PDF/ocr_dataset_1013.jsonl

{"citation_id": "19930090382", "source_url": "https://ntrs.nasa.gov/api/citations/19930090382/downloads/19930090382.pdf", "page_number": 20, "total_pages": 37, "image_filename": "19930090382_p20.jpg", "text": "22\nNACA RM L9I07\n\nCONFIDENTIAL\n\nTip Mach number, $M_t$\n1.5\n1.0\n0.5\n0\n\nEfficiency, $\\eta$\n100\n75\n50\n25\n0\n\nPower coefficient, $C_p$\n.60\n.55\n.50\n.45\n.40\n.35\n.30\n.25\n.20\n.15\n.10\n.05\n0\n\nThrust coefficient, $C_T$\n.300\n.275\n.250\n.225\n.200\n.175\n.150\n.125\n.100\n.075\n.050\n.025\n0\n\nAdvance ratio, J\n0\n.5\n1.0\n1.5\n2.0\n2.5\n3.0\n3.5\n4.0\n4.5\n\n$M_t$\n$C_p$\n$C_T$\n$\\eta$\n$\\beta_{0.75R}$ 45°\n50°\n55°\n\nCONFIDENTIAL\n\n(1) M=0.60\nFigure 5 - Continued.\n\nNACA", "timestamp": "2026-07-22T04:35:22.701180+00:00"}
{"citation_id": "19930085842", "source_url": "https://ntrs.nasa.gov/api/citations/19930085842/downloads/19930085842.pdf", "page_number": 77, "total_pages": 104, "image_filename": "19930085842_p77.jpg", "text": "NACA RM L9C29\n73\n\n[Figure: A graph plotting Lift coefficient, $C_L$ against Propeller advance-diameter ratio, $V/nD$. The graph is divided into two sections, (a) and (b). The y-axis ranges from 0 to 4.0. The x-axis ranges from 2 to 10. The graph contains multiple curves with data points marked by circles and triangles. Labels on the curves include $\\alpha, deg$ and numerical values such as 29.2, 14.4, 11.4, 5.4, -0.5, 29.1, 20.1, 11.3, 5.3, -0.6. The National Advisory Committee for Aeronautics logo is present in the bottom right corner of the graph.]\n\n(a) $\\beta = 20^\\circ$.\n(b) $\\beta = 30^\\circ$.\n\nFigure 42.- Variation of $C_L$ with $V/nD$. Basic model configuration; propellers operating; all control surfaces neutral; data $\\beta = 20^\\circ$ curves obtained with wing-tip support.", "timestamp": "2026-07-22T04:35:23.453555+00:00"}
{"citation_id": "19930082476", "source_url": "https://ntrs.nasa.gov/api/citations/19930082476/downloads/19930082476.pdf", "page_number": 7, "total_pages": 41, "image_filename": "19930082476_p7.jpg", "text": "NACA TN No. 1801\n\n$\\alpha$, degree . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 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. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .", "timestamp": "2026-07-22T04:35:24.076649+00:00"}
{"citation_id": "19930082450", "source_url": "https://ntrs.nasa.gov/api/citations/19930082450/downloads/19930082450.pdf", "page_number": 9, "total_pages": 37, "image_filename": "19930082450_p9.jpg", "text": "8\nNACA TN No. 1778\n\nREFERENCES\n\n1. Langhaar, Henry L.: Design of Hat-Type Plate-Stringer Combinations. Auto. and Aviation Ind., vol. 91, no. 11, Dec. 1, 1944, pp. 28-32 and 103-104.\n\n2. Schuette, Evan H.: Charts for the Minimum-Weight Design of 24S-T Aluminum-Alloy Flat Compression Panels with Longitudinal Z-Section Stiffeners. NACA Rep. No. 827, 1945.\n\n3. Dow, Norris F., and Hickman, William A.: Direct-Reading Design Charts for 75S-T Aluminum-Alloy Flat Compression Panels Having Longitudinal Straight-Web Y-Section Stiffeners. NACA TN No. 1640, 1948.\n\n4. Dow, Norris F., Hubka, Ralph E., and Roberts, William M.: Direct-Reading Design Charts for 24S-T Aluminum-Alloy Flat Compression Panels Having Longitudinal Straight-Web Y-Section Stiffeners. NACA TN No. 1777, 1949.\n\n5. Dow, Norris F., and Hickman, William A.: Design Charts for Flat Compression Panels Having Longitudinal Extruded Y-Section Stiffeners and Comparison with Panels Having Formed Z-Section Stiffeners. NACA TN No. 1389, 1947.", "timestamp": "2026-07-22T04:35:28.202480+00:00"}
{"citation_id": "19930085965", "source_url": "https://ntrs.nasa.gov/api/citations/19930085965/downloads/19930085965.pdf", "page_number": 45, "total_pages": 67, "image_filename": "19930085965_p45.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T04:35:28.530336+00:00"}
{"citation_id": "19930082447", "source_url": "https://ntrs.nasa.gov/api/citations/19930082447/downloads/19930082447.pdf", "page_number": 9, "total_pages": 24, "image_filename": "19930082447_p9.jpg", "text": "NACA TN No. 1775\n\nThe discrepancy between the theoretical and experimental values of the vertical-velocity ratio at the instant of rebound is attributed to friction and leakage in the compressed-air lift mechanism which balanced the weight. The leakage takes effect after the maximum draft has been reached and has a maximum effect on the motion at the instant of rebound. In addition, greater scatter is present at this instant because of the variation in the time lag of the instrumentation which was used to measure vertical velocity.\n\nIn some cases the model was immersed beyond the limits of the prismatic shape (72 inches long by 17.25 inches high - see fig. 1). The general agreement of the data with the theory, however, indicates that the effects of bow and chine immersion were of no great significance.\n\nEffect of Dead-Rise Angle on Hydrodynamic Load\n\nFrom the form of the load-factor coefficient it can be seen that, if all other parameters are held constant, the hydrodynamic load is proportional to the quantity $[f(\\beta)]^{2/3}$ where $f(\\beta) = \\frac{\\pi}{2\\beta} - 1$. If the conventional dead-rise angle of $22\\frac{1}{2}^\\circ$ is used as a base, this relationship indicates a reduction in load of 24 percent for an angle of dead rise of $30^\\circ$ and a reduction of 44 percent for an angle of dead rise of $40^\\circ$.\n\nThe validity of the theoretical variation of hydrodynamic load with dead rise was verified for angles of dead rise of $22\\frac{1}{2}^\\circ$ and $30^\\circ$ by experimental data obtained in the Langley impact basin (references 1, 3, and 4). In the present paper, the range of dead-rise angle is extended to $40^\\circ$. The data previously presented in figure 3 are further analyzed to determine an experimental value of the dead-rise function $[f(\\beta)]^{2/3}$.\n\nSolving equation (1) for $[f(\\beta)]^{2/3}$ gives\n\n$$\n[f(\\beta)]^{2/3} = \\frac{n_1 w g}{C_l \\dot{y}_o^2} \\left\\{ \\frac{w}{g} \\left[ \\frac{6 \\sin \\tau \\cos^2 \\tau}{\\gamma(A)\\pi\\rho} \\right] \\right\\}^{1/3}\n\\tag{4}\n$$\n\nFrom the theoretical relationship between $C_{l_{\\text{max}}}$ and $\\kappa$ of reference 1, a theoretical value of $C_{l_{\\text{max}}}$ is obtained corresponding to the approach parameter $\\kappa$ computed for each run (table II). Substituting this value of $C_l$ and the data of table II into equation (4) gives an experimental value of $[f(\\beta)]^{2/3}$ for each run. The distribution or resulting experimental values of $[f(\\beta)]^{2/3}$ is shown in figure 7, where the distributions are grouped as percentages of the total number of values used. This figure shows that the distribution about the average value is approximately normal and indicates that the deviation from the normal is largely random.", "timestamp": "2026-07-22T04:35:30.120534+00:00"}
{"citation_id": "19930092013", "source_url": "https://ntrs.nasa.gov/api/citations/19930092013/downloads/19930092013.pdf", "page_number": 8, "total_pages": 21, "image_filename": "19930092013_p8.jpg", "text": "4\nREPORT 948—NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\npilot still can apply in the normal manner any additional desired aileron deflection up to the maximum in either direction. The kinematics of the revised system are shown in figure 3, in which total aileron deflection $\\delta_a$ (the sum of left and right aileron angles) is plotted as a function of lateral stick deflection $\\theta$ for neutral and maximum test servo positions. These curves and additional data obtained at intermediate servo positions showed that, as is desired, the gearings $\\partial\\delta_a/\\partial\\theta$ and $(\\partial\\delta_a/\\partial\\theta)_s$ are nearly constant over the available ranges of $\\delta_a$ and $\\theta$.\n\nAileron servomechanism.—The aileron and tab servomechanisms were developed from an electric amplidyne system normally used for remote control of aircraft gun turrets. This system was chosen on the basis of signal-system and motor-output requirements, applicability to aircraft, and availability. A simplified electrical circuit diagram of the installation is given in figure 4. The error-measuring portion of the aileron servomechanism is essentially a two-potentiometer bridge circuit with a 30-volt 400-cycle power supply. One potentiometer is geared mechanically to a yaw vane located on a boom extending forward from the left wing tip of the airplane. The second potentiometer is connected to the aileron servo motor. With the yaw vane and servo motor initially neutral, the bridge circuit is at one balance point. When the vane is deflected through an angle of sideslip $\\beta$, an error signal is supplied by the bridge to the amplifier. The amplified signal, converted to direct current, is fed to the field of an amplidyne generator, the armature of which is driven continuously at constant speed by the direct-current amplidyne motor. The general output voltage, of a polarity and magnitude determined by the error signal, is applied to the armature of the reversible separately excited direct-current aileron servo motor. The generator output voltage determines the direction and speed of the servo-motor rotation, which moves the torque tube and attached potentiometer in the direction which tends to balance the bridge circuit at a new point corresponding to $\\beta$ and $(\\delta_a)_s$. The servo gearing $(\\partial\\delta_a/\\partial\\theta)_s$ can be altered through the switch $S_2$, which in effect varies the bridge unbalance voltage per degree sideslip. The sign of $(\\partial\\delta_a/\\partial\\theta)_s$, and thus $(\\Delta C_{l\\beta})_s$, can be reversed by switch $S_3$.\n\nAileron tab drive and servomechanism.—The ratio $(\\partial\\delta_t/\\partial\\delta_a)_s$ of the tab motion to aileron motion required to balance the hinge moment due to servo-actuated aileron motion was determined from preliminary flight tests. Insufficient total power of the original trim tab necessitated\n\n[Figure: Sideslip vane, Vane potentiometer, $S_2$ Servo gearing selector switch, $S_3$ Reversing switch, 30v, 400~, Aileron servo potentiometer, Aileron servo motor, Generator, Motor, To torque tube, Aileron amplifier, d.c. output, Electrical connections, Mechanical connections, 24 v d.c. supply, 115v, 400~, Aileron servo selsyn, Differential gearing, Cockpit trim tab control, Tab servo motor, To tab, Generator, Motor, Tab selsyn, Tab amplifier, Tab amplidyne, Aileron circuit, Aileron tab circuit]\n\nFIGURE 4.—Simplified electrical circuit diagram of the effective-dihedral control apparatus.", "timestamp": "2026-07-22T04:35:32.407989+00:00"}
{"citation_id": "19930086015", "source_url": "https://ntrs.nasa.gov/api/citations/19930086015/downloads/19930086015.pdf", "page_number": 39, "total_pages": 54, "image_filename": "19930086015_p39.jpg", "text": "38\nCONFIDENTIAL\nNACA RM A9E24\n\nFlow inclination, $\\epsilon$, deg\nFlow inclination, $\\epsilon$, deg\n\n| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | 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| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | |", "timestamp": "2026-07-22T04:35:35.012848+00:00"}
{"citation_id": "19930086151", "source_url": "https://ntrs.nasa.gov/api/citations/19930086151/downloads/19930086151.pdf", "page_number": 21, "total_pages": 34, "image_filename": "19930086151_p21.jpg", "text": "```markdown\nNACA RM L9J28\n19\n\nPitching-moment coefficient, $C_m$\nDrag coefficient, $C_D$\nAngle of attack, $\\alpha$, deg\nLift coefficient, $C_L$\n\nCONFIDENTIAL\n\nWing-tip | Wing aspect\n---|---\naileron | ratio\n$\\circ$ parallelogram | 1.87\n$\\triangle$ Triangular | 2.31\n\nCONFIDENTIAL\nNACA\n\nFigure 3.— The aerodynamic characteristics in pitch of the 45° sweptback wing equipped with deflectable wing-tip ailerons. Plain wing $\\delta_a=0^\\circ$\n```", "timestamp": "2026-07-22T04:35:35.787197+00:00"}
{"citation_id": "19930085930", "source_url": "https://ntrs.nasa.gov/api/citations/19930085930/downloads/19930085930.pdf", "page_number": 81, "total_pages": 92, "image_filename": "19930085930_p81.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T04:35:36.425638+00:00"}
{"citation_id": "19930086078", "source_url": "https://ntrs.nasa.gov/api/citations/19930086078/downloads/19930086078.pdf", "page_number": 32, "total_pages": 42, "image_filename": "19930086078_p32.jpg", "text": "30\nNACA RM L9H04\n\nCONFIDENTIAL\n\n<!-- Image (108, 126, 879, 904) -->\n\n(b) Aileron pivoted at 0.267-chord station.\nFigure 10.- Concluded.", "timestamp": "2026-07-22T04:35:38.549931+00:00"}
{"citation_id": "19930086081", "source_url": "https://ntrs.nasa.gov/api/citations/19930086081/downloads/19930086081.pdf", "page_number": 29, "total_pages": 44, "image_filename": "19930086081_p29.jpg", "text": "```markdown\nNACA RM L9H05\n\nCONFIDENTIAL\n\n<!-- Image (113, 155, 382, 336) -->\n\n<!-- Image (568, 245, 837, 501) -->\n\n<!-- Image (113, 433, 382, 536) -->\n\nTip\nThickness\n(percent c) Fence\n2 Off\n3 Small\n5 Large\n7 Off\n\n<!-- Image (113, 633, 382, 794) -->\n\n<!-- Image (568, 600, 837, 794) -->\n\nCONFIDENTIAL\n\n(a) Small fuselage.\n\nFigure 10.- Variation of the aerodynamic characteristics of a semispan delta wing with deflection of\nthe half-delta tip control surface. R = $4.0 \\times 10^6$; M = 1.90; $\\alpha = 0^\\circ$.\n\n27\n```", "timestamp": "2026-07-22T04:35:38.736499+00:00"}
{"citation_id": "19930082485", "source_url": "https://ntrs.nasa.gov/api/citations/19930082485/downloads/19930082485.pdf", "page_number": 6, "total_pages": 62, "image_filename": "19930082485_p6.jpg", "text": "```markdown\nNACA TN No. 1810\n5\n\n$\\frac{\\rho_s V_x}{\\rho_t V_{cr}}$ weight-flow parameter\n\n$\\frac{r}{P_t} \\frac{dP_s}{dr}$ radial pressure-gradient parameter\n\n$\\frac{\\rho_s}{\\rho_t} \\left(\\frac{V_u}{V_{cr}}\\right)^2$ centrifugal-force parameter\n\n$\\frac{r}{V_{cr}} \\frac{\\rho_s}{\\rho_t} \\frac{dV_r}{dt}$ radial-acceleration parameter\n\nBLADE DESIGN\n\nThe turbine stator considered was designed for the following conditions: inlet-gas pressure, 8200 pounds per square foot; inlet-gas temperature, 1900° R, gas flow, 40.4 pounds per second; power output, 4300 horsepower, and rotor-blade speed at the 10.35-inch radius, 1000 feet per second. The weight-flow parameter, critical velocity ratio, and gas-flow angle at three radial stations are listed in the following table:\n\n| Station | Radius r (in.) | Weight-flow parameter | | Critical velocity ratio $\\left(\\frac{V}{V_{cr}}\\right)_e$ | Gas discharge angle $\\alpha_e$ (deg) |\n| :--- | :--- | :--- | :--- | :--- | :--- |\n| | | $\\left(\\frac{\\rho_s V_x}{\\rho_t V_{cr}}\\right)_i$ | $\\left(\\frac{\\rho_s V_x}{\\rho_t V_{cr}}\\right)_e$ | | |\n| Root | 9.3 | 0.212 | 0.233 | 0.878 | 67.8 |\n| Pitch | 10.35 | .212 | .249 | .800 | 65.6 |\n| Tip | 11.5 | .212 | .260 | .737 | 63.3 |\n\nThese values were computed on the assumptions of radial equilibrium, constant gas energy and entropy, and free-vortex flow from root to tip. These assumptions result in theoretically constant axial velocity at all radii (reference 6). Flow-area blockage due to boundary-layer development at the blade root and tip was considered by assuming, in accordance with the experimental data of reference 6, a boundary-layer-displacement thickness of 0.3 and 0.1 inch at the inner and outer shroud, respectively.\n```", "timestamp": "2026-07-22T04:35:39.683799+00:00"}
{"citation_id": "19930090382", "source_url": "https://ntrs.nasa.gov/api/citations/19930090382/downloads/19930090382.pdf", "page_number": 21, "total_pages": 37, "image_filename": "19930090382_p21.jpg", "text": "NACA RM L9I07\n23\n\nCONFIDENTIAL\n\nTip Mach number, $M_t$\nEfficiency, $\\eta$\n\nPower coefficient, $C_P$\nThrust coefficient, $C_T$\n\nAdvance ratio, J\n(f) M=0.60 Concluded.\nFigure 5 - Continued.\n\nCONFIDENTIAL\n$\\beta_{0.7R} = 55^\\circ$\n\nNACA", "timestamp": "2026-07-22T04:35:43.324537+00:00"}
{"citation_id": "19930086105", "source_url": "https://ntrs.nasa.gov/api/citations/19930086105/downloads/19930086105.pdf", "page_number": 22, "total_pages": 22, "image_filename": "19930086105_p22.jpg", "text": "20\nCONFIDENTIAL\nNACA RM E9H12\n\n<!-- Image (118, 141, 888, 890) -->\n\nFigure 7. - Comparative diffuser performance with regenerative-type burner and perforated conical flame holder.\n\nCONFIDENTIAL\nNACA-Langley - 10-13-49 - 350", "timestamp": "2026-07-22T04:35:43.564649+00:00"}
{"citation_id": "19930085842", "source_url": "https://ntrs.nasa.gov/api/citations/19930085842/downloads/19930085842.pdf", "page_number": 78, "total_pages": 104, "image_filename": "19930085842_p78.jpg", "text": "74\nNACA RM L9C29\n\n<!-- Image (107, 109, 850, 862) -->\n\n(a) $\\alpha_u = 3^\\circ$.\n\nFigure 43.- Variation of $C_L$, $C_{D_R}$, and $V/nD$ with $Q_c$ for several propeller blade angles. Basic model configuration; all control surfaces neutral.", "timestamp": "2026-07-22T04:35:45.494777+00:00"}
{"citation_id": "19930082245", "source_url": "https://ntrs.nasa.gov/api/citations/19930082245/downloads/19930082245.pdf", "page_number": 11, "total_pages": 66, "image_filename": "19930082245_p11.jpg", "text": "10\nNACA TN No. 1596\n\nthan those for the airfoil with the beveled-trailing-edge aileron. The pressure diagrams (not shown) indicated that at a deflection of 18° the air flow followed the contour of the true-contour aileron quite closely at the lower values of airfoil section normal-force coefficient and the lower Mach numbers. Separation occurred at the higher values of airfoil section normal-force coefficient and the higher Mach numbers. At a deflection of 12°, separation occurred at somewhat lower values of airfoil section normal-force coefficient and Mach number for the airfoil with the true-contour aileron than for the airfoil with the beveled-trailing-edge aileron.\n\nAs mentioned previously in the section entitled \"Apparatus and Methods\" the present tests were made with unsealed ailerons. Many tests have demonstrated that the characteristics of an airfoil with a sealed aileron are generally more satisfactory than those of an airfoil with an unsealed aileron. One of the unfavorable effects of an unsealed aileron gap is that the effectiveness of the aileron is less than when the aileron is sealed. The low-speed data of references 8 and 9 indicate that an unsealed aileron gap reduces aileron effectiveness to a greater extent on an airfoil with a beveled-trailing-edge aileron than on an airfoil with a true-contour aileron.\n\nAileron Section Hinge Moments\n\nThe section hinge-moment characteristics of the beveled-trailing-edge aileron are irregular as shown by the data of figure 10. In the aileron deflection range from approximately -6° to 4°, the variation of section hinge-moment coefficient was irregular both with Mach number and aileron deflection. At larger aileron deflections the characteristics were more satisfactory.\n\nBeveling the trailing edge reduced hinge moments by more than one-half at some of the test conditions at the larger deflections. The action of the thickened trailing edge in changing the air flow about the aileron and in relieving hinge moments is illustrated by the chordwise aileron pressure distributions shown in figure 16. At positive aileron deflections, the pressures on the aileron lower surface are usually more positive than those on the aileron upper surface. The bevel on the lower surface, where the pressures are more positive, speeds up the flow to a greater extent than the bevel on the upper surface; a hinge-moment component is thus introduced which acts in a way to relieve the main hinge moment. At negative deflections, the reverse action is generally true. From these tests the action of the bevel appeared to be greater at positive deflections.\n\nLarge increases in the section hinge-moment coefficient of the true-contour aileron at a deflection of 18° occurred for some of the combinations of Mach number and airfoil section normal-force coefficient. These increases in section hinge-moment coefficient were mainly due to development of appreciable separation of the flow off the upper surface", "timestamp": "2026-07-22T04:35:48.268277+00:00"}
{"citation_id": "19930093773", "source_url": "https://ntrs.nasa.gov/api/citations/19930093773/downloads/19930093773.pdf", "page_number": 11, "total_pages": 47, "image_filename": "19930093773_p11.jpg", "text": "```markdown\n10\nNACA RM E9G09\n\nAPPENDIX - CALCULATIONS\n\nSymbols\n\nThe following symbols were used in the calculations and on the figures:\n\n| Symbol | Definition |\n| :--- | :--- |\n| A | cross-sectional area, sq ft |\n| B | thrust scale reading, lb |\n| $C_j$ | jet-velocity coefficient, ratio of actual jet velocity or thrust to ideal velocity or thrust after expansion to free-stream static pressure |\n| $C_t$ | ratio of hot exhaust-nozzle area to cold exhaust-nozzle area (1.01 at $1570^\\circ$ R) |\n| D | external drag of installation, lb |\n| $D_r$ | exhaust-nozzle tail-rake drag, lb |\n| $D_w$ | windmilling drag, lb |\n| $F_j$ | jet thrust, lb |\n| $F_n$ | net thrust, lb |\n| f/a | fuel-air ratio |\n| g | acceleration due to gravity, $32.2 \\text{ ft/sec}^2$ |\n| M | flight Mach number |\n| N | engine speed, rpm |\n| P | total pressure, lb/sq ft absolute |\n| p | static pressure, lb/sq ft absolute |\n| R | gas constant, $53.3 \\text{ ft-lb/(lb)}(^\\circ\\text{R})$ |\n| T | total temperature, $^\\circ\\text{R}$ |\n| $T_1$ | indicated temperature, $^\\circ\\text{R}$ |\n```", "timestamp": "2026-07-22T04:35:48.799143+00:00"}
{"citation_id": "19930085965", "source_url": "https://ntrs.nasa.gov/api/citations/19930085965/downloads/19930085965.pdf", "page_number": 46, "total_pages": 67, "image_filename": "19930085965_p46.jpg", "text": "NACA RM E9E06\n\nFlux\npath\n\nVan es\n\nChopper\n\nCoil\n\nSection A-A\n\nLaminations\n\nRotor\n\nFigure 2. - Schematic diagram illustrating principle of eddy-current heating\napplied to inlet guide vanes of axial-flow compressor.\n\nNACA\n\n45", "timestamp": "2026-07-22T04:35:49.334574+00:00"}
{"citation_id": "19930082450", "source_url": "https://ntrs.nasa.gov/api/citations/19930082450/downloads/19930082450.pdf", "page_number": 10, "total_pages": 37, "image_filename": "19930082450_p10.jpg", "text": "NACA TN No. 1778\n9\n\nTABLE 1.- MATERIAL PROPERTIES OF 24S-T\nALUMINUM-ALLOY PANELS HAVING FORMED\nZ-SECTION STIFFENERS\n\n| | Aluminum alloy | $\\sigma_{cy}$ (ksi) |\n| :--- | :--- | :--- |\n| Sheet | 24S-T bare | 44.0 |\n| Stiffeners | 24S-T bare sheet before forming | 44.0 |\n\n[Figure: NACA logo]", "timestamp": "2026-07-22T04:35:49.629550+00:00"}
{"citation_id": "19930086015", "source_url": "https://ntrs.nasa.gov/api/citations/19930086015/downloads/19930086015.pdf", "page_number": 40, "total_pages": 54, "image_filename": "19930086015_p40.jpg", "text": "NACA RM A59E24\nCONFIDENTIAL\n\n2\nO Experimental survey\n— Angle variation from\npressure data\n1\nFlow inclination, $\\epsilon$, deg\n0\n-1\n-2\n-32 -24 -16 -8 0 8 16 24 32\nHorizontal distance from window center line, x, in.\nz=0\nNACA\n\n(d) D=113.43; M=1.53.\n\nFigure 11.—Continued.\nCONFIDENTIAL\n39", "timestamp": "2026-07-22T04:35:59.506208+00:00"}
{"citation_id": "19930082447", "source_url": "https://ntrs.nasa.gov/api/citations/19930082447/downloads/19930082447.pdf", "page_number": 10, "total_pages": 24, "image_filename": "19930082447_p10.jpg", "text": "8\nNACA TN No. 1775\n\nFigure 8 shows the variation of the dead-rise function with the angle of dead rise. Since the dead-rise function is plotted as $[f(\\beta)]^{2/3}$ this curve also shows the variation of hydrodynamic load with angle of dead rise. The solid-line curve shows the theoretical variation given by $[f(\\beta)]^{2/3} = (\\frac{\\pi}{2\\beta} - 1)^{2/3}$. The symbols represent experimental values of the dead-rise variation determined by averaging each group of data obtained with floats of $40^\\circ$, $30^\\circ$, and $22\\frac{1}{2}^\\circ$ angles of dead rise, respectively. The value shown for a dead-rise angle of $40^\\circ$ is the average corresponding to the distribution shown in figure 7. The average values for angles of dead rise of $30^\\circ$ and $22\\frac{1}{2}^\\circ$ were obtained in a similar manner and were presented in reference 4. Figure 8 shows that the hydrodynamic load decreases appreciably (44 percent) as the angle of dead rise is increased from $22\\frac{1}{2}^\\circ$ to $40^\\circ$. The variation of the average values of the experimental data agrees well with the theoretical variation.\n\nCONCLUSIONS\n\nAn analysis of experimental data obtained by subjecting a prismatic float having an angle of dead rise of $40^\\circ$ to impacts in smooth water results in the following conclusions:\n\n1. Experimental values of the load-factor coefficient, draft coefficient, time coefficient, and vertical-velocity ratio corresponding to the instants of maximum acceleration, maximum draft, and rebound are in good agreement with values predicted by hydrodynamic impact theory.\n\n2. If all other parameters are held constant, the hydrodynamic load for a seaplane having an angle of dead rise of $40^\\circ$ is 44 percent less than the hydrodynamic load for the seaplane with a conventional dead-rise angle of $22\\frac{1}{2}^\\circ$.\n\nLangley Aeronautical Laboratory\nNational Advisory Committee for Aeronautics\nLangley Field, Va., September 28, 1948", "timestamp": "2026-07-22T04:36:02.627575+00:00"}
{"citation_id": "19930085930", "source_url": "https://ntrs.nasa.gov/api/citations/19930085930/downloads/19930085930.pdf", "page_number": 82, "total_pages": 92, "image_filename": "19930085930_p82.jpg", "text": "CONFIDENTIAL\nUNCLASSIFIED\n\nNACA RM L9G07\n\n[Figure: A shadowgraph of the flow in the passage for model 4.]\n\n(a) Passage corresponding to characteristics net shown in figure 10(a).\n\nFigure 39.— A shadowgraph of the flow in the passage for model 4.\n\nCONFIDENTIAL\nUNCLASSIFIED\n\n81", "timestamp": "2026-07-22T04:36:03.161286+00:00"}
{"citation_id": "19930086151", "source_url": "https://ntrs.nasa.gov/api/citations/19930086151/downloads/19930086151.pdf", "page_number": 22, "total_pages": 34, "image_filename": "19930086151_p22.jpg", "text": "20\nNACA RM L9J28\n\nCONFIDENTIAL\n\nPitching-moment coefficient, $C_m$\nWing-tip aileron\no Parallelogram\n$\\Delta$ Triangular\nWing aspect ratio\n1.87\n2.31\nDrag coefficient, $C_D$\nAngle of attack, $\\alpha$, deg\nLift coefficient, $C_L$\n\nCONFIDENTIAL\nNACA\n\nFigure 4.—The aerodynamic characteristics in pitch of the 45° sweptback wing equipped with deflectable wing-tip ailerons. Wing with end plate; $\\delta_a=0^\\circ$.", "timestamp": "2026-07-22T04:36:03.446633+00:00"}
{"citation_id": "19930086078", "source_url": "https://ntrs.nasa.gov/api/citations/19930086078/downloads/19930086078.pdf", "page_number": 33, "total_pages": 42, "image_filename": "19930086078_p33.jpg", "text": "NACA RM L9H04\n31\n\nCONFIDENTIAL\n\n<!-- Image (117, 161, 905, 731) -->\n\n(a) $\\delta_a = 4^\\circ$.\n\nFigure 11.- Lateral control characteristics of unswept wing with short-chord wing-tip aileron at various extensions.", "timestamp": "2026-07-22T04:36:05.883164+00:00"}
{"citation_id": "19930085838", "source_url": "https://ntrs.nasa.gov/api/citations/19930085838/downloads/19930085838.pdf", "page_number": 91, "total_pages": 118, "image_filename": "19930085838_p91.jpg", "text": "NACA RM No. L9B23\n\n89\n\n| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | 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| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | |", "timestamp": "2026-07-22T04:36:05.923054+00:00"}
{"citation_id": "19930086081", "source_url": "https://ntrs.nasa.gov/api/citations/19930086081/downloads/19930086081.pdf", "page_number": 30, "total_pages": 44, "image_filename": "19930086081_p30.jpg", "text": "28\n\nCONFIDENTIAL\n\n$C_L$\n.1\n0\n-.1\n\n$C_m$\n0\n-.04\n\n$C_D$\n.02\n0\n-4 0 4 8 12 16\n$\\delta$, deg\n\n$C_t$\n.008\n.004\n0\n-.004\n\n$C_n$\n0\n-.004\n-4 0 4 8 12 16\n$\\delta$, deg\n\nTip\nThickness\n(percent c)\nPence\n7\nLarge\nOff\n7\n\nCONFIDENTIAL\n\n(b) Large fuselage.\nFigure 10.- Concluded.\n\nNACA RM 19H05", "timestamp": "2026-07-22T04:36:07.259440+00:00"}
{"citation_id": "19930082476", "source_url": "https://ntrs.nasa.gov/api/citations/19930082476/downloads/19930082476.pdf", "page_number": 8, "total_pages": 41, "image_filename": "19930082476_p8.jpg", "text": "6\nNACA TN No. 1801\n\nmately equal to the moment of inertia about the Y-axis $I_Y$ and the value\nof the inertia yawing-moment parameter $\\frac{I_X - I_Y}{mb^2}$ was thus approximately\nzero. For loading 2, the mass distribution along the fuselage was increased\nuntil the inertia yawing-moment parameter equaled $-49 \\times 10^{-4}$; and for\nloading 3, the mass distribution along the wings was increased until the\nvalue of the inertia yawing-moment parameter was $165 \\times 10^{-4}$. For loading 4,\nthe relative density of the model was approximately doubled by increasing\nthe weight and moments of inertia, keeping the radii of gyration about the\ncenter of gravity approximately the same as for loading 1. The mass-\ndistribution parameters for the four loading conditions given in table II\nare plotted in figure 4. Because of an inadvertent error in model\nballasting calculations, loading 2, although a possible light-airplane\nloading, is not the limit of the full range possible for airplanes that\nhave the weight distributed primarily along the fuselage, whereas loading 3\nprobably exceeds the range of loadings that might be expected for single-\nengine light airplanes having the greater part of the weight distributed\nalong the wings.\n\nAll tests were conducted with the canopy closed and with a fixed\nlanding gear installed on the model.\n\nIn order to simulate two-control operation now found on some light\nairplanes, the rudder and aileron controls were considered linked for some\nof the tests. The control deflections are given in terms of a control wheel\nand are as follows:\n\n<!-- Table (80, 564, 857, 748) -->\n\\begin{tabular}{|l|c|c|c|c|}\n\\hline\n\\multicolumn{1}{|c|}{Wheel position} & \\multicolumn{2}{c|}{Rudder deflection, deg} & \\multicolumn{2}{c|}{Aileron deflection, deg} \\\\\n\\cline{2-5}\n\\multicolumn{1}{|c|}{} & Left & Right & Left & Right \\\\\n\\hline\nFull right wheel & $1\\frac{3}{4}$ right & $27\\frac{1}{2}$ right & 5 down & $51\\frac{1}{2}$ up \\\\\nOne-half right wheel & $\\frac{1}{2}$ right & $8\\frac{1}{2}$ right & $9\\frac{3}{4}$ down & $21\\frac{1}{2}$ up \\\\\nOne-third right wheel & 3 right & $4\\frac{3}{4}$ right & $8\\frac{1}{2}$ down & $11\\frac{1}{2}$ up \\\\\nOne-fourth right wheel & $2\\frac{1}{2}$ right & $3\\frac{1}{2}$ right & 7 down & 8 up \\\\\n\\hline\n\\end{tabular}\n\nPlots of the control deflections for any wheel position are shown in\nfigure 5.\n\nNormal elevator deflections for the linked-control tests were chosen\nas $13^\\circ$ up and $12^\\circ$ down. The value of $13^\\circ$ up was chosen as the probable\nminimum value that would permit the corresponding airplane to be landed", "timestamp": "2026-07-22T04:36:09.690044+00:00"}
{"citation_id": "19930090382", "source_url": "https://ntrs.nasa.gov/api/citations/19930090382/downloads/19930090382.pdf", "page_number": 22, "total_pages": 37, "image_filename": "19930090382_p22.jpg", "text": "24\nNACA RM L9I07\n\nCONFIDENTIAL\n\nThrust coefficient, $C_T$\nPower coefficient, $C_P$\nTip Mach number, $M_t$\nEfficiency, $\\eta$\nAdvance ratio, J\n($\\frac{1}{8}$) M=0.65.\nFigure 5 - Continued.\n\nCONFIDENTIAL\n\nNACA", "timestamp": "2026-07-22T04:36:11.991448+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 1, "total_pages": 99, "image_filename": "19930082511_p1.jpg", "text": "AUTHOR'S PERSONAL COPY\n\nNACA TN No. 1826\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\n\nTECHNICAL NOTE\nNo. 1826\n\nLINEAR THEORY OF BOUNDARY EFFECTS IN OPEN\nWIND TUNNELS WITH FINITE JET LENGTH\n\nBy S. Katzoff, Clifford S. Gardner,\nLeo Diesendruck, and Bertram J. Eisenstadt\n\nLangley Aeronautical Laboratory\nLangley Air Force Base, Va.\n\n[Figure: NACA logo]\n\nWashington\nMarch 1949", "timestamp": "2026-07-22T04:36:14.282413+00:00"}
{"citation_id": "19930085842", "source_url": "https://ntrs.nasa.gov/api/citations/19930085842/downloads/19930085842.pdf", "page_number": 79, "total_pages": 104, "image_filename": "19930085842_p79.jpg", "text": "NACA RM L9C29\n75\n\n[Figure: A graph plotting three curves against a grid. The x-axis is labeled \"Torque coefficient, $Q_c$\" with values 0, .004, .008, .012, .016. The left y-axis is labeled \"Lift coefficient, $C_L$\" with values 0, 1, 2, 3. The upper left y-axis is labeled \"Propeller advance-diameter ratio, $V/nD$\" with values 10, 12, 14. The right y-axis is labeled \"Resultant drag coefficient, $C_{DR}$\" with values -1, 0, 1. The curves are labeled $V/nD$, $C_{DR}$, and $C_L$. An annotation in the lower right of the graph reads \"$\\beta$, deg 30\". The National Advisory Committee for Aeronautics logo is present.]\n\n(b) $\\alpha_u = 6^\\circ$.\nFigure 43.— Continued.", "timestamp": "2026-07-22T04:36:14.740776+00:00"}
{"citation_id": "19930082485", "source_url": "https://ntrs.nasa.gov/api/citations/19930082485/downloads/19930082485.pdf", "page_number": 7, "total_pages": 62, "image_filename": "19930082485_p7.jpg", "text": "6\nNACA TN No. 1810\n\nIn establishing a velocity distribution over the blade sur-\nface, flow acceleration over as much of the blade surface as\npossible is considered desirable. Thus, blade shapes at three\nradial stations, root, pitch, and tip (table I) were designed with\nthe stream-filament theory developed in appendix A by the procedure\noutlined in appendix B. Values of local deflection angle at cas-\ncade entrance, $v_i$ and local deflection angle at cascade exit $v_e$\nof $5^\\circ$ and $2^\\circ$, respectively, were used at all radii. The values\nwere graphically obtained from previous designs based on potential-\nflow theory (reference 3).\n\nThe blade coordinates and other dimensions are shown in table I.\nBlade shapes for a typical blade section with the velocity-potential\nlines used in calculating surface velocities are shown in figure 1.\nThe calculated surface velocities are plotted against blade-surface\nlength in figure 2.\n\nAPPARATUS AND PROCEDURE\n\nThe investigation was conducted with five blades mounted in a\nsector of an annular cascade to simulate three-dimensional flow. A\nsketch of the experimental equipment is shown in figure 3.\n\nThe experimental conditions were different from the design\nconditions in that the blades were designed for an inlet-gas tem-\nperature of $1900^\\circ$ R and the experiments were conducted at an inlet-\ngas temperature of $580^\\circ$ R. The Reynolds number (based on the pitch\nof the blade) for the turbine blades at design conditions was\n600,000; whereas, for the cascade experiments, the Reynolds number\nwas 800,000. The effect of this difference was considered small.\nThe blades were, however, investigated at the design value of\ncritical velocity ratio $V/V_{cr}$ near the pitch section of the blade.\n\nThe pressure distribution about the center blade of the cas-\ncade was measured with 36 static-pressure tubes of 0.030-inch\ndiameter located as indicated in table II. Pressures were measured\nwith mercury manometers.\n\nA survey probe, which is shown in the enlarged portion of\nfigure 3, was used to determine the angle at which the gas dis-\ncharged from the cascade of blades. The gas angle, measured from\nthe axial direction, was read on a protractor. The center tube of\nthe probe was used to indicate the stagnation pressure; the static\npressure was read by static-pressure holes located 9 diameters\ndownstream of the tip of the probe. The position of the end of the", "timestamp": "2026-07-22T04:36:17.167152+00:00"}
{"citation_id": "19930085965", "source_url": "https://ntrs.nasa.gov/api/citations/19930085965/downloads/19930085965.pdf", "page_number": 47, "total_pages": 67, "image_filename": "19930085965_p47.jpg", "text": "46\nNACA RM E9E06\n\n[Figure: A schematic diagram of a section of a plate. The diagram shows a rectangular block with various dimensions and labels. The vertical dimension is labeled \"Length\" with points A and A' at the bottom and B and B' at the top. The horizontal dimension is labeled \"Width\". The depth is labeled \"d\". A vertical arrow indicates the \"Direction of flux\". The top surface has points C and C'. Dimensions on the top surface include \"1 cm\", \"ds\", \"s\", and \"l_o\". Another \"1 cm\" dimension is shown on the right side. The NACA logo is present at the bottom right of the figure.]\n\nFigure 3. - Section of plate.", "timestamp": "2026-07-22T04:36:20.128520+00:00"}
{"citation_id": "19930086015", "source_url": "https://ntrs.nasa.gov/api/citations/19930086015/downloads/19930086015.pdf", "page_number": 41, "total_pages": 54, "image_filename": "19930086015_p41.jpg", "text": "40\nCONFIDENTIAL\nNACA RM A9E24\n\n<!-- Image (128, 178, 875, 776) -->\n\nFigure 11.-Continued.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T04:36:23.050365+00:00"}
{"citation_id": "19930085930", "source_url": "https://ntrs.nasa.gov/api/citations/19930085930/downloads/19930085930.pdf", "page_number": 83, "total_pages": 92, "image_filename": "19930085930_p83.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T04:36:23.360190+00:00"}
{"citation_id": "19930093773", "source_url": "https://ntrs.nasa.gov/api/citations/19930093773/downloads/19930093773.pdf", "page_number": 12, "total_pages": 47, "image_filename": "19930093773_p12.jpg", "text": "```markdown\nNACA RM E9G09\n11\n\nt static temperature, $^\\circ$R\nV velocity, ft/sec\n$W_a$ air flow, lb/sec\n$W_f$ fuel flow, lb/hr\n$W_f/F_n$ specific fuel consumption based on net thrust, lb/(hr)\n(lb thrust)\n$\\gamma$ ratio of specific heats\n$\\delta$ ratio of tunnel static pressure $P_0$ to absolute static\npressure of NACA standard atmosphere at sea level\n$\\delta_a$ ratio of tunnel static pressure $P_0$ to absolute static\npressure of NACA standard atmosphere at desired altitude\n$\\theta$ ratio of absolute equivalent ambient static temperature to\nabsolute static temperature of NACA standard atmosphere\nat sea level\n$\\theta_a$ ratio of absolute equivalent ambient static temperature to\nabsolute static temperature of NACA standard atmosphere\nat desired altitude\n\nSubscripts:\n0 free-air stream\n1 engine inlet\n6 turbine outlet\n7 1 inch upstream of exhaust-nozzle outlet\n8 exhaust-nozzle outlet\ne equivalent\nr venturi throat rake in make-up air duct\ns scale\nx inlet duct at frictionless slip joint\n```", "timestamp": "2026-07-22T04:36:24.045657+00:00"}
{"citation_id": "19930092013", "source_url": "https://ntrs.nasa.gov/api/citations/19930092013/downloads/19930092013.pdf", "page_number": 9, "total_pages": 21, "image_filename": "19930092013_p9.jpg", "text": "APPARATUS FOR VARYING EFFECTIVE DIHEDRAL IN FLIGHT 5\n\nan increase of both the area and the control throw of the tab, which was located on the left aileron. Brief flight tests with this revised tab yielded a value of $(d\\delta_a/d\\delta_t)_s$ of $-1.15$, which was used for the present tests. Although the standard aileron trim-tab drive linkages passed near the aileron servo-motor location in the cockpit, it was not possible to utilize this servo motor in obtaining the desired tab action $(d\\delta_a/d\\delta_t)_s$ because of excessive lost motion in the tab linkages between cockpit and tab surface. Therefore, a separate servo motor was installed in the left wing to drive the tab more directly. As indicated by figure 4, the tab servo electrical circuit is similar to the aileron circuit, although selsyns are used in place of potentiometers in the signal network. Error signals arising from rotation of the selsyn attached to the aileron servo result in motions of the tab servo and selsyn which tend to reduce the error to zero. The pilot is furnished with a trim-tab control which, when rotated, acts through differential gearing to rotate the aileron motor selsyn and, hence, the aileron tab.\n\nServomechanism controls and operating procedure.—The location of the aileron drive system and associated cockpit\n\n[Figure: View of cockpit interior showing aileron servo-drive and control components. Labels: Servo control box, Aileron servo position indicator, Pilot's trim tab control, Manual control for aileron servo lock, Stick force recorder, Aileron servo drive unit, Pivot for swinging plate, NACA]\n\ncontrols is shown in figure 5. When the apparatus is operated in flight, the error-signal circuits are energized first. The desired value of $(\\partial\\delta_a/\\partial\\delta_t)_s$ then is set with the servo-gearing ratio-selector switch, which gives values ranging from maximum positive to maximum negative in six approximately equal increments. The use of ammeters which indicate the aileron servo error signal reduce the possibility of abrupt motions which might occur if the servomotor were energized with the airplane at a sizable angle of sideslip. The pilot, by use of the rudder, reduces the error signal to zero and then places the entire system in operation by switching on the aileron amplidyne. Changes in $C_{l_\\beta}$ then are easily obtainable at any time by reducing the sideslip angle to zero and moving the servo-gearing selector switch. Both the aileron and the tab drives are equipped with limit switches and with locking and emergency drive circuits which permit the pilot to lock or return to neutral the torque tube and tab in the event of malfunctioning.\n\nINSTRUMENTATION\n\nStandard NACA photographically recording instruments were used to measure as a function of time the following variables: indicated airspeed; pressure altitude, applied aileron control force; angular positions of the aileron surfaces, aileron tab, aileron servo drum, forward portion of the aileron torque tube in the horizontal plane, rudder, and stick; sideslip angle; and airplane rolling and yawing velocities. A free-swivelling pitot-static head mounted on a boom extending forward from the right wing tip was used for airspeed and altitude measurements. The recording sideslip vane was mounted on a boom from the right wing at approximately the same location relative to the wing as the vane on the left wing tip for the dihedral apparatus. (See fig. 1.)\n\n1. EVALUATION OF THE EFFECTIVE-DIHEDRAL CONTROL APPARATUS\n\nThe results presented in this part of the report are based on data obtained during the first flights of the test airplane made with the complete effective-dihedral control apparatus in operation. The primary purpose of these early tests was to determine, from recorded data and pilots' opinions, the ability of the apparatus to simulate changes in stick-fixed and stick-free dihedral effect under static and dynamic flight conditions.\n\nTESTS AND RESULTS\n\nAlthough data were obtained at several airspeeds and values of servo gearing $(\\partial\\delta_a/\\partial\\delta_t)_s$, results presented herein are confined to the normal airplane (servo inoperative) and to the maximum initial test values of $(\\partial\\delta_a/\\partial\\delta_t)_s$ or $(\\Delta C_{l_\\beta})_s$ at a nominal indicated airspeed of 300 knots. The data presented are typical, and these test conditions approximate those originally considered in the design of the apparatus. Operation under static flight conditions was studied in steady straight sideslips and under dynamic conditions in abrupt rudder kicks and cockpit-controls-fixed lateral oscillations.\n\nSteady straight sideslips.—The aileron and tab deflections supplied by servo action and the net balancing aileron deflection and stick force supplied by the pilot are plotted in figure 6 as a function of sideslip angle for the three test servo-gearing ratios. All quantities represent changes from the wings-level trim condition. Corrections for distortion in the aileron servo drive system (between the aileron servo-motor and the torque tube) have been made.", "timestamp": "2026-07-22T04:36:25.863542+00:00"}
{"citation_id": "19930086151", "source_url": "https://ntrs.nasa.gov/api/citations/19930086151/downloads/19930086151.pdf", "page_number": 23, "total_pages": 34, "image_filename": "19930086151_p23.jpg", "text": "```markdown\nNACA RM L9J28\n\nCONFIDENTIAL\n\nYawing-moment coefficient, $C_n$\nRolling-moment coefficient, $C_l$\n\n| $\\delta_a$ (deg) |\n| :--- |\n| 21 |\n| 42 |\n| 60 |\n| 96 |\n| 147 |\n| 195 |\n| 235 |\n\nCONFIDENTIAL\n\nNACA\n\nAngle of attack, $\\alpha$, deg\n\nFigure 5.— The rolling-moment and yawing-moment characteristics of the 45° sweptback wing for various deflections of the triangular wing-tip aileron. Plain wing.\n\n21\n```", "timestamp": "2026-07-22T04:36:28.209016+00:00"}
{"citation_id": "19930085838", "source_url": "https://ntrs.nasa.gov/api/citations/19930085838/downloads/19930085838.pdf", "page_number": 92, "total_pages": 118, "image_filename": "19930085838_p92.jpg", "text": "90\nNACA RM No. 19B23\n\n<!-- Image (115, 109, 846, 999) -->\n\nFlap section hinge-moment coefficient, $c_{h1}$\n\nSection angle of attack, $\\alpha_o$, deg\n\n(g) $\\delta_f = 40^\\circ$.\nFigure 11.- Continued.", "timestamp": "2026-07-22T04:36:28.429670+00:00"}
{"citation_id": "19930082447", "source_url": "https://ntrs.nasa.gov/api/citations/19930082447/downloads/19930082447.pdf", "page_number": 11, "total_pages": 24, "image_filename": "19930082447_p11.jpg", "text": "NACA TN No. 1775\n\n9\n\nREFERENCES\n\n1. Milwitzky, Benjamin: A Generalized Theoretical and Experimental Investigation of the Motions and Hydrodynamic Loads Experienced by V-Bottom Seaplanes During Step-Landing Impacts. NACA TN No. 1516, 1948.\n\n2. Batterson, Sidney A.: The NACA Impact Basin and Water Landing Tests of a Float Model at Various Velocities and Weights. NACA Rep. No. 795, 1944.\n\n3. Mayo, Wilbur L.: Theoretical and Experimental Dynamic Loads for a Prismatic Float Having an Angle of Dead Rise of $22\\frac{1}{2}^\\circ$. NACA RB No. L5F15, 1945.\n\n4. Miller, Robert W., and Leshover, Samuel: Hydrodynamic Impact Loads in Smooth Water for a Prismatic Float Having an Angle of Dead Rise of $30^\\circ$. NACA TN No. 1325, 1947.\n\n5. Mayo, Wilbur L.: Analysis and Modification of Theory for Impact of Seaplanes on Water. NACA Rep. No. 810, 1945.\n\n6. Wagner, Herbert: Über Stoss- und Gleitvorgänge an der Oberfläche von Flüssigkeiten. Z.f.a.M.M., Bd. 12, Heft 4, Aug. 1932, pp. 193-215.\n\n7. Pabst, Wilhelm: Theory of the Landing Impact of Seaplanes. NACA TM No. 580, 1930.", "timestamp": "2026-07-22T04:36:29.869420+00:00"}
{"citation_id": "19930082245", "source_url": "https://ntrs.nasa.gov/api/citations/19930082245/downloads/19930082245.pdf", "page_number": 12, "total_pages": 66, "image_filename": "19930082245_p12.jpg", "text": "NACA TN No. 1596\n\nof the aileron, the separation occurring at lower Mach numbers as the airfoil section normal-force coefficient was increased (fig. 7).\n\nThe section hinge-moment slope $\\left(\\frac{\\Delta c_{\\mathrm{h}}}{\\Delta \\delta_{\\mathrm{a}}}\\right)_{\\alpha=0^{\\circ}}$ for the true-contour aileron at moderate deflections increased in magnitude between the Mach numbers of 0.25 and 0.68 by about three-fourths of the low-speed value (fig. 15). At higher Mach numbers up to the maximum test Mach number of 0.75, there was a reduction in magnitude of this hinge-moment parameter. The effect of beveling the aileron trailing edge to an angle of $30^{\\circ}$ was to cause an overbalance of the section parameter $\\left(\\frac{\\Delta c_{\\mathrm{h}}}{\\Delta \\delta_{\\mathrm{a}}}\\right)_{\\alpha=0^{\\circ}}$ at moderate deflections similar to that shown by other tests (references 8 to 10), and this over-balance was aggravated with Mach number (fig. 15). The section hinge-moment slope $\\left(\\frac{\\Delta c_{\\mathrm{h}}}{\\Delta \\alpha}\\right)_{\\delta=0^{\\circ}}$ for the true-contour aileron was essentially constant up to a Mach number of 0.68 and then rapidly increased in magnitude at higher Mach numbers up to the maximum test Mach number of 0.75. The section parameter $\\left(\\frac{\\Delta c_{\\mathrm{h}}}{\\Delta \\alpha}\\right)_{\\delta=0^{\\circ}}$ for the beveled-trailing-edge aileron was positive in algebraic sign and appreciably increased in magnitude with Mach number.\n\nLow-speed tests (references 9 and 10) have shown that reducing the aileron gap or sealing the gap of a beveled aileron had small effect on the parameter $\\left(\\frac{\\Delta c_{\\mathrm{h}}}{\\Delta \\alpha}\\right)_{\\delta}$, but reduced the overbalance of the parameter $\\left(\\frac{\\Delta c_{\\mathrm{h}}}{\\Delta \\delta_{\\mathrm{a}}}\\right)_{\\alpha}$ at small aileron deflections.\n\nSection Normal Force\n\nAirfoil.- The section slope $\\left(\\frac{\\Delta c_{\\mathrm{n}}}{\\Delta \\alpha}\\right)_{\\delta_{\\mathrm{a}}=0^{\\circ}}$ for both configurations increased with Mach number up to a Mach number somewhat greater than the critical Mach number and then decreased with further increase in Mach number (figs. 11 and 13). The section slope $\\left(\\frac{\\Delta c_{\\mathrm{n}}}{\\Delta \\delta_{\\mathrm{a}}}\\right)_{\\alpha=0^{\\circ}}$ for both configurations was not affected very much by Mach number at subcritical speeds, and decreased at supercritical speeds (fig. 11).\n\nThe effect of beveling the trailing edge was to reduce the section slopes $\\left(\\frac{\\Delta c_{\\mathrm{n}}}{\\Delta \\alpha}\\right)_{\\delta_{\\mathrm{a}}=0^{\\circ}}$ and $\\left(\\frac{\\Delta c_{\\mathrm{n}}}{\\Delta \\delta_{\\mathrm{a}}}\\right)_{\\alpha=0^{\\circ}}$ and this effect is in qualitative agreement with low-speed two-dimensional and three-dimensional tests", "timestamp": "2026-07-22T04:36:32.288686+00:00"}
{"citation_id": "19930086081", "source_url": "https://ntrs.nasa.gov/api/citations/19930086081/downloads/19930086081.pdf", "page_number": 31, "total_pages": 44, "image_filename": "19930086081_p31.jpg", "text": "Unpublished data from rocket tests\n- - - - - Calculation based on linearized theory\n\nCONFIDENTIAL\n\nNACA RM L9H05\n\n$$\\frac{\\rho b}{2V} \\delta$$\n\n.012\n.008\n.004\n0\n\nWind-tunnel tests\n$C_{l\\delta}$ Figure 10\n$C_{lp}$ Reference 3\n\n$\\delta = 7.6^\\circ$\n$\\delta = 5.5^\\circ$\n\n1.0 1.2 1.4 1.6 1.8 2.0\nMach number\n\nCONFIDENTIAL\n\nFigure 11.- Comparison of free-flight rocket test results and wind-tunnel test results with linearized theory for half-delta tip control surfaces on a delta wing-fuselage combination.\n\n29", "timestamp": "2026-07-22T04:36:33.982231+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 2, "total_pages": 99, "image_filename": "19930082511_p2.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T04:36:35.953011+00:00"}
{"citation_id": "19930082487", "source_url": "https://ntrs.nasa.gov/api/citations/19930082487/downloads/19930082487.pdf", "page_number": 6, "total_pages": 33, "image_filename": "19930082487_p6.jpg", "text": "```markdown\n4\nNACA TN No. 1813\n\nMeasurements at Supercritical Speeds for\nNACA 23015 Airfoil at $2^\\circ$ Incidence\n\nLift- and drag-coefficient variation with Mach number for the\nNACA 23015 airfoil section at $2^\\circ$ angle of attack is presented in\nfigure 1. For this section at $2^\\circ$ angle of attack, the critical Mach\nnumber, which is the free-stream Mach number at which sonic velocity\nis first attained at some point on the airfoil surface, is 0.59.\nThe first abrupt change in the section characteristics occurs at a\nsomewhat higher free-stream Mach number, and consists of a drag-\ncoefficient increase from the low-speed value. That free-stream\nMach number at which the rate of increase of drag coefficient with\nMach number equals 0.1 is defined, as in reference 1, as the drag-\ndivergence Mach number. The free-stream Mach number at which the\nexperimental pressure distribution first indicates the presence of\nmarked boundary-layer thickness is termed the shock-stall Mach number.\nMore detailed information about the supercritical flow changes over\nthe NACA 23015 airfoil section at $2^\\circ$ angle of attack is obtained from\nthe schlieren pictures and corresponding pressure distributions pre-\nsented in figure 2. At a Mach number of 0.60 there is a small region\nof supersonic flow which contains alternate expansion and compression\nregions, as can be seen from the schlieren picture. The pressure\ndistribution resembles that for subcritical speeds. When the free-\nstream Mach number is increased to 0.65, the drag-divergence Mach\nnumber, a strong shock wave appears near the airfoil crest (the\nchordwise station at which the surface is tangent to the free-stream\ndirection) but no change in the boundary-layer thickness is apparent\nin the schlieren photograph. From the pressure distribution, it is\nseen that at the airfoil crest the local velocity is greater than\nsonic. Thus, sonic velocity is reached at the airfoil crest at a Mach\nnumber between 0.60 and 0.65. In figures 2(b), 2(c), and 2(d) the\nterminal shock wave stands close to the chordwise station at which the\nlocal pressure coefficient corresponds to sonic velocity. At 0.70 and\n0.73 free-stream Mach numbers, there is a marked thickening of the\nboundary layer ahead of the shock wave, which is located aft of the\nairfoil crest. This rapid thickening of the boundary layer appears to\nstart at the airfoil crest. The pressure distribution over the rear\nof the upper surface of the airfoil for free-stream Mach numbers of\n0.70 and 0.73 resembles that for the smaller Mach numbers and thus\ngives no indication of the presence of a thick boundary layer.\nHowever, the schlieren photograph for a free-stream Mach number of 0.73\ndoes indicate the presence of a thick boundary layer over the rear of\nthe airfoil upper surface. The drag coefficient at this Mach number is\nabout 0.04, which is comparable to the drag coefficient for low-speed\nmaximum lift at the same Reynolds number.\n\nFigure 3(a) shows the variation of pressure coefficient with Mach\n```", "timestamp": "2026-07-22T04:36:40.329837+00:00"}
{"citation_id": "19930085842", "source_url": "https://ntrs.nasa.gov/api/citations/19930085842/downloads/19930085842.pdf", "page_number": 80, "total_pages": 104, "image_filename": "19930085842_p80.jpg", "text": "76\nNACA RM L9C29\n\n<!-- Image (97, 110, 868, 800) -->\n\n(c) $\\alpha_u = 12^\\circ$.\nFigure 43.- Continued.", "timestamp": "2026-07-22T04:36:41.000882+00:00"}
{"citation_id": "19930090382", "source_url": "https://ntrs.nasa.gov/api/citations/19930090382/downloads/19930090382.pdf", "page_number": 23, "total_pages": 37, "image_filename": "19930090382_p23.jpg", "text": "NACA RM L9I07\n25\n\nCONFIDENTIAL\n\nTip Mach number, $M_t$\n1.5\n1.0\n0.5\n0\n\nEfficiency, $\\eta$\n1.00\n.75\n.50\n.25\n0\n\nPower coefficient, $C_P$\n1.20\n1.10\n1.00\n.90\n.80\n.70\n.60\n.50\n.40\n.30\n.20\n.10\n0\n\nThrust coefficient, $C_T$\n.300\n.275\n.250\n.225\n.200\n.175\n.150\n.125\n.100\n.075\n.050\n.025\n0\n\nAdvance ratio, J\n0 .5 1.0 1.5 2.0 2.5 3.0 3.5 4.0 4.5 5.0 5.5 6.0 6.5 7.0 7.5 8.0 8.5 9.0\n\nCONFIDENTIAL\n$\\beta_0.7R = .65$\n\n(g) M=0.85. Concluded.\nFigure 5 - Continued.", "timestamp": "2026-07-22T04:36:42.590963+00:00"}
{"citation_id": "19930085965", "source_url": "https://ntrs.nasa.gov/api/citations/19930085965/downloads/19930085965.pdf", "page_number": 48, "total_pages": 67, "image_filename": "19930085965_p48.jpg", "text": "1125\n\nNACA RM E9E06\n\nMagnetic-flux density, gausses\n\n15,000\n10,000\n5,000\n\n0\nHmax 1\nHmax 2\n\nMagnetic-field intensity, oersteds\n\n10\n20\n30\n40\n50\n\n[Figure: Normal magnetization curve for SAE 1020 steel annealed at 1500° F and furnace-cooled. (Data from reference 5.)]\n\nNACA\n\n47", "timestamp": "2026-07-22T04:36:42.778737+00:00"}
{"citation_id": "19930082485", "source_url": "https://ntrs.nasa.gov/api/citations/19930082485/downloads/19930082485.pdf", "page_number": 8, "total_pages": 62, "image_filename": "19930082485_p8.jpg", "text": "NACA TN No. 1810\n7\n\nprobe was varied in the discharge field by manipulation of the traversing mechanism. The static pressures at the inner and outer shrouds were read by wall taps located midway between blades in line with the trailing edge.\n\nA preliminary survey of the tunnel was conducted to assure a uniform pressure field upstream of the blades. The stagnation- and static-pressure distribution across the tunnel along the survey paths (indicated in fig. 4) is shown in figure 5 for a plane 1.5 chord lengths upstream of the blades. Inasmuch as no boundary-layer-control slots were provided, the boundary-layer development upstream of the blades along the inner and outer shrouds was investigated. The approach was long enough to provide uniform flow and short enough to prevent excessive boundary-layer growth, as seen in figure 5. The variation of weight-flow parameter at the cascade entrance with the ratio of stagnation pressure at the cascade entrance to ambient static pressure is shown in figure 6. The test condition is indicated.\n\nThe following measurements were made to determine blade performance:\n\n1. Stagnation and static-pressure surveys, both upstream and downstream of blades\n2. Stagnation upstream temperature (held constant at 580° R)\n3. Static-pressure distribution about center blade of cascade\n4. Survey of angle of discharge of blade\n\nThe downstream surveys were made in a plane 0.1 chord length behind the blades. The surveys were made close to the blades because the inner and outer shrouds extended only 1 chord length downstream of the blades.\n\nRESULTS AND DISCUSSION\n\nThe stagnation and static pressures at simulated design conditions in a survey plane 0.1 chord length downstream of cascade at points midway between blades at a number of radial stations are shown in figure 7. The construction of the survey equipment limited the positioning of the probe to radii 0.3 inch less than that of the outer shroud. The data of figure 7 are plotted as the ratio of stagnation-to-static pressure in figure 8. The design value of", "timestamp": "2026-07-22T04:36:45.457043+00:00"}
{"citation_id": "19930082476", "source_url": "https://ntrs.nasa.gov/api/citations/19930082476/downloads/19930082476.pdf", "page_number": 9, "total_pages": 41, "image_filename": "19930082476_p9.jpg", "text": "NACA TN No. 1801\n\nsatisfactorily. Elevator deflections of $20^\\circ$ and $30^\\circ$ up were also tested, however, to determine the effect of increased up elevator deflections. In addition, tests were made with the controls unlinked to determine the independent effects of the rudders and ailerons.\n\nRESULTS AND DISCUSSION\n\nThe results of the spin tests of the model with linked-control settings are presented in charts 1 to 4 and with unlinked-control settings in charts 5 to 8. The normal-spinning-control configuration for a two-control airplane having linked rudders and ailerons is different from that for an airplane utilizing a three-control system: For the two-control airplane, ailerons and rudders are both moved with the spin for normal entry into a spin; whereas, for the conventional airplane, the ailerons would be placed at neutral and only the rudders would be moved with the spin. The model data given in the charts are presented in terms of the full-scale values for a corresponding airplane at a test altitude of 5000 feet.\n\nPreliminary tests of the model showed that steady-spin data for left and right spins differed very little. Results are, therefore, arbitrarily presented in terms of equivalent right spins, that is, for the airplane turning to the pilot's right.\n\nLinked Controls\n\nNormal loading (loading 1).— The test results obtained with the model in the normal-loading condition with linked rudders and ailerons simulated are presented in chart 1. The model condition is represented by loading 1 in table II and point 1 in figure 4. For the normal-control configuration for spinning (wheel full with the spin and elevator at its normal full-up deflection of $13^\\circ$), the model did not reach a spin equilibrium but descended at a steep attitude in a wide radius in the tunnel and at a vertical velocity exceeding the maximum tunnel velocity. The motion appeared to be a steep spiral rather than a spin. Film-strip photographs of the typical model motion at this control configuration are shown in figure 6. When the wheel was set at only one-half with the spin, however, definite spins were obtainable at up elevator deflections of $8^\\circ$ and higher. Photographs of the model during a typical spin with the wheel set at this position and with the elevator set at its normal full-up deflection ($13^\\circ$) are shown in figure 7. No recoveries were attempted from these spins; but when the model was launched into the tunnel with the wheel set at neutral or against the spin at the various up elevator deflections for which spins were obtained, the original rotation imparted to the model on launching damped out rapidly; recoveries from any spins were thus indicated to be satisfactory when the wheel was moved to neutral or against the spin.", "timestamp": "2026-07-22T04:36:50.266774+00:00"}
{"citation_id": "19930086078", "source_url": "https://ntrs.nasa.gov/api/citations/19930086078/downloads/19930086078.pdf", "page_number": 34, "total_pages": 42, "image_filename": "19930086078_p34.jpg", "text": "32\nNACA RM L9H04\n\nCONFIDENTIAL\n\n<!-- Image (196, 119, 868, 935) -->\n\n(b) $\\delta_a = 8^\\circ$.\nFigure 11.- Concluded.", "timestamp": "2026-07-22T04:36:52.835185+00:00"}