PDF/ocr_dataset_1063.jsonl

{"citation_id": "19930085859", "source_url": "https://ntrs.nasa.gov/api/citations/19930085859/downloads/19930085859.pdf", "page_number": 14, "total_pages": 31, "image_filename": "19930085859_p14.jpg", "text": "```markdown\n12\n\nFloating-tail geometry\nTwice semispan area 0.0178 sq ft\nAspect ratio 4.0\nTaper ratio 0.60\n\nChord plane\n$\\alpha = 0^\\circ$\n\n1.0\n1.0\n1.0\n1.0\n\nSection B-B\n$\\frac{1}{16}$\nDouble scale\n\nStation A\n4.53\n(MAC).80\n.25(MAC) Model\nBump surface\n.40\n$\\frac{1}{8}$ Diameter\nPivot center\n\n.25-chord line\n.00\nB\nB\n1.60\n.76\n$45^\\circ$\n\n0 1 2\nScale, inches\n\n[Figure: NACA logo]\n\nFigure 3.- Details of free-floating tails used in surveys behind a model with $35^\\circ$ sweptback wing,\naspect ratio 4, taper ratio 0.6, and NACA 65A006 airfoil.\n\nNACA RM NO. L9B25\n```", "timestamp": "2026-07-22T06:22:01.605181+00:00"}
{"citation_id": "19930083221", "source_url": "https://ntrs.nasa.gov/api/citations/19930083221/downloads/19930083221.pdf", "page_number": 21, "total_pages": 47, "image_filename": "19930083221_p21.jpg", "text": "```markdown\nNACA TN No. 1824\n19\n\n$$\nN_3 = \\frac{k\\alpha}{\\pi^2 M_0} \\int_{0}^{v} \\frac{c_0 \\sqrt{2}}{M_0} \\left[ \\frac{1}{\\sqrt{(v-v_1)(u+v_1)}} \\right] \\times\n$$\n\n$$\n\\left\\{ \\sin^{-1} \\frac{(u+v_1)(1-M_0)[(v-v_1)(1+M_0)-c_0\\sqrt{2}] + 2v_1[v(1+M_0)-c_0\\sqrt{2}-u(1-M_0)]}{[v(1+M_0)-c_0\\sqrt{2} + v_1(1-M_0)] [u(1-M_0) - v_1(1+M_0)]} \\right\\} dv_1\n$$\n\nwhere\n$$\nu = \\frac{1}{\\sqrt{2}} (t-x)\n$$\n$$\nv = \\frac{1}{\\sqrt{2}} (t+x)\n$$\n\n$$\n\\left. \\begin{array}{l} F(\\psi, k) \\\\ E(\\psi, k) \\end{array} \\right\\} \\text{incomplete elliptic integrals}\n$$\n\nK, E complete elliptic integrals\n\nIn figure 9(a) the growth of pressure distribution with time is shown at subsonic speed for the period of time covered by equations (27). For purposes of comparison, pressure changes calculated from equations (18) are shown in figure 9(b) for supersonic flight velocities.\n\nEquations (27) suffice to determine the initial growth of indicial lift coefficient at subsonic speeds. Such results were given in figure 3 at $M_0=0.8$ along with the calculated growth for about one chord length of travel at $M_0=0.4$. The value of $C_{L_\\alpha}(t)$ at $t=0$ is, for all flight speeds, equal to $4/M_0$.\n\n[Figure: Two diagrams showing pressure distribution on wings. Left diagram labeled (a) Subsonic, right diagram labeled (b) Supersonic. Both show $\\Delta p/q$ on the y-axis and x on the horizontal axis, with shaded regions indicating pressure distribution over time steps.]\n\nFigure 9.- Pressure distribution on wings receiving sudden angle-of-attack change at $t = 0$.\n\nExpressions for $C_{L_\\alpha}(t)$ are as follows:\n\nFirst time interval $0 < t < \\frac{c_0}{1+M_0}$\n```", "timestamp": "2026-07-22T06:22:02.227593+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 37, "total_pages": 78, "image_filename": "19930082483_p37.jpg", "text": "NACA TN No. 1807\n35\n\nDisk loss cannot be presented clearly in figure 10 because it is too small.\n\nThe corrected specific net power, which is ideal power minus all the losses, is shown in figure 10 for each of the three turbine configurations. The specific net power when multiplied by the weight flow is in agreement with the corrected turbine output for full and $120^\\circ$ admission shown in figure 8(b).\n\nFull-Admission Performance\n\nThe turbine operating characteristics for full-admission operation are presented as a composite plot, constructed as previously outlined in the section entitled METHODS. The turbine design operating conditions are indicated in figure 8(a) as point D, which falls well within the turbine operating range covered in this investigation. The turbine design operating point D is seen to be within 1/2 point of the maximum efficiency observed. The island formed by the peak operating efficiency contour of 76 percent is centered at a corrected rotor speed of 9500 rpm at a total-pressure ratio of 2.2. The power output at this point is 412 horsepower corrected to sea level, which corresponds to approximately 3250 horsepower at full-scale operation.\n\nOperation outside of this island is accompanied by decreasing efficiencies. However, when changes in pressure ratio are accompanied by adjustments in rotor speed to maintain the most favorable blade-to-jet speed ratio, the gradient of efficiency is minimized to a drop of 3 points over the entire pressure-ratio range.\n\nThe range of turbine power output is represented in figure 8(a) as vertical distances, which are seen to increase to the right on this plot. This increase is due to the effects of pressure-ratio changes being most pronounced at higher rotor speeds.\n\nPartial-Admission Performance\n\nThe performance characteristics of this turbine at $120^\\circ$ admission over the same range of operating conditions as at full admission are presented in the form of a carpet plot in figure 8(b). This carpet is superimposed on the one previously presented in figure 8(a) for full admission to enable ready comparison of operation with full and partial admission.", "timestamp": "2026-07-22T06:22:09.805073+00:00"}
{"citation_id": "19930085485", "source_url": "https://ntrs.nasa.gov/api/citations/19930085485/downloads/19930085485.pdf", "page_number": 22, "total_pages": 26, "image_filename": "19930085485_p22.jpg", "text": "CONFIDENTIAL\n\nPitching-moment coefficient, $C_m$\n\nLift coefficient, $C_L$\n\nMach number, $M$\n\nDrag coefficient, $C_D$\n\n$\\alpha$ (deg)\n-2.0\n0\n2.0\n\nFigure 13.- Variation of the aerodynamic characteristics of the model with Mach number. Circular-arc aileron; $\\delta_a = 0^\\circ$.\n\nCONFIDENTIAL\n\nPitching-moment coefficient, $C_m$\n\nLift coefficient, $C_L$\n\nMach number, $M$\n\nDrag coefficient, $C_D$\n\n$\\alpha$ (deg)\n-2.9\n0\n2.8\n4.8\n\nFigure 14.- Variation of the aerodynamic characteristics of the model with Mach number. Flat-sided aileron; $t = 0.50$; $\\delta_a = 0^\\circ$.\n\nCONFIDENTIAL\n\nNACA RM No. L5K02", "timestamp": "2026-07-22T06:22:19.411373+00:00"}
{"citation_id": "19930082592", "source_url": "https://ntrs.nasa.gov/api/citations/19930082592/downloads/19930082592.pdf", "page_number": 37, "total_pages": 50, "image_filename": "19930082592_p37.jpg", "text": "**Page intentionally left blank**\n\n**Page intentionally left blank**", "timestamp": "2026-07-22T06:22:22.313588+00:00"}
{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 21, "total_pages": 72, "image_filename": "19930085491_p21.jpg", "text": "20 CONFIDENTIAL NACA RM No. A8J04\n\nwing-thickness distribution have also been neglected in the determination of the theoretical value of $C_{D_{\\min}}$. It is believed, however, that these factors are relatively small as compared to the effect of separation just discussed.\n\nDrag due to lift.— The experimental drag due to lift in terms of the drag-rise factor $\\Delta C_D / (\\Delta C_L)^2$ has been determined from plots of $C_D - C_{D_{\\min}}$ as a function of $(C_L - C_{L_{D=\\min}})^2$ shown by curve (1) in figure 11(a) for the basic configuration. Also shown in this figure by curve (2) is the theoretical drag due to lift and by curve (3) the drag due to lift that would result if the experimental resultant force vector was perpendicular to the wing chord. Comparison of curve (1) with (2) indicates that the experimental drag due to lift is much greater than predicted by the inviscid theory. However, curves (1) and (3) indicate that the benefits of leading-edge suction are partially realized experimentally particularly in the low lift-coefficient range. This factor is apparent by considering the variations of the parameters $\\Delta C_L / \\Delta \\alpha$ and $k_a$ which, as shown in the section Theoretical Considerations, determine the drag due to lift. These parameters may be related by the following equation which is similar to equation (1), but does not require a linear lift curve and parabolic drag curve:\n\n$$\n\\frac{\\Delta C_D}{(\\Delta C_L)^2} = \\frac{k_a}{\\Delta C_L / \\Delta \\alpha}\n\\tag{6}\n$$\n\nFigure 11(b) shows the variations of $\\Delta C_L / \\Delta \\alpha$ and $k_a$ with both $(\\Delta C_L)^2$ and $C_L$. In the range of lift coefficients from 0 to 0.09, the values of $\\Delta C_L / \\Delta \\alpha$ and $k_a$ are constant since, as accurately as could be determined, the lift and drag curves were linear and parabolic, respectively, in this range. At $C_L = 0.09$, where as previously discussed, the laminar boundary-layer flow separation line moved abruptly forward to the leading-edge region on the upper surface, there was an increase in the value of $k_a$ which indicates a loss in leading-edge suction. It is noteworthy, however, that the increase in lift-curve slope and $\\Delta C_L / \\Delta \\alpha$, because of the reattached turbulent boundary-layer flow over the rear of the wing, offsets the loss in leading-edge suction and results in a constant value of $\\Delta C_D / (\\Delta C_L)^2$ (fig. 11) up to approximately the optimum lift coefficient of 0.21. Above this value of lift coefficient, figure 11(a) shows an increase in drag-rise factor. Figure 11(b) indicates that this is due to increased values of $k_a$ which probably result from the larger areas of separated\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:22:22.786157+00:00"}
{"citation_id": "19930085862", "source_url": "https://ntrs.nasa.gov/api/citations/19930085862/downloads/19930085862.pdf", "page_number": 10, "total_pages": 60, "image_filename": "19930085862_p10.jpg", "text": "8\nNACA RM No. L9A07\n\nFrom an inspection of the basic data presented in figures 5 to 10,\nit can be seen that the rolling-moment coefficients obtained from the\nplain wing and the split-flap configurations were approximately the\nsame for either up or down aileron deflections. For the configurations having\nthe leading-edge devices, however, the rolling-moment coefficients pro-\nduced by the ailerons were larger for the up deflections than for the\ndown deflections.\n\nThe rolling-moment coefficients obtained for a total aileron deflec-\ntion of $30^\\circ$ ($15^\\circ$ up and $15^\\circ$ down) on the various wing configurations are\npresented in figure 14. The rolling-moment coefficients produced by\nlarge deflections of the ailerons varied considerably with wing configu-\nration and with angle of attack. At these large aileron deflections the\ntotal rolling-moment coefficients at low angles of attack were approxi-\nmately the same (about 0.03) for all configurations investigated. At\nhigher angles of attack the total rolling-moment coefficients produced\nby large deflections of the aileron on the different configurations varied\nin a manner similar to the aileron effectiveness at small deflections in\nthat the rolling-moment coefficients obtained with the split flap con-\nfiguration decreased rapidly with increasing angle of attack. Only moder-\nate decreases were obtained with the configurations involving the leading-\nedge and split flaps and the fences. With the leading-edge flaps and\nfences but without the split flaps the decrease was slight.\n\nAdverse yawing-moment coefficients were obtained throughout most of\nthe angle-of-attack range, the largest values of which were obtained for\nthe wing without flaps (fig. 14).\n\nPitching-moment characteristics.- The curves of pitching-moment\ncoefficient against angle of attack for the maximum aileron deflections\ninvestigated are presented in figures 5 to 10. It can be seen that\nfor the plain wing and for the wing equipped with split flaps a smaller\nincrement in the pitching-moment coefficient was obtained at positive\nangles of attack with the up aileron than with the down aileron. Con-\nversely, smaller increments in the pitching-moment coefficient were\nobtained with the down aileron than with the up aileron for the wing\nequipped with the leading-edge devices. It is estimated that about $2^\\circ$\nof elevator deflection would be needed to compensate for the maximum\nincrement in pitching-moment coefficient resulting from $25^\\circ$ up and down\ndeflection of a set of ailerons.\n\nHinge-moment characteristics.- In order to illustrate the aileron\nhinge-moment characteristics of the various configurations investigated,\nthe hinge-moment parameters $C_{h\\delta}$ and $C_{h\\alpha}$ and the balance-chamber\nresultant-pressure parameters $P_{R\\delta}$ and $P_{R\\alpha}$ were determined from the", "timestamp": "2026-07-22T06:22:24.451325+00:00"}
{"citation_id": "19930082546", "source_url": "https://ntrs.nasa.gov/api/citations/19930082546/downloads/19930082546.pdf", "page_number": 34, "total_pages": 65, "image_filename": "19930082546_p34.jpg", "text": "NACA TN No. 1870\n33\n\n[Figure: A graph plotting Blade-width ratio, b/D, and blade-thickness ratio, h/b, and Blade angle, $\\beta$, deg, against $r/R_t$. The x-axis ranges from 0.3 to 1.0. The left y-axis ranges from 0 to .20. The right y-axis ranges from 32 to 72. Three curves are plotted, labeled h/b, b/D, and $\\beta$. The NACA logo is present in the bottom right corner of the plot area.]\n\n(c) NACA 4-(5)(08)-03 propeller.\nFigure 3.- Continued.", "timestamp": "2026-07-22T06:22:32.848193+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 19, "total_pages": 149, "image_filename": "19930083192_p19.jpg", "text": "NACA TN 1976\n15\n\nto the spanwise gust distribution. The results of the evaluation of\naccelerometer records for 17 flights are shown in table V, which indi-\ncates the frequency of occurrence of given values of angular accelera-\ntion with the associated values of normal acceleration increment.\n\nLongitudinal gusts.- Rapid airspeed fluctuations have been used to\nevaluate the magnitude of longitudinal gusts. Figure 20, prepared from\nan analysis of the XC-35 airplane data, compares the maximum values of\nthe horizontal gust velocities obtained from the rapid airspeed fluctua-\ntions with the maximum vertical gust velocities. Each point represents\na separate traverse or run. The results of a detailed analysis of air-\nspeed and acceleration records in gusty air from the XBM-1 and Aeronca C-2\nairplane investigations are shown in figure 21 and indicate, for both\nairplanes, the relation of the horizontal gust velocities to the vertical\ngust velocities for equal frequencies of occurrence.\n\nACCURACY OF RESULTS\n\nConsideration of instrumental and reading errors in evaluating the\ndata from the several airplanes, together with a knowledge of the problems\ninvolved in the reactions of an airplane, indicates that the possible\nerrors in the derived quantities that are of interest are approximately\nas follows:\n\nGust velocity based on acceleration, percent . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 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. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .", "timestamp": "2026-07-22T06:22:34.039884+00:00"}
{"citation_id": "19930085890", "source_url": "https://ntrs.nasa.gov/api/citations/19930085890/downloads/19930085890.pdf", "page_number": 2, "total_pages": 26, "image_filename": "19930085890_p2.jpg", "text": "E\n\nNACA RM No. E9C11\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nRESEARCH MEMORANDUM\n\nEXPERIMENTAL INVESTIGATION OF LIQUID DIBORANE - LIQUID OXYGEN PROPELLANT COMBINATION IN 100-POUND-THRUST ROCKET ENGINE\n\nBy William H. Rowe, Paul M. Ordin, and John M. Diehl\n\nSUMMARY\n\nThe specific impulse of liquid diborane and liquid oxygen over a range of mixture ratios was determined in a 100-pound-thrust rocket engine operating at a combustion-chamber pressure of 300 pounds per square inch absolute.\n\nA faired curve through the experimental data had approximately the same shape as the theoretical curve with a maximum uncorrected experimental specific impulse of 249 pound-seconds per pound at a ratio of fuel weight to total propellant weight of 0.37. When corrected for heat rejection, this value increased to 274 pound-seconds per pound, which is 92 percent of the theoretical value of 299 pound-seconds per pound based on equilibrium expansion for the fuel and the nozzle used. The maximum experimental volume specific impulse was $182 \\times 62.4$ pound-seconds per cubic foot and occurred at a ratio of fuel weight to total propellant weight of 0.25; the corrected maximum experimental value was $199 \\times 62.4$ pound-seconds per cubic foot at the same mixture ratio. These experimental values were based on a characteristic length (ratio of combustion-chamber volume to exhaust-nozzle-throat area) of 325 inches. When the characteristic length was reduced from 325 to 159 inches, a small decrease in performance occurred. No apparent change in specific impulse was observed for the two types of injection used for the experiments.\n\nA limited number of temperature- and shock-sensitivity experiments were made with diborane. No explosions nor detonations were observed.\n\nINTRODUCTION\n\nBoron hydrides are of interest as rocket fuels principally because of the possibility of obtaining high specific impulse. As part of the NACA program on high-energy rocket fuels, theoretical", "timestamp": "2026-07-22T06:22:35.562269+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 54, "total_pages": 98, "image_filename": "19930086073_p54.jpg", "text": "52\n\nLift coefficient, $C_L$\n\n1.4\n1.2\n1.0\n.8\n.6\n.4\n.2\n0\n\n0 0 0 0 4 8 12 16 20 24 28 32 36\nAngle of attack, $\\alpha$, deg\n\n$\\circ$ $\\square$ $\\diamond$ $\\triangle$\n0.0 6.0 12.0 15.9\nAngle of sideslip, $\\beta$, deg\n\n(a) $C_L$ vs $\\alpha$.\n\nFigure 11.— Wing plus body at various angles of sideslip with flaps deflected 45.4°.\n\nNACA RM A9E04", "timestamp": "2026-07-22T06:22:37.678753+00:00"}
{"citation_id": "19930085859", "source_url": "https://ntrs.nasa.gov/api/citations/19930085859/downloads/19930085859.pdf", "page_number": 15, "total_pages": 31, "image_filename": "19930085859_p15.jpg", "text": "NACA RM No. L9B25\n13\n\n[Figure: Photograph of a model with 35° sweptback wing, aspect ratio 4, taper ratio 0.6, and NACA 65A006 airfoil showing free-floating tail mounted in fuselage.]\n\nFigure 4.- Photograph of a model with 35° sweptback wing, aspect ratio 4, taper ratio 0.6, and NACA 65A006 airfoil showing free-floating tail mounted in fuselage.", "timestamp": "2026-07-22T06:22:42.575847+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 38, "total_pages": 78, "image_filename": "19930082483_p38.jpg", "text": "36\nNACA TN No. 1807\n\nOf primary significance are the reductions in over-all turbine efficiency encountered with reduced admission. The data indicate that a drop of 16 points in the peak efficiency is obtained when the active nozzle arc is decreased to 120°. The drop is caused by the special losses that are introduced when only a fraction of the nozzle arc is active. When the turbine power output is controlled by nozzle cut-out, the velocity diagrams for equivalent operating conditions at reduced admissions remain largely unchanged, an advantageous feature in power regulation by means of active-nozzle-arc control. This effect is demonstrated in figure 8(b), where the performance patterns over the range of rotor speed and pressure ratio are similar for full and 120° admissions. Because the over-all efficiency is based on the turbine net power output, increases in the power extraction caused by the additional partial-admission losses are evidenced by a change in the position of the efficiency contours. Rapid increases of the driving-fluid and pumping losses with rotor speed cause the region of peak performance to be shifted toward lower rotor speeds with increasing nozzle cut-out, and the adjustments in rotor speed required to maintain optimum blade-to-jet speed ratios with changes in rotor speed remain virtually unchanged. In the upper speed region, the change in net turbine power per pound of driving fluid in response to a given change in pressure ratio is diminished with reduction of the active nozzle arc. The reduced net power range is attributed to the more pronounced effects of the partial-admission losses with increasing rotor speeds.\n\nPartial Admission as a Means of Power Control\n\nThe application of partial admission as a means of power control is evaluated in figure 11 in the form of a curve of over-all turbine efficiency plotted against the percentage of full-power output. At full, 180°, and 120° admission, the values expressed are based on experimental data, the remainder of the curve was constructed using the performance-prediction technique previously described. The curve indicates that, for a constant turbine operating condition, partial admission represents an effective means of power control up to 180° of nozzle-arc reduction. Power control by cut-out of more than one-half of the nozzles is accompanied by a prohibitive drop in over-all efficiency due to the increasingly severe effect of the turbine losses. For other turbine operating conditions, the general shape of this curve is similar except that the performance falls off more sharply with the intensifying effect of the partial-admission losses at higher speeds. This effect, however, is not too important because a primary feature of partial-admission power control is that the turbine operating conditions", "timestamp": "2026-07-22T06:22:43.581171+00:00"}
{"citation_id": "19930082646", "source_url": "https://ntrs.nasa.gov/api/citations/19930082646/downloads/19930082646.pdf", "page_number": 21, "total_pages": 37, "image_filename": "19930082646_p21.jpg", "text": "```markdown\n20\nNACA TN 1980\n\nO Step in\n$\\square$ Step out\n$\\triangle$ Sternpost in\n\nTrim ———\nSpeed -----\nRise - - - -\n\n<!-- Image (114, 100, 826, 803) -->\n\n(a) Warped forebody and\nextended afterbody.\n\n(b) Basic forebody and\nbasic afterbody.\n\nFigure 9.- Variation of trim, rise, and speed with time during landings\nin smooth water.\n```", "timestamp": "2026-07-22T06:22:44.494448+00:00"}
{"citation_id": "19930085847", "source_url": "https://ntrs.nasa.gov/api/citations/19930085847/downloads/19930085847.pdf", "page_number": 16, "total_pages": 32, "image_filename": "19930085847_p16.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:22:51.496658+00:00"}
{"citation_id": "19930083221", "source_url": "https://ntrs.nasa.gov/api/citations/19930083221/downloads/19930083221.pdf", "page_number": 22, "total_pages": 47, "image_filename": "19930083221_p22.jpg", "text": "20\nNACA TN No. 1824\n\n$$C_{L_\\alpha}(t) = \\frac{4}{M_o c_o} [c_o - t(1-M_o)] \\quad (28a)$$\n\nSecond time interval $\\frac{c_o}{1+M_o} < t < \\frac{c_o}{1-M_o}$\n\n$$C_{L_\\alpha}(t) = \\frac{4}{\\pi M_o c_o} \\left\\{ \\frac{4c_o - 3t(1-M_o^2)}{1+M_o} \\text{ arc tan } \\sqrt{\\frac{2c_o - t(1-M_o^2)}{2t(1+M_o) - 2c_o}} \\right.$$\n\n$$+ \\pi \\left( t - \\frac{2c_o}{1+M_o} \\right) + \\frac{1+3M_o}{(1+M_o)^2} \\sqrt{[2t(1+M_o) - 2c_o] [2c_o - t(1-M_o^2)]}$$\n\n$$+ \\frac{4c_o - 2t(1+M_o)}{1+M_o} \\text{ arc sin } \\sqrt{\\frac{t(1+M_o) - c_o}{t(1+M_o)}} - \\frac{1-M_o}{1+M_o} \\sqrt{c_o t(1+M_o) - c_o^2}$$\n\n$$+ [2c_o - t(1+M_o)] \\text{ arc tan } \\sqrt{\\frac{c_o}{t(1+M_o) - c_o}} \\Bigg\\} + \\frac{1}{c_o} \\int_{\\frac{2c_o}{1+M_o} - t}^{c_o - M_o t} \\left( \\frac{\\Delta p}{q \\alpha} \\right)_V dx \\quad (28b)$$\n\nwhere $\\left( \\frac{\\Delta p}{q \\alpha} \\right)_V$ is given by equation (27e).\n\nPART III - THREE-DIMENSIONAL LINEAR PROBLEMS FOR $M_o$ NEAR ONE\n\nSteady State\n\nGeneral solutions for arbitrary Mach numbers.- Two methods of attack are available for the solution of linearized problems at sonic speeds. In the first place, solutions to equation (13) can be written formally and the extent to which these solutions satisfy the original assumption can then be investigated. In the second place, general solutions of equation (7) can be studied in the limit as $M_o$ approaches 1. Since this latter method furnishes added information concerning the variation of the variables with $M_o$, it will be used first.", "timestamp": "2026-07-22T06:22:52.335362+00:00"}
{"citation_id": "19930085870", "source_url": "https://ntrs.nasa.gov/api/citations/19930085870/downloads/19930085870.pdf", "page_number": 4, "total_pages": 92, "image_filename": "19930085870_p4.jpg", "text": "NACA RM No. L9D07 CONFIDENTIAL 3\n\nThe present tests were made to determine the effects of giving a generous curvature to the leading edge of a series of triangular wings with the object of realizing a greater proportion of theoretical leading-edge suction and thereby increasing the wing efficiency. These tests extend the investigations initiated in reference 10 to wings of higher thickness ratio believed practical for full-scale aircraft. Two series of 11 triangular wings each were tested in the Langley 9-inch supersonic tunnel at Mach numbers of 1.62, 1.92, and 2.40. Except for leading-edge shape, the first and second series were identical. The thickness ratio of 8 percent was constant for all these wings as was the 18-percent location of maximum-thickness point. The apex half-angles ranged from $10^\\circ$ to $45^\\circ$, covering the range of conditions for the leading edge ahead of and behind the Mach cone for all test Mach numbers. A third series of eight thin flat-plate wings was tested at a Mach number of 1.92.\n\nSYMBOLS\n\n| Symbol | Definition |\n| :--- | :--- |\n| A | aspect ratio $\\left(\\frac{b^2}{S}\\right)$ |\n| $\\alpha$ | free-stream angle of attack |\n| b | wing span |\n| $\\beta = \\sqrt{M^2 - 1}$ | |\n| $c_r$ | wing root chord |\n| $\\bar{c}$ | mean aerodynamic chord, two-thirds root chord |\n| $C_L$ | lift coefficient $\\left(\\frac{\\text{Lift}}{qS}\\right)$ |\n| $C_D$ | drag coefficient $\\left(\\frac{\\text{Drag}}{qS}\\right)$ |\n| $\\Delta C_D$ | rise in drag coefficient above minimum $\\left(C_D - C_{D_{\\min}}\\right)$ |\n| $C_m$ | pitching-moment coefficient $\\left(\\frac{\\text{Moment about center of area}}{qS\\bar{c}_r}\\right)$ |\n| E | elliptic integral of second kind for $\\sqrt{1 - w^2}$ |\n| $\\epsilon$ | wing vertex half-angle |\n| L/D | ratio of lift to drag |\n| m | Mach angle $\\left(\\sin^{-1} \\frac{1}{M}\\right)$ |\n| M | Mach number |\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:22:54.935062+00:00"}
{"citation_id": "19930085485", "source_url": "https://ntrs.nasa.gov/api/citations/19930085485/downloads/19930085485.pdf", "page_number": 23, "total_pages": 26, "image_filename": "19930085485_p23.jpg", "text": "NACA RM No. L8K02\n\n$\\delta_a$\n(deg)\n$\\triangle$ 2.1\n$\\square$ 6.3\n$\\diamond$ 10.4\n\nCONFIDENTIAL\n\n$\\delta_a$\n(deg)\n$\\circ$ -2.8\n$\\triangle$ 2.3\n$\\square$ 4.6\n$\\diamond$ 10.9\n$\\nabla$ 15.3\n\nCONFIDENTIAL\n\nRolling-moment coefficient, $C_{l_a}$\n.03\n.02\n.01\n0\n-.01\n\nRolling-moment coefficient, $C_{l_a}$\n.03\n.02\n.01\n0\n-.01\n\n.5 .6 .7 .8 .9 1.0 1.1 1.2\nMach number, M\nNACA\n\n.5 .6 .7 .8 .9 1.0 1.1 1.2\nMach number, M\nNACA\n\nFigure 15.- Variation of rolling-moment coefficient\nwith Mach number for a 0.20-chord circular-arc\naileron. t = 0; $\\alpha$ = 0°; data from transonic\nbump.\nCONFIDENTIAL\n\nFigure 16.- Variation of rolling-moment coefficient\nwith Mach number for a 0.20-chord flat-sided\naileron. t = 1.00; $\\alpha$ = 0°; data from transonic\nbump.\nCONFIDENTIAL\n\n21", "timestamp": "2026-07-22T06:22:57.248977+00:00"}
{"citation_id": "19930082918", "source_url": "https://ntrs.nasa.gov/api/citations/19930082918/downloads/19930082918.pdf", "page_number": 23, "total_pages": 62, "image_filename": "19930082918_p23.jpg", "text": "22\nNACA TN 1940\n\nFurther, the maximum rupture times at each stress in figures 16 and 19 were approximately the same for aging at either $1400^\\circ$ or $1600^\\circ$ F. Examination of the other physical measurements shows that this general increase in very short-time rupture strength with increasing aging time was associated with:\n\n(1) In the case of material aged at $1400^\\circ$ F, passage through the nucleation stage prior to precipitation as revealed by the line intensity studies.\n\n(2) Removal of relatively large- or small-radius atoms through precipitate growth after nucleation as revealed by lattice-parameter measurements.\n\n(3) Progressive formation of a rather wide, continuous band of a separate grain boundary phase and with the formation of the visible precipitate particles as revealed by micrographic examination.\n\n(4) Hardening of the material through formation of strains which resulted in diffraction-line broadening in the case of aging at $1400^\\circ$ F and initial hardening and then progressive softening in the case of aging at $1600^\\circ$ F.\n\n(5) Increased ductility of the specimens. One way to consider this effect is to express the maximum true axial strain at fracture based upon the initial and final cross-sectional areas at the fracture and the assumption of constancy of volume. Figures 17 and 20 present the results of these calculations. The effect of the decrease in cross-sectional area at the fracture with aging time was to increase the true stress during the duration of the test. In general, then, the rupture tests considered were not constant-stress tests, but rather tests with true stress increasing with time along some quantitatively unknown path in a time-stress coordinate system. It is known from the time-elongation curves for the tests at 60,000 psi that reduction of area was gradual and thus the true stress increased rather gradually throughout the test, the final value of course being greater for smaller final cross-sectional area. Inspection of the geometry of ruptured specimens showed that the cross section was reduced gradually along the axis toward the fracture. From this, it can be concluded that the degree of triaxiality of the stress system at the fracture was low, as considered in the calculations of Bridgeman (see reference 11). Hence, the stress system can at least be considered uniaxial for the rupture tests covered herein. Lastly, figures 17 and 20 also show that the maximum true strain at fracture, or ductility, for any given class of specimen, decreased with increasing rupture time.\n\nAt the onset, the fact that ordinary rupture tests are not constant-stress tests makes outright analysis of the factors controlling rupture strength difficult. However, three factors stand out quite clearly:", "timestamp": "2026-07-22T06:22:58.192819+00:00"}
{"citation_id": "19930082592", "source_url": "https://ntrs.nasa.gov/api/citations/19930082592/downloads/19930082592.pdf", "page_number": 38, "total_pages": 50, "image_filename": "19930082592_p38.jpg", "text": "NACA TN 1914\n37\n\n[Figure: Micrograph showing a circular cross-section with labeled regions: Bakelite (top), Striation (middle layer), Oxide (thin layer below striation), and Unoxidized ceramal (bottom granular region). NACA logo with identifier C-22912 and date 2-7-49 is in bottom right corner of figure.]\n\nFigure 13. - Oxidation zone of 10-percent-tungsten - titanium carbide ceramal. Striations appear in oxide coatings of 5- and 10-percent-tungsten ceramals. Temperature, 1785° F; time at temperature, 30 hours; unetched; magnification, X50.", "timestamp": "2026-07-22T06:23:00.679915+00:00"}
{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 22, "total_pages": 72, "image_filename": "19930085491_p22.jpg", "text": "NACA RM No. A3J04 CONFIDENTIAL 21\n\nflow at both the leading and trailing edges. This change in flow is shown in figures 13(c) and (d) for lift coefficients of 0.21 and 0.28.³ The loss of both leading-edge suction force and pressure recovery over the rear of the wing rotates the resultant force vector rearward toward the normal to the chord and hence the experimental drag-rise variation in this range approaches that of curve (3). It should be remembered, however, in comparing the calculated and experimental results, that the theoretical values of $\\Delta C_D$ result purely from a consideration of the pressure drag; whereas the corresponding experimental values also include changes in friction drag which slightly increase the experimental values of $k_a$ and $\\Delta C_D / (\\Delta C_L)^2$.\n\nMaximum lift-drag ratio.— At a Reynolds number of 0.62 million the experimental value of maximum lift-drag ratio is 6.7 as compared to the theoretical value of 10.1. As is indicated by equation (3) and the values in table II, this difference is due to the higher experimental values of both $C_{D_{\\min}}$ and $\\Delta C_D / (\\Delta C_L)^2$. As was discussed in the preceding sections, the high experimental values of drag were due to flow separation. It appears, therefore, that any improvement in $(L/D)_{\\max}$ must come from reductions in the areas of separated flow.\n\nPitching-moment variation with lift coefficient.— The pitching-moment coefficients and the center-of-pressure positions for WF-63 are shown in figure 8(c). The center-of-pressure positions have been determined using enlarged plots of the moment data⁴ and the following equation:\n\n$$\n\\frac{X}{c} = \\frac{C_m}{C_L}\n$$\n\n(7)\n\n³Although it is not immediately apparent, a close examination of figures 13(c) and (d) reveals that at the higher lift coefficient the lengths of the attached flow areas at the leading and trailing edges are appreciably reduced.\n\n⁴The moment curves from which the center-of-pressure curves were obtained were displaced vertically by the value of the moment coefficient at zero lift which in all cases except WF-60 was small and within the limits of the experimental precision. The reason for the larger error with WF-60 was not determined. It does not, however, invalidate the variation of moment coefficient which indicates the center-of-pressure travel.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:23:08.637591+00:00"}
{"citation_id": "19930085859", "source_url": "https://ntrs.nasa.gov/api/citations/19930085859/downloads/19930085859.pdf", "page_number": 16, "total_pages": 31, "image_filename": "19930085859_p16.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:23:10.077472+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 20, "total_pages": 149, "image_filename": "19930083192_p20.jpg", "text": "16\nNACA TN 1976\n\nThe gradient distances determined from accelerometer records are a function of both the time resolution and the response characteristics of an airplane due to known gusts. Examination of gust-tunnel test results indicates that, under average conditions, the airplane response places two limits on the gust-gradient distance determined from acceleration records. The first limit, for the shorter gradient distances, is specified by the lag in lift. The gradient distances lying above the dash line in figure 4 are not generally recognizable on an accelerometer record. Thus, this limit is from 2 chords to about 4 chords for the average airplane. The second limit, for the longer gradient distances, arises from the fact that, as the airplane penetrates farther and farther into a long gradient gust, the effect of the pitching and vertical motion of the airplane, under the action of the gust, tends to counteract the contribution of lift due to each increment of gust as it is encountered.\n\nThe lateral gust-gradient distance is a more questionable quantity than the gust-gradient distance determined from accelerometer records since no single concept as to shape has been obtained. The lack of detailed data and the erratic character of the gust distributions in a spanwise direction (fig. 15) require the use of considerable judgment in arriving at any numerical values. Comparison of the actual values of the gust-gradient distance from accelerometer records and the corresponding values of the lateral distance from pressure records indicates that the data are reasonable and in agreement with the actual conditions. From the preceding remarks it is obvious that the error is significant and the value of 20 percent previously mentioned is only a crude estimate.\n\nDISCUSSION\n\nGust intensity.- Because only a small number of all the acceleration peaks in rough air can be evaluated for true gust velocities, the effective gust velocity has been used as the measure of intensity and frequency of atmospheric gusts. Although the effective gust velocity $U_e$ is a fictitious quantity, it bears a fixed relation to the gust velocity for a given airplane and given conditions of air density, gust shape, and gust size. The effective gust velocity is generally about 50 percent to 70 percent of the gust velocity.\n\nThe largest effective gust velocity recorded to date was about 55 feet per second and corresponds to a gust velocity of 100 feet per second for typical conditions. Fortunately, gust velocities this large are not frequently encountered. The distributions of effective gust velocity in figure 8 show that the limiting distributions for a wide range of airplane sizes and weather conditions are reasonably close. Reference 9 concluded, on the basis of these results, that the distribution of gust intensities was essentially independent of airplane size", "timestamp": "2026-07-22T06:23:10.483633+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 55, "total_pages": 98, "image_filename": "19930086073_p55.jpg", "text": "NACA RM A9H04\n53\n\nLift coefficient, $C_L$\nDrag coefficient, $C_D$\n\n(b) $C_L$ vs $C_D$.\n\nAngle of sideslip, $\\beta$, deg\n$\\circ$ 0.0\n$\\square$ 6.0\n$\\diamond$ 12.0\n$\\triangle$ 15.9\n\nFigure 11. - Continued.\n\n[Figure: Graph showing multiple curves of Lift coefficient vs Drag coefficient for different angles of sideslip, with data points marked by circles, squares, diamonds, and triangles. The NACA logo is visible in the bottom right corner of the graph area.]", "timestamp": "2026-07-22T06:23:11.074770+00:00"}
{"citation_id": "19930082546", "source_url": "https://ntrs.nasa.gov/api/citations/19930082546/downloads/19930082546.pdf", "page_number": 35, "total_pages": 65, "image_filename": "19930082546_p35.jpg", "text": "34\nNACA TN No. 1870\n\n[Figure: A graph plotting Blade-width ratio, b/D, and blade-thickness ratio, h/b, and Blade angle, $\\beta$, deg, against $r/R_t$. The x-axis ranges from .3 to 1.0. The left y-axis ranges from 0 to .20. The right y-axis ranges from 30 to 80. Three curves are plotted: one labeled h/b, one labeled b/D, and one labeled $\\beta$. A NACA logo is present in the bottom right corner of the plot area.]\n\n(a) Square-tip propeller.\nFigure 3.- Concluded.", "timestamp": "2026-07-22T06:23:11.864479+00:00"}
{"citation_id": "19930082646", "source_url": "https://ntrs.nasa.gov/api/citations/19930082646/downloads/19930082646.pdf", "page_number": 22, "total_pages": 37, "image_filename": "19930082646_p22.jpg", "text": "NACA TN 1980\n21\n\nTrim\nSpeed\nRise\n\nLanding trim, $\\tau_L = 5.5^\\circ$\nO Step in\n\nSpeed, mph\nRise, ft\nTrim, deg\nTime, sec\n\n(c) Extended time history for a typical landing of the modified hull.\nFigure 9.- Concluded.", "timestamp": "2026-07-22T06:23:13.136990+00:00"}
{"citation_id": "19930085862", "source_url": "https://ntrs.nasa.gov/api/citations/19930085862/downloads/19930085862.pdf", "page_number": 11, "total_pages": 60, "image_filename": "19930085862_p11.jpg", "text": "NACA RM No. L9A07\n\nbasic data of figures 5 to 10 and are presented in figure 15. It can be seen that considerable variation in the values of these parameters occurred. In order to show the effects of the variation in $C_{h_\\alpha}$, a rolling condition must be considered, and if the aileron balance is to be of the conventional, sealed, internally balanced type, the parameters $P_{R_\\delta}$ and $P_{R_\\alpha}$ also must be considered. The combined effect of these parameters for an aileron having various amounts of internal balance is shown in figure 16. The hinge-moment parameters of the aileron with varying amounts of balance were calculated by means of the following equations:\n\n$$\nC_{h_\\delta \\text{ with balance}} = C_{h_\\delta \\text{ without balance}} + \\frac{1}{2} P_{R_\\delta} \\left( \\frac{\\bar{c}_b}{\\bar{c}_s} \\right)^2 \\tag{2}\n$$\n\n$$\nC_{h_\\alpha \\text{ with balance}} = C_{h_\\alpha \\text{ without balance}} + \\frac{1}{2} P_{R_\\alpha} \\left( \\frac{\\bar{c}_b}{\\bar{c}_s} \\right)^2 \\tag{3}\n$$\n\nwhere the span of the balance was assumed equal to the span of the aileron and where the balance chord was assumed to include one-half of the gap covered by the seal.\n\nThe parameter $C'_{h_\\delta}$ is defined as the rate of change of hinge-moment coefficient in a steady roll with aileron deflection and was calculated by means of the following equation:\n\n$$\nC'_{h_\\delta} = C_{h_\\delta} + \\frac{2(\\Delta\\alpha)_p}{\\Delta\\delta_a} C_{h_\\alpha}\n$$\n\nin which the values of the parameters $C_{h_\\delta}$ and $C_{h_\\alpha}$ were computed from equations (2) and (3) for various amounts of balance and where $\\frac{2(\\Delta\\alpha)_p}{\\Delta\\delta_a} = K \\frac{C_{l_\\delta}}{C_{l_p}} = -210 C_{l_\\delta}$ and is the ratio of the effective change in angle of attack in a steady roll to the change in aileron deflection. The constant $K$ was determined by means of the charts of reference 4. The damping-in-roll coefficient $C_{l_p}$ was determined from reference 5 and had a value of 0.266.", "timestamp": "2026-07-22T06:23:13.253947+00:00"}
{"citation_id": "19930085847", "source_url": "https://ntrs.nasa.gov/api/citations/19930085847/downloads/19930085847.pdf", "page_number": 17, "total_pages": 32, "image_filename": "19930085847_p17.jpg", "text": "```markdown\nNACA RM A57D04\nCONFIDENTIAL\n\n24 x 10^6\n\nReynolds number\n\n20\n16\n12\n8\n4\n0\n.68 .70 .72 .74 .76 .78 .80 .82 .84\nMach number\n\n15,000 ft\n30,000 ft\n\no Tests with no suction\n$\\Delta$ Tests with suction\n\nNACA\n\nFigure 5.- Relation between Reynolds number and Mach number for\nthe two test altitudes.\n\nCONFIDENTIAL\n15\n```", "timestamp": "2026-07-22T06:23:18.511640+00:00"}
{"citation_id": "19930085485", "source_url": "https://ntrs.nasa.gov/api/citations/19930085485/downloads/19930085485.pdf", "page_number": 24, "total_pages": 26, "image_filename": "19930085485_p24.jpg", "text": "$\\delta_a$  \n(deg)  \n$\\triangle$ 2.0  \n$\\square$ 4.8  \n$\\diamond$ 9.8  \n$\\nabla$ 15.0  \n\nCONFIDENTIAL  \n\n$\\delta_a$  \n(deg)  \n$\\nabla$ -5.0  \n$\\circ$ -2.0  \n$\\triangle$ 2.0  \n$\\square$ 4.8  \n$\\diamond$ 9.8  \n$\\nabla$ 15.0  \n\nRolling moment coefficient, $C_{l_a}$  \nMach number, $M$  \n(a) $\\alpha = 0^\\circ$.  \n\nRolling moment coefficient, $C_{l_a}$  \nMach number, $M$  \n(b) $\\alpha = 3.4^\\circ$.  \n\nFigure 17.- Variation of rolling-moment coefficient with Mach number  \nfor a 0.20-chord flat-sided aileron. $t = 0.50$; data from  \ntransonic bump.  \n\nCONFIDENTIAL  \n\nNACA EW No. 16802", "timestamp": "2026-07-22T06:23:25.481727+00:00"}
{"citation_id": "19930082592", "source_url": "https://ntrs.nasa.gov/api/citations/19930082592/downloads/19930082592.pdf", "page_number": 39, "total_pages": 50, "image_filename": "19930082592_p39.jpg", "text": "Page intentionally left blank\n\nPage intentionally left blank", "timestamp": "2026-07-22T06:23:25.811794+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 95, "total_pages": 96, "image_filename": "19930085880_p95.jpg", "text": "NACA RM No. L9C03\n93\n\n| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | 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| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | |", "timestamp": "2026-07-22T06:23:26.173545+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 39, "total_pages": 78, "image_filename": "19930082483_p39.jpg", "text": "NACA TN No. 1807\n37\n\nare maintained at the design point and there would be little requirement for operation in the more unfavorable regions.\n\nComparison of Partial Admission with Other Power-Control Methods\n\nTurbine-inlet pressure and total-pressure ratio are evaluated as power-control parameters in figure 12. For a given admission, there appear to be about equal ranges of power control using either inlet pressure or pressure ratio as the controlling parameter. However, with respect to over-all efficiency, the curves indicate that no advantage is gained by inlet-pressure control either exclusively or in combination with pressure-ratio control. The same power reduction can be accomplished through use of pressure-ratio control independently of inlet-pressure control and at considerably higher efficiencies. Comparison of the methods shows that partial admission offers a more flexible system while maintaining a corresponding trend with regard to efficiency.\n\nTurbine-rotor speed appears as a power-control parameter in figure 13. As is evidenced in the figure, power control by means of pressure ratio offers a more flexible system than that utilizing rotor speed. Some benefit may be realized, however, particularly at lower admissions, if the rotor speed is manipulated over a range sufficient to take advantage of the gains in efficiency available at the lower pressure ratios. That the power-control curves terminate with a total-pressure ratio of 1.5 should not be taken to mean that further power reduction is not obtainable by even lower pressure ratios. In these studies, the turbine was not operated below this value and no data are available. Judging from the trends in evidence at the lowest pressure ratio shown, further power reduction by diminishing pressure ratio must be accompanied by increasingly severe speed regulation if the optimum turbine efficiencies are to obtain. Reductions in both speed and pressure ratio are attended by increased turbine-discharge temperatures, which impose, as a further limitation to this system of power control, the highest turbine-discharge temperature that may be sustained for any length of operation. The partial-admission control curve, which is superimposed on figure 13, again serves to orient the operational areas at various degrees of active nozzle arc.\n\nBecause all data were obtained in this investigation at a constant inlet temperature, little can be said concerning the effects of variable inlet temperature as a control parameter. A brief analysis would show that such a system is very restricted as regards flexibility of power output, but the small reductions permitted (on the order of 10 to 20 percent of full power) are obtained with a negligible reduction in the efficiency.", "timestamp": "2026-07-22T06:23:26.853355+00:00"}
{"citation_id": "19930083221", "source_url": "https://ntrs.nasa.gov/api/citations/19930083221/downloads/19930083221.pdf", "page_number": 23, "total_pages": 47, "image_filename": "19930083221_p23.jpg", "text": "NACA TN No. 1324\n21\n\nIn linearized theory the boundary-value problems of wing theory are concerned with two separate properties of the wing: the thickness effects and the effects produced by the twist, camber, and angle of attack. The first is called the nonlifting case and the second is the lifting case. Solutions of equation (7) for $M_0 \\ge 1$ are given in reference (13) as follows:\n\nIn the nonlifting case\n\n$$\n\\varphi(x,y,z) = - \\frac{1}{2\\pi} \\int_{\\tau} \\int \\frac{\\Delta w_0(x_1,y_1)dx_1dy_1}{\\sqrt{(x-x_1)^2 - \\beta^2(y-y_1)^2 - \\beta^2 z^2}} \\quad (29)\n$$\n\nwhere $\\beta = \\sqrt{M_0^2 - 1}$ and $\\Delta w_0 = 2w_0$ where $w_0$ is the vertical perturbation velocity on the wing and therefore related directly to the slope of the wing surface relative to the x axis. The integration region $\\tau$ is the area on the wing within the Mach forecone from the point x, y, z.\n\nIn the lifting case\n\n$$\n\\varphi(x,y,z) = \\frac{1}{2\\pi} \\sqrt{\\int_{\\tau} \\int \\frac{\\beta^2 z \\Delta \\varphi_0(x_1,y_1)dx_1dy_1}{[(x-x_1)^2 - \\beta^2(y-y_1)^2 - \\beta^2 z^2]^{3/2}}} \\quad (30)\n$$\n\nwhere $\\Delta \\varphi_0$ is the jump in the value of the velocity potential in the plane of the wing. The sign $\\sqrt{}$ denotes \"finite part\" of the integral and introduces special integration techniques. (See reference 13.)\n\nEquation (29) expresses the velocity potential for the symmetrical wing in terms of an integral involving supersonic source distributions while equation (30) employs doublet distributions. In the two cases the distributions are determined from the geometry and the load distribution over the wing, respectively.\n\nSource and doublet distribution effectiveness at infinity.— It is well known that the lift, drag, and pitching moment of a given wing may be calculated either from direct integration of the local pressures on the wing or by means of momentum considerations where the induced velocities of the wing are determined at an infinite distance and the desired forces are related to an integration over a control surface", "timestamp": "2026-07-22T06:23:32.704621+00:00"}
{"citation_id": "19930085870", "source_url": "https://ntrs.nasa.gov/api/citations/19930085870/downloads/19930085870.pdf", "page_number": 5, "total_pages": 92, "image_filename": "19930085870_p5.jpg", "text": "```markdown\n4\nCONFIDENTIAL\nNACA RM No. L9D07\n\nq dynamic pressure ($\\frac{1}{2}\\rho V^2$)\n$\\rho$ stream density\nR Reynolds number based on $\\bar{c}$\nS wing area\nt maximum wing thickness\nV free-stream velocity\nw = $\\frac{\\tan \\epsilon}{\\tan m}$\n\nAPPARATUS AND TESTS\n\nWind tunnel and model support.- The Langley 9-inch supersonic tunnel is a closed-return, direct-drive type in which the pressure and humidity of the enclosed air may be controlled. Throughout the tests the quantity of water vapor in the tunnel air was kept at sufficiently low values to insure negligible effects of condensation in the supersonic nozzle. The test Mach number is varied by means of interchangeable nozzle blocks forming test sections approximately 9 inches square. A schlieren optical system provides qualitative visual-flow observations. Eleven fine-mesh, turbulence-damping screens are installed in the settling chamber ahead of the nozzles.\n\nAs shown in figure 1 the models were mounted from the rear on very slender, tapered sting supports that passed through the sting windshield with small clearance and were attached to the scales by insertion in the model sting support. It should be noted that the forward edges of the sting windshield lay behind the sting shoulders, thus tending to avoid any impact pressures. The scales are self-balancing beam scales and measure three components, in a horizontal plane, of the total forces on the model and support system.\n\nDescription of models.- The geometric characteristics of the model wings are given in figures 2 and 3 and in table 1. Photographs of the elliptical- and wedge-leading-edge wings are shown in figure 4. These wings were constructed of highly polished, hard steel and with elliptical leading edges. The wedge-leading-edge wings were obtained from the elliptical-leading-edge wings by grinding to a wedge the region in front of the line of maximum thickness. This grinding caused no appreciable change in thickness ratio, location of maximum thickness, or vertex angle. Mirrors approximately 1/16 inch square were flush mounted in the stings just ahead of the shoulder as a part of the optical angle-of-attack system. (See fig. 1.)\n\nCONFIDENTIAL\n```", "timestamp": "2026-07-22T06:23:32.942453+00:00"}
{"citation_id": "19930085859", "source_url": "https://ntrs.nasa.gov/api/citations/19930085859/downloads/19930085859.pdf", "page_number": 17, "total_pages": 31, "image_filename": "19930085859_p17.jpg", "text": "```markdown\nNACA RM No. L9B25\n\nM=0.77\nM=0.91\n\nM=1.00\nM=1.17\n\nNominal boundary-layer thickness\n\nFigure 5.- Typical Mach number contours over transonic bump in region of model location.\n\nNACA\n15\n```", "timestamp": "2026-07-22T06:23:36.706214+00:00"}
{"citation_id": "19930085548", "source_url": "https://ntrs.nasa.gov/api/citations/19930085548/downloads/19930085548.pdf", "page_number": 19, "total_pages": 46, "image_filename": "19930085548_p19.jpg", "text": "18\nNACA RM No. E8L30\n\nt time for flow process, sec\n$V_D$ piston displacement, cu in.\n$\\Delta\\theta$ total valve- or port-opening period, deg\nN engine speed, rpm\nc velocity of sound (1100 ft/sec at inlet conditions, 2500 ft/sec at exhaust conditions, values arbitrarily chosen)\nM hypothetical average Mach number for flow through valve or port\nS average piston speed, ft/min\n$A_p$ area of piston, sq in.\n\nThe data necessary for determining the values of M for the ported and for the conventional aircraft-engine cylinder and also comparative values of M for each are as follows:\n\n| Concept | Ported cylinder | | Aircraft-engine cylinder | |\n| :--- | :---: | :---: | :---: | :---: |\n| | Inlet | Exhaust | Inlet | Exhaust |\n| S, (ft/min) | 1350 | 1350 | 1350 | 1350 |\n| $A_p$, (sq in.) | 8.3 | 8.3 | 23.7 | 23.7 |\n| c, (ft/sec) | 1120 | 2500 | 1120 | 2500 |\n| $A_{max}$(total), (sq in.) | 3.03 | 2.62 | 5.16 | 4.15 |\n| $C_a$ | 0.238 | 0.372 | 0.398 | 0.373 |\n| $\\Delta\\theta$, deg | 122 | 142 | 290 | 282 |\n| M | 0.344 | 0.0985 | 0.144 | 0.088 |\n| M (2400 rpm$^a$) | 0.624 | 0.179 | 0.263 | 0.160 |\n\n$^a$Rated speed for aircraft-engine cylinder.\n\nREFERENCES\n\n1. Teuschek, Max J., and Biermann, Arnold E.: An Analysis of a Piston-Type Gas-Generator Engine. NACA RM No. E7I10, 1948.\n2. Bogowski, A. R., and Bouchard, C. L.: Scavenging a Piston-Ported Two-Stroke Cylinder. NACA TN No. 674, 1938.\n3. Gerrish, Harold C., Meem, J. Lawrence, Jr., Scadron, Marvin D., and Colmar, Anthony: The NACA Mixture Analyzer and Its Application to Mixture-Distribution Measurement in Flight. NACA TN No. 1238, 1947.", "timestamp": "2026-07-22T06:23:36.879541+00:00"}
{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 23, "total_pages": 72, "image_filename": "19930085491_p23.jpg", "text": "22 CONFIDENTIAL NACA RM No. A8J04\n\nPositive values of X indicate center-of-pressure positions ahead of the transverse axis through the centroid of wing area which occurs at the 50-percent station of the mean aerodynamic chord.\n\nThe center-of-pressure travel associated with the variation of moment coefficient can be explained in terms of the changes in boundary-layer flow with lift coefficient previously discussed. Increasing the lift coefficient from $C_L = 0$ to $C_L = 0.09$ resulted in an increase in the area of separated flow on the inboard top surfaces and a decrease in area of separated flow on the outboard bottom surface. The loss of lift on the top surface occurs not far from the centroid of area, while the increase in lift on the bottom surface occurs considerably behind the centroid of area. The combined effect is to move the resultant center of pressure rearward from its location at zero lift.\n\nAbove $C_L = 0.09$ the flow on the bottom surface is entirely attached, but on the upper surface the line of laminar separation has moved close to the leading edge over most of its length. The corresponding reduction in the negative pressure peak near the leading edge has a tendency to move the center of pressure farther aft. However, as the lift coefficient increases, the separated area on outboard sections becomes progressively larger and since this loss of lift occurs behind the centroid of area, it has the effect of moving the center-of-pressure forward. These opposing actions limit the maximum rearward position of the center of pressure to approximately 8 percent of the mean aerodynamic chord behind the centroid of area at a lift coefficient of 0.22. Above this lift coefficient the effect of the inboard progression of separation predominates and the center of pressure moves forward.\n\nEffect of Reynolds Number on Longitudinal Characteristics\n\nBecause of the relatively small scale of the test model the effects of Reynolds number are important in an estimation of the characteristics of a full-scale configuration. Since similar Reynolds number effects may be expected with all configurations tested, the following discussion is primarily concerned with the changes observed with WF-63, the configuration which is a part of the general investigation at both subsonic and supersonic speeds.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:23:37.061431+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 56, "total_pages": 98, "image_filename": "19930086073_p56.jpg", "text": "```markdown\n54\n\nLift coefficient, $C_L$\n1.4\n1.2\n1.0\n.8\n.6\n.4\n.2\n0\n\nPitching-moment coefficient, $C_m$\n0 0 0 0 -.04 -.08 -.12 -.16 -.20 -.24 -.28 -.32\n\n[Figure: Graph plotting Lift coefficient vs Pitching-moment coefficient with four data series]\n\n$\\bigcirc$ $\\square$ $\\diamond$ $\\triangle$\n0.0 6.0 12.0 15.9\nAngle of sideslip, $\\beta$, deg\n\n(c) $C_L$ vs $C_m$.\n\nFigure 11.— Continued.\n\nNACA RM A9H04\n```", "timestamp": "2026-07-22T06:23:37.262329+00:00"}
{"citation_id": "19930082646", "source_url": "https://ntrs.nasa.gov/api/citations/19930082646/downloads/19930082646.pdf", "page_number": 23, "total_pages": 37, "image_filename": "19930082646_p23.jpg", "text": "22\nNACA TN 1980\n\nTrim at greatest cycle, deg\nTrim at contact, deg\nO Maximum trim\n$\\Delta$ Minimum trim\n\nWarped forebody and extended afterbody ———\nBasic forebody and basic afterbody - - - - -\n\nRise at greatest cycle, ft\nTrim at contact, deg\nO Maximum rise\n$\\Delta$ Minimum rise\nNACA\n\nFigure 10.- Variation in maximum and minimum trim and rise with trim at contact, for landings in smooth water.", "timestamp": "2026-07-22T06:23:39.440725+00:00"}
{"citation_id": "19930082546", "source_url": "https://ntrs.nasa.gov/api/citations/19930082546/downloads/19930082546.pdf", "page_number": 36, "total_pages": 65, "image_filename": "19930082546_p36.jpg", "text": "NACA TN No. 1870\n35\n\n[Figure: Test setup for free-space pressure measurements. A propeller mounted on a vertical pole is shown on the left, and a multi-element antenna array on a tripod is shown on the right. In the background are buildings and a large cylindrical tank. A label on the tripod reads \"NACA L-56021\".]\n\nFigure 4.- Test setup for free-space pressure measurements.", "timestamp": "2026-07-22T06:23:39.550981+00:00"}
{"citation_id": "19930082918", "source_url": "https://ntrs.nasa.gov/api/citations/19930082918/downloads/19930082918.pdf", "page_number": 24, "total_pages": 62, "image_filename": "19930082918_p24.jpg", "text": "NACA TN 1940\n\nFirst, figures 18 and 21 show that the relative initial weakness of the unaged materials was associated with weak grain boundary areas and consequent intergranular crack formation, especially on grain boundaries normal to the stress axis. Further inspection (see figs. 18 and 21) shows that this tendency for intergranular cracking was progressively removed with either increased aging time at $1400^\\circ$ or $1600^\\circ$ F prior to testing, or with decreased stress on the unaged material and thus increased time at the test temperature of $1200^\\circ$ F. This apparently was due to the formation of the grain boundary phase either prior to or during testing (see figs. 6 to 9). Thus unaged material became stronger, relatively speaking, with increased rupture time and the aged material also became relatively stronger with increased prior aging time – more slowly, however, with prior aging time at $1400^\\circ$ F than for $1600^\\circ$ F aging since the boundary phase was formed more slowly at $1400^\\circ$ F. It is then quite interesting to conclude that the matrix- and boundary-phase binding was stronger than at least certain oriented matrix-matrix bindings.\n\nSecond, inspection of figures 16 and 19 shows that, once a definite almost continuous boundary phase was formed, the rupture strengths were roughly independent of aging temperature or aging time. This indicated that the initiation and propagation of the predominantly transgranular cracks were independent of the differences in structure arising from differences in aging at $1400^\\circ$ or $1600^\\circ$ F. Aside from minor maximums or minimums, which are probably just inside or outside the experimental errors involved in determining the rupture times, the differences in lattice depletion of the large-radius atoms, and the differences in magnitude and time of occurrence of maximum hardening in specimens aged at different temperatures, had little effect. This conclusion is based, however, in part upon the fact that the degree of elongation and cross-sectional area reduction, once the grain boundary phase was formed, was approximately the same for samples having the same rupture time as the result of aging at $1400^\\circ$ or $1600^\\circ$ F. Thus the degree of deformation was also constant and did not affect to a first approximation the fracturing characteristics of the two classes of aged materials differently. It is well known, however, that, within limits, deformation in itself generally raises the resistance to rupture (see reference 12). Since this strain strengthening occurs in connection with increase in the true stress, evaluation of either of these two opposite effects is difficult along with evaluation of such things as the effect of progressive lattice depletion in general on resistance to crack propagation. Thus no further conclusions in regard to possible masked general structure factors arising with long-time aging at $1400^\\circ$ or $1600^\\circ$ F are drawn at this time.\n\nThird, the plastic strain, before rupture failure occurred, increased markedly with long aging times at either $1400^\\circ$ or $1600^\\circ$ F. This of course is a direct result of the fact that the creep resistance decreased markedly with long aging times at either aging temperature", "timestamp": "2026-07-22T06:23:44.163721+00:00"}
{"citation_id": "19930085890", "source_url": "https://ntrs.nasa.gov/api/citations/19930085890/downloads/19930085890.pdf", "page_number": 3, "total_pages": 26, "image_filename": "19930085890_p3.jpg", "text": "2\nNACA RM No. E9C11\n\nand experimental investigations are being conducted with diborane\nbecause it was the first boron hydride available. Theoretical\ncalculations of diborane with several oxidants and experimental\nresults from the operation of a 100-pound-thrust rocket engine with\nliquid diborane and hydrogen peroxide are presented in references 1\nand 2, respectively. An investigation was conducted at the NACA\nLewis laboratory from March to May 1948 to determine the performance\nof a 100-pound-thrust rocket engine using liquid diborane and liquid\noxygen. The diborane was obtained through the cooperation of the\nNavy Bureau of Aeronautics. Specific impulse was measured for a\nrange of mixture ratios with a combustion-chamber pressure of\napproximately 300 pounds per square inch absolute. Additional data\nwere obtained on the effect of reducing the ratio of combustion-\nchamber volume to exhaust-nozzle-throat area (hereinafter desig-\nnated characteristic length L*) by a factor of 2 and on the\neffect of changes in propellant injection.\n\nIn order to obtain additional information concerning the\nstability of diborane, a brief series of experiments was made to\ndetermine the sensitivity of diborane to temperature and shock.\nThese supplementary data are presented in the appendix.\n\nAPPARATUS\n\nA diagrammatic sketch of the apparatus is shown in figure 1.\nHelium pressure, controlled by two-stage regulation, was used to\nforce the propellants into the combustion chamber. The oxygen\ntank was made of stainless steel and had a vacuum jacket; the\ndiborane tank was made of a molybdenum steel. Each tank was\nmounted on a counterbalanced weighing beam for flow measurement.\nThe lines from the diborane tank and the propellant valves were\narranged for precooling with liquid nitrogen and dry ice, respec-\ntively. The rocket engine was mounted on a pivoted thrust stand\nto measure thrust and at a downward angle of 30° to prevent the\naccumulation of propellants in the combustion chamber at the\nstart of a run. A photograph of the general setup is shown in\nfigure 2.\n\nEngine Assemblies\n\nCross-sectional views of the engine assemblies are shown in\nfigure 3. The engine assemblies consisted of interchangeable\ninjection plates, combustion chambers, and exhaust nozzles.\n\nThe injection plates were made of stainless steel and used\nremovable injectors inserted from the inside. The injectors", "timestamp": "2026-07-22T06:23:49.520567+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 21, "total_pages": 149, "image_filename": "19930083192_p21.jpg", "text": "NACA TN 1976\n\nand source of turbulence. Figure 9, in contrast, shows that the XC-35 data for different altitude brackets result in distributions with slopes that vary and, when checked against figure 8, differ from those curves. This discrepancy indicates that the distributions of effective gust velocity may not be independent of the source of turbulence. Independency is a reasonable assumption at this time for general use.\n\nFigures 9 and 10 show that, for altitudes above 9000 feet, the maximum gust intensity in rough air associated with convective clouds is essentially a constant. Below 9000 feet, the XC-35 data indicate a decrease in intensity, due in part, according to reference 10, to the fact that the records were taken in clear air and clouds. Consideration of figure 10 indicates that little difference exists in the expected maximum gust velocities for all altitude ranges.\n\nGust-gradient distance.- Figure 11 and similar plots for other airplanes show that gust-gradient-distance data are of a random character. Comparisons of such plots indicate that airplane size might be a significant parameter, as in the case of boats where waves of small wave length that are of no concern to a large boat cause the small boat considerable difficulty. In a similar manner, gusts that cause a rough ride on a small airplane, such as the Aeronca C-2, would be expected to be of such small size as to have little effect on a large airplane, such as the XB-15.\n\nComparison of figures 12(a), 12(b), and 12(c) indicates that the gradient-distance data show much less scatter on the basis of the mean geometric chord than on the basis of span or feet. The scatter of points when the gust-gradient distance is defined in chords is about $1\\frac{1}{2}$ to 1 for a given value of U, while for the gust-gradient distance plotted in terms of airplane span, the scatter is greater. The data of figure 12(c) show that when the gust-gradient distance is expressed in feet, the average values of the distance vary from 40 feet to 200 feet for a given gust intensity. Inasmuch as the agreement for the different sets of data is best when the gradient distance is expressed in chords, the gust-gradient distance appears to be roughly independent of the airplane when expressed in wing mean geometric chords.\n\nThe adequacy of the average value of the gust-gradient distance for the representation of gradient-distance data deserves some consideration. The relation between the average value and the most probable value shown in figure 13 indicates that the probable gradient distance is about 1 chord to 2 chords less than the average gradient distance. Since the offset appears to be constant, the use of the average gust-gradient distance, which is generally easier to determine, is believed to be satisfactory for analysis.", "timestamp": "2026-07-22T06:23:49.698303+00:00"}
{"citation_id": "19930085485", "source_url": "https://ntrs.nasa.gov/api/citations/19930085485/downloads/19930085485.pdf", "page_number": 25, "total_pages": 26, "image_filename": "19930085485_p25.jpg", "text": "$\\delta_a$\n(deg)\n$\\Delta$ 2.0\n$\\square$ 5.0\n$\\diamond$ 10.0\n$\\nabla$ 15.0\n\nCONFIDENTIAL\n\n$\\delta_a$\n(deg)\n$\\nabla$ -5.0\n$\\circ$ -2.0\n$\\Delta$ 2.0\n$\\square$ 5.0\n$\\diamond$ 10.0\n$\\nabla$ 15.0\n\nNACA RM No. L58K02\n\nRolling-moment coefficient, $C_{l_a}$\n.03\n.02\n.01\n0\n-.01\n.5 .6 .7 .8 .9 1.0 1.1 1.2\nMach number, M\nNACA\n(a) $\\alpha = 0^\\circ$.\n\nRolling-moment coefficient, $C_{l_a}$\n.03\n.02\n.01\n0\n-.01\n.5 .6 .7 .8 .9 1.0 1.1 1.\nMach number, M\nNACA\n(b) $\\alpha = 3.5^\\circ$.\n\nFigure 18.- Variation of rolling-moment coefficient with Mach number\nfor a 0.20-chord flat-sided aileron. $t = 0.37$; data from\ntransonic bump.\nCONFIDENTIAL\n\n23", "timestamp": "2026-07-22T06:23:50.463224+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 96, "total_pages": 96, "image_filename": "19930085880_p96.jpg", "text": "94\nNACA RM No. L9C03\n\n.64\n.56\n.48\n.40\n.32\n.24\n.16\n.08\n0\n\nSpeed\n(fps)\nO 10\n□ 15\n◇ 20\n△ 25\n▽ 30\n\nDraft, ft\nWetted area, sq ft\n0 .05 .10 .15 .20 .25 .30 .35\n\n(e) $\\tau = 20^\\circ$.\nFigure 25.- Concluded.\n\nNACA\nNACA - Langley Field, Va.", "timestamp": "2026-07-22T06:23:50.643384+00:00"}
{"citation_id": "19930082566", "source_url": "https://ntrs.nasa.gov/api/citations/19930082566/downloads/19930082566.pdf", "page_number": 34, "total_pages": 44, "image_filename": "19930082566_p34.jpg", "text": "```markdown\n32\n\nMaximum principal stress, $\\sigma_2'$, psi\n\nN, cycles\n\n(b) For stress ratio $R = \\sigma_2' / \\sigma_1' = 2$.\n\nFigure 12.- Continued.\n\nNACA TN NO. 1889\n\n[Figure: A graph plotting Maximum principal stress against N, cycles. The y-axis ranges from 0 to 50 x 10^3 psi. The x-axis is logarithmic, ranging from 4 x 10^4 to 10^7 cycles. The graph contains a curve and several data points. The NACA logo is visible in the bottom right corner of the plot area.]\n```", "timestamp": "2026-07-22T06:23:53.305562+00:00"}
{"citation_id": "19930085859", "source_url": "https://ntrs.nasa.gov/api/citations/19930085859/downloads/19930085859.pdf", "page_number": 18, "total_pages": 31, "image_filename": "19930085859_p18.jpg", "text": "16\n\nReynolds number, R\n\n1.0 x 10⁶\n\n.8\n\n.6\n\n.4\n\n.6 .7 .8 .9 1.0 1.1 1.2\n\nMach number, M\n\nMean\n\nNACA\n\nFigure 6.— Variation of test Reynolds number with Mach number for a model with 35° sweptback wing, aspect ratio 4, taper ratio 0.6, and NACA 65A006 airfoil.\n\nNACA RM No. L9B25", "timestamp": "2026-07-22T06:23:53.791218+00:00"}
{"citation_id": "19930085862", "source_url": "https://ntrs.nasa.gov/api/citations/19930085862/downloads/19930085862.pdf", "page_number": 12, "total_pages": 60, "image_filename": "19930085862_p12.jpg", "text": "10\nNACA RM No. L9A07\n\nConsidering first the aileron without any internal balance, figure 16 shows that on the plain wing the aileron was more underbalanced at high angles of attack than it was at low angles of attack. The addition of the split flaps resulted in an opposite effect; the aileron was more balanced at high than at low angles of attack. The further addition of the leading-edge flaps and the stall-control fences tended to offset the effect of the split flaps and resulted in a reduction in the variation of $C'_{h\\delta}$ through the angle-of-attack range. The values of $C'_{h\\delta}$ at high angles of attack and for the flapped configurations are not necessarily correct since $C_{l_p}$, which was determined at $\\alpha = 0^\\circ$ for the plain wing and assumed constant in the determination of $C'_{h\\delta}$, probably varies with angle of attack and flap configuration. The trends, however, are considered to be indicative of the effects of the high-lift and stall-control devices.\n\nThe data presented in figure 16 indicate that on the plain wing at zero angle of attack this aileron equipped with a conventional, sealed internal balance would require a balance chord of about 30 percent of the aileron chord for $C'_{h\\delta} = 0$. As the angle of attack is increased, more balance chord is required until at about the angle of attack for $C_{L_{max}}$ a balance chord of approximately 55 percent would be required. For the split-flap configuration the amount of balance chord required for $C'_{h\\delta} = 0$ was 45 percent at $\\alpha = 0^\\circ$, increased to more than 55 percent at moderate angles of attack, and then decreased to about 45 percent at high angles of attack. For the configurations involving the leading-edge devices and the fences, the amount of balance chord required for $C'_{h\\delta} = 0$ was between 45 and 50 percent at low angles of attack and increased about 5 percent at the angles of attack corresponding to $0.85C_{L_{max}}$.\n\nIf, therefore, the aileron on the plain wing was closely balanced for the high-speed condition, it would be underbalanced at the low-speed, flaps-deflected condition although the small dynamic pressures at the low speeds would tend to prevent the occurrence of excessive control forces.\n\nThe foregoing comparison of the aileron effectiveness for the various flap configurations has been made by using slopes determined at zero aileron deflection. Since the data presented in figure 12 indicate", "timestamp": "2026-07-22T06:23:55.315873+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 40, "total_pages": 78, "image_filename": "19930082483_p40.jpg", "text": "38\nNACA TN No. 1807\n\nDISCUSSION\n\nBecause reductions in efficiency with partial-admission operation may be ascribed in large part to the induced losses, it is worthwhile examining these losses briefly with the idea of devising means to reduce or eliminate them. As has been mentioned previously, there are two principal types of loss that occur with partial admission: pumping losses and driving-fluid losses. In reference 4 (pp. 199-200), it is shown that pumping losses may be effectively reduced by closely shrouding the inactive rotor blading. Such shrouding would also serve to reduce that part of the driving-fluid loss caused by induced gas diffusion at the nozzle discharge as follows: (1) It would decrease the pressure gradient between the active and inactive flow regions of the blading, and (2) it would eliminate the flow of the diffused gases through the inactive rotor blading and the subsequent power loss. The method proposed in reference 4 (p. 221) is aimed at reducing the scavenge and eddy losses by more careful control of the flow to the rotor at the start and end of the active cycle.\n\nThese and other refinements, although not employed in this investigation, are worth consideration in the design of a partial-admission power plant.\n\nA turbine operating with partial admission is subject to induced rotor-blade vibrations in addition to those encountered with full-admission operation. During normal turbine operation, when the period of the intermittent force applications from the individual nozzles is the same as a natural vibration period of the blades, secondary resonance (reference 9) occurs. This secondary resonance is characteristic of turbines both at full admission and during the active cycle at partial admissions. Harmful stresses resulting from secondary resonance may be averted by proportioning the blades to secure a suitable mass-stiffness ratio and provide sufficient internal damping to maintain the vibration amplitudes at a safe low level.\n\nPrimary resonance, as defined in reference 9, occurs at partial admission when the forced vibration set up by the application and release of the driving-fluid force at the beginning and the end of cut-off (fig. 5) has the same period as a natural vibration frequency. One or more operating conditions will always exist during partial-admission operation at which the timing of the loading and the unloading of the blades in the manner described will be such as to cause primary resonance. Precautions must therefore be taken in the design of the rotor blading so that, under conditions of primary resonance, the maximum blade stresses induced may be withstood safely.", "timestamp": "2026-07-22T06:24:02.251366+00:00"}
{"citation_id": "19930082646", "source_url": "https://ntrs.nasa.gov/api/citations/19930082646/downloads/19930082646.pdf", "page_number": 24, "total_pages": 37, "image_filename": "19930082646_p24.jpg", "text": "```markdown\nNACA TN 1980\n23\n\nWarped forebody and extended afterbody\nBasic forebody and basic afterbody\n\nSpray on flaps\nLight\nHeavy\nFlaps clear\nClear\n\nSpray in propellers\nLight\nHeavy\nPropellers clear\nClear\n\nGross load, lb\n110 x 10³\n100\n90\n80\n70\n60\n50\n\nSpeed, mph\n0\n10\n20\n30\n40\n50\n\nFigure 11.- Variation of range of speed for spray in propellers and on flaps with gross load.\n\n[Figure: Two graphs showing spray conditions (Light, Heavy, Clear) for propellers and flaps against gross load and speed. The left graph is for \"Spray in propellers\" and the right graph is for \"Spray on flaps\". Both graphs have \"Gross load, lb\" on the y-axis (from 50 to 110 x 10³) and \"Speed, mph\" on the x-axis (from 0 to 50). The NACA logo is present in the bottom right corner of the right graph.]\n```", "timestamp": "2026-07-22T06:24:04.887627+00:00"}