PDF/ocr_dataset_1015.jsonl

{"citation_id": "19930082476", "source_url": "https://ntrs.nasa.gov/api/citations/19930082476/downloads/19930082476.pdf", "page_number": 12, "total_pages": 41, "image_filename": "19930082476_p12.jpg", "text": "10\nNACA TN No. 1801\n\nSpinproofing\n\nThe data presented in the charts indicate that at the lower of the two wing loadings tested (approx. 10 lb/sq ft) limiting the up elevator deflection to 13° (assumed to be the minimum up elevator deflection required to land the airplane satisfactorily), limiting the up aileron movement to about 5°, and limiting the outboard rudder (left rudder in a right spin) so that it can not be set with the spin would prevent the attainment of spinning equilibrium. In order to maintain satisfactory rolling characteristics in normal flight by utilizing only a 5° maximum up aileron deflection, it will be necessary to have a reverse differential aileron movement (that is, greater down aileron than up aileron deflection). Computations made by the methods outlined in reference 6 show that if the ailerons are sealed a down aileron deflection of 16° and an up aileron deflection of 5° will give a maximum value of $\\frac{pb}{2V}$ (helix angle generated by the wing tip in a roll) equivalent to 0.07, the minimum permissible value specified in reference 6. The adverse yawing moments contributed by the ailerons utilizing a 5° up and 16° down deflection were computed by methods given in references 7 and 8. Model force-test data were available for computing the yawing moments contributed by the rudder for small rudder deflections. Computations made by approximate methods to determine the yawing moments contributed by the rudders at large deflections (that is, deflecting one rudder to 45° and maintaining the other rudder at neutral) showed that the adverse yawing moments contributed by a full aileron deflection could be overcome by the rudder. The effects of slipstream rotation were neglected for these calculations. Practical considerations probably prohibit the use of a rudder deflection, however, as high as 45°; and in order to maintain satisfactory flight characteristics, it thus appears necessary to increase the size of the vertical tails so that a smaller rudder deflection could be used. On the basis of previous experience in the spin tunnel, it appears that if the size of the fin and rudder are increased in a manner to maintain the same proportions as the existing fin and rudder the airplane would probably still be spinproof.\n\nThe test data obtained during the investigation were not extensive enough to permit determination of the control limitations necessary for spin-proofing at the higher wing loading.\n\nCONCLUSIONS\n\nThe results of spin tests of a $\\frac{1}{11}$-scale model of a twin-tail low-wing personal-owner-type airplane with controls linked and unlinked indicated the following spin and recovery characteristics at a test altitude of 5000 feet:", "timestamp": "2026-07-22T04:39:47.471079+00:00"}
{"citation_id": "19930085842", "source_url": "https://ntrs.nasa.gov/api/citations/19930085842/downloads/19930085842.pdf", "page_number": 83, "total_pages": 104, "image_filename": "19930085842_p83.jpg", "text": "NACA RM L9C29\n79\n\nPropeller advance-\ndiameter ratio, $V/nD$\n\n1.6\n1.2\n.8\n.4\n0\n\n$V/nD$\n\n$C_{D_R}$\n\n1200 bhp at 1085 rpm\n1380 bhp at 1085 rpm\n\nLift coefficient, $C_L$\n\n4\n3\n2\n1\n0\n\n$C_L$\n\n$\\beta$, deg\n$\\circ$ 10\n$\\circ$ 11.5\n$\\circ$ 14\n$\\circ$ 30\n\nResultant-drag coefficient, $C_{D_R}$\n\n2\n1\n0\n-1\n-2\n-3\n\nTorque coefficient, $Q_c$\n\n0\n0.2\n0.4\n0.6\n0.8\n1.0\n\nNATIONAL ADVISORY\nCOMMITTEE FOR AERONAUTICS\n\n(f) $\\alpha_u = 30^\\circ$.\nFigure 43.— Continued.", "timestamp": "2026-07-22T04:39:49.151462+00:00"}
{"citation_id": "19930092013", "source_url": "https://ntrs.nasa.gov/api/citations/19930092013/downloads/19930092013.pdf", "page_number": 12, "total_pages": 21, "image_filename": "19930092013_p12.jpg", "text": "```markdown\n8\nREPORT 948—NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nDISCUSSION\n\nCharacteristics of effective-dihedral control in steady sideslips.—During steady sideslips, the pilots noted that the apparatus did not cause the airplane to have an artificial feel; that is, it caused no significant difference in the handling characteristics of the airplane from those which would be expected of an airplane with actual dihedral of the magnitudes being simulated by the apparatus. Figure 6 shows desirable smooth and linear variations of pilot-applied aileron deflection and stick force with sideslip angle with the apparatus in operation. The slope of the pilot-applied aileron-deflection curve varies, for the maximum test aileron servo-gearing ratios, from a large stable value to a noticeably unstable value, corresponding to a stick-fixed $C_{l_\\beta}$ range from $-0.0033$ to $0.0006$ and a $\\Gamma_v$ range from $14.9^\\circ$ to $-2.7^\\circ$. Since the normal value of $\\Gamma_v$ is $6.3^\\circ$, the servo action caused about $\\pm 9^\\circ$ change. The slope of the stick-force curve varies from a large stable value to approximately zero, indicating large changes in stick-free dihedral effect. However, the force-curve slope is zero when the aileron-deflection-curve slope is unstable; thus, the desired condition that the stick-free $C_{l_\\beta}$ and the stick-fixed $C_{l_\\beta}$ equal zero at the same gearing ratio was not accomplished for these tests. This condition could be rectified by increasing the tab servo-gearing ratio $(d\\delta_t/d\\delta_a)_s$ or the tab effectiveness. The latter solution appeared preferable in the present application because of possible loss in effectiveness of the tab at the large deflections which would result from increases in $(d\\delta_t/d\\delta_a)_s$. The tab chord was increased for the tests reported in part II of this report and the desired results were more closely approximated. The sizable $\\Gamma_v$ range from $14.9^\\circ$ to $-2.7^\\circ$, which was obtained in these tests, corresponds closely to that originally desired for the investigation of the high-speed dynamic lateral characteristics with the test airplane. However, as more data and experience were gained with the apparatus, the aileron servo-gearing ratio was increased to give about twice the initial $\\Delta\\Gamma_v$ of $\\pm 9^\\circ$ so that the characteristics of a much wider range of configurations could be studied in part II.\n\nCharacteristics in abrupt rudder kicks.—The response in roll to a given abrupt rudder deflection is one measure of dihedral effect. The rolling-velocity curves of figure 7 show that, qualitatively, the apparatus successfully simulated large changes in dihedral effect under these severe dynamic conditions; the changes in maximum rolling velocity with servo-gearing ratio (and static $\\Gamma_v$) are readily apparent. It is seen that the actual servo-applied aileron deflection is in good agreement with the values computed from the sideslip angle for an ideal servo, except for a time lag of about 0.1 second during the initial portion of the maneuver. The records and computations showed that this time lag did not have a serious effect on the rolling response in simulating positive changes in $\\Gamma_v$ (fig. 7 (b)), but that it caused an undesired initial rolling response when attempting to simulate small negative values of $\\Gamma_v$. With the apparatus set for $\\Gamma_v = -2.7^\\circ$, the measured response to left rudder deflection showed an initial small left rolling velocity prior to the development of right rolling velocity (fig. 7 (c)). The computed response showed right rolling velocity throughout the maneuver. There was good\n\n<!-- Image (58, 75, 504, 461) -->\n\nFIGURE 9.—Concluded.\n\n<!-- Image (108, 494, 428, 768) -->\n\nFIGURE 10.—Lateral-oscillation period and damping characteristics. $\\Gamma_v=300$ knots.\n\nof $P$ and $C_{l_\\beta}$ predicted by the method of reference 8 are also shown. In general, the best available data on the mass and aerodynamic characteristics were employed in the computations, but minor adjustments were made to give correlation for the normal airplane in order to facilitate comparison of the measured and predicted effects of $C_{l_\\beta}$.\n```", "timestamp": "2026-07-22T04:39:49.405951+00:00"}
{"citation_id": "19930086081", "source_url": "https://ntrs.nasa.gov/api/citations/19930086081/downloads/19930086081.pdf", "page_number": 35, "total_pages": 44, "image_filename": "19930086081_p35.jpg", "text": "CONFIDENTIAL\n\nNACA RM L9H05\n\n.1\n$C_L$ 0\n-.1\n\n.008\n$C_l$\n0\n\n.04\n$C_m$ 0\n-.04\n\n.004\n\n.02\n$C_D$\n0\n\n0\n$C_n$\n-.004\n\n-2 0 2 4 6 8 10\n$\\delta$, deg\n\n-2 0 2 4 6 8 10\n$\\delta$, deg\n\nComplete wing Average from figure 10\nControl surface Fence off\nControl surface Large fence on\n\nCONFIDENTIAL\n\nNACA\n\nFigure 15.- Comparison between aerodynamic characteristics of the half-delta tip control surface and\nof the complete wing. Large fuselage, 3-percent-thick control. Variation with $\\delta$ at $\\alpha = 0^\\circ$;\n$R = 4.0 \\times 10^6$; $M = 1.90$.\n\n33", "timestamp": "2026-07-22T04:39:50.642198+00:00"}
{"citation_id": "19930086078", "source_url": "https://ntrs.nasa.gov/api/citations/19930086078/downloads/19930086078.pdf", "page_number": 39, "total_pages": 42, "image_filename": "19930086078_p39.jpg", "text": "NACA RM L9H04\n37\n\nCONFIDENTIAL\n\n<!-- Image (101, 175, 926, 831) -->\n\nFigure 16.- Lateral control characteristics of 45° sweptback wing with short-chord wing-tip aileron at various deflections, fully extended.", "timestamp": "2026-07-22T04:39:53.157799+00:00"}
{"citation_id": "19930086151", "source_url": "https://ntrs.nasa.gov/api/citations/19930086151/downloads/19930086151.pdf", "page_number": 29, "total_pages": 34, "image_filename": "19930086151_p29.jpg", "text": "NACA RM L9J28\n27\n\nCONFIDENTIAL\n\n○ Triangular aileron\n△ Parallelogram aileron\n□ Triangular tip aileron on wing of\nreference 5 (Mach number=0.5)\n} Present\ninvestigation\n\nEstimated $C_{\\ell}$\n0\n.01\n.02\n.03\n0\n.01\n.02\n.03\nExperimentally determined $C_{\\ell}$\n\n[Figure: A scatter plot comparing estimated $C_{\\ell}$ (y-axis) against experimentally determined $C_{\\ell}$ (x-axis). The plot includes data points represented by circles, triangles, and squares, which generally follow a diagonal line from the origin. The NACA logo is visible in the bottom right corner of the plot area.]\n\nCONFIDENTIAL\n\nFigure 11.— Comparison of experimentally determined values\nof $C_{\\ell}$ with estimated values of $C_{\\ell}$ for deflectable wing-\ntip ailerons in the presence of an end plate. $\\alpha=0^{\\circ}$.", "timestamp": "2026-07-22T04:39:54.314492+00:00"}
{"citation_id": "19930085930", "source_url": "https://ntrs.nasa.gov/api/citations/19930085930/downloads/19930085930.pdf", "page_number": 88, "total_pages": 92, "image_filename": "19930085930_p88.jpg", "text": "CONFIDENTIAL\nUNCLASSIFIED\n\nNACA RM L9G07\n\n[Figure: A shadowgraph image showing airflow patterns around an aerodynamic model, with visible shock waves and flow separation.]\n\nFigure 41.— A shadowgraph of the flow at an angle of attack of $29.83^\\circ$.\n\nCONFIDENTIAL\nUNCLASSIFIED\n\nNACA\n\n87", "timestamp": "2026-07-22T04:39:55.560927+00:00"}
{"citation_id": "19930082498", "source_url": "https://ntrs.nasa.gov/api/citations/19930082498/downloads/19930082498.pdf", "page_number": 2, "total_pages": 49, "image_filename": "19930082498_p2.jpg", "text": "NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nTECHNICAL NOTE NO. 1838\n\nDYNAMOMETER-STAND INVESTIGATION OF A GROUP OF MUFFLERS\n\nBy Don D. Davis, Jr., and K. R. Czarnecki\n\nSUMMARY\n\nAs part of a research program directed toward the reduction of airplane noises, an experimental investigation has been made of a series of exhaust mufflers installed on a typical six-cylinder light-airplane engine which was mounted on a ground dynamometer stand. The results show that engine and propeller noise are about equal for this installation; therefore, both engine-exhaust and propeller noises must be reduced to obtain a sizable reduction in over-all noise. Because most of the sound energy in the exhaust is found to be concentrated at low frequencies, methods applicable only to high-frequency sound are of little value in reducing the over-all sound-pressure level. The loudest single component of engine-exhaust noise is at the fundamental firing frequency of the engine. The types of present-day commercial airplane mufflers investigated are found to produce very little reduction in the over-all sound-pressure level. From the results of this exploratory investigation, several conclusions are drawn regarding the details of muffler design. For instance, the tail-pipe length is important, but the cross-sectional shape of the muffler (oval or circular) has no measurable effect on the muffler characteristics if the cross-sectional area and all other dimensions are constant.\n\nINTRODUCTION\n\nAs part of a general research program directed toward the reduction of airplane noises, a theoretical and experimental investigation of the methods of muffler design is being conducted at the Langley full-scale tunnel. In order to obtain a sizable reduction in over-all noise, the reduction of both engine-exhaust and propeller noises is necessary. Consequently, because of the generally noisy propellers, relatively little attention has been given to airplane-engine-exhaust muffling in the past. Increasing interest in airplane quieting has made desirable an attack on both of these noise sources. The present investigation was undertaken in order to find a muffler which, in conjunction with a relatively quiet propeller, would substantially reduce the noise of a light airplane.\n\nIn the course of this muffler investigation a large number of exhaust mufflers were installed on a typical six-cylinder light-airplane engine. The tests were conducted with the engine mounted with the propeller removed in a ground dynamometer stand. Over-all noise levels", "timestamp": "2026-07-22T04:39:57.762913+00:00"}
{"citation_id": "19930085838", "source_url": "https://ntrs.nasa.gov/api/citations/19930085838/downloads/19930085838.pdf", "page_number": 96, "total_pages": 118, "image_filename": "19930085838_p96.jpg", "text": "94\nNACA RM No. L9B23\n\n[Figure: Three plots of Flap section hinge-moment coefficient, $c_{h_f}$, versus Section angle of attack, $\\alpha_0$, deg. The y-axis ranges from -0.24 to 0.16. The x-axis ranges from -20 to 20.]\n\nTop plot:\n$\\delta_a = 5^\\circ$\n$\\delta_f$ (deg)\n$\\circ$ 0\n$\\square$ -1\n$\\diamond$ -3\n\nMiddle plot:\n$\\delta_a = 10^\\circ$\n$\\delta_f$ (deg)\n$\\circ$ 0\n$\\square$ -1\n$\\diamond$ -3\n\nBottom plot:\n$\\delta_a$ (deg) $\\delta_f$ (deg)\n$\\circ$ 15 0\n$\\square$ 15 -2\n$\\diamond$ 15 -4\n\n[NACA logo]\n\n(b) $\\delta_f = 25^\\circ$.\nFigure 12.- Continued.", "timestamp": "2026-07-22T04:40:00.344613+00:00"}
{"citation_id": "19930086015", "source_url": "https://ntrs.nasa.gov/api/citations/19930086015/downloads/19930086015.pdf", "page_number": 45, "total_pages": 54, "image_filename": "19930086015_p45.jpg", "text": "44\n\nCONFIDENTIAL\n\nDistance from wall, $\\delta$, in.\n\n| | | |\n| :--- | :--- | :--- |\n| $D=165.12$<br>$M=1.23$<br>$\\Delta p_t/q = -.01$ | $D=147.20$<br>$M=1.32$<br>$\\Delta p_t/q = -.02$ | $D=129.32$<br>$M=1.43$<br>$\\Delta p_t/q = -.02$ |\n| $D=113.43$<br>$M=1.53$<br>$\\Delta p_t/q = -.05$ | $D=100.47$<br>$M=1.63$<br>$\\Delta p_t/q = -.05$ | $D=88.52$<br>$M=1.73$<br>$\\Delta p_t/q = -.02$<br>[Figure: NACA logo] |\n\nRatio of total pressures, $\\frac{H}{H_0}$\n\n(c) Lower rear rake.\n\nFigure 12.—Concluded.\n\nCONFIDENTIAL\n\nNACA RM A52B4", "timestamp": "2026-07-22T04:40:01.504311+00:00"}
{"citation_id": "19930090382", "source_url": "https://ntrs.nasa.gov/api/citations/19930090382/downloads/19930090382.pdf", "page_number": 27, "total_pages": 37, "image_filename": "19930090382_p27.jpg", "text": "NACA RM L9I07\n29\n\nCONFIDENTIAL\n\nPower coefficient, $C_P$\nThrust coefficient, $C_T$\n\nTip Mach number, $M_t$\nEfficiency, $\\eta$\n\n[Figure: Graph showing curves for $C_P$, $C_T$, and $\\eta$ versus Advance ratio, J, for various $\\beta_{0.75R}$ angles (65°, 70°). The x-axis is labeled \"Advance ratio, J\" ranging from 0 to 9.0. The left y-axis is \"Power coefficient, $C_P$\" (0.0 to 1.20) and \"Thrust coefficient, $C_T$\" (0 to 3.00). The top y-axis is \"Tip Mach number, $M_t$\" (0 to 1.5). The right y-axis is \"Efficiency, $\\eta$\" (0 to 1.00). A NACA logo is present in the upper right corner of the plot area.]\n\nCONFIDENTIAL\n(i) M=0.75 Concluded\nFigure 5 - Continued.", "timestamp": "2026-07-22T04:40:06.733287+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 5, "total_pages": 99, "image_filename": "19930082511_p5.jpg", "text": "NACA TN No. 1826\n3\n\ntogether with numerous calculated results. In part III, a method of\nsolution for the circular open tunnel is given, together with numerical\nresults for the case of a lifting element concentrated at a point on the\naxis of the tunnel. The treatment in every case is a linear one in which\ndeformation of the jet boundary is considered to be small.\n\nThe parts were essentially independently prepared. Messrs. Gardner\nand Diesendruck contributed the analysis of part II. Mr. Eisenstadt\ncontributed part III. Dr. Katzoff contributed part I, and, in the absence\nof the others, prepared the numerical results of part II, made several\nminor revisions, and served as general editor of the whole.\n\nI - B O U N D A R Y C O N D I T I O N S A N D\nE L E C T R I C A L A N A L O G I E S\n\nIn part I, boundary conditions for an open wind tunnel are discussed\nwith special reference to the effects of the closed entrance and exit\nsections. It is shown that the velocity on the free surface is not\nnecessarily equal to the velocity far upstream in the closed portion and\nthat cross-flows may exist in the free surface, unlike the case of the\ninfinitely long open jet. A basic condition - analogous to the Kutta-\nJoukowski condition for the flow at the trailing edge of an airfoil - is\nthat the velocity be continuous at the entrance lip. Electrical analogies\nthat might be used for solving the flow in open wind tunnels are outlined.\nTwo types are described - one in which electrical potential corresponds\nto velocity potential, and another in which electrical potential corre-\nsponds to acceleration potential.\n\nBOUNDARY CONDITIONS\n\nRésumé of Prandtl's theory.- In Prandtl's original discussion, in\nwhich the entrance and exit regions are neglected, the tunnel is con-\nsidered as an infinitely long cylinder on the entire surface of which the\npressure is constant, whence, by Bernoulli's law, the velocity on the\nsurface is constant. If this velocity is considered as the sum of the\nundisturbed tunnel velocity U and a small perturbation velocity (u,v,w)\ndue to the presence of a body in the jet, the condition is then\nthat $(U + u)^2 + v^2 + w^2 \\approx U^2 + 2Uu = \\text{Constant}$, from which it is concluded\nthat u is constant over the entire surface. Furthermore, since u is\nobviously zero far in front of the body, it must be zero over the entire\nsurface.\n\nA corollary is that, on the jet surface, the perturbation velocity\nnormal to u (that is, the circumferential velocity) is also zero, as is\nreadily shown from a consideration of the rectangular path SPQR on the\nsurface of the jet. (See fig. 1(a).) As has just been shown, the", "timestamp": "2026-07-22T04:40:09.157586+00:00"}
{"citation_id": "19930082487", "source_url": "https://ntrs.nasa.gov/api/citations/19930082487/downloads/19930082487.pdf", "page_number": 9, "total_pages": 33, "image_filename": "19930082487_p9.jpg", "text": "NACA TN No. 1813\n\nagreement with measured values of the drag-divergence Mach number. The experimental value of $M_{\\beta}$ at a given angle of attack was obtained from a plot of the experimentally determined pressure coefficient at the airfoil crest as a function of free-stream Mach number. The free-stream Mach number at which this curve passed through a value corresponding to a local Mach number of unity was termed $M_{\\beta \\text{exp}}$. The calculations consisted of determining the free-stream Mach number at which sonic velocity was reached at the airfoil crest by applying the Prandtl-Glauert rule to the low-speed experimental pressure coefficient at the airfoil crest. For comparison, there are also shown in figure 5 the critical Mach number $M_{cr}$ and the shock-stall Mach number $M_s$ at which marked boundary-layer thickening or separation is first indicated by negatively increasing pressure coefficients over the tail of the airfoil. It is seen that these latter two curves are not closely related to drag divergence. Similar data for several other 15-percent-thick airfoil sections are presented in figure 6. For all these cases, the Mach number for which the velocity at the crest is calculated to be sonic is close to the drag-divergence Mach number evaluated from the section drag characteristics.\n\nThe five airfoil sections, for which drag-divergence has been related to the occurrence of sonic velocity at the airfoil crest, are relatively thick sections. Since no rigorous causal relation between these phenomena has been demonstrated, it is desirable to investigate the applicability of the method for thinner airfoil sections. In figure 7, calculated values of $M_{\\beta}$ and measured drag-coefficient variation with Mach number are presented for a large number of airfoil sections. The calculations consisted of applying the Prandtl-Glauert compressibility correction to the low-speed pressure distributions in order to determine the Mach number at which sonic velocity occurs at the airfoil crest. The drag data for the 6-percent-thick symmetrical sections were obtained from reference 2 and the data for the cambered 6-series airfoils were measured in the Ames 1- by 3-1/2-foot high-speed wind tunnel on 6-inch-chord models which spanned the 1-foot dimension of the tunnel. It is seen that for each of these airfoil sections a useful measure of the free-stream Mach number at which drag divergence occurs is that Mach number at which the calculated velocity at the airfoil crest is sonic.\n\nDISCUSSION OF FLOW CHARACTERISTICS AT MACH NUMBERS\n\nABOVE THE DRAG-DIVERGENCE MACH NUMBER\n\nIn the previous sections of the report some characteristics of the supercritical flow past airfoil sections at subsonic free-stream", "timestamp": "2026-07-22T04:40:12.162848+00:00"}
{"citation_id": "19930086081", "source_url": "https://ntrs.nasa.gov/api/citations/19930086081/downloads/19930086081.pdf", "page_number": 36, "total_pages": 44, "image_filename": "19930086081_p36.jpg", "text": "```markdown\n34\nNACA RM L9H05\n\nCONFIDENTIAL\n\n<!-- Image (70, 169, 853, 766) -->\n\n(a) Normal force plotted against $\\alpha$.\n\nFigure 16.- Aerodynamic loading characteristics of a 0.03t/c half-delta control surface. Fence off. Data presented with respect to control surface axes. R = $4.0 \\times 10^6$; M = 1.90. Second series of tests.\n```", "timestamp": "2026-07-22T04:40:12.363259+00:00"}
{"citation_id": "19930085842", "source_url": "https://ntrs.nasa.gov/api/citations/19930085842/downloads/19930085842.pdf", "page_number": 84, "total_pages": 104, "image_filename": "19930085842_p84.jpg", "text": "80\nNACA RM L9C29\n\nPropeller advance-\ndiameter ratio, $V/nD$\n1.2\n8\n4\n0\n\n$V/nD$\n\n$C_{QR}$\n\nLift coefficient, $C_L$\n7\n6\n5\n4\n3\n2\n\n$C_L$\n\n$\\beta$, deg\n$\\circ$ 10\n$\\square$ 11.9\n$\\diamond$ 14\n\nResultant drag coefficient, $C_R$\n1\n0\n-1\n-2\n-3\n-4\n-5\n\n0\n0.2\n0.4\n0.6\n0.8\n1.0\n1.2\n1.4\n1.6\n1.8\nTorque coefficient, $Q_c$\n\nNATIONAL ADVISORY\nCOMMITTEE FOR AERONAUTICS\n\n(g) $\\alpha_1 = 36^\\circ$.\nFigure 43.- Continued.", "timestamp": "2026-07-22T04:40:13.816337+00:00"}
{"citation_id": "19930082476", "source_url": "https://ntrs.nasa.gov/api/citations/19930082476/downloads/19930082476.pdf", "page_number": 13, "total_pages": 41, "image_filename": "19930082476_p13.jpg", "text": "NACA TN No. 1801\n\nFor linked rudder and aileron controls:\n\n1. For the normal loading condition, spins were obtainable only when the wheel was placed approximately one-half with the spin and the elevator was deflected upward to at least $8^\\circ$. Setting the wheel farther with the spin lead to a motion that appeared to be a spiral, and setting the wheel laterally to neutral prevented the spin. Moving the elevator down was favorable in preventing the spin. Recoveries obtained by fully reversing the wheel followed by moving the elevator down would undoubtedly have been rapid from any spin.\n\n2. With the mass increased along the fuselage, more spins were obtained with the elevator between neutral and full up for wheel settings with the spin than were obtained for the normal loading condition. With the mass increased along the wings, the results were very similar to those obtained for the normal loading.\n\n3. Approximately doubling the airplane's relative density led to definite spins when the wheel was set full with the spin and the elevator was set to its normal full-up deflection (normal spinning control configuration), but for other wheel and elevator settings little effect was noted.\n\nFor unlinked controls:\n\n4. For all loadings ailerons set against the spin tended to prevent the spin; whereas ailerons set with the spin were conducive to the attainment of spinning equilibrium. Deflecting the inboard aileron up was particularly effective in maintaining the spin, especially when it was deflected from approximately three-tenths to six-tenths of its maximum full-up deflection.\n\n5. The outboard rudder was effective in terminating or maintaining the spin when the ailerons were neutral. For loadings with mass extended along the wings, rudder reversal would have to be followed by elevator reversal in order to effect recovery from the aileron-with spins. With the ailerons neutral, differential rudder deflections which maintained the outboard rudder at or near neutral were particularly effective in preventing the attainment of spinning equilibrium.\n\n6. When the corresponding full-scale wing loading of the model was 10 pounds per square foot, it was indicated that spinproofing could be obtained by limiting the aileron movement to $5^\\circ$ up, by limiting the outboard rudder movement so that it could not be deflected with the spin, and by limiting the up elevator deflection to $13^\\circ$. With the controls limited in this manner, an inboard rudder deflection of $45^\\circ$ would be required to provide satisfactory flight characteristics. Inasmuch as a rudder", "timestamp": "2026-07-22T04:40:14.067993+00:00"}
{"citation_id": "19930082245", "source_url": "https://ntrs.nasa.gov/api/citations/19930082245/downloads/19930082245.pdf", "page_number": 17, "total_pages": 66, "image_filename": "19930082245_p17.jpg", "text": "16\nNACA TN No. 1596\n\nREFERENCES\n\n1. Stevenson, David B., and Byrne, Robert W.: High-Speed Wind-Tunnel Tests of an NACA 16-009 Airfoil Having a 32.9-Percent-Chord Flap with an Overhang 20.7 Percent of the Flap Chord. NACA TN No. 1406, 1947.\n\n2. Stevenson, David B., and Adler, Alfred A.: High-Speed Wind-Tunnel Tests of an NACA 0009-64 Airfoil Having a 33.4-Percent-Chord Flap with an Overhang 20.1 Percent of the Flap Chord. NACA TN No. 1417, 1947.\n\n3. Lindsey, W. F.: Effect of Compressibility on the Pressures and Forces Acting on a Modified NACA 65,3-019 Airfoil Having a 0.20-Chord Flap. NACA ACR No. L5G31a, 1946.\n\n4. Jones, Robert T., and Ames, Milton B., Jr.: Wind-Tunnel Investigation of Control-Surface Characteristics. V - The Use of a Beveled Trailing Edge to Reduce the Hinge Moment of a Control Surface. NACA ARR, March 1942.\n\n5. Allen, H. Julian, and Vincenti, Walter G.: The Wall Interference in a Two-Dimensional-Flow Wind Tunnel with Consideration of the Effect of Compressibility. NACA Rep. No. 782, 1944.\n\n6. Vincenti, Walter G., and Graham, Donald J.: The Effect of Wall Interference upon the Aerodynamic Characteristics of an Airfoil Spanning a Closed-Throat Circular Wind Tunnel. NACA ACR No. 5D21, 1945.\n\n7. Underwood, William J., Braslow, Albert L., and Cahill, Jones F.: Two-Dimensional Wind-Tunnel Investigation of 0.20-Airfoil-Chord Plain Ailerons of Different Contour on an NACA 651-210 Airfoil Section. NACA ACR No. L5F27, 1945.\n\n8. Rogallo, F. M., and Purser, Paul E.: Wind-Tunnel Investigation of a Plain Aileron with Various Trailing-Edge Modifications on a Tapered Wing. II - Ailerons with Thickened and Beveled Trailing Edges. NACA ARR, Oct. 1942.\n\n9. Purser, Paul E., and McKee, John W.: Wind-Tunnel Investigation of a Plain Aileron with Thickened and Beveled Trailing Edges on a Tapered Low-Drag Wing. NACA ACR, Jan. 1943.\n\n10. Rogallo, F. M., and Crandall, Stewart M.: Wind-Tunnel Investigation of Trimming Tabs on a Thickened and Beveled Aileron on a Tapered Low-Drag Wing. NACA ACR, March 1943.\n\n11. Robinson, Harold L.: High-Speed Investigation of Skin Wrinkles on Two NACA Airfoils. NACA TN No. 1121, 1946.", "timestamp": "2026-07-22T04:40:14.981743+00:00"}
{"citation_id": "19930085930", "source_url": "https://ntrs.nasa.gov/api/citations/19930085930/downloads/19930085930.pdf", "page_number": 89, "total_pages": 92, "image_filename": "19930085930_p89.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T04:40:15.165948+00:00"}
{"citation_id": "19930086078", "source_url": "https://ntrs.nasa.gov/api/citations/19930086078/downloads/19930086078.pdf", "page_number": 40, "total_pages": 42, "image_filename": "19930086078_p40.jpg", "text": "```markdown\n38\nNACA RM L9H04\n\nCONFIDENTIAL\n\n<!-- Image (69, 130, 875, 838) -->\n\nFigure 17.- Lateral control characteristics of 45° sweptback wing with short-chord wing-tip aileron at various extensions. $\\delta_a = 4^\\circ$.\n```", "timestamp": "2026-07-22T04:40:16.761563+00:00"}
{"citation_id": "19930082485", "source_url": "https://ntrs.nasa.gov/api/citations/19930082485/downloads/19930082485.pdf", "page_number": 11, "total_pages": 62, "image_filename": "19930082485_p11.jpg", "text": "10\nNACA TN No. 1810\n\nRadial-equilibrium consideration. - As stated in the discussion of blade design, the blades were designed on the assumption of radial equilibrium of pressure. This condition may be stated as follows:\n\n$$\n\\frac{dP_s}{dr} = \\frac{\\rho_s V_u^2}{r}\n$$\n\nThis equation states that the centrifugal force on the air particles is balanced by a pressure difference across the particles. However, because radial flow must take place as the gases pass through the passage between blades if the exit axial velocity is to remain constant at all radii, the gas will be accelerated radially outward. These radial acceleration forces can be considered by writing\n\n$$\n\\frac{dP_s}{dr} = \\frac{\\rho_s V_u^2}{r} - \\rho_s \\frac{dV_r}{dt}\n$$\n\nThis expression may be written in dimensionless form for convenience as\n\n$$\n\\frac{\\gamma+1}{2\\gamma} \\frac{r}{P_t} \\frac{dP_s}{dr} = \\frac{\\rho_s}{\\rho_t} \\left( \\frac{V_u}{V_{cr}} \\right)^2 - \\frac{r}{V_{cr}^2} \\frac{\\rho_s}{\\rho_t} \\frac{dV_r}{dt}\n$$\n\nThe design assumption in effect neglected the radial acceleration of velocity term involving $\\frac{dV_r}{dt}$.\n\nThe magnitude of the radial-acceleration parameter involving the rate of change of radial velocity with time at the experimental conditions is shown in figure 12. The radial-pressure-gradient parameter $\\frac{r}{P_t} \\frac{dP_s}{dr}$ was calculated by graphically differentiating the lower curve of figure 7. The centrifugal-force parameter $\\frac{\\rho_s}{\\rho_t} \\left( \\frac{V_u}{V_{cr}} \\right)^2$ was computed from figures 10(b) and 7 by assuming isentropic flow.\n\nThe radial-acceleration parameter $\\frac{r}{V_{cr}^2} \\frac{\\rho_s}{\\rho_t} \\frac{dV_r}{dt}$ was then evaluated by subtracting $\\frac{\\gamma+1}{2\\gamma} \\frac{r}{P_t} \\frac{dP_s}{dr}$ from $\\frac{\\rho_s}{\\rho_t} \\left( \\frac{V_u}{V_{cr}} \\right)^2$. Although", "timestamp": "2026-07-22T04:40:18.848596+00:00"}
{"citation_id": "19930086151", "source_url": "https://ntrs.nasa.gov/api/citations/19930086151/downloads/19930086151.pdf", "page_number": 30, "total_pages": 34, "image_filename": "19930086151_p30.jpg", "text": "```markdown\n28\nNACA RM L9J28\n\nCONFIDENTIAL\n\n<!-- Image (352, 102, 728, 316) -->\n\n| Differential | $\\delta_{a_{max}}$ | $\\beta$ | $\\phi$ |\n| :--- | :--- | :--- | :--- |\n| 1:1 | 29.5°, -29.5° | 0° | 0° |\n| 2:1 | 15.5°, -29.5° | 40° | 6.9° |\n| 3:1 | 11.2°, -29.5° | 47° | 11.8° |\n\n<!-- Image (125, 490, 836, 882) -->\n\nFigure 12.— Differential linkage systems used for determining $\\frac{pb}{2V}$.\n```", "timestamp": "2026-07-22T04:40:19.543538+00:00"}
{"citation_id": "19930085965", "source_url": "https://ntrs.nasa.gov/api/citations/19930085965/downloads/19930085965.pdf", "page_number": 51, "total_pages": 67, "image_filename": "19930085965_p51.jpg", "text": "50\n\n$\\beta_2 - \\delta_2$, deg\n\n| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | |", "timestamp": "2026-07-22T04:40:22.280556+00:00"}
{"citation_id": "19930085838", "source_url": "https://ntrs.nasa.gov/api/citations/19930085838/downloads/19930085838.pdf", "page_number": 97, "total_pages": 118, "image_filename": "19930085838_p97.jpg", "text": "NACA RM No. L9B23\n\n95\n\n| | |\n| :--- | :--- |\n| $\\delta_a$ | $\\delta_s$ |\n| (deg) | (deg) |\n| $\\circ$ -5 | 0 |\n| $\\square$ -5 | +5 |\n\n| | |\n| :--- | :--- |\n| $\\delta_a = -5^\\circ$ | |\n| $\\delta_s$ | |\n| (deg) | |\n| $\\circ$ -7 | |\n| $\\square$ -10 | |\n\n| | |\n| :--- | :--- |\n| $\\delta_a = -10^\\circ$ | |\n| $\\delta_s$ | |\n| (deg) | |\n| $\\circ$ 0 | |\n| $\\square$ -7 | |\n| $\\diamond$ -10 | |\n| $\\triangle$ -15 | |\n\nFlip section hinge-moment coefficient, $c_{h_f}$\n\nSection angle of attack, $\\alpha_c$, deg\n\n(c) $\\delta_f = 25^\\circ$.\n\nFigure 12.- Continued.", "timestamp": "2026-07-22T04:40:24.411834+00:00"}
{"citation_id": "19930082450", "source_url": "https://ntrs.nasa.gov/api/citations/19930082450/downloads/19930082450.pdf", "page_number": 12, "total_pages": 37, "image_filename": "19930082450_p12.jpg", "text": "NACA TN No. 1778\n\nTABLE 2.- Z-PANEL PROPERTIES - Concluded $\\left[\\frac{t_{w}}{t_{s}}=0.51 ; \\frac{t_{b}}{t_{w}}=11.4 ; \\frac{t_{b}}{b_{w}}=0.4 ; \\frac{t_{a}}{t_{b}}=3 ; \\frac{t_{s}}{t_{B}}=4 ; \\frac{t_{s}}{t_{B}}=1.50 ; \\frac{b_{w}}{t_{s}}=10.0\\right]$\n\n| $b_{w}/t_{B}$ | 33 | 34 | 35 | 36 | 37 | 38 | 39 | 40 | 41 | 42 | 43 | 44 | 45 |\n|---|---|---|---|---|---|---|---|---|---|---|---|---|---|\n| 25 | 1.641 | 1.573 | 1.503 | 1.407 | 1.422 | 1.634 | 1.651 | 1.665 | 1.670 | 1.665 | 1.710 | 1.721 | 1.742 |\n| 26 | 1.612 | 1.546 | 1.470 | 1.384 | 1.400 | 1.612 | 1.626 | 1.640 | 1.684 | 1.655 | 1.692 | 1.703 | 1.716 |\n| 27 | 1.582 | 1.519 | 1.510 | 1.362 | 1.379 | 1.590 | 1.609 | 1.616 | 1.630 | 1.613 | 1.657 | 1.670 | 1.681 |\n| 28 | 1.554 | 1.493 | 1.420 | 1.340 | 1.359 | 1.568 | 1.581 | 1.593 | 1.606 | 1.585 | 1.625 | 1.646 | 1.659 |\n| 29 | 1.526 | 1.463 | 1.511 | 1.323 | 1.336 | 1.540 | 1.561 | 1.571 | 1.586 | 1.559 | 1.612 | 1.628 | 1.637 |\n| 30 | 1.499 | 1.438 | 1.370 | 1.303 | 1.313 | 1.513 | 1.532 | 1.542 | 1.557 | 1.531 | 1.581 | 1.597 | 1.614 |\n| 31 | 1.455 | 1.466 | 1.479 | 1.490 | 1.502 | 1.513 | 1.525 | 1.537 | 1.549 | 1.560 | 1.572 | 1.584 | 1.596 |\n| 32 | 1.441 | 1.452 | 1.464 | 1.474 | 1.486 | 1.497 | 1.509 | 1.520 | 1.532 | 1.543 | 1.554 | 1.565 | 1.577 |\n| 33 | 1.427 | 1.438 | 1.449 | 1.460 | 1.471 | 1.482 | 1.493 | 1.504 | 1.515 | 1.526 | 1.537 | 1.548 | 1.559 |\n| 34 | 1.415 | 1.425 | 1.436 | 1.446 | 1.457 | 1.468 | 1.479 | 1.490 | 1.500 | 1.511 | 1.522 | 1.532 | 1.543 |\n| 35 | 1.403 | 1.413 | 1.423 | 1.433 | 1.443 | 1.453 | 1.463 | 1.473 | 1.483 | 1.493 | 1.503 | 1.513 | 1.523 |\n| 36 | 1.391 | 1.401 | 1.411 | 1.422 | 1.432 | 1.442 | 1.452 | 1.462 | 1.472 | 1.482 | 1.493 | 1.503 | 1.513 |\n| 37 | 1.381 | 1.390 | 1.400 | 1.410 | 1.420 | 1.430 | 1.440 | 1.450 | 1.460 | 1.470 | 1.480 | 1.490 | 1.500 |\n| 38 | 1.371 | 1.380 | 1.390 | 1.399 | 1.409 | 1.419 | 1.429 | 1.438 | 1.448 | 1.457 | 1.467 | 1.476 | 1.486 |\n| 39 | 1.362 | 1.371 | 1.380 | 1.389 | 1.398 | 1.407 | 1.416 | 1.425 | 1.434 | 1.443 | 1.452 | 1.461 | 1.470 |\n| 40 | 1.352 | 1.361 | 1.370 | 1.380 | 1.389 | 1.398 | 1.407 | 1.416 | 1.425 | 1.434 | 1.443 | 1.452 | 1.462 |\n| 42 | 1.336 | 1.344 | 1.353 | 1.361 | 1.370 | 1.379 | 1.388 | 1.396 | 1.405 | 1.413 | 1.422 | 1.430 | 1.439 |\n| 44 | 1.321 | 1.328 | 1.337 | 1.345 | 1.353 | 1.362 | 1.370 | 1.378 | 1.387 | 1.395 | 1.403 | 1.411 | 1.420 |\n| 46 | 1.306 | 1.314 | 1.322 | 1.330 | 1.338 | 1.346 | 1.354 | 1.362 | 1.370 | 1.377 | 1.385 | 1.393 | 1.401 |\n| 48 | 1.294 | 1.301 | 1.309 | 1.316 | 1.324 | 1.331 | 1.339 | 1.346 | 1.354 | 1.361 | 1.368 | 1.376 | 1.383 |\n| 50 | 1.281 | 1.288 | 1.297 | 1.304 | 1.311 | 1.318 | 1.326 | 1.333 | 1.340 | 1.347 | 1.355 | 1.362 | 1.369 |\n| 52 | 1.271 | 1.278 | 1.285 | 1.292 | 1.299 | 1.306 | 1.313 | 1.320 | 1.327 | 1.334 | 1.341 | 1.348 | 1.355 |\n| 54 | 1.261 | 1.268 | 1.274 | 1.281 | 1.288 | 1.295 | 1.302 | 1.308 | 1.315 | 1.322 | 1.329 | 1.335 | 1.342 |\n| 56 | 1.252 | 1.258 | 1.265 | 1.271 | 1.278 | 1.284 | 1.291 | 1.297 | 1.304 | 1.310 | 1.317 | 1.323 | 1.330 |\n| 58 | 1.244 | 1.250 | 1.256 | 1.262 | 1.268 | 1.274 | 1.281 | 1.287 | 1.293 | 1.299 | 1.305 | 1.312 | 1.318 |\n| 60 | 1.236 | 1.242 | 1.248 | 1.254 | 1.260 | 1.266 | 1.272 | 1.278 | 1.284 | 1.290 | 1.296 | 1.302 | 1.308 |\n| 65 | 1.217 | 1.222 | 1.228 | 1.234 | 1.239 | 1.245 | 1.250 | 1.256 | 1.262 | 1.267 | 1.273 | 1.278 | 1.284 |\n| 70 | 1.201 | 1.206 | 1.211 | 1.217 | 1.222 | 1.227 | 1.233 | 1.238 | 1.243 | 1.248 | 1.253 | 1.258 | 1.264 |\n| 75 | 1.188 | 1.193 | 1.198 | 1.202 | 1.207 | 1.212 | 1.217 | 1.222 | 1.227 | 1.232 | 1.237 | 1.242 | 1.247 |\n\n| $b_{w}/t_{B}$ | 33 | 34 | 35 | 36 | 37 | 38 | 39 | 40 | 41 | 42 | 43 | 44 | 45 |\n|---|---|---|---|---|---|---|---|---|---|---|---|---|---|\n| 25 | 3.876 | 4.049 | 4.223 | 4.402 | 4.581 | 4.764 | 4.950 | 5.140 | 5.320 | 5.520 | 5.713 | 5.913 | 6.110 |\n| 26 | 3.792 | 3.961 | 4.132 | 4.306 | 4.483 | 4.663 | 4.844 | 5.028 | 5.216 | 5.405 | 5.597 | 5.790 | 5.936 |\n| 27 | 3.712 | 3.878 | 4.043 | 4.212 | 4.384 | 4.564 | 4.740 | 4.920 | 5.104 | 5.290 | 5.477 | 5.666 | 5.855 |\n| 28 | 3.637 | 3.799 | 3.962 | 4.131 | 4.301 | 4.474 | 4.649 | 4.827 | 5.007 | 5.190 | 5.375 | 5.562 | 5.751 |\n| 29 | 3.565 | 3.723 | 3.884 | 4.050 | 4.218 | 4.389 | 4.561 | 4.737 | 4.915 | 5.097 | 5.280 | 5.465 | 5.653 |\n| 30 | 3.495 | 3.650 | 3.811 | 3.970 | 4.139 | 4.301 | 4.473 | 4.640 | 4.814 | 4.991 | 5.170 | 5.351 | 5.535 |\n| 31 | 3.426 | 3.581 | 3.733 | 3.894 | 4.054 | 4.220 | 4.385 | 4.553 | 4.723 | 4.896 | 5.071 | 5.251 | 5.430 |\n| 32 | 3.362 | 3.513 | 3.665 | 3.820 | 3.980 | 4.142 | 4.303 | 4.470 | 4.636 | 4.808 | 4.982 | 5.156 | 5.334 |\n| 33 | 3.303 | 3.451 | 3.601 | 3.754 | 3.909 | 4.067 | 4.228 | 4.391 | 4.556 | 4.724 | 4.894 | 5.067 | 5.241 |\n| 34 | 3.246 | 3.390 | 3.537 | 3.689 | 3.841 | 3.997 | 4.152 | 4.315 | 4.476 | 4.644 |", "timestamp": "2026-07-22T04:40:24.617336+00:00"}
{"citation_id": "19930090382", "source_url": "https://ntrs.nasa.gov/api/citations/19930090382/downloads/19930090382.pdf", "page_number": 28, "total_pages": 37, "image_filename": "19930090382_p28.jpg", "text": "30\nNACA RM L9I07\n\nCONFIDENTIAL\n\nTip Mach number, $M_t$\n1.5\n1.0\n0.5\n0\n\nEfficiency, $\\eta$\n1.00\n.75\n.50\n.25\n0\n\nThrust coefficient, $C_T$\n.300\n.275\n.250\n.225\n.200\n.175\n.150\n.125\n.100\n.075\n.050\n.025\n0\n\nPower coefficient, $C_P$\n.60\n.55\n.50\n.45\n.40\n.35\n.30\n.25\n.20\n.15\n.10\n.05\n0.0\n\nAdvance ratio, J\n1.5\n2.0\n2.5\n3.0\n3.5\n4.0\n4.5\n\n$M_t$\n$C_P$\n$C_T$\n$\\eta$\n$\\beta_{0.75}=55^\\circ$\n$60^\\circ$\n\nCONFIDENTIAL\n\n(1) M=0.80.\nFigure 5 - Continued.", "timestamp": "2026-07-22T04:40:26.241288+00:00"}
{"citation_id": "19930082498", "source_url": "https://ntrs.nasa.gov/api/citations/19930082498/downloads/19930082498.pdf", "page_number": 3, "total_pages": 49, "image_filename": "19930082498_p3.jpg", "text": "2\nNACA TN No. 1838\n\nand noise spectrums were determined over a range of engine speed on the\nbasic engine and on the engine with various muffler configurations\nattached. The results of ground tests of a low-frequency-pass multiple-\nresonant-chamber muffler are given in reference 1. This muffler was\ninstalled and flown on a light airplane in conjunction with a relatively\nquiet five-blade propeller for a demonstration of a \"quiet\" airplane\n(reference 2). Further experimental results of the muffler investigation\nare given in this paper in order to convey an idea of the relative\nperformance of a large number of muffler configurations. The merits of\na muffler must be based on physical size, weight, back pressure, and\nthe annoyance of the exhaust noise. The annoyance depends upon the\nintensity and frequency of the noise and upon the particular person\nlistening. At the present time no way exists to evaluate accurately\nthe annoyance from objective measurements. In the present paper,\ntherefore, muffler acoustic performance is given in terms of both the\nmeasured over-all sound-pressure levels and the frequency distribution\nof the sound.\n\nMUFFLERS\n\nThe group of mufflers tested was composed of standard commercial\nmufflers, a special muffler designed for this project by a muffler\nmanufacturer, a muffler constructed from a drawing shown in reference 3,\nand mufflers designed at the Langley Laboratory. The commercial mufflers\nwere constructed of stainless steel. The mufflers built at the Langley\nLaboratory were constructed of mild steel tubing and sheet, since they\nwere not intended for actual flight use.\n\nSketches of these mufflers, all drawn to the same scale, are\nincluded in table II. Unless otherwise specified, the flow is from\nleft to right. Three-dimensional sketches showing internal details\nof several of these mufflers are given in figure 1 and photographs of\nseveral of the other mufflers are shown in figure 2.\n\nAPPARATUS AND TECHNIQUES\n\nThe test engine is a direct-drive, four-stroke, opposed six-\ncylinder engine of 435-cubic-inch displacement rated at 185 horsepower\nat 2550 rpm at sea level. This engine develops about 200 horsepower\nat 2790 rpm. The engine is equipped with two exhaust manifolds, one\non each side of the engine as shown in figure 3. In order to install\nthe mufflers the exit cones with the longitudinal slits at the ends of\nthe exhaust pipes were removed, as shown in figure 4.", "timestamp": "2026-07-22T04:40:27.265671+00:00"}
{"citation_id": "19930086015", "source_url": "https://ntrs.nasa.gov/api/citations/19930086015/downloads/19930086015.pdf", "page_number": 46, "total_pages": 54, "image_filename": "19930086015_p46.jpg", "text": "```markdown\nCONFIDENTIAL\n\nNACA RM A9E24\n\n.60\n.50\n.40\n.30\n.20\n.10\n0\n-.10\n\nLift coefficient, $C_L$\n\nDrag coefficient, $C_D$\n0 .01 .02 .03 .04 .05 .06 .07\n\nAngle of attack, $\\alpha$, deg\n0 2 4 6 8 10\n\nPitching-moment coefficient, $C_{mc_{\\bar{c}}/4}$\n0 -.04 -.08 -.12 -.16 -.20\n\no vertical\n$\\diamond$ upright\n$\\square$ inverted\n\n(a) D=165.12; M=1.23.\n\nFigure 13.-Aerodynamic characteristics of a wing having the leading edge swept back 63° from tests of the wing in the vertical, upright, and inverted positions in the Ames 6- by 6-foot supersonic wind tunnel.\n\nCONFIDENTIAL\n\n45\n```", "timestamp": "2026-07-22T04:40:29.092658+00:00"}
{"citation_id": "19930082496", "source_url": "https://ntrs.nasa.gov/api/citations/19930082496/downloads/19930082496.pdf", "page_number": 4, "total_pages": 50, "image_filename": "19930082496_p4.jpg", "text": "NACA TN No. 1836\n\nfor a number of years in the tool industry. The good strength at red heat, compared with metals, combined with extreme hardness made cemented carbides well suited for cutting tools. A large amount of research has been conducted towards obtaining fundamental data on cemented-carbide constituents in order to develop superior compositions. Fundamental data on a number of cemented-carbide constituents are given in reference 4.\n\nAn extremely important study has been the development of fabrication techniques because the characteristics of the ceramal are highly dependent upon the method of fabrication. Considerable information relative to the preparation of carbide powders and the fabrication of cemented carbides is contained in several United States patents. (For example, see reference 5.) This background in the fabrication of cemented carbides for tool applications aided in producing sound specimens for evaluation without embarking on a program of development in the laboratory. The earliest research on ceramals for gas turbines was conducted in Germany (reference 6).\n\nA material having the characteristics of high thermal conductivity, high tensile strength, low thermal expansion, and low modulus of elasticity would be expected to possess high thermal-shock resistance. Reference 2 indicates for a ceramal consisting of 80-percent titanium carbide and 20-percent cobalt (by weight) a relatively high thermal conductivity ($20.56 \\text{ Btu/(hr)(sq ft)}(^{\\circ}\\text{F/ft})$), a high tensile strength, a low thermal expansion ($5.0 \\times 10^{-6} \\text{ (in./in.)}/^{\\circ}\\text{F}$), a modulus of elasticity of 55,000,000 pounds per square inch, and good tensile strength, from which the thermal-shock resistance was concluded to be adequate for turbine-blade use. A favorable strength-to-weight ratio is also indicated in reference 2.\n\nAn investigation was conducted at the NACA Lewis laboratory to determine experimentally the resistance to thermal shock and the short-time tensile strength at elevated temperatures of a carbide-type ceramal and the performance characteristics of the carbide-type ceramal blades operated under quasi-service conditions.\n\nThe particular determinations made were:\n\n(a) Short-time tensile strength at $1800^{\\circ}$ and $2200^{\\circ} \\text{ F}$\n\n(b) Thermal-shock resistance at $1800^{\\circ}$, $2000^{\\circ}$, $2200^{\\circ}$, and $2400^{\\circ} \\text{ F}$", "timestamp": "2026-07-22T04:40:29.797289+00:00"}
{"citation_id": "19930092013", "source_url": "https://ntrs.nasa.gov/api/citations/19930092013/downloads/19930092013.pdf", "page_number": 13, "total_pages": 21, "image_filename": "19930092013_p13.jpg", "text": "```markdown\nAPPARATUS FOR VARYING EFFECTIVE DIHEDRAL IN FLIGHT\n9\n\nagreement, however, between the values of computed and measured maximum rolling velocity and the times at which they occurred. Additional step-by-step response calculations were made, and the results showed that a lag of 0.1 second would account for the undesired initial left roll. Close examination of position-recorder data showed that the lag in servo-applied aileron deflection was due partially to lag in the servomechanism response and partially to stretch in the servo-control system between the servo motor and the torque tube. Suitable minor adjustments in the electrical circuit might improve the servo response in future tests. Reduction of the lag due to stretch could be accomplished by a reduction in the flexibility and inertia of the control system, but this would necessitate major changes in the present apparatus.\n\nThe effect of this small amount of lag was noticeable to the pilots, but, in their opinion, it did not cause the airplane to have an artificial feel. To the pilots, the small amount of reverse rolling velocity appeared as an effect of yawing velocity, and their opinion was that the apparatus satisfactorily simulated changes in dihedral.\n\nA quantitative measure of the effect of the apparatus in changing the rolling response in abrupt rudder kicks is given by figure 8. The small initial undesired rolling motion at the $-2.7^\\circ \\Gamma_v$ setting was ignored in deriving these data. It is seen from figure 8 (a) that the maximum value of the rolling parameter $pb/2V$ per unit rudder deflection, a measure of dihedral effect, was varied over the wide range from the normal value of 0.005 to 0.011 and $-0.002$. Figure 8 (b) shows good agreement between the measured and predicted effects of $\\Gamma_v$ on the rolling response, and thus indicates that the effects of the desired changes in $\\Gamma_v$ are simulated by the apparatus under severe dynamic conditions.\n\nThe initial small adverse rolling motion experienced in rudder kicks when the apparatus was used to simulate the $-2.7^\\circ$ value of $\\Gamma_v$ did not prove a serious deficiency in applying the apparatus to the further dynamic-stability studies discussed in part II. This effect was most noticeable when attempting to simulate small negative dihedral effect. As more negative values of $\\Gamma_v$ were obtained by increasing the servo gearing, the initial left rolling velocity following a left rudder kick decreased and the ultimate right rolling velocity increased to the point where the undesired initial motion was not noticeable to the pilot.\n\n**Characteristics in lateral oscillations.**—As was the case for the rudder kicks, the servo-applied aileron deflection lags the sideslip angle a small amount in the lateral-oscillation time histories shown in figures 9 (b) and 9 (c). There are occasional small irregularities in the servo-output motion, but the agreement between actual and ideal aileron deflection is considered good. The time histories show that, qualitatively, the apparatus simulates the effects of changes of $C_{l_\\beta}$ on lateral-oscillation characteristics. Compared to the normal characteristics (fig. 9 (a)), the increased excitation of the rolling motion and the decreased damping of the airplane motions caused by an increase in dihedral effect (fig. 9 (b)) are apparent. With a slight negative dihedral effect (fig. 9 (c)), the rolling motion is small, and the airplane motions are rapidly damped.\n\nFair quantitative agreement between measured and predicted effects of changes in effective dihedral on oscillation period and damping are shown in figure 10. The discrepancies at the $-2.7^\\circ$ static effective-dihedral-angle setting may be due in part to difficulties in accurately evaluating the oscillation flight data when, as in this case, the damping is high, and due in part to the slight rudder motions which the pilot was unable to eliminate. Both the measured and predicted effects of $\\Gamma_v$ on the period are small.\n\nThe measured decrease in damping with increasing $\\Gamma_v$ is approximately the same as that given by the predictions, which shows that deviations of the aileron servo characteristics from the ideal did not have a serious effect on the airplane damping.\n\n## II. DETERMINATION OF TOLERABLE RANGE OF EFFECTIVE DIHEDRAL\n\nAfter the evaluation flights reported in part I the first application of the variable dihedral apparatus was a flight determination of the tolerable range of effective dihedral on the test airplane. As indicated previously, the servo-gearing ratios used for these tests were approximately twice the values used during the tests reported in part I, and the aileron tab was increased in size to give a more favorable relation between stick-fixed and stick-free dihedral effects.\n\n### PROCEDURE\n\nQuantitative data were gathered during steady straight sideslips, rudder-fixed aileron rolls, and lateral oscillations. The lateral oscillations were excited in the same manner as previously discussed in part I.\n\nA survey of pilots' opinions was made among five pilots in a series of flights separate from those during which quantitative measurements were made. Four were NACA test pilots and one was a service pilot; all were highly experienced with fighter-type aircraft. The pilots were requested to report their opinions (in the form of answers to specific questions) with regard to the damping and period of the oscillations, the response to gusts in rough air, their ability to coordinate during turn entries and exits, and the general flying qualities.\n\nThe flight conditions chosen for the investigation were as follows:\n\n**Landing-approach condition.**—In this condition the indicated airspeed was 90 knots, the flaps were extended, and the landing gear was retracted. The engine power used was that necessary for level flight. Ninety knots was about the lowest speed at which the servo-applied aileron angle caused by the wings-level sideslip angle was sufficiently small to allow reasonable maneuvers without exceeding the limits of the apparatus. The gear was retracted in order to keep the drag, the propeller loading, and, hence, the wings-level sideslip angle to a minimum.\n\n**Cruising condition.**—The indicated airspeed was 180 knots for this condition; flaps and gear were up. The engine power used was that necessary for level flight. This speed was not so high as to require diving or high engine power for level flight (as was the case at the 390-knot test speed of part I), but it was sufficiently high that further increases in speed would mean only small changes in lift and thrust coefficients.\n```", "timestamp": "2026-07-22T04:40:36.014298+00:00"}
{"citation_id": "19930086081", "source_url": "https://ntrs.nasa.gov/api/citations/19930086081/downloads/19930086081.pdf", "page_number": 37, "total_pages": 44, "image_filename": "19930086081_p37.jpg", "text": "```markdown\nNACA RM L9H05\n35\n\nCONFIDENTIAL\n\n<!-- Image (143, 109, 903, 813) -->\n\nCONFIDENTIAL\n\n(b) Chord force plotted against $\\alpha$.\n\nFigure 16.- Continued.\n```", "timestamp": "2026-07-22T04:40:38.578722+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 6, "total_pages": 99, "image_filename": "19930082511_p6.jpg", "text": "4\nNACA TN No. 1826\n\nvelocity component u parallel to the lines SP and QR is zero; hence,\nthe perturbation potentials at points P and Q are the same as at points S\nand R, respectively. If points S and R are far upstream of the body\ntheir potentials will be equal, so that the potentials at points P and Q\nare equal. The perturbation potential is thus uniform over the entire\nsurface, and only perturbation velocities normal to the surface can exist\nat the surface.\n\nModification of basic concepts.- If the closed entrance region is\nnear the body, as shown in figure I(b), the preceding discussion and\nconclusions no longer apply. Thus, although u must still be constant\nover the entire free surface, it is no longer necessarily zero; that is,\nthe total velocity on the free surface is not necessarily equal to the\nvelocity far upstream in the closed portion of the tunnel. The two\nvelocities will, in fact, generally be unequal except in special cases\nwhere equality results from geometrical symmetry of the arrangement.\n(For example, if a horseshoe vortex is located in the horizontal plane of\nsymmetry of the tunnel, the values of u at the top and bottom of the\ntunnel would be expected to be equal and opposite; but since u must be\nuniform over the surface, it follows that u = 0.) Furthermore, the\nvelocities in the jet surface normal to u (that is, the circumferential\nvelocities) are, in general, no longer zero (except for axially symmetrical\nflows, such as that produced by a source on the axis of a circular\ntunnel) so that two surface points at the same longitudinal position,\nas P' and Q' (fig. 1(b)) do not necessarily have the same values of the\nperturbation potential.\n\nEntrance-lip condition.- Consider, for simplicity, the symmetrical\ncase of figure I(b), in which the lifting element is on the horizontal\nplane of symmetry of the tunnel. Since u is zero on the free boundary,\nthe perturbation potential is constant along the elements AB, CD,\nEF, . . ., although, as just indicated, it is not necessarily the same\nfor all these elements. This one boundary condition for the open section -\nthat the potential be constant along each of these elements - does not\nsuffice, however, to define the problem uniquely. In fact, as will be\nobvious from the subsequent discussion of electrical analogies, the\npotentials of these elements may be quite arbitrarily assigned without\nviolating this condition or the boundary condition on the closed portion\nof the tunnel (that the normal derivative of the potential be zero at the\nwall). In order to avoid this lack of uniqueness, further conditions\nmust be sought. The most important of these is that the velocity be continuous\n(in particular, not infinite) at the entrance lip (points A, C,\nE, . . .). This condition takes cognizance of the fact that, because of\nviscosity, the physical flow leaves the lip smoothly, just as it leaves\nthe trailing edge of an airfoil; the condition is, in fact, strictly\nanalogous to the Kutta-Joukowski condition for the trailing-edge of an\nairfoil, which similarly takes into account the basic viscosity effect\nand provides uniqueness where otherwise an infinity of solutions would exist.\nIt is recognized that, just as the Kutta-Joukowski condition does not\nalways suffice to predict airfoil lift very accurately, the corresponding\ncondition for the open tunnel may similarly oversimplify the entrance-lip", "timestamp": "2026-07-22T04:40:39.557676+00:00"}
{"citation_id": "19930082487", "source_url": "https://ntrs.nasa.gov/api/citations/19930082487/downloads/19930082487.pdf", "page_number": 10, "total_pages": 33, "image_filename": "19930082487_p10.jpg", "text": "8\nNACA TN No. 1813\n\nspeeds have been pointed out. These characteristics will be considered further in this section of the report. Attention has already been directed to the aft boundary of the supersonic region. In figure 8, the variation of the chordwise extent of the region of supersonic flow over the NACA 23015 airfoil section is shown as a function of Mach number. It is seen that the fore and aft boundaries shift until the drag-divergence Mach number is reached. The forward sonic point remains relatively fixed for some further increase in free-stream Mach number while the aft sonic point continues to move rearward of the airfoil crest (x/c)β. At Mach numbers above that for drag divergence the rear sonic point lies close to the position of the shock wave, as determined from the schlieren photographs. Reference 3 shows that the chordwise location of the shock wave can be altered by scale effects. The question therefore arises as to whether these data, obtained at a test Reynolds number of about 2 million, apply at larger Reynolds numbers. Unpublished German pressure-distribution data, for the NACA 23015 airfoil section at 0° incidence and a Reynolds number of about 5 million, are shown for comparison in figure 8. The chordwise extent of the supersonic region is nearly the same for both sets of experimental data. However, as can be seen from figure 4, there is some small difference between the variation of local pressure coefficients with Mach number at 2 million and 5 million Reynolds number.\n\nThe appearance of a fixed chordwise location of the forward sonic point at subsonic free-stream Mach numbers above that for drag-divergence is to be expected since, as has been noted for the NACA 23015 airfoil section, the pressure-coefficient variation at these Mach numbers appears to be such as to maintain an approximately constant local Mach number distribution over that portion of the airfoil surface ahead of the crest. The location of the forward sonic point for each of the airfoils of reference 1 is shown as a function of Mach number in figure 9. The farthest forward station at which pressures were measured for that report was the 2.5-percent-chord station so that the angles of attack which could be considered were limited to those for which the sonic point was near or behind this station. It appears (fig. 9) that the forward sonic point is fixed at Mach numbers above the drag-divergence Mach number only for those cases for which the sonic point lies between the 2.5- and 15-percent-chord stations. When the sonic point is outside of this region at the drag-divergence Mach number there is some movement of the sonic point with further increase in Mach number. This movement appears to be greatest for the two low-drag airfoil sections which have, at a fixed Mach number, large variation of sonic point location with angle of attack.", "timestamp": "2026-07-22T04:40:44.014857+00:00"}
{"citation_id": "19930085842", "source_url": "https://ntrs.nasa.gov/api/citations/19930085842/downloads/19930085842.pdf", "page_number": 85, "total_pages": 104, "image_filename": "19930085842_p85.jpg", "text": "NACA RM L9C29\n81\n\n[Figure: Graph with multiple curves and axes]\n\nPropeller advance-diameter ratio, $V/nD$\nLift coefficient, $C_L$\nTorque coefficient, $Q_c$\n$\\beta$, deg\n11.5\nResultant drag coefficient, $C_R$\n\nNATIONAL ADVISORY\nCOMMITTEE FOR AERONAUTICS\n\n(h) $\\alpha_{11} = 42^\\circ$.\nFigure 43.- Continued.", "timestamp": "2026-07-22T04:40:44.702035+00:00"}
{"citation_id": "19930085930", "source_url": "https://ntrs.nasa.gov/api/citations/19930085930/downloads/19930085930.pdf", "page_number": 90, "total_pages": 92, "image_filename": "19930085930_p90.jpg", "text": "CONFIDENTIAL\n\nNACA RM L9G07\n\n[Figure: A shadowgraph of the flow at an angle of attack of 35°.]\n\nFigure 42.- A shadowgraph of the flow at an angle of attack of $35^\\circ$.\n\nCONFIDENTIAL\n\nNACA\n\n89", "timestamp": "2026-07-22T04:40:48.150663+00:00"}
{"citation_id": "19930082476", "source_url": "https://ntrs.nasa.gov/api/citations/19930082476/downloads/19930082476.pdf", "page_number": 14, "total_pages": 41, "image_filename": "19930082476_p14.jpg", "text": "```markdown\n12\nNACA TN No. 1801\n\ndeflection of this amount is probably impractical, it would appear desirable\nto increase uniformly the size of the vertical tails so that a smaller\nrudder deflection would be required.\n\nLangley Aeronautical Laboratory\nNational Advisory Committee for Aeronautics\nLangley Field, Va., November 17, 1948\n\nREFERENCES\n\n1. Neihouse, A. I.: Tail-Design Requirements for Satisfactory Spin\n   Recovery for Personal-Owner-Type Light Airplanes. NACA TN No. 1329,\n   1947.\n\n2. Zimmerman, C. H.: Preliminary Tests in the N.A.C.A. Free-Spinning Wind\n   Tunnel. NACA Rep. No. 557, 1936.\n\n3. Seidman, Oscar, and Neihouse, A. I.: Comparison of Free-Spinning Wind-\n   Tunnel Results with Corresponding Full-Scale Spin Results. NACA MR,\n   Dec. 7, 1938.\n\n4. Neihouse, A. I.: A Mass-Distribution Criterion for Predicting the Effect\n   of Control Manipulation on the Recovery from a Spin. NACA ARR, Aug. 1942.\n\n5. Seidman, Oscar, and Neihouse, A. I.: Free-Spinning Wind-Tunnel Tests of\n   a Low-Wing Monoplane with Systematic Changes in Wings and Tails.\n   V. Effect of Airplane Relative Density. NACA Rep. No. 691, 1940.\n\n6. Gilruth, R. R., and Turner, W. N.: Lateral Control Required for Satis-\n   factory Flying Qualities Based on Flight Tests of Numerous Airplanes.\n   NACA Rep. No. 715, 1941.\n\n7. Weick, Fred E., and Jones, Robert T.: Résumé and Analysis of N.A.C.A.\n   Lateral Control Research. NACA Rep. No. 605, 1937.\n\n8. Pearson, Henry A., and Jones, Robert T.: Theoretical Stability and\n   Control Characteristics of Wings with Various Amounts of Taper and\n   Twist. NACA Rep. No. 635, 1938.\n```", "timestamp": "2026-07-22T04:40:49.252017+00:00"}
{"citation_id": "19930086151", "source_url": "https://ntrs.nasa.gov/api/citations/19930086151/downloads/19930086151.pdf", "page_number": 31, "total_pages": 34, "image_filename": "19930086151_p31.jpg", "text": "NACA RM L9J28\n29\n\nCONFIDENTIAL\n\nParallelogram aileron\nTriangular aileron\nParallelogram aileron\nTriangular aileron\nPlain wing\nWing with\nend plate\n\n$C_{l_p}$\n-.3\n-.2\n-.1\n0\n0\n4\n8\n12\n16\n20\nAngle of attack, $\\alpha$, deg\n\nCONFIDENTIAL\nNACA\n\nFigure 13.— Variation of $C_{l_p}$ (used for determining $\\frac{pb}{2V}$) with angle of attack.", "timestamp": "2026-07-22T04:40:51.244824+00:00"}
{"citation_id": "19930082485", "source_url": "https://ntrs.nasa.gov/api/citations/19930082485/downloads/19930082485.pdf", "page_number": 12, "total_pages": 62, "image_filename": "19930082485_p12.jpg", "text": "NACA TN No. 1810\n\n$\\frac{r}{V_{cr}^2} \\frac{dV_r}{dt}$ is nearly as large as $\\frac{r}{P_t} \\frac{dP_s}{dr}$ near the blade tip, the effect on $\\frac{P_s}{P_t} \\left( \\frac{V_u}{V_{cr}} \\right)^2$ was small (about 4 percent) and the deviation from free-vortex flow was small.\n\nSUMMARY OF RESULTS\n\nFrom a comparison of the calculated design performance and the experimental performance of a gas-turbine stator blade, the following results were obtained:\n\n1. The stream-filament theory is adequate to determine gas velocities on the surface of high-solidity blades. For the blades investigated, the calculated surface velocities satisfactorily agreed with the experimental values.\n\n2. The experimental value of tangential component of critical velocity ratio was 3 percent greater than the design value at the blade tip and 5 percent less than the design value at the blade root. These deviations were caused by lack of radial equilibrium of pressures at 0.1 chord downstream of the turbine blades and the thick boundary layer at the inner shroud.\n\n3. Constant axial velocity at all radii was not achieved. The experimental axial velocity was 6 percent higher than the design axial velocity near the blade tip.\n\n4. Although the radial-acceleration forces were large, the deviation from free-vortex flow was small.\n\nLewis Flight Propulsion Laboratory,  \nNational Advisory Committee for Aeronautics,  \nCleveland, Ohio, September 13, 1948.", "timestamp": "2026-07-22T04:40:54.185140+00:00"}
{"citation_id": "19930085965", "source_url": "https://ntrs.nasa.gov/api/citations/19930085965/downloads/19930085965.pdf", "page_number": 52, "total_pages": 67, "image_filename": "19930085965_p52.jpg", "text": "NACA RM E9E06\n\n[Figure: Graph showing variation of $\\alpha_2$ (deg) with $H_{\\text{max } 2}/H_{\\text{max } 1}$, with curve rising from origin to approximately $\\alpha_2 = 8$ deg at $H_{\\text{max } 2}/H_{\\text{max } 1} = 200$. Axes labeled and grid lines present. NACA logo in lower right corner of plot area.]\n\nFigure 8. - Variation of $\\alpha_2$ with $H_{\\text{max } 2}/H_{\\text{max } 1}$.\n\n51", "timestamp": "2026-07-22T04:40:55.901233+00:00"}
{"citation_id": "19930085838", "source_url": "https://ntrs.nasa.gov/api/citations/19930085838/downloads/19930085838.pdf", "page_number": 98, "total_pages": 118, "image_filename": "19930085838_p98.jpg", "text": "96\nNACA RM No. L9B23\n\n<!-- Image (118, 116, 848, 878) -->\n\n(d) $\\delta_f = 25^\\circ$.\nFigure 12.- Continued.", "timestamp": "2026-07-22T04:40:56.509649+00:00"}
{"citation_id": "19930082245", "source_url": "https://ntrs.nasa.gov/api/citations/19930082245/downloads/19930082245.pdf", "page_number": 18, "total_pages": 66, "image_filename": "19930082245_p18.jpg", "text": "NACA TN No. 1596\n17\n\nTABLE I\n\nAIRFOIL ORDINATES FOR NACA 66,1-115 AIRFOIL\n[Station and ordinates in percent of wing chord]\n\n| Upper surface | | Lower surface | |\n| :--- | :--- | :--- | :--- |\n| Station | Ordinates | Station | Ordinates |\n| 1.188 | 1.851 | 1.312 | -1.744 |\n| 2.429 | 2.532 | 2.571 | -2.346 |\n| 4.922 | 3.501 | 5.078 | -3.185 |\n| 7.419 | 4.239 | 7.581 | -3.815 |\n| 9.920 | 4.843 | 10.080 | -4.325 |\n| 14.925 | 5.803 | 15.075 | -5.131 |\n| 19.932 | 6.535 | 20.068 | -5.739 |\n| 24.942 | 7.095 | 25.058 | -6.200 |\n| 29.953 | 7.505 | 30.047 | -6.533 |\n| 39.976 | 7.984 | 40.024 | -6.912 |\n| 44.988 | 8.049 | 45.012 | -6.951 |\n| 50.000 | 7.988 | 50.000 | -6.884 |\n| 60.022 | 7.434 | 59.978 | -6.362 |\n| 70.038 | 6.058 | 69.962 | -5.086 |\n| 80.040 | 4.029 | 79.960 | -3.233 |\n| 90.026 | 1.763 | 89.974 | -1.245 |\n| 95.014 | .740 | 94.986 | -.424 |\n| 100.000 | 0 | 100.000 | 0 |\n\nSlope of radius through L.E.: 0.062\nL.E. radius: 0.0161c\n\nNACA", "timestamp": "2026-07-22T04:40:58.087000+00:00"}
{"citation_id": "19930090382", "source_url": "https://ntrs.nasa.gov/api/citations/19930090382/downloads/19930090382.pdf", "page_number": 29, "total_pages": 37, "image_filename": "19930090382_p29.jpg", "text": "NACA RM L9I07\n31\n\nCONFIDENTIAL\n\nPower coefficient, $C_P$\nThrust coefficient, $C_T$\n\nTip Mach number, $M_t$\nEfficiency, $\\eta$\n\nAdvance ratio, J\n(j) M=0.80 Concluded.\nFigure 5 - Continued.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T04:40:58.991922+00:00"}
{"citation_id": "19930093773", "source_url": "https://ntrs.nasa.gov/api/citations/19930093773/downloads/19930093773.pdf", "page_number": 17, "total_pages": 47, "image_filename": "19930093773_p17.jpg", "text": "```markdown\nTABLE I - ENGINE PERFORMANCE DATA\n\n| Run | Altitude (ft) | Ram-pressure ratio $p_2/p_0$ | Flight Mach number $M$ | Ram-inlet pressure $p_2$ (lb/sq in. abs) | Ram-inlet and outlet temperature, $T_2$ ($^\\circ$R) | Engine speed, N (rpm) | Compressor-outlet temperature $T_3$ ($^\\circ$R) | Compressor-outlet pressure $p_3$ (lb/sq in. abs) | Jet thrust, $F_n$ (lb) | Jet thrust, $F_n$ (lb/sec) | Engine-inlet air flow $W_a$ (lb/sec) | Specific fuel consumption $c$ (lb/(hr)(lb thrust)) | Fuel-air ratio $f/a$ | Compressor-outlet temperature, $T_3$ ($^\\circ$R) | Turbine-outlet total pressure, $p_4$ (lb/sq in. abs) | Corrected engine speed $N/\\sqrt{\\theta_2}$ (rpm) | Corrected net thrust $F_n/\\delta_2$ (lb) | Corrected engine-inlet air flow $W_a\\sqrt{\\theta_2}/\\delta_2$ (lb/sec) | Corrected fuel consumption $W_f/\\delta_2\\sqrt{\\theta_2}$ (lb/hr) | Corrected specific fuel consumption $c$ (lb/(hr)(lb thrust)) | Corrected fuel-air ratio $f/a$ | Corrected compressor-outlet temperature, $T_3/\\theta_2$ ($^\\circ$R) | Corrected turbine-outlet temperature, $T_4/\\theta_2$ ($^\\circ$R) | Engine total-pressure ratio $p_4/p_2$ | Engine total-temperature ratio $T_4/T_2$ |\n|---|---|---|---|---|---|---|---|---|---|---|---|---|---|---|---|---|---|---|---|---|---|---|---|---|---|\n| 1 | 5,000 | 1.038 | .230 | 1740 | 507 | 7895 | 509 | 4890 | 44.37 | 81.88 | 53.00 | 1.251 | 0.0182 | 1740 | 3445 | 7690 | 5160 | 97.42 | 6523 | 1.265 | 0.0186 | 1760 | 1,865 | 3.396 |\n| 2 | 5,000 | 1.037 | .225 | 1736 | 509 | 7692 | 511 | 4593 | 39.97 | 81.07 | 4800 | 1.213 | .0164 | 1631 | 3335 | 7769 | 4770 | 96.72 | 5842 | 1.225 | .0167 | 1665 | 1,790 | 3.173 |\n| 3 | 5,000 | 1.039 | .230 | 1740 | 506 | 7500 | 510 | 4305 | 36.60 | 80.28 | 4590 | 1.196 | .0152 | 1542 | 3247 | 7598 | 4455 | 96.37 | 5406 | 1.213 | .0156 | 1585 | 1,745 | 3.012 |\n| 4 | 5,000 | 1.036 | .225 | 1736 | 508 | 7300 | 511 | 3650 | 30.53 | 76.83 | 3960 | 1.162 | .0128 | 1396 | 3162 | 7070 | 3720 | 95.46 | 4559 | 1.225 | .0130 | 1425 | 1,605 | 2.798 |\n| 5 | 5,000 | 1.034 | .215 | 1742 | 507 | 6459 | 511 | 2846 | 23.18 | 70.23 | 2710 | 1.168 | .0107 | 1268 | 2979 | 6537 | 2818 | 94.32 | 3333 | 1.183 | .0109 | 1297 | 1,454 | 2.477 |\n| 6 | 5,000 | 1.033 | .210 | 1740 | 506 | 5944 | 510 | 2080 | 16.15 | 63.46 | 2060 | 1.275 | .0090 | 1170 | 2832 | 6021 | 1965 | 94.23 | 2532 | 1.292 | .0092 | 1200 | 1,314 | 2.290 |\n| 7 | 5,000 | 1.033 | .210 | 1740 | 505 | 5024 | 509 | 1172 | 6.17 | 49.05 | 1350 | 1.654 | .0078 | 1096 | 2683 | 5094 | 994 | 97.62 | 1665 | 1.675 | .0079 | 1127 | 1,142 | 2.149 |\n| 8 | 5,000 | 1.034 | .215 | 1740 | 504 | 4091 | 509 | 669 | 4.13 | 34.21 | 1020 | 2.542 | .0065 | 1125 | 1924 | 4152 | 499 | 40.76 | 1259 | 2.565 | .0067 | 1160 | 1,059 | 2.210 |\n| 9 | 5,000 | 1.032 | .210 | 1741 | 504 | 3147 | 509 | 312 | 1.40 | 23.73 | 520 | 5.837 | .0036 | 1127 | 1832 | 3194 | 169 | 29.96 | 1009 | 5.940 | .0038 | 1028 | 1,017 | 2.233 |\n| 10 | 5,000 | 1.032 | .210 | 1738 | 504 | 2046 | 509 | 128 | 10 | 16.42 | 474 | --- | .0030 | 1134 | 1769 | 2077 | 12 | 19.69 | 655 | 48.1 | .0032 | 1168 | 987 | 2.228 |\n| 11 | 15,000 | 1.034 | .215 | 1188 | 478 | 6993 | 460 | 2733 | 23.23 | 56.41 | 2550 | 1.097 | .0126 | 1386 | 2102 | 7297 | 4137 | 96.42 | 4733 | 1.142 | .0137 | 1506 | 1,464 | 2.870 |\n| 12 | 15,000 | 1.030 | .205 | 1188 | 479 | 6459 | 461 | 2219 | 18.63 | 51.80 | 2020 | 1.065 | .0108 | 1254 | 1896 | 6724 | 3350 | 96.62 | 3745 | 1.130 | .0117 | 1358 | 1,507 | 2.996 |\n| 13 | 15,000 | 1.030 | .205 | 1188 | 475 | 5944 | 478 | 1694 | 13.91 | 46.67 | 1550 | 1.127 | .0092 | 1145 | 1707 | 6211 | 2450 | 79.54 | 2885 | 1.178 | .0100 | 1255 | 1,367 | 2.590 |\n| 14 | 15,000 | 1.030 | .205 | 1186 | 471 | 5024 | 475 | 1000 | 7.55 | 36.00 | 870 | 1.277 | .0076 | 1034 | 1454 | 5276 | 1344 | 61.17 | 1836 | 1.361 | .0083 | 1138 | 1,180 | 2.177 |\n| 15 | 15,000 | 1.028 | .195 | 1186 | 471 | 4091 | 475 | 538 | 2.76 | 24.58 | 768 | 2.043 | .0057 | 1060 | 1320 | 4296 | 670 | 41.76 | 1438 | 2.145 | .0095 | 1167 | 1,076 | 2.232 |\n| 16 | 15,000 | 1.031 | .205 | 1190 | 470 | 3147 | 474 | 304 | 1.69 | 19.45 | 605 | 2.580 | .0047 | 1118 | 1261 | 3307 | 300 | 32.85 | 1151 | 3.768 | .0095 | 1205 | 1,004 | 2.361 |\n| 17 | 15,000 | 1.028 | .195 | 1188 | 468 | 2046 | 472 | 157 | 1.00 | 8.56 | 371 | 3.675 | .0030 | 1105 | 1213 | 2154 | 179 | 14.61 | 696 | 3.870 | .0033 | 1225 | 992 | 2.341 |\n| 18 | 15,000 | 1.034 | .220 | 1196 | 474 | 7895 | 498 | 4016 | 33.36 | 65.37 | 4150 | 1.407 | .0187 | 1754 | --- | --- | --- | --- | --- | --- | --- | --- | 1,820 | 3.508 |\n| 19 | 15,000 | 1.210 | .530 | 1186 | 477 | 7692 | 502 | 3912 | 27.76 | 64.50 | 3730 | 1.344 | .0161 | 1614 | 2825 | --- | --- | --- | --- | --- | --- | --- | 1,781 | 3.202 |\n| 20 | 15,000 | 1.211 | .530 | 1187 | 474 | 7500 | 501 | 3698 | 25.63 | 64.09 | 3595 | 1.324 | .0147 | 1544 | 2549 | --- | --- | --- | --- | --- | --- | --- | 1,719 | 3.045 |\n| 21 | 15,000 | 1.205 | .525 | 1190 | 480 | 6993 | 504 | 3210 | 21.38 | 61.72 | 2720 | 1.275 | .0125 | 1365 | 2304 | --- | --- | --- | --- | --- | --- | --- | 1,577 | 2.692 |\n| 22 | 15,000 | 1.203 | .520 | 1188 | 482 | 6459 | 506 | 2430 | 14.41 | 56.77 | 1990 | 1.360 | .0097 | 1200 | 2029 | --- | --- | --- | --- | --- | --- | --- | 1,385 | 2.362 |\n| 23 | 15,000 | 1.205 | .525 | 1190 | 479 | 5944 | 505 | 1753 | 9.69 | 50.82 | 1350 | 1.500 | .0075 | 1073 | 1746 | --- | --- | --- | --- | --- | --- | --- | 1,215 | 2.081 |\n| 24 | 15,000 | 1.204 | .525 | 1185 | 479 | 5024 | 504 | 904 | 2.27 | 38.87 | 770 | 3.392 | .0055 | 914 | 1458 | --- | --- | --- | --- | --- | --- | --- | 1,015 | 1.810 |\n| 25 | 15,000 | 1.202 | .520 | 1180 | 478 | 4091 | 506 | 418 | -68 | 27.95 | 349 | --- | .0055 | 883 | 1240 | --- | --- | --- | --- | --- | --- | --- | 959 | 1.736 |\n| 26 | 15,000 | 1.203 | .520 | 1190 | 47", "timestamp": "2026-07-22T04:41:00.974119+00:00"}
{"citation_id": "19930086081", "source_url": "https://ntrs.nasa.gov/api/citations/19930086081/downloads/19930086081.pdf", "page_number": 38, "total_pages": 44, "image_filename": "19930086081_p38.jpg", "text": "36\nNACA RM L9H05\n\nCONFIDENTIAL\n\n<!-- Image (129, 110, 874, 874) -->\n\n(c) Hinge moment plotted against $\\alpha$.\nFigure 16.- Continued.", "timestamp": "2026-07-22T04:41:02.800730+00:00"}
{"citation_id": "19930086015", "source_url": "https://ntrs.nasa.gov/api/citations/19930086015/downloads/19930086015.pdf", "page_number": 47, "total_pages": 54, "image_filename": "19930086015_p47.jpg", "text": "46\n\nCONFIDENTIAL\n\n.60\n.50\n.40\n.30\n.20\n.10\n0\n-.10\n\nLift coefficient, $C_L$\n\nDrag coefficient, $C_D$\n\nAngle of attack, $\\alpha$, deg\n\nPitching-moment coefficient, $C_{m_{\\bar{c}/4}}$\n\n(b) $D=113.43$; $M=1.53$.\n\nFigure 13.—Continued.\n\nNACA RM A50E24\n\nCONFIDENTIAL\n\n46\n\n[Figure: Graph showing three plots of Lift coefficient ($C_L$) versus Drag coefficient ($C_D$), Angle of attack ($\\alpha$), and Pitching-moment coefficient ($C_{m_{\\bar{c}/4}}$). The legend indicates three configurations: vertical (circle), upright (diamond), and inverted (square). The NACA logo is present in the lower right corner of the graph area.]", "timestamp": "2026-07-22T04:41:03.714433+00:00"}
{"citation_id": "19930092013", "source_url": "https://ntrs.nasa.gov/api/citations/19930092013/downloads/19930092013.pdf", "page_number": 14, "total_pages": 21, "image_filename": "19930092013_p14.jpg", "text": "```markdown\n10\nREPORT 948—NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\n<!-- Image (87, 65, 918, 874) -->\n\n(a) $\\Gamma_n$ 28.4° ($C_{l\\beta}$ -0.0054).\n(b) $\\Gamma_n$ 22.7° ($C_{l\\beta}$ -0.0031).\n(c) $\\Gamma_n$ 14.2° ($C_{l\\beta}$ -0.0032).\n(d) $\\Gamma_n$ 3.3° ($C_{l\\beta}$ -0.0012).\n(e) $\\Gamma_n$ -3.1° ($C_{l\\beta}$ 0.0007).\n(f) $\\Gamma_n$ -10.7° ($C_{l\\beta}$ 0.0020).\n(g) $\\Gamma_n$ -18.2° ($C_{l\\beta}$ 0.0041).\n\nFIGURE 11.—Lateral stability and control characteristics during steady straight sideslips. Landing-approach condition.\n\n(a) $\\Gamma_n$ 24.4° ($C_{l\\beta}$ -0.0055).\n(b) $\\Gamma_n$ 18.2° ($C_{l\\beta}$ -0.0041).\n(c) $\\Gamma_n$ 12.9° ($C_{l\\beta}$ -0.0029).\n(d) $\\Gamma_n$ 6.2° ($C_{l\\beta}$ -0.0014).\n(e) $\\Gamma_n$ 0° ($C_{l\\beta}$ 0).\n(f) $\\Gamma_n$ -7.1° ($C_{l\\beta}$ 0.0016).\n(g) $\\Gamma_n$ -12.4° ($C_{l\\beta}$ 0.0028).\n\nFIGURE 12.—Lateral stability and control characteristics during steady straight sideslips. Cruising condition.\n```", "timestamp": "2026-07-22T04:41:08.188508+00:00"}
{"citation_id": "19930082496", "source_url": "https://ntrs.nasa.gov/api/citations/19930082496/downloads/19930082496.pdf", "page_number": 5, "total_pages": 50, "image_filename": "19930082496_p5.jpg", "text": "```markdown\n4\nNACA TN No. 1836\n\n(c) Quasi-service performance characteristics of turbine\nblades operated at shaft speeds between 10,000 and\n17,500 rpm (tip speeds of 478 and 835 ft/sec) and\nindicated inlet-gas temperatures between 1700° and\n2200° F\n\nChanges in the structure of the material resulting from operation as\na turbine blade were investigated by X-ray diffraction.\n\nAll ceramal bodies used in this investigation were fabricated by\nKennametal, Inc., who collaborated with the NACA by making available\ntheir extensive experience in the fabrication of ceramals.\n\nAPPARATUS AND PROCEDURE\n\nTensile-Strength Evaluation\n\nThe elevated-temperature, short-time, tensile-strength-evaluation\napparatus consisted essentially of a commercial tensile machine fitted\nwith a helium-atmosphere furnace (fig. 1). The 16,000-pound range of\nthe tensile machine was used. Minimum dial graduation was 20 pounds.\nThe machine was equipped with a hydraulic loading system that per-\nmitted constant rate of application of load. A platinum-wound elec-\ntric furnace was used. The furnace was installed on the machine in\nsuch a manner that it could be conveniently manipulated and located\nabout the specimen. A thermocouple attached to the specimen, by use of\na wire wrapped around the specimen and the thermocouple, indicated\nspecimen temperature. A second thermocouple near the furnace windings\nand connected to a temperature-control system so controlled the fur-\nnace as to maintain the specimen at the evaluation temperature. Platinum -\nplatinum-13-percent-rhodium thermocouples were used. The helium\natmosphere was used to minimize any oxidation effects upon the tensile-\nstrength evaluations. Adapter rods and grips of high-temperature alloys\nwere used to transmit load to the specimen. Ball-and-socket joints\nincorporated into the adapter rods were used to secure alinement and\nto obtain collinearity of specimen and load (fig. 1).\n\nThe specimens were inspected for internal and external flaws by\nradiographic and fluorescent-oil methods, respectively. Radiographic\ninspection revealed fine chemical segregation in these specimens.\nThe segregation was not considered serious.\n\nThe ends of the specimens were wrapped with a single layer of\nwoven asbestos sheet (fig. 2), which was secured with an adhesive.\nThis wrapping was used to compensate for any distortion of the grips\nand any minute surface irregularities of the specimens. The adapter\n```", "timestamp": "2026-07-22T04:41:10.319729+00:00"}
{"citation_id": "19930082498", "source_url": "https://ntrs.nasa.gov/api/citations/19930082498/downloads/19930082498.pdf", "page_number": 4, "total_pages": 49, "image_filename": "19930082498_p4.jpg", "text": "NACA TN No. 1838\n\nA photograph of the dynamometer setup is presented as figure 5. On the right-hand side of the photograph may be seen the engine and cowling which were taken as a unit from a military liaison airplane. Inasmuch as the propeller was removed for this investigation, the power from the engine was absorbed by means of an electric induction motor run as a generator with the power being fed into a variable-frequency alternator, which was utilized to absorb the output of the induction motor, feeding it into the supply line, and to supply exciting current to the motor. The motor is rated at 266 horsepower at 3500 rpm and has power-speed characteristics similar to those of the engine at full-throttle operation. Cooling air for the cylinders and for the oil cooler was supplied by means of a blower installed within a duct that guided the flow to the engine and oil-cooler cooling-air inlets. A frequency analysis of the sound of the blower showed that the sound-pressure levels were sufficiently low and the frequencies sufficiently high to cause no interference with the engine sound measurements.\n\nStandard instruments from the airplane were used to check engine operation except that engine speeds were determined with a combination of magnetic-drag aircraft tachometer generator and indicator. Thermocouples were installed in the spark-plug gaskets of the spark plugs nearest the exhaust ports to insure that the engine was not overheated during the tests. Engine back pressures were determined with a micromanometer connected to a static-pressure tap which was installed in the exhaust pipe from the left rear cylinder about 8 inches from the exhaust port.\n\nA General Radio Company type 759-A sound-level meter was used to measure the over-all sound-pressure levels and a General Radio Company type 760-A sound analyzer to determine the noise spectrums. All noise levels were measured in decibels of sound intensity referred to the Acoustical Society of America standard base pressure level of $0.000204$ dyne per square centimeter. The measurements of the over-all sound-pressure levels appear to be reliable and could be repeated. A few of the low analyzer readings, however, were lower than would be expected.\n\nNoise-level measurements and spectrum analyses, with readings taken at multiples of one-half the firing frequency, were made at a point 50 feet from the ends of the original exhaust stacks on a line $135^\\circ$ to the right and rear of a line running forward in the plane of symmetry of the engine. The sound-level meter and the analyzer were placed on a board which rested directly on the ground. No corrections for ground reflections have been applied to the data presented in this paper. Over-all noise levels and spectrum analyses were measured at one, two, or all of the following speeds: 1650, 2000, and 2790 rpm. Engine back pressure was measured for many of the configurations.", "timestamp": "2026-07-22T04:41:10.865646+00:00"}
{"citation_id": "19930085842", "source_url": "https://ntrs.nasa.gov/api/citations/19930085842/downloads/19930085842.pdf", "page_number": 86, "total_pages": 104, "image_filename": "19930085842_p86.jpg", "text": "82\nNACA RM L9C29\n\nPropeller advance-diameter ratio, $V/nD$\nLift coefficient, $C_L$\nTorque coefficient, $Q_c$\nResultant-drag coefficient, $C_{DR}$\n\n$V/nD$\n$C_{DR}$\n$C_L$\n$B, deg$\n$\\square$ 11.5\n\n1200 bhp at 1085 rpm\n1380 bhp at 1085 rpm\n\nNATIONAL ADVISORY\nCOMMITTEE FOR AERONAUTICS\n\n(i) $\\alpha_{11} = 48^\\circ$.\nFigure 43.- Continued.", "timestamp": "2026-07-22T04:41:12.138674+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 7, "total_pages": 99, "image_filename": "19930082511_p7.jpg", "text": "NACA TN No. 1826\n\nflow; however, as with the airfoil, the condition is probably adequate where the flow is not subject to an excessive pressure rise on approaching the lip. References 4, 5, and 6 used the condition, and reference 6, in addition, discussed it from the physical viewpoint and compared it with the airfoil trailing-edge condition.\n\nConcerning the downstream end of the open section, the exit lip may be considered to correspond to the leading edge of an airfoil and no effort need be made in an idealized flow analysis to eliminate infinite values of $u$ at this edge.\n\nJet contraction or expansion.- It has already been pointed out that, with a body in the jet, the velocity on the free surface is not necessarily equal to the velocity far upstream in the closed portion. During the course of the investigation, it was noted that solutions could be obtained showing a difference between these two velocities, even when there was no body in the jet. Such a flow corresponds merely to a contraction or expansion of the jet, as indicated in figure 2. Thus, in figure 2(a), the velocity on the free surface is lower than the upstream velocity and remains so even as it approaches the exit, in spite of the gradual contraction of the jet, because of the continuously increasing surface curvature. The velocity suddenly increases at the exit lip and finally is established at a value greater than that of the upstream velocity. With reasonable ratios of entrance to exit area, the flows of figure 2 may be readily obtained experimentally.\n\nThe significance of this expanding or contracting flow is that it represents a solution that satisfies all the boundary conditions previously discussed and is nevertheless undesirable. In order to avoid such solutions, a further condition must accordingly be recognized; namely, that the velocities in the closed portions far upstream and far downstream of the open section be equal.\n\nIt may be objected that in the normal design of an open wind tunnel the exit section is made larger than the entrance section. The purpose of the increased area is to allow for the reduced velocity toward the surface of the jet resulting from turbulent mixing with the surrounding still air. Increasing the exit area by other than the correct amount will result in the type of flow indicated in figures 2(a) or 2(b), with a corresponding velocity gradient along the center of the tunnel. In any potential-flow solution these viscous effects cannot be considered.\n\nSpillage.- When an airfoil is tested at a high lift coefficient in an open tunnel, the downward deflection of the jet may result in appreciable spillage from the lower lip of the exit, together with lack of contact of the main flow with the upper lip. (See fig. 3(a).) The air lost by spillage is replaced by air (of, however, a lower total pressure) entrained in the exit. Even without otherwise considering the distortion of the free surface, these flow characteristics might seem too much at variance with the previously assumed characteristics to permit application of the theories being discussed. The calculations of part II for", "timestamp": "2026-07-22T04:41:15.166363+00:00"}
{"citation_id": "19930085930", "source_url": "https://ntrs.nasa.gov/api/citations/19930085930/downloads/19930085930.pdf", "page_number": 91, "total_pages": 92, "image_filename": "19930085930_p91.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T04:41:16.441838+00:00"}