PDF/ocr_dataset_1049.jsonl

{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 83, "total_pages": 99, "image_filename": "19930082511_p83.jpg", "text": "NACA TN No. 1826\n81\n\nInsulator\nMetal\n\n(a) Closed-open tunnel.\n\n(b) Open section represented by many\nlongitudinal strips.\n\n(c) Open section represented by only\ntwo longitudinal strips.\n\nFigure 11.- Acceleration-potential analogy for three-dimensional\nclosed-open tunnel and two approximate acceleration-potential\nanalogies for three-dimensional closed-open-closed tunnels.\nThe closed-open analogy may also be considered as an approxi-\nmate analogy for the closed-open-closed tunnel.", "timestamp": "2026-07-22T05:52:16.443364+00:00"}
{"citation_id": "19930093769", "source_url": "https://ntrs.nasa.gov/api/citations/19930093769/downloads/19930093769.pdf", "page_number": 12, "total_pages": 39, "image_filename": "19930093769_p12.jpg", "text": "NACA RM No. E5L10a CONFIDENTIAL 11\n\nthe walls of both the test chamber and the tail pipe with the AN-F-58 fuel than were encountered with gasoline. A considerable amount of liquid fuel was also present on the surfaces of the tank and in puddles on the bottom of the test chamber after an engine shutdown following altitude operation with AN-F-58 fuel. Similar phenomena were not usually observed with gasoline.\n\nSUMMARY OF RESULTS\n\nThe following results were obtained in a comparison of the altitude performance of AN-F-58 fuel and gasoline in a 3000-pound-thrust turbojet engine:\n\n1. Satisfactory operation of the engine was obtained with AN-F-58 fuel over a range of engine speeds for altitudes from 5000 to 50,000 feet and for flight Mach numbers from 0.25 to 1.00. At altitudes of 45,000 and 50,000 feet, the maximum operable engine speed was limited to values less than the rated speed by excessive tail-pipe gas temperatures.\n\n2. The net thrust and average tail-pipe gas temperatures were approximately the same for both fuels for altitudes from 5000 to 50,000 feet and flight Mach numbers from 0.60 to 1.00. The specific fuel consumption and combustion efficiency at the maximum engine speeds investigated were approximately the same for both fuels at altitudes up to 35,000 feet, but at an altitude of 50,000 feet the specific fuel consumption was about 9 percent higher and the combustion efficiency correspondingly lower with the AN-F-58 fuel than with gasoline.\n\n3. The low-engine-speed blow-out limits for the two fuels at a flight Mach number of 0.60 were about the same and differed only slightly at a flight Mach number of 0.25.\n\n4. Ignition of AN-F-58 fuel with the standard spark plug was possible only with the spark plug in a clean condition; ignition was impossible at all flight conditions investigated when the plug was fouled by an accumulation of liquid fuel from a preceding false start. The use of an extended-electrode spark plug with AN-F-58 fuel provided satisfactory ignition over a slightly smaller range of altitudes and flight Mach numbers than for gasoline with the standard spark plug. Zero-ram, sea-level starts with AN-F-58 fuel were successful at inlet-air temperatures as low as -50° F.\n\n5. The radial temperature gradients at the turbine outlet were about the same with both fuels at an altitude of 20,000 feet; at an altitude of 50,000 feet, the difference in the average\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:52:16.967792+00:00"}
{"citation_id": "19930082472", "source_url": "https://ntrs.nasa.gov/api/citations/19930082472/downloads/19930082472.pdf", "page_number": 13, "total_pages": 34, "image_filename": "19930082472_p13.jpg", "text": "NACA TN No. 1797\n\noccurrence of separation might be expected to be turbulent separation over the outboard area. Again the center of pressure of the sections affected would be shifted aft resulting in more negative pitching moments. Turbulent separation would again be expected to be followed by leading-edge separation. Loss of lift due to leading-edge separation will occur first at the outboard area and, as angle of attack is further increased, will occur over sections farther inboard. Thus, loss of section lift will travel forward relative to the moment center and a tendency toward longitudinal instability will result. On this basis, alleviation of the poor characteristics of swept-back wings can be approached along the same line as previously described for the swept-forward wing.\n\nCONCLUDING REMARKS\n\nTests made on a $45^\\circ$ swept-forward wing showed the flow conditions underlying the poor longitudinal characteristics of the wing in the moderate- and high-lift range.\n\nIn the moderate-lift range ($C_L = 0.5$ to $0.7$), the occurrence of turbulent separation caused a chordwise redistribution of load over the inboard sections. This caused increases in drag and a rearward shift of the aerodynamic center (from $0.26\\bar{c}$ to $0.43\\bar{c}$) but caused no loss of lift.\n\nIn the high-lift range ($C_L = 0.7$ to $1.04$), the occurrence of leading-edge separation caused a loss of section lift that occurred first over the inboard sections and traveled outward as angle of attack was increased. This caused very large increases in drag, a decreased lift-curve slope, and, due to the changes in span-wise loading, caused an extremely large forward shift of aerodynamic center (from $0.43\\bar{c}$ to $0.05\\bar{c}$ forward of the leading edge).\n\nIn order to improve the longitudinal characteristics of the swept-forward wing, both forms of separation must be postponed. The evidence indicates that effort should be directed first toward postponing leading-edge separation. Only after leading-edge separation is adequately postponed should control of the turbulent boundary layer be attempted.\n\nAmes Aeronautical Laboratory,  \nNational Advisory Committee for Aeronautics,  \nMoffett Field, Calif.", "timestamp": "2026-07-22T05:52:19.307665+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 24, "total_pages": 98, "image_filename": "19930086073_p24.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:52:21.359930+00:00"}
{"citation_id": "19930093789", "source_url": "https://ntrs.nasa.gov/api/citations/19930093789/downloads/19930093789.pdf", "page_number": 11, "total_pages": 29, "image_filename": "19930093789_p11.jpg", "text": "10 CONFIDENTIAL NACA RM No. E8I21\n\nleakage air flow is mainly from the boundary layer at the outer radius of the flow annulus, this air contains only a small portion of the total energy passing through the turbine.\n\nFor configuration 3 (figs. 8(c) and 13), the brake efficiency was 0.82 at the total-pressure ratio and equivalent mean blade speed for which the blades were designed.\n\nSUMMARY OF RESULTS\n\nAn investigation of the effects of stator cone angle and blade-tip leakage on turbine performance was conducted with a turbine having 40-percent reaction and a design assumption of constant static pressure over the blade height. Caps at the tips of the rotor blades formed a continuous rotating shroud. The turbine was operated with a turbine-entrance temperature of approximately $660^\\circ$ R at total-pressure ratios from 1.25 to 3.70 and equivalent mean rotor-blade speeds of 188 to 855 feet per second. The following results were obtained:\n\n1. With the $0^\\circ$-cone-angle stator, the peak brake efficiency was approximately 0.04 higher than with the $70^\\circ$-cone-angle stator. With the $70^\\circ$-cone-angle stator, separation probably occurred at the tips of the rotor blades; the performance was improved by changing the cone angle to $0^\\circ$ and making the stator-blade height equal to the rotor-blade height, thereby eliminating the separation.\n\n2. Replacing the labyrinth, no-leakage shroud with a cylindrical stationary shroud that had a radial clearance of 0.016 of the blade height from the cylindrical rotating shroud and using the $0^\\circ$-cone-angle stator produced no measurable change in brake efficiency.\n\n3. The maximum brake efficiency was obtained with the blade-to-jet speed ratio between 0.50 and 0.55.\n\nLewis Flight Propulsion Laboratory,  \nNational Advisory Committee for Aeronautics,  \nCleveland, Ohio.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:52:22.692562+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 9, "total_pages": 78, "image_filename": "19930082483_p9.jpg", "text": "NACA TN No. 1807\n\nThe leakage power loss is taken as the product of the leakage gas weight flow and the isentropic enthalpy drop based on the total-pressure ratio across the turbine. The leakage loss was calculated by the expression\n\n$$\n\\text{leakage loss} = K_{T} W_{a} h'\n$$\n\nIn order to correct to sea-level conditions, the loss, as obtained by the preceding expression, is multiplied by the factor\n\n$$\n\\frac{1}{\\delta_{1} \\sqrt{\\theta_{1}}}\n$$\n\n(2) Nozzle and rotor-blading aerodynamic losses: Friction, separation, and other viscous effects result in losses as they affect the gas-velocity components in equation (2) in any operating passage. In this investigation these losses are not directly determined.\n\n(3) Shaft losses:\n\n(a) Disk-windage loss. The disk-windage (disk-friction) loss was taken as the power required to rotate the rotor disk, without blades, against the frictional drag of the relatively stagnant gases in the clearance space on each side of the rotor disk. This power loss may be represented by the following expression adapted from equation [7] of reference 3:\n\n$$\n\\text{disk power loss} = K_{II} \\left( \\frac{\\rho_{d} d_{h} u_{h}}{\\mu_{d}} \\right)^{-0.12} \\left( \\frac{N}{1000} \\right)^{3} d_{h}^{5} \\rho_{d}\n$$\n\nwhere\n\n$K_{II}$ empirical constant for disk-windage loss to be determined for a particular turbine by tests\n\n$d$ rotor-disk diameter, (ft)\n\n$\\mu$ absolute viscosity, ((lb)(sec)/sq ft)\n\n$N$ rotational speed, (rpm)\n\nSubscripts $h$ and $d$ refer to the blade-root position and the fluid surrounding the turbine-rotor disk, respectively.", "timestamp": "2026-07-22T05:52:23.051476+00:00"}
{"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 66, "total_pages": 78, "image_filename": "19930082618_p66.jpg", "text": "```markdown\n64\n\nSection lift coefficient, $c_l$\nMoment coefficient, $c_{m_{c/4}}$\nSection angle of attack, $\\alpha$, deg\n\nR\n$\\circ$ 0.7 x $10^6$\n$\\square$ 1.0\n$\\diamond$ 1.5\n$\\triangle$ 2.0\n$\\triangledown$ 5.0\nFlagged symbols denote\nstandard roughness\n\nNACA\n\n(b) Section lift and pitching-moment characteristics of the NACA 23015 airfoil section with a\n0.20c simulated split flap deflected 60°.\n\nFigure 15.— Continued.\n\nNACA TN 1945\n```", "timestamp": "2026-07-22T05:52:23.777474+00:00"}
{"citation_id": "19930085977", "source_url": "https://ntrs.nasa.gov/api/citations/19930085977/downloads/19930085977.pdf", "page_number": 27, "total_pages": 33, "image_filename": "19930085977_p27.jpg", "text": "CONFIDENTIAL\n\n| | | | |\n| :--- | :--- | :--- | :--- |\n| M = 0.60 | M = 0.80 | M = 0.85 | M = 0.90 |\n| [Graph: Downwash angle vs Tail height] | [Graph: Downwash angle vs Tail height] | [Graph: Downwash angle vs Tail height] | [Graph: Downwash angle vs Tail height] |\n\n$\\alpha$, deg -2, -1, 0, 1, 2, 3, 4, 6, 8, 10, 12\n$\\circ$ $\\square$ $\\diamond$ $\\triangle$ $\\triangleright$ $\\nabla$ $\\nabla$ $\\nabla$ $\\nabla$ $\\nabla$\n\n| | | | |\n| :--- | :--- | :--- | :--- |\n| M = 0.95 | M = 1.00 | M = 1.05 | M = 1.10 |\n| [Graph: Downwash angle vs Tail height] | [Graph: Downwash angle vs Tail height] | [Graph: Downwash angle vs Tail height] | [Graph: Downwash angle vs Tail height] |\n\nTail height, $h_t$, percent semispan\nCONFIDENTIAL\n\nFigure 11.- 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-fuselage.\n\n26\nNACA RM L9E22", "timestamp": "2026-07-22T05:52:24.040109+00:00"}
{"citation_id": "19930086097", "source_url": "https://ntrs.nasa.gov/api/citations/19930086097/downloads/19930086097.pdf", "page_number": 18, "total_pages": 36, "image_filename": "19930086097_p18.jpg", "text": "16 CONFIDENTIAL NACA RM A9H11\n\nDrag and lift forces were measured by means of a strain-gage balance which was shrouded from the external flow. Since the pressure in the balance chamber was greater than the free-stream static pressure, an appropriate correction for the \"piston effect\" has been applied to each drag measurement. This correction is based on the measured value of the pressure in the balance chamber and normally amounted to about 10 percent of the uncorrected force data. In reducing the profile drag data to coefficient form, an estimated correction of 0.0025 has been applied in each case to approximately account for the tare drag of the sting support. Because of imperfect alinement of the wings with the oncoming flow, a small lift force was measured on the symmetrical profiles with the wings nominally at zero angle of attack. Consequently, a correction based on the measured lift and linearized wing theory has been applied to the drag measurements in order to account for the small amount of drag due to lift. This latter correction usually amounted to 1 or 2 percent of the profile drag.\n\nSince the Reynolds number of each wing is about 1 million at the highest tunnel pressure, laminar flow would be expected over the entire wing surface. This expectation was verified by the liquid-film technique, the details of which have been described in reference 10. Hence, in order to simulate the case of a turbulent boundary layer approaching the base, it was necessary to add artificial roughness to the wing surfaces. This was done by applying a narrow band of salt crystals on both sides of the wing at approximately the 25-percent-chord position. It is known that the addition of artificial roughness at supersonic speeds invariably produces a certain increment of wave drag which must be accounted for if the measured drag is to correspond approximately to conditions of natural transition. This incremental wave drag was estimated from the measured increase in profile drag caused by the addition of roughness to the double-wedge profile (wing 1). The accompanying change in friction drag was approximately accounted for by assuming low-speed skin-friction coefficients and the existence of turbulent flow over the rear half of the chord. The wave drag due to roughness, as estimated in this manner, has been subtracted from all data representing cases where artificial roughness was used.\n\nThe data presented have not been corrected for nonuniformities in the free stream. The small inaccuracies in the experimental technique, together with the fact that in the present tests no corrections have been applied for the stream nonuniformities, may introduce errors of the order of ±5 percent in the absolute value of the lift-curve slopes and drag coefficients. Such uncertainties, however, will not introduce any significant error in the difference between the force coefficients of two wings of identical plan form that have the same sting support, the same artificial roughness, and are tested in same position along the nozzle axis. In view of these common test conditions the measured increments in lift and in minimum profile drag of the various wings are believed to be practically unaffected by the possible experimental errors discussed above.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:52:27.938080+00:00"}
{"citation_id": "19930086076", "source_url": "https://ntrs.nasa.gov/api/citations/19930086076/downloads/19930086076.pdf", "page_number": 24, "total_pages": 50, "image_filename": "19930086076_p24.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:52:32.964535+00:00"}
{"citation_id": "19930091993", "source_url": "https://ntrs.nasa.gov/api/citations/19930091993/downloads/19930091993.pdf", "page_number": 10, "total_pages": 21, "image_filename": "19930091993_p10.jpg", "text": "```markdown\n6\nREPORT 928—NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\n<!-- Image (107, 68, 939, 754) -->\n\nFIGURE 8.—Equivalent isentropic enthalpy drop and gas flow through turbine corrected for clearance expansion.\n\nAlso shown on the charts for comparison are the results obtained from the turbine-component experiments (reference 4). The discrepancies shown are unaccounted for. In some cases agreement is good, but where there is disagreement the data obtained in the engine experiments show a more rapid variation of efficiency with variation in enthalpy drop. The complete plot of turbine performance is shown in figure 10. A speed for maximum efficiency is 155 rps, which is closer to the design value of 136 rps than was the indicated optimum speed (180 rps) measured with cold air (reference 4). One possible reason for the shift in speed at maximum efficiency is the heat loss from the gas at high\n```", "timestamp": "2026-07-22T05:52:37.196857+00:00"}
{"citation_id": "19930085972", "source_url": "https://ntrs.nasa.gov/api/citations/19930085972/downloads/19930085972.pdf", "page_number": 34, "total_pages": 46, "image_filename": "19930085972_p34.jpg", "text": "32\nNACA RM L9B18\n\n<!-- Image (59, 161, 891, 796) -->\n\nFigure 10.- Variation of longitudinal-stability parameters with angle of sweepback for a variable-sweep model. Basic tail position; wing without cutout.", "timestamp": "2026-07-22T05:52:38.326875+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 57, "total_pages": 96, "image_filename": "19930085880_p57.jpg", "text": "NACA RM No. L9C03\n55\n\nLoad, lb\n32\n28\n24\n20\n16\n12\n8\n4\n0\n0 .05 .10 .15 .20 .25 .30 .35\nWetted area, sq ft\nSpeed\n(fps)\n35\n30\n25\n15\n10\n(a) $\\tau = 4^\\circ$.\n[Figure: NACA logo]\nFigure 18.- Variation of load with wetted area. Model 250B.", "timestamp": "2026-07-22T05:52:41.980790+00:00"}
{"citation_id": "19930086092", "source_url": "https://ntrs.nasa.gov/api/citations/19930086092/downloads/19930086092.pdf", "page_number": 20, "total_pages": 28, "image_filename": "19930086092_p20.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:52:43.414794+00:00"}
{"citation_id": "19930085938", "source_url": "https://ntrs.nasa.gov/api/citations/19930085938/downloads/19930085938.pdf", "page_number": 32, "total_pages": 42, "image_filename": "19930085938_p32.jpg", "text": "NACA RM No. L9B04\n31\n\nElevator deflection, deg\nTrim, deg\nSpeed coefficient, $C_V$\nLower limit\nC.G., 20 percent M.A.C.\n\nElevator deflection, deg\nTrim, deg\nSpeed coefficient, $C_V$\nLower limit\nC.G., 30 percent M.A.C.\n\nTrim, deg\nSpeed coefficient, $C_V$\nLower limit\nC.G., 40 percent M.A.C.\n\nTrim, deg\nSpeed coefficient, $C_V$\nLower limit\nNACA\n\n(a) Single boom.\n(b) Twin boom.\nFigure 10.— Variation of trim with speed.", "timestamp": "2026-07-22T05:52:47.106001+00:00"}
{"citation_id": "19930082474", "source_url": "https://ntrs.nasa.gov/api/citations/19930082474/downloads/19930082474.pdf", "page_number": 13, "total_pages": 21, "image_filename": "19930082474_p13.jpg", "text": "NACA TN No. 1799\n\nOne obvious disadvantage resulting from the use of the tail surface lies in the undesired vertical loads and pitching moments developed in hovering and vertical flight. For the areas mentioned these forces are actually quite small but may be further reduced, if desired, by using a biplane tail surface which would present less projected area in vertical flow or by using a free-floating tail surface arranged to be effective only in forward flight. More serious problems may arise from the fact that, in forward flight, a change from level flight to climb or to autorotation results in a sizable change in the angle of attack of the tail surface. This change occurs because the attitude angle of the helicopter remains roughly constant while the flight-path angle changes. This situation suggests that for at least the faster and more highly powered helicopters the tail surface should be made to move in conjunction with the pitch controls or should be made free-floating.\n\nThese problems and a number of details concerning the rotor downwash need further clarification before the helicopter designer can be expected to make full use of the tail surface as a cure for the angle-of-attack instability.\n\nAn improvement in the hovering characteristics should also be possible. Control sensitivity could be reduced by changing the control-system gearing, but this change is undesirable because it would limit the control available for trim unless a nonlinear system were used. A more logical solution would be to provide the pilot with a stick-force gradient which is suitably proportioned to the control sensitivity. In this regard the effects of size tend to be contradictory. In other words, the smaller the helicopter the greater its control sensitivity but the smaller the probable force gradient, and vice versa; whereas the greater sensitivity should be accompanied by a larger force gradient.\n\nControl sensitivity could also be reduced by increasing the damping and thus reducing the rate of roll. One way of making this reduction involves increasing the control lag by changing the rotor characteristics. Control lag, however, should not be increased to more than perhaps three or four times that of the subject helicopter, or more than perhaps 0.2 to 0.3 second, as it may lead to overcontrol of a different type than that mentioned previously and one which is more dangerous because of larger amplitude. A better solution would be to increase damping without changing lag.\n\nFriction in the control system should be kept to a minimum or to a value which will permit good self-centering characteristics. Undesirable transient control forces in maneuvers, as well as excessive vibratory stick forces, should be prevented from reaching the pilot by means of irreversible mechanisms rather than by introduction of large amounts of friction. The desired control feel can then be introduced on the pilot's side of the irreversible mechanism.", "timestamp": "2026-07-22T05:52:49.176470+00:00"}
{"citation_id": "19930091987", "source_url": "https://ntrs.nasa.gov/api/citations/19930091987/downloads/19930091987.pdf", "page_number": 11, "total_pages": 12, "image_filename": "19930091987_p11.jpg", "text": "Positive directions of axes and angles (forces and moments) are shown by arrows\n\n| Axis | | Force (parallel to axis) symbol | Moment about axis | | | Angle | | Velocities | |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| Designation | Symbol | | Designation | Symbol | Positive direction | Designation | Symbol | Linear (component along axis) | Angular |\n| Longitudinal<br>Lateral<br>Normal | $X$<br>$Y$<br>$Z$ | $X$<br>$Y$<br>$Z$ | Rolling<br>Pitching<br>Yawing | $L$<br>$M$<br>$N$ | $Y \\rightarrow Z$<br>$Z \\rightarrow X$<br>$X \\rightarrow Y$ | Roll<br>Pitch<br>Yaw | $\\phi$<br>$\\theta$<br>$\\psi$ | $u$<br>$v$<br>$w$ | $p$<br>$q$<br>$r$ |\n\nAbsolute coefficients of moment\n$C_l = \\frac{L}{qbS}$ (rolling)\n$C_m = \\frac{M}{qcS}$ (pitching)\n$C_n = \\frac{N}{qbS}$ (yawing)\n\nAngle of set of control surface (relative to neutral position), $\\delta$. (Indicate surface by proper subscript.)\n\n**4. PROPELLER SYMBOLS**\n\n$D$ Diameter\n$p$ Geometric pitch\n$p/D$ Pitch ratio\n$V'$ Inflow velocity\n$V_s$ Slipstream velocity\n$T$ Thrust, absolute coefficient $C_T = \\frac{T}{\\rho n^2 D^4}$\n$Q$ Torque, absolute coefficient $C_Q = \\frac{Q}{\\rho n^2 D^5}$\n\n$P$ Power, absolute coefficient $C_P = \\frac{P}{\\rho n^3 D^5}$\n$C_s$ Speed-power coefficient $= \\sqrt[5]{\\frac{\\rho V^3}{P n^3}}$\n$\\eta$ Efficiency\n$n$ Revolutions per second, rps\n$\\Phi$ Effective helix angle $= \\tan^{-1} \\left( \\frac{V}{2\\pi r n} \\right)$\n\n**5. NUMERICAL RELATIONS**\n\n1 hp = 76.04 kg-m/s = 550 ft-lb/sec\n1 metric horsepower = 0.9863 hp\n1 mph = 0.4470 mps\n1 mps = 2.2369 mph\n\n1 lb = 0.4536 kg\n1 kg = 2.2046 lb\n1 mi = 1,609.35 m = 5,280 ft\n1 m = 3.2808 ft", "timestamp": "2026-07-22T05:52:51.083350+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 84, "total_pages": 99, "image_filename": "19930082511_p84.jpg", "text": "82\nNACA TN No. 1826\n\nminus\nequals =\n(a) Representation of lifting element.\n\nminus\nequals\nequals\n(b) Representation of open boundary in\nclosed-open-closed tunnel.\n\nFigure 12.- Acceleration-potential analogies as the\ndifference between two velocity-potential analogies\nslightly shifted relative to each other.", "timestamp": "2026-07-22T05:52:51.323358+00:00"}
{"citation_id": "19930082472", "source_url": "https://ntrs.nasa.gov/api/citations/19930082472/downloads/19930082472.pdf", "page_number": 14, "total_pages": 34, "image_filename": "19930082472_p14.jpg", "text": "12\nNACA TN No. 1797\n\nREFERENCES\n\n1. DeYoung, John: Theoretical Additional Span Loading Characteristics of Wings With Arbitrary Sweep, Aspect Ratio, and Taper Ratio. NACA TN No. 1491, 1947.\n\n2. Jones, Robert T.: Effect of Sweepback on Boundary Layer and Separation. NACA TN No. 1402, 1947.\n\n3. Betz, A.: Applied Airfoil Theory. Unsymmetrical and Non-Steady Types of Motion. Vol. IV of Aerodynamic Theory, div. J, ch. IV, sec. 4, W. F. Durand, ed., Julius Springer (Berlin), 1935, pp. 94 - 107.", "timestamp": "2026-07-22T05:52:51.976429+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 25, "total_pages": 98, "image_filename": "19930086073_p25.jpg", "text": "NACA RM A9H04\n23\n\n[Figure: A model of a wing, body, and vertical tail mounted on a stand in a wind tunnel. The model is oriented vertically. There are markings on the wall behind the model. A logo with \"NACA\" and \"A-12196\" is visible in the upper right corner of the photograph.]\n\n(c) Wing plus body and vertical tail.\n\nFigure 3.— Concluded.", "timestamp": "2026-07-22T05:52:54.304980+00:00"}
{"citation_id": "19930085962", "source_url": "https://ntrs.nasa.gov/api/citations/19930085962/downloads/19930085962.pdf", "page_number": 39, "total_pages": 51, "image_filename": "19930085962_p39.jpg", "text": "38\nCONFIDENTIAL\nNACA RM A9E05\n\nLift coefficient, $C_L$\nElevator deflection, $\\delta_e$, deg\n(h) M, 0.92.\n\nFigure 12. — Continued.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:52:57.763920+00:00"}
{"citation_id": "19930093789", "source_url": "https://ntrs.nasa.gov/api/citations/19930093789/downloads/19930093789.pdf", "page_number": 12, "total_pages": 29, "image_filename": "19930093789_p12.jpg", "text": "NACA RM No. E8I21 CONFIDENTIAL 11\n\nREFERENCES\n\n1. Moore, Charles S., Biermann, Arnold E., and Voss, Fred: The NACA Balanced-Diaphragm Dynamometer-Torque Indicator. NACA RB No. 4C28, 1944.\n\n2. Anon.: Fluid Meters, Their Theory and Application. A.S.M.E. Res. Pub., pub. by Am. Soc. Mech. Eng. (New York), 4th ed., 1937.\n\n3. NACA Subcommittee on Compressors: Standard Procedures for Rating and Testing Multistage Axial-Flow Compressors. NACA TN No. 1138, 1946.\n\n4. Stodola, A.: Steam and Gas Turbines. Vol. I. McGraw-Hill Book Co., Inc., 1927, p. 188. (Reprinted, Peter Smith (New York), 1945.)\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:53:03.961066+00:00"}
{"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 67, "total_pages": 78, "image_filename": "19930082618_p67.jpg", "text": "```markdown\nNACA TN 1945\n\nR\n0 0.7 x 10^6\n□ 1.0\n△ 1.5\n▽ 2.6\n◇ 8.9\nFlagged symbols denote\nstandard roughness\n\nR\n▽ 2.6 x 10^6\n◇ 6.0\n△ 8.9\nFlagged symbols denote\nstandard roughness\n\nSection drag coefficient, c_d\nSection lift coefficient, c_l\n\nR a.c. position\nx/c y/c\n0 0.7 x 10^6 .211 .136\n□ 1.0 .239 .100\n○ 1.5 .259 .062\n△ 2.6 .258 .054\n▽ 2.6 .251 .050\n◇ 6.0 .239 .045\n△ 8.9 .243 .050\n\nNACA\n\nMoment coefficient, c_m\nSection lift coefficient, c_l\n\n(c) Section drag characteristics and section pitching-moment characteristics about the aerodynamic\ncenter of the plain NACA 23015 airfoil section.\n\nFigure 15.— Concluded.\n\n65\n```", "timestamp": "2026-07-22T05:53:04.702417+00:00"}
{"citation_id": "19930093769", "source_url": "https://ntrs.nasa.gov/api/citations/19930093769/downloads/19930093769.pdf", "page_number": 13, "total_pages": 39, "image_filename": "19930093769_p13.jpg", "text": "```markdown\n12\nCONFIDENTIAL\nNACA RM No. E8L10a\n\ntemperature between the blade tip and the root was about $240^\\circ$ F\ngreater for AN-F-58 fuel than for gasoline and the spread between\nthe maximum and minimum temperatures at a given radial location\nwas from $150^\\circ$ to $300^\\circ$ F less for AN-F-58 fuel than for gasoline,\nthe difference being greatest at the blade-root section.\n\n6. During the investigation with AN-F-58 fuel, which involved\nan operating time of 30 hours and 11 minutes, about 240 grams of\nhard carbon were found to have been deposited on the combustor\nbasket near the fuel nozzles.\n\nLewis Flight Propulsion Laboratory,\nNational Advisory Committee for Aeronautics,\nCleveland, Ohio.\n\nREFERENCE\n\n1. Gooding, Richard M., and Hopkins, Ralph L.: The Determination\nof Aromatics in Petroleum Distillates. Paper presented before\nDiv. Petroleum Chem., Am. Chem. Soc. (Chicago, Ill.),\nSept. 9-13, 1946, pp. 131-141.\n\nCONFIDENTIAL\n1070\n```", "timestamp": "2026-07-22T05:53:07.397120+00:00"}
{"citation_id": "19930086076", "source_url": "https://ntrs.nasa.gov/api/citations/19930086076/downloads/19930086076.pdf", "page_number": 25, "total_pages": 50, "image_filename": "19930086076_p25.jpg", "text": "NACA RM E9F09\n23\n\n[Figure: Cutaway view of flame holder 7 after 10 minutes of operation. The image shows two metal components with multiple horizontal slots or vanes. The left component has markings \"NACA 7054 662\" and \"2x3\". The right component has markings \"NACA 7054 662\" and \"2x3\". A scale bar labeled \"INCHES\" is present at the bottom center.]\n\nFigure 7. Cutaway view of flame holder 7 after 10 minutes of operation.\n\nNACA\nC-23391\n5-4-48", "timestamp": "2026-07-22T05:53:08.155188+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 10, "total_pages": 78, "image_filename": "19930082483_p10.jpg", "text": "```markdown\n8\nNACA TN No. 1807\n\nBy use of equations (5) to (10) and the relation $\\left(\\frac{\\mu_i}{\\mu_0}\\right) \\sim \\left(\\frac{T_i}{T_0}\\right)^{0.6}$, which is valid for the range of temperatures normally encountered in turbine operation, the power loss as corrected to standard sea-level conditions may be written as\n\n$$ \\text{power loss} = K_{II} \\left(\\frac{\\rho_d d_h u_h}{\\mu_d}\\right)^{-0.12} \\left(\\frac{N}{1000}\\right)^3 d_h^5 \\rho_d \\frac{1}{\\delta_i^{0.88} \\theta_i^{0.632}} \\quad (15) $$\n\n(b) Bearing loss. It was assumed that the entire frictional loss of the bearings appeared as the observed temperature rise of the lubricating oil.\n\n$$ \\text{bearing power loss} = W_{oil} c_{p,oil} \\Delta T_{oil} \\quad (16) $$\n\nwhere\n\n$c_p$ specific heat at constant pressure, Btu/(lb)($^\\circ$F)\n\n$\\Delta T_{oil}$ temperature rise of lubricating oil in bearings, $^\\circ$F\n\nThis power loss was considered to be constant at any altitude (and degree admission); therefore, no correction factor is used.\n\nTurbine over-all efficiency. - The over-all efficiency of a turbine may be defined as\n\n$$ \\eta' = \\frac{\\text{net power}}{\\text{ideal power}} \\quad (17) $$\n\nwhere $\\eta'$ is the efficiency based on total-pressure ratio.\n\nPartial-Admission Turbine Performance\n\nThe ideal power in the driving fluid of any turbine operating with partial admission may be expressed by\n\n$$ (\\text{ideal power})_F = W_F \\Delta_s h' \\quad (18) $$\n\nwhere the subscript F refers to partial admission.\n\nA basic assumption made in the estimation of turbine performance at various degrees of admission is that the weight flow of\n```", "timestamp": "2026-07-22T05:53:09.438582+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 23, "total_pages": 114, "image_filename": "19930086061_p23.jpg", "text": "NACA RM L9J07\n\nThe longitudinal stability of the three wings increased slightly with aspect ratio between lift coefficients of about 0.15 and 0.4; however, above that lift coefficient the longitudinal stability decreased with increased wing aspect ratio (fig. 55). Wings 2 and 3 had good stability throughout the $C_L$ range including a stable break at stall. Although wing 1 had a stable break at the stall, it had a strong destabilizing shift in the pitching-moment curve at about $C_L = 0.6$, where the rapid loss of outboard lift was noted. The excellent agreement between the longitudinal stability of wing 2 and that for the comparable large-scale wing of reference 3 again illustrates the validity of the low Reynolds number data for configurations having sharp-edged sections.\n\nLift and moment characteristics in yaw.- The effect of yaw on $C_L$, $C_m$, and $C_l$ of the three wings is given in figures 56 to 58. As is the case for conventional wings, the decrease in $C_L$ with yaw was more pronounced as the angle of attack increased. Also the decrease in $C_L$ with yaw was generally greater for all angles of attack as the wing aspect ratio increased. The general effect of yaw on the curves of $C_L$ against $\\alpha$ (figs. 27 to 29) was to decrease the lift-curve slope and make it less linear and to broaden and lower the curve in the region of $C_{L_{max}}$. The trends of decreasing $C_{L_{max}}$ and increasing $C_{l_\\alpha}$ with increased wing aspect ratio noted at zero yaw also generally prevailed in yaw. The effect of angle of yaw on $C_m$ was insignificant for all wings. The variation of $C_l$ with yaw was erratic, but generally at $\\psi = 0^\\circ$ it indicated dihedral effect which varied from positive or approximately zero values at low angles of attack to strong negative values as the angle of attack increased to $\\alpha_{C_{L_{max}}}$. For angles of yaw greater than about $10^\\circ$, the dihedral effect generally increased negatively with yaw for all wings in the angle-of-attack range investigated below stall.\n\nSUMMARY OF RESULTS\n\nThe significant results of the low-speed pressure-distribution and flow investigation of three small-scale low-aspect-ratio pointed wings having 10-percent-thick biconvex sections, $60^\\circ$ sweptback leading edge, and $0^\\circ$, $30^\\circ$, and $-30^\\circ$ trailing-edge sweep may be summarized as follows:\n\n1. At zero yaw each wing had conical separation vortices that emanated in the region of the apex and increased in size and were swept back farther from the leading edge along the span as the angle of attack was increased. Flow observations showed that the center of vortex rotation coincided with the maximum depth of a region of turbulent", "timestamp": "2026-07-22T05:53:09.644218+00:00"}
{"citation_id": "19930085977", "source_url": "https://ntrs.nasa.gov/api/citations/19930085977/downloads/19930085977.pdf", "page_number": 28, "total_pages": 33, "image_filename": "19930085977_p28.jpg", "text": "CONFIDENTIAL\nWing alone\nWing-fuselage\n\nNACA RM L9H22\n\n$\\left(\\frac{\\partial \\epsilon}{\\partial \\alpha}\\right)_M$\n\n| | | | | |\n| :--- | :--- | :--- | :--- | :--- |\n| | | | | .8 |\n| | | | | .4 |\n| | | M = 0.90 | | 0 |\n| | | | | .8 |\n| | | | | .4 |\n| | | M = 0.85 | | 0 |\n| | | | | .8 |\n| | | | | .4 |\n| | | M = 0.80 | | 0 |\n| | | | | .8 |\n| | | | | .4 |\n| | | M = 0.60 | | 0 |\n| -40 | -20 | 0 | 20 | 40 |\n\n$\\left(\\frac{\\partial \\epsilon}{\\partial \\alpha}\\right)_M$\n\n| | | | | |\n| :--- | :--- | :--- | :--- | :--- |\n| | | | | .8 |\n| | | | | .4 |\n| | | M = 1.10 | | 0 |\n| | | | | .8 |\n| | | | | .4 |\n| | | M = 1.05 | | 0 |\n| | | | | .8 |\n| | | | | .4 |\n| | | M = 1.00 | | 0 |\n| | | | | .8 |\n| | | | | .4 |\n| | | M = 0.95 | | 0 |\n| -40 | -20 | 0 | 20 | 40 |\n\nTail height, $h_t$, percent semispan\nCONFIDENTIAL\n\nFigure 12.— Variation of downwash gradient with tail height and Mach number for a model with $0^\\circ$ sweptback wing, aspect ratio 4, taper ratio 0.6, and NACA 65A006 airfoil section.\n\n27", "timestamp": "2026-07-22T05:53:10.345132+00:00"}
{"citation_id": "19930085972", "source_url": "https://ntrs.nasa.gov/api/citations/19930085972/downloads/19930085972.pdf", "page_number": 35, "total_pages": 46, "image_filename": "19930085972_p35.jpg", "text": "NACA RM L9B18\n33\n\nDownwash angle, $\\epsilon$, deg\nAngle of attack, $\\alpha$, deg\n\nTail-off lift-curve slope, $(\\frac{\\partial C_L}{\\partial \\alpha})_o$\nHorizontal-tail effectiveness, $\\frac{\\partial C_{m_t}}{\\partial \\alpha_t}$\nNo cutout\nFaired cutout\n\nNeutral-point and tail-off aerodynamic-center location, $\\eta_p$ and $\\eta_o$, percent $\\bar{c}$ ($\\Lambda=0$)\nLift coefficient, $C_L$\n\nDownwash gradient, $\\frac{\\partial \\epsilon}{\\partial \\alpha}$\nTail-off lift coefficient, $C_{L_o}$\n\nFigure 11.- Longitudinal-stability parameters of a variable-sweep model with and without faired wing cutout. $\\Lambda = 0^\\circ$.", "timestamp": "2026-07-22T05:53:11.648557+00:00"}
{"citation_id": "19930086097", "source_url": "https://ntrs.nasa.gov/api/citations/19930086097/downloads/19930086097.pdf", "page_number": 19, "total_pages": 36, "image_filename": "19930086097_p19.jpg", "text": "NACA RM A9H11 CONFIDENTIAL 17\n\nEXPERIMENTAL RESULTS AND DISCUSSION\n\nDrag Measurements at Zero Lift\n\nThe results of drag measurements at zero lift for wings 1 and 2 at a Mach number of 1.5 are shown in figure 9. These data were taken with the wing surfaces smooth and represent the case of laminar flow in the boundary layer. In accordance with the theoretical expectations, the measurements at this Mach number show that the blunt-trailing-edge airfoil has a significantly lower drag than the double-wedge airfoil of the same section modulus. The drag reduction varies from 15 to 23 percent over the Reynolds number range encountered in the tests. The results of measurements on wings 1 and 2 at a Mach number of 2.0 are shown in figure 10. Also shown in this figure are the results for wings 3 and 4, which were obtained from wing 2 by modifying the base contour. At this Mach number, wing 2 has from 17- to 25-percent lower drag than wing 1. Wing 4 has from 25- to 31-percent lower drag than wing 1. The measured reductions in minimum drag with laminar boundary-layer flow approaching the base, therefore, are in satisfactory agreement with the theoretical considerations both at a Mach number of 1.5 and 2.0.\n\nSome indication of the effect of finite span is given by the data for wing 1. The sum of the theoretical wave drag of this wing as calculated by the shock-expansion method, and the laminar skin-friction drag as calculated from low-speed values, is shown by the dotted lines in figures 9 and 10. These lines representing the theoretical values for two-dimensional flow are several percent higher than the corresponding measured values for the double-wedge profile. The direction of this discrepancy is the same as would occur if the flow separated from the surface downstream of the maximum thickness location. Such separation, which would tend to reduce the profile drag, was clearly shown to exist near the wing tips by the liquid-film technique.\n\nThe experimental values of minimum profile drag for wings 1 and 2 with artificial roughness added are shown in figure 11. These data, which have been corrected for the wave drag due to roughness, are for a Mach number of 2.0 and are representative of the case of turbulent flow approaching the trailing edge. The data for M=1.5 are not presented as they show essentially the same characteristics as the curves in figure 11. It is apparent from this figure that the drag reduction of wing 2 as compared to wing 1 is not as great as for the case of laminar flow approaching the base. This result indicates that on wing 2, which does not have appreciable boattailing, the base drag for turbulent boundary-layer flow is greater than for laminar boundary-layer flow.\n\nAs was discussed earlier, the experimental results for axially symmetric supersonic flow (reference 8) have shown that, with turbulent flow approaching the base, the base drag is greatly reduced by employing a moderate amount of boattailing. In view of this known result for bodies of revolution, the angle of boattailing at the base of wing 2 was\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:53:11.847589+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 58, "total_pages": 96, "image_filename": "19930085880_p58.jpg", "text": "56\nNACA RM No. L9C03\n\nLoad, lb\nSpeed\n(fps)\n30\n25\n20\n15\n10\nWetted area, sq ft\n(b) $\\tau=8^\\circ$.\nNACA\nFigure 18.- Continued.", "timestamp": "2026-07-22T05:53:12.522858+00:00"}
{"citation_id": "19930085938", "source_url": "https://ntrs.nasa.gov/api/citations/19930085938/downloads/19930085938.pdf", "page_number": 33, "total_pages": 42, "image_filename": "19930085938_p33.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:53:14.485573+00:00"}
{"citation_id": "19930085914", "source_url": "https://ntrs.nasa.gov/api/citations/19930085914/downloads/19930085914.pdf", "page_number": 39, "total_pages": 42, "image_filename": "19930085914_p39.jpg", "text": "38\nNACA RM A9D25\n\nPitching-moment coefficient, $C_m$\nfor $M=.20$\n\n<!-- Image (142, 89, 842, 839) -->\n\nAngle of attack, $\\alpha$, deg\n(c) $C_m$ vs $\\alpha$.\n\nFigure 15. - Concluded.", "timestamp": "2026-07-22T05:53:14.834146+00:00"}
{"citation_id": "19930091993", "source_url": "https://ntrs.nasa.gov/api/citations/19930091993/downloads/19930091993.pdf", "page_number": 11, "total_pages": 21, "image_filename": "19930091993_p11.jpg", "text": "```markdown\n# ANALYSIS OF PERFORMANCE OF JET ENGINE FROM CHARACTERISTICS OF COMPONENTS\n\ntemperatures of operation. With heat loss through the wheel as well as to the casing, the volumetric expansion of the gas after passing through the first turbine stage would be less than when the engine was run with cold air. Consequently, the second-stage gas velocities would be decreased and lower wheel speeds would be required for optimum operation of the second stage. This explanation does not account for the local efficiency maximum in the region of an equivalent speed of 85 rps.\n\n<!-- Image (50, 198, 433, 536) -->\n\nFIGURE 9.—Ideal and actual work output of turbine at equivalent turbine speeds between 120 and 180 rps. Solid curve fitted for 185 rps.\n\nBecause of this effect of heat transfer on efficiency, characteristics obtained with hot gas are desirable if turbine-component experiments are considered necessary.\n\nThis shift in efficiency will probably be very much smaller for a single-stage turbine than for a multistage turbine because the volumetric change caused by heat loss primarily affects operation of second or later stages.\n\nIt will be obvious from subsequent analysis that with a turbine, which operates choked or nearly so in the first stage, accurate estimates may be made of compressor operation in the engine from component data even if the turbine efficiencies are in error, provided that accurate mass-flow characteristics of the turbine are available.\n\n## COMBUSTION CHAMBER\n\nThe other significant component of the engine is the combustion chamber. Two variables depending on combustion-chamber performance are needed to determine the over-all engine performance, the pressure ratio $p_{T,3}/p_{T,2}$, and the combustion-chamber efficiency $\\eta$.\n\n**Pressure loss.**—The combustion-chamber pressure loss consists of two components: (1) the friction loss, which is independent of the burning process, and (2) the momentum\n\n<!-- Image (467, 77, 925, 324) -->\n\nFIGURE 10.—Corrected performance of turbine in parameters for engine computations.\n\npressure loss, which depends on the over-all temperature ratio $T_{T,3}/T_{T,2}$ of the combustion chamber. An analysis of these losses based on an idealized form of the combustion chamber is made in reference 8. The data obtained could not be correlated on the basis of this analysis. Aside from possible inaccuracies of the data, another probable cause of the discrepancy is indicated by the fact that some of the data showed lower pressure loss with increased combustion rates for comparable air flows. This anomaly is impossible with the assumption of a tubular combustion chamber of uniform flow area, as is indicated in reference 8. In this engine, the combustion-chamber flow area expanded and the flow diffused between the compressor and the combustion zone and, to some extent, in the combustion zone. The introduction of a resistance, such as a screen at the outlet of a diffuser, increases the efficiency of the diffuser (reference 9). Combustion in the chamber of the engine acts as a resistance to the flow in a manner similar to a screen with a greater pressure drop in the combustion region of the higher local velocities. With such an effect, the data could not be satisfactorily correlated on a simple basis, and for the analysis, the assumption was made that the empirical relation\n\n$$p_{T,2}-p_{T,3}=0.4665 \\frac{W_1^2}{p_{T,2}} \\text{ (lb/sq ft)}$$\n\ncould be used, where $\\rho$ is the gas density in pounds per cubic foot. This constant was obtained from the data and yielded values that show extreme variations of 2.5 percent in absolute pressure at the turbine inlet. For the highest gas flow, the mean deviation of the absolute pressure is only 0.8 percent.\n\n**Combustion efficiency.**—Combustion efficiency was computed from the equation\n\n$$\\eta=\\frac{(H_{T,3}-H_{T,2})+f(H_{T,3}-H_f)}{fh}$$\n\nwhere\n$f$ fuel-air ratio\n$h$ heating value of fuel, (ft-pound)/lb)\n$H_f$ enthalpy of incoming fuel, (ft-pound)/lb)\n\n888267—50—2\n```", "timestamp": "2026-07-22T05:53:21.678985+00:00"}
{"citation_id": "19930086092", "source_url": "https://ntrs.nasa.gov/api/citations/19930086092/downloads/19930086092.pdf", "page_number": 21, "total_pages": 28, "image_filename": "19930086092_p21.jpg", "text": "NACA RM A9F14 CONFIDENTIAL 19\n\n[Figure: Geometric diagram of a swept-back wing-fuselage combination with vertical tail, including dimensions and annotations]\n\nWing  \nSweep $63^\\circ$  \nAspect ratio 3.5  \nTaper ratio .25  \nTwist 0  \nDihedral 0  \nIncidence 0  \nAirfoil section NACA 64A006  \nArea 208.3 sq ft  \n\nFuselage  \nFineness ratio 12.5  \nOrdinate at station, x $1.840(1 - (\\frac{x}{23} - 1)^2)^{3/4}$ ft  \n\nVertical tail  \nSweep $63^\\circ$  \nAspect ratio 1.75  \nTaper ratio .25  \nAirfoil section NACA 64A006  \nArea 35.14 sq ft  \nTail length 20.75 ft  \n\n[NACA logo]\n\nFigure 3. — Geometric characteristics of $63^\\circ$ swept-back wing-fuselage combination with $63^\\circ$ swept-back vertical tail.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:53:22.479453+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 85, "total_pages": 99, "image_filename": "19930082511_p85.jpg", "text": "NACA TN No. 1826\n83\n\n$\\zeta$-plane\n\nA'\nB'\nC'\n\nA\nB\nC\n\n$i\\eta$\n$\\xi_r$\n$\\beta$\n\n$z_1$\n\nz-plane\n$z = e^{\\pi\\zeta}$\n\nC'\nB'\nA'\nA\nB\nC\n\n-1\n$\\frac{1}{2}$\n+1\n\n$\\bar{z}_1$\n\nNACA\n\nFigure 13.- Physical and transformed spaces for two-dimensional closed-open tunnel of unit height.", "timestamp": "2026-07-22T05:53:23.675588+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 26, "total_pages": 98, "image_filename": "19930086073_p26.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:53:24.225445+00:00"}
{"citation_id": "19930082474", "source_url": "https://ntrs.nasa.gov/api/citations/19930082474/downloads/19930082474.pdf", "page_number": 14, "total_pages": 21, "image_filename": "19930082474_p14.jpg", "text": "12\nNACA TN No. 1799\n\nIn order to reduce the tendency of the machine to react to horizontal\ngusts in hovering, the stability with speed could be reduced as by the\nuse of a linkage such that flapping causes corrective feathering.\n\nCONCLUDING REMARKS\n\nFlight investigations of a helicopter have been made to help in\nclarifying the outstanding flying-qualities problems and have lead to the\nfollowing observations:\n\nThe forward-flight instability with angle of attack of the rotor and\nthe fuselage is of greatest concern. The rotor instability is considered\nto arise as a result of flapping and increases in severity with increasing\nspeed. This instability may result in the loss of control in rough air,\nin maneuvers, or during instrument flight. The possibility of allevi-\nating this difficulty by means of a tail surface is briefly discussed.\n\nIn hovering, neither the period of the stick-fixed oscillations nor\nthe lag in the response of the rotor to control application - both of\nwhich have at times been suspected of making hovering difficult for the\nbeginner - was found objectionable. The smaller helicopters, however,\nhave been found to develop high rates of roll per unit stick displacement,\nand this sensitivity results in a tendency for an inexperienced pilot to\novercontrol, particularly during hovering. Reduction in sensitivity by\nchanging the control-system gearing is not feasible because of require-\nments for trim in forward flight. The situation can be alleviated,\nhowever, by increasing the rotor damping, although caution must be used\nto prevent introducing excessive control lag as a result. A further\nmeans for reduction of the control difficulties caused by high sensitivity\nlies in the providing of an appropriate stick-force gradient.\n\nIt is difficult with any whirling rotor system, and particularly\nwith the larger and faster machines, to prevent the occurrence of undesira-\nble control-system forces. In several cases movement of the control stick\nwas found to result in transient forces of an unstable nature or in\nforces out of phase with the direction of stick motion. These phenomena\nwere noted in hovering as well as in forward flight. Such forces were\nfound to increase the difficulty of control greatly and therefore indicate\nthe desirability of irreversible control systems with the desired feel\nintroduced on the pilot's side of the irreversible mechanism. Friction\nhas been used as a cure but in itself has been found very undesirable.\n\nLangley Aeronautical Laboratory\nNational Advisory Committee for Aeronautics\nLangley Air Force Base, Va., November 10, 1948", "timestamp": "2026-07-22T05:53:30.953776+00:00"}
{"citation_id": "19930082472", "source_url": "https://ntrs.nasa.gov/api/citations/19930082472/downloads/19930082472.pdf", "page_number": 15, "total_pages": 34, "image_filename": "19930082472_p15.jpg", "text": "NACA TN No. 1797\n13\n\nTABLE I\n\nORDINATES OF NACA 64A112 a=0.8 (MODIFIED) AIRFOIL SECTION\n[Stations and ordinates given in percent of airfoil chord]\n\n| Upper Surface | | Lower Surface | |\n| :--- | :--- | :--- | :--- |\n| **Station** | **Ordinate** | **Station** | **Ordinate** |\n| 0 | 0 | 0 | 0 |\n| .454 | .988 | .988 | -.932 |\n| .699 | 1.197 | .801 | -1.117 |\n| 1.192 | 1.523 | 1.308 | -1.403 |\n| 2.433 | 2.123 | 2.567 | -1.911 |\n| 4.924 | 2.967 | 5.076 | -2.607 |\n| 7.421 | 3.606 | 7.579 | -3.120 |\n| 9.921 | 4.136 | 10.079 | -3.540 |\n| 14.924 | 4.969 | 15.076 | -4.189 |\n| 19.931 | 5.597 | 20.069 | -4.667 |\n| 24.940 | 6.060 | 25.060 | -5.008 |\n| 29.950 | 6.383 | 30.050 | -5.235 |\n| 34.961 | 6.577 | 35.039 | -5.353 |\n| 39.973 | 6.632 | 40.027 | -5.354 |\n| 44.985 | 6.520 | 45.015 | -5.206 |\n| 44.997 | 6.270 | 50.003 | -4.940 |\n| 55.007 | 5.907 | 54.993 | -4.581 |\n| 60.017 | 5.452 | 59.983 | -4.150 |\n| 65.025 | 4.916 | 64.975 | -3.662 |\n| 70.032 | 4.312 | 69.968 | -3.130 |\n| 75.038 | 3.658 | 74.962 | -2.578 |\n| 80.045 | 2.967 | 79.955 | -2.033 |\n| 85.044 | 2.242 | 84.956 | -1.520 |\n| 90.031 | 1.508 | 89.969 | -1.018 |\n| 95.016 | .767 | 94.984 | -.521 |\n| 100.000 | 0 | 100.000 | 0 |\n\nL.E. radius: 0.994\nSlope of radius through L.E.: 0.0475\n\n[Figure: NACA logo]", "timestamp": "2026-07-22T05:53:38.813161+00:00"}
{"citation_id": "19930086076", "source_url": "https://ntrs.nasa.gov/api/citations/19930086076/downloads/19930086076.pdf", "page_number": 26, "total_pages": 50, "image_filename": "19930086076_p26.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:53:39.151189+00:00"}
{"citation_id": "19930093789", "source_url": "https://ntrs.nasa.gov/api/citations/19930093789/downloads/19930093789.pdf", "page_number": 13, "total_pages": 29, "image_filename": "19930093789_p13.jpg", "text": "Scale:\n$1 \\text{mm} = 20 \\text{ ft/sec}$\n\n$\\frac{U}{V_3} = \\frac{1210}{2840} = .431$\n\n$\\frac{\\Delta V_u}{\\Delta V_{ju}} = \\frac{2350}{2660} = .885$\n\n[Figure: Velocity-vector diagram with vectors $V_2$, $V_3$, $W_2$, $W_3$, $U$, $V_{u,2}$, $V_{x,2}$, angles $\\alpha_2$, $\\alpha_3$, $\\beta_2$, $\\beta_3$, and annotations including $W_{2u}=795$, $V_{x,2}=755$, $40\\%$ Reaction assumed, $\\beta_1/\\beta_3=4.0$, $\\beta_1'/\\beta_3'=?$]\n\nCONFIDENTIAL\n\nEquivalent velocities\nat NACA S.L. condition\n\n| Test condition | $T_1 = 519$ | $T_1 = 660$ |\n| :--- | :--- | :--- |\n| | $U = 645$ | $U = 728$ |\n| | $V_2 = 1160$ | $V_2 = 1308$ |\n| | $V_3 = 578$ | $V_3 = 652$ |\n| | $W_2 = 598$ | $W_2 = 675$ |\n| | $W_3 = 980$ | $W_3 = 1106$ |\n| | | $\\Delta V_u = 1410$ |\n\n$H_1 = 31900$\n$\\Delta H_2 = 171$\n$V_{x2} = 454$\n\nDesign Conditions,\n\n| Velocity (ft/sec) | Angle (deg) |\n| :--- | :--- |\n| U 1210 | $\\alpha_2$ 20.0 |\n| $V_2$ 2177 | $\\alpha_3$ 106.4 |\n| $V_3$ 281/2087 | $\\beta_2$ 41.7 |\n| $W_2$ 1120 | $\\beta_3$ 145.5 |\n| $W_3$ 2081/1840 | |\n\nAll angles are correct.\n\n$h_c = \\frac{V_2^2 - V_3^2}{2g} = 54800$\n$h_p = \\frac{W_3^2 - W_2^2}{2g} = \\frac{33300}{88100}$\n$\\Delta h = \\frac{33300}{88100} = .38$\n\n$\\Delta V_u = 2350$ (scaled)\n$U = 1210$\n$\\frac{\\Delta V_u U}{g} = 88500$\n\nCONFIDENTIAL\n\nNACA RM NO. E8121\n\nFigure 1. - Velocity-vector diagram for mean section.\n\n1031", "timestamp": "2026-07-22T05:53:43.693710+00:00"}
{"citation_id": "19930085962", "source_url": "https://ntrs.nasa.gov/api/citations/19930085962/downloads/19930085962.pdf", "page_number": 40, "total_pages": 51, "image_filename": "19930085962_p40.jpg", "text": "NACA RM A9E05 CONFIDENTIAL 39\n\n$$\n\\begin{array}{c}\n\\text{Lift coefficient, } C_L \\\\\n\\end{array}\n$$\n\n[Figure: Graph showing lift coefficient vs. elevator deflection for various angles of attack (α), with data points marked by symbols and labeled curves for α = -6°, -4°, -2°, 0°, 2°, 4°, 6°, and α_approx.]\n\n$$\n\\begin{array}{c}\n\\text{Elevator deflection, } \\delta_e \\text{, deg} \\\\\n(i) M, 0.94.\n\\end{array}\n$$\n\nFigure 12.— Concluded.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:53:45.523342+00:00"}
{"citation_id": "19930085977", "source_url": "https://ntrs.nasa.gov/api/citations/19930085977/downloads/19930085977.pdf", "page_number": 29, "total_pages": 33, "image_filename": "19930085977_p29.jpg", "text": "```markdown\nM = 0.70\nM = 0.80\nCONFIDENTIAL\nM = 0.85\n\n| | | |\n| :--- | :--- | :--- |\n| $\\frac{q_{wake}}{q}$ | $\\alpha = 10^\\circ$ | $\\alpha = 10^\\circ$ |\n| | | |\n| $\\frac{q_{wake}}{q}$ | $\\alpha = 6^\\circ$ | $\\alpha = 6^\\circ$ |\n| | | |\n| $\\frac{q_{wake}}{q}$ | $\\alpha = 4^\\circ$ | $\\alpha = 4^\\circ$ |\n| | | |\n| $\\frac{q_{wake}}{q}$ | $\\alpha = 0^\\circ$ | $\\alpha = 0$ |\n| | | |\n\n- 80 - 40 0 40 80\n- 80 - 40 0 40 80\n- 80 - 40 0 40 80\n\nTail height, $h_t$, percent semispan\nCONFIDENTIAL\n\n— Wing alone\n--- Wing-fuselage\n\nNACA\n\nFigure 13.— Dynamic-pressure surveys 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.\n\n28\nNACA RM L9E22\n```", "timestamp": "2026-07-22T05:53:46.665326+00:00"}
{"citation_id": "19930085972", "source_url": "https://ntrs.nasa.gov/api/citations/19930085972/downloads/19930085972.pdf", "page_number": 36, "total_pages": 46, "image_filename": "19930085972_p36.jpg", "text": "34\nNACA RM L9B18\n\nDownwash angle, $\\epsilon$, deg\nAngle of attack, $\\alpha$, deg\n\nTail-off lift-curve slope, $\\left(\\frac{dC_L}{d\\alpha}\\right)_o$\nHorizontal-tail effectiveness, $\\frac{dC_{m_t}}{d\\epsilon}$\n\nNeutral-point and tail-off aerodynamic-center location, $\\eta_p$ and $\\eta_o$, percent\n$c'(\\Lambda=0)$\nLift coefficient, $C_L$\n\nDownwash gradient, $\\frac{d\\epsilon}{d\\alpha}$\nTail-off lift coefficient, $C_{L_o}$\n\n(a) Basic tail position.\nFigure 12.- Longitudinal-stability parameters of a variable-sweep model with and without wing cutouts. $\\Lambda = 15^\\circ$.", "timestamp": "2026-07-22T05:53:47.588147+00:00"}
{"citation_id": "19930093769", "source_url": "https://ntrs.nasa.gov/api/citations/19930093769/downloads/19930093769.pdf", "page_number": 14, "total_pages": 39, "image_filename": "19930093769_p14.jpg", "text": "NACA RM No. E6L10a CONFIDENTIAL 13\n\nTABLE I - DESCRIPTION OF FUELS USED\n\n| | AN-F-58 specification | Analysis | | |\n| :--- | :--- | :--- | :--- | :--- |\n| | | AN-F-58 | | Gasoline |\n| | | NACA fuel number | | |\n| | | 48-206 | 48-210 | |\n| A.S.T.M. distillation D 96-46, °F | | | | |\n| Initial boiling point | --------- | 100 | 102 | 106 |\n| Percentage evaporated | | | | |\n| 5 | --------- | 125 | 150 | 124 |\n| 10 | --------- | 145 | 174 | 141 |\n| 20 | --------- | 195 | 234 | 160 |\n| 30 | --------- | 306 | 286 | 176 |\n| 40 | --------- | 355 | 322 | 190 |\n| 50 | --------- | 380 | 360 | 203 |\n| 60 | --------- | 406 | 390 | 211 |\n| 70 | --------- | 430 | 412 | 221 |\n| 80 | --------- | 459 | 444 | 231 |\n| 90 | 425(min.) | 501 | 480 | 250 |\n| Final boiling point | 600(max.) | 568 | 546 | 340 |\n| Residue, (percent) | 1.5(max.) | 1.0 | 0.8 | --------- |\n| Loss, (percent) | 1.5(max.) | 0.2 | 0.2 | --------- |\n| Freezing point, °F | -76(max.) | < -76 | < -76 | --------- |\n| Aromatics, (percent by volume) | | | | |\n| A.S.T.M. D-875-46T | 30(max.) | 23 | 23 | --------- |\n| Silica gel$^a$ | --------- | 23 | 29 | --------- |\n| Viscosity, (centistokes at -40° F) | 10.0(max.) | 3.93 | 4.26 | --------- |\n| (centistokes at 100° F) | --------- | 0.93 | 0.89 | --------- |\n| Bromine number | 14.0(max.) | 4.0 | --------- | --------- |\n| Reid vapor pressure (lb/sq in.) | 5-7(max.) | 5.1 | 5.7 | --------- |\n| Hydrogen-carbon ratio | --------- | 0.161 | 0.153 | 0.183 |\n| Heat of combustion (Btu/lb) | 18,200(min.) | 18,630 | 18,480 | 18,635 |\n| Specific gravity | --------- | 0.786 | 0.794 | 0.706 |\n\n$^a$Reference 1.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:53:47.777814+00:00"}
{"citation_id": "19930091987", "source_url": "https://ntrs.nasa.gov/api/citations/19930091987/downloads/19930091987.pdf", "page_number": 12, "total_pages": 12, "image_filename": "19930091987_p12.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:53:48.572281+00:00"}
{"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 68, "total_pages": 78, "image_filename": "19930082618_p68.jpg", "text": "66\nNACA TN 1945\n\nSection drag coefficient at experimental $c_l$,\n\n.010\n.008\n.006\n.004\n$\\circ$ NACA 64-409\n$\\square$ NACA 641-412\n$\\diamond$ NACA 642-415\n$\\triangle$ NACA 645-418\n\n.008\n.006\n.004\n$\\circ$ NACA 641-012\n$\\square$ NACA 641A212\n$\\diamond$ NACA 641-412\n$\\triangle$ NACA 641-612\n\n.010\n.008\n.006\n.004\n.002\n$\\circ$ NACA 652-415\n$\\square$ NACA 642-415\n$\\diamond$ NACA 652-415\n$\\triangle$ NACA 662-415\n\n.010\n.008\n.006\n.004\n$\\circ$ NACA 0012\n$\\square$ NACA 4412\n$\\diamond$ NACA 4415\n$\\triangle$ NACA 23012\n$\\nabla$ NACA 23015\n\n.5 1.0 2.0 3.0 4.0 5.0 10.0 x $10^6$\nReynolds number, R\n\n(a) Airfoils with smooth surfaces.\n\nFigure 16.— Variation with Reynolds number of section drag coefficient at design lift coefficient for the 15 plain airfoils.", "timestamp": "2026-07-22T05:53:49.130272+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 59, "total_pages": 96, "image_filename": "19930085880_p59.jpg", "text": "NACA RM No. L9C03\n57\n\nLoad, lb\n32\n28\n24\n20\n16\n12\n8\n4\n0\n0 .05 .10 .15 .20 .25 .30 .35\nWetted area, sq ft\n\nSpeed\n(fps)\n30\n25\n20\n15\n10\n\n(c) $\\tau = 12^\\circ$.\nFigure 18.- Continued.\nNACA", "timestamp": "2026-07-22T05:53:49.334580+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 11, "total_pages": 78, "image_filename": "19930082483_p11.jpg", "text": "```markdown\nNACA TN No. 1807\n9\n\ndriving fluid is directly proportional to the number of active\nnozzle passages, that is, to the fraction of active nozzle arc.\n(Data obtained at partial admission bear this out experimentally.)\n\nOf the major turbine losses that occur with full peripheral\nadmission, the rotor-tip leakage loss and the nozzle and rotor-\nblading aerodynamic losses as defined are proportional to the\nnumber of active nozzle passages; whereas the shaft losses are\nindependent of the degree of admission.\n\nThen, under fixed conditions of inlet-gas total temperature,\ninlet total pressure, total-pressure ratio across the turbine\nwheel, and rotative speed of the wheel, the gross power output\nwith no losses must be proportional to the weight flow and hence\nto the degree of admission.\n\n$$(\\text{gross power})_F = F (\\text{gross power})_{360^\\circ} \\quad (19)$$\n\nwhere F is a fraction of the active nozzle arc and the sub-\nscript $360^\\circ$ refers to full admission.\n\nLosses. - A turbine operating with partial admission is sub-\nject to certain inherent losses in addition to the normal losses\nencountered in a full-admission turbine.\n\nAdditional losses encountered with partial admission include:\n\n(1) pumping or windage loss in inactive rotor blading\n\n(2) driving-fluid losses, which include\n\n(a) scavenge and eddy losses during filling and\nemptying of rotor blading\n\n(b) loss due to diffusion of gases at nozzle discharge\n(primarily encountered in reaction-type machines)\n\nFor these studies the scavenge, eddy, and diffusion losses\nhave been combined and considered in the aggregate as driving-\nfluid losses.\n\nCorrelation of data for runs with different degrees of admis-\nsion requires a quantitative statement of these combined losses,\nthe refinement of the correlation being dependent on the accuracy\nand the comprehensiveness of the survey of these quantities.\n```", "timestamp": "2026-07-22T05:53:49.889555+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 24, "total_pages": 114, "image_filename": "19930086061_p24.jpg", "text": "20\nNACA RM L9J07\n\nseparated flow and with a negative pressure peak. Behind the center of\nvortex rotation a negative-pressure dip occurred as the depth of the\nturbulent region diminished rather rapidly.\n\n2. In yaw at moderate and high angles of attack, the vortex increased\nin size and assumed the characteristics of a trailing vortex on the\ntrailing semispan but apparently transformed into a bound vortex on the\nleading semispan.\n\n3. The pressure distributions and flow characteristics of the\nthree wings were similar in nature except that (a) the regions of\nincreased and decreased negative pressure extended farther, in percent\nof chord, at comparable spanwise stations of the wings with higher aspect\nratio because the vortex location was generally about the same absolute\ndistance from the leading edge of each wing at equal angles of attack and\nyaw, (b) the highest negative pressure coefficient decreased with\nincreasing wing aspect ratio, and (c) the area of decreased tip effective-\nness increased with wing aspect ratio.\n\n4. At low angles of attack and zero yaw, the spanwise-load distri-\nbutions agreed fairly well with those predicted by the Weissinger\nlifting-line theory. With increasing angle of attack, however, the\ncenter of semispan loading shifted inboard because of the increasing\nextent of the stalled area at the tip and because of the development\nof a pronounced hump in the spanwise-loading curve just inboard of the\nstalled area. The inboard movement of the semispan center of pressure\nwas generally greater for the wings of higher aspect ratio.\n\n5. Yawing the wings shifted more of the spanwise loading to the\nleading semispan, especially at low angles of attack.\n\n6. The local chordwise center of pressure at zero yaw was\ngenerally at about 35 to 40 percent of the chord at the plane of\nsymmetry. Outward along the span from the plane of symmetry at each\nangle of attack the center of pressure was first closer to or even\nahead of the quarter-chord line where the vortex was on the forward part\nof the chord and then finally farther behind the quarter-chord line near\nthe midspan as the vortex moved to the rear of the wing.\n\n7. The wing lift-curve slopes increased and the values of maximum\nlift coefficient decreased with increased wing aspect ratio.\n\n8. All wings had a stable pitching-moment break at stall, but for\nlift coefficients above 0.4 the longitudinal stability decreased with\naspect ratio, especially for the highest aspect-ratio wing.", "timestamp": "2026-07-22T05:53:51.735977+00:00"}