PDF/ocr_dataset_1065.jsonl

{"citation_id": "19930085890", "source_url": "https://ntrs.nasa.gov/api/citations/19930085890/downloads/19930085890.pdf", "page_number": 6, "total_pages": 26, "image_filename": "19930085890_p6.jpg", "text": "NACA RM No. E9C11\n\nPROCEDURE\n\nThe rocket engine was operated by first injecting oxygen until the flow system was cold enough for oxygen to emerge as a liquid. Then diborane was injected and ignition immediately resulted. The average length of a run was 6 seconds. Immediately after the run, the diborane system was purged with helium. The first series of runs was made with an I* of 325 inches and the eight-hole injection system (combustion chamber A and injection plate A); the second series of runs was made with an I* of 159 inches and the eight-hole injection system (combustion chamber B and injection plate A); the third series of runs was made with an I* of 325 inches and the four-hole injection system (combustion chamber A and injection plate B). The large I* value was chosen for most of the runs to insure complete combustion. A combustion-chamber pressure of approximately 500 pounds per square inch absolute was maintained.\n\nWith the comparatively large I* engines used, a considerable amount of heat was lost to the engine walls. The heat absorbed by the engine was determined by the product of the temperature rise of the copper mass, the average specific heat of the copper, and the weight of the engine. The mean temperature rise was determined after the temperature of the copper engine, as measured at several places, had reached a nearly uniform value. This temperature was reached shortly after the end of the run. The value of the temperature rise was corrected for heat loss to the atmosphere between the end of the combustion time and the time that nearly constant temperature was reached.\n\nIn order to present the performance of diborane and liquid oxygen without the penalty of high heat losses, the experimental specific-impulse values were corrected for the heat lost. As the determination of the heat loss is only approximate, the heat-loss correction was made in a simple manner by assuming that the heat loss, if available, could be converted into kinetic energy by the nozzle at the ideal cycle efficiency $\\eta$. Thus\n\n$$\nI \\text{ (corrected)} = \\sqrt{I^2 \\text{ (experimental)} + 48.37 \\ Q\\eta}\n$$\n\nwhere\n\n| I | specific impulse, lb-sec/lb |\n|---|-----------------------------|\n| Q | heat loss, Btu/lb           |\n| $\\eta$ | $1 - T_e/T_c$             |", "timestamp": "2026-07-22T06:26:38.102242+00:00"}
{"citation_id": "19930085900", "source_url": "https://ntrs.nasa.gov/api/citations/19930085900/downloads/19930085900.pdf", "page_number": 2, "total_pages": 33, "image_filename": "19930085900_p2.jpg", "text": "NACA RM L9D20 CONFIDENTIAL\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nRESEARCH MEMORANDUM\n\nTHE EFFECT OF AIR-JET AND STRIP MODIFICATIONS ON THE HYDRODYNAMIC CHARACTERISTICS OF THE STREAMLINE FUSELAGE OF A TRANSONIC AIRPLANE\n\nBy Bernard Weinflash, Kenneth W. Christopher, and Charles L. Shuford, Jr.\n\nSUMMARY\n\nSpecific free-to-trim tests were made on a $\\frac{1}{12}$-size model of a streamline fuselage modified by patterns of air jets or strips on the fuselage bottom. The effects of spacing of jets, length of jet rows, and direction of jets were determined for a simulated chine configuration. Tests were also made of a simulated multiple-step configuration. The effect of air flow on both the chine and step configurations was studied. In addition, the effect of substituting narrow breaker strips for the rows of jets in the chine configuration and in three multiple-step configurations was investigated.\n\nData are presented on resistance, trim, effective hydrodynamic lift, and spray. The resistance was reduced by decreasing the jet spacing, increasing the length of rows of jets, and increasing the air flow. In the chine configuration, the strips gave about the same results as the $\\frac{1}{4}$-inch-spaced jets. Strips in the form of multiple V-steps pointed forward gave the highest resistance and strips in the form of multiple V-steps pointed aft resulted in the lowest resistance of all the jet and strip configurations tested.\n\nINTRODUCTION\n\nWhen a fuselage having a circular or oval cross section moves along a water surface at high speeds, the water flowing up around the convex bottom and sides of the fuselage creates a suction force which keeps the hull low in the water and causes a large hydrodynamic resistance which increases rapidly with speed. Results reported in reference 1 showed that the very high hydrodynamic resistance was greatly reduced when air was ejected at high velocity through fine jets in the fuselage bottom. In that investigation various patterns of jets simulating chines and multiple steps were explored.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:26:38.520155+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 42, "total_pages": 78, "image_filename": "19930082483_p42.jpg", "text": "40\nNACA TN No. 1807\n\n6. Considerable utility of the performance data is afforded by\nuse of the convenient method presented for plotting the turbine\ncharacteristics. The method further serves to condense large\nquantities of data into a single composite plot.\n\n7. In addition to secondary rotor-blade vibration encountered\nin any operating gas turbine, provisions are required to avoid or\nwithstand the effects of primary rotor-blade vibrations induced by\npartial-admission operation.\n\nLewis Flight Propulsion Laboratory,\nNational Advisory Committee for Aeronautics,\nCleveland, Ohio, October 5, 1948.", "timestamp": "2026-07-22T06:26:41.687132+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 53, "total_pages": 114, "image_filename": "19930086061_p53.jpg", "text": "NACA RM L9J07\n49\n\nLeft semispan\nRight semispan\nUpper\nLower\n(c) $\\psi = 20^\\circ$\n\nLeft semispan\nRight semispan\nUpper\nLower\n(d) $\\psi = 35^\\circ$\n\nFigure 13.- Concluded.", "timestamp": "2026-07-22T06:26:42.474931+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 59, "total_pages": 98, "image_filename": "19930086073_p59.jpg", "text": "NACA RM A9E04\n\nLift coefficient, $C_L$\n\nDrag coefficient, $C_D$\n\n$\\delta_{a_L} = +10.8$ $\\quad$ $\\delta_{a_R} = -10.8$\n\nAileron deflection, $\\delta_a$, deg\n\n(b) $C_L$ vs $C_D$.\n\nFigure 12.— Continued.\n\n57", "timestamp": "2026-07-22T06:26:43.688241+00:00"}
{"citation_id": "19930082566", "source_url": "https://ntrs.nasa.gov/api/citations/19930082566/downloads/19930082566.pdf", "page_number": 38, "total_pages": 44, "image_filename": "19930082566_p38.jpg", "text": "$50 \\times 10^3$\n\nMaximum principal stress, $\\sigma_1'$ or $\\sigma_2'$, psi\n\n$\\circ$ R = $\\sigma_2' / \\sigma_1' = 0$\n\n$\\varnothing$ R = $\\sigma_2' / \\sigma_1' = 0.5$\n\n$\\square$ R = $\\sigma_2' / \\sigma_1' = 1$\n\n$\\triangle$ R = $\\sigma_2' / \\sigma_1' = 2$\n\nN, cycles\n\n[Figure: S-N curves for 24S-T aluminum-alloy tubing under different stress ratios]\n\nNACA\n\nFigure 14.- S-N curves of figure 12 shown in a single plot. 24S-T aluminum-alloy tubing.\n\nNACA TN NO. 1889", "timestamp": "2026-07-22T06:26:45.303648+00:00"}
{"citation_id": "19930085870", "source_url": "https://ntrs.nasa.gov/api/citations/19930085870/downloads/19930085870.pdf", "page_number": 9, "total_pages": 92, "image_filename": "19930085870_p9.jpg", "text": "8\nCONFIDENTIAL\nNACA RM No. L9D07\n\nand that for triangular wings with leading edges behind the Mach cone\nthis value becomes\n\n$$\n\\frac{dC_L}{d\\alpha} = \\frac{2\\pi \\frac{\\tan \\epsilon}{\\tan m}}{E\\sqrt{M^2 - 1}} \\quad (2)\n$$\n\nThe lift-curve slopes are shown in figure 9 and are plotted as a\nratio to the theoretical two-dimensional slope, given by equation (1),\nagainst the parameter $\\tan \\epsilon/\\tan m$. The ratio of the measured lift-\ncurve slope to the two-dimensional value is, for any given relation of\nthe Mach line and leading edge, relatively independent of Mach number,\nmore so for the wedge than for the elliptical-leading-edge series. In\nthe lower range of values of $\\tan \\epsilon/\\tan m$, 0 to 0.5, the elliptical- and\nwedge-leading-edge series give approximately the same value of lift-\ncurve-slope ratio, though somewhat higher than that predicted by the\nlinear theory. At values of $\\tan \\epsilon/\\tan m$ between 0.5 and 0.6, the\ncurves of both series cross the theoretical curve and give values\nconsiderably less than the theoretical value in the vicinity of $\\frac{\\tan \\epsilon}{\\tan m} = 1$.\nAs the leading edge becomes coincident with and moves well ahead of the\nMach cone, the lift-curve slopes exhibit a tendency to increase. This\neffect is much more marked for the wedge-leading-edge series and indicates\na more rapid lift recovery, probably due to a more rapid approach to\nattachment of the shock wave to the wedge leading edge. At a value\nof $\\tan \\epsilon/\\tan m$ of 2.19, the lift-curve slope of the wedge-leading-\nedge series attains 98 percent of the two-dimensional value. It was\nnoted that the present tests showed none of the marked breaks in the\nvicinity of $\\frac{\\tan \\epsilon}{\\tan m} = 1$ as were obtained in the tests of reference 10\non a series of thin, flat-plate triangular wings; and to ascertain\nwhether the thicker nature of the present wing series might possibly\nhave eliminated such breaks, eight thin-plate wings of comparable thick-\nness to those tested in reference 10 were tested at a Mach number of 1.92.\nFigure 9 shows that no breaks or abrupt changes in lift-curve slopes were\nobtained from these wings. However, in contrast to the results for the\nthicker triangular-wing series, at values of $\\tan \\epsilon/\\tan m$ less than 1 the\nthin wings gave slightly higher lift-curve slopes for the sharp leading-\nedge configuration than for the round leading edges. Figure 10 is a\ncompilation of several existing results of tests on triangular wings.\nThe faired curves of the present tests are included for comparison.\nExcept for the present tests and the tests of reference 10, the wings\nwere subject to effects of the body on which they were mounted.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:26:51.010680+00:00"}
{"citation_id": "19930082613", "source_url": "https://ntrs.nasa.gov/api/citations/19930082613/downloads/19930082613.pdf", "page_number": 34, "total_pages": 46, "image_filename": "19930082613_p34.jpg", "text": "NACA TN 1938\n\n33\n\n[Figure: Micrograph showing scale formations at a crack. Labels point to “Subsurface scale”, “Surface scale”, and “Crack”. A small box contains: “NACA C-22630 12-9-48”]\n\nFigure 10. - Scale formations at crack. Type-B liner etchant, 5-percent aqua regia in water, electrolytic condition, cracked according to procedure. Scale is not uniform in color and surface scale zone. Scale almost appears to be two-phased because dark zone is dull purple and surface scale is gray.", "timestamp": "2026-07-22T06:26:51.113461+00:00"}
{"citation_id": "19930085899", "source_url": "https://ntrs.nasa.gov/api/citations/19930085899/downloads/19930085899.pdf", "page_number": 4, "total_pages": 29, "image_filename": "19930085899_p4.jpg", "text": "NACA RM No. L9A21\n3\n\n$C_B$\nbending-moment coefficient at plane of symmetry\n$$ \\left( \\frac{\\text{Root bending moment}}{q \\left( \\frac{S}{2} \\right) \\left( \\frac{b}{2} \\right)} \\right) $$\n\n$q$\neffective dynamic pressure over span of model, pounds per square foot $\\left( \\frac{1}{2} \\rho V^2 \\right)$\n\n$S$\ntwice wing area of semispan model, 0.1250 square foot\n\n$\\bar{c}$\nmean aerodynamic chord of wing, 0.181 ft; based on relationship $\\frac{2}{S} \\int_0^{b/2} c^2 dy$ (using the theoretical tip)\n\n$c$\nlocal wing chord\n\n$b$\ntwice span of semispan model\n\n$y$\nspanwise distance from plane of symmetry\n\n$\\rho$\nair density, slugs per cubic foot\n\n$V$\nairspeed, feet per second\n\n$M$\neffective Mach number over span of model\n\n$M_a$\naverage chordwise local Mach number\n\n$M_l$\nlocal Mach number\n\n$R$\nReynolds number of wing based on $\\bar{c}$\n\n$\\alpha$\nangle of attack, degrees\n\n$\\epsilon$\neffective downwash angle, degrees\n\n$q_{\\text{wake}}/q$\nratio of point dynamic pressure at quarter chord of tail mean aerodynamic chord to free-stream dynamic pressure\n\n$(L/D)_{\\text{max}}$\nmaximum ratio of lift to drag\n\n$y_{C_p}$\nlateral center of pressure, percent semispan $\\left( 100 C_B / C_L \\right)$\n\n$h_t$\ntail height relative to wing chord plane extended, percent semispan; positive for tail positions above the chord plane extended", "timestamp": "2026-07-22T06:27:00.274396+00:00"}
{"citation_id": "19930085862", "source_url": "https://ntrs.nasa.gov/api/citations/19930085862/downloads/19930085862.pdf", "page_number": 16, "total_pages": 60, "image_filename": "19930085862_p16.jpg", "text": "14\nNACA RM No. L9A07\n\nThe foregoing comparison was based on the assumption that the\nstick force of the aileron control would be held within range of pilot\ncapabilities through the use of an internal aerodynamic balance and\nthat the aileron deflection would be limited to $\\pm 15^\\circ$ by this balance.\nUnder such conditions the maximum effectiveness of the spoilers can be\nexpected to be as good as or superior to that of the aileron except for\nthe plain wing at high angles of attack. If, however, by employment of\nsome means of power boost aileron deflections up to $\\pm 25^\\circ$ could be\nobtained, the aileron rolling effectiveness would be considerably\nsuperior to that of the spoiler.\n\nCONCLUSIONS\n\nThe results of an investigation in the Langley 19-foot pressure\ntunnel of the characteristics of two types of lateral-control devices\non a $42^\\circ$ sweptback wing with circular-arc airfoil sections and various\nhigh-lift and stall-control devices indicated the following conclusions:\n\n1. The effectiveness of the aileron $C_{l\\delta}$ on the plain wing decreased\nslightly at high angles of attack. At low angles of attack the effective-\nness of the aileron was approximately the same regardless of the flap\nconfiguration. As the angle of attack was increased, however, deflection\nof inboard-located half-span split flaps resulted in a considerable loss\nof aileron effectiveness. The combination of leading-edge flaps and\nstall-control fences tended to offset the detrimental effects which\nresulted when the split flaps were deflected.\n\n2. The aileron hinge-moment characteristics were such that a con-\nventional, sealed, internal aerodynamic balance of approximately 30 per-\ncent of the aileron chord would be required to completely balance the\naileron at low angles of attack of the plain wing configuration. With\nthis amount of balance the aileron probably would be underbalanced at\nhigh angles of attack of the plain wing and at all angles of attack of\nthe flapped configurations.\n\n3. The rolling effectiveness of the spoiler at high angles of\nattack appears largely dependent upon the spoiler location with respect\nto the areas of separated flow on the wing. When stalling occurred\non the outboard portions of the wing an inboard spoiler location was\nmore effective. When stalling occurred inboard, as it did with the\nwing equipped with the stall-control devices, the outboard spoiler\nlocation was more effective.", "timestamp": "2026-07-22T06:27:00.476654+00:00"}
{"citation_id": "19930082546", "source_url": "https://ntrs.nasa.gov/api/citations/19930082546/downloads/19930082546.pdf", "page_number": 41, "total_pages": 65, "image_filename": "19930082546_p41.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:27:00.722314+00:00"}
{"citation_id": "19930082646", "source_url": "https://ntrs.nasa.gov/api/citations/19930082646/downloads/19930082646.pdf", "page_number": 28, "total_pages": 37, "image_filename": "19930082646_p28.jpg", "text": "NACA TN 1980\n27\n\n[Figure: Photograph of an aircraft on water, showing spray from propellers]\n$\\tau = 3.2^\\circ \\quad V = 28 \\text{ mph}$\n\n[Figure: Photograph of an aircraft on water, showing spray from propellers]\n$\\tau = 3.2^\\circ \\quad V = 30.2 \\text{ mph}$\n\n[Figure: Photograph of an aircraft on water, showing spray from propellers]\n$\\tau = 3.4^\\circ \\quad V = 32.3 \\text{ mph}$\n\n[Figure: Photograph of an aircraft on water, showing spray from propellers]\n$\\tau = 3.6^\\circ \\quad V = 34.5 \\text{ mph}$\nNACA\nL-59839\n\nFigure 13.- Spray in propellers during take-off at gross load of 65,000 pounds (modified hull). $\\delta_e = -10^\\circ$.", "timestamp": "2026-07-22T06:27:01.122219+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 25, "total_pages": 149, "image_filename": "19930083192_p25.jpg", "text": "NACA TN 1976\n21\n\nlength of 20 to 28 chords. While such a velocity distribution may\nrepresent the average gust, because of the wide variations in shape, a\nwedge-shaped velocity distribution such that the velocity distribution\nalong the span is uniform and the longitudinal shape is either triangular\nor sinusoidal is recommended for load calculations.\n\nCONCLUDING REMARKS\n\nAlthough much of the information on gust structure can be ration-\nalized to obtain a \"standard\" gust, this material should be used with\ncaution for unconventional types of aircraft. The gust structure of\ninterest is, of course, dependent to some degree on the airplane charac-\nteristics. Many other sizes of gusts exist and those that affect the\nairplane may be only a small part of the random variations that exist in\nthe atmosphere.\n\nAvailable information on the structure of atmospheric gusts has\nshown that, when the gust size and the gust-gradient distance are expressed\nin mean geometric chords, the gust size is independent of airplane,\nweather, topography, and altitude. The probable size of the gust is\n20 chords, and the probable gust-gradient distance for the standard gust\nis about 10 chords. A wedge-shaped gust with the gust velocity uniform\nacross the span and either triangular or sinusoidal in shape with a base\nof 20 chords is believed to be the proper type for most load calculations.\n\nFor a sequence of gusts, the gusts may be of either like or unlike\nsign and will be continuous, with the gust peaks spaced 20 chords apart.\n\nSince the maximum horizontal and vertical gust intensities and\nfrequency distributions are essentially the same within any region of\nrough air, the gust structure obtained for vertical gusts should apply\nequally well for horizontal gusts. Such information is pertinent where\ngust loads are under consideration for diving airplanes or missiles.\n\nSince the data are influenced by the reactions of the airplane, the\napplication to unconventional configurations should be investigated by\na detailed analysis of several possible combinations of gust shape and\nsize.\n\nAIRPLANE REACTIONS\n\nThe second phase of the gust-load problem is that of determining\nthe reaction or forces imposed on an airplane due to a known gust. The\nfactors considered are the aerodynamic coefficients and the possible", "timestamp": "2026-07-22T06:27:02.511599+00:00"}
{"citation_id": "19930085548", "source_url": "https://ntrs.nasa.gov/api/citations/19930085548/downloads/19930085548.pdf", "page_number": 24, "total_pages": 46, "image_filename": "19930085548_p24.jpg", "text": "NACA RM No. E8L30\n23\n\n[Figure: General view of engine setup.]\n\nFigure 3. - General view of engine setup.\n\nNACA\nC-20857\n3-1-48", "timestamp": "2026-07-22T06:27:03.258431+00:00"}
{"citation_id": "19930082592", "source_url": "https://ntrs.nasa.gov/api/citations/19930082592/downloads/19930082592.pdf", "page_number": 44, "total_pages": 50, "image_filename": "19930082592_p44.jpg", "text": "NACA TN 1914\n43\n\n<!-- Image (43, 161, 887, 702) -->\n\nFigure 16. - Oxidation interface of 10-percent-tungsten - titanium carbide ceramal.\nNo preferential oxide penetration along grain boundaries is visible. Temperature,\n1785° F; time at temperature, 50 hours; etchant, potassium hydroxide plus potassium\nferricyanide KOH+K$_3$Fe(CN)$_6$; magnification, X750.\n\nNACA\nC-22915\n2-7-49", "timestamp": "2026-07-22T06:27:05.092468+00:00"}
{"citation_id": "19930082918", "source_url": "https://ntrs.nasa.gov/api/citations/19930082918/downloads/19930082918.pdf", "page_number": 28, "total_pages": 62, "image_filename": "19930082918_p28.jpg", "text": "NACA TN 1940\n27\n\n(7) Bars are numbered 1 through 56; bar 1 represents the extreme bottom\nof ingot and bar 56 the extreme top position\nAll billets were kept in number sequence throughout all processing,\nso that ingot position of any bar can be determined by its number\n\n(8) All bars were cooled on the bed and no anneal or stress relief was\napplied after rolling", "timestamp": "2026-07-22T06:27:05.263927+00:00"}
{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 26, "total_pages": 72, "image_filename": "19930085491_p26.jpg", "text": "NACA RM No. A8J04 CONFIDENTIAL 25\n\nthat the theoretical value should be more closely approached at Reynolds numbers somewhat above those attainable at the small scale of the present tests.\n\nThe value of the lift coefficient for maximum lift-drag ratio of 0.21 is independent of Reynolds number in the range investigated. This result is associated with the increased rate of drag rise near a lift coefficient of 0.20 which, because of the severeness of the adverse pressure gradient due to angle of attack is also independent of Reynolds number.\n\nEffect of Reynolds number on center-of-pressure travel with lift coefficient.— The effect of increased Reynolds number in reducing the center-of-pressure travel is shown in figures 8(c) and (d) where a decrease in total travel of approximately 8 percent of the mean aerodynamic chord is indicated as a result of increasing the Reynolds number from 0.62 to 0.84 million. (The data for a Reynolds number of 0.31 million were omitted since, for this test condition, the temperature effects on the moment strain gage in combination with the relatively small moments result in excessive scatter in the plotted data.) It appears that the favorable decrease in total center-of-pressure travel with increased Reynolds number, like the increase in maximum lift-drag ratio with increased Reynolds number, is due to the decreased areas of separated flow.\n\nProbable effects of higher Reynolds numbers.— The probable changes in the aerodynamic characteristics due to increases in Reynolds numbers above those obtained in the present study may be discussed best by considering two ranges of lift coefficient; namely, the range near zero lift where laminar separation occurs near midchord and the higher lift-coefficient range where laminar separation occurs near the wing leading edge.\n\nIn the lower range of lift coefficients, the line of laminar separation is determined by the rate of pressure recovery behind the line of minimum pressure and the energy level of the laminar boundary layer. If the boundary layer remains laminar, a continued decrease in minimum drag coefficient with increased Reynolds number may be expected for the reasons previously discussed. If, however, boundary-layer transition to turbulent flow occurs ahead of the observed laminar separation line, reference 18 indicates that the value of the right-hand side of equation (5) would become 0.5; that is, the turbulent boundary layer can theoretically withstand a pressure recovery about five times greater than that of the laminar boundary layer before separation occurs. Thus with the pressure\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:27:06.321975+00:00"}
{"citation_id": "19930083221", "source_url": "https://ntrs.nasa.gov/api/citations/19930083221/downloads/19930083221.pdf", "page_number": 27, "total_pages": 47, "image_filename": "19930083221_p27.jpg", "text": "NACA TN No. 1824\n25\n\n$$D = \\frac{1}{2} \\rho_0 \\int_{II} \\int \\left[ (M_0^2 - 1)u^2 + v^2 + w^2 \\right] dydz - \\rho_0 \\int_{III} \\int uv_r r d\\theta dx \\quad (38)$$\n\nwhere $v_r$ is the radial component of the perturbation velocity. No loss in generality results, moreover, if the surface II is moved infinitely distant downstream and the radius of the cylinder is made arbitrarily large. The notation II and III will henceforth refer to this particular configuration.\n\nIf the drag of a lifting surface is to be calculated, the surface and its vorticity wake are replaced by doublet distributions and in that case the integral over region II in equation (38) is called the vortex drag of the body while region III yields the wave drag. It has also been shown (see, for example, reference 17) that the vortex drag of a supersonic wing is a function only of its span load distribution and is equal to the induced drag at subsonic speeds for the same span loading. If a finite nonlifting body is considered, each of the velocity components in region II is attenuated in such a manner that its contribution to the vortex drag is zero. The integration over region III again provides the wave drag for the nonlifting body.\n\nThe combination of the results given in this and the last section provides a method for finding the wave drag of an arbitrary body. The first step is the determination of the source-sink or doublet distribution corresponding to the body and then, by means of the principle of equivalent positions, the sources or doublets are moved to the x axis. The wave drag is then calculated from equation (38) once the induced velocities on the control surface are known. In the next section the wave drag will be written in a different form and the drag at sonic speeds will be investigated.\n\n[Figure: System of axes in transformation equation (39). Axes labeled y, x, $\\xi$, $\\eta$, and angle $\\mu$.]\n\nFigure 12.— System of axes in transformation equation (39).\n\nThis analysis will also provide some insight into the range of validity of the sonic theory.\n\nEvaluation of wave drag as $M_0$ approaches one.— In order to study the drag of a symmetrical body at zero angle of attack, it is convenient to consider the general expression for the velocity potential given in equation (29). Introducing first the transformation (fig. 12)", "timestamp": "2026-07-22T06:27:11.132094+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 60, "total_pages": 98, "image_filename": "19930086073_p60.jpg", "text": "```markdown\n58\nNACA RM A9H04\n\n<!-- Image (113, 96, 866, 828) -->\n\n$$ \\delta_{a_L} = +10.8 $$\n$$ \\delta_{a_R} = -10.8 $$\nAileron deflection, $\\delta_a$, deg\n\n(c) $C_L$ vs $C_m$.\n\nFigure 12.— Continued.\n```", "timestamp": "2026-07-22T06:27:17.506269+00:00"}
{"citation_id": "19930085890", "source_url": "https://ntrs.nasa.gov/api/citations/19930085890/downloads/19930085890.pdf", "page_number": 7, "total_pages": 26, "image_filename": "19930085890_p7.jpg", "text": "6\nNACA RM No. E9C11\n\n$T_e$ exit temperature, $^\\circ$K (reference 1)\n$T_c$ combustion temperature, $^\\circ$K (reference 1)\n\nPressure corrections were made by determining from theoretical data the effect on performance of small deviations in combustion pressure from the design value of 300 pounds per square inch absolute. In the pressure region of 300 pounds per square inch absolute a 1-pound-per-square-inch change in pressure from the design value results in a change in specific impulse of approximately 0.106 pound-second per pound. For combustion pressures of approximately 300 $\\pm$15 pounds per square inch absolute, the maximum pressure correction of the specific-impulse value is less than 1 percent.\n\nIt is desirable to compare experimental performance with theoretically obtained values. Theoretical values of specific impulse are usually determined for such ideal conditions as pure compounds for propellants, parallel flow of exhaust gases in nozzles, equilibrium composition or frozen composition expansion, isentropic expansion, and no effect of the jet on external pressure. None of these ideal conditions exist in the actual case. Estimation of some of the more obvious deviations between the assumed and actual conditions, however, is possible. The diborane used for the experimental investigation consisted of approximately 95-percent diborane with the remainder probably ethane and ethyl ether. The theoretical performance of pure diborane with liquid oxygen is reported in reference 1. The theoretical performance of ethane and ethyl ether with liquid oxygen was unavailable. In order to correct the theoretical specific impulse of the fuel used, however, the impurity was considered to be propane, a related hydrocarbon, for which theoretical performance calculations were available. If it is assumed that the fuel used was 95-percent diborane and 5-percent propane, the theoretical specific impulse was estimated as 2 percent lower than the performance of pure diborane with liquid oxygen. The correction for the deviation of the flow from the assumed axial direction through the exhaust nozzle reduces the theoretical performance by approximately 2 percent (reference 3). These two factors would therefore reduce the theoretical values of specific impulse by about 4 percent.\n\nFor the purpose of determining the volume specific impulse, the density for diborane was chosen from the expression (reference 4, p. 559)\n\n$$d = 0.3140 - 0.001296t \\ ^\\circ C$$", "timestamp": "2026-07-22T06:27:17.767792+00:00"}
{"citation_id": "19930082566", "source_url": "https://ntrs.nasa.gov/api/citations/19930082566/downloads/19930082566.pdf", "page_number": 39, "total_pages": 44, "image_filename": "19930082566_p39.jpg", "text": "NACA TN No. 1889\n37\n\nStrength ratio, $\\sigma_1'/\\sigma_{1u}$\nStrength ratio, $\\sigma_2'/\\sigma_{2u}$\n\n$\\circ$ N = $1 \\times 10^5$\n$\\triangle$ N = $5 \\times 10^5$\n$\\square$ N = $1 \\times 10^6$\n$\\oslash$ N = $5 \\times 10^6$\n\nPrincipal stress ratio\n$\\sigma_2/\\sigma_1 \\longrightarrow$\n$\\longleftarrow \\sigma_1/\\sigma_2$\n\n$\\sigma_1'$\n$\\sigma_2'$\n\n$\\sigma_{1u}$\n$\\sigma_{2u}$\n\nNACA\n\nFigure 15.- Comparison of biaxial fatigue and biaxial ultimate strengths.", "timestamp": "2026-07-22T06:27:18.262277+00:00"}
{"citation_id": "19930085900", "source_url": "https://ntrs.nasa.gov/api/citations/19930085900/downloads/19930085900.pdf", "page_number": 3, "total_pages": 33, "image_filename": "19930085900_p3.jpg", "text": "2\nCONFIDENTIAL\nNACA RM L9D20\n\nIn the present investigation, the effects of spacing of jets, length\nof jet rows, direction of jets, and amount of air flow on the hydro-\ndynamic characteristics were determined for one of the better chine con-\nfigurations of reference 1. Jets in the form of V-steps pointed forward\nwere also investigated for a closer jet spacing than that used in\nreference 1. The effects of substituting narrow breaker strips for the\nrows of jets in the chine configuration and for the rows of jets in the\nthree multiple-step configurations reported in reference 1 were also\ninvestigated.\n\nDESCRIPTION OF MODEL\n\nThe model was a $\\frac{1}{12}$-size model of the streamline fuselage of a hypo-\nthetical transonic airplane (see figs. 1 and 2) and was the same model\ndescribed in reference 1. The center of gravity was located 0.43 inch\nbelow the center line at station 21.22. (Distances from the nose\nmeasured along the center line are designated as stations.) The length\nof the model was 42.22 inches and the maximum diameter was 5 inches.\n\nStainless-steel tubes of 0.026-inch inside diameter and spaced\n1/4 inch apart were inserted into the bottom of the model in rows\nsimulating chines as shown in figures 1 and 2(a). Two sets of tubes\nwere inserted; one in which the tubes were perpendicular to the center\nline and one in which they were slanted aft at an angle of 45°. A plan\nview of these simulated chine configurations is shown in figure 2(a).\nAdditional rows of perpendicular jets were inserted to form the pattern\nsimulating the multiple steps shown in figure 2(b). These jets were\nalso spaced 1/4 inch apart.\n\nThe basic model was also modified by $\\frac{1}{16}$-inch wide strips of tri-\nangular cross section arranged in all the patterns shown in figure 2.\nThe size of the strips relative to the model is shown in figure 3.\n\nTwo types of strips were arranged in each multiple-step pattern.\nIn one type, the forward side of the strip was perpendicular to the\nsurface of the fuselage with the hypotenuse of the cross section forming\nthe after side; in the other type, these conditions were reversed.\n\nFigure 2 shows the multiple-step configurations arranged as\neight V-steps pointed forward, as eight V-steps pointed aft, and as\nnine transverse steps having no V-angle.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:27:22.870768+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 54, "total_pages": 114, "image_filename": "19930086061_p54.jpg", "text": "50\nNACA RM L9J07\n\nLeft semispan\nRight semispan\nUpper\nLower\nP\n(a) $\\psi = 0^\\circ$\n\nLeft semispan\nRight semispan\nUpper\nLower\nP\n$10^\\circ$\n(b) $\\psi = 10^\\circ$\nNACA\n\nFigure 14.- Pressure distribution about wing 1 at various angles of yaw;\n$\\alpha = 24.1^\\circ$.", "timestamp": "2026-07-22T06:27:24.368542+00:00"}
{"citation_id": "19930082613", "source_url": "https://ntrs.nasa.gov/api/citations/19930082613/downloads/19930082613.pdf", "page_number": 35, "total_pages": 46, "image_filename": "19930082613_p35.jpg", "text": "Page intentionally left blank\n\nPage intentionally left blank", "timestamp": "2026-07-22T06:27:28.061499+00:00"}
{"citation_id": "19930085859", "source_url": "https://ntrs.nasa.gov/api/citations/19930085859/downloads/19930085859.pdf", "page_number": 21, "total_pages": 31, "image_filename": "19930085859_p21.jpg", "text": "NACA RM No. L9B25\n\n| Angle of attack, $\\alpha$, deg | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | 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| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | |", "timestamp": "2026-07-22T06:27:36.763272+00:00"}
{"citation_id": "19930085870", "source_url": "https://ntrs.nasa.gov/api/citations/19930085870/downloads/19930085870.pdf", "page_number": 10, "total_pages": 92, "image_filename": "19930085870_p10.jpg", "text": "NACA RM No. L9D07 CONFIDENTIAL 9\n\nDrag\n\nThe minimum drag coefficients for the 8-percent-thick triangular-wing series are presented in figure 11 for the three Mach numbers and compared with the theoretical pressure drag as predicted from linear theory. The wave drag of the triangular wings of double-wedge section was computed by the method of reference 7 for the three positions of the Mach line, namely, ahead of, between, and behind leading edge and ridge line. The equations used are included in appendix A. Below a value of $\\tan \\epsilon / \\tan m$ of approximately 1.6 the elliptical leading edge produces the lower minimum drag. Above this value the converse is true. This effect might be expected in view of the lessening of the adverse pressure gradient aft of the ridge line predicted by theory for high values of $\\tan \\epsilon / \\tan m$. A similar effect was noted in the lift results (fig. 9) in that the lift-curve slopes of the wedge-leading-edge wings became greater than those of the elliptical-leading-edge wings beyond a value of $\\tan \\epsilon / \\tan m$ of approximately 1.6. Unusually low values of the minimum drags of wing 7 at all Mach numbers were due to the fact that the thickness of this model was only 97 percent of the specified amount. The curves have been faired through a point corrected for this thickness error. It should be noted that for wings of this thickness ratio and range of Reynolds numbers the linear theory is in poor agreement with the test results. As can be seen by adding a reasonable skin-friction-drag increment to the linear-theory values, the best correlation of actual test values and theory occurs at values of $\\tan \\epsilon / \\tan m$ less than 0.7. In any case it is very doubtful that actual test results will achieve the characteristic peaks indicated by the linear theory as the Mach line successively passes over the ridge line and behind the leading edge; rather, a much smoother curve appears to be the physical result.\n\nDrag-Rise Factor\n\nReference 4 shows the theoretical value of the drag-rise factor $\\Delta C_D / C_L^2$ for triangular wings having a subsonic leading edge (velocity component normal to leading edge is subsonic) and realizing leading-edge suction as\n\n$$\n\\frac{\\Delta C_D}{C_L^2} = \\frac{1}{\\left(\\frac{dC_L}{d\\alpha}\\right)} - \\frac{\\beta \\sqrt{1 - w^2}}{4\\pi w}\n$$\n\nwhere $\\alpha$ is in radians.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:27:38.738547+00:00"}
{"citation_id": "19930085899", "source_url": "https://ntrs.nasa.gov/api/citations/19930085899/downloads/19930085899.pdf", "page_number": 5, "total_pages": 29, "image_filename": "19930085899_p5.jpg", "text": "```markdown\n4\nNACA RM No. L9A21\n\nSubscripts and abbreviations:\n\nW wing\n\nWF wing-fuselage\n\na.c. aerodynamic center\n\nTESTS\n\nThe tests were made in the Langley high-speed 7- by 10-foot tunnel utilizing an adaptation of the NACA wing-flow technique for obtaining transonic speeds. The technique used involves placing the model in the high-velocity flow field generated over the curved surface of a bump on the tunnel floor. (See reference 1.)\n\nTypical contours of local Mach number in the vicinity of the model location on the bump are shown in figure 5. It is seen that there is a Mach number gradient of about 0.04 over the 1/4 span at low Mach numbers and from 0.06 to 0.07 at the highest Mach numbers. The chordwise Mach number gradient is generally less than 0.01. No attempt has been made to evaluate the effects of this chordwise and spanwise Mach number variation. Note that the long dashed lines shown near the root of the wing (fig. 5) indicate a local Mach number 5 percent below the maximum value and represent a nominal extent of the bump boundary layer. The effective test Mach number was obtained from contour charts similar to those presented in figure 5 using the relationship\n\n$$ M = \\frac{2}{S} \\int_{0}^{b/2} c M_a dy $$\n\nThe variation of mean test Reynolds number with Mach number is shown in figure 6. The boundaries on the figure are an indication of the probable range in Reynolds number caused by variations in test conditions in the course of the investigation.\n\nForce and moment data, effective downwash angles, and the ratio of dynamic pressure at 25 percent of the tail mean aerodynamic chord to free-stream dynamic pressure were obtained for various model configurations through a Mach number range of 0.60 to 1.18 and an angle-of-attack range of -2° to 10°.\n\nNo terms have been applied to the data to account for the presence of the end plates on the models. Jet-boundary corrections have not been evaluated because the boundary conditions to be satisfied are not rigorously defined. However, inasmuch as the effective flow field is large compared with the span and chord of the model the corrections are believed to be small.\n```", "timestamp": "2026-07-22T06:27:41.085938+00:00"}
{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 27, "total_pages": 72, "image_filename": "19930085491_p27.jpg", "text": "26 CONFIDENTIAL NACA RM No. A6J04\n\nfield shown in figure 6 and a turbulent boundary layer near midchord, no flow separation would be likely to occur on the wing of WF-63 at zero angle of attack. In tests at larger scale of a wing with approximately $63^\\circ$ leading-edge sweepback having a biconvex section and taper ratio of one, Frick and Boyd (reference 20) have shown, through both pressure-distribution measurements and liquid-film studies at a Reynolds number of approximately 1.8 million, that natural boundary-layer transition did occur near midchord. Hence, a similar condition may be expected with the present wing at higher Reynolds numbers. This will cause a reduction in pressure drag, but will also be accompanied by an increase in skin-friction drag. Thus an estimation of the drag of a full-scale configuration operating near zero lift at a Mach number of 1.53 depends upon a comparison of the laminar skin-friction drag and the separation drag at the test Reynolds number with the laminar and turbulent skin-friction drags of the full-scale Reynolds number. Theodorsen and Regier in reference 22 have shown that skin-friction coefficients are independent of Mach number up to at least 1.69. Therefore, at relatively high Reynolds numbers, since the laminar and turbulent skin-friction coefficients both decrease with increasing Reynolds number, it may be expected that the turbulent skin-friction coefficient will be of the same order of magnitude as the laminar skin-friction coefficient at the test Reynolds number. In this higher range of Reynolds numbers, because the separation area and associated drag will have disappeared, it is probable the drag values near the minimum will be less than that for a similar configuration in this study.\n\nIn the range of higher lift coefficients, the pressures due to angle of attack predominate and the flow-separation line in the present tests moved close to the wing leading edge. The most important effect of increasing the Reynolds number in this range of lift coefficients is that of reducing the chordwise extent of the separated bubble which exists immediately behind the line of separation. The possibility of obtaining transition in the boundary layer ahead of the line of separation and thus removing completely the separated bubble at full-scale Reynolds numbers will depend upon the length of run, leading-edge-surface condition, and the strength of the adverse pressure gradient due to the lifting pressure distribution. The reduction or disappearances of the separated area near the leading edge would probably result in an increase in the leading-edge suction and a decrease in the drag-rise factor. This decrease in the magnitude of the drag-rise factor associated with this improvement of flow in conjunction with the probable decrease in minimum drag coefficient would result in a further increase in the maximum lift-drag ratio.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:27:43.381235+00:00"}
{"citation_id": "19930082546", "source_url": "https://ntrs.nasa.gov/api/citations/19930082546/downloads/19930082546.pdf", "page_number": 42, "total_pages": 65, "image_filename": "19930082546_p42.jpg", "text": "NACA TN No. 1870\n41\n\n[Figure: A black and white photograph showing a side view of a large circular steel wall structure mounted on a wooden frame. To the left, a microphone on a stand is positioned near the opening of the wall. In the background, a building and trees are visible. A label in the bottom right corner of the photo reads \"NACA L-56020\".]\n\n(c) Circular steel wall (side view with end stiffener removed) showing reinforcement and microphone supports.\n\nFigure 5.- Concluded.", "timestamp": "2026-07-22T06:27:43.620482+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 61, "total_pages": 98, "image_filename": "19930086073_p61.jpg", "text": "```markdown\nNACA RM A59E04\n\nLift coefficient, $C_L$\n\nRolling-moment coefficient, $C_l$\n\nYawing-moment coefficient, $C_n$\n\nSide-force coefficient, $C_Y$\n\n$\\delta_a$, deg\n$\\square$ $\\delta_{a_L} = +10.8$\n$\\circ$ $0$\n$\\diamond$ $\\delta_{a_R} = -10.8$\n\n[Figure: Three graphs plotting Lift coefficient ($C_L$) against Rolling-moment coefficient ($C_l$), Yawing-moment coefficient ($C_n$), and Side-force coefficient ($C_Y$).]\n\n(d) $C_L$ vs $C_l$, $C_n$ and $C_Y$.\n\nFigure 12.— Concluded.\n\n59\n```", "timestamp": "2026-07-22T06:27:44.089701+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 26, "total_pages": 149, "image_filename": "19930083192_p26.jpg", "text": "22\nNACA TN 1976\n\neffects of stability and elasticity. The purpose of this section is to indicate what is known about these factors by coordinating the available information. In most cases, specific procedures or numbers to be used in load calculations are not obtained. In regard to some factors, the state of available information is unsatisfactory but the factor is considered and its status is indicated.\n\nMETHODS\n\nThe three methods utilized in the study of airplane reactions were an analysis and two experimental methods. The experimental methods consist of tests of models in the Langley gust tunnel and flight tests with full-scale airplanes.\n\nAnalysis\n\nInspection of the equations given in references 14 and 15 indicates that the inclusion of unsteady lift leads to variable coefficients, which result in integral equations when the unknown appears under the integral sign. In all cases the integrals that appear are of the same form and represent the application of the principle of superposition to unit-jump solutions to obtain the response to arbitrary or known disturbances.\n\nThe principle cited is illustrated in figure 22 for a linear variation in angle of attack for which the lift on an airfoil after $s_1$ chords of penetration is desired. In figure 22(a) the angle of attack is assumed to vary directly with $s$ and the development of lift per unit change in angle of attack is assumed similar to that shown in figure 3. The angle-of-attack variation is assumed to be approximated by a series of unit changes in angle of attack superimposed in the $s$ direction. For steady lift, the corresponding variations of lift with $s$ are indicated by the straight line and steps in figure 22(b). The unsteady lift develops for each step in angle of attack according to the dash curves and the approximate lift at $s_1$ is the sum of the contributions of each step at $s_1$. The result at $s_1$ can be written as\n\n$$\n\\Delta C_{L_1} = \\frac{dC_L}{d\\alpha} \\sum_{0}^{s_1} C_{L_\\alpha}(s_1 - s) \\Delta\\alpha = \\frac{dC_L}{d\\alpha} \\sum_{0}^{s_1} C_{L_\\alpha}(s_1 - s) \\frac{\\Delta\\alpha}{\\Delta s} \\Delta s\n$$", "timestamp": "2026-07-22T06:27:46.701992+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 43, "total_pages": 78, "image_filename": "19930082483_p43.jpg", "text": "NACA TN No. 1807\n41\n\nAPPENDIX A\n\nSYMBOLS\n\nThe following symbols are used in this report:\n\n| | |\n| :--- | :--- |\n| A | area, sq ft |\n| a | velocity of sound at observed conditions, ft/sec |\n| B | number of active rotor blades at any instant |\n| b | total number of blades in turbine rotor |\n| C | blade radial tip clearance, ft |\n| c | absolute velocity, ft/sec |\n| $c_p$ | specific heat at constant pressure, Btu/(lb)($^\\circ$F) |\n| D | pitch-line diameter, ft |\n| d | rotor-disk diameter, ft |\n| F | fraction of active nozzle arc |\n| G | fraction of active nozzle arc other than F |\n| g | acceleration due to gravity, 32.174, ft/sec$^2$ |\n| $\\Delta h'$ | enthalpy drop based on ratio of inlet stagnation condition to discharge stagnation conditions, Btu/lb |\n| $\\Delta_g h$ | isentropic enthalpy drop based on ratio of inlet stagnation conditions to discharge static conditions, Btu/lb |\n| $\\Delta_g h'$ | isentropic enthalpy drop based on ratio of inlet stagnation conditions to discharge stagnation conditions, Btu/lb |\n| J | mechanical equivalent of heat, 778, ft-lb/Btu |\n| $K_I$ | percentage of active gas through rotor-tip clearance space, constant for any particular turbine |\n| $K_{II}$ | empirical constant for disk-windage loss |\n| $K_{III}$ | empirical constant for driving-fluid losses |", "timestamp": "2026-07-22T06:27:46.846563+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 55, "total_pages": 114, "image_filename": "19930086061_p55.jpg", "text": "NACA RM L9J07\n51\n\nUpper\nLower\n\nLeft semispan\nRight semispan\n\n(c) $\\psi = 20^\\circ$\n\nUpper\nLower\n\nLeft semispan\nRight semispan\n\n(d) $\\psi = 35^\\circ$\n\nFigure 14.- Concluded.", "timestamp": "2026-07-22T06:27:59.631096+00:00"}
{"citation_id": "19930082646", "source_url": "https://ntrs.nasa.gov/api/citations/19930082646/downloads/19930082646.pdf", "page_number": 29, "total_pages": 37, "image_filename": "19930082646_p29.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:27:59.884381+00:00"}
{"citation_id": "19930085548", "source_url": "https://ntrs.nasa.gov/api/citations/19930085548/downloads/19930085548.pdf", "page_number": 25, "total_pages": 46, "image_filename": "19930085548_p25.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:28:03.674098+00:00"}
{"citation_id": "19930085890", "source_url": "https://ntrs.nasa.gov/api/citations/19930085890/downloads/19930085890.pdf", "page_number": 8, "total_pages": 26, "image_filename": "19930085890_p8.jpg", "text": "```markdown\nNACA RM No. E9C11\n7\n\nusing the temperature of a mixture of dry ice and alcohol ($-72^\\circ$ C)\n(reference 5, p. 1764), which approximately corresponded to the\nexperimental conditions. The density of liquid oxygen was obtained\nfrom reference 5. Values of characteristic velocity and thrust\ncoefficient were obtained from the results of the experimental data\nusing the equations\n\n$$\nC^* = P_c A_t g / W\n$$\n\nand\n\n$$\nC_F = T / P_c A_t\n$$\n\nwhere\n\n$C^*$ characteristic velocity\n\n$P_c$ combustion-chamber pressure, (lb/sq in. absolute)\n\n$A_t$ exhaust-nozzle-throat area, (sq in.)\n\n$g$ gravitational constant, ($ft/sec^2$)\n\n$W$ total propellant flow, (lb/sec)\n\n$C_F$ thrust coefficient\n\n$T$ thrust, (lb)\n\nThese two parameters are frequently used to evaluate rocket per-\nformance.\n\nRESULTS AND DISCUSSION\n\nTypical experimental data obtained during the experiments are\nshown in figure 4. Records of the thrust and the combustion-\nchamber pressure showed that combustion was stabilized between the\nfirst and second seconds of operation, after which time there was\nvery little deviation in the values of thrust and combustion-\nchamber pressure. Figure 5 presents the experimental specific\nimpulse plotted against the ratio of fuel weight to total pro-\npellant weight. Only runs that were steady in operation and free\nof engine failures are presented. A curve is drawn through the\npoints that were obtained with the engine having an $L^*$ of 325 inches\nand an eight-hole injection system. Also shown in figure 5 is one\nspecific-impulse value for this engine operating at a combustion\npressure of 343 pounds per square inch absolute. Other data taken\nwith an engine having a smaller $L^*$ and with one having a different\n```", "timestamp": "2026-07-22T06:28:04.830302+00:00"}
{"citation_id": "19930085900", "source_url": "https://ntrs.nasa.gov/api/citations/19930085900/downloads/19930085900.pdf", "page_number": 4, "total_pages": 33, "image_filename": "19930085900_p4.jpg", "text": "NACA RM L9D20 CONFIDENTIAL 3\n\nAPPARATUS AND PROCEDURE\n\nThe tests were conducted in Langley tank no. 2. The model was arranged on the staff of the towing gear as shown in figure 4. The model was supported at the center of gravity and towed free to rise and free to trim between $0^\\circ$ and $20^\\circ$. A dashpot was used to damp out oscillations in trim. The load on the water was varied with speed assuming a constant aerodynamic lift coefficient for a hypothetical wing. Measurements were taken of resistance, trim, and rise at constant speeds up to 60 feet per second. The effective hydrodynamic lift was calculated by subtracting from the load on the water the static buoyancy corresponding to the immersed volume of the model at rest for the trim and rise measured when up to speed. No data are presented between 60 feet per second and the assumed take-off speed of 70 feet per second, because at these speeds practically all of the model was out of the water and slight variations in wetted surface caused the readings to become erratic.\n\nThe average air flow per jet for the jet configuration was $11 \\times 10^{-5}$ pounds per second (0.055 lb/sec, full-size) except when varied for a few representative speeds to determine the effect of air flow on resistance. The full-scale air flow was computed by dimensionally scaling up the model air flow assuming that all forces varied in the same way as the gravitational forces.\n\nThe jets perpendicular to the center line and arranged in rows simulating chines extending from station 10 to the aft end of the model, were tested with jet spacings of 2 inches, 1 inch, 1/2 inch, and 1/4 inch. The 1/4-inch-spaced jets were also tested for three other lengths extending from the after end of the fuselage forward to stations 18, 26, and 34. The jets slanted aft and arranged in rows simulating chines and the rows of jets simulating multiple steps were tested with the 1/4-inch spacing.\n\nStrips simulating chines were tested for the same lengths as the rows of jets. Strips placed in the multiple-step configuration were tested for V-steps pointed forward, V-steps pointed aft, and transverse steps.\n\nRESULTS AND DISCUSSION\n\nBasic or Unmodified Model\n\nThe resistance, trim, and effective hydrodynamic lift of the basic or unmodified model are shown in figure 5. (See reference 1.) The resistance increased rapidly to 19.5 pounds at 40 feet per second with\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:28:05.344215+00:00"}
{"citation_id": "19930085862", "source_url": "https://ntrs.nasa.gov/api/citations/19930085862/downloads/19930085862.pdf", "page_number": 17, "total_pages": 60, "image_filename": "19930085862_p17.jpg", "text": "NACA RM No. L9A07\n\n4. The spoilers on the plain wing became ineffective in the maximum lift range. The maximum rolling effectiveness of the 0.10c spoilers on the wing equipped with the high-lift and stall-control devices was equivalent to that produced by a total aileron deflection of approximately $35^\\circ$.\n\nLangley Aeronautical Laboratory\nNational Advisory Committee for Aeronautics\nLangley Air Force Base, Va.\n\nREFERENCES\n\n1. Neely, Robert H., and Koven, William: Low-Speed Characteristics in Pitch of a $42^\\circ$ Sweptback Wing with Aspect Ratio 3.9 and Circular-Arc Airfoil Sections. NACA RM No. L7E23, 1947.\n\n2. Fischel, Jack, and Schneiter, Leslie E.: An Investigation at Low Speed of a $51.3^\\circ$ Sweptback Semispan Wing Equipped with 16.7-Percent-Chord Plain Flaps and Ailerons Having Various Spans and Three Trailing-Edge Angles. NACA RM No. L8H20, 1948.\n\n3. Weick, Fred E., and Jones, Robert T.: Résumé and Analysis of N.A.C.A. Lateral Control Research. NACA Rep. No. 605, 1937.\n\n4. Langley Research Department (Compiled by Thomas A. Toll): Summary of Lateral-Control Research. NACA Rep. No. 868, 1947.\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.", "timestamp": "2026-07-22T06:28:07.754137+00:00"}
{"citation_id": "19930082592", "source_url": "https://ntrs.nasa.gov/api/citations/19930082592/downloads/19930082592.pdf", "page_number": 45, "total_pages": 50, "image_filename": "19930082592_p45.jpg", "text": "**Page intentionally left blank**\n\n**Page intentionally left blank**", "timestamp": "2026-07-22T06:28:07.992678+00:00"}
{"citation_id": "19930082566", "source_url": "https://ntrs.nasa.gov/api/citations/19930082566/downloads/19930082566.pdf", "page_number": 40, "total_pages": 44, "image_filename": "19930082566_p40.jpg", "text": "38\nNACA TN No. 1889\n\nStrength ratio, $\\sigma_1'/\\sigma_{1y}$\nStrength ratio, $\\sigma_2'/\\sigma_{2y}$\n\n$\\circ$ N = $1 \\times 10^5$\n$\\triangle$ N = $5 \\times 10^5$\n$\\square$ N = $1 \\times 10^6$\n$\\oslash$ N = $5 \\times 10^6$\n\n$\\sigma_2/\\sigma_1 \\longrightarrow$\n$\\longleftarrow \\sigma_1/\\sigma_2$\nPrincipal stress ratio\n\n$\\sigma_1'$\n$\\sigma_2'$\n\n$\\sigma_{1y}$\n$\\sigma_{2y}$\n\nNACA\n\nFigure 16.- Comparison of biaxial fatigue and biaxial yield strengths.", "timestamp": "2026-07-22T06:28:10.332303+00:00"}
{"citation_id": "19930082546", "source_url": "https://ntrs.nasa.gov/api/citations/19930082546/downloads/19930082546.pdf", "page_number": 43, "total_pages": 65, "image_filename": "19930082546_p43.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:28:10.521091+00:00"}
{"citation_id": "19930082613", "source_url": "https://ntrs.nasa.gov/api/citations/19930082613/downloads/19930082613.pdf", "page_number": 36, "total_pages": 46, "image_filename": "19930082613_p36.jpg", "text": "NACA TN 1938\n35\n\n[Figure: Micrograph showing a cross-section of a cracked material. Labels point to: Nickel plate, Solid scale (surface), Main crack, Branch crack solid scale]\n\nNACA\nC-22631\n12-9-48\n\nFigure 11. - Branch of main crack. Type-B liner; etchant, none; condition, cracked during accelerated-life determination; magnification, X1500. Scale appears to be solid, similar to surface scale. Edge of main crack is plated with nickel.", "timestamp": "2026-07-22T06:28:12.886758+00:00"}
{"citation_id": "19930085870", "source_url": "https://ntrs.nasa.gov/api/citations/19930085870/downloads/19930085870.pdf", "page_number": 11, "total_pages": 92, "image_filename": "19930085870_p11.jpg", "text": "10 CONFIDENTIAL NACA RM No. L9D07\n\nThe last term of this equation accounts for the forward inclination of the resultant force on the wing due to the presence of leading-edge suction. For the case of the triangular wing with supersonic leading edge, this latter term will vanish and the drag-rise factor becomes merely the reciprocal of the lift-curve slope. The difference between the reciprocal of the lift-curve slope and the value $\\Delta C_D/C_L^2$ represents the increment of drag rise due to leading-edge suction. The drag-rise factors for the triangular-wing series are presented in figure 12 for the three Mach numbers and are compared with theory. Experimental values of $\\Delta C_D/C_L^2$ were obtained from the parabola which appeared to fit best the variation of $\\Delta C_D$ with $C_L$. The test results given by the reciprocal of the individual lift-curve slopes are compared with the experimental values of $\\Delta C_D/C_L^2$. For all Mach numbers the experimental $\\Delta C_D/C_L^2$ curves were higher than the theory with leading-edge suction and gave lower values than, but exhibited the same general trend as, the curves of the reciprocal lift-curve slopes. As previously stated, the difference between the experimental $\\Delta C_D/C_L^2$ values and the reciprocal of the lift-curve slopes indicates, according to equation (3), leading-edge suction. On this basis, but contrary to expectations, the greater suction is realized by the wedge-leading-edge wings. The extensive change in leading-edge shape probably introduced phenomena other than leading-edge suction, having such a large effect as to mask the effects of the suction. The method of indicating leading-edge suction based on equation (3) is apparently inadequate for the wings tested. Although leading-edge suction would not be expected for thin, uncambered wings of sharp leading edge, it is possible that the wedge-leading-edge wings may realize some leading-edge suction because of the well-forward location of the maximum-thickness point, the large absolute thickness of the wings, and the resulting large included angle of the wedge leading edge.\n\nThe experimental $\\Delta C_D/C_L^2$ curves for the wedge-leading-edge wings gave a lower value of drag rise, departing from the elliptical-leading-edge values very noticeably as the Mach cone is swept behind the leading edge. Such an effect might possibly be expected from theoretical drag considerations as the elliptical leading edge creates a stronger bow wave or unattached shock. At Mach numbers of 1.92 and 2.40 the experimental curves of $\\Delta C_D/C_L^2$ for the wedge-leading-edge wings show less drag rise at high values of $\\tan \\epsilon/\\tan \\mu$, roughly 1.4 and higher, than that predicted by theory. However, the fact that the theoretical curve assumes no change in the basic form drag and friction drag with angle of attack and does not include viscous effects must, of course, be considered in making any comparison with theory.\n\nLift-Drag Ratio\n\nThe maximum values of lift-drag ratio $(L/D)_{max}$ are presented in figure 13 for the three Mach numbers and compared with the linear theory\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:28:13.512252+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 62, "total_pages": 98, "image_filename": "19930086073_p62.jpg", "text": "```markdown\n1.4\n1.2\n1.0\n.8\n.6\n.4\n.2\n0\n-.2\n\nLift coefficient, $C_L$\n\n0 0 4 8 12 16 20 24 28 32 36\nAngle of attack, $\\alpha$, deg\n\n$\\circ$ $\\square$\n$\\delta_{a_L}=+10.8, \\& \\delta_{a_R}=-10.8$\nAileron deflection, $\\delta_a$, deg\n\n(a) $C_L$ vs $\\alpha$.\n\nFigure 13.— Wing plus body at $00^\\circ$ angle of sideslip with both ailerons deflected.\n\n[NACA logo]\n\n69\nNACA RM A57E04\n```", "timestamp": "2026-07-22T06:28:13.856055+00:00"}
{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 28, "total_pages": 72, "image_filename": "19930085491_p28.jpg", "text": "NACA RM No. A8J04 CONFIDENTIAL 27\n\nSince the distance from the moment axis to the center of pressure at low lift coefficients was reduced by reducing the area of laminar separation, the reduction or disappearance of the separated flow at full-scale Reynolds numbers should result in a more rearward center-of-pressure position near zero lift. This effect is illustrated in figures 8(c) and (d) for WF-63 at Reynolds numbers of 0.62 and 0.84 million. The increase in Reynolds numbers in the higher range of lift coefficients where the line of laminar separation is close to the leading edge will decrease the extent of the laminar bubble. It is probable, therefore, that the center of pressure will have a more forward position in this lift-coefficient range with increasing Reynolds number. Thus it is to be expected that the decrease in total center-of-pressure travel with increase in Reynolds number within the range of the present investigation will be continued to higher Reynolds numbers.\n\nBecause of the high induced angles of attack on the outboard wing sections and the associated highly adverse-pressure gradients (fig. 5), full benefit of increased Reynolds number may not be achieved at lift coefficients near the optimum; that is, the flow may separate even at full-scale Reynolds numbers. A possible solution to this problem has been indicated by Jones in reference 1 where it is shown that camber and washout may be utilized at supersonic speeds to obtain a uniform lift distribution at a specific lift coefficient. Hence, the lifting pressure gradients are neutral and should not promote separation.\n\nEffect of Sweep on Longitudinal Characteristics\n\nThe longitudinal characteristics presented in figure 7 for the various sweep angles investigated are summarized in figure 10 for purposes of discussion. These data were obtained at a constant Reynolds number of 0.62 million, the highest possible that permitted the determination of the maximum lift-drag ratio of each angle of sweep without exceeding the limits of the balance. As in the preceding sections, the effects of sweep will be considered on lift-curve slope, minimum drag coefficient, drag-rise factor, maximum lift-drag ratio, and pitching moment. Because the sweep angle was varied by rotating the wing panels about the midpoint of the root chord, there is an accompanying change in aspect ratio and thickness-chord ratio measured parallel to the plane of symmetry. These changes, it should be noted, very nearly represent the relation which must be followed in practical wing construction to maintain a given structural strength and stiffness. In the following discussion,\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:28:19.821729+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 44, "total_pages": 78, "image_filename": "19930082483_p44.jpg", "text": "```markdown\n42\nNACA TN No. 1807\n\n| Symbol | Definition |\n| :--- | :--- |\n| $l$ | rotor-blade length, ft |\n| M | Mach number |\n| N | rotational speed, rpm |\n| n | thickness coefficient (unity for reaction turbines) |\n| P | turbine-shaft power output, corrected to sea-level conditions, hp |\n| p | absolute pressure, lb/sq ft or in. Hg |\n| R | gas constant, 53.345, ft-lb/(lb)($^\\circ$R) |\n| Re | Reynolds number |\n| r | radius, ft |\n| T | temperature, $^\\circ$R |\n| $\\Delta T_{oil}$ | temperature rise of lubricating oil in bearings |\n| u | velocity, ft/sec |\n| V | gas velocity, ft/sec |\n| $V_j$ | theoretical jet velocity based on ratio of inlet total pressure to discharge static pressure, ft/sec |\n| $V_j'$ | theoretical jet velocity based on ratio of inlet total pressure to discharge total pressure, ft/sec |\n| W | weight flow, lb/sec |\n| z | total number of blades in turbine rotor |\n| $\\beta_2$ | rotor-blade exit angle relative to plane of rotor disk measured at pitch line, deg |\n| $\\gamma$ | ratio of specific heats of gas |\n| $\\delta$ | pressure correction ratio, $p/p_0$ |\n| $\\eta$ | turbine over-all efficiency based on ratio of inlet stagnation conditions to discharge static conditions |\n```", "timestamp": "2026-07-22T06:28:20.031004+00:00"}
{"citation_id": "19930085548", "source_url": "https://ntrs.nasa.gov/api/citations/19930085548/downloads/19930085548.pdf", "page_number": 26, "total_pages": 46, "image_filename": "19930085548_p26.jpg", "text": "154-986-B\n1077\n\nNACA RM No. E8L30\n\nThermocouples\nThermocouple\nExhaust tank\nInlet\nsurge\ntank\nExperi-\nmental\ncylinder\nExhaust-\ngas cool-\ning tank\nExhaust-\npressure\nregulator\nAir-measuring\norifice\nThermocouple\nAir heater\nInlet\nTo\natmosphere\nInlet-air pressure regulator\nNACA\nFigure 4. - Engine setup.\n25", "timestamp": "2026-07-22T06:28:20.647222+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 27, "total_pages": 149, "image_filename": "19930083192_p27.jpg", "text": "```markdown\nNACA TN 1976\n23\n\nor for an analytic expression of differential increments\n\n$$\n\\Delta C_{L_1} = \\frac{dC_L}{d\\alpha} \\int_0^{s_1} C_{L_\\alpha}(s_1 - s) \\frac{d\\alpha}{ds} ds\n$$\n\nThis expression is commonly known as Duhamel's integral and is illustrated by Berg in reference 16. The integral can be evaluated analytically step by step as indicated in figure 22 or graphically by Carson's theorem (reference 17).\n\nAs previously mentioned, the type of equation obtained is of the form\n\n$$\nC_{L_{w1}} = A \\int_0^{s_1} C_{L_g}(s_1 - s) \\frac{d\\alpha}{ds} ds - B \\int_0^{s_1} C_{L_\\alpha}(s_1 - s) \\frac{dC_{L_w}}{ds} ds\n$$\n\nwhere $C_{L_{w1}}$ is the local value of $C_{L_w}$ at $s_1$. The difficulty in solving the equation arises from the fact that the relation between $C_{L_w}$ and $s$ must be known before the second integral on the right side can be evaluated. The equation shown is a simple case and the more complete equations of motion consist of additional integrals and, in some cases, derivatives of integrals.\n\nSolutions of equations of the type shown have been obtained by methods of successive approximation, Fredholm's solution (reference 6), Laplace transforms (reference 14), or by assuming, as in reference 18, that the function is known. The need for solutions for arbitrary disturbances and the complicated nature of solutions generally lead to graphical or approximate methods in all but a few cases.\n\n### Gust-Tunnel Testing\n\nThe Langley gust tunnel (fig. 23) was built to permit the determination of airplane reactions and other pertinent quantities under controlled and known conditions. It consists of a catapult to launch a dynamically scaled airplane model into steady level flight through a vertical jet of air having characteristics that are under control, a means of catching the model, and, finally, suitable equipment to record the required\n```", "timestamp": "2026-07-22T06:28:22.299944+00:00"}
{"citation_id": "19930085862", "source_url": "https://ntrs.nasa.gov/api/citations/19930085862/downloads/19930085862.pdf", "page_number": 18, "total_pages": 60, "image_filename": "19930085862_p18.jpg", "text": "```markdown\n16\n\n68.25\n\nRoot\n25-(50)(05)-(50)(05)\n\n42.05°\n\n21.95\n36.80\n42.00\n\nLine of maximum\nthickness\n\nA\n\n38.19°\n\n90°\n\n5/4\n\nA\n\n26.25\n\nTip\n25-(50)(03.2)-(50)(03.2)\n\n87.18\n\nSection A-A (enlarged)\n\nFigure 1.- Geometry of wing. All dimensions in inches.\n\nNACA\n\nNACA RM NO. L9H07\n```", "timestamp": "2026-07-22T06:28:24.508681+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 56, "total_pages": 114, "image_filename": "19930086061_p56.jpg", "text": "52\nNACA RM L9J07\n\nUpper\nLower\n\nLeft semispan\nRight semispan\n\n(a) $\\psi = 0^\\circ$\n\nUpper\nLower\n\nLeft semispan\nRight semispan\n\n(b) $\\psi = 10^\\circ$\n\nFigure 15.- Pressure distribution about wing 1 at various angles of yaw;\n$\\alpha = 34.1^\\circ$.", "timestamp": "2026-07-22T06:28:25.500645+00:00"}