PDF/ocr_dataset_1048.jsonl

{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 53, "total_pages": 96, "image_filename": "19930085880_p53.jpg", "text": "NACA RM No. L9C03\n51\n\n[Figure: A graph plotting Draft, ft (y-axis) against Wetted area, sq ft (x-axis). The y-axis ranges from 0 to .64. The x-axis ranges from 0 to .35. A legend indicates Speed (fps) with symbols: 10 O, 15 □, 20 ◇, 25 △, 30 ▽. Data points follow a generally increasing trend. A NACA logo is present near the x-axis label.]\n\n(b) $\\tau = 8^\\circ$.\nFigure 17.- Continued.", "timestamp": "2026-07-22T05:49:52.356039+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 20, "total_pages": 98, "image_filename": "19930086073_p20.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:49:54.302534+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 6, "total_pages": 78, "image_filename": "19930082483_p6.jpg", "text": "4\nNACA TN No. 1807\n\nSubscripts m, u, 1, and 2 refer to pitch line, tangential,\nrotor inlet, and rotor discharge, respectively. Barred symbols\nrefer to average values.\n\nBlade power and gross power are the same for a full-admission\nturbine. They are different, however, for a turbine operating with\npartial admission as will be subsequently shown.\n\nA part of the blade power produced is so dissipated by shaft\nlosses that the available net power for a full-admission turbine\nmay be written\n\nnet power = ideal power - rotor-tip leakage losses\n- aerodynamic nozzle and rotor-blading losses\n- shaft losses\n(3)\n\nBlade power can be described by\n\nblade power = net power + shaft losses\n(4)\n\nSummarized in schematic form in figure 1(a) are the relations\nbetween the various concepts of turbine power for full admission.\n\nFor greater utility, the performance data are presented as\ncorrected to a constant reference state. The reference state com-\nmonly chosen and used herein is standard sea-level conditions.\n\nAt constant Mach number\n\n$$\n\\frac{V}{a} = \\frac{V}{\\sqrt{\\gamma gRT}} = \\frac{V_0}{a_0} = \\frac{V_0}{\\sqrt{\\gamma gRT_0}}\n$$\n(5)\n\nwhere\n\n| | |\n| :--- | :--- |\n| V | any gas velocity term observed, (ft/sec) |\n| a | velocity of sound at observed conditions, (ft/sec) |\n| $\\gamma$ | ratio of specific heats of gas |\n| g | acceleration due to gravity, 32.174, (ft/sec$^2$) |\n| R | gas constant, 53.345, (ft-lb)/(lb)($^\\circ$F) |\n| T | temperature observed, ($^\\circ$R) |", "timestamp": "2026-07-22T05:49:55.074289+00:00"}
{"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 62, "total_pages": 78, "image_filename": "19930082618_p62.jpg", "text": "```markdown\nR\n0.7 x 10⁶\n○ 1.0\n□ 1.5\n◇ 2.0\n△ 3.0\n▽ 4.0\n◁ 5.0\n▷ 6.0\n◁ 9.0\n\nFlagged symbols denote standard roughness\n\nSection lift coefficients, cₗ\n2.0\n1.6\n1.2\n.8\n.4\n0\n-.4\n-.8\n\nMoment coefficients, cₘ\n.2\n0\n-.1\n-.2\n-.3\n\nSection angle of attack, αₛ, deg\n-24 -16 -8 0 8 16 24\n\n(a) Section lift and pitching-moment characteristics of the plain airfoil section.\nFigure 14.— Aerodynamic characteristics of the NACA 23012 airfoil section, 24-inch chord.\n\nNACA\nNACA TN 1945\n60\n```", "timestamp": "2026-07-22T05:49:55.259016+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 80, "total_pages": 99, "image_filename": "19930082511_p80.jpg", "text": "78\nNACA TN No. 1826\n\n[Figure: (a) Vortex on the center line.]\n\n[Figure: (b) Unsymmetrically located vortex (incomplete representation).]\n\n[Figure: (c) Expanding or contracting jet.]\n\nFigure 8.- Velocity-potential analogies for the two-dimensional closed-open-closed tunnel.", "timestamp": "2026-07-22T05:50:00.886085+00:00"}
{"citation_id": "19930082474", "source_url": "https://ntrs.nasa.gov/api/citations/19930082474/downloads/19930082474.pdf", "page_number": 10, "total_pages": 21, "image_filename": "19930082474_p10.jpg", "text": "8\nNACA TN No. 1799\n\nfound to be actually less than 0.1 second for the subject rotors, a time period too short for perception by the pilot. Correspondingly, after the stick reaches its position following an abrupt lateral deflection only about 0.1 second elapses before the fuselage attains maximum angular acceleration in roll. Experience gained from airplane handling-qualities studies indicates that this is a satisfactory response; in fact, airplane requirements allow 0.2 second (reference 5). The helicopter also approaches a steady rate of roll in about the same time as does an airplane. The impression of lag when hovering, therefore, seems to arise from the fact that velocity changes or displacement of the helicopter in space do not follow the inclination of the thrust vector immediately, because of the mass of the machine. A similar apparent lag effect occurs in airplane formation flying where the problem is to control the rate of closure. The pilot overcomes his first impressions of lag during training by learning to control the helicopter's accelerations.\n\nIn hovering the helicopter also drifts back and forth as a result of the motions of the air. Some drift has to be expected of any aircraft since it is supported by the air. The stability of the machine with respect to speed and the directional stability in connection with yawing motions, both of which are desirable in other respects, increase the tendency to move or yaw with changes in wind velocity or direction. In this respect, reduction of stability can be beneficial.\n\nIn hovering, control-fixed lateral and longitudinal oscillations have been found to build up rapidly in amplitude per cycle. Since the machine performs an oscillation, a restoring tendency following a disturbance exists due to stability with speed. The restoring tendency itself is beneficial, provided the period of the motion is long enough to allow for the pilot's reaction time in perceiving and correcting the motion. The longitudinal period for the helicopter was found to be about 14 seconds, while the lateral period was about 6 seconds, a considerably shorter time. From experience gained from airplane handling-qualities studies and from personal experience with this and some other helicopters the period of the lateral motion is considered great enough to eliminate it as a control problem.\n\nIsolated Flight Phenomena\n\nEarly in the rotor performance investigations a phenomenon in connection with vertical flight was encountered. In determining the power required at zero airspeed with varying rates of descent, a region was encountered in which control of the machine could not be maintained. The descents were entered from forward flight with fixed power, and when zero airspeed was reached the rate of descent was low. If the power was insufficient to maintain descent at less than 500 feet per minute (as indicated by a standard rate-of-climb indicator) the machine would slowly increase its vertical velocity. At an indicated rate of descent of about 500 feet per minute, shaking of the machine became quite pronounced. Rather", "timestamp": "2026-07-22T05:50:06.581860+00:00"}
{"citation_id": "19930093789", "source_url": "https://ntrs.nasa.gov/api/citations/19930093789/downloads/19930093789.pdf", "page_number": 9, "total_pages": 29, "image_filename": "19930093789_p9.jpg", "text": "8\nCONFIDENTIAL\nNACA RM No. E8I21\n\n(5) Ratio of equivalent mean rotor-blade speed to equivalent\nmass flow, $\\frac{U_{D1}'}{w_{P0}}$ (This is the reciprocal of mass flow\nper unit blade speed.)\n(6) Ratio of mean rotor-blade speed to ideal jet speed (blade-\nto-jet speed ratio), $U/V_j$ (U/Vj)\n\nThe brake efficiency of the turbine, which includes the mechan-\nical efficiency of the turbine, is defined as\n$$\n\\eta = \\frac{E}{\\Delta_s h_{1,3}'}\n$$\nThe weight flow of air $w$ was determined from the orifice measure-\nments and the data of reference 2.\n\nThe ideal enthalpy drop $\\Delta_s h_{1,3}'$ was computed using the\nchart of air properties in reference 3 and the values of entrance\ntotal pressure and temperature and exit total pressure, $P_1'$, $T_1'$,\nand $P_3'$, respectively. Because the exit total pressure $P_3'$\nwas not directly measured, this value was computed. The exit\nstatic pressure $P_3$, the exit total temperature $T_3'$, the weight\nflow $w$, and the annular area between the exhaust-guide shells\nwere used to determine the exit total pressure $P_3'$ by adding to\nthe measured static pressure a dynamic pressure computed from con-\ntinuity with the assumptions that the tangential velocity is zero\nand the axial velocity is constant; these assumptions result in\nefficiency values that are always conservative.\n\nThe ideal jet speed $V_j$ is the jet speed for an isentropic\nexpansion from the entrance total state to the exit static pres-\nsure, that is,\n$$\nV_j = \\sqrt{2Jg (h_1' - h_3)_s}\n$$\nThe isentropic enthalpy drop $(h_1' - h_3)_s$ was computed from refer-\nence 3 using the entrance total temperature and pressure and the\nexit static pressure, $P_1'$, $T_1'$, and $P_3$, respectively.\n\nRESULTS AND DISCUSSION\n\nThe performance characteristics over the range of operation\ncovered in this investigation are shown for configurations 1, 2,\nand 3 in figures 11, 12, and 13, respectively. For each of the\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:50:10.190453+00:00"}
{"citation_id": "19930086076", "source_url": "https://ntrs.nasa.gov/api/citations/19930086076/downloads/19930086076.pdf", "page_number": 22, "total_pages": 50, "image_filename": "19930086076_p22.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:50:22.390911+00:00"}
{"citation_id": "19930085977", "source_url": "https://ntrs.nasa.gov/api/citations/19930085977/downloads/19930085977.pdf", "page_number": 24, "total_pages": 33, "image_filename": "19930085977_p24.jpg", "text": "CONFIDENTIAL\n\nNACA RM L9H22\n\nM\n1.15 △\n1.10 ▽\n1.08 ⊗\n1.05 △\n1.03 ▽\n1.00 ⊗\n.98 △\n.95 ▽\n.93 ⊗\n.90 △\n.88 ▽\n.85 ⊗\n.80 △\n.70 ▽\n.60 ⊗\n\nAngle of attack, $\\alpha$, deg\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n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"timestamp": "2026-07-22T05:50:25.372757+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 54, "total_pages": 96, "image_filename": "19930085880_p54.jpg", "text": "52\nNACA RM No. L9C03\n\n.64\nSpeed\n(fps)\n10 O\n15 □\n20 ◇\n25 △\n30 ▽\n.56\n.48\n.40\nDraft, ft\n.32\n.24\n.16\n.08\n0\n0 .05 .10 .15 .20 .25 .30 .35\nWetted area, sq ft\nNACA\n(c) $\\tau = 12^\\circ$.\nFigure 17.- Continued.", "timestamp": "2026-07-22T05:50:26.606698+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 21, "total_pages": 114, "image_filename": "19930086061_p21.jpg", "text": "NACA RM L9J07\n17\n\nCENTERS OF PRESSURE AT ZERO YAW\n\nThe local center-of-pressure variation with angle of attack (figs. 49 to 51) depended primarily upon the change in size, strength, and location of the vortex. At $\\alpha = 4.1^\\circ$ the center of pressure was generally in the vicinity of the quarter chord from station 3 outboard except for a slight rearward displacement from the quarter-chord line at the outboard spanwise location where the chordwise negative pressure dip behind the vortex weakened. Each center-of-pressure curve for higher angles of attack had a rearward displacement with reference to the quarter-chord line, with the most rearward point generally moving inboard as the angle of attack was increased. This maximum rearward displacement of local center of pressure generally occurred at the spanwise location where the vortex was on the rear of the section chords. Farther inboard for each angle of attack of each wing above $4.1^\\circ$, the local center of pressure was closer to or even ahead of the quarter-chord line where the negative-chordwise-pressure dips behind the vortex were located on the rear of the section chords and the negative pressure peaks in the vortex region were on the forward part of the chords. The distance from the plane of symmetry of the described regions of rearward and forward center-of-pressure displacement from the quarter-chord line varied approximately inversely with the wing aspect ratio. Although its variation with angle of attack was erratic, the center of pressure at station 1 was always at a greater percent of the local chord behind the leading edge, generally between 0.35c and 0.40c, than the center of pressure of station 2.\n\nAt an angle of attack of approximately $4^\\circ$, as shown in figure 52, the lateral center of pressure of the three wings was about 42 percent of the semispan, which is only about 1 percent higher than that predicted by the Weissinger theory in reference 13. With increased angle of attack there was a gradual inboard movement of the lateral center of pressure for each wing as the outboard sections progressively became less effective. The distance of the spanwise center of pressure from the plane of symmetry varied among the wings basically as an inverse function of aspect ratio, although a greater successive change in position at a given angle of attack was noted between wings 1 and 2 than between wings 2 and 3 because of the more rapid loss of outboard effectiveness for wing 1 as noted in the section entitled \"Section Lift Characteristics at Zero Yaw.\"", "timestamp": "2026-07-22T05:50:27.876835+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 21, "total_pages": 98, "image_filename": "19930086073_p21.jpg", "text": "NACA RM A9H04\n19\n\n[Figure: A large triangular wing mounted on a support structure inside a wind tunnel. A person in a white suit is standing near the base of the structure.]\n\n(a) Wing alone.\n\nFigure 3.- Triangular plan-form wing as mounted for investigation in the Ames 40- by 80-foot wind tunnel.\n\nNACA\nA-11220", "timestamp": "2026-07-22T05:50:29.645425+00:00"}
{"citation_id": "19930085972", "source_url": "https://ntrs.nasa.gov/api/citations/19930085972/downloads/19930085972.pdf", "page_number": 31, "total_pages": 46, "image_filename": "19930085972_p31.jpg", "text": "NACA RM L9B18\n29\n\nPitching-moment coefficient, $C_m$\n4\n.2\n0\n-.2\n-.4\n-.6\n\n$i_t$ (deg)\n-0.83 No cutout\n3 tail off\n0 Faired cutout\n3 tail off\n\nLongitudinal-force coefficient, $C_X$\n.3\n.2\n.1\n0\n\nAngle of attack, $\\alpha$, deg\n16\n8\n0\n-8\n\nLift coefficient, $C_L$\n-4 0 4 8 12\n\nNACA\n\nFigure 8.- Aerodynamic characteristics of a variable-sweep model with and without faired wing cutout. $\\Lambda = 30^\\circ$.", "timestamp": "2026-07-22T05:50:32.142996+00:00"}
{"citation_id": "19930093769", "source_url": "https://ntrs.nasa.gov/api/citations/19930093769/downloads/19930093769.pdf", "page_number": 10, "total_pages": 39, "image_filename": "19930093769_p10.jpg", "text": "NACA RM No. E5L10a CONFIDENTIAL 9\n\nignition limits and the areas in figure 9(b) that separate the regions of successful and unsuccessful attempts are drawn primarily for ease of interpreting the data. At 30,000 feet, successful starts could be made only at a flight Mach number of 0.40, whereas at altitudes from sea level to 10,000 feet starts were successful at flight Mach numbers from 0 to 0.50. The results of the starting characteristics of gasoline, which were investigated with the standard spark plug, are presented in figure 9(c) in a similar manner to that used in figure 9(b) for AN-F-58 fuel. The dashed lines for starting limits of the AN-F-58 fuel from figure 9(b) are also included for comparison. For flight Mach numbers below 0.50, the altitude starting limit for gasoline was approximately 10,000 feet higher than for AN-F-58 fuel; satisfactory starts were also obtained over a wider range of flight Mach numbers at low altitudes for gasoline than for AN-F-58 fuel. Even with the extended-electrode spark plug, the starting characteristics of AN-F-58 fuel were slightly poorer than those of gasoline with the standard spark plug and further development of the spark-plug design is therefore required to obtain ignition of AN-F-58 fuel comparable to that of gasoline in this engine.\n\nCold-Weather, Sea-Level Starting Characteristics\n\nwith AN-F-58 Fuel\n\nCold-weather starting experiments were simulated at zero-ram, sea-level conditions with AN-F-58 fuel 48-210 in engine A at inlet-air temperatures of -30° and -50° F. Comparative experiments with gasoline were not conducted. For these starting experiments, the engine was run from zero speed with the standard cranking motor after it had been previously windmilled for a sufficient period of time to bring the temperatures throughout the engine into equilibrium. Both the fuel and the lubricating oil were supplied to the engine from outside tanks at a temperature of about 70° F and therefore only the fuel and the oil contained in the engine parts were at the reduced temperature of the experiment. At the conditions of -30° and -50° F, the engine was successfully started and accelerated on the first attempt. At -50° F, however, a delay of nearly 30 seconds occurred before the fuel ignited and the resulting start was very hot.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:50:33.094423+00:00"}
{"citation_id": "19930082472", "source_url": "https://ntrs.nasa.gov/api/citations/19930082472/downloads/19930082472.pdf", "page_number": 11, "total_pages": 34, "image_filename": "19930082472_p11.jpg", "text": "NACA TN No. 1797\n\nThese pressure distributions are typical of those obtained at the higher angles of attack over all except the tip sections of the wing. It will be observed that above $16.6^\\circ$ the suction pressures over the leading edge began to decrease. This caused a loss of lift over the section when the angle of attack was further increased. This loss of lift occurred first over the inboard sections and, as angle of attack was further increased, occurred over sections farther outboard. For example (fig. 7), loss of lift occurred at the 20.9-percent semispan section at about $16.6^\\circ$ angle of attack, had progressed out to 41.7-percent semispan at $20^\\circ$ angle of attack, and did not occur at 80-percent semispan until about $30^\\circ$ angle of attack. The loss of lift over the inboard sections caused a change in the spanwise loading in which the spanwise center of load was shifted outward. The outward movement of the center of load, due to the forward sweep of the wing, caused positive pitching moments. As a result the aerodynamic center moved forward to 5 percent of the mean aerodynamic chord forward of the leading edge.\n\nThe loss of the leading-edge suction peak evidently is the result of a permanent separation of the laminar boundary layer at the airfoil nose with no subsequent reattachment. It is apparently largely independent of the turbulent separation that occurs over the rear portion of the airfoil. This is evidenced by its sudden appearance and its rapid spread beyond the area affected by turbulent separation. The boundary-layer measurements were not sufficiently detailed to completely verify the foregoing inferences. The evidence strongly indicates, however, that the greatly increased drag, the decreased lift-curve slope, and the forward shift of aerodynamic center were caused primarily by a leading-edge type of separation.\n\nThe section lift characteristics (fig. 7) show the influence of the spanwise boundary-layer drain over the swept-forward wing. The maximum lift coefficients of sections perpendicular to the quarter-chord line varied from 1.01 at 28.1-percent semispan to 1.5 at 71.4-percent semispan. Two-dimensional data for the airfoil section used (NACA 64A112 perpendicular to the quarter-chord line) show that a maximum lift coefficient of 1.5 is attainable in a comparable range of Reynolds number. In comparing two-dimensional and three-dimensional values, however, account must be taken of the effects of wing sweep. If the lift coefficients shown in figure 7 had been based on the velocity component perpendicular to the leading edge in accordance with the concepts of simple sweep theory (reference 3), section maximum lift would vary from 2.02 at 28.1-percent semispan to 3.0 at 71.4-percent semispan. On this basis, the sections attained considerably higher maximum lift coefficients than are attainable in two-dimensional flow. From this it can be concluded that insofar as", "timestamp": "2026-07-22T05:50:34.803070+00:00"}
{"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 63, "total_pages": 78, "image_filename": "19930082618_p63.jpg", "text": "```markdown\nNACA TN 1945\n\nSection lift coefficient, $c_l$\n2.0\n1.6\n1.2\n.8\n.4\n0\n-.4\n-.8\n-1.2\n-1.6\n-2.0\n\nMoment coefficient, $C_{m_{c/4}}$\n-.1\n-.2\n-.3\n-.4\n\nSection angle of attack, $\\alpha$, deg\n-16 -8 0 8 16 24\n\nR\n$\\circ$ 0.7 x $10^6$\n$\\square$ 1.0\n$\\diamond$ 1.5\n$\\nabla$ 2.0\n$\\triangle$ 6.0\n\nFlagged symbols denote\nstandard roughness\n\nNACA\n\n(b) Section lift and pitching-moment characteristics of the NACA 23012 airfoil section with a\n0.20c simulated split flap deflected 60°.\n\nFigure 14.— Continued.\n\n61\n```", "timestamp": "2026-07-22T05:50:36.903114+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 7, "total_pages": 78, "image_filename": "19930082483_p7.jpg", "text": "NACA TN No. 1807\n5\n\nThe subscript 0 refers to corrected conditions at NACA sea level\nin each case.\n\nTherefore\n\n$$V_0 = \\frac{V}{\\sqrt{\\theta}} \\tag{6}$$\n\nwhere $\\theta$ is the temperature correction ratio, $T/T_0$.\n\n$$\\rho_0 = \\rho \\frac{\\theta}{\\delta} \\tag{7}$$\n\nwhere\n$\\rho$ mass density, (slugs/cu ft)\n$\\delta$ pressure correction ratio, $p/p_0$\n$p$ absolute pressure, (lb/sq ft) or (in. Hg)\n\n$$W_0 = \\frac{W\\sqrt{\\theta}}{\\delta} \\tag{8}$$\n\n$$\\Delta_s h_0' = \\frac{\\Delta_s h'}{\\theta} \\tag{9}$$\n\n$$(\\text{power})_0 = \\frac{(\\text{power})}{\\delta\\sqrt{\\theta}} \\tag{10}$$\n\nCorrection of any performance or loss expression can be made\nby correcting the terms within that expression which are variable\nwith changes in altitude, as indicated in equations (5) to (10).\n\nIf the pressure relations throughout the turbine are main-\ntained constant, the correction of performance or loss expressions\nto equivalent sea-level values may be made by using correction\nfactors based on the gas properties at some convenient reference\npoint in the turbine. The reference point selected is the turbine\ninlet. That the performance or loss expression may contain terms\ninvolving gas properties at locations other than at the reference\npoint does not alter the validity of the procedure for it can be\nshown that changes of temperature, pressure, and Mach number at the\nreference point are generally reflected by similar changes through-\nout the turbine in most practical applications.", "timestamp": "2026-07-22T05:50:37.124702+00:00"}
{"citation_id": "19930086097", "source_url": "https://ntrs.nasa.gov/api/citations/19930086097/downloads/19930086097.pdf", "page_number": 16, "total_pages": 36, "image_filename": "19930086097_p16.jpg", "text": "14 CONFIDENTIAL NACA RM A9H11\n\n$$\n\\alpha^2 = \\frac{c_{d_{\\min}}}{2C_1 \\left(1 + \\frac{3C_2h}{2C_1c}\\right)}\n$$\n\nand the maximum lift-drag ratio for small values of $ h/c $ is, accordingly,\n\n$$\n\\left(\\frac{L}{D}\\right)_{\\max} = \\left(\\frac{C_1}{2c_{d_{\\min}}}\\right)^{1/2} \\left(1 + \\frac{1}{4} \\frac{C_2h}{C_1c}\\right)\n$$\n\nIf the changes in minimum drag are small, this leads to the following approximate result: The percentage improvement in maximum lift-drag ratio is equal to one-half the percentage improvement in minimum drag plus one-fourth the percentage improvement in lift-curve slope. Consequently, it is possible for a blunt-trailing-edge airfoil to have a higher minimum drag coefficient, yet still have a higher maximum lift-drag ratio than a corresponding sharp-trailing-edge airfoil. For such a case to occur, it is necessary that the percentage increase in lift-curve slope exceed twice the percentage increase in minimum drag.\n\n### Parameters Affecting the Theoretical Characteristics of Blunt-Trailing-Edge Airfoils\n\nThe preceding theoretical calculations apply strictly only for two-dimensional flow. On the basis of existing knowledge it would be expected that the calculations of lift-curve slope would represent actual conditions reasonably well as long as three-dimensional effects, such as tip effects, are not large. In general, variations in airfoil-thickness ratio, type of boundary-layer flow, or shape of the airfoil contour forward of the base should not have an appreciable effect on the lift characteristics in two-dimensional flow. Such variations, however, may have a pronounced effect on the drag. The calculations made earlier, which illustrated lower drag for blunt-trailing-edge sections, were concerned only with specific flow conditions; namely, airfoil contours of straight sides, thickness ratio of 10 percent, and laminar flow in the boundary layer. Since the analysis has shown that sizable drag reductions may result under these specific conditions, the question immediately arises as to what may be expected when other conditions exist.\n\nOne parameter that is expected to have a significant effect on the drag of blunt-trailing-edge airfoils is the condition of the boundary layer just forward of the base. A change from laminar to turbulent boundary-layer flow is known to have a large effect on the base drag of bodies of revolution. In fact, negative base drag coefficients have actually been measured (reference 8) on certain highly boattailed bodies having a turbulent boundary layer approaching the base. This phenomenon\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:50:38.619826+00:00"}
{"citation_id": "19930085938", "source_url": "https://ntrs.nasa.gov/api/citations/19930085938/downloads/19930085938.pdf", "page_number": 29, "total_pages": 42, "image_filename": "19930085938_p29.jpg", "text": "28\n\n11 1/4\n18\n25 1/2\n\n7\n9\n-2\n0\n1\n3\n5\n\nFP\n-2 0 1 3 5 7 9 11 1/4\n18\n25 1/2\n\nNACA\n\nFigure 7.- Hull lines of model 237-7TB.\n\nNACA RM No. 19B04", "timestamp": "2026-07-22T05:50:39.035618+00:00"}
{"citation_id": "19930091987", "source_url": "https://ntrs.nasa.gov/api/citations/19930091987/downloads/19930091987.pdf", "page_number": 9, "total_pages": 12, "image_filename": "19930091987_p9.jpg", "text": "CHARACTERISTICS OF LOW-ASPECT-RATIO WINGS AT SUPERCRITICAL MACH NUMBERS 5\n\n[Figure: Six plots labeled (a) through (f), showing moment coefficient $C_{m_{c/4}}$ versus lift coefficient $C_L$. Vertical axis labeled “Moment coefficient, $C_{m_{c/4}}$” with scale from 0 to .1. Horizontal axis labeled $C_L$ with scale from 0 to .6. Each plot contains multiple curves corresponding to different aspect ratios A = 8, 5, 3, 2 as indicated by legend on right side of plot (c). Below plots, legend indicates Mach numbers: (a) M=0.5, (b) M=0.7, (c) M=0.8, (d) M=0.85, (e) M=0.875, (f) M=0.9.]\n\nFIGURE 5.—Moment-lift curves for various aspect ratios and Mach numbers. NACA 0012 section; rectangular plan form and tip.\n\nDISCUSSION\n\nFigures 3 and 4 show a most pronounced change in finite wing characteristics as the critical Mach number of the sections, approximately 0.72 for the NACA 0012 sections used herein, is exceeded. For conventional and higher aspect ratios, the lift curves (fig. 3) show irregularities in slope and effectively discontinuous slope changes at supercritical speeds. These irregularities, which are the principal cause of the stability difficulties that have been encountered at supercritical speed, appear first at the higher lift coefficients encountered in the pull-out condition, but as the speed is increased they occur at progressively lower lift coefficients until finally irregularities occur in the low-lift region around zero lift coefficient. For the higher-aspect-ratio wings tested in the present investigation, the lift-curve slope decreases almost to zero in the low angle-of-attack range at Mach numbers between 0.85 and 0.875.\n\nThe low-aspect-ratio wings (aspect ratios 2 and 3), however, show none of the characteristic lift-curve irregularities at the high Mach numbers. The lift-curve slopes for the low-aspect-ratio wings also show relatively little change with Mach number. The usual rise of lift-curve slope with Mach number through the subcritical speed range is absent as is the abrupt fall in slope at supercritical speeds. A partial explanation for the absence of the increase of lift-curve slope with Mach number in the subcritical range is given by considering the finite-wing characteristics to be composed of the infinite-wing or section characteristics and the induced characteristics. The induced characteristics are determined principally by the lift coefficient and are, in first-order approximation, independent of the Mach number. Hence, when the induced characteristics are large, as for the low-aspect-ratio wings, a given change in section characteristics produces less relative change of lift-curve slope than is usually expected or obtained for wings of high or conventional aspect ratios for which the induced characteristics are relatively small.\n\nThe effects of aspect ratio on the drag characteristics as shown by the polar diagrams (fig. 4) indicate very marked departure from the usual low-speed characteristics when the speeds are increased to supercritical values. The results presented in figure 4, as previously noted, include the induced drag. At the lower speeds the low-aspect-ratio wings have the highest drag, as could be determined by theory. As the critical speed of the basic section of the wings is exceeded, however, the differences in drag diminish and the polar curves approach coincidence (fig. 4 (d)). With still further increase of speed, the order of the variation of drag with aspect ratio reverses; the low-aspect-ratio wings, even including the induced drag as in figure 4, have markedly reduced drag as compared with the high-aspect-ratio wings. This change in characteristics is associated with delayed and less rapid rise of drag as the aspect ratio is decreased. Both the delayed drag rise and the slower rate of drag rise are illustrated for the minimum drag attitude ($0^\\circ$ angle of attack and zero lift for the symmetrical section) in figure 6. The section critical Mach number is given in the figure for comparison. For the wings of aspect ratios 2 and 3, the Mach number for significant drag rise is approximately 0.1 higher than for the infinite- or high-aspect-ratio wings and the initial rate of drag rise is much less.\n\n[Figure: Plot of minimum drag coefficient $C_{D_0}$ versus Mach number $M$. Vertical axis labeled $C_{D_0}$ with scale from 0 to .06. Horizontal axis labeled “Mach number, M” with scale from .4 to 1.0. Multiple curves correspond to aspect ratios A = 8, 5, 3, 2 as indicated by legend on left. Dashed vertical line labeled “Section critical Mach number” near M = 0.72.]\n\nFIGURE 6.—Variation with Mach number of the minimum drag coefficient for wings of various aspect ratios.", "timestamp": "2026-07-22T05:50:40.556465+00:00"}
{"citation_id": "19930091993", "source_url": "https://ntrs.nasa.gov/api/citations/19930091993/downloads/19930091993.pdf", "page_number": 8, "total_pages": 21, "image_filename": "19930091993_p8.jpg", "text": "```markdown\n4\nREPORT 928—NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nThis computation with measured values for $f$ and $p_{T,3}$ completely established the state of the gas at station 3. The isentropic enthalpy drop through the turbine was then computed from $H_{T,2}$, $f$, $p_{T,3}$, $p_{T,4}$, and the charts of reference 6.\n\nIn reference 4, the range of turbine operation obtainable from engine experiments was predicted to be insufficient for accurate determination of efficiency contours and location of region of highest efficiency. The prediction was verified by the results on the engine. The range of turbine operation was then extended by resetting the stator blades of the compressor for a lower air flow and the experiments were then repeated. Only one resetting of the compressor stator blades was necessary to cover a sufficient range of turbine operation.\n\noperation in a region that is normally the surge region although no fluctuation in compressor-outlet-air pressure was noticed during operation of the engine.\n\n<!-- Image (514, 140, 864, 425) -->\nFIGURE 5.—Selected compressor data.\n\nPERFORMANCE OF COMPONENTS\n\nFrom the engine experiments, the performance of the compressor, the combustion chamber, and the turbine components were separately evaluated. Curves faired through these data were then used to evaluate the engine performance and to examine the interaction of the engine components.\n\nCOMPRESSOR PERFORMANCE\n\nIf a small Reynolds number effect is assumed, the complete compressor performance is determined by any two independent parameters, provided all parameters are dimensionless or are given in equivalent values. If the independent variables chosen are the equivalent speed $n/\\sqrt{\\theta_{T,1}}$ and the equivalent air weight flow $W_1/(\\sigma_{T,1}\\sqrt{\\theta_{T,1}})$, these variables determine the dependent parameters $\\Delta H_{i,c}/\\theta_{T,1}$ and $\\Delta H_i/\\theta_{T,1}$,\nwhere\n$\\Delta H_{i,c}$ isentropic rise in stagnation enthalpy, (ft-pound/lb)\n$\\Delta H_i$ stagnation enthalpy rise of air in compressor, $H_{T,2}-H_{T,1}$, (ft-pound/lb)\n\nThe compressor efficiency is the ratio of these two parameters. The pressure ratio $p_{T,2}/p_{T,1}$ is a function only of the equivalent isentropic enthalpy rise $\\Delta H_{i,c}/\\theta_{T,1}$ and the temperature ratio is a function only of the equivalent enthalpy rise $\\Delta H_i/\\theta_{T,1}$. In order to obtain good fairing of the compressor data, functions of the four fundamental parameters just described were used instead of the original parameters. These parameters were constructed to reduce the dependence of the resultant functions on compressor speed. With the parameters $\\Delta H_{i,c}/n^2$ and $\\Delta H_i/n^2$ as ordinate and abscissa, curves were faired through the data for constant speeds and not only the data for one particular speed but also the data for other speeds were considered. A systematic effect of inlet-air temperature was noticeable. A plot of data selected for mean compressor-air temperature equal to the room temperature showed much better correlation, which indicates the effect of heat lost to the room air. The resultant plot is shown in figure 5. The straight lines are contours of constant efficiency. The discontinuity, which may be seen in all curves, indicates the incidence of compressor surge. The portion of the curves in the region of low efficiency represents\n\nAn additional chart is required to express the relation between the compressor air flow and the other compressor performance variables. This chart is a plot of pressure ratio $p_{T,2}/p_{T,1}$ against the air-flow parameter $W_1/(\\sigma_{T,1}\\theta_{T,1})$ for constant values of the equivalent compressor speed $n/\\sqrt{\\theta_{T,1}}$, as shown in figure 6. Only the data used in figure 5 (mean compressor air temperature equal to room temperature) are plotted on this chart. When any uncertainty existed as to the fairing of the curves, the data for several of the speeds were plotted on a chart that showed little speed effect ($\\Delta H_{i,c}/n^2$ and $W_1/(\\sigma_{T,1}\\theta_{T,1})$) as coordinates with contours for constant speeds $n/\\sqrt{\\theta_{T,1}}$. Thus, the data allowed several speeds to be considered in fairing a curve for any one speed. The region beyond the surge line for each curve in figure 6 is shown connected with the normal operating portion by a dashed line indicating the absence of any data for that section of the curve.\n\n<!-- Image (514, 666, 864, 892) -->\nFIGURE 6.—Selected compressor-pressure-ratio and air-flow data.\n```", "timestamp": "2026-07-22T05:50:41.321284+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 81, "total_pages": 99, "image_filename": "19930082511_p81.jpg", "text": "NACA TN No. 1826\n79\n\n[Figure: Diagram (a) showing a rectangular channel with hatched boundaries and a central lifting element symbol.]\n\n(a) Lifting element on the center line.\n\n[Figure: Diagram (b) showing a similar rectangular channel with a lifting element symbol, a voltmeter symbol labeled 'V' connected to the top boundary, and a point labeled 'P' on the left boundary. The NACA logo is present below the diagram.]\n\n(b) Unsymmetrical location of the lifting element.\n\nFigure 9.- Acceleration-potential analogies for the two-dimensional closed-open tunnel.", "timestamp": "2026-07-22T05:50:43.660553+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 22, "total_pages": 98, "image_filename": "19930086073_p22.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:50:52.269550+00:00"}
{"citation_id": "19930085977", "source_url": "https://ntrs.nasa.gov/api/citations/19930085977/downloads/19930085977.pdf", "page_number": 25, "total_pages": 33, "image_filename": "19930085977_p25.jpg", "text": "24\nNACA RM L9H22\n\nCONFIDENTIAL\n\nBending-moment coefficient, $C_B$\nLift coefficient, $C_L$\n\n| M | |\n| :--- | :--- |\n| 1.15 | $\\triangleleft$ |\n| 1.10 | $\\triangleright$ |\n| 1.08 | $\\triangle$ |\n| 1.05 | $\\triangleright$ |\n| 1.03 | $\\diamond$ |\n| 1.00 | $\\square$ |\n| .98 | $\\triangle$ |\n| .95 | $\\triangleright$ |\n| .93 | $\\square$ |\n| .90 | $\\triangleright$ |\n| .88 | $\\triangleright$ |\n| .85 | $\\triangledown$ |\n| .80 | $\\diamond$ |\n| .70 | $\\square$ |\n| .60 | $\\circ$ |\n\n[Graph plotting Bending-moment coefficient, $C_B$ against Lift coefficient, $C_L$ for various M values]\n\nNACA\n\nCONFIDENTIAL\nFigure 9.— Concluded.", "timestamp": "2026-07-22T05:50:52.805507+00:00"}
{"citation_id": "19930082474", "source_url": "https://ntrs.nasa.gov/api/citations/19930082474/downloads/19930082474.pdf", "page_number": 11, "total_pages": 21, "image_filename": "19930082474_p11.jpg", "text": "NACA TN No. 1799\n\nviolent, random yawing motions would then occur with some roll, the rate of descent would apparently increase rapidly, the rotational speed of the rotor would vary noticeably, and more often than not the machine would eventually pitch nose down and recover by gaining speed, despite application of considerable rearward control. This behavior had many variations which apparently depended on small horizontal velocities and on power conditions. In some cases similar shaking of the machine was encountered at indicated rates of descent of only 300 feet per minute. The loss of control appeared most severe when the power was as high as possible at the required rate of descent. As power was progressively reduced during successive trials the difficulties were reduced to the point at which no trouble was encountered for the power settings permitting steady descents of about 1500 feet per minute and higher. These descents were always performed with a margin of altitude and no difficulty was ever encountered in recovering at any stage desired.\n\nThe yawing motions and inadvertent recovery mentioned previously are possibly affected by rearward velocity. Nevertheless, the fundamental cause of the phenomenon appears to be an irregular flow of air through the rotor. In hovering, a definite downward flow of air through the rotor occurs, and in descent with the power completely off an upward flow of air through the rotor takes place; but in this intermediate condition the air tends to move with the rotor. A logical assumption is that when the air attempts to stay with the rotor, it might actually mix in turbulent and erratic fashion with the air outside the rotor disk. Motion-picture studies of tufted blades during some of these cases have thus far shown no stalling but have shown pronounced, but irregular, blade bending. The presence of this irregular bending tends to support the irregular-flow explanation, but much remains to be learned about this region of operation.\n\nAnother phenomenon has been encountered following take-off. The machine was being accelerated rapidly horizontally from hovering and, at 20 to 30 miles per hour, it pitched up abruptly. In several cases it was necessary to have the control against the forward stop for a short interval of time to check the motion. This same tendency has been noticed in other helicopters. The horizontal acceleration is normally low enough that full control deflection is not required. This characteristic may be due to the dynamic stability characteristics in pitch and to the rapid entry into the higher speed range. This condition should be investigated, however, as a possible critical one in determining the required control range.\n\nThe preceding sections have pointed out some of the stability and control characteristics found for a particular helicopter type. They appear to be applicable to other types, however, in whole or in part.", "timestamp": "2026-07-22T05:50:53.024281+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 55, "total_pages": 96, "image_filename": "19930085880_p55.jpg", "text": "NACA RM No. L9C03\n53\n\n| Speed (fps) | |\n| :--- | :--- |\n| 10 | $\\bigcirc$ |\n| 15 | $\\square$ |\n| 20 | $\\diamond$ |\n| 25 | $\\triangle$ |\n| 30 | $\\nabla$ |\n\nDraft, ft\n.64\n.56\n.48\n.40\n.32\n.24\n.16\n.08\n0\n\nWetted area, sq ft\n0\n.05\n.10\n.15\n.20\n.25\n.30\n.35\n\n(d) $\\tau = 16^\\circ$.\n\n[NACA logo]\n\nFigure 17.- Continued.", "timestamp": "2026-07-22T05:50:54.540179+00:00"}
{"citation_id": "19930093789", "source_url": "https://ntrs.nasa.gov/api/citations/19930093789/downloads/19930093789.pdf", "page_number": 10, "total_pages": 29, "image_filename": "19930093789_p10.jpg", "text": "NACA RM No. E5I21 CONFIDENTIAL 9\n\nthree configurations, the curves of constant equivalent blade speed $U \\left( \\frac{a_0}{a_1} \\right)$ are vertical above a pressure ratio $p_1'/p_3'$ of approximately 2.25 because the flow through the stator blades was choking and the equivalent weight flow therefore remained constant. The over-all performance characteristics are similar for the three configurations and each configuration has an efficiency near the maximum over the range of pressure ratio from 1.25 to 3.70.\n\nBrake efficiency $\\eta$ plotted against the ratio of blade-to-jet speed ratio $U/V_j$ (fig. 14) is nearly independent of total-pressure ratio $p_1'/p_3'$ for configuration 2. Because this independence of pressure ratio also exists for configurations 1 and 3, the comparison of the three configurations in figure 15 at a pressure ratio of 3.00 is representative of the entire range of pressure ratios. In addition to showing the same changes in brake efficiency as figures 11 to 13, figures 14 and 15 indicate that the maximum brake efficiency was obtained with a blade-to-jet speed ratio $U/V_j$ between 0.50 and 0.55. The maximum brake efficiency was 0.795 for configuration 1 and 0.833 for configuration 2, a difference of approximately 0.04. With the 70°-cone-angle stator, separation probably occurred at the tips of the rotor blades; the performance was improved by changing the cone angle to 0° and making the stator-blade height equal to the rotor-blade height, thereby eliminating the separation (figs. 8(a) and 8(b)). If the efficiency were computed using the total-to-static pressure ratio $p_1'/p_3$ instead of the total-pressure ratio $p_1'/p_3'$, the relative advantage of configuration 2 over configuration 1 would be substantially unchanged.\n\nFor configuration 3 (fig. 8(c)), the maximum brake efficiency was 0.835, an increase of less than 0.005 over the maximum efficiency of configuration 2. Although any tip leakage introduced by replacing the labyrinth no-leakage shroud with the cylindrical shroud would cause a reduction in the working fluid passing through the rotor blades, the measured efficiency slightly increased. Within the probable reproducibility of the data, the performance of configuration 2 is identical to the performance of configuration 3.\n\nIn order to determine the approximate magnitude of the leakage air flow between the blade caps and the cylindrical stationary shroud, the weight flow through a thin-plate orifice of equal area was computed and multiplied by a factor of 0.7 (reference 4). The conditions for maximum leakage flow occurred at the highest blade speed and pressure ratio investigated; at these conditions the computed leakage was about 1.7 percent of the air flow entering the turbine. Because the\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:50:54.790193+00:00"}
{"citation_id": "19930085972", "source_url": "https://ntrs.nasa.gov/api/citations/19930085972/downloads/19930085972.pdf", "page_number": 32, "total_pages": 46, "image_filename": "19930085972_p32.jpg", "text": "30\nNACA RM L9B18\n\n<!-- Image (193, 116, 712, 874) -->\n\n(a) Basic tail position.\nFigure 9.- Aerodynamic characteristics of a variable sweep model with two horizontal-tail positions. $\\Lambda = 45^\\circ$.", "timestamp": "2026-07-22T05:51:00.070067+00:00"}
{"citation_id": "19930085962", "source_url": "https://ntrs.nasa.gov/api/citations/19930085962/downloads/19930085962.pdf", "page_number": 38, "total_pages": 51, "image_filename": "19930085962_p38.jpg", "text": "NACA RM A9E05 CONFIDENTIAL 37\n\nLift coefficient, $C_L$\n\n| $a_{approx.}$ |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  |  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{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 22, "total_pages": 114, "image_filename": "19930086061_p22.jpg", "text": "18\nNACA RM L9J07\n\nWING FORCE AND MOMENT CHARACTERISTICS\n\nLift and Pitching-Moment Characteristics at Zero Yaw\n\nThe lift-curve slopes of the three wings were practically linear but increased slightly with aspect ratio below the angle-of-attack range of $12^\\circ$ to $14^\\circ$ (fig. 53). At higher angles of attack, however, the slopes in contrast decreased with aspect ratio. The values of $C_{L_{max}}$, $\\alpha_{C_{L_{max}}}$, and $C_{L_\\alpha}$ (measured at $C_L = 0.2$) for the three wings and the theoretical values of $C_{L_\\alpha}$ (obtained from reference 13 at $C_L = 0$ using the Weissinger theory) are presented in the following table:\n\n| Wing | A | $C_{L_{max}}$ | $\\alpha_{C_{L_{max}}}$ (deg) | Measured $C_{L_\\alpha}$ at $C_L = 0.2$ (per deg) | Theoretical $C_{L_\\alpha}$ at $C_L = 0$ (per deg) |\n| :--- | :--- | :--- | :--- | :--- | :--- |\n| 1 | 3.46 | 0.98 | 34.1 | 0.043 | 0.046 |\n| 2 | 2.31 | 1.16 | 36.1 | .041 | .042 |\n| 3 | 1.73 | 1.17 | 38.5 | .037 | .037 |\n\nThe experimental values of $C_{L_\\alpha}$ were measured at $C_L = 0.2$ due to insufficient data at zero lift for all three wings. The experimental values agree well with the theoretical values and tend to increase with increased wing aspect ratio. A comparison is made in figure 54 of the lift of wing 2 at Reynolds numbers of $0.85 \\times 10^6$ and $1.42 \\times 10^6$ with that of the large-scale wing of reference 3 (identical in plan form and section to wing 2) at a Reynolds number of $2.91 \\times 10^6$. The wing of reference 3, which had negligible scale effect from Reynolds numbers of $2.91 \\times 10^6$ to $9.61 \\times 10^6$, generally had a slightly higher lift-curve slope than wing 2 and a more gradual stall at a lower $C_{L_{max}}$ (1.03) and $\\alpha_{C_{L_{max}}}$ ($33^\\circ$). Consistent with these trends, increasing the Reynolds number of wing 2 increased the lift-curve slope except at low angles of attack and produced a more gradual stall. The inclusion of the 1.1-percent increment in $C_L$ due to chord force mentioned in the section entitled \"Corrections to Data\" for wing 2 at $\\alpha = 24.1^\\circ$ and $R = 0.85 \\times 10^6$ would give better agreement than noted in figure 54.", "timestamp": "2026-07-22T05:51:02.790403+00:00"}
{"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 64, "total_pages": 78, "image_filename": "19930082618_p64.jpg", "text": "R\n○ 0.7 x 10⁶\n□ 1.0\n△ 1.5\n▲ 2.0\nFlagged symbols denote\nstandard roughness\n\nSection drag coefficient, c_d\n.024\n.020\n.016\n.012\n.008\n.004\n\nSection lift coefficient, c_l\n-1.2  -.8  -.4   0   .4   .8  1.2  1.6\n\nR\n▽ 3.0 x 10⁶\n◁ 6.0\n◁ 9.0\nFlagged symbols denote\nstandard roughness\n\nSection drag coefficient, c_d\n.024\n.020\n.016\n.012\n.008\n.004\n\nSection lift coefficient, c_l\n-1.2  -.8  -.4   0   .4   .8  1.2  1.6\n\nMoment coefficient, c_m\n.0\n-.04\n-.08\n-.12\n\nSection lift coefficient, c_l\n-1.2  -.8  -.4   0   .4   .8  1.2  1.6\n\nR          a.c. position\n           x/c     y/c\n○ 0.7 x 10⁶  .231    .152\n□ 1.0        .230    .139\n△ 1.5        .233    .131\n▽ 3.0        .232    .098\n◁ 6.0        .241    .064\n◁ 9.0        .241    .035\n◁ 9.0        .247    .035\n\nNACA\n\n(c) Section drag characteristics and section pitching-moment characteristics about the aerodynamic center of the plain NACA 23012 airfoil section.\n\nFigure 14.— Concluded.\n\n62\n\nNACA TN 1945", "timestamp": "2026-07-22T05:51:11.293117+00:00"}
{"citation_id": "19930093769", "source_url": "https://ntrs.nasa.gov/api/citations/19930093769/downloads/19930093769.pdf", "page_number": 11, "total_pages": 39, "image_filename": "19930093769_p11.jpg", "text": "10 CONFIDENTIAL NACA RM No. E3L10a\n\nComparison of Turbine-Outlet-Gas-Temperature Distribution\nwith AN-F-58 Fuel and Gasoline\n\nThe radial temperature distribution at the turbine outlet with\nboth AN-F-58 fuel and gasoline are shown in figures 10(a) and\n10(b) for a flight Mach number of 0.85 and altitudes of 20,000 and\n50,000 feet, respectively. These data were obtained at an engine\nspeed of approximately 12,025 rpm at an altitude of 20,000 feet\nand at approximately 12,050 rpm at an altitude of 50,000 feet.\nCurves are shown for maximum, average, and minimum temperatures\nwith the maximum and minimum temperatures being the highest and\nlowest temperatures, respectively, at each radial thermocouple\nlocation and are not necessarily at the same circumferential posi-\ntion. For both altitudes, the average turbine-outlet temperatures\nfor the two fuels are about the same, as previously indicated in\nfigure 5. At an altitude of 20,000 feet, there is very little dif-\nference in the temperature distribution for the two fuels. At\n50,000 feet, the highest maximum temperature, which occurred near\nthe blade roots, was about the same for both fuels (fig. 10(b)).\nThe difference in average temperature between the blade tip and\nthe root, however, was about $240^\\circ$ F greater for AN-F-58 fuel than\nfor gasoline and the spread between the maximum and minimum tem-\nperatures is about $300^\\circ$ F less at the blade roots and about $150^\\circ$ F\nless at the blade tips for the AN-F-58 fuel than for gasoline.\n\nCarbon Deposition\n\nThe completion of the program with AN-F-58 fuel in engine A\nrequired about 30 hours and 11 minutes of engine operation after\nwhich time the engine was disassembled for inspection. Hard car-\nbon deposits were built up on the inside of both the inner and\nouter annuli of the combustor basket near the fuel nozzles to a\nheight of about 1/2 inch. A photograph of these carbon deposits\nis shown in figure 11. The total weight of the carbon deposits\nwas 240 grams. No serious warpage or other deterioration of the\ncombustor basket was observed. Because engine B was operated on\nboth AN-F-58 fuel and gasoline without disassembly for inspection,\nno evaluation of carbon deposits formed by gasoline was made; car-\nbon deposits formed by gasoline during other investigations, how-\never, have been very small or completely absent.\n\nThe exhaust jet from the engine when using AN-F-58 fuel,\nwhich was observed at sea-level conditions, was characterized by\na faint haze very similar to that which usually occurs with gaso-\nline. Much heavier carbon deposits were, however, built up on\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:51:11.918337+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 82, "total_pages": 99, "image_filename": "19930082511_p82.jpg", "text": "80\nNACA TN No. 1826\n\n<!-- Image (207, 112, 740, 177) -->\n(a) Lifting element on the center line.\n\n<!-- Image (211, 277, 730, 393) -->\n(b) Tunnel arrangement that corresponds to omitting the additional\nshort strips.\n\n<!-- Image (195, 503, 750, 615) -->\n(c) Expanding or contracting jet.\n\n<!-- Image (195, 696, 738, 840) -->\n(d) Curving jet.\nNACA\n\nFigure 10.- Acceleration-potential analogies for the two-dimensional\nclosed-open-closed tunnel.", "timestamp": "2026-07-22T05:51:13.731427+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 8, "total_pages": 78, "image_filename": "19930082483_p8.jpg", "text": "```markdown\n6\nNACA TN No. 1807\n\nFull-Admission Turbine Performance\n\nIn the case of a turbine operating with full-peripheral admission, the ideal power output corrected to sea level is expressed as\n\n$$\n(\\text{ideal power})_C = \\frac{W \\Delta_s h'}{\\delta_i \\sqrt{\\theta_i}} \\quad (11)\n$$\n\nwhere subscript $i$ refers to the conditions at the inlet measuring station.\n\nLosses. - The principle turbine losses with full admission that must be subtracted from the ideal power are:\n\n1. rotor-tip leakage loss\n2. nozzle and rotor-blading aerodynamic losses\n3. shaft losses, which for a full-admission turbine include\n    (a) disk-windage loss\n    (b) bearing loss\n\n(1) Rotor-tip leakage loss: The percentage of active gas flow that passes over the rotor-blade tips may be estimated by the following expression as indicated in references 1 and 2:\n\npercentage of active gas flow through tip clearance space\n\n$$\n= \\frac{100 \\text{ C}}{\\text{nl} \\sin \\beta_2 + \\text{C}} \\quad (12)\n$$\n\n$$\n= K_I\n$$\n\nwhere\n\nC blade radial tip clearance, (ft)\n\nn thickness coefficient (unity for reaction turbines)\n\nl rotor-blade length, (ft)\n\n$\\beta_2$ rotor-blade exit angle relative to plane of rotor disk measured at pitch line, (deg)\n\n1032\n```", "timestamp": "2026-07-22T05:51:15.631283+00:00"}
{"citation_id": "19930086076", "source_url": "https://ntrs.nasa.gov/api/citations/19930086076/downloads/19930086076.pdf", "page_number": 23, "total_pages": 50, "image_filename": "19930086076_p23.jpg", "text": "NACA RM E9F09\n\n[Figure: Schematic diagram of flame holder 7. The figure shows a rectangular duct with two rows of V-shaped flame holders arranged in staggered pattern. Dimensions are annotated: overall height 8\", divided into two sections of 2 5/8\" and 2 3/8\". Each V-shape has leg length 1\" and spacing 1/2\" between adjacent vertices. Side view shows thickness 1/16\" and width 4\". NACA logo is present below the main diagram.]\n\nFigure 6. - Schematic diagram of flame holder 7.\n\n21", "timestamp": "2026-07-22T05:51:22.203975+00:00"}
{"citation_id": "19930085938", "source_url": "https://ntrs.nasa.gov/api/citations/19930085938/downloads/19930085938.pdf", "page_number": 30, "total_pages": 42, "image_filename": "19930085938_p30.jpg", "text": "```markdown\nNACA RM No. L9B04\n\nMaximum elevator deflection\nBoth configurations\n\nElevator deflection, deg\n\nStable\n\nSingle boom\n\nUnstable\n\nTwin boom\n\nLess than $2^\\circ$ trim\nat take-off\n\nCenter of gravity, percent M.A.C.\n\nNACA\n\nFigure 8.— Center-of-gravity limits of stability. Gross load coefficient 3.87; full power.\n\n29\n```", "timestamp": "2026-07-22T05:51:22.357847+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 23, "total_pages": 98, "image_filename": "19930086073_p23.jpg", "text": "NACA RM A9E04\n\n[Figure: A large white model of a wing plus body mounted on supports in a wind tunnel, with split flaps deflected. A person stands near the base for scale. The NACA logo and identifier \"A-11948\" are visible in the lower right corner of the image.]\n\n(b) Wing plus body; split flaps deflected $45.4^\\circ$.\n\nFigure 3.— Continued.\n\n21", "timestamp": "2026-07-22T05:51:28.622859+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 56, "total_pages": 96, "image_filename": "19930085880_p56.jpg", "text": "54\nNACA RM No. L9C03\n\n.64\n.56\n.48\n.40\n.32\n.24\n.16\n.08\n0\n\nSpeed\n(fps)\n10 O\n15 □\n20 ◇\n25 △\n30 ▽\n\nDraft, ft\n\n0 .05 .10 .15 .20 .25 .30 .35\nWetted area, sq ft\n\n(e) $\\tau = 20^\\circ$.\nFigure 17.- Concluded.", "timestamp": "2026-07-22T05:51:30.251503+00:00"}
{"citation_id": "19930086092", "source_url": "https://ntrs.nasa.gov/api/citations/19930086092/downloads/19930086092.pdf", "page_number": 18, "total_pages": 28, "image_filename": "19930086092_p18.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:51:31.834300+00:00"}
{"citation_id": "19930082472", "source_url": "https://ntrs.nasa.gov/api/citations/19930082472/downloads/19930082472.pdf", "page_number": 12, "total_pages": 34, "image_filename": "19930082472_p12.jpg", "text": "10\nNACA TN No. 1797\n\nmaximum lift is concerned boundary-layer effects resulting from wing sweep are not detrimental. It must be noted that the application of simple sweep theory in the above manner is not intended as a precise correction, but is used only to enable an approximate comparison between two-dimensional and three-dimensional values.\n\nAlleviation of Separation Effects\n\nThe problems involved in alleviating the poor characteristics of the swept-forward wing are evident from the foregoing discussion. These consist of a postponement of turbulent separation in the moderate lift range and a postponement of leading-edge separation in the high-lift range. Both postponements, of course, should be to an angle of attack at least as high as the maximum that can be encountered in flight. Beyond the flight range of angle of attack the stall progression must be such that no longitudinal instability results, since instability would possibly curtail the usable lift range.\n\nOf the two types of separation encountered, effort should first be directed toward delaying leading-edge separation. There are several reasons for this. The range over which leading-edge separation must be delayed is fairly small, from $16^\\circ$ to somewhere in the neighborhood of, say, $20^\\circ$ (a possible maximum ground angle). Hence, simple mechanical nose modifications, such as a plain leading-edge flap or a Kruger flap, should provide adequate control. Furthermore, any beneficial changes in leading-edge flow will be reflected as beneficial changes in trailing-edge flow and consequently delays in turbulent separation should result. On the other hand, the influence of trailing-edge flow on leading-edge separation is uncertain, and any benefits obtained by attempting to control turbulent separation first might soon be overshadowed by the detrimental effects of leading-edge separation.\n\nApplication to General Case of Swept Wings\n\nThe present investigation was concerned primarily with a particular configuration of a swept-forward wing. However, if reasonable consideration is given to the effects of physical changes, certain inferences can be drawn as to the behavior of other swept wings whether swept forward or swept back. The effects of separation and section stall on the characteristics of swept-back wings should be quite similar to the corresponding effects on the swept-forward wing. For a swept-back wing with like airfoil sections the first", "timestamp": "2026-07-22T05:51:32.041284+00:00"}
{"citation_id": "19930086097", "source_url": "https://ntrs.nasa.gov/api/citations/19930086097/downloads/19930086097.pdf", "page_number": 17, "total_pages": 36, "image_filename": "19930086097_p17.jpg", "text": "NACA RM A9H11 CONFIDENTIAL 15\n\nis accompanied by marked changes in the schlieren photographs of the axially symmetric flow. (See reference 8.) The trailing shock wave, which normally stands downstream of the boattailed base for laminar flow, moves upstream as transition is effected and attaches to the rim of the base, thereby reducing the base drag. The condition of the boundary-layer flow, therefore, should definitely be viewed as an important variable.\n\nAnother parameter that is expected to be important is the airfoil thickness ratio. If the thickness ratio is decreased, equations (9) and (10) indicate that the percentage drag reduction will also decrease substantially. This is easily explained on physical grounds since the drag reduction ultimately is obtained by a decrease in wave drag. The pressure drag, of course, progressively becomes a smaller fraction of the profile drag as the thickness ratio approaches zero. For very thin profiles, however, the boundary layer becomes thick compared to the trailing-edge height, and this should tend to reduce the base drag. The extent to which the profile drag can be reduced for airfoil ratios of, say, 5-percent-thickness ratio will have to be determined by future experiments.\n\nSince the ambient air can flow laterally around the wing tip and into the dead-air region behind the base, there probably is a tip-relieving effect of a finite span. This inflow would be expected to reduce the base drag, particularly at high supersonic Mach numbers, and hence it would appear that a finite aspect ratio would be more favorable for blunt-trailing-edge wings than an infinite aspect ratio. Again, experiments are needed to establish the importance of this variable.\n\nSome of the foregoing is, of course, conjectural in nature. The discussion of the various parameters that may affect the drag of blunt-trailing-edge wings has been given in order to emphasize the fact that there is as yet no simple answer to the question of whether blunt-trailing-edge airfoils can always be designed to have significantly lower drag than corresponding sharp-trailing-edge airfoils.\n\nTEST METHODS\n\nA description of the apparatus and the general procedure for testing wing models in the Ames 1- by 3-foot supersonic wind tunnel No. 1 may be found in reference 9. In order to simplify model construction as well as test methods, constant-chord wings of finite span were employed throughout the experimental phase of the present investigation. Each wing had an aspect ratio of 4 and was sting supported from the rear in the manner shown by the photograph in figure 7. The profile shape was the sole variable for the different wings tested. The dimensions of the various airfoil contours are given in figure 8. Wings 1, 2, 3, and 4, which have essentially the same section modulus, were tested only at zero angle of attack; whereas wings 5, 6, and 7, which have the same thickness ratio, were tested through the available angle-of-attack range.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:51:36.113692+00:00"}
{"citation_id": "19930085977", "source_url": "https://ntrs.nasa.gov/api/citations/19930085977/downloads/19930085977.pdf", "page_number": 26, "total_pages": 33, "image_filename": "19930085977_p26.jpg", "text": "CONFIDENTIAL\n\nDownwash angle, $\\epsilon$, deg\n\nM = 0.60\n\nM = 0.80\n\nM = 0.85\n\nM = 0.90\n\n$\\alpha$, deg -2, -1, 0, 1, 2, 3, 4, 6, 8, 10, 12\n\no □ ◇ △ ▲ ▽ ∇ ⊙\n\nDownwash angle, $\\epsilon$, deg\n\nM = 0.95\n\nM = 1.00\n\nM = 1.05\n\nM = 1.10\n\nTail height, $h_t$, percent semispan\n\n-80 -40 0 40 80\n\n-80 -40 0 40 80\n\n-80 -40 0 40 80\n\n-80 -40 0 40 80\n\nNACA\n\nCONFIDENTIAL\n\nFigure 10.- Effective downwash angles in region of tail plane for a model with $0^\\circ$ sweptback wing, aspect ratio 4, taper ratio 0.6, and NACA 65A006 airfoil section. Wing alone.\n\nNACA RM L9E22\n\n25", "timestamp": "2026-07-22T05:51:37.792448+00:00"}
{"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 65, "total_pages": 78, "image_filename": "19930082618_p65.jpg", "text": "NACA TN 1945\n\nSection lift coefficient, $c_l$\n\nSection pitching-moment coefficient, $c_{m_{c/4}}$\n\nSection angle of attack, $\\alpha$, deg\n\nR\n- 0.7 x $10^6$\n- 1.0\n- 1.5\n- 2.0\n- 2.6\n- 4.0\n- 6.0\n- 8.9\n\nFlagged symbols denote standard roughness\n\nNACA\n\n(a) Section lift and pitching-moment characteristics of the plain airfoil section.\n\nFigure 15.— Aerodynamic characteristics of the NACA 23015 airfoil section, 24-inch chord.\n\n63", "timestamp": "2026-07-22T05:51:39.502820+00:00"}
{"citation_id": "19930082474", "source_url": "https://ntrs.nasa.gov/api/citations/19930082474/downloads/19930082474.pdf", "page_number": 12, "total_pages": 21, "image_filename": "19930082474_p12.jpg", "text": "10\nNACA TN No. 1799\n\nDISCUSSION AND POSSIBLE SOLUTIONS\n\nThe basic purposes for making flying-qualities studies are to isolate the characteristics most in need of improvement and to find means for achieving these improvements. A discussion of a few examples of the lines of development which are suggested by the evaluation of flying qualities which have been given therefore seems in order.\n\nIn the authors' opinion the problem which seems to need investigation most urgently is the instability with angle of attack. One proposed solution to this problem is to provide stick forces in the proper direction, or stick-free stability. This proposal means that in maneuvers at constant speed pull forces are required to hold constant positive acceleration and push forces to hold negative acceleration. This solution does not alter the fact that the control moves in the wrong direction as the maneuver develops. Stick-free stability is considered to be essential for a completely satisfactory solution but is not, in itself, sufficient. First, the stick is never actually free because of friction; also, the pilot imposes some restraint on the stick, either consciously or unconsciously, because the stick will tend to move noticeable amounts in counteracting the stick-fixed instability. Second, and most important, the stick-free stability does not alter the fact that maneuvers (either intentional or due to gusts) can be severe enough that insufficient control for prompt recovery exists.\n\nIf the machine could be provided with stick-fixed stability with respect to angle of attack, the danger of loss of control would be virtually eliminated, and friction or pilot restraint of the stick would not affect the machine's tendency to maintain steady flight. Maneuvers could be executed without reversing the stick motion, and recovery could be made by simply returning the stick to the trim position. Stick-free stability could be provided in this case by mechanical means such as simple springs.\n\nSince the instability with angle of attack arises as a result of forward speed and is greatest at the highest speeds, to attempt to obtain the desired stabilizing forces by using some form of horizontal tail surface mounted on the fuselage seems logical. This use of a horizontal tail surface is particularly valid, of course, for overcoming the instability of the fuselage itself. Rotor instability could more logically be eliminated by self-contained means, but the more practical immediate solution may nevertheless lie in the use of some form of horizontal tail surface. Preliminary calculations indicate that a rather small tail area should suffice; for example, calculations for a sample two-place helicopter indicated that about 4 square feet would be needed to stabilize the fuselage and that an additional area of about 4 square feet should serve to stabilize the rotor.", "timestamp": "2026-07-22T05:51:42.418101+00:00"}
{"citation_id": "19930085914", "source_url": "https://ntrs.nasa.gov/api/citations/19930085914/downloads/19930085914.pdf", "page_number": 38, "total_pages": 42, "image_filename": "19930085914_p38.jpg", "text": "NACA RM A9D25\n37\n\nLift coefficient, $C_L$\nfor M=20\n\n| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | |", "timestamp": "2026-07-22T05:51:42.625200+00:00"}
{"citation_id": "19930091987", "source_url": "https://ntrs.nasa.gov/api/citations/19930091987/downloads/19930091987.pdf", "page_number": 10, "total_pages": 12, "image_filename": "19930091987_p10.jpg", "text": "```markdown\n6\nREPORT 922—NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nThe pitching-moment-coefficient results, though not so complete as the lift-coefficient and drag-coefficient data, show changes as the speed is increased from subcritical to supercritical values that are in general of the same character as the lift-coefficient and drag-coefficient changes. At subcritical Mach numbers the aspect ratio has but little influence on the values of the moment about the wing quarter-chord axis. The slope of the curve of moment against lift is slightly positive, indicating that for the section used, the aerodynamic center is slightly ahead of the wing quarter-chord axis. For all Mach numbers up to 0.7, as indicated by figures 5 (a) and 5 (b), the aspect ratio has but little influence on the moment coefficient for a given lift coefficient. Increase of Mach number up to the critical value causes, in accordance with the known theories, a small increase in the moment coefficient. When large supercritical values of the Mach number are reached, however, drastic changes in the wing quarter-chord moment coefficient are found for the high-aspect-ratio wings. Abrupt changes in the variation of the moment coefficient with the lift coefficient occur at very low lift coefficients and at higher lift coefficients these curves tend to give a stable slope as distinguished from the usual unstable slope characteristics of low-speed data. This change in moment characteristics has been shown previously and is due to the movement of the shock on the wing.\n\nReduction of aspect ratio reduces the changes in moment coefficient, as is shown in figure 5 by the data for wings of aspect ratios 2 and 3. The slope remains positive for all Mach numbers up to 0.9, and the change in moment coefficient from the values obtained at or near the wing-section critical Mach number (0.72) as compared with the changes for the wings of conventional aspect ratios is relatively small up to the highest speeds investigated.\n\nThe over-all effects of reducing the aspect ratio on improving the undesirable wing characteristics are very great. The absence of irregularities in the lift curve, the indicated freedom of the lift curve from drastic slope changes, and the similar effects for the wing moment curve indicate that the serious stability changes which have occurred with conventional aircraft when flown in the supercritical region may be alleviated in large degree. Likewise the delayed drag rise and the less rapid rate of drag increase at the high supercritical Mach numbers permit increased speed.\n\nThe improved supercritical-speed characteristics found for the low-aspect-ratio wings are a consequence of the three-dimensional type of flow at the tip. Because the effects of the flow at the tip are quite large, the tip shape is likely to be of great importance. In the present experiments the tip shape was made square principally as a matter of convenience in using an existing model to investigate the over-all effect. The square tip leads to large local velocities and at low speeds is known to produce undesirable disturbances. It is likely, therefore, that an appreciable improvement in the low-aspect-ratio-wing characteristics may be obtained by suitably shaping the tip. Large local velocities that would occur over the forward and middle parts of the tip can lead to large disturbances, probably involving shock, which might produce at least partly separated flows over the rear part of the wing. General considerations of the flow about the tip indicate that a change of plan form from the square type used in these illustrative experiments to a tapered plan form giving reduced chord at the tip and a rounded or elliptical tip shape may produce further favorable effects. Likewise, a thinner section of late critical Mach number type can be expected to delay the onset of the drag rise until much higher speeds. These experiments indicate that the serious adverse compressibility phenomena, particularly as regards drag and lift, are delayed to Mach numbers exceeding 0.9 by a low-aspect-ratio wing of rectangular plan form with a conventional 12-percent-thick section. Use of a 10-percent-thick wing of late-critical-speed type will, on the basis of two-dimensional data for the wing sections, give a further rise of 0.08 in the critical Mach number. This change together with an improved tip shape and plan form appears to offer a new possibility of overcoming the existing problems of flight in the transonic speed range.\n\nThough not specifically shown by the present results, two other advantages are offered by the low-aspect-ratio wing. First, thin sections giving high critical speeds may be used without the usually imposed condition of inadequate wing depth for an efficient structure, and second, the spanwise center-of-pressure shift in the supercritical region will be much reduced because of the short spanwise length.\n\nCONCLUSIONS\n\nThe detrimental effects of compressibility in the supercritical speed range on the stability and performance of aircraft are alleviated to very great degree by the use of low-aspect-ratio lifting surfaces. Further consideration of the advantages of this type of configuration is warranted.\n\nLANGLEY MEMORIAL AERONAUTICAL LABORATORY\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS,\nLANGLEY FIELD, VA., September 6, 1945.\n\nREFERENCES\n\n1. Stack, John: Compressible Flows in Aeronautics. Jour. Aero. Sci., vol. 12, no. 2, April 1945, pp. 127-143.\n2. Stack, John, Lindsey, W. F., and Littell, Robert E.: The Compressibility Burble and the Effect of Compressibility on Pressures and Forces Acting on an Airfoil. NACA Rep. 646, 1938.\n3. Hilton, W. F.: The Photography of Aircrew Sound Waves. Proc. Roy. Soc. (London), ser. A, vol. 169, no. 937, Dec. 22, 1938, pp. 174-190.\n4. Stack, John: Tests of Airfoils Designed to Delay the Compressibility Burble. NACA Rep. 763, 1943.\n5. Allen, H. Julian, and Vincenti, Walter G.: The Wall Interference in a Two-Dimensional-Flow Wind Tunnel with Consideration of the Effect of Compressibility. NACA Rep. 782, 1944.\n6. Ferri, Antonio: Investigations and Experiments in the Guidonia Supersonic Wind Tunnel. NACA TM 901, 1939.\n7. Byrne, Robert W.: Experimental Construction Effects in High-Speed Wind Tunnels. NACA ACR L4L07a, 1944.\n\nU. S. GOVERNMENT PRINTING OFFICE: 1949\n```", "timestamp": "2026-07-22T05:51:53.222611+00:00"}
{"citation_id": "19930091993", "source_url": "https://ntrs.nasa.gov/api/citations/19930091993/downloads/19930091993.pdf", "page_number": 9, "total_pages": 21, "image_filename": "19930091993_p9.jpg", "text": "ANALYSIS OF PERFORMANCE OF JET ENGINE FROM CHARACTERISTICS OF COMPONENTS 5\n\nTURBINE PERFORMANCE\n\nFor correlation of the turbine performance, the corrected turbine gas flow was plotted with the equivalent isentropic enthalpy drop. No effect of speed was observed; a definite grouping of the data, however, took place according to the turbine-inlet temperature. This grouping is shown in figure 7 where the data for turbine-inlet temperatures of $190^\\circ$, $500^\\circ$, and $1250^\\circ$ F are plotted. The curve for an inlet temperature of $190^\\circ$ F, shown for comparison, was obtained from the data of reference 4, which reports the turbine performance obtained from calibrating the turbine component alone by the use of air at $190^\\circ$ F. The gas flow at $1250^\\circ$ F is from 6 to 17 percent higher than the gas flow for the same pressure ratio at $190^\\circ$ F and was believed to be caused by variation in the leakage flow through the blade-tip clearance space, which changed with gas temperature. In order to account for this discrepancy, a very crude method was used to estimate the leakage. According to reference 7, the leakage flow may be estimated by assuming twice the mass flow per unit area in the clearance space as for the rest of the annular area of the turbine. If a linear variation of the temperature of the wheel and the blades varying from $200^\\circ$ F at the center of the wheel to the turbine-inlet stagnation gas temperature at the tips of the blades and a casing temperature equal to turbine-inlet stagnation gas temperature are assumed, the clearance flow space could be computed for each temperature. The wheel runs cooler than the insulated case and as a result of the metal expansion, the clearance will expand with increasing temperature. When the gas-flow data were corrected for clearance variations from measured values of $W_2$ to values of $W_2'$, which correspond to operation at a turbine-inlet temperature of $190^\\circ$ F, the data showed no variation of gas flow for temperature, rotative speed of the turbine, or indicated whether the data were obtained from engine- or turbine-component experiments (fig. 8). All turbine-performance parameters for which the symbols are primed have thus been corrected for the effect of leakage. The enthalpy drop is corrected by assuming that no work was obtained from the leakage air.\n\nIn order to obtain complete correlation of the turbine performance, an additional chart presenting corrected work output at constant values of equivalent speed is required.\n\nBecause of the inability to measure accurately the turbine-inlet temperature, the turbine equivalent speed could not be accurately set at preassigned values during engine operation and turbine data could not be directly plotted for constant equivalent speeds. An auxiliary chart was therefore prepared by plotting curves of $W_2'\\Delta H_{t,2}/(n\\sigma_{r,2}\\delta_{r,2})$ against $W_2'n/(\\sigma_{r,2}\\delta_{r,2})$ for constant assumed values of $n/\\sqrt{\\theta_{r,2}}$. These curves were directly obtained from the faired curve of figure 8 (b), which showed no speed effect and which was considered to have perfect correlation. The efficiency contours as indicated by the data were then drawn and the corrected work output was determined from the isentropic work and efficiency values. A typical curve of corrected equivalent enthalpy drop against the equivalent isentropic enthalpy drop for an equivalent speed of 135 rps is shown by the solid curve in figure 9. Also shown are data obtained for equivalent turbine speeds between 130 and 140 rps. There are two sets of engine data shown: one set from the engine with the original settings of the compressor blades that correspond to the performance curves in figures 5 and 6; and the second blade settings, which were merely used to explore a larger operational range of the turbine. The importance of this procedure may be seen from figure 9, which shows that the data obtainable with the original blade settings are insufficient to give reliable efficiency curves because of the short range and the experimental errors.\n\n<!-- Image (496, 270, 891, 785) -->\n\nFIGURE 7.—Equivalent isentropic enthalpy drop and gas flow through turbine with various gas temperatures.", "timestamp": "2026-07-22T05:51:53.685727+00:00"}
{"citation_id": "19930085972", "source_url": "https://ntrs.nasa.gov/api/citations/19930085972/downloads/19930085972.pdf", "page_number": 33, "total_pages": 46, "image_filename": "19930085972_p33.jpg", "text": "NACA RM L9B18\n31\n\n[Figure: A graph plotting aerodynamic coefficients. The x-axis represents Lift coefficient, $C_L$. The left y-axis represents Angle of attack, $\\alpha$, deg. The right y-axis represents Longitudinal-force coefficient, $C_X$. The top y-axis represents Pitching-moment coefficient, $C_m$. The graph contains three sub-plots with data points marked by circles, squares, and triangles. A legend indicates symbols for $i_t$ (deg): 0, 3, and tail off. The NACA logo is present at the bottom of the graph.]\n\n(b) Alternate tail position.\nFigure 9.- Concluded.", "timestamp": "2026-07-22T05:51:55.735308+00:00"}
{"citation_id": "19930086092", "source_url": "https://ntrs.nasa.gov/api/citations/19930086092/downloads/19930086092.pdf", "page_number": 19, "total_pages": 28, "image_filename": "19930086092_p19.jpg", "text": "NACA RM A9F14 CONFIDENTIAL 17\n\n[Figure: A black and white photograph showing a close-up side view of a model aircraft or missile mounted on a test stand. The model has a pointed nose cone, swept-back wings, and tail fins. It is held in place by several thin rods attached to its body. In the upper right corner of the photo, there is a small triangular logo with \"NACA\" written inside it and the number \"A-12456\" below it.]\n\n(b) Close-up side view.\n\nFigure 2.- Concluded.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:52:13.460726+00:00"}
{"citation_id": "19930085938", "source_url": "https://ntrs.nasa.gov/api/citations/19930085938/downloads/19930085938.pdf", "page_number": 31, "total_pages": 42, "image_filename": "19930085938_p31.jpg", "text": "```markdown\n30\n\n18\nMaximum trim\n16\nTwin boom\n14\nSingle boom\n12\nSingle-boom lower-limit origin\nTrim, deg\n10\nTwin-boom lower-limit origin\nStable\n8\n6\nBoth configurations\nUnstable\n4\n2\n0\n0 1.0 2.0 3.0 4.0 5.0 6.0 7.0 8.0 9.0 10.0\nSpeed coefficient, $C_V$\n\nFigure 9.- Trim limits of stability.\n\nNACA\nNACA TM NO. 10904\n```", "timestamp": "2026-07-22T05:52:13.783010+00:00"}