PDF/ocr_dataset_1076.jsonl

{"citation_id": "19930085983", "source_url": "https://ntrs.nasa.gov/api/citations/19930085983/downloads/19930085983.pdf", "page_number": 2, "total_pages": 46, "image_filename": "19930085983_p2.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:46:54.974752+00:00"}
{"citation_id": "19930085970", "source_url": "https://ntrs.nasa.gov/api/citations/19930085970/downloads/19930085970.pdf", "page_number": 5, "total_pages": 30, "image_filename": "19930085970_p5.jpg", "text": "NACA RM A9E09 CONFIDENTIAL 3\n\n$\\Delta C_L$ increment in lift coefficient $\\left( C_L - C_L C_{D_{min}} \\right)$\n\n$\\frac{\\Delta C_D}{\\Delta C_L^2}$ drag-rise factor\n\n$C_m$ pitching-moment coefficient $\\left[ \\frac{\\text{moment about } (\\bar{c}/4)}{qS\\bar{c}} \\right]$\n\n$(L/D)_{\\text{max}}$ maximum lift-drag ratio\n\nM free-stream Mach number\n\nq free-stream dynamic pressure, pounds per square foot\n\nR Reynolds number, based on mean aerodynamic chord\n\nS wing area, square feet\n\ny lateral coordinate measured from the plane of symmetry, feet\n\n$\\alpha$ angle of attack, degrees\n\n$\\Delta \\alpha$ jet boundary correction to angle of attack, degrees\n\n$\\lambda$ taper ratio $\\left( \\frac{\\text{tip chord}}{\\text{root chord}} \\right)$\n\nAPPARATUS AND TEST METHODS\n\nThe tests were conducted in the Ames 1- by 3-1/2-foot high-speed wind tunnel which is equipped with a flexible nozzle permitting tests at both subsonic and supersonic Mach numbers.\n\nThe model, which was constructed of steel, was of the same basic configuration as that of the tests of reference 2. The wing consisted of NACA 64A-006 sections in the streamwise direction. Major dimensions of the model and the meridian curve of the body are shown in figure 1. The model was supported from the rear of the body by a sting that was shielded from direct air loads. (See fig. 2.)\n\nLift, drag, and pitching moment were measured on a three-component strain-gage balance at angles of attack varied in approximately $1^\\circ$ increments from $-2^\\circ$ to $7^\\circ$ and at Mach numbers from 0.5 to 0.95 and from 1.09 to 1.51. The Reynolds number varied from 0.35 to 0.52 million as shown in figure 3.\n\nSchlieren observations of the flow field about the model were\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:46:55.420648+00:00"}
{"citation_id": "19930085966", "source_url": "https://ntrs.nasa.gov/api/citations/19930085966/downloads/19930085966.pdf", "page_number": 6, "total_pages": 55, "image_filename": "19930085966_p6.jpg", "text": "NACA RM L9B17 CONFIDENTIAL 5\n\nmain fuel was initially heated by the pilot flame before entering the main boiler. It received further heating in the boiler and then issued from the vapor-jet orifices. The small main vapor jets were ignited by the sheet of flame from the pilot burner and then in turn ignited the large main vapor jets, which had greater penetration. The heat for the main boiler was supplied largely by the burning of the large main vapor jets. A common-rail igniter tube interconnected the five pilot housings so that, if the flame faltered in one housing, ignition would be provided by flame traveling from other pilot housings under the back pressure due to combustion. The small blocking of the thin-plate burner was achieved by the manner of mixing the fuel with the air during the combustion process, which allowed a burner of small frontal area to serve a combustion region of relatively large cross section.\n\nDuring preliminary runs the pilot burners as shown in figure 2 did not function properly at the higher combustion-chamber inlet velocities. To correct this shortcoming, the pilots were modified as shown in figure 3: the velocity in the pilot housing was reduced by partly closing off the pilot air inlets with drilled rivets and flaring the pilot skirts. The flared skirts probably also induced a certain amount of turbulence in the region of the pilot-housing trailing edge which may have aided the combustion. The modified configuration operated satisfactorily up to full blower capacity with 3100 pounds of fuel per hour and a combustion-chamber inlet velocity of 145 feet per second. Combustion-chamber inlet velocities up to 300 feet per second were obtained by removing the jet discharge nozzle and decreasing the rate of fuel injection in order to reduce the resistance to flow.\n\nThe proportions of the diffuser tested are shown diagrammatically in the sketch in figure 4. The configuration of the diffuser exit was determined by the combustion-chamber shape. The inlet shape was made similar to the exit shape except for fillets located in the two upper corners. The diffuser had an equivalent angle of expansion of $16^\\circ$ and a ratio of exit to inlet area of 2.14. The diffuser installation is shown in the photograph of figure 5.\n\nThe exhaust-nozzle cross sections were made geometrically similar to that of the combustion-chamber cross section. The nozzle exit area of $1\\frac{1}{2}$ square feet was chosen as a result of preliminary calculations to determine that area which would produce the maximum thrust at a fuel-air ratio of 0.03 with the blower air flow and pressure rises available. The nozzle walls were designed to produce an area variation approaching zero at the exit.\n\nThe air flow was supplied by two 1000-horsepower centrifugal blowers, series connected, which made available 38,000 cubic feet per minute\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:46:57.433082+00:00"}
{"citation_id": "19930085862", "source_url": "https://ntrs.nasa.gov/api/citations/19930085862/downloads/19930085862.pdf", "page_number": 50, "total_pages": 60, "image_filename": "19930085862_p50.jpg", "text": "48\nNACA RM No. L9A07\n\n.01\n0\n$C_n$\n-.01\n-.02\n\nSpoiler span\n$\\diamond$ 0.775b/2\n$\\diamond$ .775b/2\n$\\square$ .40b/2\n\nLocation of outboard end\n0.975b/2\n.975b/2\n.60b/2\n\n.01\n0\n$C_l$\n-.01\n-.02\n-.03\n\n-8 -4 0 4 8 12 16 20 24\n$\\alpha$, deg\n\n[Figure: NACA logo]\n\n(a) $C_l$ and $C_n$ against $\\alpha$.\n\nFigure 18.— Effects of 0.10c projection step spoilers on characteristics of plain wing.", "timestamp": "2026-07-22T06:46:58.614268+00:00"}
{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 51, "total_pages": 72, "image_filename": "19930085491_p51.jpg", "text": "```markdown\n50\n\n$M_o = 1.53$\n$\\Lambda_{LE} = 57.0^\\circ$\n— Linear theory\n$\\circ$ $R = 0.62$ million\n\nCONFIDENTIAL\n\n<!-- Image (138, 333, 419, 756) -->\n\n<!-- Image (501, 333, 828, 756) -->\n\n(a) WF-57\n\nNACA\n\nFigure 7.- Characteristics of swept-back wing and fuselage configurations.\n\nCONFIDENTIAL\n\nWF-57\n\nNACA RM No. A8J04\n```", "timestamp": "2026-07-22T06:46:59.823008+00:00"}
{"citation_id": "19930085951", "source_url": "https://ntrs.nasa.gov/api/citations/19930085951/downloads/19930085951.pdf", "page_number": 13, "total_pages": 92, "image_filename": "19930085951_p13.jpg", "text": "NACA RM L9D29\nCONFIDENTIAL\n11\n\nefficiency of about 6 percent. A small reduction in thickness (from 6.2 to 5 percent at the 0.7 radius) of the blade sections between the shank and the tip had little effect on the maximum propeller efficiency for the conditions of operation tested.\n\nAn examination of the thrust and power coefficients of the propellers operating when the effects of compressibility are present may provide a better understanding of the results. In figure 32 the thrust and power coefficients for maximum efficiency are shown plotted against helical-tip Mach number for the test propellers at a blade angle of $45^\\circ$ at the 0.75 radius. It should be realized that the scarcity of data prevents a definite establishment of the critical Mach numbers, but the curves in figure 32 are shown to illustrate the trends indicated by the data. The curves are somewhat similar to plots of airfoil lift coefficient against Mach number for constant angles of attack and show that increases in thrust and power coefficient occur before the critical Mach number is reached. The critical Mach number is higher for the propellers having the thinner blade sections. After the critical Mach number is reached, there is a marked decrease in both thrust and power coefficients up to a helical-tip Mach number of approximately 1.0. At this helical-tip Mach number near 1.0 the power coefficients begin to increase again, and the thrust coefficients either level off or, in the case of the thinner blades, begin to increase again. These changes in thrust and power coefficients which occur with changes in helical-tip Mach number are, with one exception, less abrupt for the propellers having the thicker blade sections. The single exception is the NACA 10-(3)(08)-045 propeller which has relatively thin shank sections but thick outboard blade sections. The curves in figures 31 and 32 show that the radial distribution of blade-section thickness ratios has a pronounced effect on the characteristics of propellers operating at helical-tip Mach numbers above the critical value.\n\nAny efficiency comparisons of the propellers in the 0.03-solidity group with those in the 0.045-solidity group to study the effects of thickness ratio will include the effects of solidity; however, these effects are small (of the order of 2 percent), and some generalization may be permitted despite the variation in solidity. The NACA propellers discussed in this paper are closely related, and it may be assumed that the radial distribution of camber, solidity, and blade-section thickness is a reasonable optimum for the better propellers. With the generalization in mind, the data for the NACA propellers may serve to indicate the compressibility losses to be expected for propellers having various blade-section thickness ratios.\n\nThe curves in figure 33 show the variation of maximum propeller efficiency with thickness ratio at the 0.7 radius for constant values of helical-tip Mach Number. At a helical-tip Mach number of 0.9 and air-stream Mach number of 0.520 the maximum efficiency of a propeller may be increased approximately 7 percent by reducing the blade-section thickness from 12 to 5 percent at the 0.7 radius. For this Mach number gradient along the blade the rate of change of maximum propeller efficiency with\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:47:03.674921+00:00"}
{"citation_id": "19930085870", "source_url": "https://ntrs.nasa.gov/api/citations/19930085870/downloads/19930085870.pdf", "page_number": 36, "total_pages": 92, "image_filename": "19930085870_p36.jpg", "text": "NACA RM No. L9D07\n37\n\nCONFIDENTIAL\n\nElliptical L.E. $\\circledcirc C_L$ $\\square C_m$\nWedge L.E. $\\triangle C_L$ $\\diamond C_m$\n\n$C_L$\n$C_m$\n\n$\\alpha, deg$\n\nElliptical L.E. $\\circledcirc C_D$ $\\square L/D$\nWedge L.E. $\\triangle C_D$ $\\diamond L/D$\n\n$C_D$\n$L/D$\n\nNACA\n\n(g) Wing 7.w=0.801; R = 940,000.\nFigure 5. - Continued.\nCONFIDENTIAL", "timestamp": "2026-07-22T06:47:05.310062+00:00"}
{"citation_id": "19930085975", "source_url": "https://ntrs.nasa.gov/api/citations/19930085975/downloads/19930085975.pdf", "page_number": 4, "total_pages": 30, "image_filename": "19930085975_p4.jpg", "text": "2\nCONFIDENTIAL\nNACA RM L9E10\n\nrange from 0.40 to 0.91. This paper also includes a comparison with theoretical results computed from Weissinger's theory as presented in reference 2 and from the more approximate, but more convenient, theory of reference 3.\n\nCOEFFICIENTS AND SYMBOLS\n\n| | |\n| :--- | :--- |\n| A | wing aspect ratio |\n| a | speed of sound, feet per second |\n| b | wing span, (3.000 ft on model) |\n| c' | mean aerodynamic chord (M.A.C.; 0.765 ft on model) |\n| $C_l$ | rolling-moment coefficient ($L/qSb$) |\n| $C_{l_p}$ | coefficient of damping in roll $\\left( \\frac{\\partial C_l}{\\partial \\frac{pb}{2V}} \\right)$ |\n| L | rolling moment, ft-lb |\n| M | free-stream Mach number ($V/a$) |\n| p | rate of roll, radians per second |\n| q | dynamic pressure, pounds per square foot ($\\rho V^2/2$) |\n| $\\Lambda_{c/4}$ | sweep angle, degrees (referred to 25 percent chord) |\n| R | Reynolds number ($\\rho V c' / \\mu$) |\n| S | wing area (2.25 sq ft on model) |\n| V | free-stream velocity, feet per second |\n| $\\rho$ | mass density of air, slugs per cubic foot |\n| $\\mu$ | absolute viscosity, pound-seconds per square foot |\n| $\\alpha$ | angle of attack of wing, degrees |\n| $\\delta$ | control-surface deflection with reference to wing chord line parallel to plane of symmetry, degrees |\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:47:19.416063+00:00"}
{"citation_id": "19930085922", "source_url": "https://ntrs.nasa.gov/api/citations/19930085922/downloads/19930085922.pdf", "page_number": 22, "total_pages": 22, "image_filename": "19930085922_p22.jpg", "text": "NACA-Langley - 6-29-55 - 75\n4C\nNACA RM No. L9C23\n\nCoefficient of damping in roll, $C_{l_p}$\n\nUnstable\nStable\nFins off\nFins on\n\nMach number, M\n\nNACA\n\nFigure 12.- Effect of the vertical fins on the coefficient of damping in roll $C_{l_p}$; $\\alpha = 0.30^\\circ$.\n\n21", "timestamp": "2026-07-22T06:47:28.476903+00:00"}
{"citation_id": "19930085913", "source_url": "https://ntrs.nasa.gov/api/citations/19930085913/downloads/19930085913.pdf", "page_number": 17, "total_pages": 34, "image_filename": "19930085913_p17.jpg", "text": "```markdown\nTABLE I.- EXPERIMENTAL DATA - Continued\n\n| Model | Run | $q_r$ (lb/sq ft) | $v_r$ (fps) | Mach number | Distance of weight from root (percent l) | Frequencies (cps) | | | | Phase-angle relationship of bending and torsional stresses. (Ref indicates reference strain-gage trace) | | | | | | | | | | | | | |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| | | | | | | **Natural** | | | **Flutter** | **2nd natural mode** | | | | **3rd natural mode** | | | | **Flutter mode** | | | |\n| | | | | | | 1st | 2nd | 3rd | | 1 | 2 (deg) | 3 (deg) | 4 (deg) | 1 | 2 (deg) | 3 (deg) | 4 (deg) | 1 | 2 (deg) | 3 (deg) | 4 (deg) |\n| Model B<br>Swept untapered wing;<br>$\\Lambda = 45^\\circ$<br>Weight moved along<br>midchord line; $e_w = 0$<br>Reynolds number $\\cong 5117.5v_r$ | 49 | 27.02 | 157.4 | 0.1355 | 0 | 2.50 | 14.89 | 20.18 | 11.11 | Ref | --- | - | 0 | Ref | 0 | 0 | 180 | Ref | 24 | 0 | 24 |\n| | 50 | 26.75 | 156.9 | .1350 | 11.11 | 2.49 | 14.58 | 20.25 | 11.01 | Ref | 0 | - | 0 | Ref | 0 | 0 | 180 | Ref | 0 | 0 | 15 |\n| | 51 | 27.00 | 157.7 | .1355 | 16.66 | 2.48 | 14.08 | 20.00 | 10.80 | Ref | 0 | - | 0 | Ref | 0 | 0 | 180 | Ref | 0 | 0 | 0 |\n| | 52 | 26.66 | 156.6 | .1345 | 27.77 | 2.47 | 10.51 | 20.00 | 9.24 | Ref | 0 | - | 0 | Ref | 0 | 0 | 180 | Ref | 350 | 0 | 343 |\n| | 53 | 28.33 | 161.7 | .1389 | 33.33 | 2.40 | 9.45 | 18.87 | 8.60 | Ref | 0 | - | 0 | Ref | 0 | 0 | --- | Ref | 0 | 0 | 0 |\n| | 54 | 35.99 | 182.6 | .1569 | 38.90 | 2.31 | 8.97 | 18.57 | 8.33 | Ref | 0 | - | 0 | Ref | 0 | 0 | --- | Ref | 0 | 0 | 0 |\n| | 55 | 41.03 | 195.1 | .1675 | 44.40 | 2.20 | 8.87 | 18.00 | 14.71 | Ref | 0 | - | 0 | Ref | 0 | 0 | --- | Ref | 0 | 0 | 0 |\n| | 56 | 41.03 | 195.1 | .1675 | 50.00 | 2.06 | 9.33 | 17.19 | 13.56 | Ref | 0 | - | 0 | Ref | 0 | 0 | --- | Ref | 0 | 0 | 0 |\n| | 57 | 37.68 | 187.1 | .1605 | 55.50 | 1.93 | 9.78 | 16.81 | $^a$13.13 | Ref | --- | - | 0 | Ref | 0 | 0 | --- | Ref | 0 | 0 | --- |\n| | 58 | 48.25 | 212.2 | .1820 | 61.11 | 1.80 | 10.86 | 16.39 | $^a$9.89 | Ref | --- | - | 0 | Ref | 0 | 0 | 180 | Ref | 0 | 0 | --- |\n| | 59 | 43.70 | 201.9 | .1730 | 66.66 | 1.67 | 12.31 | 16.22 | 7.72 | Ref | --- | - | 0 | Ref | 0 | 0 | 180 | Ref | 0 | 0 | --- |\n| | 60 | 40.80 | 194.9 | .1670 | 72.20 | 1.55 | 13.50 | 16.11 | 7.50 | Ref | --- | - | 0 | Ref | 0 | 0 | 180 | Ref | 0 | 0 | --- |\n| | 61 | 37.71 | 187.5 | .1605 | 77.80 | 1.45 | 14.06 | 16.33 | 6.99 | Ref | 0 | 0 | 0 | Ref | 0 | 0 | 180 | Ref | 0 | 0 | --- |\n| | 62 | 35.82 | 182.8 | .1565 | 83.30 | 1.35 | 13.75 | 16.13 | $^a$5.71 | Ref | 0 | 0 | 0 | Ref | 0 | 0 | 180 | Ref | 0 | 0 | --- |\n| | 63 | 31.00 | 170.0 | .1455 | 94.40 | 1.16 | 11.90 | 15.28 | 4.35 | Ref | 180 | 0 | 0 | Ref | 0 | 0 | 180 | Ref | 0 | 0 | 0 |\n\n[Figure: Diagram of wing model showing sweep angle $\\Lambda$, length $l$, chord $c$, and distance $2b$]\n\n$^a$Note oscillograph record, figure 1.\n\nNACA\n\n16\n\nNACA RM L59E24\n```", "timestamp": "2026-07-22T06:47:31.173800+00:00"}
{"citation_id": "19930085862", "source_url": "https://ntrs.nasa.gov/api/citations/19930085862/downloads/19930085862.pdf", "page_number": 51, "total_pages": 60, "image_filename": "19930085862_p51.jpg", "text": "NACA RM No. L9A07\n\n49\n\n$C_m$\n\n$C_D$\n\n$C_L$\n\n$\\alpha, deg$\n\n(b) $C_L$, $C_D$, and $C_m$ against $\\alpha$.\n\nFigure 18.— Concluded.\n\nSpoiler span | Location of outboard end\n---|---\n$\\triangle$ | Spoiler off\n$\\circ$ | 0.775b/2 | 0.975b/2\n$\\diamond$ | .575b/2 | .975b/2\n$\\square$ | .400b/2 | .800b/2\n\n[Figure: Three graphs showing $C_m$, $C_D$, and $C_L$ versus $\\alpha$ with various spoiler configurations indicated by symbols.]", "timestamp": "2026-07-22T06:47:44.391488+00:00"}
{"citation_id": "19930085983", "source_url": "https://ntrs.nasa.gov/api/citations/19930085983/downloads/19930085983.pdf", "page_number": 3, "total_pages": 46, "image_filename": "19930085983_p3.jpg", "text": "NACA RM A9I27 CONFIDENTIAL\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nRESEARCH MEMORANDUM\n\nAERODYNAMIC STUDY OF A WING-FUSELAGE COMBINATION EMPLOYING A WING SWEPT BACK $63^\\circ$.— EFFECTS AT SUBSONIC SPEEDS OF A CONSTANT-CHORD ELEVON ON A WING CAMBERED AND TWISTED FOR A UNIFORM LOAD AT A LIFT COEFFICIENT OF 0.25\n\nBy J. Lloyd Jones and Fred A. Demele\n\nSUMMARY\n\nA cambered and twisted wing having a leading edge swept back $63^\\circ$ and equipped with constant-chord elevons was tested in combination with a slender fuselage to determine the longitudinal and lateral control afforded by the elevons from a Mach number of 0.20 up to a Mach number of 0.93. The tests were performed at a Reynolds number of 2.0 million. Data are presented showing lift, drag, pitching-moment, and rolling-moment characteristics of the model for various elevon deflections, and hinge-moment characteristics of the elevon. Data from the tests have been applied to the calculation of the longitudinal-stability and -control characteristics of a hypothetical airplane geometrically similar to the model.\n\nWith the elevons undeflected, the model was longitudinally unstable about the one-quarter point of the wing mean aerodynamic chord at lift coefficients above about 0.50. The elevons had sufficient pitching-moment and rolling-moment effectiveness for all lift coefficients at which the model was longitudinally stable. At low speeds, the lift coefficient at which static longitudinal instability occurred was decreased by increasing negative elevon deflection. Increasing the Mach number increased the pitching-moment effectiveness at lift coefficients above 0.20, but reduced the rolling-moment effectiveness of the elevons.\n\nINTRODUCTION\n\nA coordinated research program has been undertaken by the Ames Aeronautical Laboratory for an aerodynamic investigation of a wing-fuselage combination employing a wing having the leading edge swept back $63^\\circ$. Aerodynamic characteristics of such a wing with no camber or twist have been presented in references 1, 2, 3, and 4. Reference 1 includes low-speed data on the effectiveness of a constant-chord elevon, and reference 2 reports the Mach number and Reynolds number effects on the effectiveness of the same elevon.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:47:45.116945+00:00"}
{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 52, "total_pages": 72, "image_filename": "19930085491_p52.jpg", "text": "$M_o = 1.53$\n\n$\\Lambda_{LE} = 60.4^\\circ$\n\n— Linear theory\n\n○ $R=0.62$ million\n\nCONFIDENTIAL\n\nPitching-moment coefficient, $C_{m_{\\frac{1}{2}}}$\n\nNACA RM No. A8J04\n\nLift coefficient, $C_L$\n\nAngle of attack, $\\alpha$, deg\n\nDrag coefficient, $C_D$\n\nLift coefficient, $C_L$\n\nLift-drag ratio, $\\frac{L}{D}$\n\nCONFIDENTIAL\n\n(b) WF-60\n\nNACA\n\nWF-60\n\nFigure 7- Continued.\n\n51", "timestamp": "2026-07-22T06:47:51.626870+00:00"}
{"citation_id": "19930085966", "source_url": "https://ntrs.nasa.gov/api/citations/19930085966/downloads/19930085966.pdf", "page_number": 7, "total_pages": 55, "image_filename": "19930085966_p7.jpg", "text": "6 CONFIDENTIAL NACA RM L9B17\n\nat a pressure rise of 150 inches of water. The flow conditions desired at the ram-jet air intake were obtained through use of a fine mesh screen in the low-velocity ducting upstream of the intake, and a boundary-layer bleed gap immediately preceding the intake. A photograph of the ram-jet unit installed in the test cell is presented as figure 6.\n\nThe fuel burned in the ram jet was an unleaded 65-octane gasoline. The fuel was pumped through a 30-gallon surge tank, a filter, rotameters, control valves, and the burners. The fuel pressure in the boilers was of the order of 2 to 5 psi gage. The ignition was operated from a 12-volt power source, and a single-electrode 10-millimeter spark plug was used to produce a spark from the center of each pilot housing to the pilot wall.\n\nThe instrumentation on the ram-jet unit consisted primarily of pressure tubes and thermocouples, layouts of which are shown in figure 7. In addition to the pressure tubes shown there were three rows of wall static-pressure orifices on the diffuser and a single row along the top of the combustion chamber. All the pressure tubes at station 5 were externally water-cooled. A self-balancing potentiometer accurate to $\\pm 1^\\circ$ F was used to read the thermocouple temperatures. All the pressures were made to indicate on a 72-tube manometer board through the use of pneumatically operated pinchboards, and the pressures were photographically recorded. All indicating instruments, controls, and test personnel were housed in a soundproof operating booth.\n\nTESTS\n\nThe main program was divided into two series of tests consisting of constant-fuel-flow and variable-blower-speed runs, and vice versa. Each of the series covered the same variable ranges so that a direct cross check was obtained on the reproducibility of the data. The fuel flow range extended from 0 to 3100 pounds per hour, and the blower speed range extended from maximum rotational speed to the lowest rotational speed at which the nozzle-exit gas temperature did not exceed approximately $2400^\\circ$ F. Higher temperatures than $2400^\\circ$ F produced failures in the setup from overheating. The runs lasted about 2 minutes during which fuel flows, temperatures, blower-drive power frequency, and manometer photographic data were recorded. A running record of certain key pressures obtained through the use of airspeed indicators was maintained to insure uniform test results. The ranges of variables covered in the program are listed in the following table.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:47:52.444301+00:00"}
{"citation_id": "19930085951", "source_url": "https://ntrs.nasa.gov/api/citations/19930085951/downloads/19930085951.pdf", "page_number": 14, "total_pages": 92, "image_filename": "19930085951_p14.jpg", "text": "```markdown\n12\nUNCLASSIFIED\nNACA RM L9D29\n\nblade-section thickness is small for thicknesses up to 12 percent at the\n0.7 radius, and figure 33 indicates that reductions in blade-section\nthickness below 5 percent at the 0.7 radius will probably increase the\nmaximum efficiency very little. However, at a helical-tip Mach number\nof 1.1 and air-stream Mach number of 0.625 the maximum efficiency of a\npropeller may be increased approximately 20 percent by reducing the blade-\nsection thickness from 12 to 5 percent at the 0.7 radius. For this higher\nMach number gradient along the blade the rate of change of maximum pro-\npeller efficiency with blade-section thickness is greater, and for thick-\nnesses between 12 and 8 percent at the 0.7 radius the maximum efficiency\nincreases approximately 3 percent for each decrease in thickness of 1 per-\ncent at the 0.7 radius. For thicknesses between 8 and 5 percent at the\n0.7 radius the rate of increase in propeller efficiency with reductions\nin blade-section thickness is smaller. Figure 33 indicates that further\nreductions in thickness may still improve the maximum efficiency of pro-\npellers operating at helical-tip Mach numbers as high as 1.1, particularly\nfor conditions of operation where the air-stream Mach number is high so\nthat large portions of the blades are subjected to the effects of air\ncompressibility.\n\nCONCLUSIONS\n\nAn investigation of a series of 10-foot-diameter two-blade NACA pro-\npellers differing in blade-section thickness has been completed for a\nrange of blade angles from 20° to 55° at airspeeds up to 500 miles per hour.\nThe results of these investigations have been compared to afford an evalu-\nation of the effects of blade-section thickness ratios on propeller aero-\ndynamic characteristics, and the following conclusions may be drawn:\n\n1. The envelope efficiencies of all the NACA propellers are high at\nthe lower Mach numbers where the adverse effects of compressibility are\nsmall. The higher envelope efficiencies, however, are attained by the\npropellers having the thinner blade sections. The highest efficiencies\n(about 93 percent at a helical-tip Mach number of 0.9 and 84 percent at a\nhelical-tip Mach number of 1.1) reflect the importance of using thin,\nefficient airfoil sections throughout the blade.\n\n2. For propeller operation at constant rotational speed (1140 rpm)\nand power ($C_P = 0.15$) at helical-tip Mach numbers below 0.8,\n\n    (a) A reduction in blade-section thickness from 12 to 8 percent at\n    the 0.7 radius, or approximately one-third all along the radius, results\n    in gains in propeller efficiency up to 10 percent.\n\n    (b) A reduction in blade-section thickness of only the inboard blade\n    sections (from 30 to 13 percent at the 0.3 radius) results in gains in\n    propeller efficiency up to 10 percent.\n\nUNCLASSIFIED\n```", "timestamp": "2026-07-22T06:47:53.111634+00:00"}
{"citation_id": "19930085899", "source_url": "https://ntrs.nasa.gov/api/citations/19930085899/downloads/19930085899.pdf", "page_number": 28, "total_pages": 29, "image_filename": "19930085899_p28.jpg", "text": "NACA RM No. L9A21\n27\n\nDownwash angle, $\\epsilon$, deg\nM = 1.00\n\nDownwash angle, $\\epsilon$, deg\nM = 0.90\n\nDownwash angle, $\\epsilon$, deg\nM = 0.80\n\n$\\alpha$, deg Symbol\n-2 $\\circ$\n-1 $\\triangle$\n0 $\\square$\n1 $\\nabla$\n2 $\\diamond$\n3 $\\blacktriangleleft$\n4 $\\odot$\n6 $\\blacktriangle$\n8 $\\otimes$\n10 $\\blacklozenge$\n\nTail height, $h_t$, percent semispan\n-80 -40 0 40 80\n\nNACA\n\nFigure 12.— Effective downwash angles in region of tail plane for a model with 45° sweptback wing, aspect ratio 4, taper ratio 0.6, and NACA 65A006 airfoil. Wing fuselage.", "timestamp": "2026-07-22T06:47:53.321586+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 76, "total_pages": 78, "image_filename": "19930082483_p76.jpg", "text": ".90\n.70\n.60\n.50\n.40\n.30\n.20\n.10\n0\n10\n20\n30\n40\n50\n60\n70\n80\n90\n100\nPercentage full-power output\n\nOver-all turbine efficiency, $\\eta$\n\nAdmission\n(deg)\n90°\n120°\n150°\n180°\n240°\n300°\n360°\n\nTotal-pressure\nratio,\n$p_i/p_e$\n20\n30\n40\n45\n\nInlet total\npressure, $p_i$\n(in. Hg abs.)\n20\n22\n24\n40\n45\n\n120° admission\nFull admission\n\nPartial-admission power control\nfor turbine operating at design\nconditions\nConstant inlet total pressure, $p_i$\nConstant total-pressure ratio, $p_i/p_e$\n\nFigure 12. - Comparison of inlet pressure, pressure ratio, and active nozzle arc as gas-turbine-power-control parameters. Inlet total temperature, 800° R; corrected rotor speed, 8650 rpm.\n\nNACA\nNACA TN NO. 1807\n74\n1032", "timestamp": "2026-07-22T06:47:54.902711+00:00"}
{"citation_id": "19930085975", "source_url": "https://ntrs.nasa.gov/api/citations/19930085975/downloads/19930085975.pdf", "page_number": 5, "total_pages": 30, "image_filename": "19930085975_p5.jpg", "text": "NACA RM L9E10 CONFIDENTIAL 3\n\n$\\epsilon$ angle of attack of wing-tip chord relative to root chord, radians\n\npb/2V wing-tip helix angle, radians\n\n$C_{l\\delta} = \\frac{\\partial C_l}{\\partial \\delta}$\n\n$\\left(\\frac{pb}{2V}\\right)_\\delta = \\frac{\\partial \\left(\\frac{pb}{2V}\\right)}{\\partial \\delta}$\n\nK correction factor for wing distortion due to bending\n\nSubscripts:\n\n$a_l$ left aileron\n\n$a_r$ right aileron\n\ntest measured values, uncorrected for distortion due to bending\n\nMODEL AND APPARATUS\n\nThe pertinent dimensions of the three wings used in the present investigation are given in figure 1. The wings were constructed of an aluminum alloy. The sweptback wings were designed by shearing the unswept wing; that is, the chordwise elements of the unswept wing parallel to the plane of symmetry were moved rearward until the desired sweep angle of the 25-percent-chord line was obtained. Thus, all wing sections parallel to the plane of symmetry are NACA 65A006 sections. The ailerons were true-contour, sealed-gap, plain flaps of 20 percent chord and 40 percent span.\n\nThe wings were supported by a sting extending forward into the test section from a vertical strut located behind the model. The vertical strut was part of the wind-tunnel balance system and both the strut and a portion of the sting were shielded from the air stream by a fairing. A schematic drawing of the support system and rolling apparatus is shown in figure 2. The angle of attack of the model was changed by varying the angle of incidence of the wing relative to the sting. This was accomplished by utilizing various incidence blocks fitted into the sting. A photograph of the installation is shown in figure 3. The rolling-moment data were obtained from wind-tunnel balance measurements with the sting restrained in roll. When the model was permitted to\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:47:55.281662+00:00"}
{"citation_id": "19930085919", "source_url": "https://ntrs.nasa.gov/api/citations/19930085919/downloads/19930085919.pdf", "page_number": 15, "total_pages": 47, "image_filename": "19930085919_p15.jpg", "text": "The total boundary-induced vertical velocity was then\n\n$$\n\\frac{w}{\\Gamma} = \\frac{1}{4\\pi} \\sum_{n=-\\infty}^{\\infty} \\sum_{m=-\\infty}^{\\infty} (-1)^m \\left\\{ \\frac{2na-y_1-y}{(2na-y_1-y)^2+(mh)^2} \\left[ 1 + \\frac{x}{\\sqrt{(2na-y_1-y)^2+x^2+(mh)^2}} \\right] \\right.\n$$\n\n$$\n\\left. - \\frac{2na+y_1-y}{(2na+y_1-y)^2+(mh)^2} \\left[ 1 + \\frac{x}{\\sqrt{(2na+y_1-y)^2+x^2+(mh)^2}} \\right] + \\frac{x}{x^2+(mh)^2} \\left[ \\frac{2na-y_1-y}{\\sqrt{(2na-y_1-y)^2+x^2+(mh)^2}} \\right. \\right.\n$$\n\n$$\n\\left. \\left. - \\frac{2na+y_1-y}{\\sqrt{(2na+y_1-y)^2+x^2+(mh)^2}} \\right] \\right\\}\n$$\n\nwhere\n\n| | |\n| :--- | :--- |\n| a | 7-foot dimension of wind tunnel |\n| h | 10-foot dimension of wind tunnel |\n| m | number of image patterns in the Z direction |\n| n | number of image patterns in the Y direction |\n| $\\frac{w}{\\Gamma}$ | boundary-induced vertical velocity |\n\nThe remaining symbols in the above equation are defined in figure 20.\n\n14\nCONFIDENTIAL\nCONFIDENTIAL\nNACA RM No. A9C21", "timestamp": "2026-07-22T06:47:55.618328+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 48, "total_pages": 149, "image_filename": "19930083192_p48.jpg", "text": "44\nNACA TN 1976\n\nthe horizontal surfaces can be best estimated as the tail load due to the\ngust alone multiplied by the factor $1 - \\frac{d\\epsilon}{d\\alpha}$. For the canard airplane,\nthe downwash factor would be assumed to be zero.\n\nVertical tail loads.- The suggestion has been made that alleviating\neffects be neglected in the calculation of vertical tail loads due to\nthe action of side gusts. On this basis, the effective gust velocity\nfor the vertical tail would be related to the vertical velocity for the\nwing by the relation\n\n$$U_{e_{\\text{vert. tail}}} = \\frac{U_{e_{\\text{wing}}}}{\\left(\\frac{\\Delta n}{\\Delta n_B}\\right)_{\\text{airplane}}}$$\n\nThe experimental data in table XV show excellent agreement between the\nratio of effective gust velocities and the reciprocal of the accelera-\ntion ratio. The relation given is equivalent to the calculation of the\nvertical tail load for a true gust velocity with no alleviation due to\nunsteady-lift effects or airplane motion. The suggestion appears satis-\nfactory and it would presumably also apply to the vertical tail on the\ncanard airplane.\n\nSteady lift in contrast to unsteady lift.- Consideration of the\nequations of motion for steady and unsteady flow indicates that the use\nof steady-flow lift functions for generalized studies of gust-load\nproblems is not warranted. The ratio of the \"steady-flow\" acceleration\nto the \"unsteady-flow\" acceleration in the same gust is\n\n$$\\frac{\\Delta n_{\\text{steady}}}{\\Delta n_{\\text{unsteady}}} = \\frac{1 - \\frac{1}{\\mu_g} \\int_0^{s_1} \\Delta n(s) ds}{C - \\frac{1}{\\mu_g} \\int_0^{s_1} C_{L_\\alpha}(s_1 - s) \\Delta n(s) ds}$$", "timestamp": "2026-07-22T06:48:06.237304+00:00"}
{"citation_id": "19930085928", "source_url": "https://ntrs.nasa.gov/api/citations/19930085928/downloads/19930085928.pdf", "page_number": 16, "total_pages": 22, "image_filename": "19930085928_p16.jpg", "text": "NACA RM No. A9A31 CONFIDENTIAL 15\n\nMaximum total-pressure ratio, $(H_3/H_0)_{max}$\n\nNormal shock wave\n\nModel of reference 1\nModel A\nModel B\n\nMach number, $M_0$\n\nFigure 3.—Variation of maximum total-pressure ratio with Mach number for models without slots.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:48:14.588571+00:00"}
{"citation_id": "19930085862", "source_url": "https://ntrs.nasa.gov/api/citations/19930085862/downloads/19930085862.pdf", "page_number": 52, "total_pages": 60, "image_filename": "19930085862_p52.jpg", "text": "50\nNACA RM No. L9A07\n\n.01\n$C_n$ 0\n-.01\n\nSpoiler span\n$\\circ$ 0.775b/2\n$\\diamond$ .375b/2\n$\\square$ .40b/2\n\nLocation of outboard end\n0.975b/2\n.975b/2\n.60b/2\n\n.01\n$C_l$ 0\n-.01\n-.02\n\n-8 -4 0 4 8 12 16 20 24\n$\\alpha$, deg\n\n(a) $C_l$ and $C_n$ against $\\alpha$.\n\nFigure 19.- Effects of 0.05c projection step spoilers on characteristics of plain wing.", "timestamp": "2026-07-22T06:48:18.035295+00:00"}
{"citation_id": "19930085970", "source_url": "https://ntrs.nasa.gov/api/citations/19930085970/downloads/19930085970.pdf", "page_number": 6, "total_pages": 30, "image_filename": "19930085970_p6.jpg", "text": "4 CONFIDENTIAL NACA RM A9E09\n\nmade at supersonic Mach numbers. Several representative photographs are presented in figure 4.\n\nREDUCTION OF DATA\n\nAll forces and moments were measured about the wind axes and are presented in the conventional coefficient form. At subsonic Mach numbers the following jet boundary corrections to the angle of attack and drag due to lift, determined by the methods of reference 5, were applied to the data:\n\n$$\n\\Delta \\alpha = 0.398 \\ C_L\n$$\n\n$$\n\\Delta C_D = 0.007 \\ C_L^2\n$$\n\nBlockage corrections were found to be negligible for the model investigated and were not applied to the data.\n\nPossible interference effects between the support system and the model were eliminated by correcting the measured drag for the force resulting from the difference between the pressure measured at the base of the body and the free-stream static pressure. By this means, which was also employed in reference 2, the base drag of the body is subtracted from the total drag of the model. The measured drag values were further corrected for the effects on the body of the static pressure gradients of the free stream.\n\nAlthough zero lift at zero angle of attack was obtained at all subsonic Mach numbers, where the stream inclination is known to be negligible, this was not true at several supersonic Mach numbers. In these instances, the angles of attack were corrected by the amount required to shift the angle of zero lift to the origin. The drag coefficients were correspondingly corrected for the corrections to the angles of attack. The stream-angle correction did not exceed $1^\\circ$.\n\nRESULTS AND DISCUSSION\n\nLift\n\nThe curves of lift coefficient as a function of angle of attack for the model at all test Mach numbers are presented in figure 5. These have been drawn as straight lines although the test points at the lower Mach numbers indicate actual variations that are somewhat nonlinear. At the low test Reynolds numbers the nonlinear characteristics could be caused by large differences in the thicknesses of the\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:48:18.819020+00:00"}
{"citation_id": "19930082918", "source_url": "https://ntrs.nasa.gov/api/citations/19930082918/downloads/19930082918.pdf", "page_number": 58, "total_pages": 62, "image_filename": "19930082918_p58.jpg", "text": "NACA TN 1940\n69\n\n<!-- Image (43, 43, 435, 402) -->\n(a) Unaged: failed in 0.083 hour\nunder a stress of 70,000 psi.\n\n<!-- Image (536, 46, 932, 405) -->\n(b) Unaged: failed in 34 hours\nunder a stress of 55,000 psi.\n\n<!-- Image (43, 479, 435, 839) -->\n(c) Aged 1 hour: failed in 0.22 hour\nunder a stress of 75,000 psi.\n\n<!-- Image (536, 480, 932, 840) -->\n(d) Aged 100 hours: failed in 0.05\nhour under a stress of 80,000 psi.\n\nNACA\nFigure 18.- Effect of aging at 1400° F on fracture characteristics at\n1200° F of low-carbon N-155 alloy solution-treated 10 hours at\n2200° F and water-quenched.", "timestamp": "2026-07-22T06:48:19.062602+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 82, "total_pages": 114, "image_filename": "19930086061_p82.jpg", "text": "```markdown\n78\nNACA RM L9J07\n\n<!-- Image (59, 86, 861, 188) -->\n\nUnsteady flow\nStalled\n\n| $\\alpha$, deg | $C_L$ | $\\alpha$, deg | $C_L$ | $\\alpha$, deg | $C_L$ | $\\alpha$, deg | $C_L$ |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| 8.1 | 0.28 | 8.1 | 0.26 | 8.1 | 0.22 | 8.1 | 0.18 |\n| 14.1 | 0.50 | 14.1 | 0.44 | 14.1 | 0.38 | 14.1 | 0.34 |\n| 29.1 | 0.99 | 29.1 | 0.91 | 29.1 | 0.76 | 29.1 | 0.59 |\n| 39.1 | 1.17 | 39.1 | 1.10 | 39.1 | 0.94 | 39.1 | 0.74 |\n\n<!-- Image (66, 212, 832, 876) -->\n\n(a) $\\psi=0^\\circ$\n(b) $\\psi=10^\\circ$\n(c) $\\psi=20^\\circ$\n(d) $\\psi=35^\\circ$\n\nFigure 29.- Flow characteristics over wing 3 as indicated by surface tufts.\n```", "timestamp": "2026-07-22T06:48:23.680942+00:00"}
{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 53, "total_pages": 72, "image_filename": "19930085491_p53.jpg", "text": "```markdown\n52\n\n$M_0 = 1.53$\n$\\Lambda_{LE} = 63.0^\\circ$\n— Approximate linear theory\n$\\square$ $R=0.31 \\text{ million}$\n$\\circ$ $R=0.62 \\text{ million}$\n$\\nabla$ $R=0.84 \\text{ million}$\n\nCONFIDENTIAL\n\nLift coefficient, $C_L$\nAngle of attack, $\\alpha$, deg\n\nPitching-moment coefficient, $C_{m_{\\frac{1}{4}}}$\n\nDrag coefficient, $C_D$\nLift coefficient, $C_L$\nLift-drag ratio, $\\frac{L}{D}$\n\nCONFIDENTIAL\n\n(c) WF-63\n\nFigure 7- Continued.\n\nNACA\n\nWF-63\nNACA RM No. A8J04\n```", "timestamp": "2026-07-22T06:48:24.362643+00:00"}
{"citation_id": "19930085979", "source_url": "https://ntrs.nasa.gov/api/citations/19930085979/downloads/19930085979.pdf", "page_number": 3, "total_pages": 25, "image_filename": "19930085979_p3.jpg", "text": "2\nNACA RM E9E12\n\nfor introducing hot gas into the inlet. Data were obtained to\ndetermine the effect of plenum-chamber-gas temperature and pressure,\ntunnel-air velocity, and angle of attack on the temperature distri-\nbution at the simulated engine-inlet screen. The icing investi-\ngation to determine the minimum heat requirements was conducted over\na range of liquid-water content from 0.7 to 1.4 grams per cubic\nmeter with an average drop diameter of 15 microns. The tunnel total\ntemperature was $0^\\circ$ F and the model was investigated at angles of\nattack of $0^\\circ$ and $8^\\circ$.\n\nAPPARATUS\n\nThe nacelle investigated was similar to the nacelle described\nin references 1 and 2, but with a short, circular, and straight\nair inlet. The model was two-thirds full scale and was constructed\nof steel, Inconel, and aluminum. The model as installed in the\ntunnel test section is shown in figure 1. The inlet length from\nthe lip to the accessory housing was 14.0 inches. The inlet area\nat the lip (leading edge) was 1.952 square feet and the area at\nthe minimum section was 1.124 square feet. The orifices through\nwhich the hot gas was discharged were located at the minimum section\n3.0 inches from the inlet lip. Hot gas was obtained by passing high-\npressure air through a combustion heater and by ducting the gas to\nthe model. A 1/4-inch-mesh, 0.050-inch-diameter wire screen was\nmounted in the model (fig. 2) to simulate a protective screen instal-\nlation and to provide a means of indicating icing.\n\nINSTRUMENTATION\n\nThe model instrumentation used in the investigation is shown\nin figure 2. Measurements were made of mass flow, ram-pressure\nrecovery, pressure drop across the screen, and temperature distri-\nbution upstream of the inlet screen.\n\nThe average model-air temperature and the temperature distri-\nbution inside the model were measured by means of four thermocouple\nrakes located 33 inches downstream of the orifices and 12 inches\nupstream of the inlet screen, as shown in figure 2. The thermo-\ncouple rakes were spaced at approximately $90^\\circ$ intervals around the\nmodel and each rake consisted of 16 thermocouple probes spaced\n1/4 inch apart.\n\nAft of the accessory-housing tip the model is the same as that\nused in the offset and straight air inlet investigations (references 1\nand 2). A detailed description of the instrumentation is presented\nin reference 1.", "timestamp": "2026-07-22T06:48:28.264638+00:00"}
{"citation_id": "19930085988", "source_url": "https://ntrs.nasa.gov/api/citations/19930085988/downloads/19930085988.pdf", "page_number": 1, "total_pages": 17, "image_filename": "19930085988_p1.jpg", "text": "NACA RM L9H30\n\nCONFIDENTIAL\n\nNACA\n\nRESEARCH MEMORANDUM\n\nFLIGHT INVESTIGATION FROM HIGH SUBSONIC TO SUPERSONIC\n\nSPEEDS TO DETERMINE THE ZERO-LIFT DRAG OF A\n\nTRANSONIC RESEARCH VEHICLE HAVING\n\nWINGS OF 45° SWEEPBACK, ASPECT RATIO 4,\n\nTAPER RATIO 0.6, AND NACA 65A006 AIRFOIL SECTIONS\n\nBy Ellis Katz\n\nLangley Aeronautical Laboratory\n\nLangley Air Force Base, Va.\n\nCLASSIFIED DOCUMENT\n\nThis document contains classified information affecting the National Defense of the United States within the meaning of the Espionage Act, USC 50:31 and 32. The transmission or the revelation of its contents in any manner to an unauthorized person is prohibited by law. Information so classified may be imparted only to persons in the military and naval services of the United States, appropriate civilian officers and employees of the Federal Government who have a legitimate interest therein, and to United States citizens of known loyalty and discretion who of necessity must be informed thereof.\n\nNATIONAL ADVISORY COMMITTEE\n\nFOR AERONAUTICS\n\nWASHINGTON\n\nOctober 27, 1949\n\nCONFIDENTIAL\n\nCopy\n\nRM L9H30\n\nCLASSIFICATION CHANGED TO UNCLASSIFIED\n\nAUTHORITY: NACA RESEARCH ABSTRACT NO. 105\n\nDATE: AUGUST 28, 1956\n\nWHL", "timestamp": "2026-07-22T06:48:31.350290+00:00"}
{"citation_id": "19930085951", "source_url": "https://ntrs.nasa.gov/api/citations/19930085951/downloads/19930085951.pdf", "page_number": 15, "total_pages": 92, "image_filename": "19930085951_p15.jpg", "text": "NACA RM L9D29\nCONFIDENTIAL\n13\n\n(c) A reduction in blade-section thickness of only the outboard blade sections (from 8 to 5 percent at the 0.7 radius) results in gains in efficiency up to 4 percent.\n\n3. For operation at a blade angle of $45^\\circ$ at the 0.75 radius and a helical-tip Mach number of 1.1 the loss in maximum propeller efficiency due to compressibility amounts to 26 percent for the NACA propeller having a blade-section thickness of 12 percent at the 0.7 radius. The corresponding loss in maximum propeller efficiency amounts to only 9 percent for the NACA propeller having a blade-section thickness of 5 percent at the 0.7 radius.\n\n4. At a helical-tip Mach number of 0.9 and air-stream Mach number of 0.520 the rate of change of maximum propeller efficiency with blade-section thickness is small for thicknesses up to 12 percent at the 0.7 radius, and reductions in blade-section thickness below 5 percent at the 0.7 radius will probably increase the maximum efficiency very little.\n\n5. At a helical-tip Mach number of 1.1 and air-stream Mach number of 0.625 the maximum efficiency of a propeller may be increased approximately 20 percent by reducing the blade-section thickness from 12 to 5 percent at the 0.7 radius. For blade-section thicknesses between 12 and 8 percent at the 0.7 radius the maximum efficiency increases approximately 3 percent for each decrease in thickness of 1 percent at the 0.7 radius. For blade-section thicknesses between 8 and 5 percent at the 0.7 radius the rate of increase in propeller efficiency with reductions in blade-section thickness is smaller, but further reductions in thickness may still improve the maximum efficiency of propellers operating at high forward speeds with helical-tip Mach numbers as high as 1.1.\n\nLangley Aeronautical Laboratory\nNational Advisory Committee for Aeronautics\nLangley Air Force Base, Va.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:48:32.645852+00:00"}
{"citation_id": "19930085899", "source_url": "https://ntrs.nasa.gov/api/citations/19930085899/downloads/19930085899.pdf", "page_number": 29, "total_pages": 29, "image_filename": "19930085899_p29.jpg", "text": "```markdown\n$C_{D_{L=0}}$\n.04\n0\n\n$\\left(\\frac{L}{D}\\right)_{\\text{Max.}}$\n20\n10\n0\n\n$\\left(\\frac{\\partial C_L}{\\partial \\alpha}\\right)_M$\n.12\n$C_L = 0$\n$C_L = 0.4$\n.08\n.04\n\n$y_{cp}$\n60\n$C_L = 0.4$\n40\n.6 .7 .8 .9 1.0 1.1 1.2\nMach number, M\n\nWing alone\nWing-fuselage\n\n$\\left(\\frac{\\partial C_m}{\\partial C_L}\\right)_M$\n.2\n$C_L = 0.4$\n0\n$C_L = 0$\n-.2\n\n$\\left(\\frac{\\partial \\epsilon}{\\partial \\alpha}\\right)_M$\n.4\n$h_t$ 0 -30 30\n$C_L = 0$\n0\n\n$\\frac{q_{\\text{wake}}}{q}$\n1.2\n$h_t$ W and WF $C_L = 0$\n$\\pm 30$\n0\n.8\n.6 .7 .8 .9 1.0 1.1 1.2\nMach number, M\n\nNACA\n\nFigure 13.— Summary of aerodynamic characteristics for a model with 45° sweptback wing, aspect ratio 4, taper ratio 0.6, and NACA 65A006 airfoil.\n\nNACA - Langley Field, Va.\nNACA RM No. L9A21\n82\n```", "timestamp": "2026-07-22T06:48:32.969725+00:00"}
{"citation_id": "19930085870", "source_url": "https://ntrs.nasa.gov/api/citations/19930085870/downloads/19930085870.pdf", "page_number": 37, "total_pages": 92, "image_filename": "19930085870_p37.jpg", "text": "38\nNACA RM No. L9D07\n\nCONFIDENTIAL\n\n.24\nElliptical L.E. {C_L} {C_m}\nWedge L.E. {C_L} {C_m}\n.16\n.08\nC_L\n0\n-.08\n-.16\n-.24\n01\nC_m\n0\n-.01\n\n.06\nElliptical L.E. {C_D} {L/D}\nWedge L.E. {C_D} {L/D}\n.04\nC_D\n.02\n0\n-8\n-6\n-4\n-2\n0\n2\n4\n6\n8\n6\n4\n2\n0\nL/D\nNACA\n\n$\\alpha$, deg\n\n(h) Wing 8. w = 0.899. R = 860,000.\nFigure 5. - Continued.\nCONFIDENTIAL", "timestamp": "2026-07-22T06:48:35.044866+00:00"}
{"citation_id": "19930085928", "source_url": "https://ntrs.nasa.gov/api/citations/19930085928/downloads/19930085928.pdf", "page_number": 17, "total_pages": 22, "image_filename": "19930085928_p17.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:48:37.424399+00:00"}
{"citation_id": "19930085975", "source_url": "https://ntrs.nasa.gov/api/citations/19930085975/downloads/19930085975.pdf", "page_number": 6, "total_pages": 30, "image_filename": "19930085975_p6.jpg", "text": "4\nCONFIDENTIAL\nNACA RM L9E10\n\nroll freely under the moment created by the deflected ailerons, the\nrate of roll was recorded electrically.\n\nTESTS AND PROCEDURE\n\nScope\n\nFor each wing, static rolling-moment data and rates of roll were\nobtained through a Mach number range of 0.4 to 0.91 at angles of attack\nof $0.30^\\circ$, $3.45^\\circ$, and $6.50^\\circ$ and for aileron deflections of $0^\\circ$, $\\pm4^\\circ$,\nand $\\pm8^\\circ$ in a plane parallel to the plane of symmetry. The ailerons\nwere deflected oppositely so that the total differential aileron\ndeflections used were $0^\\circ$, $8^\\circ$, and $16^\\circ$.\n\nThe size of the model used in the present investigation resulted\nin an estimated choking Mach number of 0.94, and the data are believed\nto be reliable to a corrected Mach number of about 0.91. The variation\nof test Reynolds number with Mach number for average test conditions is\npresented in figure 4.\n\nCorrections\n\nA small tare correction in the form of bearing friction was deter-\nmined by forced rotation of the rolling apparatus, under both vertical\nand horizontal loads, for the range of angular velocities encountered\nin the tests. This bearing friction has been applied to the results in\nthe form of an increment of damping-in-roll coefficient equal to a\nvalue of $C_{l_p} = -0.005$.\n\nThe rolling moment and Mach numbers have been corrected for blocking\nby the model and its wake by the method of reference 5. The jet-boundary\neffects were estimated and found to be negligible.\n\nThe aluminum-alloy wings were known to bend under load. Accordingly,\nthe effect of wing distortion on the test results was investigated. The\npossible sources of error considered were: (1) deflection of the\nailerons under load; (2) twist of the wing about its elastic axis due\nto the aerodynamic forces being applied at some distance from the\nelastic axis; and, (3) the spanwise change in angle of attack due to\nbending of the wing panel under the span-load distribution. The error\ndue to this last consideration is essentially zero, of course, for an\nunswept wing but increases very markedly as the sweep angle is increased.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:48:38.199082+00:00"}
{"citation_id": "19930082918", "source_url": "https://ntrs.nasa.gov/api/citations/19930082918/downloads/19930082918.pdf", "page_number": 59, "total_pages": 62, "image_filename": "19930082918_p59.jpg", "text": "Page intentionally left blank\n\nPage intentionally left blank", "timestamp": "2026-07-22T06:48:40.745420+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 49, "total_pages": 149, "image_filename": "19930083192_p49.jpg", "text": "NACA TN 1976\n\nwhere C is the unsteady lift coefficient for a gradient gust. For infinite mass parameter, the ratio of accelerations varies from infinity for zero gradient distance (C = 0 at H = 0) to 1.0 for infinite gradient distance (steady flow, C = 1.0). The variation with mass parameter would not be so drastic because it affects only the alleviation term, but, since the integral in the denominator contains an unsteady-lift function, considerable deviation would be expected at small values of mass parameter. Therefore, the trends indicated by steady-flow calculations are not necessarily expected to be those for unsteady-lift calculations.\n\nThe calculated results shown in figure 38(b), in which the unsteady-lift functions for A = 6 and ∞ were utilized, indicate the importance of unsteady-lift-curve shape at large gust-gradient distances. As indicated in the figure, for gradient distance of about 18 chords, the substitution of unsteady-lift functions of aspect ratio 6 for those of aspect ratio ∞ changed the calculated acceleration increment by 50 percent. It is apparent that in the case in which the pitching motion of the airplane is significant and the detailed response of the airplane is to be calculated, minor changes in the unsteady-lift functions themselves can lead to serious discrepancies. The neglect of unsteady-lift relations should involve more radical deviations than the substitution of one unsteady-lift function for another and does not appear to be valid.\n\nGust shape.- Figure 40 shows that the sinusoidal gust can be substituted for the triangular gust with a small relative error. The curves indicate that the relative error is about 2 percent for a range of mass parameter from 5 to 80, with the absolute error varying from 5 to about 8 percent high. The results indicate, therefore, that the trends predicted for the one gust shape should apply equally well to the other gust shape.\n\nCalculated and experimental results.- Inspection of figure 28 indicates that, for the case in which the pitch effects are negligible, the scatter in computing the acceleration ratio appears to be random and the experimental results fall on each side of the calculated curve. Some of the scatter noted is caused by pitching of the model and some is undoubtedly caused by difficulties in making the experimental measurements. On the whole, the calculated results appear to agree in a satisfactory manner with experiment. Similar agreement is indicated in figures 35, 38(a), 38(b), and 32 for H = 0 where pitch is considered negligible. It is believed, therefore, on the basis of these results, that the unsteady-lift functions, the slope of the lift curve, and similar items involved in the calculations are known with an accuracy comparable to that for other elements of gust-load calculations.", "timestamp": "2026-07-22T06:48:42.553797+00:00"}
{"citation_id": "19930085983", "source_url": "https://ntrs.nasa.gov/api/citations/19930085983/downloads/19930085983.pdf", "page_number": 4, "total_pages": 46, "image_filename": "19930085983_p4.jpg", "text": "2\nCONFIDENTIAL\nNACA RM A9I27\n\nCamber and twist have been incorporated in the wing in an effort to\nimprove the flow near the wing tips where, as was evident from early\ninvestigations, loss of lift occurred even at very low angles of attack.\nAerodynamic characteristics of such a wing, cambered and twisted to\nsupport a uniform distribution of lift over its surface at a lift coef-\nficient of 0.25 and a Mach number of 1.5, have been presented in refer-\nences 5 and 6.\n\nThis report presents the results of tests in the Ames 12-foot pres-\nsure wind tunnel of the effectiveness and hinge moments of constant-\nchord elevons at Mach numbers ranging up to 0.93. The elevons extended\nover the outer 50 percent of the span of the cambered and twisted wing,\nwhich is described in reference 6, and had the same plan form as the\nelevons on the model used for the tests reported in references 1 and 2.\n\nNOTATION\n\na speed of sound, feet per second\n\nb wing span measured perpendicular to plane of symmetry, feet\n\nc local chord measured parallel to plane of symmetry, feet\n\n$\\bar{c}$ wing mean aerodynamic chord $\\left( \\frac{\\int_{0}^{b/2} c^2 dy}{\\int_{0}^{b/2} c dy} \\right)$, feet\n\n$C_D$ drag coefficient $\\left( \\frac{\\text{drag}}{qS} \\right)$\n\n$C_h$ hinge-moment coefficient $\\left( \\frac{\\text{hinge moment}}{2q \\times \\text{area moment of elevon about elevon hinge axis}} \\right)$\n\n$C_L$ lift coefficient $\\left( \\frac{\\text{lift}}{qS} \\right)$\n\n$C_l$ rolling-moment coefficient $\\left( \\frac{\\text{rolling moment}}{qSb} \\right)$\n\n$C_{l_p}$ damping-moment coefficient in roll; the rate of change of\nrolling-moment coefficient $C_l$ with wing-tip helix angle\npb/2V, per radian\n\n$C_m$ pitching-moment coefficient about the one-quarter point of the\nwing mean aerodynamic chord $\\left( \\frac{\\text{pitching moment}}{qS\\bar{c}} \\right)$\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:48:44.687486+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 83, "total_pages": 114, "image_filename": "19930086061_p83.jpg", "text": "NACA RM L9J07\n79\n\n[Figure: Diagram (a) showing a top-down view of a triangle with a downward arrow at the top vertex. Two wavy lines originate from the top vertex, flow downwards along the sides of the triangle, and curl outwards at the bottom.]\n\n(a) Three-dimensional top view normal\nto plane of chord lines; $\\alpha=20^\\circ$; $\\psi=0^\\circ$.\n\n[Figure: Diagram (b) showing a rear view of a triangle. Two circular vortex loops are depicted on either side of the centerline, originating from the top vertex and curling downwards and outwards.]\n\n(b) Three-dimensional rear view parallel\nto air stream; $\\alpha=20^\\circ$; $\\psi=0^\\circ$.\n\nNACA\n\nFigure 30.- Typical vortex flow as observed by smoke-flow and\ntuft-probing studies over wing 2.", "timestamp": "2026-07-22T06:48:45.656181+00:00"}
{"citation_id": "19930085966", "source_url": "https://ntrs.nasa.gov/api/citations/19930085966/downloads/19930085966.pdf", "page_number": 8, "total_pages": 55, "image_filename": "19930085966_p8.jpg", "text": "NACA RM L9B17 CONFIDENTIAL 7\n\n| | Fuel flow (lb/hr) | Combustion-chamber inlet velocity (fps) | Fuel Air | Free-stream Mach number |\n|---|---|---|---|---|\n| Minimum | 0 | 40 | 0 | 0.20 |\n| Maximum | 3100 | 195 | .049 | .55 |\n\nThe temperatures and static pressures at the combustion-chamber inlet ranged from $85^\\circ$ F to $125^\\circ$ F and from 2170 to 2600 pounds per square foot absolute, respectively.\n\nCOMPUTATION METHODS\n\nA diagrammatic sketch of the simulated ram-jet configuration is presented in figure 7. Stations 0 and 7 are by definition stations at which the static pressure is equal to the free-stream static pressure; adiabatic flow was assumed between stations 0 and 1 and between stations 5 and 7. In order to calculate the parameters presented in this paper it was necessary to determine almost all quantities identifying the flow at all stations except station 4. The methods used in obtaining these quantities are outlined in the following paragraphs.\n\nThe static-pressure variations at stations 2, 5, and 6 and the total-temperature variations at stations 2 and 6 were so small that arithmetic averages of the data reading could be used. An exact determination of the average total pressure at station 5 would have required knowledge of both total-pressure and temperature variations across the section. Data from preliminary tests in which thermocouple measurements were taken at station 5 were used to obtain an indication of the order of magnitude of the discrepancies between arithmetic and weighted averages of $(Pt_5 - P_0)$. The arithmetic average differed from the weighted by less than 5 percent for all cases. These inaccuracies were not considered of sufficient magnitude to justify the added work of data reduction and complication in instrumentation necessary to measure temperatures, especially in view of the irregularity of the total-pressure gradients, examples of which are presented in figure 8. Therefore, the average total pressure at station 5 was obtained by arithmetic averages of the tube readings. The static pressure at station 1 was determined through the use of a calibration of the three static-tube readings taken at station 1. The calibration was obtained by surveying the static pressure at the station over a range of air flows for several resistances to flow obtained by installing\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:48:47.837071+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 77, "total_pages": 78, "image_filename": "19930082483_p77.jpg", "text": "```markdown\n1032\n\nNACA TN NO. 1807\n\nOver-all turbine efficiency, $\\eta'$\n\nCorrected rotor speed, $N/\\sqrt{\\theta_i}$ (rpm)\n\nAdmission (deg)\n\nTotal-pressure ratio, $p_i/p_e$\n\nPercentage full-power output\n\nPartial-admission power control for turbine operating at design conditions\n\nConstant corrected rotor speed, $N/\\sqrt{\\theta_i}$\n\nConstant total-pressure ratio, $p_i/p_e$\n\nFull admission\n\n120° admission\n\nFigure 13. - Comparison of rotor speed, total-pressure ratio, and active nozzle arc as gas-turbine-power-control parameters. Inlet total temperature, 600° R; inlet total pressure, 45 inches mercury absolute.\n\nNACA\n\n75\n```", "timestamp": "2026-07-22T06:48:48.951580+00:00"}
{"citation_id": "19930085919", "source_url": "https://ntrs.nasa.gov/api/citations/19930085919/downloads/19930085919.pdf", "page_number": 16, "total_pages": 47, "image_filename": "19930085919_p16.jpg", "text": "NACA RM No. A9C21 CONFIDENTIAL 15\n\nThe values of $\\frac{W}{T}$ calculated by the above equation were then used in the basic equation given in reference 6 to obtain the actual wind-tunnel-wall corrections listed in a previous section of this report.\n\nREFERENCES\n\n1. Jones, Robert T.: Estimated Lift-Drag Ratios at Supersonic Speed. NACA TN No. 1350 1947.\n\n2. Madden, Robert T.: Aerodynamic Study of a Wing-Fuselage Combination Employing a Wing Swept Back $63^\\circ$.— Characteristics at a Mach Number of 1.53 Including Effect of Small Variations of Sweep. NACA RM No. A8J04, 1949.\n\n3. Reynolds, Robert M., and Smith, Donald W.: Aerodynamic Study of a Wing-Fuselage Combination Employing a Wing Swept Back $63^\\circ$.— Subsonic Mach and Reynolds Number Effects on the Characteristics of the Wing and on the Effectiveness of an Elevon. NACA RM No. A8D20, 1948.\n\n4. McCormack, Gerald M., and Walling, Walter C.: Aerodynamic Study of a Wing-Fuselage Combination Employing a Wing Swept Back $63^\\circ$.— Investigation of a Large-Scale Model at Low Speed. NACA RM No. A8D02, 1949.\n\n5. Gustafson, F. B., and O'Sullivan, William J., Jr.: The Effect of High Wing Loading on Landing Technique and Distance, with Experimental Data for the B-26 Airplane. NACA ARR No. L4K07, 1945.\n\n6. Swanson, Robert S., and Toll, Thomas A.: Jet-Boundary Corrections for Reflection-Plane Models in Rectangular Wind Tunnels. NACA Rep. No. 770, 1943.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:48:49.745512+00:00"}
{"citation_id": "19930085870", "source_url": "https://ntrs.nasa.gov/api/citations/19930085870/downloads/19930085870.pdf", "page_number": 38, "total_pages": 92, "image_filename": "19930085870_p38.jpg", "text": "```markdown\nNACA RM No. L9D07\n39\n\nCONFIDENTIAL\n\n.24\nElliptical L.E. {O CL, □ Cm}\nWedge L.E. {△ CL, ◇ Cm}\n.16\n.08\nCL\n0\n-.08\n-.16\n-.24\n.01\nCm\n0\n-.01\n\n.06\nElliptical L.E. {O CD, □ L/D}\nWedge L.E. {△ CD, ◇ L/D}\n.04\nCD\n.02\n6\n4\nL/D\n2\n0.8 -6 -4 -2 0 2 4 6 8\nα, deg.\nNACA\n\n(i) Wing 9. w = 0.996; R = 780,000.\nFigure 5. - Continued.\nCONFIDENTIAL\n```", "timestamp": "2026-07-22T06:48:55.570612+00:00"}
{"citation_id": "19930085928", "source_url": "https://ntrs.nasa.gov/api/citations/19930085928/downloads/19930085928.pdf", "page_number": 18, "total_pages": 22, "image_filename": "19930085928_p18.jpg", "text": "NACA RM No. A9A31\n\nCONFIDENTIAL\n\n(a) $M_0$, 1.70; $m_1/m_0$, 0.801; $H_3/H_0$, 0.814.\n\n(b) $M_0$, 1.70; $m_1/m_0$, 0.987; $H_3/H_0$, 0.677.\n\n(c) $M_0$, 2.01; $m_1/m_0$, 0.872; $H_3/H_0$, 0.692.\n\n(d) $M_0$, 2.01; $m_1/m_0$, 1.007; $H_3/H_0$, 0.548.\n\nFigure 4.— Schlieren photographs of the flow about model A without slots.\n\nNACA\nA-13839\n\n17", "timestamp": "2026-07-22T06:48:56.881975+00:00"}
{"citation_id": "19930085990", "source_url": "https://ntrs.nasa.gov/api/citations/19930085990/downloads/19930085990.pdf", "page_number": 1, "total_pages": 132, "image_filename": "19930085990_p1.jpg", "text": "NACA RM A9I01\n\nCONFIDENTIAL\n\nCopy 327\nRM A9I01\n\nNACA\n\nRESEARCH MEMORANDUM\n\nINVESTIGATION OF A THIN WING OF ASPECT RATIO 4 IN THE AMES\n12-FOOT PRESSURE WIND TUNNEL. V - STATIC LONGITUDINAL\nSTABILITY AND CONTROL THROUGHOUT THE SUBSONIC SPEED\nRANGE OF A SEMISPAN MODEL OF A SUPERSONIC AIRPLANE\n\nBy Ben H. Johnson, Jr., and Francis W. Rollins\n\nAmes Aeronautical Laboratory\nMoffett Field, Calif.\n\nCASE FILE\nCOPY\n\nCLASSIFIED DOCUMENT\nThis document contains classified information affecting the National Defense of the United States within the meaning of the Espionage Act, 50 U.S.C. 31 and 32. Its transmission or the revelation of its contents in any manner to an unauthorized person is prohibited by law. Information so classified may be imparted only to persons in the military and naval services of the United States, to appropriate civilian officers and employees of the Federal Government who have a legitimate interest therein, and to United States citizens of known loyalty and discretion who of necessity must be informed thereof.\n\nCLASSIFICATION CHANGED TO UNCLASSIFIED\nAUTHOR: J. W. CROWLEY DATE: 9-15-55\nCHARGE NO. 306\nWHL\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\nWASHINGTON\nDecember 8, 1949\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:49:00.227824+00:00"}
{"citation_id": "19930082918", "source_url": "https://ntrs.nasa.gov/api/citations/19930082918/downloads/19930082918.pdf", "page_number": 60, "total_pages": 62, "image_filename": "19930082918_p60.jpg", "text": "NACA TN 1940\n71\n\n<!-- Image (63, 110, 902, 986) -->\n\nFigure 19.- Effect of aging at 1600° F on 1200° F rupture strength of low-carbon N-155 alloy solution-treated 10 hours at 2200° F and water-quenched.", "timestamp": "2026-07-22T06:49:01.030470+00:00"}
{"citation_id": "19930085911", "source_url": "https://ntrs.nasa.gov/api/citations/19930085911/downloads/19930085911.pdf", "page_number": 25, "total_pages": 52, "image_filename": "19930085911_p25.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:49:07.988896+00:00"}
{"citation_id": "19930085951", "source_url": "https://ntrs.nasa.gov/api/citations/19930085951/downloads/19930085951.pdf", "page_number": 16, "total_pages": 92, "image_filename": "19930085951_p16.jpg", "text": "14\nUNCLASSIFIED\nCONFIDENTIAL\nNACA RM L9D29\n\nREFERENCES\n\n1. Stack, John: Tests of Airfoils Designed to Delay the Compressibility Burble. NACA Rep. No. 763, 1943.\n\n2. Maynard, Julian D., and Salters, Leland B., Jr.: Aerodynamic Characteristics at High Speeds of Related Full-Scale Propellers Having Different Blade-Section Cambers. NACA RM No. L8E06, 1948.\n\n3. Corson, Blake W., Jr., and Maynard, Julian D.: The NACA 2000-Horsepower Propeller Dynamometer and Tests at High Speed of an NACA 10-(3)(08)-03 Two-Blade Propeller. NACA RM No. L7L29, 1948.\n\n4. Davidson, Robert E.: Aerodynamic Characteristics of a Three-Blade Propeller Having NACA 10-(3)(08)-03 Blades. NACA RM No. L8H16, 1948.\n\n5. Evans, Albert J., and Salters, Leland B., Jr.: Aerodynamic Characteristics of a Two-Blade NACA 10-(3)(08)-03R Propeller. NACA RM No. L8E24, 1948.\n\n6. Maynard, Julian D.: Aerodynamic Characteristics at High Speeds of Full-Scale Propellers Having Different Shank Designs. NACA RM No. L6I27a, 1947.\n\n7. Solomon, William: Aerodynamic Characteristics at High Speeds of a Two-Blade NACA 10-(3)(062)-045 Propeller and of a Two-Blade NACA 10-(3)(08)-045 Propeller. NACA RM No. L8E26, 1948.\n\n8. Gray, W. H., and Allis, A. E.: Aerodynamic Characteristics of a Two-Blade NACA 10-(3)(12)-03 Propeller. NACA RM No. L8D01, 1948.\n\n9. Johnson, Peter J.: Aerodynamic Characteristics at High Speeds of Full-Scale Propellers Having Clark Y Blade Sections. NACA RM No. L8E07, 1948.\n\n10. Crigler, John L., and Talkin, Herbert W.: Charts for Determining Propeller Efficiency. NACA ACR No. I4I29, 1944.\n\n11. Lindsey, W. F., Stevenson, D. B., and Daley, Bernard N.: Aerodynamic Characteristics of 24 NACA 16-Series Airfoils at Mach Numbers between 0.3 and 0.8. NACA TN No. 1546, 1948.\n\nUNCLASSIFIED\nCONFIDENTIAL", "timestamp": "2026-07-22T06:49:08.242753+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 84, "total_pages": 114, "image_filename": "19930086061_p84.jpg", "text": "```markdown\n80\n\n-3\n-2\nP -1\n0\nV\n\nPressure distribution\nRegion of turbulent separated flow\nApproximate observed streamline at boundary of separated region\n\n0 .2 .4 .6 .8 1.0\nx/c\n\nNACA\n\n(c) Pressure distribution and cross-sectional view of flow over station 2.\nof wing 2 at $\\alpha=24.1^\\circ$; $\\psi=0^\\circ$.\n\nFigure 30.- Continued.\n\nNACA RM L9J07\n```", "timestamp": "2026-07-22T06:49:08.786704+00:00"}
{"citation_id": "19930085988", "source_url": "https://ntrs.nasa.gov/api/citations/19930085988/downloads/19930085988.pdf", "page_number": 2, "total_pages": 17, "image_filename": "19930085988_p2.jpg", "text": "NACA RM L9H30 CONFIDENTIAL\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nRESEARCH MEMORANDUM\n\nFLIGHT INVESTIGATION FROM HIGH SUBSONIC TO SUPERSONIC\n\nSPEEDS TO DETERMINE THE ZERO-LIFT DRAG OF A\n\nTRANSONIC RESEARCH VEHICLE HAVING\n\nWINGS OF $45^\\circ$ SWEEPBACK, ASPECT RATIO 4,\n\nTAPER RATIO 0.6, AND NACA 65A006 AIRFOIL SECTIONS\n\nBy Ellis Katz\n\nSUMMARY\n\nRocket-powered flight tests were made from high subsonic to supersonic speeds and at high Reynolds numbers to determine the zero-lift drag of a transonic wing-body and body-alone configuration. The test wing was of $45^\\circ$ sweepback, aspect ratio 4, taper ratio 0.6, and NACA 65A006 airfoil section in the free-stream direction. The body had a fineness ratio of 10 and a frontal area equal to 6.06 percent of the wing-plan-form area.\n\nThe test results indicated that at supersonic speeds, the drag coefficient based on total wing area was approximately 0.015 for the body and 0.027 for the body-plus-wing configuration; at subsonic speeds, the drag coefficient was approximately 0.008 for the body and 0.013 for the body-plus-wing configuration. The force-break Mach number was 0.98 for the body and 0.95 for the body-plus-wing configuration. The base contributed very little to the total drag of the test models but indicated a possible interference effect in that the addition of the wing and removal of two stabilizing fins increased the base drag coefficient by 0.002 at a Mach number of 0.95.\n\nINTRODUCTION\n\nAs a part of an NACA program of transonic research, the Langley Pilots Aircraft Research Division is making a series of flight tests at its Wallops Island facility to investigate the aerodynamic characteristics of several rocket-powered wing-body configurations. These tests\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:49:09.082362+00:00"}
{"citation_id": "19930085970", "source_url": "https://ntrs.nasa.gov/api/citations/19930085970/downloads/19930085970.pdf", "page_number": 7, "total_pages": 30, "image_filename": "19930085970_p7.jpg", "text": "NACA RM A9E09 CONFIDENTIAL 5\n\nboundary layers or in the respective extent of the separated flow regions on the upper and lower surfaces of the wing at the trailing edges. In either case the camber of the wing sections would be effectively altered.\n\nIn figure 6, the variation of the mean lift-curve slope with Mach number for the model is compared with the theoretical variation for the wing and with the results of the tests of references 2, 3, and 4. The calculated values were obtained by the methods of references 6 and 7.\n\nAt subsonic Mach numbers it is seen that no lift divergence occurs and that good agreement exists with the type of variation of lift-curve slope with Mach number predicted by the use of lifting-line theory and the extensions of the Prandtl-Glauert rule described in reference 6. The agreement of the present results with those of reference 3, at a much greater Reynolds number, is also good. The increment of lift-curve slope contributed by the body is appreciable, as is indicated in figure 6 by the results of reference 4. This body lift accounts for a major portion of the difference between the present results and the calculated results for the wing alone.\n\nAt supersonic Mach numbers the agreement between the results obtained with the model and the calculated values varies from good to fair with increasing Mach number. It should be noted that the effect of the body has not been considered in the calculations. At 1.5 Mach number, agreement of the present result with that of reference 2¹ is excellent, but both results are somewhat smaller than the value calculated at that Mach number. As was pointed out in reference 2 the lack of agreement with inviscid theory is associated with the extensive laminar separation existing over the aft sections of the upper surface of the wing at moderate angles of attack. This flow separation results in an effective change in the airfoil camber that decreases the lift.\n\nDrag\n\nThe variations of drag coefficient with lift coefficient of the model at the several test Mach numbers are presented in figure 7. In order to facilitate a study and comparison of experimental and calculated drag characteristics, it is convenient to separate the total drag into two components: minimum drag and drag due to lift.\n\n---\n\n¹ The Reynolds number of $0.69 \\times 10^6$ indicated on the figure for the results of reference 2 is based upon the mean aerodynamic chord of the model and corresponds to the value of $0.62 \\times 10^6$ as used in the reference report based upon the mean geometric chord.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:49:12.153254+00:00"}
{"citation_id": "19930085919", "source_url": "https://ntrs.nasa.gov/api/citations/19930085919/downloads/19930085919.pdf", "page_number": 17, "total_pages": 47, "image_filename": "19930085919_p17.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:49:13.808891+00:00"}