PDF/ocr_dataset_1062.jsonl

{"citation_id": "19930085870", "source_url": "https://ntrs.nasa.gov/api/citations/19930085870/downloads/19930085870.pdf", "page_number": 1, "total_pages": 92, "image_filename": "19930085870_p1.jpg", "text": "CONFIDENTIAL\nCopy No. 367\nRM No. L9D07\n\nMueller\nNACA\nRep. 1238\n\nNACA\n\nRESEARCH MEMORANDUM\n\nINVESTIGATIONS AT SUPERSONIC SPEEDS OF 22 TRIANGULAR\nWINGS REPRESENTING TWO AIRFOIL SECTIONS FOR\nEACH OF 11 APEX ANGLES\n\nBy\nEugene S. Love\n\nLangley Aeronautical Laboratory\nLangley Air Force Base, Va.\n\nAUTHOR'S PERSONAL COPY\nCLASSIFIED DOCUMENT\n\nCLASSIFICATION CHANGED TO\nUNCLASSIFIED\nDATE 8-19-58\n\nThis document contains classified information\naffecting the National Defense of the United\nStates within the meaning of the Espionage Laws,\nUSC 50:31 and 32. Its transmission or the\nrevelation of its contents in any manner to an\nunauthorized person is prohibited by law.\nInformation so classified may be imparted\nonly to persons in the military and naval\nservices of the United States, appropriate\ncivilian officers and employees of the Federal\nGovernment who have a legitimate interest\ntherein, and to United States citizens of known\nloyalty and discretion who of necessity must be\ninformed thereof.\n\nAUTHORITY MR. J. W. CROWLEY\nCHANGE NO. 2430\nE.L.B.\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\nWASHINGTON\nMay 10, 1949\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:19:09.635018+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 51, "total_pages": 98, "image_filename": "19930086073_p51.jpg", "text": "```markdown\nNACA RM A9E04\n\nLift coefficient, $C_L$\n\nDrag coefficient, $C_D$\n\nAngle of sideslip, $\\beta$, deg\n\n(b) $C_L$ vs $C_D$.\n\nFigure 10.— Continued.\n\n49\n```", "timestamp": "2026-07-22T06:19:10.640647+00:00"}
{"citation_id": "19930085847", "source_url": "https://ntrs.nasa.gov/api/citations/19930085847/downloads/19930085847.pdf", "page_number": 13, "total_pages": 32, "image_filename": "19930085847_p13.jpg", "text": "NACA RM A9D04 CONFIDENTIAL 11\n\n$38' \\cdot 10 \\frac{1}{2}''$\n\n$8' 1 \\frac{1}{4}''$\n\nAirspeed boom\n\nSlot\n\nExit\n\n$11 \\frac{1}{4}''$\n\nProfile drag rake\n\n[Figure: Plan form of test airplane showing location of suction slot and duct exit.]\n\nFigure 2.- Plan form of test airplane showing location of suction slot and duct exit.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:19:11.148845+00:00"}
{"citation_id": "19930082918", "source_url": "https://ntrs.nasa.gov/api/citations/19930082918/downloads/19930082918.pdf", "page_number": 19, "total_pages": 62, "image_filename": "19930082918_p19.jpg", "text": "18 NACA TN 1940\n\n(3) Broadening of the (220) line and hardening in the case of material aged at $1400^\\circ$ and $1600^\\circ$ F. The point is not known for aging at $1200^\\circ$ F since it was not possible to obtain either appreciable hardening or broadening at this temperature because of the excessive aging time required to reach high hardnesses.\n\nConversely, then, for materials aged at $1200^\\circ$, $1400^\\circ$, or $1600^\\circ$ F, retention of creep strength was associated with:\n\n(1) As-solution-treated material or aged material still in the short-period nucleation state.\n\n(2) Relatively little or no visible general matrix precipitation and incomplete boundary reaction in the case of material aged at $1200^\\circ$ and $1400^\\circ$ F. In all cases pertaining to material aged at $1400^\\circ$ F precipitate particles and the boundary phase, if present at all, had concentration gradients surrounding them after the aging periods associated with high creep resistance.\n\n(3) Low hardness and the unbroadened (220) diffraction line in the case of unaged material or material aged at $1400^\\circ$ and $1200^\\circ$ F.\n\nIt would seem then that, as far as creep resistance at $1200^\\circ$ F and at the 30,000-psi stress level was concerned, the removal of either the precipitant atoms from random solid solution or small nuclei made up of them by subsequent precipitate growth led to a progressive lowering of the creep resistance. Further, the probable long-period strains associated with line broadening and relatively high hardness resulting from large nuclei or precipitate particles did not control the creep strength. Rather, the continued removal of the atoms required to make up the precipitate, which in turn caused the long-period internal strain, resulted in still lower creep strength. It was the strain associated with the individual precipitant atoms, while still in random solid solution or in small nuclei, which controlled the creep strength at $1200^\\circ$ F. and a stress of 30,000 psi.\n\nFrom figure 15 it can be seen that when considering creep resistance at 60,000 psi and $1200^\\circ$ F the effect of long-time aging was qualitatively the same as at 30,000 psi, but that the material unaged and short-time aged at $1400^\\circ$ F had relatively less creep resistance than the same material at 30,000 psi. It thus appeared that, while long-time aging and concomitant matrix depletion of the precipitant atoms resulted in lowering of the creep strength, there was an optimum state of precipitation or nucleation and corresponding internal strain for an optimum creep resistance. However, an additional factor applied here. The relatively high stress of 60,000 psi resulted in fracture of all the creep specimens within a maximum of 32 hours. Figure 18 shows that the material unaged or short-time aged at $1400^\\circ$ F failed with general intergranular cracking.", "timestamp": "2026-07-22T06:19:11.411286+00:00"}
{"citation_id": "19930082592", "source_url": "https://ntrs.nasa.gov/api/citations/19930082592/downloads/19930082592.pdf", "page_number": 33, "total_pages": 50, "image_filename": "19930082592_p33.jpg", "text": "**Page intentionally left blank**\n\n**Page intentionally left blank**", "timestamp": "2026-07-22T06:19:12.166009+00:00"}
{"citation_id": "19930085862", "source_url": "https://ntrs.nasa.gov/api/citations/19930085862/downloads/19930085862.pdf", "page_number": 6, "total_pages": 60, "image_filename": "19930085862_p6.jpg", "text": "4\nNACA RM No. L9A07\n\n$C_{l\\delta}$ rate of change of rolling-moment coefficient with aileron deflection (aileron effectiveness)\n\n$C_{h\\delta}$ rate of change of hinge-moment coefficient with aileron deflection\n\n$C_{h\\alpha}$ rate of change of hinge-moment coefficient with angle of attack\n\n$P_{R\\delta}$ rate of change of aileron-balance-chamber pressure coefficient with aileron deflection\n\n$P_{R\\alpha}$ rate of change of aileron-balance-chamber pressure coefficient with angle of attack\n\n$C'_{h\\delta}$ rate of change of hinge-moment coefficient in a steady roll with aileron deflection\n\nMODEL\n\nThe principal dimensions of the model are shown in figure 1. Photographs of the model mounted in the Langley 19-foot pressure tunnel are shown in figure 2. The wing was of solid steel construction and had an aspect ratio of 3.94 and a taper ratio of 0.625. A straight line connecting the leading edge of the root and theoretical tip chords was swept back 42.05°. The symmetrical circular-arc airfoil sections were fabricated with a constant radius of 83.26 inches in a plane perpendicular to the line of maximum thickness. As a result, the leading and trailing edges were slightly curved in plan form. The maximum divergence from a straight line connecting the root and theoretical tip chords at the leading and trailing edges was about 0.4 inch. The airfoil sections, taken normal to the line of maximum thickness, had a maximum thickness of 10 percent of the chord at the root and 6.4 percent of the chord at the tip. Parallel to the plane of symmetry the maximum thickness was 7.9 percent of the chord at the root and 5.2 percent of the chord at the tip.\n\nThe high-lift and stall-control devices used on the model are shown in figure 3. The drooped-nose flaps extended over the outer 60 percent of the wing, had a chord of 0.184c, and were deflected 30° measured in a plane normal to the hinge line. The amount of flap deflection was based upon unpublished data which indicated 30° to be optimum for this wing from considerations of pitching moment and maximum lift. The extensible leading-edge flaps had a span of $0.55\\frac{b}{2}$ and extended from $0.425\\frac{b}{2}$ to $0.975\\frac{b}{2}$ (beginning of rounded tip). The chord was constant and amounted", "timestamp": "2026-07-22T06:19:15.860328+00:00"}
{"citation_id": "19930085548", "source_url": "https://ntrs.nasa.gov/api/citations/19930085548/downloads/19930085548.pdf", "page_number": 15, "total_pages": 46, "image_filename": "19930085548_p15.jpg", "text": "```markdown\n14\nNACA RM No. ESL30\n\nThe two most serious deficiencies of the experimental cylinder with regard to gas-generator application are its low charging efficiency and its high heat losses. The low charging efficiency causes a loss in power output, which may limit the manifold pressure at which the gas-generator engine can operate. This limitation in turn leads to higher specific weight and specific fuel consumption in the gas-generator engine.\n\nThe nature of this power limit is shown in figure 23. The cylinder power output is shown for constant exhaust-gas temperature and constant maximum cylinder pressure. The operating point of the gas-generator engine obviously is that point where the power-available curve intersects the power-required curve. Because all the curves are so nearly parallel, a small drop in power output resulting from inadequate scavenging results in a large loss in manifold pressure.\n\nThe gas-generator engine ordinarily operates at fuel-air ratios of about 0.03 to 0.035 and if the scavenging is adequate to keep the cylinder fuel-air ratio somewhat below stoichiometric, complete scavenging is no longer so important. The present cylinder is incapable of accomplishing this end. Revision of the porting scheme with particular emphasis on increasing the exhaust lead may effect a satisfactory improvement.\n\nThe second fault of the present cylinder, that of excessive heat losses, is attributed to the use of a small-scale cylinder with a low coolant temperature. In this investigation, overcooling the cylinder rather than developing a cylinder that would operate well with a minimum of cooling was expedient. The use of a full-scale cylinder with a certain amount of development to permit the use of higher coolant temperatures should be effective in reducing the heat losses. A reduction of manifold cooling area would also be made possible with a full-scale cylinder. Despite poor scavenging and high heat losses, the performance of the experimental cylinder confirmed the assumptions used in the previous analysis sufficiently well to indicate that a reasonable approach could be made to the gas-generator-engine performance calculated in reference 1.\n\nMECHANICAL PERFORMANCE OF ASSEMBLY\n\nThe operating conditions used in the investigation represent quite a radical departure from conventional practice. Accordingly, the practicability of operating at these conditions may be questioned.\n\n1077\n```", "timestamp": "2026-07-22T06:19:18.429893+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 93, "total_pages": 96, "image_filename": "19930085880_p93.jpg", "text": "NACA RM No. L9C03\n91\n\n[Figure: A graph plotting Draft, ft. against Wetted area, sq ft. The y-axis ranges from 0 to .64. The x-axis ranges from 0 to .35. The graph contains a legend for Speed (fps) with symbols: circle for 10, square for 15, diamond for 20, triangle for 25, and inverted triangle for 30. The data points show a curve rising from the origin and leveling off around .16 ft draft. A small diagram of a triangle is shown in the legend box.]\n\n(b) $\\tau = 8^\\circ$.\nNACA\nFigure 25.- Continued.", "timestamp": "2026-07-22T06:19:35.034828+00:00"}
{"citation_id": "19930082592", "source_url": "https://ntrs.nasa.gov/api/citations/19930082592/downloads/19930082592.pdf", "page_number": 34, "total_pages": 50, "image_filename": "19930082592_p34.jpg", "text": "NACA TN 1914\n33\n\nLine following contour\nof oxidation interface\nBakelite\nOxidation\ninterface\n\n[Figure: Micrograph showing oxidation zone of 10-percent-molybdenum - titanium carbide ceramal. Line within oxide coating follows contour of oxidation interface. Temperature, 1785° F; time at temperature, 40 hours; unetched; magnification, X50.]\n\nNACA\nC-22910\n2-7-49\n\nFigure 11. - Oxidation zone of 10-percent-molybdenum - titanium carbide ceramal. Line within oxide coating follows contour of oxidation interface. Temperature, 1785° F; time at temperature, 40 hours; unetched; magnification, X50.", "timestamp": "2026-07-22T06:19:43.199362+00:00"}
{"citation_id": "19930082703", "source_url": "https://ntrs.nasa.gov/api/citations/19930082703/downloads/19930082703.pdf", "page_number": 26, "total_pages": 28, "image_filename": "19930082703_p26.jpg", "text": "24\nNACA TN 1983\n\nStick motion\nfrom trim, in.\nRearward 5\n0\nForward 5\n-Rearward stop\n-Forward stop\n\nPitching velocity,\ndeg/sec\nNose up 20\n0\nNose down 20\n\nNormal\nacceleration,\ng\n1.5\n1.0\n.5\n0\n0 2 4 6 8\nTime, sec\nNACA\n\nFigure 7.- Time history of a pull-up maneuver for helicopter B\nat 80 miles per hour.\n\nStick motion\nfrom trim, in.\nRearward 5\n0\nForward 5\n-Rearward stop\n-Forward stop\n\nPitching velocity,\ndeg/sec\nNose up 20\n0\nNose down 20\n\nNormal\nacceleration,\ng\n1.5\n1.0\n.5\n0\n0 2 4 6 8 10 12\nTime, sec\nNACA\n\nFigure 8.- Time history of a pull-up maneuver for helicopter C\nat 80 miles per hour.", "timestamp": "2026-07-22T06:19:48.977592+00:00"}
{"citation_id": "19930082646", "source_url": "https://ntrs.nasa.gov/api/citations/19930082646/downloads/19930082646.pdf", "page_number": 17, "total_pages": 37, "image_filename": "19930082646_p17.jpg", "text": "16\nNACA TN 1980\n\nElevator deflection, $\\delta_e$, deg\n0 ———\n-10 - - - -\n-20 — —\n\n[Figure: Six graphs arranged in three rows and two columns. The left column is labeled (a) and the right column is labeled (b). The vertical axis for all graphs is labeled \"Trim, $\\delta_e$, deg\" and ranges from 0 to 12. The horizontal axis for all graphs is labeled \"Speed, mph\" and ranges from 0 to 80. The top row of graphs is labeled \"c.g., 24 percent M.A.C.\". The middle row is labeled \"c.g., 30 percent M.A.C.\". The bottom row is labeled \"c.g., 36 percent M.A.C.\". The bottom right graph contains a NACA logo.]\n\n(a) Warped forebody and extended afterbody.\n(b) Basic forebody and basic afterbody.\n\nFigure 5.- Variation of trim with speed during take-off.", "timestamp": "2026-07-22T06:19:49.654114+00:00"}
{"citation_id": "19930085485", "source_url": "https://ntrs.nasa.gov/api/citations/19930085485/downloads/19930085485.pdf", "page_number": 18, "total_pages": 26, "image_filename": "19930085485_p18.jpg", "text": "16\n\n$1.4 \\times 10^6$\n\nCONFIDENTIAL\n\nReynolds number, R\n\n1.2\n\n.10\n\n.8\n\n.6\n\n0 .2 .4 .6 .8 1.0 1.2\n\nMach number, M\n\nNACA\n\nFigure 7.- Variation of Reynolds number with test Mach number. Reynolds number based on model mean geometric chord of 0.25 foot.\n\nCONFIDENTIAL\n\nNACA RM No. L8E02", "timestamp": "2026-07-22T06:19:51.019834+00:00"}
{"citation_id": "19930083221", "source_url": "https://ntrs.nasa.gov/api/citations/19930083221/downloads/19930083221.pdf", "page_number": 19, "total_pages": 47, "image_filename": "19930083221_p19.jpg", "text": "NACA TN No. 1824\n17\n\nThe load distributions\nand indicial lift functions\nwill be given here for values\nof t up to the time when\nthe characteristic from the\ntrailing edge first crosses\nthe leading-edge trace, that\nis, for\n\n$$0 \\le t \\le \\frac{c_o}{1-M_o}$$\n\nor\n\n$$0 \\le t' \\le \\frac{c_o}{a_o(1-M_o)}$$\n\nThis period covers the time\nin which the wing travels\n$2M_o/(1-M_o)$ half chords and,\nsince the present analysis\nis concerned with values of Mach number near one, will in some cases\nextend beyond the range of the linear theory.\n\n<!-- Image (512, 106, 877, 385) -->\n\nFigure 8.- Regions used in the\nstudy of subsonic unsteady lift.\n\nThe following results are obtained for load coefficient:\n\nRegion I (between lines x=t, t=0, and x=$c_o$-t)\n\n$$\\frac{\\Delta p}{q} = \\frac{4\\alpha}{M_o} \\quad (27a)$$\n\nRegion II (between lines x=$-M_o$t, x=t, and x=$c_o$-t)\n\n$$\\frac{\\Delta p}{q} = \\frac{8\\alpha}{\\pi M_o} \\left( \\frac{M_o}{1+M_o} \\sqrt{\\frac{t-x}{M_o t+x}} + \\text{arc tan} \\sqrt{\\frac{M_o t+x}{t-x}} \\right) \\quad (27b)$$\n\nRegion III (between lines x=$c_o$-t, x=t, x=-t + $\\frac{2c_o}{1+M_o}$, and t = $\\frac{c_o}{1-M_o}$)", "timestamp": "2026-07-22T06:19:51.405849+00:00"}
{"citation_id": "19930085862", "source_url": "https://ntrs.nasa.gov/api/citations/19930085862/downloads/19930085862.pdf", "page_number": 7, "total_pages": 60, "image_filename": "19930085862_p7.jpg", "text": "NACA RM No. L9A07\n\nto about 18 and 13 percent of the wing chord at the outboard and inboard ends, respectively. The deflection was $37^\\circ$, measured in the manner shown in figure 3. The 0.20c trailing-edge split flaps extended over the inboard 50 percent of the wing semispan and were deflected $60^\\circ$ from the lower surface of the wing.\n\nThe upper-surface fences (fig. 3) were mounted normal to the wing surface and parallel to the plane of symmetry. They projected 0.6 of the maximum thickness of the root section above the wing surface. When used in conjunction with the drooped-nose flaps, the fences extended from the wing trailing edge to about the 0.18c point and were located $0.05\\frac{b}{2}$ outboard of the inboard ends of the drooped-nose flaps. For the configurations with the extensible leading-edge flaps, the fences extended from the trailing edge to the leading edge of the wing and were located $0.025\\frac{b}{2}$ outboard of the inboard ends of the flaps.\n\nOnly the left side of the wing was equipped with the sealed, unbalanced, contour aileron. The aileron chord was about 0.18c, and the span was $0.475\\frac{b}{2}$, with the inboard end located at $0.5\\frac{b}{2}$. Resistance-type electrical strain gages were employed to measure the aileron normal forces and hinge moments. The aileron seal, which was designed in a manner so that no moments and negligible forces were transferred from it to the aileron, extended the full span of the aileron except for cut-outs to allow for the mounting of the strain-gage beams. Pressure orifices were installed in the aileron balance chamber to enable the pressure differences across the seal to be determined. The details of the aileron are given in figure 4.\n\nThe spoilers used were of the step type. The span of each step was $0.10\\frac{b}{2}$ with the exception of the outboard one which was $0.075\\frac{b}{2}$. With all steps in place, the spoilers extended from the $0.20\\frac{b}{2}$ station outboard to the $0.975\\frac{b}{2}$ station. Spoiler projections of 0.05c and 0.10c were tested. The spoilers were normal to the wing surface and to the plane of symmetry. They were located on the 0.70c line of the left wing panel in the manner shown in figure 4.\n\nTESTS\n\nThe tests were made in the Langley 19-foot pressure tunnel with the air in the tunnel compressed to approximately $2\\frac{1}{3}$ atmospheres. Measurements of the lift and drag and the pitching, rolling, and yawing moments", "timestamp": "2026-07-22T06:19:55.161034+00:00"}
{"citation_id": "19930085548", "source_url": "https://ntrs.nasa.gov/api/citations/19930085548/downloads/19930085548.pdf", "page_number": 16, "total_pages": 46, "image_filename": "19930085548_p16.jpg", "text": "NACA RM No. E8L30\n15\n\nMany of the anticipated difficulties, such as, roughness of the engine, broken and stuck piston rings, broken cylinders, connecting rods, and pistons, and rapid wear of parts failed to materialize. The engine operated satisfactorily throughout the entire investigation. Ring sticking was not a problem, nor was combustion roughness or knock encountered. The average rate of combustion-pressure rise was 32.5 pounds per square inch per degree at 1800 rpm or 352,000 pounds per square inch per second as determined from indicator cards. A minimum rate of pressure rise of about 50 pounds per square inch per degree will usually cause engine roughness. It should be pointed out, however, that no attempt has been made to reduce the high heat losses from the cylinder; the higher cylinder temperatures that would accompany such an attempt have not been investigated.\n\nSUMMARY OF RESULTS\n\nThe results of an investigation of the performance of a small-scale, two-stroke-cycle, compression-ignition cylinder of the loop-scavenged type operated under simulated piston-type gas-generator-engine conditions may be summarized as follows:\n\n1. Charging and scavenging of the cylinder was inadequate because of unsatisfactory porting. The poor charging and scavenging were traced to inadequate exhaust lead, which was found to be only one-sixth of that required. The charging efficiency adversely affected power output and thermal efficiency of the cylinder at overall fuel-air ratios in excess of 0.03.\n\n2. The thermal efficiency and the power output at over-all fuel-air ratios less than 0.03 checked satisfactorily with anticipated values. Small improvements may be obtained by optimizing the fuel-injection-system characteristics.\n\n3. Heat losses from the cylinder were excessive. Part of the reason for these large losses was the necessarily large surface-volume ratio of the cylinder, which was about $2\\frac{1}{2}$ times that of a current full-scale two-stroke-cycle cylinder. A contributing factor was the low coolant temperature used to expedite the investigation.\n\n4. Measured exhaust-gas temperatures checked anticipated values reasonably well when the experimentally determined heat losses and combustion efficiency were considered.\n\n5. The combustion pressure rise was about one-half that to be expected with constant-volume combustion. The combustion pressure", "timestamp": "2026-07-22T06:19:56.379842+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 35, "total_pages": 78, "image_filename": "19930082483_p35.jpg", "text": "NACA TN No. 1807\n\nnot a second-power variation, as would be expected, is attributed to inaccuracy of the experimental method employed. The error involved, however, is slight, if not negligible.\n\nPumping loss. - Pumping losses are shown in figure 9(c) for the range of admissions considered. The curves serve to demonstrate the increase of this partial-admissions loss with increased blocking. As is indicated by equation (21), this loss varies as the third power of the rotor speed.\n\nThe importance of this loss in a particular application is determined largely by the pitch-line diameter of the turbine because the pumping loss varies as the fourth power of the pitch-line diameter.\n\nDriving-fluid losses. - The driving-fluid losses are presented in figure 9(d) for several degrees of admission. Because these losses are directly proportional to rotor speed, the curves appear as straight lines with a slope of unity. For this turbine, it is noted that the driving-fluid losses are quite large as compared with the other losses considered.\n\nAnalysis of the turbine data confirms that the driving-fluid losses due to partial admission as expressed by equation (27) are representative of actual performance. The power difference representing the driving-fluid losses is experimentally found to be proportional to the reduction of active nozzle arc and may be correlated by the product $K_{III} p_1 N$ over most of the operating range. Verification of this analysis is provided in figure 14, where the driving-fluid losses for $180^\\circ$ and $120^\\circ$ admission are experimentally determined. From full-admission data by use of equation (19), the gross powers at $180^\\circ$ and $120^\\circ$ admission are obtained. The blade power for each of these partial admissions is found, as in equation (4), from the observed net power output plus the shaft losses obtained from figures 9(b) and 9(c). These power terms are then plotted against corrected rotor speed and, according to equation (26), (blade power at full admission being equal to gross power at full admission) the driving-fluid losses, identified by the shaded areas, are obtained as the difference between the gross powers and the blade powers for each of these two admissions, respectively.\n\nIt can be seen from figure 14 that the driving-fluid losses are proportional to the rotor speed and likewise proportional to the amount of nozzle blocking. Driving-fluid losses may therefore be properly expressed by equation (27) and related at various degrees of admission by equation (28).", "timestamp": "2026-07-22T06:20:02.172804+00:00"}
{"citation_id": "19930082918", "source_url": "https://ntrs.nasa.gov/api/citations/19930082918/downloads/19930082918.pdf", "page_number": 20, "total_pages": 62, "image_filename": "19930082918_p20.jpg", "text": "NACA TN 1940\n\nThis formation of cracks resulted undoubtedly in greater creep rates than would be established with the same material with greater boundary strength. Hence, it was felt that the general effects of matrix depletion of the precipitant atoms observed at a lower stress level still apply but that the unaged and short-time-aged material has its creep resistance lowered, relatively speaking, at higher stresses by virtue of a weak grain boundary area.\n\nThe fact that the line-broadening strains, most probably associated with the individual precipitate particles, or large nuclei, which resulted in high hardness, did not have any effect on increasing creep resistance is quite surprising at first thought. However, since the short-period strains surrounding the individual precipitant atoms when in random or nearly random solid solution do give high creep resistance, it appears that the answer, as far as the alloy considered here is concerned, lies in the periodicity and size of the two types of strained regions. The atoms most likely to be the source of the short-period strains, in the solution-treated material, are tungsten, molybdenum, columbium, carbon, and nitrogen. The total atomic fraction of these atoms is $0.042$, or one atom in 24. (See table 6.) In the absence of further data this can be interpreted to mean that one atom in 24 is the center of a strained region, the period of such strains being of the order of $6 \\times 10^{-7}$ centimeter. Since interatomic forces of solids in general extend over only a few atoms (see reference 8), the size of the strained regions might be assumed to be of the order of $1 \\times 10^{-7}$ centimeter. This leaves a mean strain-free-path of the order of $5 \\times 10^{-7}$ centimeter in the solution-treated material. In the aged material it seems plausible to assign the periodicity of the precipitate particles as the periodicity of the line-broadening strains. This spacing was of the order of $10^{-4}$ centimeter in the hardest sample prepared, that is, material aged 1000 hours at $1400^\\circ$ F.\n\nIt appears, unfortunately, that at present it is impossible to estimate closely the actual size, at any given aging time, of the strained areas associated with each of the precipitant particles or the average size of the precipitant-atom-free-paths. This is because enough is not yet known to separate the line broadening due to elastic strains from the line broadening due to concentration gradients in the matrix. However, it can be said that the mean precipitant-atom-free-path approached $10^{-4}$ centimeter as diffusion and precipitation approached completion. This path was more than two orders of magnitude greater than the original mean precipitant-atom-free-path.\n\nThe net result of aging, then, in solution-treated low-carbon N-155 alloy, was probably to replace a relatively short-period small-strain system containing a short mean strain-free-path ($5 \\times 10^{-7}$ cm) with another larger-strain system of longer period and much longer mean strain-free-path (approaching $10^{-4}$ cm). Since creep as here", "timestamp": "2026-07-22T06:20:15.915615+00:00"}
{"citation_id": "19930085859", "source_url": "https://ntrs.nasa.gov/api/citations/19930085859/downloads/19930085859.pdf", "page_number": 11, "total_pages": 31, "image_filename": "19930085859_p11.jpg", "text": "NACA RM No. L9B25\n9\n\nTABLE I.- FUSELAGE ORDINATES\n[Basic fineness ratio 12; actual fineness ratio 10\nachieved by cutting off the rear one-sixth of\nthe body; $\\bar{c}/4$ located at $l/2$]\n\n[Figure: Diagram of fuselage with dimensions $l=14.14$, $\\frac{5}{6}l$, $\\frac{l}{2}$, $x$, $r$, and $D(Max)$]\n\n| Ordinates | | | |\n| :--- | :--- | :--- | :--- |\n| $x/l$ | $r/l$ | $x/l$ | $r/l$ |\n| 0 | 0 | 0 | 0 |\n| .005 | .00231 | .4500 | .04143 |\n| .0075 | .00298 | .5000 | .04167 |\n| .0125 | .00428 | .5500 | .04130 |\n| .0250 | .00722 | .6000 | .04024 |\n| .0500 | .01205 | .6500 | .03842 |\n| .0750 | .01613 | .7000 | .03562 |\n| .1000 | .01971 | .7500 | .03128 |\n| .1500 | .02593 | .8000 | .02526 |\n| .2000 | .03090 | .8338 | .02000 |\n| .2500 | .03465 | .8500 | .01852 |\n| .3000 | .03741 | .9000 | .01125 |\n| .3500 | .03933 | .9500 | .00439 |\n| .4000 | .04063 | 1.0000 | 0 |\n\nL. E. radius = 0.00051\n\nNACA", "timestamp": "2026-07-22T06:20:19.008937+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 18, "total_pages": 149, "image_filename": "19930083192_p18.jpg", "text": "14\nNACA TN 1976\n\nData on the spanwise distribution were sorted according to the shape of the distribution as indicated in figure 15. Six shapes ranging from the uniform or rectangular distribution to the double triangular distribution are shown, and under each shape is noted the total number of gusts classified in that category. Figure 15 also indicates possible variations included under each heading.\n\nThe spanwise gust distributions represented in figure 15 were, in turn, evaluated graphically to obtain values of the lateral gust-gradient distance H. The results are shown in figure 16 where the average values of the gradient distance in chords are given for a range of values of gust velocity. The associated gust velocity U was obtained from the synchronized acceleration data, and the faired curve shown in figure 16 corresponds to that given in figure 12(a).\n\nThe significance of the values of $U_e$ obtained from pressure measurements as compared to those determined from acceleration measurements has been examined in reference 12. Figure 17, reproduced from that reference, shows the average value of the local effective gust velocity as a function of the effective gust velocity for a given gust as determined from acceleration records. The average value of $U_e$ from the pressure measurements is the average value weighted according to the amount of wing area represented by the particular orifice or measuring station. Also shown in the figure is the line for equality of the two measures of $U_e$ and the error band.\n\nSince, in some cases, it is convenient to divide the lateral gust distribution into symmetrical and unsymmetrical components, the gust shapes listed in figure 15 were classified according to whether they were symmetrical or unsymmetrical and then represented as having a linear variation of gust velocity across the span of the airplane. The results are shown in figure 18. The inset in figure 18 indicates the type of variation assumed and the two quantities obtained from the evaluation. The average gust velocity $U_{e_{av}}$ and the increment in gust velocity at each wing tip $\\Delta U_e$ were taken as the symmetrical and unsymmetrical components, respectively. Many small values have been omitted from the figure.\n\nAs in the case of the gust-gradient distance, the data have been further analyzed to find an average unsymmetrical gust increment. Figure 19 has been prepared from the summary of data for the different gust shapes to show the average gust velocity for a gust of the unsymmetrical shape as a function of the average gust velocity of a triangular gust for equal frequencies of occurrence.\n\nIn addition to the determination of spanwise gust distribution from local pressure distributions, synchronized records from accelerometers located at the center of gravity of the airplane and in the wing were evaluated to obtain the angular acceleration in roll, which is related", "timestamp": "2026-07-22T06:20:19.539281+00:00"}
{"citation_id": "19930085485", "source_url": "https://ntrs.nasa.gov/api/citations/19930085485/downloads/19930085485.pdf", "page_number": 19, "total_pages": 26, "image_filename": "19930085485_p19.jpg", "text": "```markdown\nCONFIDENTIAL\n\n(a) Circular-arc aileron, t=0\n\n(b) Flat-sided aileron, t=.50\n\n(c) Flat-sided aileron, t=1.00\n\nLift coefficient, $C_L$\n\n$\\alpha$ (deg)\n$\\diamond$ 6.0\n$\\square$ 4.1\n$\\triangle$ 2.6\n$\\circ$ 0\n$\\nabla$ -2.9\n\nMach number, M\n\nNACA\n\nFigure 8.- Variation of lift coefficient with Mach number for the model with several aileron profiles. Data from transonic bump.\n\nCONFIDENTIAL\n\nCONFIDENTIAL\nAileron\n—— Flat-sided, t = .50\n--- Flat-sided, t = 1.00\n\n$\\alpha$ (deg)\n\nIncrement of drag coefficient, $\\Delta C_D$\n\nMach number, M\n\nNACA\n\nFigure 9.- Variation with Mach number of increment of drag coefficient caused by changing aileron contour from circular arc. Data from transonic bump.\n\nCONFIDENTIAL\n\nNACA RM No. L8K02\n\n17\n```", "timestamp": "2026-07-22T06:20:26.839602+00:00"}
{"citation_id": "19930085862", "source_url": "https://ntrs.nasa.gov/api/citations/19930085862/downloads/19930085862.pdf", "page_number": 8, "total_pages": 60, "image_filename": "19930085862_p8.jpg", "text": "6\nNACA RM No. I9A07\n\nwere made for each configuration through an angle-of-attack range\nextending from -4° to beyond maximum lift. For the aileron tests the\nnormal forces, hinge moments, and balance-chamber pressures of the aileron\nwere determined for aileron deflections ranging from 25° to -25°. The\nspoiler tests were made by using various spans of step spoilers, in incre-\nments of approximately 20 percent of the semispan, starting either\nfrom $0.975\\frac{b}{2}$ or $0.20\\frac{b}{2}$ spanwise stations. The stall studies were made by\nvisual observation and from motion-picture records of the behavior of wool\ntufts attached to the upper surface of the wing.\n\nAll of the spoiler tests and the stall studies were conducted at a\nMach number of 0.15 and a Reynolds number of $6.9 \\times 10^6$, based on the\nwing mean aerodynamic chord. The aileron tests, with the exception of\nthe plain wing configuration which was tested at a Reynolds number\nof $6.9 \\times 10^6$, were conducted at a Reynolds number of $5.3 \\times 10^6$ and a\nMach number of 0.11. Scale-effect tests were not made, since reference 1\nhas indicated no appreciable scale effect in this Reynolds number range.\n\nREDUCTION OF DATA\n\nAll data have been reduced to standard nondimensional coefficients.\nCorrections have been applied to the force and moment data to account\nfor the tare and interference effects of the model support system.\nStream-inclination and jet-boundary corrections have been applied to the\nangle of attack and to the drag and pitching-moment coefficients. Jet-\nboundary corrections to the rolling- and yawing-moment coefficients were\nfound to be negligible and, therefore, were not applied to the data.\n\nThe aileron hinge-moment coefficients presented herein are based\nupon the product of the aileron span and the square of the root-mean-\nsquare chord. Some recent practice (reference 2) has based the hinge-\nmoment coefficients upon twice the area moment of the aileron. The\ncoefficients presented herein may be converted to this base by means of\nthe following equation:\n\n$$C_{h_a} \\text{ (based on twice area moment)} = 0.952 C_{h_a} \\text{ (presented herein)} \\quad (1)$$\n\nAs a result of the interference of the strain-gage beams, the aileron\nseal was incomplete and a small amount of leakage across it occurred. A\ncalibration of the leakage was made, and the resultant pressure coefficients\ncorrected to a no-leakage condition. The effect of the leakage on the\nrolling-moment and hinge-moment coefficients is believed to be small and\nhas been neglected.", "timestamp": "2026-07-22T06:20:34.119910+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 94, "total_pages": 96, "image_filename": "19930085880_p94.jpg", "text": "92\nNACA RM No. L9C03\n\n.64\nSpeed\n(fps)\nO 10\n□ 15\n◇ 20\n△ 25\n▽ 30\n\n.56\n\n.48\n\n.40\n\n.32\nDraft, ft.\n\n.24\n\n.16\n\n.08\n\n0\n0 .05 .10 .15 .20 .25 .30 .35\nWetted area, sq ft\n\n(c) $\\tau = 12^\\circ$.\nNACA\nFigure 25.- Continued.", "timestamp": "2026-07-22T06:20:38.173796+00:00"}
{"citation_id": "19930082546", "source_url": "https://ntrs.nasa.gov/api/citations/19930082546/downloads/19930082546.pdf", "page_number": 31, "total_pages": 65, "image_filename": "19930082546_p31.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:20:40.770038+00:00"}
{"citation_id": "19930082646", "source_url": "https://ntrs.nasa.gov/api/citations/19930082646/downloads/19930082646.pdf", "page_number": 18, "total_pages": 37, "image_filename": "19930082646_p18.jpg", "text": "NACA TN 1980\n17\n\n<!-- Image (160, 100, 826, 882) -->\n\n(a) Lower-limit porpoising.\n\n(b) Upper-limit porpoising.\n\nFigure 6.- Maximum amplitude of porpoising at different positions of center of gravity.", "timestamp": "2026-07-22T06:20:48.597506+00:00"}
{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 19, "total_pages": 72, "image_filename": "19930085491_p19.jpg", "text": "18 CONFIDENTIAL NACA RM No. A3J04\n\nminimum drag coefficient and drag due to lift.\n\n2. The experimental value of lift-curve slope is less than that predicted by theory.\n\n3. With regard to longitudinal stability, theory indicates a linear variation of $C_m$ with $C_L$; whereas experiment shows a nonlinear variation which indicates an appreciable center-of-pressure travel.\n\nThese differences between experiment and theory are attributed to flow separation which is not considered in the linear theory but which was observed to exist on the model. The effect of this separation on the aerodynamic parameters $dC_L/d\\alpha$, $C_{D_{min}}$, $\\Delta C_D/(\\Delta C_L)^2$, $(L/D)_{max}$, and center-of-pressure location are discussed in the sections immediately following.\n\nLift-curve slope.— The lift curve in figure 7(c) is composed of two linear portions that join near $C_L = 0.09$. In the lower range, the slope is 0.038 and in the range above $C_L = 0.09$, the slope increases to 0.045, both values being less than the theoretical value of 0.051 which excludes the effects of viscosity. Some insight into the effects of viscosity at a Reynolds number of 0.62 million is possible through a correlation of the lift characteristics with the boundary-layer flow as observed by liquid-film tests.\n\nAt zero lift (fig. 13(b)) on both upper and lower surfaces laminar separation occurred at approximately 60-percent chord over the outboard sections. Theory (fig. 6) predicts this separation, since the pressure recovery over the rear of these sections is sufficiently large to cause the laminar boundary layer to separate. Although theory also indicates separation should occur on the inboard sections, the experimental result at this Reynolds number revealed no such separation. This disagreement on these sections probably results in part from the departure of the flow from the two-dimensional oblique cylindrical flow assumed to exist on all sections when calculating the theoretical location of laminar separation.\n\nAs the angle of attack is increased, the separation area on the upper surface expanded to include the area on the inboard sections and the separated area on the lower surface contracted to include only the tip sections.² This change in boundary-layer flow was\n\n---\n\n²These results are based on visual observations made with the model mounted horizontally in the tunnel. Hence, it was not possible to obtain plan-form photographs of the surface flow patterns.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:20:57.167793+00:00"}
{"citation_id": "19930085859", "source_url": "https://ntrs.nasa.gov/api/citations/19930085859/downloads/19930085859.pdf", "page_number": 12, "total_pages": 31, "image_filename": "19930085859_p12.jpg", "text": "```markdown\nTabulated wing data\nArea (Twice semispan) 0.1250 sq ft\nMean aerodynamic chord 0.1805 ft\nAspect ratio 4.0\nTaper ratio 0.6\nIncidence 0.0°\nDihedral 0.0°\nAirfoil section parallel to free stream NACA 65A006\n\n0.25-Chord line\n35°\nc̄ = 2.186\n4.243\n1.044\n90°\n2.652\nClearance 1/8\nReference center line\nBump surface\n11.8\nCenterline of balance normal to bump surface\n7.07\n1.16 D Max\n.225\n.56\nWing-alone end plate\nWing-fuselage end plate\n\nNACA\n0 1 2\nScale, inches\n\nFigure 1.- General arrangement of a model with 35° sweptback wing, aspect ratio 4, taper ratio 0.6, and NACA 65A006 airfoil.\n\nNACA RM No. L9B25\n10\n```", "timestamp": "2026-07-22T06:20:57.578099+00:00"}
{"citation_id": "19930082703", "source_url": "https://ntrs.nasa.gov/api/citations/19930082703/downloads/19930082703.pdf", "page_number": 27, "total_pages": 28, "image_filename": "19930082703_p27.jpg", "text": "NACA TN 1983\n25\n\nNormal\nacceleration,\ng\n3.0\n2.5\n2.0\n1.5\n1.0\n.5\n0\n(a) Helicopter A.\n\nNormal\nacceleration,\ng\n1.5\n1.0\n.5\n0\n0 2 4 6\nTime, sec\n(b) Helicopter B.\n\nFigure 9.- Theoretical time histories of normal acceleration following\na sudden rearward displacement of the control stick at 80 miles\nper hour. A change of $1^{\\circ}$ in longitudinal cyclic pitch is used.\n\nNACA-Langley - 11-16-49 - 900", "timestamp": "2026-07-22T06:20:57.844860+00:00"}
{"citation_id": "19930082918", "source_url": "https://ntrs.nasa.gov/api/citations/19930082918/downloads/19930082918.pdf", "page_number": 21, "total_pages": 62, "image_filename": "19930082918_p21.jpg", "text": "```markdown\n20\nNACA TN 1940\n\nconsidered is essentially a small-strain phenomenon, the few slip\nsystems in operation had a considerably larger probability of running\ninto a strained area per unit shear strain in unaged material than in\naged material with any appreciable amount of precipitate, despite the\nlarge average strains in the latter system. This means lower creep\nresistance in the aged material. Hardness as commonly measured,\nhowever, is a large-strain phenomenon. The whole volume of the\nmatrix is in a hardness test eventually filled with the many slip\nsystems formed. The relatively large strains associated with the\nactual precipitate particles act very effectively to hinder the slip\nsystems which eventually must form and try to pass through the large-\nstrain areas during a hardness test. Thus, aged low-carbon N-155 is\nharder than unaged stock. A small-strain hardness test would probably\ngive a much more accurate indication of the relative creep resistance\nof aged alloys of the low-carbon N-155 type.\n\nTungsten, molybdenum, and columbium can be considered sources of\nthese short-period lattice strains since they are uniformly atoms of\nlarger radii than those of the matrix and aging in solution-treated\nlow-carbon N-155 resulted in a decrease in lattice parameter, pre-\nsumably by rejection from solution of these large-radius atoms. The\nmatrix is assumed to be composed of iron, cobalt, nickel, and chromium,\nall atoms with approximately the same atomic radius (2.5Å). (See\ntable 6.) Carbon and nitrogen present interstitially also act to\nexpand the lattice and their removal through precipitation would make\nthe lattice contract. Thus these atoms must be considered along with\ntungsten, molybdenum, and columbium as sources of short-period strains\nwhile in random solid solution. The fact that solute atoms of\nconsiderably smaller or larger radius than those of the solvent\nintroduce strain into the system when in solid solution is well\nestablished generally, as witness the Hume-Rothery rules for solid-\nsolution limitation (reference 9). To quote another source, Sir\nLaurence Bragg, \"Most engineering alloys of importance are the ones\nderiving their strength, at least in part, from the modulation of the\nlattice due to the presence of foreign atoms of different size.\"$^1$\nFurther inspection of table 6 shows manganese and silicon to be atoms\nof smaller radius than the average radius for N-155; and these also\ncould be sources of short-period strains. Manganese and silicon are\nprobably in the lattice substitutionally and would tend to make its\nparameter smaller than normal and to make it increase with aging\ntime if they were precipitating out. Preliminary chemical investi-\ngations indicated that only a phase or phases containing at most\ncarbon, nitrogen, tungsten, molybdenum, columbium, chromium, and iron\nwere precipitating. Hence it is believed that manganese and silicon\nplay no part in the aging process, merely remaining in solution in the\nlattice throughout the aging process. It appears from this same\npreliminary chemical data that the carbon and nitrogen were partly\nin the form of an inert carbonitride of columbium in the solution-\ntreated material. The fraction thus present was unknown but even by\n\n$^1$Lecture given at the University of Michigan, Fall 1948.\n```", "timestamp": "2026-07-22T06:21:00.906744+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 52, "total_pages": 98, "image_filename": "19930086073_p52.jpg", "text": "50\nNACA RM A9H04\n\nLift coefficient, $C_L$\nPitching-moment coefficient, $C_m$\n\n| | | | |\n| :--- | :--- | :--- | :--- |\n| $\\bigcirc$ | $\\square$ | $\\diamond$ | $\\triangle$ |\n| 0.0 | 6.0 | 12.0 | 15.9 |\n\nAngle of sideslip, $\\beta$, deg\n\n(c) $C_L$ vs $C_m$.\n\nFigure 10.— Continued.", "timestamp": "2026-07-22T06:21:06.797263+00:00"}
{"citation_id": "19930085870", "source_url": "https://ntrs.nasa.gov/api/citations/19930085870/downloads/19930085870.pdf", "page_number": 2, "total_pages": 92, "image_filename": "19930085870_p2.jpg", "text": "NACA RM No. L9D07\nCONFIDENTIAL\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\nRESEARCH MEMORANDUM\nINVESTIGATIONS AT SUPERSONIC SPEEDS OF 22 TRIANGULAR\nWINGS REPRESENTING TWO AIRFOIL SECTIONS FOR\nEACH OF 11 APEX ANGLES\nBy Eugene S. Love\n\nSUMMARY\n\nInvestigations of two series of 11 triangular wings were conducted\nat Mach numbers of 1.62, 1.92, and 2.40 to determine the effect of\nleading-edge shape and to compare actual test values with the nonviscous\nlinear theory. The two series of wings had identical plan forms, a\nconstant thickness ratio of 8 percent, a constant location of maximum-\nthickness point of 18 percent, and a range of apex half-angles from 10°\nto 45°. The first series had an elliptical leading edge and the second\nseries, a wedge leading edge. Measurements were made of lift, drag,\npitching moment, and pressure distribution, the latter being confined\nto three wings at one Mach number.\n\nThe results indicated that the ratio of the lift-curve slope to the\ntheoretical two-dimensional lift-curve slope was, for any given ratio of\nthe tangent of the wing vertex half-angle to the tangent of the Mach\nangle ($\\tan \\epsilon/\\tan m$), relatively independent of Mach number for each\nseries; and in the case of the wedge-leading-edge wings for which the\nleading edge lies well ahead of the Mach cone, this ratio approached\nvery nearly one. For the range of vertex angles in the vicinity of the\nMach cone, the theoretical drag was in poor agreement with the test\nvalues, the test values being much lower. Except for cases with the\nMach cone well behind the leading edge, the elliptical-leading-edge\nconfiguration gave lower minimum drag. Any leading-edge suction achieved\nby the elliptical-leading-edge wings was evidently of such magnitude as\nto be overshadowed by other effects. The largest value of maximum lift-\ndrag ratio was obtained by the elliptical-leading-edge configuration.\nBoth series of wings showed a forward travel of the center of pressure\nwith increase in aspect ratio. Schlieren photographs, liquid-film tests,\nand pressure distributions indicated that the shocks arising on the wing\nsurfaces, the boundary-layer transition lines, and the steep adverse\npressure gradients were practically coincident.\n\nIt was concluded that, for triangular wings of this thickness ratio,\nthe aerodynamic gains experienced by the elliptical-leading-edge wings\nas compared with the wedge-leading-edge wings were not a result of any\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:21:11.404861+00:00"}
{"citation_id": "19930083221", "source_url": "https://ntrs.nasa.gov/api/citations/19930083221/downloads/19930083221.pdf", "page_number": 20, "total_pages": 47, "image_filename": "19930083221_p20.jpg", "text": "18\nNACA TN No. 1824\n\n$$\n\\frac{\\Delta p}{q} = \\frac{8\\alpha}{\\pi(1+M_0)} \\sqrt{\\frac{t-x}{x+M_0 t}} + \\frac{4\\alpha}{\\pi M_0} \\left[ \\arcsin \\frac{2x-t(1-M_0)}{(1+M_0)t} + \\arcsin \\frac{2(c_0-x)-t(1+M_0)}{t(1-M_0)} \\right] \\quad (27c)\n$$\n\nRegion IV (between lines $x=t$, $x=c_0-M_0 t$, and $x=c_0-t$)\n\n$$\n\\frac{\\Delta p}{q} = \\frac{8\\alpha}{\\pi M_0} \\arcsin \\sqrt{\\frac{x-c_0+M_0 t}{t(M_0-1)}} \\quad (27d)\n$$\n\nRegion V (between lines $x=-t + \\frac{2c_0}{1+M_0}$, $x=c_0-M_0 t$, and $t = \\frac{c_0}{1-M_0}$)\n\n$$\n\\frac{\\Delta p}{q} = \\frac{16\\alpha}{\\pi^2(1+M_0)} \\sqrt{\\frac{t-x}{x+M_0 t}} \\left[ \\frac{\\pi}{2} - EF(\\psi, k') - KE(\\psi, k') + KF(\\psi, k') \\right]\n$$\n\n$$\n- \\frac{2\\alpha}{\\pi M_0} \\arcsin \\frac{x}{t} + \\frac{4\\alpha}{\\pi M_0} \\arcsin \\frac{2x-t(1-M_0)}{(1+M_0)t} + \\frac{32\\alpha K}{\\pi^2(1+M_0)} \\sqrt{\\frac{c_0-x-M_0 t}{(1-M_0^2)(x+t)}}\n$$\n\n$$\n- \\frac{2\\alpha}{\\pi M_0} \\arcsin \\frac{2c_0-t(1+M_0)}{t(1+M_0)} + N_3 \\quad (27e)\n$$\n\nwhere\n\n$$\nk' = \\sqrt{1-k^2}\n$$\n\n$$\nk = \\sqrt{1 - \\frac{2c_0}{(t+x)(1+M_0)}}\n$$\n\n$$\n\\psi = \\arcsin \\sqrt{\\frac{x+M_0 t}{c_0}}\n$$", "timestamp": "2026-07-22T06:21:13.109958+00:00"}
{"citation_id": "19930085485", "source_url": "https://ntrs.nasa.gov/api/citations/19930085485/downloads/19930085485.pdf", "page_number": 20, "total_pages": 26, "image_filename": "19930085485_p20.jpg", "text": "```markdown\n18\n\nCONFIDENTIAL\n\nPitching-moment\ncoefficient, $C_m$\nAngle of attack, $\\alpha$, deg\nLift coefficient, $C_L$\nDrag coefficient, $C_D$\n\nCONFIDENTIAL\n\nPitching-moment\ncoefficient, $C_m$\nAngle of attack, $\\alpha$, deg\nLift coefficient, $C_L$\nDrag coefficient, $C_D$\n\nFigure 10.- Aerodynamic characteristics of the model\nwith circular-arc contour aileron. $\\delta_a = 0^\\circ$.\nCONFIDENTIAL\n\nFigure 11.- Aerodynamic characteristics of the model\nwith flat-sided aileron. $t = 0.50$; $\\delta_a = 0^\\circ$.\nCONFIDENTIAL\n\nNACA RM NO. L58C02\n```", "timestamp": "2026-07-22T06:21:27.058414+00:00"}
{"citation_id": "19930082646", "source_url": "https://ntrs.nasa.gov/api/citations/19930082646/downloads/19930082646.pdf", "page_number": 19, "total_pages": 37, "image_filename": "19930082646_p19.jpg", "text": "18\nNACA TN 1980\n\nWarped forebody and extended afterbody ———\nBasic forebody and basic afterbody — — — —\n\n<!-- Image (183, 129, 736, 569) -->\n\n(a) Lower-limit porpoising.\n\n<!-- Image (183, 613, 736, 786) -->\n\n(b) Upper-limit porpoising.\nFigure 7.- Comparison of maximum amplitude\nof porpoising between basic and modified\nhulls.", "timestamp": "2026-07-22T06:21:28.616867+00:00"}
{"citation_id": "19930082703", "source_url": "https://ntrs.nasa.gov/api/citations/19930082703/downloads/19930082703.pdf", "page_number": 28, "total_pages": 28, "image_filename": "19930082703_p28.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:21:30.887647+00:00"}
{"citation_id": "19930082592", "source_url": "https://ntrs.nasa.gov/api/citations/19930082592/downloads/19930082592.pdf", "page_number": 35, "total_pages": 50, "image_filename": "19930082592_p35.jpg", "text": "Page intentionally left blank\n\nPage intentionally left blank", "timestamp": "2026-07-22T06:21:31.891178+00:00"}
{"citation_id": "19930085859", "source_url": "https://ntrs.nasa.gov/api/citations/19930085859/downloads/19930085859.pdf", "page_number": 13, "total_pages": 31, "image_filename": "19930085859_p13.jpg", "text": "NACA RM No. 19B25\n\n4.53\n1.0\n4.3\n1.60\n.76\n.40\n\nCenter line of balance\nnormal to bump surface\n\n1/10. Max\n2.50\n\nEnd plate used with\nfloating tail in fuselage\n\n1.25 R.\n\n0 1 2\nScale, inches\n\nNACA\n\nEnd plate used with floating\ntails 1 inch from chord plane at $\\alpha=0^\\circ$\n\nFigure 2.- Details of free-floating tail mounted in fuselage of a model with $35^\\circ$ sweptback wing,\naspect ratio 4, taper ratio 0.6, and NACA 65A006 airfoil.\n\n11", "timestamp": "2026-07-22T06:21:32.464585+00:00"}
{"citation_id": "19930082546", "source_url": "https://ntrs.nasa.gov/api/citations/19930082546/downloads/19930082546.pdf", "page_number": 32, "total_pages": 65, "image_filename": "19930082546_p32.jpg", "text": "NACA TN No. 1870\n31\n\nBlade-width ratio, b/D, and blade-thickness ratio, h/b\n\nBlade angle, $\\beta$, deg\n\n| $r/R_t$ | 0 | .1 | .2 | .3 | .4 | .5 | .6 | .7 | .8 | .9 | 1.0 |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| **Left Axis** | 0 | .02 | .04 | .06 | .08 | .10 | .12 | .14 | .16 | .18 | .20 | .22 | .24 | .26 | .28 |\n| **Right Axis** | 0 | 2 | 4 | 6 | 8 | 10 | 12 | 14 | 16 | 18 | 20 | 22 | 24 | 26 | 28 |\n\n[Figure: Graph showing three curves labeled $\\beta$, h/b, and b/D plotted against $r/R_t$. The NACA logo is visible in the bottom right corner of the plot area.]\n\n(a) Clark Y propeller.\n\nFigure 3.- Blade-form curves for test propellers.", "timestamp": "2026-07-22T06:21:32.627882+00:00"}
{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 20, "total_pages": 72, "image_filename": "19930085491_p20.jpg", "text": "NACA RM No. A8J04 CONFIDENTIAL 19\n\ngradual in the range of lift coefficients below 0.09. These asymmetrical separation areas on the upper and lower surfaces effectively reduced the wing angle of attack over the sections affected and are therefore undoubtedly responsible for the low value of $dC_L/d\\alpha$ obtained experimentally in this range.\n\nAbove $C_L = 0.09$, there was an abrupt change in the liquid-film pattern on the upper surface. (See figs. 13(c) and (d).) The line of laminar separation moved from approximately midchord to the region near the leading edge because of the influence of the highly adverse pressure gradients shown in figure 5. After separating, the boundary layer reattached as a turbulent flow on the inboard sections (where the pressure gradients behind the leading edge were not severely adverse). The elimination of the separation near the trailing edge on these sections increases the effective angle of attack as is indicated by the increase in lift-curve slope from 0.038 to 0.045. This change suggests that, if the flow would reattach on the tip sections, the experimental value of $dC_L/d\\alpha$ would closely approach that of 0.051 predicted by the inviscid linear theory.\n\nAt $C_L = 0.21$ and 0.28 (figs. 13(c) and (d)), the line of laminar separation is very close to the leading edge except for a small length near the two-thirds semispan location. The rearward displacement of the line on these sections may be due to a localized supercritical flow based on the velocity component ($M_n = 0.69$) and wing section (similar to an NACA 0012 section) perpendicular to the wing leading edge. This condition would displace the minimum pressure point and consequently the leading-edge adverse pressure gradient region rearward as was observed in the tests of reference 19. However, no reason for the restriction of this flow alteration to only a part of the wing is apparent at the present time.\n\nMinimum drag coefficient.— The value of $C_{Dmin}$ obtained experimentally at a Reynolds number of 0.62 million is 0.0175 which is somewhat greater than the theoretical value of 0.0133. Several factors may contribute to this discrepancy, the most important being the increased pressure drag component included in the experimental value which results from laminar boundary-layer separation. The previously discussed liquid-film result of figure 13(b) shows that a large separated area exists at minimum drag ($C_L = 0$). A similar condition observed in the tests of reference 20 with a swept-back-wing pressure-distribution model revealed that the pressures behind the line of separation are constant (as in subsonic flow) and more negative than indicated by theory, thereby increasing the experimental pressure drag increment. The effects of wing-fuselage interference and\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:21:35.471148+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 36, "total_pages": 78, "image_filename": "19930082483_p36.jpg", "text": "34\nNACA TN No. 1807\n\nBy dimensional analysis, it can be seen that only the driving-fluid losses vary as $K_{III}P_1N$ and therefore shaft or tip leakage losses cannot be made to correlate the power differentials observed. This fact rules out the possibility that the power differences can be accounted for by errors in the magnitudes of the shaft losses as calculated by the empirical equations presented.\n\nNozzle and rotor-blading losses. - The nozzle and rotor-blading aerodynamic losses are presented for the range of admissions in figure 9(e). These losses may be seen in the figure to be directly proportional to the degree of admission. By comparison with figures 8(a) and 8(b), it can be seen that the nozzle and rotor-blading aerodynamic losses are minimal at the speed range corresponding to the peak operating efficiencies of the turbine.\n\nLoss relations. - Represented in figure 10 is a quantitative analysis of the manner in which the ideal power per pound (specific ideal power) is dissipated by the various losses at 360° (full), 180°, and 120° admissions. All the quantities are corrected to sea-level conditions. Because the ideal power per pound of driving fluid is independent of the degree of admission, comparison of the magnitudes of these losses at the various degrees of admission enables a direct evaluation of the relative effect of each of the losses on the net power output with reduction of the active nozzle arc.\n\nFrom figure 10 it can be seen that the specific rotor-tip leakage loss is constant for all degrees of admission and for all rotor speeds.\n\nThe bearing power loss of the turbine is independent of the percentage of active nozzle arc; therefore the specific bearing loss is inversely proportional to the weight flow and hence to the degree of admission. As indicated earlier, the bearing losses are proportional approximately to the 1.4 power of the rotor speed.\n\nAt any given degree of admission, the pumping losses are proportional to the cube of the rotor speed, whereas the driving-fluid losses are directly proportional to rotor speed.\n\nThe pumping and driving-fluid losses were shown, in the ANALYSIS section, to be proportional to the inactive nozzle arc and in figure 10 may be seen to vary in this manner.\n\nThe nozzle and rotor aerodynamic losses per pound of driving fluid are independent of the degree of admission and consequently are the same for the three admissions presented in figure 10.", "timestamp": "2026-07-22T06:21:36.536138+00:00"}
{"citation_id": "19930085862", "source_url": "https://ntrs.nasa.gov/api/citations/19930085862/downloads/19930085862.pdf", "page_number": 9, "total_pages": 60, "image_filename": "19930085862_p9.jpg", "text": "NACA RM No. L9A07\n7\n\nRESULTS AND DISCUSSION\n\nAileron Characteristics\n\nRolling characteristics.- The basic aileron data are shown in figures 5 to 10. Several representative crossplots of $C_l$ and $C_{h_a}$ against aileron deflection are presented in figures 11 and 12, respectively. In order to show the aileron effectiveness $C_{l_\\delta}$ determined for a small range of aileron deflections through $\\delta_a = 0^\\circ$, the variation of $C_{l_\\delta}$ with angle of attack is presented in figure 13 for the several flap arrangements tested.\n\nIt can be seen that $C_{l_\\delta}$ has a value of approximately 0.00100 at low angles of attack for all flap configurations. The value of $C_{l_\\delta}$ of 0.00105 obtained at $\\alpha = 0^\\circ$ for the wing without flaps was about the same as that (0.00102) determined by means of the charts of reference 3 and reduced by $\\cos^2\\Lambda$ to account for the effects of sweep. The rate of change of rolling-moment coefficient with aileron deflection $C_{l_\\delta}$ for the wing without flaps and for the wing equipped with the extensible leading-edge flaps and fences remained approximately constant as the angle of attack was increased up to that corresponding to $0.85C_{L_{max}}$. The lift coefficients corresponding to $0.85C_{L_{max}}$ are used herein as a basis for comparison since they might be considered as representative of those for the landing-approach condition. The addition of the split flaps to the plain wing resulted in a 25-percent decrease in $C_{l_\\delta}$ at $0.85C_{L_{max}}$. Furthermore at $C_{L_{max}}$, the value of $C_{l_\\delta}$ was only 0.00040. The further addition of the leading-edge flaps did not prevent the large reduction caused by the split flaps but, with the leading-edge flaps in combination with the stall-control fences, the values of $C_{l_\\delta}$ were comparable to those obtained for the plain wing. The aileron effectiveness of the wing equipped with the drooped-nose and split flaps and fences was approximately the same as that for the configuration with the extensible leading-edge and split flaps and fences.", "timestamp": "2026-07-22T06:21:37.995619+00:00"}
{"citation_id": "19930082918", "source_url": "https://ntrs.nasa.gov/api/citations/19930082918/downloads/19930082918.pdf", "page_number": 22, "total_pages": 62, "image_filename": "19930082918_p22.jpg", "text": "NACA TN 1940\n\ndisregarding entirely the carbon, nitrogen, and columbium content (which is most probably not correct) the conclusions as to the period of the strains involved are still warranted.\n\nIt appears that a logical way to increase the creep strength of low-carbon N-155 would be to add elements with either larger atomic radius than molybdenum, columbium, and tungsten or smaller atomic radius than carbon and nitrogen. Those suggested in table 7 are of the former type. Only boron, among elements of small atomic radius, appears promising at the moment. The usefulness of boron seems to be substantiated by the high creep strength of boron modifications of low-carbon N-155 recently investigated by the Union Carbide and Carbon Research Laboratories, Inc.\n\nTable 8 lists elements which have atomic radii similar to molybdenum; tungsten, and columbium and which might be expected to act as substitutes. The elements suggested in these tables are not the only metallic elements with the desired atomic size but are thought to be the most promising. With the exception of aluminum and silver, all the proposed additions or substitutions are transition elements. Some evidence (see reference 10) is available which indicates that such transition elements have abnormally high binding energies. Aluminum and silver are considered only on the basis of having the desired apparent atomic size.\n\nIt is of course desirable that the materials suggested in tables 7 and 8 go readily into solution at some relatively high (solution-treating) temperature, and either stay in solution (or the nucleated state) or precipitate very slowly at the lower temperatures of service. It is possible that any or all of the proposed elements may show an increased tendency to come out of supersaturated solid solution and thus cause rapid loss of creep strength or show a wrong or poorer type of solubility-temperature characteristic. Other metallurgical characteristics (i.e., ductility) must necessarily be satisfied before such modifications could be considered satisfactory.\n\nFactors Controlling Rupture Strength and Ductility\n\nInspection of figures 16 and 19 shows in general that aging at either $1400^\\circ$ or $1600^\\circ$ F resulted in a progressive increase in very short-time rupture strength. For a given increase in rupture strength, less aging time was required at $1600^\\circ$ F than at $1400^\\circ$ F. For somewhat longer rupture times, the relative increase in the rupture strength of the unaged material was quite striking. In fact, the solution-treated stock rapidly became equal in strength to the aged material at the increased rupture times. Previous experience has indicated that the rupture strength of the unaged material will actually exceed the rupture strength of the aged material when considering longer rupture times than were used in this investigation.", "timestamp": "2026-07-22T06:21:43.177194+00:00"}
{"citation_id": "19930085548", "source_url": "https://ntrs.nasa.gov/api/citations/19930085548/downloads/19930085548.pdf", "page_number": 17, "total_pages": 46, "image_filename": "19930085548_p17.jpg", "text": "16\nNACA RM No. E8L30\n\nrise apparently had little bearing on the performance of the\ncylinder when the limitations were maximum cylinder pressure and\nexhaust-gas temperature.\n\n6. An analysis of the data showed that inadequate cylinder\ncharging (charge air insufficient to maintain the cylinder fuel-\nair ratio somewhat below stoichiometric) limited the manifold pres-\nsure at which a gas-generator engine using this cylinder may oper-\nate. The analysis indicates, however, that if the cylinder charging\nwas adequate, only small increases in allowable manifold pressure\nwould accompany further improvement in charging efficiency.\n\n7. The unusual operating conditions had no harmful effects on\nthe mechanical operation of the engine; the operation was quite\nsmooth because of the low rate of combustion-pressure rise in the\ncylinder.\n\nLewis Flight Propulsion Laboratory,\nNational Advisory Committee for Aeronautics,\nCleveland, Ohio.", "timestamp": "2026-07-22T06:21:43.274923+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 53, "total_pages": 98, "image_filename": "19930086073_p53.jpg", "text": "```markdown\nNACA RM A59E04\n\nLift coefficient, $C_L$\n\n| $\\beta$, deg |\n| :--- |\n| $\\circ$ 0.0 |\n| $\\square$ 6.0 |\n| $\\diamond$ 12.0 |\n| $\\triangle$ 15.9 |\n\nRolling-moment coefficient, $C_l$\n\nYawing-moment coefficient, $C_n$\n\nSide-force coefficient, $C_Y$\n\n(d) $C_L$ vs $C_l$, $C_n$ and $C_Y$.\n\n[Figure: Three graphs plotting Lift coefficient ($C_L$) against Rolling-moment coefficient ($C_l$), Yawing-moment coefficient ($C_n$), and Side-force coefficient ($C_Y$) for various sideslip angles ($\\beta$).]\n\nFigure 10.— Concluded.\n\nNACA\n\n51\n```", "timestamp": "2026-07-22T06:21:43.459721+00:00"}
{"citation_id": "19930085847", "source_url": "https://ntrs.nasa.gov/api/citations/19930085847/downloads/19930085847.pdf", "page_number": 15, "total_pages": 32, "image_filename": "19930085847_p15.jpg", "text": "NACA RM A57D04\nCONFIDENTIAL\nTEST\nNACA\nA-12654\nFigure 4.- Test equipment showing wing root bump installed for the tests.\nCONFIDENTIAL\n13", "timestamp": "2026-07-22T06:21:48.886186+00:00"}
{"citation_id": "19930085485", "source_url": "https://ntrs.nasa.gov/api/citations/19930085485/downloads/19930085485.pdf", "page_number": 21, "total_pages": 26, "image_filename": "19930085485_p21.jpg", "text": "NACA RM No. L8K02\n19\n\nCONFIDENTIAL\n\nPitching-moment coefficient, $C_m$\nDrag coefficient, $C_D$\nAngle of attack, $\\alpha$, deg\nLift coefficient, $C_L$\n\nM\no .597\n□ .888\n◇ .965\n\nNACA\n\nFigure 12.- Aerodynamic characteristics of the model with flat-sided aileron. $t = 1.00$; $\\delta_a = 0^\\circ$.\nCONFIDENTIAL", "timestamp": "2026-07-22T06:21:49.901284+00:00"}
{"citation_id": "19930082646", "source_url": "https://ntrs.nasa.gov/api/citations/19930082646/downloads/19930082646.pdf", "page_number": 20, "total_pages": 37, "image_filename": "19930082646_p20.jpg", "text": "NACA TN 1980\n19\n\nWarped forebody and extended afterbody\nBasic forebody and basic afterbody\n\nFull up elevators\n\nUnstable\nStable\n\nElevator deflection, deg\n-30\n-25\n-20\n-15\n-10\n-5\n0\n\nCenter of gravity, percent M.A.C.\n18\n20\n22\n24\n26\n28\n30\n32\n34\n36\n38\n40\n\nFigure 8.- Variation of center-of-gravity limits of stability, with elevator deflection, for 2° amplitude of porpoising.", "timestamp": "2026-07-22T06:21:53.178716+00:00"}
{"citation_id": "19930085870", "source_url": "https://ntrs.nasa.gov/api/citations/19930085870/downloads/19930085870.pdf", "page_number": 3, "total_pages": 92, "image_filename": "19930085870_p3.jpg", "text": "2\nCONFIDENTIAL\nNACA RM No. L9D07\n\nappreciable realization of leading-edge suction but the favorable\neffect of the gentle or easy curvature of the ridge line common to\nthe elliptical-leading-edge shape.\n\nINTRODUCTION\n\nThe wing of triangular plan form has received much attention lately\nas a possible efficient wing for supersonic flight. Reference 1 pointed\nout that L/D ratios of configurations employing sweepback as outlined in\nreference 2 could be improved upon provided the wing lay well within the\nMach cone. Later, the theory of small disturbances was applied to the\ncase of finite aspect ratios (references 3 and 4) and a theory was developed\nfor computing the L/D ratios for practical configurations. Recently,\nseveral different authors have developed methods independently for calcu-\nlating the lift and drag of triangular and sweptback wings (references 5\nto 9).\n\nAn experimental investigation of triangular wings was undertaken in\n1945 in the Langley model supersonic tunnel, forerunner of the present\nLangley 9-inch supersonic tunnel (reference 10). These tests were pri-\nmarily a preliminary investigation of flat-plate triangular wings (thick-\nness ratio, approx. $\\frac{1}{2}$ percent) to determine the limits of Jones' slender-\nwing theory and to ascertain the highest values of maximum L/D. In the\nrange of low aspect ratios the results confirmed Jones' original theory\nbut exhibited some unusual breaks when the leading edge lay near the Mach\ncone. In addition, the tests showed that the center of area of the wing\nand the center of pressure were coincident. Although the absolute values\nof the drag were in doubt, as stated by the authors, a maximum L/D of\nabout 7 was obtained.\n\nIn order to further the study of triangular-wing characteristics at\nsupersonic speeds, a series of tests was conducted on three triangular-\nwing models at a Mach number of 1.53 in the Ames 1- by 3-foot supersonic\ntunnel (reference 11). The models had a thickness ratio of 5 percent,\nan aspect ratio of 2, and were designed to study the effects of variation\nin thickness distribution and camber with the wing apex both leading and\ntrailing. These tests indicated that, for the apex-forward condition,\nthe highest value of maximum L/D is obtained with the maximum-thickness\npoint well forward and a slightly rounded leading edge. With the maximum-\nthickness point at 20 percent, maximum L/D was increased from 6.4 for\nthe sharp leading edge to 6.8 for the rounded leading edge, indicating\nthe possible existence of leading-edge suction predicted by theory. The\ndrag relief from rounding the leading edge fell short of that predicted\nfrom theoretical considerations.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:21:55.570478+00:00"}
{"citation_id": "19930082546", "source_url": "https://ntrs.nasa.gov/api/citations/19930082546/downloads/19930082546.pdf", "page_number": 33, "total_pages": 65, "image_filename": "19930082546_p33.jpg", "text": "32\nNACA TN No. 1870\n\nBlade-width ratio, b/D, and blade-thickness ratio, b/b\n\nBlade angle, $\\beta$, deg\n\n<!-- Image (109, 151, 831, 827) -->\n\n(b) NACA 4-(3)(06,3)-06 propeller.\nFigure 3.- Continued.", "timestamp": "2026-07-22T06:21:56.245024+00:00"}
{"citation_id": "19930082592", "source_url": "https://ntrs.nasa.gov/api/citations/19930082592/downloads/19930082592.pdf", "page_number": 36, "total_pages": 50, "image_filename": "19930082592_p36.jpg", "text": "NACA TN 1914\n35\n\n[Figure: Micrograph showing a circular cross-section with three distinct layers. The top layer is labeled \"Bakelite\". The middle layer is labeled \"Oxide\". The bottom layer is labeled \"Unoxidized ceramal\". A NACA logo with \"C-22911\" and \"2-7-49\" is in the bottom right corner of the figure.]\n\nFigure 12. - Oxidation zone of 30-percent-tungsten - titanium carbide ceramal. Under low magnification oxide penetration is regular and line of demarcation between oxide and ceramal is even. Temperature, $1785^\\circ$ F; time at temperature, 30 hours; unetched; magnification, X50.", "timestamp": "2026-07-22T06:21:56.730055+00:00"}
{"citation_id": "19930085890", "source_url": "https://ntrs.nasa.gov/api/citations/19930085890/downloads/19930085890.pdf", "page_number": 1, "total_pages": 26, "image_filename": "19930085890_p1.jpg", "text": "NACA RM No. E9C11\n\nRM No. E9C11\n\n[Figure: NACA logo with wings]\n\nRESEARCH MEMORANDUM\n\nEXPERIMENTAL INVESTIGATION OF LIQUID DIBORANE - LIQUID OXYGEN\nPROPELLANT COMBINATION IN 100-POUND-THRUST ROCKET ENGINE\n\nBy William H. Rowe, Paul M. Ordin, and John M. Diehl\n\nLewis Flight Propulsion Laboratory\nCleveland, Ohio\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\nWASHINGTON\n\nMay 9, 1949\nDeclassified March 19, 1957", "timestamp": "2026-07-22T06:21:58.013583+00:00"}