PDF/ocr_dataset_1030.jsonl

{"citation_id": "19930085471", "source_url": "https://ntrs.nasa.gov/api/citations/19930085471/downloads/19930085471.pdf", "page_number": 28, "total_pages": 28, "image_filename": "19930085471_p28.jpg", "text": "UNCLASSIFIED\n\nUNCLASSIFIED", "timestamp": "2026-07-22T05:10:45.578869+00:00"}
{"citation_id": "19930082614", "source_url": "https://ntrs.nasa.gov/api/citations/19930082614/downloads/19930082614.pdf", "page_number": 28, "total_pages": 36, "image_filename": "19930082614_p28.jpg", "text": "```markdown\n25\n\nWing loading, lb/sq ft\n30\n50\n\nNet drag coefficient for a vertical dive, $C_{Dn}$\n.5\n.4\n.3\n.2\n.1\n0\n\nTrue airspeed, V = 400 ft/sec\n500\n400\n600\n500\n600\n800\n1000\n800\n1000\n\nVertical dive\n\nAltitude, h, ft\n0\n10\n20\n30\n40 x 1000\n\n[NACA logo]\n\nAngle of dive, deg\n0\n30\n60\n90\n\n.52\n.48\n.44\n.40\n.36\n.32\n.28\n.24\n.20\n.16\n.12\n.08\n.04\n$C_{Dn} = .02$\n\nFigure 4. — Drag coefficient required for no longitudinal acceleration in a dive.\n\nNACA TN 1939\n```", "timestamp": "2026-07-22T05:10:47.049885+00:00"}
{"citation_id": "19930082245", "source_url": "https://ntrs.nasa.gov/api/citations/19930082245/downloads/19930082245.pdf", "page_number": 54, "total_pages": 66, "image_filename": "19930082245_p54.jpg", "text": "```markdown\nNACA TN NO. 1596\n\n1.6\n1.4\n1.2\n1.0\n.8\n.6\n.4\n.2\n0\n-.2\n-.4\n-.6\n.1 .2 .3 .4 .5 .6 .7 .8 .9\nMach number, M\nAileron section normal-force coefficient, $C_{n_a}$\n\n$\\delta_a$ (deg)\n30\n18\n12\n6\n4\n2\n0\n-2\n-4\n-6\n-12\n\n18\n12\n4\n2\n-2\n-4\n-6\n-12\n\n.16\n.12\n.08\n.04\n0\n-.04\n-.08\n-.12\n-.16\n-.20\n-.24\n-.28\n.1 .2 .3 .4 .5 .6 .7 .8 .9\nMach number, M\nAileron section hinge-moment coefficient, $C_h$\n\n$\\delta_a$ (deg)\n-12\n-6\n-4\n-2\n0\n2\n4\n6\n12\n18\n30\n\n-6\n-4\n-2\n0\n2\n4\n6\n-12\n-6\n-4\n-2\n0\n2\n4\n6\n-12\n-6\n-4\n-2\n0\n2\n4\n6\n\n(d) $C_h = 0.2$.\nFigure 10.—Continued.\n\n[Figure: NACA logo]\n\n53\n```", "timestamp": "2026-07-22T05:10:47.991890+00:00"}
{"citation_id": "19930082496", "source_url": "https://ntrs.nasa.gov/api/citations/19930082496/downloads/19930082496.pdf", "page_number": 40, "total_pages": 50, "image_filename": "19930082496_p40.jpg", "text": "NACA TN No. 1836\n39\n\n[Figure: Photograph of a fractured ceramic blade assembly. The central blade is broken, showing a rough fracture surface. Other blades are visible in the background. A NACA logo with the text \"C-20115 12-3-47\" is in the bottom right corner of the image.]\n\nFigure 10. - Typical fracture of ceramic blade experienced during phase 1 of quasi-service evaluation.", "timestamp": "2026-07-22T05:10:50.938412+00:00"}
{"citation_id": "19930082914", "source_url": "https://ntrs.nasa.gov/api/citations/19930082914/downloads/19930082914.pdf", "page_number": 29, "total_pages": 66, "image_filename": "19930082914_p29.jpg", "text": "28\nNACA TN No. 1857\n\n3. The turbulence in the mixing region was less than that for incompressible jets.\n\n4. The rates of spread of the mixing region, both into the jet and into the ambient air, were less than those of incompressible jets.\n\nLangley Aeronautical Laboratory\nNational Advisory Committee for Aeronautics\nLangley Air Force Base, Va., January 21, 1949", "timestamp": "2026-07-22T05:10:52.559573+00:00"}
{"citation_id": "19930085544", "source_url": "https://ntrs.nasa.gov/api/citations/19930085544/downloads/19930085544.pdf", "page_number": 12, "total_pages": 33, "image_filename": "19930085544_p12.jpg", "text": "NACA RM No. L8K26\n\nFor the case being considered it is assumed that $\\alpha_{W_{ut}} = \\alpha_{u_P} = \\omega$. Therefore $k_{W_{ut}} = k_{\\alpha_P} = k$ and $k_{W_{ut}} + k_{\\alpha_P} = k_1 = 2k$. Accordingly the functions $C(k_{W_{ut}})$ and $C(k_{\\alpha_P})$ are equal and are denoted by $C(k) = F + iG$ while the function $C(k_{W_{ut}} + k_{\\alpha_P})$ is denoted by $F_1 + iG_1$.\n\nIt should be noted that some of the real terms in equation (8) are not multiplied by the $C(k)$ functions. These real terms can be interpreted as giving the total force on the airfoil and the imaginary terms the force due to perturbation velocities only.\n\nTaking the real part of equation (8), the resulting total force coefficient is given by\n\n$$\n\\begin{aligned}\nL_c &= \\frac{L}{\\frac{\\rho W_o^2}{2} c} = \\frac{\\alpha}{\\alpha_{P_o}} + \\left[ F\\left(1 + \\epsilon \\frac{\\alpha}{\\alpha_{P_o}}\\right) - \\frac{k}{2} G + \\epsilon \\frac{\\alpha}{\\alpha_{P_o}} \\right] \\sin \\omega t \\\\\n&\\quad + \\epsilon \\left[ k\\left(1 + \\frac{F}{2}\\right) + G\\left(1 + \\epsilon \\frac{\\alpha}{\\alpha_{P_o}}\\right) + G_1 \\right] \\sin 2\\omega t - \\epsilon^2 F_1 \\sin 3\\omega t \\\\\n&\\quad + \\left[ \\frac{k}{2}\\left(1 + F + \\epsilon \\frac{\\alpha}{\\alpha_{P_o}}\\right) + G\\left(1 + \\epsilon \\frac{\\alpha}{\\alpha_{P_o}}\\right) \\right] \\cos \\omega t \\\\\n&\\quad - \\epsilon \\left[ F\\left(1 + \\epsilon \\frac{\\alpha}{\\alpha_{P_o}}\\right) - \\frac{k}{2} G + F_1 \\right] \\cos 2\\omega t - \\epsilon^2 G_1 \\cos 3\\omega t\n\\end{aligned}\n\\tag{9}\n$$\n\nThe lift-force coefficient $L_c$ for propeller blade computations applies to only one blade element. If the curves from all blades are plotted and integrated and the ordinates summed, the curves thus obtained may be used to determine the turning moment on the propeller shaft as a function of time. The turning moment on the shaft (yawing moment for pitched propeller) for any position is found by plotting the lift force at each blade element times its moment arm and integrating graphically.\n\nCALCULATION OF FORCES AND DISCUSSION OF RESULTS\n\nSteady state.- The calculations were made for a 4-foot-diameter propeller having an NACA 4-(3.9)(07)-0345-B blade design of NACA 16-series", "timestamp": "2026-07-22T05:10:53.083697+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 43, "total_pages": 99, "image_filename": "19930082511_p43.jpg", "text": "NACA TN No. 1826\n\n$\\rho$ density of fluid  \n$L$ lift of lifting element  \n$w$ tunnel-induced velocity normal to the $\\xi\\eta$-plane  \n$V$ free-stream velocity  \n$R$ tunnel radius  \n\nANALYSIS\n\nIntroduction\n\nIn the analysis of the three-dimensional, circular, closed-open-closed tunnel, an appreciable simplification results when the tunnel axis lies in the plane of the horseshoe vortex. For off-center locations of the horseshoe vortex, or for a source-sink body on the axis, or for the general unsymmetrical disturbance, certain complications arise that are related to the fact that the pressure on the free boundary is then not equal to the pressure at $\\pm \\infty$ in the closed parts of the tunnel. That is, for these cases, if the net perturbation velocity is zero far upstream and downstream in the closed parts of the tunnel, a constant longitudinal perturbation velocity $u \\neq 0$ will exist on the free surface. (See parts I and II.) A similar complication results for a source in a completely closed tunnel.\n\nThe analysis described in the following section is applicable directly to the case in which the tunnel axis lies in the plane of the horseshoe vortex and for which the longitudinal perturbation velocity on the free surface is zero. (See part I.) In the succeeding section are derived the additional terms needed for the solution of the more general problem. The significance of the titles of these two sections will become clear in the analysis.\n\nCylindrically Symmetric Term Omitted\n\nBoundary conditions and formal expression for $\\phi_A$.— The solution is developed in cylindrical coordinates $(\\xi,\\rho,\\theta)$ where the $\\xi$-axis coincides with the tunnel axis and $\\theta$ is measured from the horizontal plane. The relations of these coordinates to the rectangular coordinates $(\\xi,\\eta,\\zeta)$ are indicated in figure 23. The distance variables $\\xi$, $\\eta$, $\\zeta$, and $\\rho$ are considered in units of the tunnel radius.\n\nLet $\\phi_0(\\xi,\\rho,\\theta)$ be the disturbance velocity potential associated with the lifting body in unlimited space (in particular, the velocity", "timestamp": "2026-07-22T05:10:57.968363+00:00"}
{"citation_id": "19930085542", "source_url": "https://ntrs.nasa.gov/api/citations/19930085542/downloads/19930085542.pdf", "page_number": 15, "total_pages": 46, "image_filename": "19930085542_p15.jpg", "text": "NACA RM No. L8I29\n13\n\nREFERENCES\n\n1. Jones, Robert T.: Properties of Low-Aspect-Ratio Pointed Wings at Speeds below and above the Speed of Sound. NACA Rep. No. 835, 1946.\n\n2. Ribner, Herbert S.: The Stability Derivatives of Low-Aspect-Ratio Triangular Wings at Subsonic and Supersonic Speeds. NACA TN No. 1423, 1947.\n\n3. DeYoung, John: Theoretical Additional Span Loading Characteristics of Wings with Arbitrary Sweep, Aspect Ratio, and Taper Ratio. NACA TN No. 1491, 1947.\n\n4. Bird, John D.: Some Theoretical Low-Speed Span Loading Characteristics of Swept Wings in Roll and Sideslip. NACA TN No. 1839, 1949.\n\n5. Toll, Thomas A., and Queijo, M. J.: Approximate Relations and Charts for Low-Speed Stability Derivatives of Swept Wings. NACA TN No. 1581, 1948.\n\n6. MacLachlan, Robert, and Letko, William: Correlation of Two Experimental Methods of Determining the Rolling Characteristics of Unswept Wings. NACA TN No. 1309, 1947.\n\n7. Letko, William, and Jaquet, Byron M.: Effect of Airfoil Profile of Symmetrical Sections on the Low-Speed Static-Stability and Yawing Derivatives of 45° Sweptback Wing Models of Aspect Ratio 2.61. NACA RM No. L8H10, 1948.\n\n8. Tosti, Louis P.: Low-Speed Static Stability and Damping-in-Roll Characteristics of Some Swept and Unswept Low-Aspect-Ratio Wings. NACA TN No. 1468, 1947.\n\n9. Anderson, Adrien E.: An Investigation at Low Speed of a Large-Scale Triangular Wing of Aspect Ratio Two.- II. The Effect of Airfoil Section Modifications and the Determination of the Wake Downwash. NACA RM No. A7H28, 1947.", "timestamp": "2026-07-22T05:10:58.169920+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 4, "total_pages": 96, "image_filename": "19930085880_p4.jpg", "text": "2\nNACA RM No. L9C03\n\nMODELS AND APPARATUS\n\nThe principal details of the models tested are given in figures 1 to 3. Model 250A had a flat bottom surface and a rectangular plan form. Model 250B had a rectangular plan form but the bottom was curved in cross section. Model 250D had a flat bottom and a triangular plan form; it was tested with the base of the triangle forward (leading edge). All models had the same plan-form area (0.347 sq ft) and were made of solid mahogany. The upper surface was arbitrarily faired by making all the longitudinal sections circular arcs with a height at the center of 5 percent of the chord which forms the bottom of the section.\n\nThe tests were made on the small model towing gear in Langley tank no. 2. The test setup is shown in figure 4.\n\nTwo lenticular struts supported the models from a trimming moment dynamometer which was fastened to the towing staff. A phosphor bronze strap in the dynamometer restrained the model in trim. Electrical strain gages fastened to this strap indicated the trimming moment encountered. The towing staff was free only in rise and the vertical load was varied by counterbalancing. Changes in draft were read by means of a disc and pointer arrangement which mechanically magnified changes in the vertical position of the staff. The guides for the staff were connected to the resistance dynamometer. This dynamometer consisted of a cantilever spring, the deflections of which were magnified by an optical system.\n\nPROCEDURE\n\nGeneral\n\nThe tests consisted of towing the models at various speeds and loads, at fixed trims of 4°, 8°, 12°, 16°, and 20°. A sufficient number of loads were chosen at each trim to define the variations of resistance, trimming moment, and draft with wetted length. The maximum speed was determined by the measuring limits of the equipment and ranged from 30 to 35 feet per second. The minimum speed was set at 10 feet per second since indications were that below this speed satisfactory planing data could not be obtained with these models. Resistance, trimming moment, draft, and wetted length were read. Draft is the depth of the trailing edge of the model below the free-water surface. Trimming moment was measured about an arbitrary point above the model and from the measured results the trimming moment about the trailing edge at the center line of the model was calculated.\n\nWetted Length\n\nThe wetted length read was the distance from the trailing edge of the model to the intersection of the dynamic solid water boundary with", "timestamp": "2026-07-22T05:10:58.485053+00:00"}
{"citation_id": "19930082498", "source_url": "https://ntrs.nasa.gov/api/citations/19930082498/downloads/19930082498.pdf", "page_number": 37, "total_pages": 49, "image_filename": "19930082498_p37.jpg", "text": "Page intentionally left blank\n\nPage intentionally left blank", "timestamp": "2026-07-22T05:11:02.139419+00:00"}
{"citation_id": "19930085881", "source_url": "https://ntrs.nasa.gov/api/citations/19930085881/downloads/19930085881.pdf", "page_number": 2, "total_pages": 31, "image_filename": "19930085881_p2.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:11:06.423381+00:00"}
{"citation_id": "19930085869", "source_url": "https://ntrs.nasa.gov/api/citations/19930085869/downloads/19930085869.pdf", "page_number": 8, "total_pages": 36, "image_filename": "19930085869_p8.jpg", "text": "6\nCONFIDENTIAL\nNACA RM L9D15\n\nTypical photographs of the model during these take-off runs are presented in figure 10. The deep draft and small bow clearance at low speeds are shown in figure 10(a). The high trims beyond hump speed, shown in figure 10(a), enables the low bow to be well clear of the water, but keeps the tail boom in the forebody spray. At high speeds and low trims, figure 10(c) shows that only the point of the step is in the water and the bow has ample clearance.\n\nThe range of elevator deflection available for take-offs is plotted against location of the center of gravity in figure 11. A large range of elevator deflection was available at all center-of-gravity locations. Neither lower-limit nor upper-limit porpoising determined these elevator limits. Maximum up-elevator deflection resulted in no upper-limit porpoising greater than $2^\\circ$ amplitude. It is apparent from figure 9 that $2^\\circ$ minimum trim at high speed will be reached before $2^\\circ$ amplitude of lower-limit porpoising. The minimum elevator deflection for take-off was therefore determined by the minimum trim of $2^\\circ$ rather than by the lower-limit porpoising.\n\nLanding Stability\n\nThe amplitudes of the maximum oscillations of trim during landing are plotted against contact trims in figure 12(a). In figure 12(b) the amplitudes of the maximum vertical motions, at the center of gravity, are plotted against contact trims. From these plots it is seen that there was little change in rise or trim during any landing and all landings were considered stable. The model trimmed down at contact, since the center of gravity was located well forward of the step point. This contact rotation is plotted against contact trim in figure 13 for the center-of-gravity locations tested.\n\nResistance\n\nIn figure 14, resistance coefficient, load-resistance ratio, load coefficient, and trim are plotted against speed coefficient. The hump load-resistance ratio of 3.1 at a speed coefficient of 3.25 is lower than that obtained with well-designed conventional hulls. The power in the model would not be sufficient for take-off, but in a high-performance airplane considerably more thrust would be available.\n\nDirectional Stability\n\nThe model was attached to a tubular staff which was slightly flexible torsionally and a tendency to yaw was noticed over a range of speed coefficient from 2.9 to 4.2. Apparently the roach, which impinged on the boom throughout this range, caused the same instability as found in reference 2 on an unswept single-boom hull. The results of unpublished tests with the model in a free, self-propelled condition indicate\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:11:06.798849+00:00"}
{"citation_id": "19930085487", "source_url": "https://ntrs.nasa.gov/api/citations/19930085487/downloads/19930085487.pdf", "page_number": 28, "total_pages": 36, "image_filename": "19930085487_p28.jpg", "text": "```markdown\n26\nNACA RM No. EB122\n\n100,000\n\n10,000\n\nFrequency, cps\n\nRear stator row, 38 blades\n\nFront stator row, 22 blades\n\nFirst torsional mode\nSecond bending mode\n\n1,000\n\nFourth order, front bearing support\n\nFirst bending mode\n\nSecond order, split compressor case\n\nFirst order, rotor speed\n\n100\n4 6 8 10 12 14 16x10³\nRotor speed, rpm\n\n(b) Second stage.\n\nFigure 9. - Continued. Critical-speed diagrams for 10 stages of\ncompressor rotor.\n\nNACA\n```", "timestamp": "2026-07-22T05:11:07.497910+00:00"}
{"citation_id": "19930082496", "source_url": "https://ntrs.nasa.gov/api/citations/19930082496/downloads/19930082496.pdf", "page_number": 41, "total_pages": 50, "image_filename": "19930082496_p41.jpg", "text": "40\n\nPage intentionally left blank\n\nPage intentionally left blank", "timestamp": "2026-07-22T05:11:12.810086+00:00"}
{"citation_id": "19930085626", "source_url": "https://ntrs.nasa.gov/api/citations/19930085626/downloads/19930085626.pdf", "page_number": 9, "total_pages": 24, "image_filename": "19930085626_p9.jpg", "text": "Rocket motor\nCONFIDENTIAL\nSpin-sonde\n5.00 Diam.\n39.00\n56.00\n14.00\nDimensions are in inches\n.24\n.93\n4.05\n.93\nc/4 line\nA\nA\n40°\n1.09\n8.48\n1.09\n7.00\n4.50\n.020 steel plates\nSection A-A\nNACA\nCircular-arc section normal\nto c/4 line. Thickness ratio, 0.10\n(NACA 25-(50)(05)-(50)(05))\nFigure 1.— General arrangement of test vehicles. Original aileron configuration.\nCONFIDENTIAL\nNACA RM No. L58C03\n8", "timestamp": "2026-07-22T05:11:12.981165+00:00"}
{"citation_id": "19930085879", "source_url": "https://ntrs.nasa.gov/api/citations/19930085879/downloads/19930085879.pdf", "page_number": 7, "total_pages": 29, "image_filename": "19930085879_p7.jpg", "text": "NACA RM L9D11\n\nDISCUSSION\n\nThe flights of RM-11 models indicate that the nose section should be jettisoned after the thrust has been cut off. This might not be difficult to do by having the escape device also cut off the fuel supply, but the time lag between fuel cut-off and complete loss of thrust is very important. Even if the design were such that the rear body became unstable after jettisoning the nose, because of the short distance of forcible separation, the shielding effect of the nose on the rear body, and the large moment of inertia of the rear body, it seems improbable that ejection of the nose during power-on flight could be accomplished safely. The flight of model A indicates that the three fins remaining on the model after collision with the wing were sufficient to damp out oscillations but the asymmetry after loss of one fin caused the model to follow a helical path. The accelerations during the flight indicate that escape might have been possible.\n\nThe flight of model B indicates that the separation phase of escape in this case would not have caused any great discomfort to a pilot. While an effort was made to make the ratio between the drag and weight of the nose and rear sections of the order of that of a full-scale airplane, this test represents an extreme case in that the aft body was stable after removal of the nose section. In the case of a conventional airplane, the aft fuselage after nose release would probably be unstable, and, because of this, its drag would be greatly increased.\n\nThe longitudinal acceleration-time curves shown in figure 9 indicate that the initial push given the nose section was necessary. When the two bodies were 4 to 5 inches apart, the shielding effect of the nose on the aft body caused the drag of the aft body to be so low that its deceleration was less than that of the nose. The initial added velocity given the nose section by the jettison charge was sufficient to widen the gap until the drag-weight ratios became favorable. Whether or not this would be true in the case of a full-scale airplane would depend upon the individual configuration of the airplane and nose section.\n\nOne factor that must be given consideration at high Mach numbers is the deceleration due to drag experienced after ejection. In figure 15 the instantaneous deceleration calculated for a 1500-pound nose section is plotted for three different altitudes using the RM-11 configuration but having linear dimensions five times those of the RM-11. In addition, the average decelerations experienced over a given elapsed time as computed from these instantaneous values are also presented for comparison with human-tolerance values. The line showing human tolerance is taken from unpublished data and is for a human body fully extended with the accelerations transverse to the human body. While the acceleration on the pilot in a jettisoned nose is applied in the same direction, some variation must be expected because of a pilot having his legs forward in a sitting position. The plot indicates that the deceleration due to drag is not serious at a Mach number of 2.0 at altitudes above 30,000 feet.", "timestamp": "2026-07-22T05:11:13.172627+00:00"}
{"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 28, "total_pages": 78, "image_filename": "19930082618_p28.jpg", "text": "26\n\nSection lift coefficient, $c_l$\nMoment coefficient, $c_{m_{c/4}}$\nSection angle of attack, $\\alpha_{or}$ deg\n\nR\nO 0.7 x $10^6$\nO 1.0\n$\\diamond$ 1.5\n$\\triangle$ 2.0\n$\\nabla$ 5.0\nFlagged symbols denote\nstandard roughness\n\nNACA\n\n(b) Section lift and pitching-moment characteristics of the NACA 64$_1$-412 airfoil section with a\n0.20c simulated split flap deflected 60$^\\circ$.\n\nFigure 2.- Continued.\n\nNACA TN 1945", "timestamp": "2026-07-22T05:11:16.705456+00:00"}
{"citation_id": "19930085519", "source_url": "https://ntrs.nasa.gov/api/citations/19930085519/downloads/19930085519.pdf", "page_number": 25, "total_pages": 46, "image_filename": "19930085519_p25.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:11:25.561556+00:00"}
{"citation_id": "19930082914", "source_url": "https://ntrs.nasa.gov/api/citations/19930082914/downloads/19930082914.pdf", "page_number": 30, "total_pages": 66, "image_filename": "19930082914_p30.jpg", "text": "NACA TN No. 1857\n29\n\nREFERENCES\n\n1. Tollmien, Walter: Calculation of Turbulent Expansion Processes.\nNACA TM No. 1085, 1945.\n\n2. Görtler, H.: Berechnung von Aufgaben der freien Turbulenz auf\nGrund eines neuen Näherungsansatzes. Z.f.a.M.M., Bd. 22, Nr. 5,\nOct. 1942, pp. 244-254.\n\n3. Abramovich, G. N.: The Theory of a Free Jet of a Compressible\nGas. NACA TM No. 1058, 1944.\n\n4. Kuethe, Arnold M.: Investigations of the Turbulent Mixing\nRegions Formed by Jets. Jour. Appl. Mech., vol. 2, no. 3,\nSept. 1935, pp. A-87 - A-95.\n\n5. Förthmann, E.: Turbulent Jet Expansion. NACA TM No. 789, 1936.\n\n6. Reichardt, H.: Laws of Free Turbulence. R.T.P. Translation No. 1752,\nBritish Ministry of Aircraft Production. (From V.D.I. Forschungsheft\nNo. 414, May/June 1942, pp. 1-22.)\n\n7. Corrsin, Stanley: Investigation of Flow in an Axially Symmetrical\nHeated Jet of Air. NACA ACR No. 3L23, 1943.\n\n8. Liepmann, Hans Wolfgang, and Laufer, John: Investigations of Free\nTurbulent Mixing. NACA TN No. 1257, 1947.\n\n9. Zobel, Th.: Development and Construction of an Interferometer\nfor Optical Measurements of Density Fields. NACA TM No. 1184,\n1947.\n\n10. Ladenburg, R., Van Voorhis, C. C., and Winckler, J.: Interferometric\nStudy of Supersonic Phenomena. Part I: A Supersonic Air Jet\nat 60 LB/IN² Tank Pressure. NAVORD Rep. 69-46, Bur. Ordnance,\nNavy Dept., April 17, 1946; Part II: The Gas Flow around Various\nObjects in a Free Homogeneous Supersonic Air Stream. NAVORD\nRep. 93-46, Sept. 2, 1946.\n\n11. Schardin, H.: Theory and Applications of the Mach-Zehnder\nInterference-Refractometer. Univ. of Texas, Defense Res. Lab.,\n1946.", "timestamp": "2026-07-22T05:11:27.985453+00:00"}
{"citation_id": "19930082498", "source_url": "https://ntrs.nasa.gov/api/citations/19930082498/downloads/19930082498.pdf", "page_number": 38, "total_pages": 49, "image_filename": "19930082498_p38.jpg", "text": "NACA TN No. 1838\n\n[Figure: A black-and-white photograph showing two cylindrical metal objects lying on a concrete surface. The object on the left is labeled with a white circular tag marked \"45\". The object on the right is labeled with a white circular tag marked \"62\" and has \"REB 176\" stenciled on its side. A rod connects the two objects. In the bottom right corner of the photo, there is a NACA stamp with the number \"L-53828\".]\n\n(e) Configuration 68, which consists of configurations 45 and 62 in combination.\n\nFigure 2.- Concluded.\n\n37", "timestamp": "2026-07-22T05:11:28.782291+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 44, "total_pages": 99, "image_filename": "19930082511_p44.jpg", "text": "42\nNACA TN No. 1826\n\npotential of a horseshoe vortex). It is desired to find a function $\\phi(\\xi, \\rho, \\theta)$, harmonic inside the cylinder $\\rho = 1$, for which (see part I)\n\n$$\n\\left. \\frac{\\partial(\\phi_0 + \\phi)}{\\partial\\rho} \\right|_{\\rho=1} = 0 \\quad (\\xi < a, \\xi > b)\n$$\n\n$$\n\\left. \\frac{\\partial(\\phi_0 + \\phi)}{\\partial\\xi} \\right|_{\\rho=1} = 0 \\quad (a < \\xi < b)\n$$\n\nwhere the region $\\xi < a, \\xi > b$ is the closed portion of the tunnel and $a < \\xi < b$ is the open portion of the tunnel. In addition, according to the condition of continuity at the entrance lip (part I), the derivative $\\left. \\frac{\\partial\\phi}{\\partial\\rho} \\right|_{\\rho=1}$ must be continuous at $\\xi = a$. The function $\\phi$ is then the velocity potential of the additional flow due to the tunnel boundary.\n\nThe function $\\phi$ is conveniently considered in two parts:\n\n$$\n\\phi = \\phi_C + \\phi_A\n$$\n\nwhere $\\phi_C$ is the known tunnel-induced potential for the same vortex system in a completely closed circular tunnel (reference 13). The determining conditions for $\\phi_A$ are then\n\n(1) $\\Delta\\phi_A = 0 \\quad (\\rho < 1)$\n\n(2) $\\left. \\frac{\\partial\\phi_A}{\\partial\\rho} \\right|_{\\rho=1} = 0 \\quad (\\xi < a, \\xi > b)$\n\n(3) $\\left. \\frac{\\partial\\phi_A}{\\partial\\xi} \\right|_{\\rho=1} = - \\left. \\frac{\\partial(\\phi_0 + \\phi_C)}{\\partial\\xi} \\right|_{\\rho=1} \\quad (a < \\xi < b)$\n\n(4) $\\left. \\frac{\\partial\\phi_A}{\\partial\\rho} \\right|_{\\rho=1} = 0 \\quad (\\xi = a)$", "timestamp": "2026-07-22T05:11:29.728987+00:00"}
{"citation_id": "19930082496", "source_url": "https://ntrs.nasa.gov/api/citations/19930082496/downloads/19930082496.pdf", "page_number": 42, "total_pages": 50, "image_filename": "19930082496_p42.jpg", "text": "NACA TN No. 1836\n41\n\n[Figure: Two fractured ceramic blade fragments shown side-by-side against a dark background. Below the image is a scale bar labeled \"1 INCH\". To the right of the scale bar is a NACA logo with the text \"C-20994\" and \"4-1-48\" underneath.]\n\nFigure 11. - Ceramal blade inadvertently fractured after 9 hours and 42 minutes of operation.", "timestamp": "2026-07-22T05:11:33.182347+00:00"}
{"citation_id": "19930085487", "source_url": "https://ntrs.nasa.gov/api/citations/19930085487/downloads/19930085487.pdf", "page_number": 29, "total_pages": 36, "image_filename": "19930085487_p29.jpg", "text": "NACA RM No. EBJ22\n27\n\n100,000\n\n10,000\n\nFrequency, cps\n\n1,000\n\n100\n\n4 6 8 10 12 14 16x10³\nRotor speed, rpm\n\nRear stator row, 40 blades\nFront stator row, 38 blades\nFirst torsional mode\nSecond bending mode\nFirst bending mode\nSecond order, split compressor base\nFirst order, Rotor speed\n\n(c) Third stage.\nFigure 9. - Continued. Critical-speed diagrams for 10 stages of compressor rotor.\n\n[Figure: NACA logo]", "timestamp": "2026-07-22T05:11:34.002007+00:00"}
{"citation_id": "19930085869", "source_url": "https://ntrs.nasa.gov/api/citations/19930085869/downloads/19930085869.pdf", "page_number": 9, "total_pages": 36, "image_filename": "19930085869_p9.jpg", "text": "NACA RM L9D15 CONFIDENTIAL 7\n\nthat the model can be directionally controlled without the use of asymmetric power.\n\nCONCLUSIONS\n\nThe results of tests to determine the hydrodynamic characteristics of the swept planing-tail flying-boat lead to the following conclusions:\n\n1. With vertical spray strips, only light spray struck the propellers over a short speed range before the hump. No spray came over the bow. Heavy spray from the forebody roach struck the tail surfaces over a short speed range near the hump speed. Light spray wetted the under surface of the wing at high speeds.\n\n2. Although the peak of the lower limit was high ($10.5^\\circ$), a minimum trim of $2^\\circ$ at high speed rather than lower-limit porpoising determined the minimum elevator deflection for take-off.\n\n3. Upper-limit porpoising occurred over a short speed range near take-off.\n\n4. A large range of elevator deflection was available for take-offs over a wide range of center-of-gravity locations.\n\n5. All landings at center-of-gravity locations from 0.20$\\bar{c}$ to 0.40$\\bar{c}$ were stable.\n\n6. The hump load-resistance ratio of 3.1 was lower than ratios obtained with conventional hulls.\n\nLangley Aeronautical Laboratory\nNational Advisory Committee for Aeronautics\nLangley Air Force Base, Va.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:11:34.794542+00:00"}
{"citation_id": "19930085881", "source_url": "https://ntrs.nasa.gov/api/citations/19930085881/downloads/19930085881.pdf", "page_number": 3, "total_pages": 31, "image_filename": "19930085881_p3.jpg", "text": "NACA RM L9D12 CONFIDENTIAL\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nRESEARCH MEMORANDUM\n\nEFFECTS OF SOME AIRFOIL-SECTION VARIATIONS ON WING-AILERON ROLLING EFFECTIVENESS AND DRAG AS DETERMINED IN FREE FLIGHT AT TRANSONIC AND SUPERSONIC SPEEDS\n\nBy Carl A. Sandahl, William M. Bland, Jr., and H. Kurt Strass\n\nSUMMARY\n\nAn experimental investigation has been made in free flight of the rolling effectiveness of plain ailerons in conjunction with wings having $0^\\circ$ and $45^\\circ$ sweepback employing several airfoil sections. The total drag of the test vehicles was also obtained. Positive control effectiveness over the Mach number range investigated was obtained for the configurations which employed airfoil sections with trailing-edge angles of the order of $10^\\circ$. Reversal of effectiveness was encountered for configurations with trailing-edge angles of the order of $20^\\circ$. The aileron effectiveness was not appreciably affected by changes in the shape of the forward part of the airfoil section. For the rectangular wings near Mach numbers of unity, the blunt-nose sections had slightly lower drag than did the sharp-nose sections. At higher supersonic Mach numbers the drag of the rectangular wings having sharp-nose sections approached that of the sweptback wings. For the sweptback wings, variations of airfoil section produced no measurable differences in drag for the Mach number range investigated.\n\nINTRODUCTION\n\nAt the present time the Pilotless Aircraft Research Division of the Langley Laboratory is engaged in an investigation of wing-aileron rolling-effectiveness characteristics at transonic and supersonic speeds utilizing rocket-propelled test vehicles in free flight. A considerable amount of systematic experimental information relating to plain ailerons on wings of various plan forms having NACA 65A009 airfoil sections has been obtained by the techniques described in references 1 and 2. The results obtained have been summarized in reference 3.\n\nIn order to evaluate the effects of some major variations in airfoil section on the rolling effectiveness of plain ailerons, tests were made of configurations having NACA 65A009, NACA 16-009, symmetrical double-wedge and circular-arc airfoil sections of 9-percent thickness ratio. These four sections were tested with wings of aspect ratio 3.71,\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:11:38.768759+00:00"}
{"citation_id": "19930085889", "source_url": "https://ntrs.nasa.gov/api/citations/19930085889/downloads/19930085889.pdf", "page_number": 1, "total_pages": 37, "image_filename": "19930085889_p1.jpg", "text": "NACA RM L9F14\n198\nCopy\nRM L9F14\n\nCONFIDENTIAL\n\nNACA\n\nRESEARCH MEMORANDUM\n\nEFFECT OF SWEEPBACK ON THE LOW-SPEED STATIC AND\nROLLING STABILITY DERIVATIVES OF THIN TAPERED\nWINGS OF ASPECT RATIO 4\n\nBy\nWilliam Letko and Walter D. Wolhart\n\nLangley Aeronautical Laboratory\nLangley Air Force Base, Va.\n\nCLASSIFICATION CHANGED TO\nUNCLASSIFIED\nDATE 8-18-54\n\nCLASSIFIED DOCUMENT\nThis document contains classified information\naffecting the National Defense of the United\nStates within the meaning of the Espionage\nLaws, U.S.C. 50 and 50. Its transmission or the\nrevelation of its contents in any manner to an\nunauthorized person is prohibited by law.\nInformation to classified may be imparted\nonly to persons in the military and naval\nservices of the United States, appropriate\ncivilian officers and employees of the Federal\nGovernment who have a legitimate interest\ntherein, and to United States citizens of known\nloyalty and discretion who of necessity must be\ninformed thereof.\n\nAUTHORITY J.W.CROWLEY\nCHANGE # 2440\nW.H.L.\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\nWASHINGTON\nAugust 9, 1949\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:11:44.932708+00:00"}
{"citation_id": "19930082245", "source_url": "https://ntrs.nasa.gov/api/citations/19930082245/downloads/19930082245.pdf", "page_number": 55, "total_pages": 66, "image_filename": "19930082245_p55.jpg", "text": ".16\n.14\n.12\n1.0\n.8\n.6\n.4\n.2\n0\n-.2\n-.4\n-.6\n.1 .2 .3 .4 .5 .6 .7 .8 .9\nMach number, M\nAileron section normal-force coefficient, $c_{n\\alpha}$\n$\\delta_a$\n(deg)\n30\n18\n12\n6\n4\n2\n0\n-2\n-4\n-6\n-12\n18\n12\n2\n-4\n\n.16\n.12\n.08\n.04\n0\n-.04\n-.08\n-.12\n-.16\n-.20\n-.24\n-.28\n.1 .2 .3 .4 .5 .6 .7 .8 .9\nMach number, M\nAileron section hinge-moment coefficient, $c_h$\n$\\delta_a$\n(deg)\n-12\n-6\n-4\n-2\n0\n2\n4\n6\n12\n18\n30\n12\n-6\n-4\n-2\n0\n2\n4\n6\n12\n18\n30\n\n(b) $c_n = 0.4$.\nFigure 10.-Continued.\n\nNACA\nNACA TN No. 1596\n54", "timestamp": "2026-07-22T05:11:46.344385+00:00"}
{"citation_id": "19930085626", "source_url": "https://ntrs.nasa.gov/api/citations/19930085626/downloads/19930085626.pdf", "page_number": 10, "total_pages": 24, "image_filename": "19930085626_p10.jpg", "text": "NACA RM No. L8K23\n9\n\nCONFIDENTIAL\n\n[Figure: A photograph of a white rocket model with a pointed nose cone and four fins at the base, standing vertically against a dark background. A label in the bottom right corner of the photo reads \"NACA L-55427\".]\n\n(a) Original aileron configuration.\n\nFigure 2.- Photographs of test vehicles.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:11:47.756947+00:00"}
{"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 29, "total_pages": 78, "image_filename": "19930082618_p29.jpg", "text": "```markdown\nNACA TN 1945\n\n<!-- Image (79, 108, 878, 806) -->\n\n(c) Section drag characteristics and section pitching-moment characteristics about the aerodynamic center of the plain NACA 641-412 airfoil section.\n\nFigure 2.- Concluded.\n\n27\n```", "timestamp": "2026-07-22T05:11:47.926579+00:00"}
{"citation_id": "19930082614", "source_url": "https://ntrs.nasa.gov/api/citations/19930082614/downloads/19930082614.pdf", "page_number": 29, "total_pages": 36, "image_filename": "19930082614_p29.jpg", "text": "NACA TN 1939\n27\n\nExample 2. For n=6.6,\nV=600 ft/sec, $\\gamma$=30°\n\nExample 1.\n(See table III)\n$\\gamma$=\n0°\n30°\n40°\n50°\n60°\n70°\n80°\n90°\n\nr=1,400,000\n2\n3\n4\n6\n8\n10\n15\n20\n25\n30\n40\n80\n\nn-cos $\\gamma$\nNormal acceleration factor, n\n7\n6\n5\n4\n3\n2\n1\n0\n\nAverage airspeed, V, ft/sec\n0 200 400 600 800 1000 2000\n\n$\\Delta \\gamma$=10.8 sec\n0.4\n0.6\n0.8\n1.0\n1.2\n1.4\n1.6\n1.8\n2.0\n2.2\n2.4\n2.6\n3.0\n4.0\n5.0\n\n$\\Delta \\gamma$=10.8 sec\n400\n500\n600\n700\n800\n900\n1000\n1200\n1400\n\n$\\Delta \\gamma$, deg/sec\n28-1\n24-2\n20-3\n16-4\n12\n8\n4\n0\n\nChange in $\\gamma$, $\\Delta \\gamma$, deg\n28 24 20 16 12 8 4 0\n\nn-cos $\\gamma$\nr=1,400,000\n2\n3\n4\n6\n8\n10\n15\n20\n30\n40\n80\n$\\infty$\n\nr=1,400,000\n2\n3\n4\n6\n8\n10\n15\n20\n30\n40\n80\n$\\infty$\n\ncos $\\gamma$-cos($\\gamma$-$\\Delta \\gamma$)\n0.5\n0.4\n0.3\n0.2\n0.1\n0\n-0.01\n-0.02\n-0.03\n-0.04\n\nIf $\\Delta \\gamma$ is positive\n$\\gamma$=\n90 deg\n80\n70\n60\n50\n40\n30\n20\n10\n0\n-10\n-20\n-30\n-40\n-50\n-60\n-70\n-80\n-90\n\nIf $\\Delta \\gamma$ is negative\n$\\gamma$=\n90 deg\n80\n70\n60\n50\n40\n30\n20\n10\n0\n-10\n-20\n-30\n-40\n-50\n-60\n-70\n-80\n-90\n\nNACA\n\nChange in altitude, $\\Delta h$, ft\n0 0.4 0.8 1 2 3 4x1000 2 8 24 20 16 12 8 4 0\n\nChange in $\\gamma$, $\\Delta \\gamma$, deg\n0 0.4 0.8 1 2 3 4x1000 2 8 24 20 16 12 8 4 0\n\nFigure 5.— Graphical solution for changes of altitude and flight-path angle.\n(A larger copy of this figure is enclosed in an envelope at the end of the report.)", "timestamp": "2026-07-22T05:11:51.495085+00:00"}
{"citation_id": "19930085544", "source_url": "https://ntrs.nasa.gov/api/citations/19930085544/downloads/19930085544.pdf", "page_number": 13, "total_pages": 33, "image_filename": "19930085544_p13.jpg", "text": "12\nNACA RM No. L8K26\n\nsections. A description of the propeller and blade-form curves is given in reference 3. The calculations made, assuming steady flow, were for a two-blade propeller for blade angles of $26^\\circ$ and $53^\\circ$ measured at the 0.75 radius, for a free-stream Mach number of 0.30, and for the propeller thrust axis inclined at an angle of $4^\\circ$. For each operating condition of the propeller there is a variation of Mach number along the blade which must be taken into account. The airfoil data used were taken from reference 5 but, since the highest Mach number covered in the report was 0.7, extrapolation of the airfoil data to Mach numbers as high as 0.9 was necessary. The lift characteristics for a Mach number of 0.6 were used to extrapolate to higher Mach numbers. The extrapolation was made by holding the angle of zero lift obtained at M = 0.6 constant and changing the slope of the lift curve by the Prandtl-Glauert relationship\n\n$$\n\\left(\\frac{dc_l}{d\\alpha}\\right)_M = \\left(\\frac{dc_l}{d\\alpha}\\right)_{M=0.6} \\frac{\\sqrt{1 - (0.6)^2}}{\\sqrt{1 - M^2}}\n$$\n\nFigure 5 gives a comparison of the variation of the calculated and the wake-survey thrust coefficients with respect to the blade position at at J = 1.2 for a blade-angle setting of $26^\\circ$ and a thrust-axis angle $\\alpha_T$ of $4^\\circ$. Figure 6 gives a similar comparison at J = 2.8, $\\beta = 53^\\circ$, and $\\alpha_T = 4^\\circ$, and figure 7 at J = 3.1 for $\\beta = 53^\\circ$ and $\\alpha_T = 4^\\circ$. The maximum and minimum calculated instantaneous thrust coefficients differ slightly from the measured values in all cases. There is also a rather large phase difference in the position of the maximum calculated and measured thrust coefficient at the value J = 1.2. The position of the survey rake for the experimental data of reference 3 was changed to get each individual point rather than making simultaneous measurements of a number of points for each particular operating condition.\n\nAt the same time that the experimental rake survey data were taken force-test thrust coefficients were measured. A study of these data revealed that the force-test thrust coefficients ($C_T$) at the operating V/nD of 1.2 for the $26^\\circ$ blade-angle setting varied from $C_T = 0.0125$ to $C_T = 0.0170$ depending on the position of the survey rake. The force-test thrust coefficient for the condition of V/nD of 2.8 and the blade-angle setting of $53^\\circ$ varied from 0.0725 to 0.0800. This change in thrust coefficient with rake position suggests that there is a blocking effect which changes with rake position. This blocking effectively changes the velocity in the plane of the propeller and thus the operating V/nD of the propeller. Therefore comparisons made at the same V/nD with theoretical calculations based on free-air conditions would not be expected to", "timestamp": "2026-07-22T05:11:53.033792+00:00"}
{"citation_id": "19930085519", "source_url": "https://ntrs.nasa.gov/api/citations/19930085519/downloads/19930085519.pdf", "page_number": 26, "total_pages": 46, "image_filename": "19930085519_p26.jpg", "text": "NACA RM No. L8K19\n\n[Figure: Diagram showing locations and details of flap-slot flow-control vanes B on a 42° sweptback wing, with dimensions labeled in inches. Includes Section A-A view and annotations such as .18 b/2, .35 b/2, .53 b/2, .70 b/2, .88 b/2, 2.2, 5.2, 4.0.]\n\nSection A-A\n\nFigure 6.- Locations and details of flap-slot flow-control vanes B on the 42° sweptback wing. All dimensions in inches unless otherwise indicated.\n\nNACA\n\n25", "timestamp": "2026-07-22T05:11:55.259208+00:00"}
{"citation_id": "19930082914", "source_url": "https://ntrs.nasa.gov/api/citations/19930082914/downloads/19930082914.pdf", "page_number": 31, "total_pages": 66, "image_filename": "19930082914_p31.jpg", "text": "30\n\nPage intentionally left blank\n\nPage intentionally left blank", "timestamp": "2026-07-22T05:11:56.219780+00:00"}
{"citation_id": "19930085879", "source_url": "https://ntrs.nasa.gov/api/citations/19930085879/downloads/19930085879.pdf", "page_number": 8, "total_pages": 29, "image_filename": "19930085879_p8.jpg", "text": "6\nNACA RM L9D11\n\nIt has been found that nose fins produce a detrimental effect on the stability of a complete airplane. The problem of making retractable fins that are of sufficient stiffness, take up little space, and are still capable of being extended almost instantaneously is very great. One possible method of compromise might be to have the fins permanently attached but able to float freely. They could be locked in position of zero incidence quickly and with a fairly simple mechanism. By using this method the drag of the fins is present but it would probably eliminate the destabilizing effect. An investigation of the effect of such fins on stability and of susceptibility to flutter would be necessary.\n\nSince the terminal velocity of most jettisoned nose sections would probably be too high for direct escape into the air stream, some means of slowing down the nose will probably be necessary. In view of the fact that it is desirable to keep any escape device as simple as possible, it might be desirable to use a drag parachute to decrease the terminal velocity rather than drag flaps.\n\nCONCLUSIONS\n\nResults obtained from tests of rocket-propelled models to test the jettisonable-nose method of pilot escape indicated the following:\n\n(1) If the nose section is released during power-on flight, there is danger of collision with the main body of the airplane.\n\n(2) With suitable stabilizing fins the nose section may be forcibly ejected at a Mach number of 0.87 during power-off stable flight without producing accelerations dangerous to a pilot.\n\n(3) The drag-weight ratio of such a nose section should be made sufficiently less than the drag-weight ratio of the main body so that collision of the two after ejection is impossible.\n\n(4) The shielding effect of the nose section on the main body is considerable, and forcible ejection seems necessary for smooth separation.\n\nCalculations for the configuration tested indicated that the deceleration due to drag on the nose section after separation will not be dangerous to a pilot at Mach number 2.0 at altitudes above 30,000 feet.\n\nLangley Aeronautical Laboratory\nNational Advisory Committee for Aeronautics\nLangley Air Force Base, Va.", "timestamp": "2026-07-22T05:11:57.558115+00:00"}
{"citation_id": "19930082485", "source_url": "https://ntrs.nasa.gov/api/citations/19930082485/downloads/19930082485.pdf", "page_number": 50, "total_pages": 62, "image_filename": "19930082485_p50.jpg", "text": "NACA TN No. 1810\n49\n\n$\\sqrt{I_2}/\\sqrt{I_m}$\n$\\frac{C_1}{\\Delta C}$\n\n| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | |", "timestamp": "2026-07-22T05:12:03.378803+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 5, "total_pages": 96, "image_filename": "19930085880_p5.jpg", "text": "```markdown\nNACA RM No. L9C03\n3\n\nthe chine of the model. (See fig. 5.) Forward of this boundary there\nwas a region of loose spray which seemed to fan out from the boundary.\nBecause of the transverse curvature of model 250B some difficulty was\nencountered in visually reading wetted lengths particularly at the lower\ntrims. A few underwater photographs (see fig. 5) showed that the\nvisually read wetted lengths were satisfactory except at $4^\\circ$ trim. There-\nfore, additional underwater photographs were taken at $4^\\circ$ trim and the\nwetted lengths were obtained from these photographs.\n\nDue to a forward curvature of the solid water boundary the wetted\nlength at the center line of all models was greater than the observed\nwetted length. The manner in which this curvature varied with the models\nand their wetted length, trim, and speed was determined from underwater\nphotographs.\n\nThe slight curvature of the solid water boundary on model 250A\n(rectangular plan form, flat bottom) gave only a small difference between\nthe wetted length at the center line and the observed wetted length.\nThis difference was within the experimental scatter of the test data\nobtained, and therefore was not considered.\n\nDue to the transverse bottom curvature of model 250B, the difference\nin the two wetted lengths varied considerably with trim. However, the\ndifference was less than that which would be indicated by the intersection\nof the water surface with the curved bottom surface of the model at rest\nas is shown in figure 5. Figure 6 indicates the difference in the wetted\nlength at the center line and the observed wetted length at the chine.\n\nFor model 250D (triangular plan form, flat bottom) the dynamic water\nline was in the form of an arc having a ratio of mid-ordinate to chord\n(beam) equal to 0.10. Within the range of trims and speeds covered in\nthese tests this ratio did not vary appreciably. Because of the trian-\ngular plan form of the model the chord of this arc and therefore the mid-\nordinate, varied with wetted length or wetted area. Figure 7 shows the\nresulting difference in the wetted length at the center line and the\nobserved wetted length at the chine.\n\nWetted Area\n\nThe wetted area is defined as the wetted plan-form area. This area\nwas determined from the plan form of the models, the observed wetted\nlength at the chine, and the additional wetted area forward of the\nobserved wetted length. The additional wetted area forward of the\nobserved wetted length was determined from the underwater photographs.\n\nThe wetted area forward of the observed wetted length varied in\nthe same manner as the wetted length forward of the observed wetted\nlength. The additional area involved was neglected for model 250A. The\narea involved for model 250B varied appreciably only with trim and is\n```", "timestamp": "2026-07-22T05:12:07.218347+00:00"}
{"citation_id": "19930082496", "source_url": "https://ntrs.nasa.gov/api/citations/19930082496/downloads/19930082496.pdf", "page_number": 43, "total_pages": 50, "image_filename": "19930082496_p43.jpg", "text": "42\n\nPage intentionally left blank\n\nPage intentionally left blank", "timestamp": "2026-07-22T05:12:07.430773+00:00"}
{"citation_id": "19930082498", "source_url": "https://ntrs.nasa.gov/api/citations/19930082498/downloads/19930082498.pdf", "page_number": 39, "total_pages": 49, "image_filename": "19930082498_p39.jpg", "text": "Page intentionally left blank\n\nPage intentionally left blank", "timestamp": "2026-07-22T05:12:07.605001+00:00"}
{"citation_id": "19930082617", "source_url": "https://ntrs.nasa.gov/api/citations/19930082617/downloads/19930082617.pdf", "page_number": 27, "total_pages": 58, "image_filename": "19930082617_p27.jpg", "text": "```markdown\n26\nNACA TN 1962\n\n10\n9\n8\n7\n6\n5\n4\n3\n2\n1\n0\n1\n2\n3\n4\n5\n6\n7\n8\n9\n10\n\nDistance from horizontal diameter, in.\n\nStringers\no 1 to 9\nx 10 to 16\n\nMoment\n(in. - lb)\n1 54.0 X $10^3$\n2 108.0 X $10^3$\n3 162.0 X $10^3$\n4 216.0 X $10^3$\n5 270.0 X $10^3$\n\n3.00\"\nA\nBand H\nA\n45° < 45°\nA-A\n\n16 12 8 4 0 -4 -8 -12 -16 -20 X $10^{-4}$\nStrain\n\n1 2 3 4 5\n\nNACA\n\nFigure 15.- Strain diagram of cylinder 76. Band H.\n```", "timestamp": "2026-07-22T05:12:16.027423+00:00"}
{"citation_id": "19930085869", "source_url": "https://ntrs.nasa.gov/api/citations/19930085869/downloads/19930085869.pdf", "page_number": 10, "total_pages": 36, "image_filename": "19930085869_p10.jpg", "text": "8\nCONFIDENTIAL\nNACA RM L9D15\n\nREFERENCES\n\n1. Riebe, John M., and Naeseth, Rodger L.: Aerodynamic Characteristics\nof a Refined Deep-Step Planing-Tail Flying-Boat Hull with Various\nForebody and Afterbody Shapes. NACA RM No. L8F01, 1944.\n\n2. McKann, Robert, and Coffee, Claude W.: Hydrodynamic Characteristics\nof Aerodynamically Refined Planing-Tail Hulls. NACA RM No. L9B04,\n1949.\n\n3. Naeseth, Rodger L., and MacLeod, Richard G.: Aerodynamic Characteristics\nof an Airfoil-Forebody Swept Flying-Boat Hull with a Wing and Tail\nSwept Back $51.3^\\circ$ at the Leading Edge. NACA RM No. L9F08, 1949.\n\n4. Riebe, John M., and MacLeod, Richard G.: Preliminary Wind-Tunnel\nInvestigation at High-Subsonic Speeds of Planing-Tail, Blended,\nand Airfoil-Forebody Swept Hulls. NACA RM No. L9D01, 1949.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:12:17.262533+00:00"}
{"citation_id": "19930085626", "source_url": "https://ntrs.nasa.gov/api/citations/19930085626/downloads/19930085626.pdf", "page_number": 11, "total_pages": 24, "image_filename": "19930085626_p11.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:12:18.647929+00:00"}
{"citation_id": "19930085542", "source_url": "https://ntrs.nasa.gov/api/citations/19930085542/downloads/19930085542.pdf", "page_number": 16, "total_pages": 46, "image_filename": "19930085542_p16.jpg", "text": "14\nNACA RM No. L8L29\n\nTABLE I.- PERTINENT MODEL DIMENSIONS AND TEST CONDITIONS\n\n| Model | Profile | Plan form | $\\Lambda_{T.E.}$ (deg) | $\\Lambda_{L.E.}$ (deg) | Aspect ratio | Span (in.) | Root chord (in.) | M.A.C. (in.) | $\\bar{c}$ (in.) | Area (sq in.) | R | $\\frac{pb}{2V}$ (radians) | Taper ratio |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| 1 | Flat plate | [Triangle] | 60 | 58.2 | 2.31 | 36.07 | 31.23 | 20.82 | 15.62 | 564.0 | $1.603 \\times 10^6$ | $\\pm .005$ <br> $\\pm .075$ | 0 |\n| 2 | NACA 0012 | [Triangle] | 60 | 58.2 | 2.31 | 36.00 | 31.60 | 21.10 | 15.83 | 576.0 | 1.624 | $\\pm .005$ <br> $\\pm .075$ | 0 |\n| 3 | Biconvex 12 percent | [Triangle] | 60 | 58.2 | 2.31 | 36.00 | 31.60 | 21.10 | 15.83 | 576.0 | 1.624 | $\\pm .005$ <br> $\\pm .075$ | 0 |\n| 4 | NACA 0012 | [Triangle] | 75 | 70.4 | 1.07 | 24.85 | 46.37 | 30.90 | 23.20 | 576.0 | 2.180 | $\\pm .017$ <br> $\\pm .051$ | 0 |\n| 5 | NACA 0012 and flat-plate fin | [Delta wing with fin] | Wing, 60 <br> Fin, 68.9 | 58.2 | Wing, 2.31 <br> Fin, 0.77 | Wing, 36.0 <br> Fin, 12.10 | 31.60 | 21.10 | 15.83 | Wing, 576 <br> Fin, 192 | 1.624 | $\\pm .005$ <br> $\\pm .075$ | 0 |\n| 6 | NACA 0012 and flat-plate fin | [Delta wing with fin] | Wing, 60 <br> Fin, 60 | 58.2 | Wing, 2.31 <br> Fin, 1.45 | Wing, 36.0 <br> Fin, 18.25 | 31.60 | 21.10 | 15.83 | Wing, 576 <br> Fin, 288 | 1.624 | $\\pm .005$ <br> $\\pm .075$ | 0 |\n| 7 | NACA 0012 | [Triangle] | 45 | 36.9 | 4.0 | 48.00 | 24.00 | 16.00 | 12.00 | 576.0 | 1.232 | $\\pm .033$ <br> $\\pm .099$ | 0 |\n| 8 | NACA 0012 | [Trapezoid] | 45 | 36.9 | 3.0 | 41.20 | 24.00 | 16.30 | 11.77 | 563.6 | 1.254 | $\\pm .028$ <br> $\\pm .084$ | 0.15 |\n| 9 | NACA 0012 | [Pentagon] | 45 | 36.9 | 2.0 | 31.80 | 24.00 | 17.10 | 10.95 | 507.8 | 1.335 | $\\pm .022$ <br> $\\pm .066$ | 0.36 |\n| 10 | NACA 0012 | [Hexagon] | 45 | 36.9 | 1.0 | 18.80 | 24.00 | 19.60 | 9.13 | 355.8 | 1.510 | $\\pm .013$ <br> $\\pm .039$ | 0.58 |\n\nNACA", "timestamp": "2026-07-22T05:12:19.439830+00:00"}
{"citation_id": "19930085889", "source_url": "https://ntrs.nasa.gov/api/citations/19930085889/downloads/19930085889.pdf", "page_number": 2, "total_pages": 37, "image_filename": "19930085889_p2.jpg", "text": "NACA RM L9F14\nCONFIDENTIAL\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\nRESEARCH MEMORANDUM\nEFFECT OF SWEEPBACK ON THE LOW-SPEED STATIC AND\nROLLING STABILITY DERIVATIVES OF THIN TAPERED\nWINGS OF ASPECT RATIO 4\nBy William Letko and Walter D. Wolhart\nSUMMARY\nA low-speed investigation was made in the Langley stability tunnel\nto determine the effect of sweepback on the static and rolling stability\nderivatives of a series of wings, each of which had a taper ratio\nof 0.6 and an aspect ratio of 4. The wings were of NACA 65A006 section\nin planes parallel to the axis of symmetry and had sweepback angles of\ntheir quarter-chord line of $3.6^\\circ$, $32.6^\\circ$, and $46.7^\\circ$. Most of the tests\nwere made with the wings in combination with a fuselage.\nResults of the investigation indicate that the maximum lift coef-\nficient of the wing-fuselage combinations increased as the angle of\nsweepback increased. The usual effect of sweepback in reducing the\nlift-curve slope was confined to the lift-coefficient range between\nabout -0.2 and 0.2 but was less than expected, probably because the\nusual effect of sweepback was masked by a variable influence of the\nfuselage. As the sweepback was increased, there was a rearward shift of\nthe aerodynamic center which was greater than indicated by the theory.\nThis shift is believed to be caused by a large destabilizing effect of\nthe fuselage on the $3.6^\\circ$ sweptback wing, while tests showed practically\nno effect for the $46.7^\\circ$ sweptback wing.\nAt low lift coefficients the derivative of rolling moment caused by\nyaw varied linearly with lift coefficient, and the rate of variation was\nincreased with an increase of sweep angle in very much the same manner\nthat is predicted by theory. Because the linear variations were\nmaintained over only very small ranges of lift coefficients for the more\nhighly swept wings, the maximum positive values of the derivatives of\nrolling moment due to yaw for the $32.6^\\circ$ and $46.7^\\circ$ sweptback wings were\nsmaller than the values of this derivative for the $3.6^\\circ$ sweptback wing\nat lift coefficients greater than 0.6.\nCONFIDENTIAL", "timestamp": "2026-07-22T05:12:24.368942+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 45, "total_pages": 99, "image_filename": "19930082511_p45.jpg", "text": "NACA TN No. 1826\n\nThe function $\\phi_A$ may be expressed formally (see reference 13) as\n\n$$\n\\phi_A = \\sum_{m=1}^{\\infty} \\left[ \\sin m\\theta \\frac{1}{\\pi} \\int_0^{\\infty} \\frac{J_m(i\\rho q)}{i q J_m'(iq)} dq \\int_a^b \\varepsilon_m^{(1)}(\\beta) \\cos q(\\beta - \\xi) d\\beta \\right.\n$$\n$$\n\\left. + \\cos m\\theta \\frac{1}{\\pi} \\int_0^{\\infty} \\frac{J_m(i\\rho q)}{i q J_m'(iq)} dq \\int_a^b \\varepsilon_m^{(2)}(\\beta) \\cos q(\\beta - \\xi) d\\beta \\right]\n\\tag{26}\n$$\n\nwhere $\\varepsilon_m^{(1)}(\\xi)$ and $\\varepsilon_m^{(2)}(\\xi)$ are, respectively, the $m^{\\text{th}}$ sine and cosine coefficients of the Fourier series for $\\left. \\frac{\\partial \\phi_A}{\\partial \\rho} \\right|_{\\rho=1}$ and $q$ and $\\beta$ are variables of integration. The integrations over $\\beta$ would, in general, have the range $-\\infty$ to $+\\infty$; however, condition (2) shows that the functions $\\varepsilon_m^{(j)}(\\beta)$, $j = 1, 2$, are zero from $-\\infty$ to $a$ and from $b$ to $+\\infty$. The convergence of this function and its derivatives to the desired function $\\phi_A$ and its derivatives is discussed in the appendix of reference 13. A modification is necessary because of the discontinuity in $\\left. \\frac{\\partial \\phi_A}{\\partial \\rho} \\right|_{\\rho=1}$ that may exist at $\\xi = b$. For this case, the desired convergence may be proved for regions bounded away from the circle $\\rho = 1$, $\\xi = b$.\n\nThe assumption of zero perturbation velocity on the surface of the jet is equivalent to the assumption that the expansion of $-\\left. \\frac{\\partial (\\phi_0 + \\phi_c)}{\\partial \\xi} \\right|_{\\rho=1}$ in a Fourier series in $\\theta$ contains no term independent of $\\theta$. For this reason no $m = 0$ term appears in expansion (26). The next section discusses the somewhat special treatment that is required when the Fourier series contains a term independent of $\\theta$.\n\nEvaluation of $\\varepsilon_m^{(j)}(\\xi)$.— The function $\\phi_A$ given in the preceding equation satisfies conditions (1) and (2) regardless of the precise form of the functions $\\varepsilon_m^{(j)}(\\xi)$. It is now desired to find the functions $\\varepsilon_m^{(j)}(\\xi)$ for $a < \\xi < b$ such that $\\phi_A$ will satisfy conditions (3) and (4). To this end the functions are represented by infinite series of the form\n\n$$\n\\varepsilon_m^{(j)}(\\xi) = h_{m0}^{(j)} \\sin \\frac{\\pi}{2} \\frac{\\xi - a}{b - a} + \\sum_{n=1}^{\\infty} h_{mn}^{(j)} \\sin n\\pi \\frac{\\xi - a}{b - a}\n\\tag{27}\n$$", "timestamp": "2026-07-22T05:12:26.714517+00:00"}
{"citation_id": "19930082614", "source_url": "https://ntrs.nasa.gov/api/citations/19930082614/downloads/19930082614.pdf", "page_number": 30, "total_pages": 36, "image_filename": "19930082614_p30.jpg", "text": "```markdown\n28\nNACA TN 1939\n\nTypical Section\nPlan View\nFront View\n\n<!-- Image (105, 105, 838, 828) -->\n\n| Brake | Chordwise location | | Brake flap chord | |\n| :--- | :--- | :--- | :--- | :--- |\n| | $d_u$ | $d_l$ | $c_{B_u}$ | $c_{B_l}$ |\n| A | $0.80c$ | $0.80c$ | $0.200c$ | $0.200c$ |\n| B | $.80c$ | $.80c$ | $.200c$ | $.200c$ |\n| C | $.64c$ | $.64c$ | $.117c$ | $.117c$ |\n| D | $.86c$ | $.88c$ | $.144c$ | $.125c$ |\n| E | — | $.76c$ | — | $.240c$ |\n\n(a) Split-flap-type air brakes.\n\nFigure 6.— Aerodynamic brake installations for which drag characteristics are summarized in Table I.\n```", "timestamp": "2026-07-22T05:12:28.860534+00:00"}
{"citation_id": "19930085881", "source_url": "https://ntrs.nasa.gov/api/citations/19930085881/downloads/19930085881.pdf", "page_number": 4, "total_pages": 31, "image_filename": "19930085881_p4.jpg", "text": "```markdown\n2\nCONFIDENTIAL\nNACA RM L9D12\n\ntaper ratio 1.0 having $0^\\circ$ and $45^\\circ$ sweepback. The purpose of this paper\nis to present these additional results.\n\nThe tests were made by means of the technique described in reference 1\nwhich permits the evaluation of the wing-aileron rolling effectiveness\ncontinuously over the Mach number range from about 0.6 to 1.9 at relatively\nlarge scale. The variation of total drag coefficient with Mach number\nwas also obtained.\n\nSYMBOLS\n\n| Symbol | Definition |\n| :--- | :--- |\n| $\\frac{pb}{2V}$ | wing-tip helix angle, radians |\n| p | rolling velocity, radians per second |\n| b | diameter of circle swept by wing tips, feet |\n| V | flight-path velocity, feet per second |\n| $C_D$ | total drag coefficient based on exposed wing area (1.563 sq ft) |\n| M | Mach number |\n| R | Reynolds number based on wing chord (0.59 ft) |\n| A | aspect ratio $\\left(\\frac{b}{c}\\right)$ |\n| c | chord of wing parallel to model center line, feet |\n| $\\delta_a$ | deflection of each aileron measured in plane normal to chord plane and parallel to fuselage center line, degrees (see fig. 1(b)) |\n| $i_w$ | wing incidence measured in plane of $\\delta_a$, degrees (see fig. 1(b)) |\n\nTEST VEHICLES AND TESTS\n\nThe general arrangement of the test vehicles is shown in figures 1\nand 2. Additional pertinent information is contained in tables I and II.\n\nThe test vehicles, which were relatively simple, inexpensive, and\nexpendable, consisted of a pointed cylindrical wooden body to which the\nparticular wing-aileron configuration under investigation was attached\nin a three-panel arrangement. Unpublished tests of three- and four-panel\narrangements indicate that, with regard to the rolling-effectiveness\n\nCONFIDENTIAL\n```", "timestamp": "2026-07-22T05:12:29.057602+00:00"}
{"citation_id": "19930082914", "source_url": "https://ntrs.nasa.gov/api/citations/19930082914/downloads/19930082914.pdf", "page_number": 32, "total_pages": 66, "image_filename": "19930082914_p32.jpg", "text": "NACA TN No. 1857\n31\n\n[Figure: Supersonic nozzles]\n\nFigure 1.- Supersonic nozzles.\n\nNACA\nL-55229", "timestamp": "2026-07-22T05:12:34.788377+00:00"}
{"citation_id": "19930085519", "source_url": "https://ntrs.nasa.gov/api/citations/19930085519/downloads/19930085519.pdf", "page_number": 27, "total_pages": 46, "image_filename": "19930085519_p27.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:12:37.442829+00:00"}
{"citation_id": "19930085487", "source_url": "https://ntrs.nasa.gov/api/citations/19930085487/downloads/19930085487.pdf", "page_number": 30, "total_pages": 36, "image_filename": "19930085487_p30.jpg", "text": "28\nNACA RM No. E8J22\n\n<!-- Image (111, 110, 881, 904) -->\n\nFigure 9. - Continued. Critical-speed diagrams for 10 stages of\ncompressor rotor.", "timestamp": "2026-07-22T05:12:39.770058+00:00"}
{"citation_id": "19930082485", "source_url": "https://ntrs.nasa.gov/api/citations/19930082485/downloads/19930082485.pdf", "page_number": 51, "total_pages": 62, "image_filename": "19930082485_p51.jpg", "text": "50\nNACA TN No. 1810\n\n[Figure: Diagram showing two camber lines with labeled parameters: Pitch, T; Chord; angles $\\alpha_1$, $c_e$; velocities $v_1$, $v_e$; and ratio $\\left(\\frac{V}{V_{cr}}\\right)_e$. The NACA logo is present at the bottom right.]\n\nFigure 19.- Assumed camber lines.", "timestamp": "2026-07-22T05:12:42.185355+00:00"}