1-inch water line | 2-inch water line | 3-inch water line | 4-inch water line | 5-inch water line | 6-inch water line | 7-inch water line | 8-inch water line | 9-inch water line | 10-inch water line | 11-inch water line | 12-inch water line |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| | | | | | | | | Upper | Lower | Upper | Lower | | | | | | | | | | | | |\n| F.P. | 0 | 2.80 | 2.80 | 0 | 2.80 | | | 2.89 | 2.56 | | | 2.26 | 0.92 | | | | | | | | | |\n| 1 | 1.66 | 1.80 | 2.64 | 1.08 | 3.76 | | | 4.34 | 1.72 | | | 1.64 | 1.51 | 1.15 | 0.58 | | | | | | | |\n| 2 | 4.31 | .76 | 2.38 | 1.64 | 5.28 | | | 5.86 | 1.17 | 2.37 | 2.05 | 1.89 | 1.88 | 1.64 | 1.36 | 0.96 | | | | | | |\n| 3 | 6.96 | .20 | 2.11 | 2.01 | 6.84 | | | 7.44 | .82 | 3.91 | 1.67 | 2.24 | 2.16 | 1.99 | 1.79 | 1.52 | 1.20 | 0.65 | | | | |\n| 4 | 9.62 | 0 | 1.84 | 2.24 | 8.35 | | | 8.78 | .64 | 5.48 | 1.36 | 2.35 | 2.30 | 2.20 | 2.08 | 1.88 | 1.66 | 1.36 | .86 | | | |\n| 5 | 12.27 | 0 | 1.60 | 2.36 | 9.72 | | | 9.92 | .56 | 7.00 | 1.15 | 2.40 | 2.39 | 2.33 | 2.26 | 2.14 | 2.00 | 1.78 | 1.46 | .91 | | |\n| 6 | 14.95 | 0 | 1.34 | 2.41 | 10.80 | | | 10.72 | .49 | 8.62 | 1.00 | 2.37 | 2.41 | 2.39 | 2.34 | 2.26 | 2.20 | 2.06 | 1.82 | 1.44 | .80 | |\n| 7 | 17.18 | 0 | 1.14 | 2.34 | 11.52 | | | 11.53 | .40 | 9.58 | .76 | 2.15 | 2.28 | 2.36 | 2.41 | 2.40 | 2.36 | 2.32 | 2.24 | 2.10 | 1.83 | 1.36 | .48 |\n| 8 | 20.25 | 0 | .82 | 2.12 | 12.18 | | | 12.04 | .24 | 10.32 | 1.73 | 1.85 | 2.04 | 2.20 | 2.28 | 2.36 | 2.40 | 2.40 | 2.36 | 2.31 | 2.08 | 1.70 | 1.04 |\n| 9 | 22.91 | 0 | .58 | 1.73 | 12.56 | | | 12.42 | .26 | 10.98 | 3.23 | 1.13 | 1.68 | 1.94 | 2.12 | 2.24 | 2.32 | 2.36 | 2.40 | 2.40 | 2.24 | 1.98 | 1.47 |\n| 10 | 25.55 | 0 | .28 | 1.04 | 12.78 | | | 12.60 | 2.08 | 11.58 | 5.18 | .47 | .93 | 1.36 | 1.70 | 1.96 | 2.12 | 2.24 | 2.34 | 2.40 | 2.36 | 2.18 | 1.70 |\n| 11 | 28.94 | 0 | 0 | 0 | 12.85 | | | 12.62 | 3.27 | 11.72 | 6.28 | | .44 | .88 | 1.30 | 1.68 | 1.96 | 2.14 | 2.24 | 2.32 | 2.38 | 2.29 | 1.75 |\n| 12 | 30.88 | | | | 12.80 | | | 12.46 | 4.78 | 11.58 | 7.79 | | | .16 | .64 | 1.08 | 1.48 | 1.80 | 2.04 | 2.20 | 2.28 | 2.20 | 1.64 |\n| 13 | 33.53 | 2.62 | | | 12.66 | | | | 6.28 | 11.25 | 9.30 | | | | | .58 | .86 | 1.30 | 1.66 | 1.92 | 2.18 | |\n| 14 | 36.18 | 4.15 | | | 12.52 | 2.21 | 10.31 | | 7.82 | | | | | | | .14 | .60 | 1.06 | 1.60 | 2.06 | | |\n| 15 | 38.84 | 5.68 | | | 12.39 | 2.08 | 10.31 | | | | | | | | | | .31 | 1.40 | 1.92 | | | |\n| 16 | 41.50 | 7.24 | | | 12.24 | 1.94 | 10.31 | | | | | | | | | | | | | | | |\n| 17 | 44.15 | | | | | 1.80 | 10.31 | | | | | | | | | | | | | | | |\n| 18 | 46.81 | | | | | 1.64 | 10.31 | | | | | | | | | | | | | | | |\n| 19 | 49.47 | | | | | 1.50 | 10.31 | | | | | | | | | | | | | | | |\n| 20 | 52.13 | | | | | 1.37 | 10.31 | | | | | | | | | | | | | | | |\n| 21 | 54.78 | | | | | 1.22 | 10.31 | | | | | | | | | | | | | | | |\n| 22 | 57.44 | | | | | 1.08 | 10.31 | | | | | | | | | | | | | | | |\n| 23 | 60.09 | | | | | .94 | 10.31 | | | | | | | | | | | | | | | |\n| 24 | 62.75 | | | | | .80 | 10.31 | | | | | | | | | | | | | | | |\n| A.P. | 64.34 | | | | | .71 | 10.31 | | | | | | | | | | | | | | | |\n\nCONFIDENTIAL\nNACA\n9\nNACA RM L9D15", "timestamp": "2026-07-22T05:13:56.860902+00:00"}
{"citation_id": "19930082914", "source_url": "https://ntrs.nasa.gov/api/citations/19930082914/downloads/19930082914.pdf", "page_number": 34, "total_pages": 66, "image_filename": "19930082914_p34.jpg", "text": "NACA TN No. 1857\n33\n\n[Figure: Basic arrangement of Mach-Zehnder interferometer.]", "timestamp": "2026-07-22T05:13:58.859772+00:00"}
{"citation_id": "19930085519", "source_url": "https://ntrs.nasa.gov/api/citations/19930085519/downloads/19930085519.pdf", "page_number": 29, "total_pages": 46, "image_filename": "19930085519_p29.jpg", "text": "28\n\n$\\frac{1}{16}$\n\n$0.70c$\n\n$0.022c$\n\n$\\delta_P$\n\n$0.002c$\n\n$\\frac{1}{8}$\n\n$0.10c$\n\n$0.01$\n\n$\\frac{1}{4}$\n\n$0.002c$\n\n$0.013c$\n\n$0.036c$\n\n$\\frac{1}{32} \\text{ gap}$\n\nNACA\n\nFigure 8.- Section dimensions of the plug aileron with the sharp plug-slot lower lip tested on the $42^\\circ$ sweptback wing. All dimensions are in inches unless otherwise noted.\n\nNACA RM No. L58L19", "timestamp": "2026-07-22T05:14:02.129020+00:00"}
{"citation_id": "19930085487", "source_url": "https://ntrs.nasa.gov/api/citations/19930085487/downloads/19930085487.pdf", "page_number": 32, "total_pages": 36, "image_filename": "19930085487_p32.jpg", "text": "```markdown\n30\nNACA RM No. E8J22\n\n100,000\n\nFrequency, cps\n\n10,000\n\nRear stator row, 56 blades\nFront stator row, 52 blades\n\nFirst torsional mode\nSecond bending mode\n\n1,000\n\nFirst bending mode\n\nSecond order, split compressor case\n\nFirst order, rotor speed\n\n100\n4 6 8 10 12 14 16x10³\nRotor speed, rpm\n(f) Sixth stage.\n\nFigure 9. - Continued. Critical-speed diagrams for 10 stages of\ncompressor rotor.\n\nNACA\n\n601\n```", "timestamp": "2026-07-22T05:14:03.906943+00:00"}
{"citation_id": "19930085889", "source_url": "https://ntrs.nasa.gov/api/citations/19930085889/downloads/19930085889.pdf", "page_number": 4, "total_pages": 37, "image_filename": "19930085889_p4.jpg", "text": "NACA RM L9F14 CONFIDENTIAL 3\n\nSYMBOLS\n\nThe data are presented in the form of standard NACA coefficients of forces and moments which are referred in all cases to the stability axes, with the origin at the quarter-chord point of the mean aerodynamic chord of the models tested. The positive directions of the forces, moments, and angular displacements are shown in figure 1. The coefficients and symbols used herein are defined as follows:\n\n| Symbol | Definition |\n| :--- | :--- |\n| $C_L$ | lift coefficient ($L/qS$) |\n| $C_D$ | drag coefficient ($-X/qS$) |\n| $C_{D_0}$ | drag coefficient at zero lift |\n| $C_Y$ | lateral-force coefficient ($Y/qS$) |\n| $C_l$ | rolling-moment coefficient ($L'/qSb$) |\n| $C_m$ | pitching-moment coefficient ($M/qS\\bar{c}$) |\n| $C_n$ | yawing-moment coefficient ($N/qSb$) |\n| $L$ | lift |\n| $X$ | longitudinal force |\n| $Y$ | lateral force |\n| $L'$ | rolling moment about X-axis |\n| $M$ | pitching moment about Y-axis |\n| $N$ | yawing moment about Z-axis |\n| $q$ | dynamic pressure $\\left(\\frac{1}{2}\\rho V^2\\right)$ |\n| $\\rho$ | mass density of air |\n| $V$ | free-stream velocity |\n| $R$ | Reynolds number |\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:14:07.562113+00:00"}
{"citation_id": "19930082498", "source_url": "https://ntrs.nasa.gov/api/citations/19930082498/downloads/19930082498.pdf", "page_number": 42, "total_pages": 49, "image_filename": "19930082498_p42.jpg", "text": "NACA TN No. 1838\n41\n\n[Figure: Engine with exhaust exit cones removed. (Propeller was removed for configuration 2, table II.)]\n\nFigure 4.- Engine with exhaust exit cones removed. (Propeller was removed for configuration 2, table II.)", "timestamp": "2026-07-22T05:14:08.597723+00:00"}
{"citation_id": "19930082614", "source_url": "https://ntrs.nasa.gov/api/citations/19930082614/downloads/19930082614.pdf", "page_number": 32, "total_pages": 36, "image_filename": "19930082614_p32.jpg", "text": "30\nNACA TN 1939\n\nP\nO\nQ\nR\n\n[Figure: Four aircraft illustrations labeled P, O, Q, and R, showing different configurations of fuselage air brakes.]\n\n(c) Fuselage air brakes.\n\nFigure 6.— Concluded.", "timestamp": "2026-07-22T05:14:10.811660+00:00"}
{"citation_id": "19930082617", "source_url": "https://ntrs.nasa.gov/api/citations/19930082617/downloads/19930082617.pdf", "page_number": 28, "total_pages": 58, "image_filename": "19930082617_p28.jpg", "text": "```markdown\nNACA TN 1962\n27\n\n| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | |", "timestamp": "2026-07-22T05:14:14.807510+00:00"}
{"citation_id": "19930085544", "source_url": "https://ntrs.nasa.gov/api/citations/19930085544/downloads/19930085544.pdf", "page_number": 15, "total_pages": 33, "image_filename": "19930085544_p15.jpg", "text": "14\nNACA RM No. L8K26\n\n$L_c$ against $\\omega t$ at $k = 0$ shows the results obtained for assumed steady-state conditions in potential flow. The curves in figure 9 for other values of $k$ show that the effect of the oscillations is to modify the forces as obtained from the assumption of steady-state conditions. Presumably, a similar effect would occur in a compressible flow.\n\nThe asymmetry of the $L_c$ curves in figure 9 is caused primarily by the variation in the dynamic pressure $\\frac{1}{2}\\rho W_{wt}^2$ during the cycle. This asymmetry may be also seen in figure 10, which shows the variation at $k = 0.10$ of $L_c$ with $\\omega t$ with $\\epsilon$ as parameter for several values of $\\alpha/\\alpha_{P_o}$. The curves for $\\epsilon = 0$ are the most nearly symmetrical; this condition corresponds to Theodorsen's case of an airfoil oscillating in a steady flow (reference 1). It can be seen from figure 10 that the amplitude of the lift variation tends to increase as $\\epsilon$ increases and also that this amplitude increase is further accentuated by increasing the initial load at $\\omega t = 0$ ($\\alpha/\\alpha_{P_o}$ increasing). Figure 11 shows the variation of the force coefficients for $\\omega t = 90^\\circ$ and for $\\omega t = 270^\\circ$ for several values of $\\epsilon$ for $\\frac{\\alpha}{\\alpha_{P_o}} = 0$. On the same figure are values of $(L_{c90} - L_{c270})$ which are a measure of the maximum bending moment on the shaft axis of a two-blade propeller (divided by 2 for convenience of plotting). The absolute magnitude of the maximum force coefficients varies greatly as the value of $\\epsilon$ is increased, but the variation of the difference in the forces from the heavily loaded side to the lightly loaded side which gives the bending moment is small.\n\nFigure 12 shows the variation of $L_c$ at $90^\\circ$ and $270^\\circ$ with $k$ at several values of $\\alpha/\\alpha_{P_o}$ with $\\epsilon = 0$. In this special case it is seen that the magnitude of the loading increases as $\\alpha/\\alpha_{P_o}$ is increased but that the bending load factor $\\frac{L_{c90} - L_{c270}}{2}$ is independent of initial loading on the blade section. This independence does not hold in a pulsating flow field as may be seen in figure 13, which shows the variation of the bending load factor $\\frac{L_{c90} - L_{c270}}{2}$ with $k$ at several values of $\\epsilon \\frac{\\alpha}{\\alpha_{P_o}}$.\n\nFigure 14 shows the variation in the turning-moment coefficient on the inclined propeller shaft (yawing moment for pitched propeller) of the NACA 4-(3.9)(07)-0345-B propeller. These coefficients were calculated by the oscillating-flow theory and are considerably lower than the moments calculated by assuming steady-state conditions at each phase angle and using compressible airfoil characteristics. The curves are only useful", "timestamp": "2026-07-22T05:14:15.557586+00:00"}
{"citation_id": "19930082496", "source_url": "https://ntrs.nasa.gov/api/citations/19930082496/downloads/19930082496.pdf", "page_number": 46, "total_pages": 50, "image_filename": "19930082496_p46.jpg", "text": "NACA TN No. 1836\n45\n\n[Figure: A photograph of a mechanical component, likely a turbine disk, showing a series of blades or vanes attached to a central hub. A ruler labeled \"INCHES\" is placed next to the component for scale. An arrow points to a specific area on the hub, labeled \"Wheel dovetail distortion\". In the bottom right corner of the image, there is a NACA logo with the text \"C-21153\" and \"4-13-48\".]\n\n(b) Distortion 33 minutes before failure.\nFigure 12. - Concluded. Disk-dovetail failure.", "timestamp": "2026-07-22T05:14:15.806724+00:00"}
{"citation_id": "19930085626", "source_url": "https://ntrs.nasa.gov/api/citations/19930085626/downloads/19930085626.pdf", "page_number": 14, "total_pages": 24, "image_filename": "19930085626_p14.jpg", "text": "NACA RM No. L8K23\n13\n\nCONFIDENTIAL\n\n[Figure: A black and white photograph of a model aircraft or missile in a wind tunnel. The model has a cylindrical body with a large, open circular intake at the front. It features a high-mounted vertical tail fin and two horizontal wings or stabilizers. The model is mounted on a support structure. A small plaque is visible on the side of the fuselage. In the bottom right corner of the photograph, there is a label with the text \"NACA L-56078\".]\n\n(c) Parallel-sided aileron configuration.\n\nFigure 2.- Concluded.\nCONFIDENTIAL", "timestamp": "2026-07-22T05:14:16.174242+00:00"}
{"citation_id": "19930082245", "source_url": "https://ntrs.nasa.gov/api/citations/19930082245/downloads/19930082245.pdf", "page_number": 57, "total_pages": 66, "image_filename": "19930082245_p57.jpg", "text": ".16\n.12\n.08\n.04\n0\n-.04\n-.08\n-.12\n-.16\n-.20\n-.24\n-.28\n.1 .2 .3 .4 .5 .6 .7 .8 .9\nMach number, M\n$\\delta_a$\n(deg)\n-12\n-6\n-4\n-2\n0\n2\n4\n6\n12\n18\n30\nAileron section hinge-moment coefficient, $c_h$\nNACA\n56\nNACA TN No. 1996\n\n1.6\n1.4\n1.2\n1.0\n.8\n.6\n.4\n.2\n0\n-.2\n-.4\n-.6\n.1 .2 .3 .4 .5 .6 .7 .8 .9\nMach number, M\n$\\delta_a$\n(deg)\n30\n18\n12\n6\n4\n2\n0\n-2\n-4\n-6\n-12\nAileron section normal-force coefficient, $c_{n\\alpha}$\n\n(g) $c_n = 0.7$.\nFigure 10.—Continued.", "timestamp": "2026-07-22T05:14:17.271546+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 7, "total_pages": 96, "image_filename": "19930085880_p7.jpg", "text": "```markdown\nNACA RM No. L9C03\n5\n\nRESULTS\n\nThe results are presented in the form of plots of the load on the\nwater, resistance, trimming moment, and draft against total wetted area\nwith speed and trim as parameters. Figures 14 to 17 give the results\nfor the rectangular surface with a flat bottom (model 250A). Figures 18\nto 21 are for the rectangular surface with curved bottom (model 250B)\nand figures 22 to 25 are for the triangular surface with a flat bottom\n(model 250D).\n\nFrom the procedure described, the quantities in the figures are\ndefined as follows:\n\n(a) Resistance is the measured resistance less the aerodynamic\ndrag of the towing gear less model (fig. 9) and the estimated aero-\ndynamic drag of the unwetted portion of the model (figs. 10 and 11).\n\n(b) Trimming moment is the measured trimming moment referred to\nthe trailing edge of the model. The aerodynamic moment tare was\nnegligible.\n\n(c) The load on the water is the unbalanced weight of the model\nand gear less the estimated aerodynamic lift of the unwetted portion\nof the model (figs. 12 and 13). The aerodynamic lift tare on the gear\nalone was negligible.\n\n(d) Draft is the depth of the trailing edge of the model below\nthe free water surface.\n\n(e) Wetted area is the wetted plan-form area computed from the plan\nform and the observed wetted length at the chine plus the wetted area\nforward of the observed wetted length (fig. 8). The latter was\nnegligible for model 250A.\n\n(f) Wetted length at the chine is the observed length from the\ntrailing edge of the model to the intersection of the dynamic solid\nwater boundary with the chine. The wetted length at the center line\nwas appreciably greater for models 250B and 250D. (See figs. 6 and 7.)\n\nLangley Aeronautical Laboratory\nNational Advisory Committee for Aeronautics\nLangley Air Force Base, Va.\n```", "timestamp": "2026-07-22T05:14:20.782764+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 47, "total_pages": 99, "image_filename": "19930082511_p47.jpg", "text": "NACA TN No. 1826\n45\n\nThe constants $h_{mn}^{(j)}$ are to be determined so that condition (3) is satisfied. For this purpose the function $- \\left. \\frac{\\partial(\\phi_0 + \\phi_C)}{\\partial \\xi} \\right|_{\\rho=1}$ is expanded in a Fourier series in $\\theta$:\n\n$$\n- \\left. \\frac{\\partial(\\phi_0 + \\phi_C)}{\\partial \\xi} \\right|_{\\rho=1} = \\sum_{m=1}^{\\infty} \\left[ r_m^{(1)}(\\xi) \\sin m\\theta + r_m^{(2)}(\\xi) \\cos m\\theta \\right] \\quad (30)\n$$\n\nEquating coefficients in equations (29) and (30) in order to satisfy condition (3) then gives\n\n$$\nr_m^{(j)}(\\xi) = \\sum_{n=0}^{\\infty} h_{mn}^{(j)} \\frac{\\partial P_{mn}(\\xi, 1)}{\\partial \\xi} \\quad (j = 1, 2)\n$$\n\nIt is assumed that the functions $r_m^{(j)}(\\xi)$ can be satisfactorily approximated by a finite number of terms of these series. This assumption seems reasonable, inasmuch as $\\frac{\\partial P_{mn}(\\xi, 1)}{\\partial \\xi}$ is bounded as $n$ approaches infinity (see appendix B) and $h_{mn}^{(j)}$ approaches zero as $n$ approaches infinity. Thus,\n\n$$\nr_m^{(j)}(\\xi) \\approx \\sum_{n=0}^{N} h_{mn}^{(j)} \\frac{\\partial P_{mn}(\\xi, 1)}{\\partial \\xi}\n$$\n\nThe functions $r_m^{(j)}(\\xi)$ and $\\frac{\\partial P_{mn}(\\xi, 1)}{\\partial \\xi}$ are computed at a set of points $\\{\\xi_i\\}$, $i = 0, 1, 2, \\dots I$, where $I \\ge N$. The coefficients $h_{mn}^{(j)}$ are then determined (method of least squares) so that the expression\n\n$$\n\\sum_{i=0}^{I} \\left[ r_m^{(j)}(\\xi_i) - \\sum_{n=0}^{N} h_{mn}^{(j)} \\frac{\\partial P_{mn}(\\xi_i, 1)}{\\partial \\xi} \\right]^2\n$$\n\nis a minimum for all values of $m$ and $j$. For each pair of values of $m$ and $j$, this condition gives $N + 1$ equations for the $N + 1$", "timestamp": "2026-07-22T05:14:24.469989+00:00"}
{"citation_id": "19930085869", "source_url": "https://ntrs.nasa.gov/api/citations/19930085869/downloads/19930085869.pdf", "page_number": 12, "total_pages": 36, "image_filename": "19930085869_p12.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:14:31.235561+00:00"}
{"citation_id": "19930082914", "source_url": "https://ntrs.nasa.gov/api/citations/19930082914/downloads/19930082914.pdf", "page_number": 35, "total_pages": 66, "image_filename": "19930082914_p35.jpg", "text": "34\nNACA TN No. 1857\n\n[Figure: Diagram showing the production of fringes with monochromatic parallel light. The diagram includes two trains of waves from sources $I_1$ and $I_2$, with wavelength $\\lambda$. The intensity of the resulting wave pattern is shown as a graph with peaks and troughs.]\n\nFigure 3.—Production of fringes with monochromatic parallel light. $\\lambda$, wavelength of light.", "timestamp": "2026-07-22T05:14:34.844286+00:00"}
{"citation_id": "19930082485", "source_url": "https://ntrs.nasa.gov/api/citations/19930082485/downloads/19930082485.pdf", "page_number": 53, "total_pages": 62, "image_filename": "19930082485_p53.jpg", "text": "52\nNACA TN No. 1810\n\n1026\n\n$\\left(\\frac{V}{V_{cr}}\\right)_i$\n\n$\\phi_{10}$\n$\\phi_9$\n$\\phi_8$\n$\\phi_7$\n$\\psi_5$\n$\\psi_4$\n$\\psi_3$\n$\\psi_2$\n$\\psi_1$\n$\\phi_6$\n$\\phi_5$\n$\\phi_4$\n$\\phi_3$\n$\\phi_2$\n$\\phi_1$\n\n$\\left(\\frac{V}{V_{cr}}\\right)_e$\n\n$\\mu n_o$\n\nNACA\n\nFigure 21. - Flow network in channel between blades and assumed distribution of weight-flow parameter.\n\n132-1448", "timestamp": "2026-07-22T05:14:35.116865+00:00"}
{"citation_id": "19930082542", "source_url": "https://ntrs.nasa.gov/api/citations/19930082542/downloads/19930082542.pdf", "page_number": 26, "total_pages": 53, "image_filename": "19930082542_p26.jpg", "text": "NACA TN No. 1867\n25\n\nTABLE II.—RUPTURE TEST CHARACTERISTICS AT 1200° F OF LOW-CARBON B-155 BAR STOCK — Continued\n\n<!-- Table (93, 119, 891, 904) -->\n\\begin{tabular}{|c|c|c|c|c|c|c|c|c|c|c|c|c|c|}\n\\hline\n\\multicolumn{3}{|c|}{Heat treatment} & \\multicolumn{2}{c|}{Aging treatment (a)} & \\multicolumn{2}{c|}{Hot-cold-rolling (b)} & \\multicolumn{7}{c|}{Rupture properties at 1200° F} \\\\\n\\cline{1-14}\n\\multicolumn{3}{|c|}{Solution treatment} & \\multicolumn{2}{c|}{} & \\multicolumn{2}{c|}{} & \\multicolumn{7}{c|}{} \\\\\n\\cline{1-14}\nTemper- & Time & Method & Temper- & Time & Temper- & Percent & Stress & Rupture & Elongation & Reduction & \\multicolumn{2}{c|}{Rupture strength} \\\\\nature & (hr) & of & ature & (hr) & ature & reduction & (psi) & time & in 1 in. & of area & \\multicolumn{2}{c|}{(psi)} \\\\\n($^\\circ$F) & & cooling & ($^\\circ$F) & & ($^\\circ$F) & & & (hr) & (percent) & (percent) & 100 hr & 1000 hr \\\\\n& & (c) & & & & & & & & & & \\\\\n\\hline\n\\multicolumn{14}{|c|}{Solution-treated at 2050° F} \\\\\n\\hline\n\\multicolumn{14}{|c|}{Cold-working amount and temperature} \\\\\n\\hline\n2050 & 2 & W.Q. & ----- & -- & 75 & 10 & 60,000 & 24 & $e_1$ & 5.3 & 56,000 & 50,000 \\\\\n& & & & & & (9.8) & 55,000 & 275 & 4 & 11.2 & & \\\\\n& & & & & & & 50,000 & 1068 & 5 & 20.6 & & \\\\\n\\hline\n2050 & 2 & W.Q. & ----- & -- & 1200 & 5 & 55,000 & 46 & 2 & 5.6 & 52,000 & 44,000 \\\\\n& & & & & & & 52,500 & 86 & 2 & 9.7 & & \\\\\n& & & & & & & 50,000 & 261 & $e_4$ & 8.5 & & \\\\\n& & & & & & & 45,000 & 821 & 18 & 23.3 & & \\\\\n\\hline\n2050 & 2 & W.Q. & ----- & -- & 1200 & 10 & 60,000 & 69 & $e_1$ & 2.7 & 59,000 & 52,000 \\\\\n& & & & & & & 55,000 & 614 & 4 & 11.0 & & \\\\\n& & & & & & & 50,000 & 1465 & 8.5 & 23.2 & & \\\\\n\\hline\n2050 & 2 & W.Q. & 1400 & 24 & 1200 & 10 & 55,000 & 88 & 22 & 39.8 & 54,000 & 43,500 \\\\\n& & & & & & & 50,000 & 268 & 18 & 31.3 & & \\\\\n& & & & & & & 47,000 & 302 & $e_{15}$ & 23.3 & & \\\\\n& & & & & & & 44,000 & 920 & 15 & 40.0 & & \\\\\n\\hline\n2050 & 2 & W.Q. & ----- & -- & 1200 & 15 & 60,000 & 167 & 1 & 1.5 & 62,000 & 53,500 \\\\\n& & & & & & (14.3) & 55,000 & 389 & 8 & 21.0 & & \\\\\n& & & & & & & 52,000 & 1596 & 4 & 16.4 & & \\\\\n\\hline\n2050 & 2 & W.Q. & 61400 & 24 & 1200 & 15 & 55,000 & 110 & 18 & 37.9 & 55,500 & 48,000 \\\\\n& & & & & & & 52,500 & 365 & 19 & 37.9 & & \\\\\n& & & & & & & 50,000 & 576 & 13.5 & 37.9 & & \\\\\n\\hline\n2050 & 2 & W.Q. & ----- & -- & 1200 & 20 & 65,000 & 102 & 1.5 & 2.8 & 62,000 & 56,000 \\\\\n& & & & & & (18.9) & 60,000 & 95 & 2 & 6.9 & & \\\\\n& & & & & & & 60,000 & 243 & 1.5 & 4.1 & & \\\\\n& & & & & & & 57,500 & 773 & 3 & 7.0 & & \\\\\n& & & & & & & 55,000 & 1348 & 3 & 7.8 & & \\\\\n\\hline\n2050 & 2 & W.Q. & ----- & -- & 1200 & 25 & 70,000 & 19 & 2 & 5.7 & 66,000 & 56,000 \\\\\n& & & & & & & 65,000 & 200 & 4 & 11.0 & & \\\\\n& & & & & & & 60,000 & 423 & 5 & 18.9 & & \\\\\n& & & & & & & 55,000 & 1142 & 5 & 21.6 & & \\\\\n\\hline\n2050 & 2 & W.Q. & ----- & -- & 1000 & 15 & 65,000 & 79 & 2 & 2.5 & 63,500 & 53,000 \\\\\n& & & & & & (15.6) & 60,000 & 208 & 1.5 & 2.8 & & \\\\\n& & & & & & & 55,000 & 389 & 9 & 21.3 & & \\\\\n& & & & & & & 50,000 & 2087 & 4.5 & 19.1 & & \\\\\n\\hline\n2050 & 2 & W.Q. & ----- & -- & 1400 & 15 & 65,000 & 66 & 3 & 10.9 & 62,500 & 52,000 \\\\\n& & & & & & (15.3) & 60,000 & 176 & 3.5 & 13.2 & & \\\\\n& & & & & & & 55,000 & 551 & 6 & 25.6 & & \\\\\n\\hline\n2050 & 2 & W.Q. & ----- & -- & 1600 & 15 & 60,000 & 40 & 15 & 29.8 & 56,000 & 45,000 \\\\\n& & & & & & (14.3) & 55,000 & 188 & 11 & 37.2 & & \\\\\n& & & & & & & 50,000 & 321 & 13 & 40.2 & & \\\\\n& & & & & & & 45,000 & 1013 & 14 & 29.5 & & \\\\\n\\hline\n2050 & 2 & W.Q. & ----- & -- & 1700 & 15 & 55,000 & 61 & 24.5 & 36.9 & 52,000 & 39,000 \\\\\n& & & & & & (15.7) & 50,000 & 123 & 25 & 36.9 & & \\\\\n& & & & & & & 45,000 & 342 & 28 & 39.8 & & \\\\\n\\hline\n2050 & 2 & W.Q. & ----- & -- & 1800 & 15 & 54,000 & 48.5 & 25 & 36.9 & 51,000 & 43,000 \\\\\n& & & & & & (15.0) & 50,000 & 125 & 19 & 35.0 & & \\\\\n& & & & & & & 45,000 & 535 & 20 & 33.0 & & \\\\\n\\hline\n\\end{tabular}\n\n$^a$All aging treatments preceded hot-cold-rolling except where noted.\n$^b$All hot-cold-rolled material was given a final stress relief at 1200° F for 1 hr.\n$^c$W.Q., water-quenched; A.C., air-cooled.\n$^d$Fractured in gage mark.\n$^e$Aged after rolling.\n\nNACA", "timestamp": "2026-07-22T05:14:35.528667+00:00"}
{"citation_id": "19930085487", "source_url": "https://ntrs.nasa.gov/api/citations/19930085487/downloads/19930085487.pdf", "page_number": 33, "total_pages": 36, "image_filename": "19930085487_p33.jpg", "text": "NACA RM No. E8J22\n31\n\nFrequency, cps\n100,000\n10,000\n1,000\n100\n\nRear stator row, 58 blades\nFront stator row, 56 blades\nFirst torsional mode\nSecond bending mode\nFirst bending mode\nFourth order required\nto cause failure in\nfundamental mode\nSecond order, split compressor case\nFirst order, rotor speed\n\n4 6 8 10 12 14 16x10³\nRotor speed, rpm\n\nNACA\n\n(g) Seventh stage.\nFigure 9. - Continued. Critical-speed diagrams for 10 stages of\ncompressor rotor.", "timestamp": "2026-07-22T05:14:39.174451+00:00"}
{"citation_id": "19930082498", "source_url": "https://ntrs.nasa.gov/api/citations/19930082498/downloads/19930082498.pdf", "page_number": 43, "total_pages": 49, "image_filename": "19930082498_p43.jpg", "text": "Page intentionally left blank\n\nPage intentionally left blank", "timestamp": "2026-07-22T05:14:43.826885+00:00"}
{"citation_id": "19930085889", "source_url": "https://ntrs.nasa.gov/api/citations/19930085889/downloads/19930085889.pdf", "page_number": 5, "total_pages": 37, "image_filename": "19930085889_p5.jpg", "text": "4\nCONFIDENTIAL\nNACA RM L9F14\n\nS wing area\nb span of wing, measured perpendicular to plane of symmetry\nc chord of wing, measured parallel to plane of symmetry\n$\\overline{c}$ mean aerodynamic chord $\\left(\\frac{2}{S} \\int_{0}^{b/2} c^2 dy\\right)$\ny distance measured perpendicular to the plane of symmetry\n$a_o$ slope of section lift curve per radian\nA aspect ratio $(b^2/S)$\n$\\alpha$ angle of attack measured in plane of symmetry, degrees\n$\\delta$ aileron deflection measured in plane normal to aileron hinge axis, degrees\n$\\psi$ angle of yaw, degrees\n$\\Lambda$ angle of sweepback of quarter-chord line, degrees\n$\\frac{pb}{2V}$ wing-tip helix angle, radians\np rolling velocity, radians per second\n$\\left(\\frac{pb}{2V}\\right)_\\delta$ rate of change of wing-tip helix angle per degree of total aileron deflection\n$C_{l_\\delta}$ rate of change of rolling-moment coefficient per degree of total aileron deflection\n$C_{L_\\alpha} = \\frac{\\partial C_L}{\\partial \\alpha}$\n$C_{l_\\psi} = \\frac{\\partial C_l}{\\partial \\psi}$\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:14:46.308101+00:00"}
{"citation_id": "19930082614", "source_url": "https://ntrs.nasa.gov/api/citations/19930082614/downloads/19930082614.pdf", "page_number": 33, "total_pages": 36, "image_filename": "19930082614_p33.jpg", "text": "```markdown\nNACA TN 1939\n31\n\nUpper-and lower-surface split flaps at\n80-percent chord, brake type A\nUpper-surface spoiler at 20-percent\nchord, brake type I\nLower-surface spoiler at 20-percent\nchord, brake type I\nUpper-surface spoiler at 75-percent\nchord, brake type J\nUpper-surface spoiler at 63-percent\nchord, brake type K\nFuselage side brake, brake type R\n\nIncremental drag coefficient, $\\Delta C_{DB}$\n2.0\n1.6\n1.2\n.8\n.4\n0\n\n0\n20\n40\n60\n80\n100\nAir-brake deflection, percent\n\n[Figure: NACA logo]\n\nFigure 7.— The variation of air-brake drag\ncoefficient with percent of full brake\ndeflection.\n```", "timestamp": "2026-07-22T05:14:46.937184+00:00"}
{"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 32, "total_pages": 78, "image_filename": "19930082618_p32.jpg", "text": "```markdown\n30\n\n.028\n.024\n.020\n.016\n.012\n.008\n.004\nSection drag coefficient, $c_d$\n\nR\n$\\circ$ 0.7 x $10^6$\n$\\square$ 1.0\n$\\triangle$ 1.5\n$\\nabla$ 2.0\nFlagged symbols denote\nstandard roughness\n\n-.8 -.4 0 .4 .8 1.2\nSection lift coefficient, $c_l$\n\n.028\n.024\n.020\n.016\n.012\n.008\n.004\nSection drag coefficient, $c_d$\n\nR\n$\\nabla$ 3.1 x $10^6$\n$\\triangleright$ 6.0\n$\\triangleleft$ 9.0\nFlagged symbols denote\nstandard roughness\n\n-1.2 -.8 -.4 0 .4 .8 1.2 1.6\nSection lift coefficient, $c_l$\n\n.02\n.01\n0\n-.01\n-.02\n-.03\nMoment coefficient, $c_{mac}$\n\nR a.c. position\nx/c y/c\n$\\circ$ 0.7 x $10^6$ .265 -.060\n$\\square$ 1.0 .268 -.087\n$\\triangle$ 1.5 .270 -.084\n$\\nabla$ 2.0 .271 -.077\n$\\triangleright$ 3.1 .264 -.070\n$\\triangleleft$ 6.0 .265 -.051\n$\\triangleleft$ 9.0 .264 -.040\n\nNACA\n\n-.8 -.4 0 .4 .8 1.2 1.6\nSection lift coefficient, $c_l$\n\n(c) Section drag characteristics and section pitching-moment characteristics about the aerodynamic\ncenter of the plain NACA 64$_2$-415 airfoil section.\n\nFigure 3.— Concluded.\n\nNACA TN 1945\n```", "timestamp": "2026-07-22T05:14:52.616722+00:00"}
{"citation_id": "19930085626", "source_url": "https://ntrs.nasa.gov/api/citations/19930085626/downloads/19930085626.pdf", "page_number": 15, "total_pages": 24, "image_filename": "19930085626_p15.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:14:52.847751+00:00"}
{"citation_id": "19930082617", "source_url": "https://ntrs.nasa.gov/api/citations/19930082617/downloads/19930082617.pdf", "page_number": 29, "total_pages": 58, "image_filename": "19930082617_p29.jpg", "text": "28\nNACA TN 1962\n\nStringers\nO 1 to 9\nX 10 to 16\n\nMoment\n(in. - lb)\n1 36.0 X $10^3$\n2 72.0 X $10^3$\n3 108.0 X $10^3$\n4 144.0 X $10^3$\n5 180.0 X $10^3$\n6 216.0 X $10^3$\n7 288.0 X $10^3$\n\n2.57\"\n1\"\nBand B\nA\nA-A\n45°\n\nDistance from horizontal diameter, in.\n10\n9\n8\n7\n6\n5\n4\n3\n2\n1\n0\n1\n2\n3\n4\n5\n6\n7\n8\n9\n10\n\n20 16 12 8 4 0 -4 -8 -12 -16 -20 X $10^{-4}$\nStrain\n\n1 2 3 4 5 6 7\n\nNACA\n\nFigure 17.- Strain diagram of cylinder 77. Band B.", "timestamp": "2026-07-22T05:14:54.532543+00:00"}
{"citation_id": "19930085879", "source_url": "https://ntrs.nasa.gov/api/citations/19930085879/downloads/19930085879.pdf", "page_number": 10, "total_pages": 29, "image_filename": "19930085879_p10.jpg", "text": "8\nNACA RM L9D11\n\nTABLE I\n\nMASS-BALANCE CHARACTERISTICS OF MODELS\n\n| | Model A | Model B |\n| :--- | :---: | :---: |\n| **Complete model:** | | |\n| Weight at launching . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 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{"citation_id": "19930085544", "source_url": "https://ntrs.nasa.gov/api/citations/19930085544/downloads/19930085544.pdf", "page_number": 16, "total_pages": 33, "image_filename": "19930085544_p16.jpg", "text": "NACA RM No. L8K26\n\nin showing the variation of the moments with the time variable $\\omega t$ and not the absolute magnitude. The variation of the turning-moment coefficient with time for a two-blade propeller operating at a $V/nD$ of 1.2, $\\alpha_T = 4^\\circ$, and $\\beta = 26^\\circ$ is shown by the solid curve and for a three-blade propeller by the dashed curve. It is seen that the moment coefficient for a two-blade propeller varies from approximately zero when the blades are in the vertical position to a maximum of 0.0021 when the blades are approximately on the horizontal (twice per revolution). For the same operating condition for the three-blade propeller the moment coefficient remains very nearly constant at approximately 0.0016 (varying between 0.0015 and 0.0017).\n\nCombined steady and oscillating forces.- From the standpoint of theory a combination of the steady-state and the oscillating-airfoil theories approaches the actual operating conditions of the pitched or yawed propeller. The forces or moments are computed by the steady-state methods including compressibility and downwash. The oscillating-airfoil theory is then used to modify these forces. For a two-blade propeller, the maximum force difference $(L_{90} - L_{270})$ at each section as computed from steady-state calculations with compressible airfoil characteristics would be reduced by a factor which is the ratio of $\\left(\\frac{L_{c90} - L_{c270}}{L_{90} - L_{270}}\\right)$ at the operating values of $k$ to its value at $k = 0$ (from fig. 13). From figure 13 it will be noticed that, properly, the operating value of $\\epsilon \\frac{\\alpha}{\\alpha_{p_o}}$ should also be taken into account; within the limits given in the figure, however, the ratio\n\n$$\n\\frac{(L_{c90} - L_{c270})_k}{(L_{c90} - L_{c270})_{k=0}}\n$$\n\nchanges only slightly with $\\epsilon \\frac{\\alpha}{\\alpha_{p_o}}$.\n\nFigure 15 shows the calculated distribution of the moment coefficient along the radius for the two-blade NACA 4-(3.9)(07)-0345-B propeller with the blades in the horizontal position, with the thrust axis inclined at $4^\\circ$, with the propeller blade angle set at $26^\\circ$ at the 0.75 radius, and operating at a $V/nD$ of 1.2. The moment coefficients are computed for steady-state conditions in compressible flow and for an oscillating airfoil by the oscillating-airfoil theory. An integration of these curves gives the total turning moment on the propeller shaft. The correction for each radius as obtained from figure 13 has been applied to the steady-state calculations. This correction does not bring the curves into agreement because of the differences in the airfoil characteristics used.", "timestamp": "2026-07-22T05:15:01.121786+00:00"}