PDF/ocr_dataset_1006.jsonl

{"citation_id": "19930085952", "source_url": "https://ntrs.nasa.gov/api/citations/19930085952/downloads/19930085952.pdf", "page_number": 42, "total_pages": 50, "image_filename": "19930085952_p42.jpg", "text": "NACA RM L9C24\n41\n\n$C_{ha}$\n.1\n0\n-.1\n-.2\n4.2\n3.8\n\n$C_L$\n3.4\n3.0\n2.6\n\n$C_m$\n.1\n0\n-.1\n-.2\n-.3\n\n$\\circ$ Articulated propellers; V/nD, 0.37\n$\\square$ Rigid propellers; V/nD, 0.29\n\n-4.8 -4.0 -3.2 -2.4 -1.6 -8 0\n$\\delta_a$, deg\n\n(b) $\\alpha \\approx 46^\\circ$.\nFigure 20.— Continued.", "timestamp": "2026-07-22T04:21:17.842427+00:00"}
{"citation_id": "19930086078", "source_url": "https://ntrs.nasa.gov/api/citations/19930086078/downloads/19930086078.pdf", "page_number": 11, "total_pages": 42, "image_filename": "19930086078_p11.jpg", "text": "NACA RM L9H04 CONFIDENTIAL 9\n\nincrease with increasing $\\alpha$ even at the large aileron deflections, probably because as the wing continued to load up, mutual interference between the wing and the aileron (induction effects) tended to increase further the loading on the wing.\n\nExcept at some negative values of $\\alpha$, an increase of aileron extension at constant aileron deflection caused an increase in rolling moment for all angles of attack and for all aileron configurations (figs. 7, 9, and 11).\n\nSeveral tests performed with the fully extended short-chord aileron moved forward on the wing so the aileron midchord would coincide with the $0.267c$ line of the unswept wing showed that the aileron produced approximately the same lateral control characteristics on the wing as when in the normal position investigated (fig. 10).\n\nThe short-chord aileron-wing configuration reported herein was geometrically comparable to the extensible wing-tip aileron described in reference 10, but comparison between the results of the two investigations was available for only $\\alpha = 0^\\circ$. Although larger values of $C_l$ were obtained for various aileron extensions in the present case, both investigations showed the same general variation of rolling moment with aileron extension.\n\nComparing the three ailerons of this investigation on the basis of equal values of $S_a/S$, it can be noted that for similar aileron deflections each of the ailerons on the unswept wing configuration generally produced about the same amount of rolling moment, except at high angles of attack where the short-chord aileron did not produce rolling moments as great as those produced by the large-chord or triangular ailerons (figs. 7, 9, and 11).\n\nThe yawing moments produced by each of the aileron configurations were generally adverse over the entire $\\alpha$ range and became more adverse with increase in $\\alpha$, aileron deflection, and/or aileron extension. The adverse $C_n/C_l$ ratio was large for all aileron configurations at large angles of attack, but was largest for the short-chord aileron, being larger than 0.25.\n\n45° sweptback wing.— As was the case of the unswept wing, the rolling moments produced by the various ailerons on the 45° swept wing generally increased with increase in $\\alpha$ (figs. 12 to 17). The rolling-moment data show a reversal of roll direction at some negative angles of attack, as did the aileron on the unswept wing.\n\nDeflecting each of the ailerons effected fairly linear increases in rolling moments at all positive angles of attack for the aileron-deflection range investigated (figs. 12, 14, and 16). Unit deflection-\n\nCONFIDENTIAL", "timestamp": "2026-07-22T04:21:21.810946+00:00"}
{"citation_id": "19930085842", "source_url": "https://ntrs.nasa.gov/api/citations/19930085842/downloads/19930085842.pdf", "page_number": 56, "total_pages": 104, "image_filename": "19930085842_p56.jpg", "text": "52\nNACA RM L9029\n\n[Figure: Graph plotting Allavator deflection, $\\delta(C_{m0}=0)$, deg (y-axis) against Lift coefficient, $C_L$ (x-axis). The y-axis ranges from -32 to 8. The x-axis ranges from -2 to 8. A curve starts near (0, 0) and trends downwards to approximately (7, -24). The graph includes a stamp: \"NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\".]\n\nFigure 23.- Variation of allavator deflection required for trim with lift coefficient. Complete model configuration; propellers removed.", "timestamp": "2026-07-22T04:21:22.193131+00:00"}
{"citation_id": "19930086060", "source_url": "https://ntrs.nasa.gov/api/citations/19930086060/downloads/19930086060.pdf", "page_number": 13, "total_pages": 20, "image_filename": "19930086060_p13.jpg", "text": "CONFIDENTIAL\n\nConfiguration\n9\n10\n11\n12\n\nGeneral body-\nprofile equations\n$$\n\\begin{cases}\nY_1 = 3.75 - a(KL - X)^2 & [\\text{from } X = 0 \\text{ to } X = KL] \\\\\nY_2 = 3.75 - b(X - KL)^2 & [\\text{from } X = KL \\text{ to } X = L = 45.32]\n\\end{cases}\n$$\n\n| Configuration | K | $a \\times 10^3$ | $b \\times 10^3$ |\n| :--- | :--- | :--- | :--- |\n| 9 | .2 | 45.64 | 1.61 |\n| 10 | .4 | 11.41 | 2.85 |\n| 11 | .6 | 5.07 | 6.42 |\n| 12 | .8 | 2.85 | 25.68 |\n\n[Figure: NACA logo]\n\nFigure 3.- Profiles and general equations for test configurations. Configuration 12 is shown as reference for theoretical calculations.\n\nCONFIDENTIAL\n\nNACA RM L59E02\n11", "timestamp": "2026-07-22T04:21:34.147476+00:00"}
{"citation_id": "19930085965", "source_url": "https://ntrs.nasa.gov/api/citations/19930085965/downloads/19930085965.pdf", "page_number": 29, "total_pages": 67, "image_filename": "19930085965_p29.jpg", "text": "28\nNACA RM E9E06\n\n1. Heating of axial-flow-compressor blades by eddy currents appeared to be applicable for protection against icing.\n\n2. The surface average heat requirement for the inlet guide vanes in conditions of icing of a jet engine of the 4000-pound-thrust class was calculated to be maximum at rated-speed conditions and equal to 5400 Btu per hour per square foot or approximately 11 watts per square inch. Creation of heat densities of this magnitude for ice protection of the inlet guide vanes of a compressor appeared possible.\n\n3. The shaft-power requirement for the generation of eddy currents to produce the heat needed for protection of the inlet guide vanes is small, and for the proposed design is only 0.20 percent of the turbine load. Compared to this power, the electrical power required for the magnetizing field coils, however, is only of the order of 1 percent of the shaft power required for generation of the eddy currents or 0.002 percent of the turbine load.\n\n4. In an alternating-flux eddy-current system at heat densities necessary for ice protection, the eddy-current power generated is proportional to the square roots of frequency, material resistivity, and saturation flux density, and to the 3/2 power of the maximum field intensity at the vane surface. The blade material chosen should therefore have the highest resistivity and the greatest saturation flux density.\n\n5. In a pulsating-unidirectional flux eddy-current system in which the flux varies from zero to a maximum value and which can be simulated by superimposing a direct-current flux on a sine-wave alternating-current flux, the heat generated does not change appreciably with application of the direct-current flux. Heat calculations can therefore be made on the basis of the alternating-current component. In the production of the pulsating flux by means of a chopper and the application of a direct-current magnetomotive force, however, the required direct-current ampere turns will be greater than the root-mean-square ampere turns of the alternating-current component by a factor of $2\\sqrt{2}$.\n\n6. The vane surfaces will probably operate well past the knee of the magnetization curve of the blade material, and the magnetomotive-force drop along the blade will be great in comparison with air gaps of practical sizes, which may lead to difficulty in obtaining sufficient chopping action at high frequencies.", "timestamp": "2026-07-22T04:21:34.881410+00:00"}
{"citation_id": "19930086003", "source_url": "https://ntrs.nasa.gov/api/citations/19930086003/downloads/19930086003.pdf", "page_number": 22, "total_pages": 36, "image_filename": "19930086003_p22.jpg", "text": "```markdown\n20\n\nCONFIDENTIAL\n\n1.0 x 10^6\n\nReynolds number, R\n.8\n.6\n.4\n\nMean\n\n.6 .7 .8 .9 1.0 1.1 1.2\nMach number, M\n\nCONFIDENTIAL\n\nNACA\n\nFigure 7.- Variation of test Reynolds number with Mach number for a model with 45° sweptback wing,\naspect ratio 6, taper ratio 0.6, and NACA 65A006 airfoil section.\n\nNACA RM L9108\n```", "timestamp": "2026-07-22T04:21:35.413989+00:00"}
{"citation_id": "19930085982", "source_url": "https://ntrs.nasa.gov/api/citations/19930085982/downloads/19930085982.pdf", "page_number": 26, "total_pages": 32, "image_filename": "19930085982_p26.jpg", "text": "```markdown\n24\n\nTotal-pressure ratio, $P_2/P_1$\nAdiabatic temperature-rise efficiency\n\nCorrected rotor speed $N/\\sqrt{\\theta}$, rpm\no 9,030\n□ 13,545\n△ 18,060\n\nCorrected weight flow, $W\\sqrt{\\theta}/\\delta$, lb/sec\n\n<!-- Image (59, 178, 885, 738) -->\n\nFigure 4. - Over-all performance of typical inlet stage with NACA 65-(2)10 blade section.\n\nNACA RM E9E13\n```", "timestamp": "2026-07-22T04:21:39.415776+00:00"}
{"citation_id": "19930086081", "source_url": "https://ntrs.nasa.gov/api/citations/19930086081/downloads/19930086081.pdf", "page_number": 9, "total_pages": 44, "image_filename": "19930086081_p9.jpg", "text": "NACA RM L9H05 CONFIDENTIAL 7\n\n| Wing (figs. 4 to 9) | | Control surface (initial series of tests) (figs. 13 and 14) | | Control surface (second series of tests) (figs. 15 and 16) | |\n| :--- | :--- | :--- | :--- | :--- | :--- |\n| **Variable** | **Error** | **Variable** | **Error** | **Variable** | **Error** |\n| $\\alpha$ | $\\pm 0.05^\\circ$ | $\\alpha$ | $\\pm 0.05^\\circ$ | $\\alpha$ | $\\pm 0.05^\\circ$ |\n| $\\delta$ | $\\pm .2^\\circ$ | $\\delta$ | $\\pm .3^\\circ$ | $\\delta$ | $\\pm .2^\\circ$ |\n| $C_L$ | $\\pm .003$ | $C_L$ | $\\pm .001$ | $C_{N_F}$ | $\\pm .005$ |\n| $C_D$ | $\\pm .001$ | $C_D$ | $\\pm .001$ | $C_{C_F}$ | $\\pm .010$ |\n| $C_m$ | $\\pm .001$ | $C_m$ | $\\pm .001$ | $C_{M_F}$ | $\\pm .008$ |\n| $C_l$ | $\\pm .0004$ | $C_l$ | $\\pm .0002$ | $C_{BM_F}$ | $\\pm .015$ |\n| $C_n$ | $\\pm .0003$ | $C_n$ | $\\pm .0001$ | | |\n\nIt should be noted that different geometric parameters and different axes were used in reducing the data for the two series of control-surface tests. The electrical system of the balance was arranged to permit direct moment measurements about the wing axes in the first series of tests and about the control axes in the second series. This technique was found necessary to avoid the introduction of considerable scatter in the moment data which appeared when an attempt was made to transfer the data to axes far distant from the point of measurement. From one model set-up to another (change in fence, fuselage, or control-surface thickness) the angle of attack could have differed by $\\pm 0.1^\\circ$, the control deflection could have differed by $\\pm 0.4^\\circ$, and the fuselage incidence with respect to the wing could have varied by $\\pm 0.3^\\circ$. Repeat tests were made for each configuration to assess the magnitude of errors. Static calibration indicated no measurable change in control-surface deflection caused by control-surface loading.\n\nRESULTS\n\nFigures 4 to 9 present test data of the complete wing as plots of the aerodynamic coefficients plotted against angle of attack for each of the various deflection angles of the tip control surface. Figure 10 presents cross plots of these data in which the coefficients are plotted against control deflection at zero angle of attack. Fuselage incidence is believed to have caused the displacement in the curves of figure 8 at zero deflection.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T04:21:40.244250+00:00"}
{"citation_id": "19930085936", "source_url": "https://ntrs.nasa.gov/api/citations/19930085936/downloads/19930085936.pdf", "page_number": 45, "total_pages": 59, "image_filename": "19930085936_p45.jpg", "text": "44\nNACA RM No. E9B03\n\nPressure coefficient, $C_p$\n\n| Angle of yaw (deg) | |\n| :--- | :--- |\n| $\\square$ | -12 |\n| $\\circ$ | -6 |\n| $\\diamond$ | 0 |\n| $\\triangle$ | 6 |\n| $\\nabla$ | 12 |\n\nDistance from tip, $x/L$\n\n(c) $\\theta = 180^\\circ$ longitudinal plane.\n\nFigure 8. - Continued. Pressure distributions along longitudinal planes at $10^\\circ$ angle of attack for range of yaw angles.", "timestamp": "2026-07-22T04:21:40.978888+00:00"}
{"citation_id": "19930082090", "source_url": "https://ntrs.nasa.gov/api/citations/19930082090/downloads/19930082090.pdf", "page_number": 107, "total_pages": 122, "image_filename": "19930082090_p107.jpg", "text": "NACA TN No. 1455\n105\n\nVentilating air\n5.0\"\nI.D.\nApprox. 10'\n30\"\n7.83\"\nI.D.\nExhaust gas\n22\"\n26\"\n7.83\"\nI.D.\nExhaust\ngas\n13.5\"\n17.5\"\n40\"\n60\"\n5.0\"\nI.D.\nVentilating air\n\nx Temperature traverse\no Static-pressure tap\n\nNACA\n\nFigure 54.- Schematic diagram of test setup of heat exchanger N and air\nshroud, showing location of static-pressure and temperature measuring\nstations.", "timestamp": "2026-07-22T04:21:41.284041+00:00"}
{"citation_id": "19930086015", "source_url": "https://ntrs.nasa.gov/api/citations/19930086015/downloads/19930086015.pdf", "page_number": 28, "total_pages": 54, "image_filename": "19930086015_p28.jpg", "text": "NACA RM A9E24\nCONFIDENTIAL\n\nVertical distance from tunnel center line, z, in.\n24\n16\n8\n0\n-8\n-16\n-24\n\nFlagged symbols\ndenote survey\noff center line\n\nx=-24\nx=0\nx=21\n\n-0.04 0 .04\n-0.08 -0.04 0 .04 .08\n-0.08 -0.04 0 .04\n\nStream pressure coefficient, $\\Delta P/q$\n\n(c) D=129.32; M=1.43. $H_0 = 0.688$\n$17_c = 0.007$\n\nFigure 7.— Continued.\n\nCONFIDENTIAL\n27", "timestamp": "2026-07-22T04:21:41.541352+00:00"}
{"citation_id": "19930086151", "source_url": "https://ntrs.nasa.gov/api/citations/19930086151/downloads/19930086151.pdf", "page_number": 4, "total_pages": 34, "image_filename": "19930086151_p4.jpg", "text": "2\nCONFIDENTIAL\nNACA RM L9J28\n\nthe wing and is deflected about a spanwise hinge axis approximately\nnormal to the plane of symmetry to produce rolling moment. The ailerons\nare, of course, deflected oppositely on each semispan of a complete wing\nin a manner similar to conventional ailerons.\n\nPrevious investigations of wing-tip ailerons deflected from a free-\nfloating position have been made on more conventional (unswept) wings,\nand have shown adequate lateral control obtainable with this type of\naileron (references 1 to 4). The results of a preliminary investigation\nof a triangular wing-tip aileron deflected $30^\\circ$ at an angle of attack\nof $0^\\circ$ on a $42^\\circ$ sweptback wing showed that this control surface provided\nlarge rolling moments at both subsonic and transonic speeds (refer-\nence 5). In addition, data obtained in an investigation of various\nextensible-type wing-tip ailerons at several small deflections on a\n$45^\\circ$ sweptback wing showed that a deflectable wing-tip aileron offered\npromise of providing large rolling moments on a sweptback wing\n(reference 6).\n\nThe present investigation on an untapered $45^\\circ$ sweptback semispan\nwing was performed in the Langley 300 MPH 7- by 10-foot tunnel in order\nto determine the lateral control characteristics of deflectable-type\nwing-tip ailerons on a sweptback wing. A parallelogram- and a triangular-\nplan-form wing-tip aileron having flat-plate profiles and equal areas\nwere investigated on the wing model through a large wing-angle-of-attack\nrange and at aileron deflections up to $30^\\circ$. These ailerons were\ninvestigated with and without a large end plate (simulating a vertical\nfin) mounted on the wing inboard of the aileron in order to determine\nthe effect of the end plate on both the plain-wing and aileron\ncharacteristics.\n\nSYMBOLS\n\nInasmuch as the span of the wing equipped with the parallelogram\nand triangular ailerons differed appreciably (fig. 1), all data presented\nare based on the dimensions of each complete-wing configuration.\n\nThe forces and moments measured on the wings are presented about\nthe wind axes, which, for the conditions of these tests (zero yaw),\ncorrespond to the stability axes. The X-axis is in the plane of symmetry\nof the models and is parallel to the tunnel free-stream air flow. The\nZ-axis is in the plane of symmetry of the models and is perpendicular to\nthe X-axis. The Y-axis is mutually perpendicular to the X-axis and\nZ-axis. All three axes intersect at the intersection of the chord plane\nand the 25-percent station of the mean aerodynamic chord at the root of\nthe models (fig. 1).\n\nCONFIDENTIAL", "timestamp": "2026-07-22T04:21:42.452048+00:00"}
{"citation_id": "19930085551", "source_url": "https://ntrs.nasa.gov/api/citations/19930085551/downloads/19930085551.pdf", "page_number": 74, "total_pages": 82, "image_filename": "19930085551_p74.jpg", "text": "```markdown\nNACA RM No. L8K30\n73\n\n[Figure: Time history of airplane motions after loss of power on No. 1 engine in the cruise condition. C-54D airplane; clean power for level flight before No. 1 engine cut to idling at start of record; corrective control started at 10.5 seconds.]\n\nFigure 19.-- Time history of airplane motions after loss of power on\nNo. 1 engine in the cruise condition. C-54D airplane; clean power\nfor level flight before No. 1 engine cut to idling at start of record;\ncorrective control started at 10.5 seconds.\n```", "timestamp": "2026-07-22T04:21:42.705992+00:00"}
{"citation_id": "19930090382", "source_url": "https://ntrs.nasa.gov/api/citations/19930090382/downloads/19930090382.pdf", "page_number": 2, "total_pages": 37, "image_filename": "19930090382_p2.jpg", "text": "NACA RM L9I07\nCONFIDENTIAL\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\nRESEARCH MEMORANDUM\nINVESTIGATION OF THE NACA 4-(4)(06)-04 TWO-BLADE\nPROPELLER AT FORWARD MACH NUMBERS TO 0.925\nBy James B. Delano and Daniel E. Harrison\nSUMMARY\nInvestigations of the NACA 4-(4)(06)-04 two-blade propeller (design\nthickness ratio, 0.06 at 0.7-radius station) have been made in the\nLangley 8-foot high-speed tunnel for blade angles from $20^{\\circ}$ to $70^{\\circ}$ for\nforward Mach numbers up to 0.925.\nIn general, the effects of compressibility on maximum efficiency are\nsimilar to those reported earlier for the NACA 4-(5)(08)-03 propeller.\nThe envelope efficiency for the NACA 4-(4)(06)-04 propeller is 10 percent\nhigher than that for the NACA 4(5)(08)-03 propeller at a forward Mach\nnumber of 0.8.\nINTRODUCTION\nResults of the first two phases of a general investigation to study\nthe effects of compressibility, design camber, blade sweep, thickness\nratio, and dual rotation on propeller performance at transonic speeds\nwere presented in references 1 and 2. The effects of blade sweep on\npropeller performance constitute the next phase of this general\ninvestigation.\nIn the course of the investigation to study the effect of blade\nsweep on propeller performance a thin (6-percent thickness ratio)\nstraight blade was tested to provide a comparison of results for swept\nand unswept blades. It was believed that the results obtained for this\nthin propeller would be of immediate interest and consequently are\npresented herein. The investigation was made for a range of blade angle\nfrom $20^{\\circ}$ to $70^{\\circ}$ for a forward Mach number range from 0.175 to 0.925.\nCONFIDENTIAL", "timestamp": "2026-07-22T04:21:43.092536+00:00"}
{"citation_id": "19930086022", "source_url": "https://ntrs.nasa.gov/api/citations/19930086022/downloads/19930086022.pdf", "page_number": 23, "total_pages": 34, "image_filename": "19930086022_p23.jpg", "text": "NACA RM L9E24\n21\n\n<!-- Image (279, 109, 781, 882) -->\n\n(b) $P_R$ and $C_{ha}$ plotted against $\\alpha$.\nFigure 5.- Continued.", "timestamp": "2026-07-22T04:21:51.672116+00:00"}
{"citation_id": "19930085958", "source_url": "https://ntrs.nasa.gov/api/citations/19930085958/downloads/19930085958.pdf", "page_number": 34, "total_pages": 43, "image_filename": "19930085958_p34.jpg", "text": "NACA RM No. L9B11\n33\n\n<!-- Image (153, 115, 876, 826) -->\n\nFigure 14.- Effects of wing-fuselage position on the characteristics of a 42° sweptback wing with 0.70b extensible leading-edge flaps, split flaps, and upper-surface fences.", "timestamp": "2026-07-22T04:21:55.098258+00:00"}
{"citation_id": "19930086078", "source_url": "https://ntrs.nasa.gov/api/citations/19930086078/downloads/19930086078.pdf", "page_number": 12, "total_pages": 42, "image_filename": "19930086078_p12.jpg", "text": "10 CONFIDENTIAL NACA RM L9H04\n\nof the short-chord aileron produced larger incremental roll at low angles of attack than at high angles of attack (fig. 16). This phenomenon concerning the short-chord aileron was also noted on the unswept wing configuration and was attributed earlier in the paper to separation of flow over the aileron.\n\nExcept at some negative values of $\\alpha$, an increase of aileron extension at constant deflection for each aileron plan form caused an increase in rolling moment for all angles of attack (figs. 13, 15, and 17).\n\nOn the basis of equal values of $S_a/S$ and at the same deflection, the short-chord aileron produced rolling moments over the entire $\\alpha$ range comparable to the rolling moments produced by the large-chord and triangular ailerons (figs. 13, 15, and 17).\n\nAdverse yawing moments were produced by the ailerons for all positive angles of attack, generally becoming more severe with increased $\\alpha$ (figs. 12 to 17). Yawing moments also became more adverse with an increase of aileron deflection or extension. The adverse $C_n/C_l$ ratio amounted to as much as 0.6 for some aileron configurations at large angles of attack near $C_{l_{max}}$.\n\nComparison of the unswept and sweptback wing configurations.-\n\nComparison of the lateral-control data for the unswept and sweptback wing configurations shows that the rolling moments produced by each of the ailerons generally exhibited similar variations with change in angle of attack, aileron extension, and/or aileron deflection.\n\nFor any given value of $C_l$ the triangular aileron on the unswept wing configuration generally produced greater rolling moments than the corresponding triangular aileron on the sweptback configuration, whereas the large-chord and short-chord ailerons on the unswept wing configuration generally produced smaller rolling moments than the corresponding ailerons on the sweptback configurations. However, because the damping-in-roll coefficient $C_{l_p}$ is smaller for the sweptback wing than for the unswept wing (primarily because of the smaller aspect ratio of the sweptback wing (reference 11)), the values of the wing-tip helix angle $pb/2V$ produced by each of the ailerons on the sweptback wing configuration were considerably greater than the values of $pb/2V$ produced by the respective ailerons on the unswept wing configuration.\n\nIn order to compare the rolling effectiveness of the various ailerons, the variation of wing-tip helix angle $pb/2V$ with lift coefficient, estimated for the unswept and sweptback wing configurations, is given in figures 18 and 19, respectively. Values of the damping-in-roll coefficient $C_{l_p}$ used in computing $pb/2V$ were 0.27 and 0.13 for\n\nCONFIDENTIAL", "timestamp": "2026-07-22T04:22:08.876601+00:00"}
{"citation_id": "19930085965", "source_url": "https://ntrs.nasa.gov/api/citations/19930085965/downloads/19930085965.pdf", "page_number": 30, "total_pages": 67, "image_filename": "19930085965_p30.jpg", "text": "NACA RM E9E06\n29\n\n7. Provided that the blade is sufficiently thick to insure a negligible flux at the center, the heat generated per unit surface is independent of vane thickness or the total heat generated is a function of only the vane surface.\n\n8. The disadvantages anticipated in the application of eddy-current heating would be the necessity for including magnetic material in the flux circuit and the corresponding weight increase of these materials and of the coil and the chopper.\n\nLewis Flight Propulsion Laboratory,\nNational Advisory Committee for Aeronautics,\nCleveland, Ohio.", "timestamp": "2026-07-22T04:22:14.401538+00:00"}
{"citation_id": "19930085529", "source_url": "https://ntrs.nasa.gov/api/citations/19930085529/downloads/19930085529.pdf", "page_number": 79, "total_pages": 85, "image_filename": "19930085529_p79.jpg", "text": "```markdown\n78\nNACA RM No. L8A30a\n\nTABLE T2\n$$[\\lambda = -30^\\circ, \\delta_{a_L} = 10.0^\\circ, \\alpha = 7^\\circ]$$\nCONFIDENTIAL\n\n| | | UPPER SURFACE | | | | | | | | LOWER SURFACE | | | | | |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| | | | **Mach Number** | | | | | | | | | **Mach Number** | | | |\n| **Tube** | **Per-cent chord** | **0.60** | **0.80** | **0.85** | **0.89** | **0.925** | **0.96** | **Tube** | **Per-cent chord** | **0.60** | **0.80** | **0.85** | **0.89** | **0.925** | **0.96** |\n| A 1 | 2.0 | - | - | - | - | - | - | 86 | 3.0 | - | - | - | - | - | - |\n| 2 | 6.0 | - | - | - | - | - | - | 87 | 10.0 | - | - | - | - | - | - |\n| 3 | 15.0 | - | - | - | - | - | - | 88 | 25.0 | - | - | - | - | - | - |\n| 4 | 27.5 | - | - | - | - | - | - | 89 | 41.0 | - | - | - | - | - | - |\n| 5 | 40.0 | - | - | - | - | - | - | 90 | 52.5 | -0.002 | -0.088 | -0.124 | -0.161 | -0.206 | -0.277 |\n| 6 | 50.0 | -0.368 | -0.561 | -0.569 | -0.560 | - | -0.618 | 91 | 62.5 | -0.011 | -0.081 | -0.124 | -0.159 | -0.196 | -0.238 |\n| 7 | 59.0 | -0.439 | -0.590 | -0.601 | -0.601 | -0.611 | -0.683 | 92 | 72.5 | -0.027 | -0.095 | -0.117 | -0.141 | -0.140 | -0.123 |\n| 8 | 67.5 | -0.361 | -0.537 | -0.577 | -0.595 | -0.611 | -0.584 | 93 | 84.0 | - | - | - | - | - | - |\n| 9 | 77.5 | - | - | - | - | - | - | 94 | 94.0 | - | - | - | - | - | - |\n| 10 | 87.5 | - | - | - | - | - | - | | | | | | | | |\n| 11 | 96.0 | - | - | - | - | - | - | | | | | | | | |\n| | | | | | | | | | | | | | | | |\n| 812 | 2.0 | -0.651 | -0.943 | -0.957 | -0.952 | -0.981 | -0.979 | 95 | 3.0 | -0.739 | -0.755 | -0.739 | -0.709 | -0.646 | -0.708 |\n| 13 | 6.0 | -0.670 | -0.997 | -0.971 | -0.971 | -0.607 | -0.807 | 96 | 10.0 | -0.399 | -0.458 | -0.397 | -0.375 | -0.360 | -0.387 |\n| 14 | 15.0 | -0.677 | -0.940 | -0.906 | -0.947 | -0.968 | -0.847 | 97 | 25.0 | -0.110 | -0.098 | -0.071 | -0.056 | -0.036 | -0.062 |\n| 15 | 27.5 | -0.720 | -0.939 | -0.987 | -0.961 | -0.971 | -0.860 | 98 | 41.0 | -0.021 | -0.054 | -0.080 | -0.109 | -0.136 | -0.136 |\n| 16 | 40.0 | -0.700 | -0.609 | -0.609 | -0.571 | -0.581 | -0.801 | 99 | 52.5 | -0.051 | -0.096 | -0.136 | -0.179 | -0.216 | -0.227 |\n| 17 | 50.0 | -0.656 | -0.616 | -0.618 | -0.577 | -0.588 | -0.818 | 100 | 62.5 | -0.041 | -0.081 | -0.122 | -0.151 | -0.197 | -0.232 |\n| 18 | 59.0 | -0.546 | -0.611 | -0.625 | -0.593 | -0.594 | -0.786 | 101 | 72.5 | -0.006 | -0.047 | -0.075 | -0.095 | -0.113 | -0.125 |\n| 19 | 67.5 | -0.493 | -0.586 | -0.609 | -0.592 | -0.599 | -0.708 | 102 | 86.3 | -0.005 | -0.029 | -0.040 | -0.046 | -0.049 | -0.049 |\n| 20 | 77.5 | -0.344 | -0.517 | -0.567 | -0.581 | -0.595 | -0.694 | 103 | 94.5 | -0.028 | -0.051 | -0.061 | -0.081 | -0.074 | -0.074 |\n| 21 | 88.0 | -0.287 | -0.448 | -0.448 | -0.466 | -0.505 | -0.626 | | | | | | | | |\n| 22 | 95.3 | - | - | - | - | - | - | | | | | | | | |\n| | | | | | | | | | | | | | | | |\n| C23 | 2.0 | -1.304 | -1.419 | -1.249 | -1.096 | -0.889 | -0.790 | 104 | 3.0 | -0.721 | -0.726 | -0.706 | -0.689 | -0.681 | -0.691 |\n| 24 | 6.0 | -1.040 | -1.357 | -1.161 | -1.065 | -0.939 | -0.813 | 105 | 10.0 | -0.456 | -0.439 | -0.425 | -0.411 | -0.406 | -0.425 |\n| 25 | 15.0 | -0.969 | -1.079 | -1.066 | -0.993 | -0.923 | -0.812 | 106 | 25.0 | -0.169 | -0.177 | -0.157 | -0.141 | -0.136 | -0.163 |\n| 26 | 27.5 | -0.681 | -0.739 | -0.774 | -0.944 | -0.849 | -0.849 | 107 | 41.0 | -0.098 | -0.016 | -0.012 | -0.037 | -0.054 | -0.010 |\n| 27 | 40.0 | -0.566 | -0.662 | -0.610 | -0.745 | -0.854 | -0.803 | 108 | 52.5 | -0.009 | -0.045 | -0.061 | -0.120 | -0.154 | -0.134 |\n| 28 | 50.0 | -0.504 | -0.491 | -0.493 | -0.708 | -0.864 | -0.864 | 109 | 62.5 | -0.011 | -0.045 | -0.061 | -0.121 | -0.164 | -0.195 |\n| 29 | 59.0 | -0.429 | -0.428 | -0.464 | -0.467 | -0.648 | -0.873 | 110 | 72.5 | -0.017 | -0.006 | -0.036 | -0.070 | -0.106 | -0.110 |\n| 30 | 67.5 | -0.325 | -0.360 | -0.391 | -0.370 | -0.490 | -0.835 | 111 | 85.1 | -0.047 | -0.000 | -0.015 | -0.015 | -0.043 | -0.060 |\n| 31 | 77.5 | -0.233 | -0.325 | -0.372 | -0.365 | -0.424 | -0.813 | 112 | 94.6 | -0.066 | -0.059 | -0.012 | -0.009 | -0.005 | -0.002 |\n| 32 | 88.0 | -0.182 | -0.263 | -0.312 | -0.325 | -0.390 | -0.680 | | | | | | | | |\n| 33 | 95.3 | - | - | - | - | - | - | | | | | | | | |\n| | | | | | | | | | | | | | | | |\n| D34 | 2.0 | -1.006 | -1.443 | -1.182 | -0.981 | -0.825 | -0.698 | 113 | 3.0 | -0.713 | -0.698 | -0.679 | -0.663 | -0.654 | -0.665 |\n| 35 | 15.0 | -1.007 | -1.115 | -1.154 | -1.006 | -0.895 | -0.788 | 114 | 10.0 | -0.456 | -0.444 | -0.427 | -0.415 | -0.411 | -0.430 |\n| 36 | 27.5 | -0.672 | -1.028 | -1.085 | -1.018 | -0.925 | -0.831 | 115 | 25.0 | -0.185 | -0.195 | -0.177 | -0.162 | -0.158 | -0.183 |\n| 37 | 40.0 | -0.567 | -0.661 | -0.610 | -0.744 | -0.854 | -0.804 | 116 | 41.0 | -0.098 | -0.016 | -0.012 | -0.037 | -0.054 | -0.010 |\n| 38 | 50.0 | -0.473 | -0.643 | -0.805 | -0.953 | -0.913 | -0.895 | 117 | 52.5 | -0.015 | -0.010 | -0.020 | -0.054 | -0.081 | -0.051 |\n| 39 | 59.0 | -0.375 | -0.496 | -0.604 | -0.701 | -0.871 | -0.944 | 118 | 62.5 | -0.011 | -0.011 | -0.020 | -0.066 | -0.101 | -0.076 |\n| 40 | 67.5 | - | - | - | - | - | - | 119 | 72.5 | -0.040 | -0.041 | -0.003 | -0.015 | -0.062 | -0.047 |\n| 41 | 77.5 | -0.176 | -0.263 | -0.312 | -0.325 | -0.390 | -0.564 | 120 | 87.4 | -0.046 | -0.046 | -0.052 | -0.021 | -0.015 | -0.007 |\n| 42 | 87.5 | -0.097 | -0.108 | -0.177 | -0.262 | -0.323 | -0.342 | 121 | 94.2 | -0.063 | -0.067 | -0.018 | -0.006 | -0.006 | -0.045 |\n| 43 | 94.2 | -0.004 | -0.074 | -0.160 | -0.247 | -0.309 | -0.323 | | | | | | | | |\n| | | | | | | | | | | | | | | | |\n| E44 | 2.0 | -1.1", "timestamp": "2026-07-22T04:22:15.706919+00:00"}
{"citation_id": "19930085982", "source_url": "https://ntrs.nasa.gov/api/citations/19930085982/downloads/19930085982.pdf", "page_number": 27, "total_pages": 32, "image_filename": "19930085982_p27.jpg", "text": "NACA RM E9E13\n25\n\n11331\n\nRelative Mach number\nat rotor inlet\n.6\n.5\n.4\n\nEquivalent angle of attack, $\\alpha_e$, deg\n16\n12\n8\n4\n0\n.5 .6 .7 .8 .9 1.0\nRadius ratio, $r/r_t$\n\n[Figure: NACA logo]\n\nFigure 5. - Angle of attack and Mach number at\ncorrected weight flow of 21.1 pounds per second\nfor three-fourths design speed.", "timestamp": "2026-07-22T04:22:18.270527+00:00"}
{"citation_id": "19930082090", "source_url": "https://ntrs.nasa.gov/api/citations/19930082090/downloads/19930082090.pdf", "page_number": 108, "total_pages": 122, "image_filename": "19930082090_p108.jpg", "text": "106\nNACA TN No. 1455\n\n<!-- Image (127, 114, 768, 907) -->\n\nFigure 55.- Thermal output and isothermal frictional pressure drops of aluminum-alloy fin heat exchanger N.", "timestamp": "2026-07-22T04:22:22.685929+00:00"}
{"citation_id": "19930085930", "source_url": "https://ntrs.nasa.gov/api/citations/19930085930/downloads/19930085930.pdf", "page_number": 59, "total_pages": 92, "image_filename": "19930085930_p59.jpg", "text": "```markdown\n58\n\nCONFIDENTIAL\nUNCLASSIFIED\n\n$$ \\frac{p_2}{p_0} $$\nLocal static-pressure ratio,\n\n$$ \\bigcirc \\frac{p_1}{p_A} = 1.28 $$\n$$ \\square \\frac{p_1}{p_A} = 1.00 $$\n$$ \\diamond \\frac{p_1}{p_A} = .85 $$\n\n.28\n.26\n.24\n.22\n.20\n.18\n\n0 .2 .4 .6 .8 1.0 1.2 1.4 1.6\nDistance from convex surface\n\nConcave surface\n\nNACA\n\n(a) Local static-pressure ratio.\n\nFigure 24.- Pressure distribution at 50-percent-span station for three static-pressure ratios at an area ratio of 1.077 for model 2.\n\nNACA RM L9G07\n```", "timestamp": "2026-07-22T04:22:31.378290+00:00"}
{"citation_id": "19930086151", "source_url": "https://ntrs.nasa.gov/api/citations/19930086151/downloads/19930086151.pdf", "page_number": 5, "total_pages": 34, "image_filename": "19930086151_p5.jpg", "text": "NACA RM L9J28 CONFIDENTIAL 3\n\nThe symbols used in the presentation of results are as follows:\n\n$C_L$ lift coefficient (twice lift of semispan model/qS)\n\n$C_D$ drag coefficient (D/qS)\n\n$C_m$ pitching-moment coefficient (M/qSc)\n\n$C_l$ rolling-moment coefficient (L/qSb)\n\n$C_n$ yawing-moment coefficient (N/qSb)\n\npb/2V wing-tip helix angle, radians\n\n$C_{l_p}$ damping-in-roll coefficient; that is, rate of change of rolling-moment coefficient with wing-tip helix angle $\\left( \\partial C_l / \\partial \\left( \\frac{pb}{2V} \\right) \\right)$\n\n$\\overline{c}$ wing mean aerodynamic chord $\\left( \\frac{2}{S} \\int_0^{b/2} c^2 dy \\right)$\n\n(wing with parallelogram-plan-form aileron, 3.42 ft; wing with triangular-plan-form aileron, 3.36 ft)\n\nc local wing chord, feet\n\nb twice span of each semispan model, including aileron\n\n(wing with parallelogram-plan-form aileron, 6.28 ft; wing with triangular-plan-form aileron, 6.97 ft)\n\ny lateral distance from plane of symmetry, feet\n\nS twice area of each semispan model, including aileron\n\n(21.02 sq ft)\n\nD twice drag of semispan models, pounds\n\nM twice pitching moment of semispan model about Y-axis, foot-pounds\n\nL rolling moment, resulting from aileron deflection, about X-axis, foot-pounds\n\nN yawing moment, resulting from aileron deflection, about Z-axis, foot-pounds\n\nCONFIDENTIAL", "timestamp": "2026-07-22T04:22:32.167259+00:00"}
{"citation_id": "19930090382", "source_url": "https://ntrs.nasa.gov/api/citations/19930090382/downloads/19930090382.pdf", "page_number": 3, "total_pages": 37, "image_filename": "19930090382_p3.jpg", "text": "```markdown\n2\nCONFIDENTIAL\nNACA RM L9I07\n\nForce-test results and a limited analysis for only the unswept propeller are presented at this time to expedite publication of this information. Large-scale plots of the basic propeller characteristics (fig. 5) are available on request to the NACA.\n\nSYMBOLS\n\n| | |\n| :--- | :--- |\n| b | blade width, feet |\n| $c_{l_d}$ | blade-section design lift coefficient |\n| $C_P$ | power coefficient $\\left(P/\\rho n^3 D^5\\right)$ |\n| $C_T$ | thrust coefficient $\\left(T/\\rho n^2 D^4\\right)$ |\n| D | propeller diameter, feet |\n| b/D | blade width ratio |\n| h | maximum thickness of blade section, feet |\n| h/b | blade thickness ratio |\n| J | advance ratio $\\left(V_o/nD\\right)$ |\n| M | tunnel-datum (forward) Mach number (tunnel Mach number uncorrected for tunnel-wall constraint) |\n| $M_t$ | helical tip Mach number $\\left(M\\sqrt{1 + \\frac{\\pi^2}{J^2}}\\right)$ |\n| n | propeller rotational speed, revolutions per second |\n| P | power, foot-pounds per second |\n| q | dynamic pressure, pounds per square foot $\\left(\\rho V^2/2\\right)$ |\n| R | propeller-tip radius, feet |\n| r | blade-section radius, feet |\n| T | thrust, pounds |\n\nCONFIDENTIAL\n```", "timestamp": "2026-07-22T04:22:35.480456+00:00"}
{"citation_id": "19930085842", "source_url": "https://ntrs.nasa.gov/api/citations/19930085842/downloads/19930085842.pdf", "page_number": 57, "total_pages": 104, "image_filename": "19930085842_p57.jpg", "text": "NACA RM L9C29\n53\n\n[Figure: Graph plotting Pitching moment coefficient, Cm (vertical axis) against Lift coefficient, CL (horizontal axis). The vertical axis ranges from -2 to 2. The horizontal axis ranges from 0 to 8. A curve starts near the origin, rises to a peak around CL=1.5, then decreases and flattens out. A white rectangular box obscures part of the lower grid area. The text \"NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\" is visible in the lower right corner of the graph.]\n\nFigure 24.- Variation of pitching-moment coefficient with lift coefficient.\n$C_{h_a} = 0$. Model in complete configuration; propellers removed.", "timestamp": "2026-07-22T04:22:35.717187+00:00"}
{"citation_id": "19930086081", "source_url": "https://ntrs.nasa.gov/api/citations/19930086081/downloads/19930086081.pdf", "page_number": 10, "total_pages": 44, "image_filename": "19930086081_p10.jpg", "text": "8 CONFIDENTIAL NACA RM L9H05\n\nFigure 11 presents the variation of the rolling-effectiveness parameter $\\frac{pb}{2V}$ with Mach number as obtained from free-flight rocket tests at two control deflections, wind-tunnel tests, and calculations based upon linearized theory. The rocket configuration (unpublished data) was tested by the Langley Pilotless Aircraft Research Division by the same technique and subject to the same limitations as the investigation of reference 1. The wing of rocket configuration had the large fence installed and operated at a Reynolds number of about $10 \\times 10^6$ for the maximum Mach number of 1.5. In the region of the wing, the rocket fuselage diameter relative to the exposed wing span lay between the small fuselage and the large fuselage combinations of the wind-tunnel configurations. The fuselage nose of the rocket vehicle extended much farther ahead of the wing. The wind-tunnel test point was obtained from an average value of the experimental rolling effectiveness of the small fuselage configuration (reported herein) divided by an experimental damping coefficient $C_{l_p}$ (from reference 3). To account for difference in fuselage diameters, the experimental rolling effectiveness was multiplied by a factor of 1.02, which is the theoretical ratio between the spanwise location in percent semispan of the control-surface loading for the rocket and the wind-tunnel configurations. The experimental damping coefficient included the effect of a fuselage (having about the same diameter relative to the wing span as did the present configuration) and had a value of 85 percent of that calculated for a flat-plate delta wing by linearized theory (reference 4). The calculations of $\\frac{pb}{2V}$ by linearized theory utilized the method of reference 5 to obtain the rolling moment caused by control deflection and the method of reference 4 to obtain the damping coefficient (ignoring fuselage effects). Figure 12 presents the lift-drag curves of several configurations differing in tip thickness and in the fairing of the airfoil contours.\n\nFigures 13 and 14 present the data first obtained for the control surface alone tested in the presence of the wing panel both with fence off and with large fence on. The coefficients in this figure are based on the wing dimensions and the moments are taken about the wing wind axes. Cross plots of these data at zero angle of attack are shown in figure 15 along with comparable data for the complete wing.\n\nThe second series of control-surface tests were made after first stiffening the balance structure (to increase the angle-of-attack range) and shifting the electrical center of the balance moment measuring components to the axes of the control surface. The data for these tests are presented in coefficient form (figs. 16 and 17) and include\n\nCONFIDENTIAL", "timestamp": "2026-07-22T04:22:35.949087+00:00"}
{"citation_id": "19930086022", "source_url": "https://ntrs.nasa.gov/api/citations/19930086022/downloads/19930086022.pdf", "page_number": 24, "total_pages": 34, "image_filename": "19930086022_p24.jpg", "text": "22\nNACA RM L9E24\n\n<!-- Image (228, 123, 717, 778) -->\n\n(c) $C_m$ and $C_L$ plotted against $\\alpha$.\nFigure 5.- Concluded.", "timestamp": "2026-07-22T04:22:39.798106+00:00"}
{"citation_id": "19930085551", "source_url": "https://ntrs.nasa.gov/api/citations/19930085551/downloads/19930085551.pdf", "page_number": 75, "total_pages": 82, "image_filename": "19930085551_p75.jpg", "text": "```markdown\n74\nNACA RM No. L8K30\n\n<!-- Image (171, 116, 786, 823) -->\n\n(a) Wave-off from final approach condition; flaps full down; gear down;\nNo. 1 engine idling; No. 2, 3, and 4 engines 45 in. Hg; 2550 rpm.\nTrimmed for symmetrical power in approach condition.\n\nFigure 20.- Lateral and longitudinal trim characteristics with asymmetric\npower. C-54D airplane.\n```", "timestamp": "2026-07-22T04:22:39.999789+00:00"}
{"citation_id": "19930085958", "source_url": "https://ntrs.nasa.gov/api/citations/19930085958/downloads/19930085958.pdf", "page_number": 35, "total_pages": 43, "image_filename": "19930085958_p35.jpg", "text": "34\nNACA RM No. L9B11\n\n<!-- Image (125, 78, 812, 832) -->\n\n(a) Flaps off. (b) $0.60\\frac{b}{2}$ drooped nose flaps deflected $30^\\circ$. Split flaps on. (c) $0.60\\frac{b}{2}$ drooped nose flaps deflected $30^\\circ$. Split flaps and upper surface fences on.\n\nFigure 15.- Stalling characteristics of a $42^\\circ$ sweptback wing.", "timestamp": "2026-07-22T04:22:40.234756+00:00"}
{"citation_id": "19930086105", "source_url": "https://ntrs.nasa.gov/api/citations/19930086105/downloads/19930086105.pdf", "page_number": 6, "total_pages": 22, "image_filename": "19930086105_p6.jpg", "text": "4\nCONFIDENTIAL\nNACA RM E9H12\n\nHigh-speed schlieren photographs of the shock pattern at the\ndiffuser inlet were taken with a 16-millimeter camera operating at\napproximately 2500 frames per second.\n\nSYMBOLS\n\nThe following symbols are used in this report:\n\n| | |\n| :--- | :--- |\n| $A_1$ | diffuser-inlet area with cone removed |\n| $A_4$ | outlet-nozzle area |\n| $P_0$ | free-stream total pressure |\n| $P_3$ | total pressure at combustion-chamber inlet (diffuser outlet) |\n| $p_3$ | static pressure at combustion-chamber inlet (diffuser outlet) |\n| $f/a$ | fuel-air ratio |\n\nRESULTS AND DISCUSSION\n\nPerforated Conical Flame Holder\n\nTypical of the instantaneous data taken with the variable-\ninductance pressure pickup is the pressure-time diagram in figure 2.\nThe pressure ratio indicated by static tubes is superimposed on this\ninstantaneous pressure signal. Within the accuracy of the curve, the\nmeasured mean pressure from the static-pressure tubes coincides with\nthe integrated mean of the instantaneous pressures between points A\nand B.\n\nThe wave pattern (fig. 2) was the result of pressure fluctuations\nin the combustion chamber when the shock was oscillating in and out of\nthe diffuser inlet. During this unstable operation, a distinct change\nin the slope of the wave form between the conditions of increasing and\ndecreasing pressure was noted. As the back pressure on the diffuser\nincreased, the normal shock moved upstream of the inlet until enough\nair flow was spilled that the burner presumably approached or reached\nthe rich blow-out limit. Thereupon, the shock rapidly reentered the\ndiffuser (as indicated by the nearly vertical slope) and the conditions\nagain became favorable for combustion.\n\nCONFIDENTIAL\n1179", "timestamp": "2026-07-22T04:22:41.860797+00:00"}
{"citation_id": "19930086003", "source_url": "https://ntrs.nasa.gov/api/citations/19930086003/downloads/19930086003.pdf", "page_number": 23, "total_pages": 36, "image_filename": "19930086003_p23.jpg", "text": "```markdown\nNACA RM L5J108\n\nCONFIDENTIAL\n\n| Angle of attack, $\\alpha$, deg | Lift coefficient, $C_L$ | Pitching-moment coefficient, $C_m$ | Drag coefficient, $C_D$ | Lift coefficient, $C_L$ |\n| :--- | :--- | :--- | :--- | :--- |\n| 0 | -0.2 | M | 0 | -0.2 |\n| 0 | 0 | 1.18 $\\nabla$ | 0 | 0 |\n| 0 | 0.2 | 1.15 $\\diamond$ | 0 | 0.2 |\n| 0 | 0.4 | 1.10 $\\diamond$ | 0 | 0.4 |\n| 0 | 0.6 | 1.08 $\\diamond$ | 0 | 0.6 |\n| 0 | 0.8 | 1.05 $\\triangle$ | 0 | 0.8 |\n| 12 | | 1.03 $\\nabla$ | 0.20 | |\n| 0 | | 1.00 $\\diamond$ | 0.16 | |\n| 0 | | .98 $\\diamond$ | 0.12 | |\n| 0 | | .95 $\\diamond$ | 0.08 | |\n| 0 | | .93 $\\diamond$ | 0.04 | |\n| 0 | | .90 $\\diamond$ | 0 | |\n| 8 | | .88 $\\diamond$ | | |\n| 0 | | .85 $\\triangle$ | | |\n| 0 | | .80 $\\diamond$ | | |\n| 4 | | .70 $\\square$ | | |\n| 0 | | .60 $\\diamond$ | | |\n| -4 | | | | |\n\n[Figure: Three graphs showing aerodynamic characteristics. The left graph plots Angle of attack vs Lift coefficient. The middle graph plots Pitching-moment coefficient vs Lift coefficient. The right graph plots Drag coefficient vs Lift coefficient. Each graph contains multiple curves corresponding to different Mach numbers (M) as listed in the legend. The NACA logo is visible in the bottom right corner of the rightmost graph.]\n\nFigure 8.- Aerodynamic characteristics for a model with $45^\\circ$ sweptback wing, aspect ratio 6, taper ratio 0.6, and NACA 65A006 airfoil section. Wing alone.\n\n12\n```", "timestamp": "2026-07-22T04:22:44.518312+00:00"}
{"citation_id": "19930085838", "source_url": "https://ntrs.nasa.gov/api/citations/19930085838/downloads/19930085838.pdf", "page_number": 73, "total_pages": 118, "image_filename": "19930085838_p73.jpg", "text": "NACA RM No. L9B23\n\n71\n\nAileron section hinge-moment coefficient, $c_{h_a}$\n\nSection angle of attack, $\\alpha_0$, deg\n\n$\\delta_1 = -2^\\circ$\n\n$\\delta_a$ (deg)\n\n0\n\n5\n\n10\n\n15\n\nNACA\n\n(b) $\\delta_f = 40^\\circ$.\n\nFigure 9.- Continued.", "timestamp": "2026-07-22T04:22:45.825489+00:00"}
{"citation_id": "19930085982", "source_url": "https://ntrs.nasa.gov/api/citations/19930085982/downloads/19930085982.pdf", "page_number": 28, "total_pages": 32, "image_filename": "19930085982_p28.jpg", "text": "26\nNACA RM E9E13\n\n1131\n\n[Figure: Graph plotting Corrected mean total enthalpy addition vs. Corrected weight flow. The graph includes a curve with data points, a marked \"Design point\", and a NACA logo in the bottom left corner.]\n\nCorrected mean total enthalpy addition, ft-lb/slug\n14 x 10^4\n12\n10\n8\n6\n4\n14 16 18 20 22\nCorrected weight flow, W√θ/δ, lb/sec\n\nFigure 6. - Energy addition across rotor\nat three-fourths design speed.", "timestamp": "2026-07-22T04:22:46.266249+00:00"}
{"citation_id": "19930082090", "source_url": "https://ntrs.nasa.gov/api/citations/19930082090/downloads/19930082090.pdf", "page_number": 109, "total_pages": 122, "image_filename": "19930082090_p109.jpg", "text": "NACA TN No. 1455\n107\n\nExhaust gas\n50.5\"\nVentilating air\nB\n15.7\"\nVentilating air\nB\n2.25\"\n8.38\"\nExhaust gas\n\nSection B-B\nUC-1 air shroud\n8.38\"\n5.1\"\n\n13.5\"\n3.1\"\n11\"\n8.0\"\nNACA\nCopper and stainless-steel\nheat exchanger\n\n27 equally spaced\ninternal fins\n27 double rows of\nexternal fins\n(equally spaced\n62 rows)\n5.6\"\n\nFigure 56.- Schematic diagram of fin heat exchanger O and air shroud. Weight\nof heat exchanger, 24.5 pounds.\n\n| | Air side | Gas side |\n| :--- | :--- | :--- |\n| Cross-sectional area, sq ft | 0.262 | 0.142 |", "timestamp": "2026-07-22T04:22:51.843623+00:00"}
{"citation_id": "19930085965", "source_url": "https://ntrs.nasa.gov/api/citations/19930085965/downloads/19930085965.pdf", "page_number": 31, "total_pages": 67, "image_filename": "19930085965_p31.jpg", "text": "30\nNACA RM E9E06\n\nAPPENDIX A\n\nDERIVATION OF ANGLES $\\delta_2$ AND $\\beta_2$\n\nFrom figure 6,\n$$ \\Phi_{\\text{max}} \\text{d}\\delta = B_{\\text{max}} \\text{1 dl sin } (45^\\circ+\\beta-\\delta) \\tag{A1} $$\n\nor, using equation (10),\n$$ \\text{d}\\delta = \\frac{B_{\\text{max}} \\text{1 dl sin } (45^\\circ+\\beta-\\delta)}{\\frac{B_{\\text{max}} \\text{1}}{\\sqrt{2} \\text{ c}} + \\int B_{\\text{max}} \\text{1 dl cos } (45^\\circ+\\beta-\\delta)} $$\n$$ = \\frac{\\text{sin } (45^\\circ+\\beta-\\delta)}{\\text{cos } (45^\\circ+\\beta-\\delta)} \\frac{\\sqrt{2} \\text{ c dl cos } (45^\\circ+\\beta-\\delta)}{1 + \\int \\sqrt{2} \\text{ c dl cos } (45^\\circ+\\beta-\\delta)} \\tag{A2} $$\n\nIntegrating in the limits 0 to $\\delta_2$ and 0 to $l_2$,\n$$ \\delta_2 = \\text{tan } (45^\\circ+\\alpha_2) \\text{log}_e \\left[ 1 + \\sqrt{2} \\text{ c cos } (45^\\circ+\\alpha_2) l_2 \\right] \\tag{A3} $$\n\nwhere $\\alpha_2$ is the mean value of $\\beta_2-\\delta_2$ for the given value of $l_2$, which gives from equation (15a)\n$$ \\delta_2 = \\frac{\\text{tan } (45^\\circ+\\alpha_2)}{2} \\text{log}_e \\left[ 1 + 2 \\text{ cos } (45^\\circ+\\alpha_2) \\left( \\frac{H_{\\text{max}} \\text{2}}{H_{\\text{max}} \\text{1}} - 1 \\right) \\right] \\tag{A4} $$\n\nThe derivation of $\\beta_2$ follows a similar procedure. In figure 6, the increment of current in each new sheet of material of thickness dl is equal to the product of induced voltage and the conductance of the sheet, or $\\gamma E_{\\text{max}} \\text{dl}$; therefore,\n$$ I_{\\text{max}} = I_{\\text{max}} \\text{1} + \\int \\gamma E_{\\text{max}} \\text{dl sin } (45^\\circ+\\beta-\\delta) $$\n$$ = \\frac{H_{\\text{max}} \\text{1}}{0.4\\pi} + \\int \\gamma E_{\\text{max}} \\text{dl sin } (45^\\circ+\\beta-\\delta) \\tag{A5} $$", "timestamp": "2026-07-22T04:22:52.846507+00:00"}
{"citation_id": "19930085936", "source_url": "https://ntrs.nasa.gov/api/citations/19930085936/downloads/19930085936.pdf", "page_number": 46, "total_pages": 59, "image_filename": "19930085936_p46.jpg", "text": "NACA RM No. E9B03\n45\n\nPressure coefficient, $C_p$\n\nAngle\nof yaw\n(deg)\n$\\circ$ -12\n$\\square$ -6\n$\\diamond$ -0\n$\\triangle$ 6\n$\\nabla$ 12\n\nDistance from tip, x/L\n\n(d) $\\theta = 225^\\circ$ longitudinal plane.\n\nFigure 8. - Continued. Pressure distributions along longitudinal\nplanes at $10^\\circ$ angle of attack for range of yaw angles.", "timestamp": "2026-07-22T04:22:55.187182+00:00"}
{"citation_id": "19930086078", "source_url": "https://ntrs.nasa.gov/api/citations/19930086078/downloads/19930086078.pdf", "page_number": 13, "total_pages": 42, "image_filename": "19930086078_p13.jpg", "text": "NACA RM L9H04 CONFIDENTIAL 11\n\nthe unswept and sweptback wing configurations, respectively, and were obtained from reference 11. These values of $C_{T_p}$ pertain to low-speed data for the wing alone and do not account for the increase in aspect ratio resulting from extending the aileron. All values of $pb/2V$ are probably high since rolling due to sideslip, yawing, and wing twist were neglected.\n\nExcept for the short-chord aileron on the unswept wing, for which the values of $pb/2V$ are low over the entire $C_L$ range, aileron effectiveness available at moderate and large lift coefficients with each of the ailerons investigated on both wing configurations (figs. 18 and 19) would easily satisfy requirements of reference 12. However, at small lift coefficients, aileron effectiveness appears to be inadequate for application to an airplane. It is thought that the extensible ailerons may be sufficiently effective for some types of missiles.\n\nAdverse yawing moments produced by the ailerons on the unswept and sweptback configurations generally varied in the same manner with changes in angle of attack, aileron deflection, and/or aileron extension, but yawing moments were generally larger for the sweptback configurations. These yawing moments were comparable to those produced by conventional flap-type ailerons.\n\nThe rudder deflection required in a roll to correct for adverse yawing moments due to aileron extension and yawing moments due to rolling was computed for an assumed airplane utilizing the 45° swept-back wing with triangular tip ailerons. The vertical tail of the assumed airplane had 45° of sweepback, an aspect ratio of 1.0, an area of 0.15 of the wing area, a rudder chord of 0.25 of the vertical-tail chord, and a tail length of 2.5$\\bar{c}$. For sweptback wings, yawing moments due to roll $C_{n_p}$ are adverse at low lift coefficients and favorable ($C_n$ same sign as $C_l$) at high lift coefficients (reference 13). For $\\frac{pb}{2V} = 0.15$ and $C_L = 0.5$ (fig. 19(b), 1/2 aileron extension) and $C_{n_p} = -0.04$ (reference 13), a rudder deflection of less than 10° would maintain a coordinated roll. At high lift coefficients rudder deflections would be small since the adverse yawing moments due to aileron extension are counteracted by the favorable yawing moments due to roll.\n\nCONCLUSIONS\n\nA low-speed wind-tunnel investigation, made to determine the lateral control characteristics of extensible wing-tip ailerons on an untapered semispan wing at 0° and 45° sweepback, led to the following conclusions:\n\nCONFIDENTIAL", "timestamp": "2026-07-22T04:22:58.070439+00:00"}
{"citation_id": "19930085952", "source_url": "https://ntrs.nasa.gov/api/citations/19930085952/downloads/19930085952.pdf", "page_number": 43, "total_pages": 50, "image_filename": "19930085952_p43.jpg", "text": "42\nNACA RM L9C24\n\n<!-- Image (199, 154, 708, 814) -->\n\n(c) $\\alpha \\approx 58^\\circ$.\nFigure 20.- Continued.", "timestamp": "2026-07-22T04:23:03.402503+00:00"}
{"citation_id": "19930085930", "source_url": "https://ntrs.nasa.gov/api/citations/19930085930/downloads/19930085930.pdf", "page_number": 60, "total_pages": 92, "image_filename": "19930085930_p60.jpg", "text": "```markdown\nCONFIDENTIAL\n\nNACA RM L9G07\n\nLocal stagnation-pressure recovery, $\\frac{P_2}{P_0}$\n\n1.0\n.8\n.6\n.4\n.2\n0\n0 .2 .4 .6 .8 1.0 1.2 1.4 1.6\nDistance from convex surface\n\n$\\circ \\quad \\frac{P_1}{P_A} = 1.28$\n$\\square \\quad \\frac{P_1}{P_A} = 1.00$\n$\\diamond \\quad \\frac{P_1}{P_A} = .85$\n\nConcave surface\n\n[Figure: Graph showing local stagnation-pressure recovery vs. distance from convex surface for three pressure ratios, with data points marked by circles, squares, and diamonds.]\n\n(b) Local stagnation-pressure recovery.\n\nFigure 24.- Concluded.\n\nNACA\n\n59\n```", "timestamp": "2026-07-22T04:23:10.810816+00:00"}
{"citation_id": "19930085842", "source_url": "https://ntrs.nasa.gov/api/citations/19930085842/downloads/19930085842.pdf", "page_number": 58, "total_pages": 104, "image_filename": "19930085842_p58.jpg", "text": "54\nNACA RM L9C29\n\n<!-- Image (100, 109, 842, 839) -->\n\nFigure 25.- Variation of $C_L$, $C_D$, and $C_m$ with $\\alpha$ of the model with the horizontal tail installed and removed. Basic model configuration; propellers removed.", "timestamp": "2026-07-22T04:23:11.877619+00:00"}
{"citation_id": "19930086022", "source_url": "https://ntrs.nasa.gov/api/citations/19930086022/downloads/19930086022.pdf", "page_number": 25, "total_pages": 34, "image_filename": "19930086022_p25.jpg", "text": "NACA RM L9E24\n23\n\n<!-- Image (250, 119, 771, 836) -->\n\n(a) $C_{Z_a}$, $C_n$, and $C_l$ plotted against $\\alpha$.\n\nFigure 6.— Aileron characteristics of wing with leading and trailing flaps deflected and fences installed.", "timestamp": "2026-07-22T04:23:13.232654+00:00"}
{"citation_id": "19930086151", "source_url": "https://ntrs.nasa.gov/api/citations/19930086151/downloads/19930086151.pdf", "page_number": 6, "total_pages": 34, "image_filename": "19930086151_p6.jpg", "text": "4\nCONFIDENTIAL\nNACA RM L9J28\n\nq free-stream dynamic pressure, pounds per square foot ($\\frac{1}{2} \\rho V^2$)\n\nV free-stream velocity, feet per second\n\n$\\rho$ mass density of air, slugs per cubic foot\n\n$\\alpha$ angle of attack with respect to chord plane at root of models, degrees\n\n$\\delta_a$ aileron deflection, measured between wing chord plane and aileron chord plane (positive when trailing edge is down), degrees\n\n$\\delta_{at}$ total aileron deflection\n\nA wing aspect ratio ($b^2/S$)\n(wing with parallelogram-plan-form aileron, 1.87; wing with triangular-plan-form aileron, 2.31)\n\n$C_{l\\delta_a}$ rate of change of rolling-moment coefficient with aileron deflection ($\\partial C_l / \\partial \\delta_a$)\n\nCORRECTIONS\n\nThe angle-of-attack and the drag data have been corrected for jet-boundary (induced-upwash) effects according to the methods outlined in reference 7. Blockage corrections were applied to the test data by the methods of reference 8.\n\nReflection-plane corrections were not applied to the rolling-moment and yawing-moment data because available correction data did not apply to the configurations of this investigation. However, by extrapolation of the correction data of reference 9, it is estimated that the values of $C_l$ presented herein were approximately 10 percent too high for both wing-aileron configurations. In addition, the yawing moments, if corrected, would be generally more adverse than the data show.\n\nMODEL AND APPARATUS\n\nThe right semispan wing model was mounted vertically in the Langley 300 MPH 7- by 10-foot tunnel with the root chord of the model adjacent\n\nCONFIDENTIAL", "timestamp": "2026-07-22T04:23:13.717786+00:00"}
{"citation_id": "19930090382", "source_url": "https://ntrs.nasa.gov/api/citations/19930090382/downloads/19930090382.pdf", "page_number": 4, "total_pages": 37, "image_filename": "19930090382_p4.jpg", "text": "```markdown\nNACA RM L9I07 CONFIDENTIAL 3\n\n$T_c$ thrust disk-loading coefficient ($T/2qD^2$)\n\nV tunnel-datum velocity (tunnel velocity uncorrected for tunnel-wall constraint), feet per second\n\n$V_o$ equivalent free-air velocity (tunnel-datum velocity corrected for tunnel-wall constraint), feet per second\n\nx blade-section station (r/R)\n\n$\\beta$ section blade angle, degrees\n\n$\\beta_{0.7R}$ section blade angle at 0.7 tip radius, degrees\n\n$\\eta$ efficiency ($\\frac{C_T}{C_P} J$)\n\n$\\eta_{max}$ maximum efficiency\n\n$\\rho$ air density, slugs per cubic foot\n\nAPPARATUS, METHODS, AND TESTS\n\nThe apparatus and methods described in reference 1 were used in this investigation which was conducted in the Langley 8-foot high-speed tunnel. A sketch of the 800-horsepower dynamometer installed in the tunnel is shown as figure 1.\n\nThe NACA 4-(4)(06)-04 two-blade propeller used in this investigation is the unswept propeller of a family of propellers designed to study the effect of blade sweep on propeller characteristics at transonic speeds. It utilizes NACA 16-series airfoil sections and was designed as a two-blade propeller to produce minimum energy losses (profile drag assumed equal to zero) at a blade angle of $60^\\circ$ at the 0.7-radius station and at an advance ratio of 3.65. The gaps between the spinner and blades were sealed for all operating conditions. Blade-form curves are given in figure 2, and a photograph of the blades is shown in figure 3.\n\nThrust, torque, and rotational speed were measured throughout the complete operating range of the propeller. For each tunnel Mach number the propeller was run at a constant blade angle and the rotational speed\n\nCONFIDENTIAL\n```", "timestamp": "2026-07-22T04:23:13.920053+00:00"}
{"citation_id": "19930085958", "source_url": "https://ntrs.nasa.gov/api/citations/19930085958/downloads/19930085958.pdf", "page_number": 36, "total_pages": 43, "image_filename": "19930085958_p36.jpg", "text": "NACA RM No. L9B11\n35\n\n<!-- Image (194, 80, 861, 839) -->\n\n(a) Low wing.\n(b) Midwing.\n(c) High wing.\n\nFigure 16.- Stalling characteristics of a 42° sweptback wing-fuselage combination with 0.60c drooped nose flaps, split flaps, and upper surface fences.", "timestamp": "2026-07-22T04:23:15.576648+00:00"}
{"citation_id": "19930085551", "source_url": "https://ntrs.nasa.gov/api/citations/19930085551/downloads/19930085551.pdf", "page_number": 76, "total_pages": 82, "image_filename": "19930085551_p76.jpg", "text": "NACA RM No. L8K30\n75\n\nAngle of bank, deg\nSideslip angle, deg\nLeft Right\n5\n0\n5\nSideslip\nAngle of bank\n\nRudder force, lb\nLeft Right\n200\n100\n0\n100\nWheel force, lb\nPush Pull\nLeft Right\n80\n40\n0\n40\nRudder\nAileron\nElevator\n\nControl position, deg\nDown Up\nLeft Right\n20\n10\n0\n10\nRudder\nTotal aileron\nElevator\nMaximum rudder deflection; no load\nNACA\n\n100 120 140\nIndicated airspeed, mph\n\n(b) Take-off condition; flaps 20°; gear down; No. 1 engine idling; No. 2, 3, and 4 engines 45 in. Hg; 2550 rpm. Trimmed for symmetrical power in take-off condition.\n\nFigure 20.— Concluded.", "timestamp": "2026-07-22T04:23:16.840821+00:00"}
{"citation_id": "19930085838", "source_url": "https://ntrs.nasa.gov/api/citations/19930085838/downloads/19930085838.pdf", "page_number": 74, "total_pages": 118, "image_filename": "19930085838_p74.jpg", "text": "72\nNACA RM No. L9B23\n\n<!-- Image (99, 119, 833, 926) -->\n\nSection angle of attack, $\\alpha_o$, deg\n(1) $\\delta_F = 40^\\circ$.\nFigure 9.- Continued.", "timestamp": "2026-07-22T04:23:22.154483+00:00"}
{"citation_id": "19930086003", "source_url": "https://ntrs.nasa.gov/api/citations/19930086003/downloads/19930086003.pdf", "page_number": 24, "total_pages": 36, "image_filename": "19930086003_p24.jpg", "text": "22\nNACA RM L9I08\n\nCONFIDENTIAL\n\nBending-moment coefficient, $C_B$\nLift coefficient, $C_L$\n\n| M | |\n| :--- | :--- |\n| 1.18 | $\\nabla$ |\n| 1.15 | $\\diamond$ |\n| 1.10 | $\\square$ |\n| 1.08 | $\\nabla$ |\n| 1.05 | $\\nabla$ |\n| 1.03 | $\\nabla$ |\n| 1.00 | $\\triangle$ |\n| .98 | $\\diamond$ |\n| .95 | $\\diamond$ |\n| .93 | $\\square$ |\n| .90 | $\\triangle$ |\n| .88 | $\\triangle$ |\n| .85 | $\\triangle$ |\n| .80 | $\\diamond$ |\n| .70 | $\\square$ |\n| .60 | $\\circ$ |\n\n[Figure: Graph plotting Bending-moment coefficient ($C_B$) against Lift coefficient ($C_L$) for various Mach numbers (M). The graph contains multiple data series represented by different symbols corresponding to the legend on the right. A NACA logo is visible at the bottom right of the plot area.]\n\nNACA\n\nCONFIDENTIAL\n\nFigure 8.- Concluded.", "timestamp": "2026-07-22T04:23:22.654174+00:00"}
{"citation_id": "19930082090", "source_url": "https://ntrs.nasa.gov/api/citations/19930082090/downloads/19930082090.pdf", "page_number": 110, "total_pages": 122, "image_filename": "19930082090_p110.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T04:23:24.784277+00:00"}
{"citation_id": "19930086081", "source_url": "https://ntrs.nasa.gov/api/citations/19930086081/downloads/19930086081.pdf", "page_number": 11, "total_pages": 44, "image_filename": "19930086081_p11.jpg", "text": "NACA RM L9H05 CONFIDENTIAL 9\n\nnormal force $C_{N_p}$, chord force $C_{C_p}$, pitching moment about the control pivot axis $C_{M_p}$, and the bending moment about the root chord of the control $C_{BM_p}$. Cross-plots of these data are presented in figure 18 for angles of attack of $0^\\circ$, $2^\\circ$, and $4^\\circ$. Since the model had symmetrical airfoil sections, all angles and coefficients can arbitrarily be reversed in sign. This change in sign makes possible the application of the test data to cover the condition of negative deflection angles for the control. This procedure has been followed in presenting the cross-plot data of figure 18 to show the nature of the curve shapes in the negative range of control deflections. In going from negative to positive deflections, a discontinuity exists in most curves as a result of inaccuracies in the test measurements.\n\nDISCUSSION\n\nWing Characteristics\n\nControl undeflected.- With the 3-percent-thick tip, the value of the wing lift-curve slope $C_{L_\\alpha}$ for both fuselage conditions was about 0.040. The calculated value based on flat-plate theory (reference 6) corrected by an estimate of the additional lift resulting from fuselage upwash (reference 7) was 0.047 for the small fuselage and 0.049 for the large fuselage. The minimum drag coefficient was about 0.012 with the large fuselage and 0.013 with the small fuselage. Based on the lift and pitching-moment data of figures 4 to 9, the chordwise location of the aerodynamic center was 7 percent of the wing mean aerodynamic chord ahead of the center of area. Similarly the lift and rolling-moment slopes indicated the spanwise center of pressure to be located about 40 percent of the exposed half-span outboard of the wing-fuselage juncture.\n\nControl surface deflected.- Deflecting the control surface in the positive direction tended to increase the value of minimum drag coefficient and to displace negatively the curves of pitching moment plotted against angle of attack. The drag and yawing-moment curves were shifted in the negative angle-of-attack direction since the wing drag load at negative angles of attack tended to be counteracted by a decreased control-deflection loading (control more alined with the air stream).\n\nWithin the accuracy of the test data, the various coefficients varied linearly with control deflection at $\\alpha = 0^\\circ$ (fig. 10). The values of $C_{L_\\delta}$ and $C_{m_\\delta}$ were about 0.004 and -0.0013 as compared with\n\nCONFIDENTIAL", "timestamp": "2026-07-22T04:23:26.212638+00:00"}
{"citation_id": "19930086060", "source_url": "https://ntrs.nasa.gov/api/citations/19930086060/downloads/19930086060.pdf", "page_number": 14, "total_pages": 20, "image_filename": "19930086060_p14.jpg", "text": "12\n\nCONFIDENTIAL\n\n$C_D$\n\n| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | 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