PDF/ocr_dataset_1018.jsonl

{"citation_id": "19930092013", "source_url": "https://ntrs.nasa.gov/api/citations/19930092013/downloads/19930092013.pdf", "page_number": 18, "total_pages": 21, "image_filename": "19930092013_p18.jpg", "text": "```markdown\n14\nREPORT 948—NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nthe approach condition, although the period-damping com-\nbination produced by this dihedral (fig. 19) was well within\nthe unsatisfactory area as defined by references 2 and 3. It\nwould appear, then, that the period-damping requirements of\nreferences 2 and 3 are too severe in this case.\n\nThe maximum tolerable effective dihedral in the cruising\nand high-speed conditions is seen in figure 18 to be about 22°.\nFigure 19 shows that, for the cruising condition and the high-\nspeed condition, the good configurations satisfied the require-\nments of references 2 and 3, but the intolerable configurations\ndid not. With 24.4° effective dihedral in the cruising con-\nfiguration, the oscillations set up in rough air were difficult\nto control. Some of the pilots attributed this difficulty to the\nshort period in combination with the low damping. The\nmeasurements showed the natural period to be about 3.0\nseconds for $\\Gamma_e=24.4^\\circ$ in the cruising condition and 3.6\nseconds for $\\Gamma_e=28.4^\\circ$ in the landing-approach condition.\nThe 3.0-second period in the cruising condition was intoler-\nable, and the 3.6-second period in the approach condition was\ntolerable—yet the latter was oscillatory unstable. When\npresented with the results of the measurements, the pilots\nagreed that they probably could not detect the difference\nbetween a 3.0-second period and a 3.6-second period, at least\nnot definitely enough to enable them to classify one as toler-\n\n<!-- Image (51, 78, 500, 653) -->\n\n(a) $\\Gamma_e$ 24.4°.\n(c) $\\Gamma_e$ 12.9°.\n(b) $\\Gamma_e$ 18.2°.\n(d) $\\Gamma_e$ 6.2° (normal airplane).\nFIGURE 17.—Time histories of typical rudder-fixed aileron rolls. Cruising condition.\n\nFigure 19 shows how the various configurations compare\nwith the period-damping requirements of references 2 and 3.\nThe data of figure 15 were used to plot time to damp to half\namplitude against period, and the points were labeled with\nthe opinions shown in figure 18 and the corresponding effec-\ntive dihedrals.\n\nThe maximum tolerable effective dihedral in the landing-\napproach condition was not reached. Although the highest\ndihedral used (28.4°) produced oscillatory instability in the\napproach condition, the oscillations were relatively easy to\ncontrol because, according to the pilots, the period was long\nand the rolling velocities were not too high. In fact, the\npilots considered an effective dihedral of 22.7° to be good in\n\n<!-- Image (527, 439, 950, 906) -->\n\nFIGURE 18.—Variation of pilots' opinions of lateral-handling characteristics with effective dihedral.\n```", "timestamp": "2026-07-22T04:44:55.978020+00:00"}
{"citation_id": "19930086015", "source_url": "https://ntrs.nasa.gov/api/citations/19930086015/downloads/19930086015.pdf", "page_number": 53, "total_pages": 54, "image_filename": "19930086015_p53.jpg", "text": "52\nCONFIDENTIAL\nNACA RM A9E24\n\nTheory\nExperiment : $\\circ \\Delta \\alpha = 5.35^\\circ$ (Model Vertical)\n$\\Delta \\alpha = 370^\\circ$ (Model Horizontal)\n$\\square \\Delta \\alpha = 5.74^\\circ$ (Model Horizontal)\n\nLoading coefficient\nper unit angle of attack, $h_2$, per deg\n.18\n.16\n.14\n.12\n.10\n.08\n.06\n.04\n.02\n0\n\n20 40 60 80 100\nPercent of local chord\n\n(e) M = 1.70.\nFigure 14- concluded.\n\nCONFIDENTIAL\nNACA-Langley - 9-15-49 - 375", "timestamp": "2026-07-22T04:44:59.157681+00:00"}
{"citation_id": "19930085842", "source_url": "https://ntrs.nasa.gov/api/citations/19930085842/downloads/19930085842.pdf", "page_number": 93, "total_pages": 104, "image_filename": "19930085842_p93.jpg", "text": "NACA RM L9C29\n89\n\nPropeller advance-diameter ratio, V/nD\nLift coefficient, $C_L$\nTorque coefficient, $Q_c$\nResultant-drag coefficient, $C_{DR}$\n\n[Figure: Graph showing three curves plotted against Torque coefficient ($Q_c$) on the x-axis. The top curve represents Propeller advance-diameter ratio ($V/nD$), the middle curve represents Resultant-drag coefficient ($C_{DR}$), and the bottom curve represents Lift coefficient ($C_L$).]\n\nNATIONAL ADVISORY\nCOMMITTEE FOR AERONAUTICS\n\n(c) $\\alpha_u = 14^\\circ$.\nFigure 44.- Continued.", "timestamp": "2026-07-22T04:44:59.711057+00:00"}
{"citation_id": "19930082542", "source_url": "https://ntrs.nasa.gov/api/citations/19930082542/downloads/19930082542.pdf", "page_number": 4, "total_pages": 53, "image_filename": "19930082542_p4.jpg", "text": "NACA TN No. 1867\n3\n\nThis report is based on a systematic study of the effects of heat treatment and hot-cold-work on the properties of low-carbon N-155 alloy. Bar stock from one heat of the alloy was subjected to 67 different treatments. The criterions used for evaluating the effect of the treatments were physical properties at room temperature and rupture test characteristics at 1200° F.\n\nLow-carbon N-155 was one of the outstanding alloys which evolved from the initial work in this field by the NACA. The alloy was known to have a large range in properties depending on the forging and heat-treating conditions used in its production. (See reference 4.) For these reasons and because of the potential usefulness of the alloy at high temperatures it was selected as an acceptable alloy for fundamental studies. It is expected, however, that the principles developed from the work will be applicable to other alloys with considerably less experimental work.\n\nThe criterions used were selected primarily because they were believed to be indicative of the suitability of alloys for service in the discs of rotors for gas turbines. The types of heat treatment and processing which should be used for producing such discs have not been clear and it seemed worth while to start the investigations on the effect of heat treatment and other processing variables on disc criterions.\n\nThis work was conducted at the University of Michigan under the sponsorship and with the financial assistance of the National Advisory Committee for Aeronautics.\n\nTEST MATERIAL\n\nThe following information concerning the low-carbon N-155 alloy used in this investigation was supplied by the alloy producers:\n\nAlloy producers:\n\nA special heat was melted by the Union Carbide and Carbon Research Laboratories, Inc.\n\nThe ingot was processed by the Universal-Cyclops Steel Corporation\n\nHeat designation:\n\nLot 30276", "timestamp": "2026-07-22T04:45:00.135759+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 12, "total_pages": 99, "image_filename": "19930082511_p12.jpg", "text": "```markdown\n10\nNACA TN No. 1826\n\nsignificance, provided the gradients are determined at reasonable\ndistances from it. In general, boundary interference at the vortex itself\ncannot be found directly by this method, but may be determined by\ninterpolation between or extrapolation from neighboring points.\n\nTwo-dimensional closed-open tunnel.- For simplification of the\nnomenclature, the open tunnel with closed upstream region but without a\nclosed exit is designated the closed-open tunnel. The open tunnel with\nclosed upstream and downstream regions is designated the closed-open-\nclosed tunnel.\n\nFigure 7(a) illustrates the setup for a two-dimensional closed-open\ntunnel with a vortex on its center line. Shaded lines indicate\ninsulating boundaries, where $\\frac{\\partial \\phi}{\\partial n} = 0$, and heavy unshaded lines indicate\nmetal boundaries on which $\\phi$ is constant. The upstream closed portion\nshould be so long that the potential is essentially uniform at its\nupstream end; the length indicated on the figure should suffice. The\nopen region should similarly be so long that the vertical flow between\nthe vortex strips and the boundary strips no longer changes with distance\ndownstream; again, the length indicated on the figure should suffice.\nFrom the condition of velocity continuity at the entrance lips and the\nfact that $\\frac{\\partial \\phi}{\\partial x}$ is zero along the free boundaries, it follows that $\\frac{\\partial \\phi}{\\partial x}$\nmust be zero at the edges of the two closed boundaries. The potentials\non the two free boundaries must, therefore, be adjusted until the differ-\nence between the potential of each and the potential of a thin feeler\nelectrode just upstream of its edge is zero. For the symmetrical condi-\ntion shown, the single variable voltage source indicated will provide\nzero $\\frac{\\partial \\phi}{\\partial x}$ at both edges simultaneously.\n\nFigure 7(b) illustrates the setup for the two-dimensional closed-\nopen tunnel with the vortex in an off-center position. A single variable\nvoltage source across the two free boundaries is now no longer capable\nof simultaneously satisfying the continuity condition at both edges, so\nthat an additional variable voltage source and an upstream electrode\nare required. The current in the closed part of the tunnel flowing into\nthis upstream electrode corresponds to an upstream perturbation velocity.\nThis upstream perturbation velocity constitutes the previously mentioned\ndifference between the velocity far upstream in the closed part and the\nvelocity on the free surface. The concept here is slightly at variance\nwith previous discussion, which considered a perturbation velocity along\nthe free surface, with the far upstream velocity appearing as the\nundisturbed velocity U. As the analogy is set up, however, no pertur-\nbation velocity may appear along the free surfaces because they are at\nconstant potential; hence, the total velocity on the free surfaces must\nbe considered as the undisturbed velocity U and any difference between\nthis velocity and the velocity far upstream appears as an upstream\nperturbation velocity. As appears in part II, this viewpoint is also\nfound convenient in the analytical solution of these problems.\n```", "timestamp": "2026-07-22T04:45:00.494983+00:00"}
{"citation_id": "19930082476", "source_url": "https://ntrs.nasa.gov/api/citations/19930082476/downloads/19930082476.pdf", "page_number": 19, "total_pages": 41, "image_filename": "19930082476_p19.jpg", "text": "NACA TN No. 1801\n17\n\nCHART 2.- SPIN CHARACTERISTICS OF MODEL WITH MASS DISTRIBUTION INCREASED ALONG\nTHE FUSELAGE (LINKED RUDDER AND AILERON CONTROLS)\n\n$$ \\left[ \\frac{I_x - I_y}{mb^2} = -49 \\times 10^{-4}; \\mu = 5.29 \\text{ (loading 2 in table II and point 2 in fig. 4); right erect spins} \\right] $$\n\nWheel setting\n\n| | Left | | | | Right | | Full | |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| | | 0 | $\\frac{1}{4}$ | $\\frac{1}{2}$ | | | | |\n| **Up** | | | | | **a, b** | | **b** | |\n| | 20 | No spin | | | 18<br>30 | 9D<br>18D | 13<br>29 | 10D |\n| | | | | | 142 | 0.57 | 145 | 0.50 |\n| | | | **c** | | **b** | | **d** | |\n| | 13 | No spin | | | 15<br>22 | 10D<br>16D | | |\n| | | | >212 | | 175 | 0.68 | | |\n| | | | | | **b** | | | |\n| | 5 | | | | 15<br>26 | 12D | | |\n| | | | | | 179 | 0.71 | | |\n| | | | | | | | **c** | |\n| | 0 | No spin | | | No spin | | | |\n| | | | | | | | >215 | |\n| **Down** | | | | | | | | |\n| | 12 | No spin | | | No spin | | No spin | |\n\nElevator setting, degrees\n\n$^a$Spin has a whipping motion.\n$^b$Oscillatory spin, range of values or average value given.\n$^c$Steep spin, velocity too high to permit obtaining test data.\n$^d$Steep spiral.\n\nModel values converted to corresponding full-scale values.\nU Inner wing up\nD Inner wing down\n\n| $\\alpha$ (deg) | $q$ (deg) |\n| :--- | :--- |\n| V (fps) | $\\omega$ (rps) |\n\nNACA", "timestamp": "2026-07-22T04:45:01.947781+00:00"}
{"citation_id": "19930082487", "source_url": "https://ntrs.nasa.gov/api/citations/19930082487/downloads/19930082487.pdf", "page_number": 18, "total_pages": 33, "image_filename": "19930082487_p18.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T04:45:02.210407+00:00"}
{"citation_id": "19930082617", "source_url": "https://ntrs.nasa.gov/api/citations/19930082617/downloads/19930082617.pdf", "page_number": 1, "total_pages": 58, "image_filename": "19930082617_p1.jpg", "text": "NATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\n\nTECHNICAL NOTE 1962\n\nSTRESSES IN AND GENERAL INSTABILITY OF\nMONOCOQUE CYLINDERS WITH CUTOUTS\nVII - EXPERIMENTAL INVESTIGATION OF CYLINDERS\nHAVING EITHER LONG BOTTOM CUTOUTS OR\nSERIES OF SIDE CUTOUTS\n\nBy N. J. Hoff, Bruno A. Boley, and Joseph J. Mele\nPolytechnic Institute of Brooklyn\n\n[Figure: NACA logo]\n\nWashington\nOctober 1949", "timestamp": "2026-07-22T04:45:04.282057+00:00"}
{"citation_id": "19930082614", "source_url": "https://ntrs.nasa.gov/api/citations/19930082614/downloads/19930082614.pdf", "page_number": 2, "total_pages": 36, "image_filename": "19930082614_p2.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T04:45:08.879102+00:00"}
{"citation_id": "19930082496", "source_url": "https://ntrs.nasa.gov/api/citations/19930082496/downloads/19930082496.pdf", "page_number": 10, "total_pages": 50, "image_filename": "19930082496_p10.jpg", "text": "NACA TN No. 1836\n9\n\n## RESULTS AND DISCUSSION\n\n### Tensile-Strength Evaluation\n\nThe results obtained in the short-time tensile-strength evaluation of the ceramal at $1800^\\circ$ and $2200^\\circ$ F are presented in table I. At $1800^\\circ$ F, the average short-time tensile strength obtained with specimens 3D1 and 3D5 was 33,200 pounds per square inch based upon the original test-section area. Specimen 3D1 underwent negligible oxidation; specimen 3D5 underwent a 0.2-percent reduction in area because of oxidation, which can also be considered negligible. Specimen 3D2 yielded a short-time tensile strength of 7800 pounds per square inch at $2200^\\circ$ F. This specimen, however, underwent a 12-percent reduction in area because of oxidation subsequent to a failure of supply of the helium atmosphere. On the assumption that practically all the oxidation occurred during the soaking period, the tensile evaluation was in reality conducted upon a specimen test-section area of 0.1712 square inch (measured after removal of oxide film) rather than of the 0.1955-square-inch original test-section area. The oxide scale was not considered as possessing any significant load-carrying capacity and therefore was not considered in determining fracture stresses. Correction for this change in area gave a tensile strength of 8900 pounds per square inch at $2200^\\circ$ F.\n\nSpecimen 3D3 failed during the soaking period. Specimen 3D4 was soaked 4 hours at $2300^\\circ$ F and yielded a short-time tensile strength of 12,000 pounds per square inch at $2200^\\circ$ F. Correction for oxidation gave a tensile strength of 13,200 pounds per square inch at $2200^\\circ$ F. The assumption was made that the 4-hour soaking at $2300^\\circ$ F used for specimen 3D4 would result in adequate alignment and stress relieving and would minimize any detrimental effects of the longer soaking and would thereby result in a more reliable short-time tensile-stress value. These tensile investigations had the features of a stress-rupture test; in particular, specimen 3D2 was subjected to a load of 1000 pounds per square inch for $13\\frac{1}{2}$ hours at $2300^\\circ$ F and specimen 3D4 was subjected to a load of 1000 pounds per square inch for 4 hours at $2300^\\circ$ F. No data are available for alloys with which to compare these stress-rupture values.\n\nThe short-time tensile strength of 33,200 pounds per square inch at $1800^\\circ$ F compares favorably with that of the heat-resistant alloys", "timestamp": "2026-07-22T04:45:09.924245+00:00"}
{"citation_id": "19930093773", "source_url": "https://ntrs.nasa.gov/api/citations/19930093773/downloads/19930093773.pdf", "page_number": 22, "total_pages": 47, "image_filename": "19930093773_p22.jpg", "text": "NACA RM E9G09\n\nSlip joint\nInlet-air duct\nAir flow\nStation r, venturi throat\nStation 1, engine inlet\nStation 2, compressor inlet\nCompressor\nStation 2a, compressor stator stages\nStation 3, compressor outlet\nCombustion chamber\nTurbine\nStation 4, turbine inlet\nStation 5, turbine outlet\nStation 6, turbine outlet (10 1/2 in. behind turbine)\nStation 7, 1 in. upstream of exhaust-nozzle outlet\n\nNACA\n\n| Station | Total pressure tubes | Static-pressure tubes | Wall static-pressure orifices | Thermo-couples |\n|---|---|---|---|---|\n| r | 12 | 4 | 4 | 6 |\n| 1 | 40 | 4 | 0 | 8 |\n| 2 | 24 | 0 | 4 | 0 |\n| 2a | 0 | 0 | 13 | 0 |\n| 3 | 20 | 0 | 4 | 6 |\n| 4 | 5 | 0 | 0 | 0 |\n| 5 | 0 | 0 | 0 | 8 |\n| 6 | 30 | 0 | 2 | 33 |\n| 7 | 18 | 5 | 4 | 14 |\n\nFigure 2. - Cross section of turbojet-engine installation showing sections at which instrumentation was installed.\n\n21", "timestamp": "2026-07-22T04:45:11.343475+00:00"}
{"citation_id": "19930082447", "source_url": "https://ntrs.nasa.gov/api/citations/19930082447/downloads/19930082447.pdf", "page_number": 20, "total_pages": 24, "image_filename": "19930082447_p20.jpg", "text": "18\nNACA TN No. 1775\n\n<!-- Image (109, 188, 849, 754) -->\n\nFigure 4.-Variation of draft coefficient with approach parameter.", "timestamp": "2026-07-22T04:45:14.821693+00:00"}
{"citation_id": "19930085965", "source_url": "https://ntrs.nasa.gov/api/citations/19930085965/downloads/19930085965.pdf", "page_number": 59, "total_pages": 67, "image_filename": "19930085965_p59.jpg", "text": "58\n\nCalculated Measured\na.c. a.c.\nQ O\n∇ ∇\nΔ Δ\n\nTotal\nair gap\n(in.)\nSAE 1020 steel 0.04\nArmco Magnetic .04\nIngot Iron\nHipernik .04\n\nVoltage per cm of blade width\n0\n.02\n.04\n.06\n.08\n\n0 100 200 300 400\nEffective ampere turns\n\nNACA\n\nFigure 15. - Variation of voltage of self-induction per centimeter of blade width with\nampere turns. Vane, $\\frac{45}{8}$ by 1 by $\\frac{1}{8}$ inch (Hipernik, $\\frac{45}{8}$ by 1 by 0.094 in.); frequency,\n6100 cycles per second.\n\nNACA RM E9E06\n1125", "timestamp": "2026-07-22T04:45:15.414416+00:00"}
{"citation_id": "19930082245", "source_url": "https://ntrs.nasa.gov/api/citations/19930082245/downloads/19930082245.pdf", "page_number": 26, "total_pages": 66, "image_filename": "19930082245_p26.jpg", "text": "```markdown\n8\n6\n4\n2\n0\n-2\n-4\n-6\n-8\n-10\n-12\n-14\n.1 .2 .3 .4 .5 .6 .7 .8 .9\nMach number, M\n\n.16\n.12\n.08\n.04\n0\n-.04\n-.08\n-.12\n-.16\n-.20\n-.24\n-.28\n.1 .2 .3 .4 .5 .6 .7 .8 .9\nMach number, M\n\nSection angle of attack, $\\alpha$, deg\n$\\delta_a$ (deg)\n-12\n-6\n-4\n-2\n0\n2\n4\n12\n18\n\nSection pitching-moment coefficient, $C_m$\n$\\delta_a$ (deg)\n-12\n-6\n-4\n-2\n0\n2\n4\n12\n18\n\n(a) $C_n = -0.4$.\n\nNACA TN NO. 1596\n\nFigure 6.- Variation of angle of attack and pitching-moment coefficient with Mach number for an NACA 66,1-115 airfoil section equipped with an unsealed 0.20c plain aileron of true-airfoil-contour profile.\n\n25\n```", "timestamp": "2026-07-22T04:45:23.518518+00:00"}
{"citation_id": "19930082485", "source_url": "https://ntrs.nasa.gov/api/citations/19930082485/downloads/19930082485.pdf", "page_number": 19, "total_pages": 62, "image_filename": "19930082485_p19.jpg", "text": "18\nNACA TN No. 1810\n\nThe incompressible value of $\\sqrt{Z_m}$ is determined by setting $g = 0$ that is,\n$$ \\frac{\\mu}{J} = f = \\sqrt{Z_m} \\quad \\left( \\frac{\\mu}{J} < 0.25 \\right) $$\n\nThen\n$$ \\frac{W}{\\rho_t n_o J} = V_m $$\n\nInasmuch as the flow is considered incompressible\n$$ \\rho_t = \\rho_s $$\n$$ V_m = \\frac{W}{\\rho_s n_o J} $$\n\nEvaluation of $J$ and $K$. - A chart for the evaluation of $J$ and $K/J$ to be used in finding $\\sqrt{Z_m}$ would be a simplification over evaluation from tables. A chart can be constructed where $J$ and $K/J$ are functions of the surface curvature and channel width, inasmuch as from equation (26)\n$$ J_1 = \\frac{\\exp(-t_m^2)}{\\sqrt{-\\frac{\\Delta C n_o}{2}}} [F(t_1) - F(t_2)] \\quad (28) $$\n\nLet\n$$ y = \\sqrt{-\\frac{n_o \\Delta C}{2}} \\quad (29) $$\n\nThen,\n$$ t_m = \\frac{-y}{\\Delta C} \\left( \\frac{C_2 + C_1}{2} \\right) \\quad (30) $$\n\nand\n$$ F(t_1) = F \\left( t_m + \\frac{y}{2} \\right) \\quad (31) $$", "timestamp": "2026-07-22T04:45:24.583322+00:00"}
{"citation_id": "19930082498", "source_url": "https://ntrs.nasa.gov/api/citations/19930082498/downloads/19930082498.pdf", "page_number": 9, "total_pages": 49, "image_filename": "19930082498_p9.jpg", "text": "8\nNACA TN No. 1838\n\neffectiveness of the muffler and reduces the sound-pressure level to 86 decibels at 2000 rpm; therefore, the results indicate that the muffler acts more as a volume resonator than as a sound-energy absorber. The results obtained with the steel wool removed (muffler 21, table II) strengthen this view, because the low-frequency effectiveness is again increased. No increase in the over-all effectiveness occurred, however, because the second harmonic intensity increased. Various lengths of tail pipe were attached to this muffler (configurations 22 to 29, table II). The results show the large effect which the tail pipe may exert on the exhaust-system noise characteristics. For example, figure 8 shows that at 2000 rpm the sound-pressure levels of the first three harmonics vary continuously as the tail-pipe length is varied. At the same time the over-all level varies from 83 to 86 decibels. In addition, certain \"other sounds\" of various frequencies become more or less prominent as the tail-pipe length is altered. These other sounds all occur approximately at multiples of $16\\frac{2}{3}$ cycles per second, which is the firing frequency of one cylinder of the engine at 2000 rpm. Although the tail pipe can produce considerable attenuation at certain frequencies, it can also reduce the muffler effectiveness at other frequencies, so the choice of tail-pipe length is very important in the design of a muffler installation. Data from configuration 21 are not plotted in figure 8 because of changes in the length of the inlet wye between the tests of muffler 21 and the tests of configurations 22 to 28.\n\nMuffler 30 was designed, by using the formulas of reference 1, to give good attenuation over the complete frequency band (up to 1000 cps) which would be encountered at 2000 rpm. Asbestos packing was placed around the outside of the muffler to reduce the noise radiation from the muffler shell. (See fig. 1.) Although the muffler is quite effective it does not meet the original expectations, particularly at the second harmonic. In the design of this muffler the connecting holes in the large chambers were made small in an attempt to hold the chamber size required for the chosen value of cut-off frequency to a minimum. The area of the connecting orifices, however, may have been so small as to impair the silencing characteristics. Additional connecting orifices were drilled and the results of the tests (muffler 31, table II) support this view, because the second harmonic is reduced 28 decibels by this change. The low-frequency characteristics are slightly impaired, but nevertheless the over-all sound-pressure level of 82 decibels at 2000 rpm is the lowest attained by any muffler thus far discussed. The reduction of the exhaust-pipe diameter to 2 inches in the design of this muffler results in excessive back pressure. The over-all sound level of muffler 31 at 2000 rpm is not changed by removing the asbestos jacket (muffler 32, table II), nor is it changed by removing the high-frequency chambers (muffler 33, table II). Removal of one of the three large low-frequency chambers results in a 4-decibel noise increase (muffler 34, table II). Removal of all four medium-frequency chambers results in only a 1.5-decibel further noise increase (mufflers 35 and 36, table II). Elimination of one of the two low-frequency chambers", "timestamp": "2026-07-22T04:45:26.137036+00:00"}
{"citation_id": "19930085838", "source_url": "https://ntrs.nasa.gov/api/citations/19930085838/downloads/19930085838.pdf", "page_number": 106, "total_pages": 118, "image_filename": "19930085838_p106.jpg", "text": "104\nNACA RM No. L9B23\n\n[Figure: A graph plotting Section drag coefficient, $c_d$ (y-axis) against Section lift coefficient, $c_l$ (x-axis). The y-axis ranges from .004 to .028. The x-axis ranges from -.8 to 1.6. A curve with circular data points shows a U-shape, reaching a minimum drag coefficient near $c_l = 0.2$. A legend box inside the graph indicates the data corresponds to $\\delta_f = 0$ (deg), $\\delta_a = 0$ (deg), and $\\delta_l = 0$ (deg). The NACA logo is visible in the bottom right corner of the plot area.]\n\nFigure 14.- Drag characteristics of the approximately 15.4-percent-chord thick NACA 7-series-type airfoil with double slotted flap, straight-sided Frise aileron, and flap.\nR = 6 x $10^6$ (approx.).", "timestamp": "2026-07-22T04:45:29.908197+00:00"}
{"citation_id": "19930082585", "source_url": "https://ntrs.nasa.gov/api/citations/19930082585/downloads/19930082585.pdf", "page_number": 4, "total_pages": 30, "image_filename": "19930082585_p4.jpg", "text": "NACA TN 1907\n\nVelocities\n\nV true airspeed of helicopter along flight path, feet per second\n\nVv vertical component of V (positive down, as in descent\n\nΩ rotor angular velocity, radians per second\n\nv induced velocity at rotor (always positive), feet per second\n\nλ inflow ratio (assuming v constant over the disk) $\\left(\\frac{V_{v}-v}{\\Omega R}\\right)$\n\nBlade-Element Aerodynamic Characteristics\n\n$c_{d_{o}}$ section profile-drag coefficient\n\n$c_{d_{o}}^{\\prime}$ $c_{d_{o}}$ corrected for friction torque, auxiliary mechanisms, and so forth\n\n$\\delta_{0}, \\delta_{0}^{\\prime}, \\delta_{1}, \\delta_{2}$ coefficients in power series for $c_{d_{o}}$ as a function of angle of attack\n\n$\\left(c_{d_{o}}=\\delta_{0}+\\delta_{1} \\alpha_{r}+\\delta_{2} \\alpha_{r}^{2}\\right)$\n\nor\n\n$\\left(c_{d_{o}}^{\\prime}=\\delta_{0}^{\\prime}+\\delta_{1} \\alpha_{r}+\\delta_{2} \\alpha_{r}^{2}\\right)$\n\n$\\Delta c_{d_{o}}, \\Delta \\delta_{0}$ increment in $c_{d_{o}}$ to account for friction torque, and so forth\n\n$\\left(\\Delta c_{d_{o}}=c_{d_{o}}^{\\prime}-c_{d_{o}}=\\Delta \\delta_{0}=\\delta_{0}^{\\prime}-\\delta_{0}\\right)$\n\na slope of lift curve, per radian", "timestamp": "2026-07-22T04:45:34.604372+00:00"}
{"citation_id": "19930086015", "source_url": "https://ntrs.nasa.gov/api/citations/19930086015/downloads/19930086015.pdf", "page_number": 54, "total_pages": 54, "image_filename": "19930086015_p54.jpg", "text": "UNCLASSIFIED\n\nThis document contains information affecting the National Defense of the United States, within the meaning of the Espionage Laws, Title 18, U.S.C., Sections 793 and 794. Its transmission or the revelation of its contents in any manner to an unauthorized person is prohibited by law.\n\nUNCLASSIFIED", "timestamp": "2026-07-22T04:45:35.991150+00:00"}
{"citation_id": "19930085842", "source_url": "https://ntrs.nasa.gov/api/citations/19930085842/downloads/19930085842.pdf", "page_number": 94, "total_pages": 104, "image_filename": "19930085842_p94.jpg", "text": "90\nNACA RM L9C29\n\nPropeller advance-diameter ratio, $V/nD$\nLift coefficient, $C_L$\nTorque coefficient, $Q_c$\nResultant drag coefficient, $C_{DR}$\n\nNATIONAL ADVISORY\nCOMMITTEE FOR AERONAUTICS\n\n(d) $\\alpha_u = 29^\\circ$.\nFigure 44.- Continued.", "timestamp": "2026-07-22T04:45:36.050887+00:00"}
{"citation_id": "19930082542", "source_url": "https://ntrs.nasa.gov/api/citations/19930082542/downloads/19930082542.pdf", "page_number": 5, "total_pages": 53, "image_filename": "19930082542_p5.jpg", "text": "4\nNACA TN No. 1867\n\nChemical composition:\nThe chemical composition was reported to be the following percentages:\n\n| C | Mn | Si | S | P | Cr | Ni | Co | Mo | W | Cb | N |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| 0.12 | 1.64 | 0.39 | 0.003 | 0.026 | 21.33 | 18.88 | 18.60 | 3.21 | 1.97 | 1.10 | 0.12 |\n\nFabrication procedure:\nThe approximately 6000-pound (8- by 7-inch) ingot was hammer cogged without difficulty to 2-inch-square billets with $2050^\\circ$ as maximum and $1750^\\circ$ F as minimum working temperatures. The 2-inch-square billets were hot-rolled to 7/8-inch-square bars in one heat from $2075^\\circ$ F to a finishing temperature of $1725^\\circ$ F. One hundred feet of the 7/8-inch-square bar stock was furnished for this investigation.\n\nStructure:\nAs received, the bar stock had a Brinell hardness of 233 and a 0.02-percent-offset yield strength of 72,500 psi at room temperature. The grain size was finer than 8 and there was relatively little grain-boundary precipitation, as is shown by figure 1. Considerable center segregation of excess constituents was observed however as is shown by the macrographs. After considerable study it was decided that the segregated condition was not too severe and that the stock was typical of the material being produced.\n\nEXPERIMENTAL PROCEDURE\n\nThe experimental procedure was designed to show the effects of the following treatments on the properties of low-carbon N-155 alloy at room temperature and the rupture test characteristics at $1200^\\circ$ F:\n\n(1) Hot-rolled stock\n(a) Aging\n(b) Hot-cold-work\n(c) Agglomeration of excess constituents", "timestamp": "2026-07-22T04:45:47.626645+00:00"}
{"citation_id": "19930082476", "source_url": "https://ntrs.nasa.gov/api/citations/19930082476/downloads/19930082476.pdf", "page_number": 20, "total_pages": 41, "image_filename": "19930082476_p20.jpg", "text": "18\nNACA TN No. 1801\n\nCHART 3.-SPIN CHARACTERISTICS OF MODEL WITH MASS DISTRIBUTION INCREASED ALONG\nTHE WINGS (LINKED RUDDER AND AILERON CONTROLS)\n\n$$ \\left[ \\frac{I_x - I_y}{mb^2} = 165 \\times 10^{-4}; \\mu = 5.32 \\text{ (loading 3 in table II and point 3 in fig. 4); right erect spins} \\right] $$\n\n| | Wheel setting | | |\n| :--- | :---: | :---: | :---: |\n| | **Left** | **0** | **Right** |\n| | | | **1/2** |\n| | | | **Full** |\n| **Elevator setting, degrees** | | | |\n| **Up** | | | |\n| 20 | | a | b |\n| | | 21 2D | |\n| | | 30 8D | |\n| | | 140 0.56 | |\n| 13 | No spin | a | c |\n| | | 18 2D | |\n| | | 24 8D | |\n| | | 151 0.66 | |\n| 8 | | b | |\n| | | | |\n| 5 | | No spin | |\n| | | | |\n| 0 | No spin | | No spin |\n| **Down** | | | |\n| 12 | No spin | No spin | No spin |\n\n$^a$Oscillatory spin, range of values or average value given.\n$^b$Steep spin, vertical velocity too high to permit obtaining test data.\n$^c$Steep spiral.\n\nModel values converted to corresponding full-scale values.\nU Inner wing up\nD Inner wing down\n\nNACA\n\n| $\\alpha$ (deg) | $\\theta$ (deg) |\n| :---: | :---: |\n| V (fps) | $\\Omega$ (rps) |", "timestamp": "2026-07-22T04:45:48.914288+00:00"}
{"citation_id": "19930082487", "source_url": "https://ntrs.nasa.gov/api/citations/19930082487/downloads/19930082487.pdf", "page_number": 19, "total_pages": 33, "image_filename": "19930082487_p19.jpg", "text": "NACA TN No. 1813\n17\n\n[Figure: Graph showing Pressure coefficient, P vs Mach number, M_o. The graph contains multiple curves for different x/c values (0.05, 0.10, 0.15, 0.20, 0.25, 0.30, 0.40, 0.60, 0.85). Annotations include M, 1.0; M, 1.4; P_{M_o=0} \\sqrt{1-M_o^2}; M_{cr}; M_d; M_s.]\n\n(a) Upper surface.\n[NACA logo]\nFigure 3.- Variation of pressure coefficient with Mach number at various chordwise stations of NACA 23015 airfoil section; $\\alpha$, 2°.", "timestamp": "2026-07-22T04:45:50.308473+00:00"}
{"citation_id": "19930093773", "source_url": "https://ntrs.nasa.gov/api/citations/19930093773/downloads/19930093773.pdf", "page_number": 23, "total_pages": 47, "image_filename": "19930093773_p23.jpg", "text": "22\n\nExhaust-nozzle jet-velocity coefficient, $C_j$\n\n1.04\n\n1.00\n\n.96\n\n.92\n\n.88\n\n1.0 1.4 1.8 2.2 2.6 3.0 3.4 3.8\n\nExhaust-nozzle pressure ratio, $P_7/P_0$\n\nFigure 3. - Variation of exhaust-nozzle jet-velocity coefficient with exhaust-nozzle pressure ratio.\n\nNACA\n\nNACA RM E9609", "timestamp": "2026-07-22T04:45:50.800051+00:00"}
{"citation_id": "19930082617", "source_url": "https://ntrs.nasa.gov/api/citations/19930082617/downloads/19930082617.pdf", "page_number": 2, "total_pages": 58, "image_filename": "19930082617_p2.jpg", "text": "```markdown\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nTECHNICAL NOTE 1962\n\nSTRESSES IN AND GENERAL INSTABILITY OF\nMONOCOQUE CYLINDERS WITH CUTOUTS\n\nVII - EXPERIMENTAL INVESTIGATION OF CYLINDERS\nHAVING EITHER LONG BOTTOM CUTOUTS OR\nSERIES OF SIDE CUTOUTS\n\nBy N. J. Hoff, Bruno A. Boley, and Joseph J. Mele\n\nSUMMARY\n\nEight 24S-T$^1$ Alclad cylinders 20 inches in diameter and 0.012 inch\nin sheet thickness were tested in pure bending. They were between 57.87\nand 77.16 inches long and were reinforced with 16 stringers and from 14\nto 29 rings of various cross-sectional areas and spacings. Of these,\nfive had a 45° cutout on the compression side extending over 33.44,\n34.72, or 51.44 inches in the axial direction. The remaining three\nhad a series of five 45° cutouts on each side.\n\nAll but one cylinder failed in general instability. The one that\ndid not show the general-instability pattern failed in tension.\n\nINTRODUCTION\n\nThe effect of cutouts on the stress distribution in end on the\ngeneral-instability buckling load of monocoque cylinders has been\ndiscussed in references 1 to 6. These reports describe theoretical\nand experimental investigations which were carried out at the\nPolytechnic Institute of Brooklyn Aeronautical Laboratories. All the\ncylinders studied were loaded in pure bending. The cutout was located\neither on the compression side (symmetric cutout) symmetrically with\nrespect to the most highly compressed stringer, or on a side (side\ncutout) in the neighborhood of the neutral axis.\n\nTwo main effects of the cutouts were studied. The first is the\nstress concentration which arises because of the discontinuities\nintroduced by the cut members of the cylinder. The second is the\ndecrease in the general-instability buckling load of the cylinders\nbecause of the increased freedom of motion of the cut members in the\nneighborhood of the cutout.\n\n$^1$New designations: Alclad 24S-T sheet, now Alclad 24S-T3;\n24S-T extrusions, now 24S-T4; Al7S-T rivets, now Al7S-T4.\n```", "timestamp": "2026-07-22T04:45:55.725543+00:00"}
{"citation_id": "19930085965", "source_url": "https://ntrs.nasa.gov/api/citations/19930085965/downloads/19930085965.pdf", "page_number": 60, "total_pages": 67, "image_filename": "19930085965_p60.jpg", "text": "```markdown\n1160\n\nNACA RM E9E06\n\n| Calculated a.c. | Measured a.c. | Measured pulsating d.c. | | Total air gap (in.) |\n| :---: | :---: | :---: | :--- | :---: |\n| $\\diamondsuit$ | $\\circ$ | | SAE 1020 steel | 0.04 |\n| $\\nabla$ | $\\nabla$ | $\\mathbb{V}$ | Armco Magnetic Ingot Iron | .04 |\n| | $\\square$ | | Armco Magnetic Ingot Iron | .06 |\n| $\\Delta$ | $\\Delta$ | | Hipernik | .04 |\n\nPower, watts/sq in.\n\n[Graph plotting Power vs Effective ampere turns with data points corresponding to the legend above]\n\nNACA\n\nFigure 16. - Ampere turns required for eddy-current heating. Vane, $4\\frac{5}{8}$ by 1 by $\\frac{1}{8}$ inch (Hipernik, $4\\frac{5}{8}$ by 1 by 0.094 in.); frequency, 6100 cycles per second.\n\n59\n```", "timestamp": "2026-07-22T04:45:56.395414+00:00"}
{"citation_id": "19930092013", "source_url": "https://ntrs.nasa.gov/api/citations/19930092013/downloads/19930092013.pdf", "page_number": 19, "total_pages": 21, "image_filename": "19930092013_p19.jpg", "text": "APPARATUS FOR VARYING EFFECTIVE DIHEDRAL IN FLIGHT 15\n\n[Figure: Graph with vertical axis labeled \"Time to double amplitude, sec\" and horizontal axis labeled \"Period, sec\". Data points and curves are plotted with labels including: \"Landing approach\", \"Cruising\", \"High speed\", \"Good\", \"Tolerable\", \"Intolerable\", \"Region of approximately neutral stability\", \"T(Ts = 28.4°; Ts = 38 sec)\", \"G(Ts = 22.7°)\", \"I(Ts = 24.4°)\", \"T(Ts = 18.2°)\", \"G(Ts = 12.9°)\", \"G(Ts = 14.2°)\", \"G(Ts = 5.3°)\", \"G(Ts = 0°)\", \"G(Ts = 6.2°)\", and a diagonal line labeled \"Used by pilots for field landings (refs 2 & 3)\".]\n\nFIGURE 19.—Comparison of pilots’ opinions of over-all lateral characteristics with period-damping requirements of references 2 and 3.\n\nable and the other intolerable. The difficulty in controlling the oscillations in the cruising condition was finally attributed by the pilots to the high rolling velocities. These higher rolling velocities are apparent when figure 14 (a) is compared with figure 13 (a). It seems, then, that a period-damping relationship cannot, in itself, define all of a pilot’s concepts of the lateral-dynamic-stability characteristics, at least when extreme values of effective dihedral are considered. It would seem that a limitation should be placed on the rolling response to some form of yawing or sideslipping disturbance. The reduction of these concepts to a concrete, numerical criterion is a problem which deserves consideration in future work.\n\nA possible criterion on which a limitation might be placed is the ratio of the amplitude of the rolling velocity to the amplitude of the sideslip angle in the oscillatory mode as measured in lateral oscillations such as were made for this investigation. Another possible criterion worthy of future study is the ratio of the amplitude of angle of bank to that of the angle of sideslip, perhaps as a function of period. For purposes of future reference, the above-mentioned quantities were evaluated from the data gathered during this investigation and are presented in table I together with the periods, the effective dihedrals, and the pilots’ opinions of the over-all lateral handling characteristics.\n\nThe minimum tolerable effective dihedral in the landing-approach condition is seen in figure 18 to be about $-7^\\circ$. With $\\Gamma_s = -10.7^\\circ$, the adverse rolling response to rudder control (left roll with right rudder) was considered by the pilots to be intolerably rapid for a landing approach. It should be noted here, however, that, although all flights were made at altitude, the pilots based their opinions on the consideration of the use of the airplane for field landings. It is believed that, due to lower approach speeds and the necessity for rapid maneuvers during wave-off, the minimum tolerable effective dihedral for carrier landings would be less negative.\n\nThe minimum tolerable effective dihedral in the cruising and high-speed conditions is shown in figure 18 to be about $-5^\\circ$. With $\\Gamma_s = -7.1^\\circ$, the rolling response to gusts and the adverse rolling response to rudder control when corrections were made were so rapid that the pilot had to be constantly on the controls, a situation which, the pilots believed, would be intolerable from the standpoint of fatigue on flights of normal duration.\n\nIt was the opinion of the pilots that the optimum values of effective dihedral investigated were $6.2^\\circ$ (normal airplane without apparatus) for the cruising and high-speed conditions and $14.2^\\circ$ for the landing-approach condition. They thought more than normal amounts of dihedral were desirable in the approach condition because of the good response in roll to rudder control. It is noteworthy that this is the direction of the variation of effective dihedral with lift coefficient for swept-back wings; that is, increasing lift coefficient results in increasing effective dihedral.\n\nConsideration of the results with respect to the flying-qualities specifications.—Examination of references 2 and 3 indicates that the requirements which probably limit the designer’s choice of effective dihedral in most cases are, for the lower limit, the requirement that static effective dihedral be positive and, for the upper limit, the prohibition of rolling-velocity reversal during aileron rolls and the oscillation period-damping requirement (fig. 19). Information gathered during this investigation has indicated that, if these requirements are met with an airplane similar to the test airplane, the resultant lateral-stability characteristics will certainly be satisfactory. The investigation has further indicated, however, that, if necessary, small negative values of effective dihedral can be tolerated and that the upper limit of dihedral is determined by some criterion other than a restriction against rolling-velocity reversal during aileron rolls or a period-damping relationship. The tolerable amount of negative dihedral is apparently related to the growth of rolling motion following a yawing-moment disturbance.\n\nThe specific values of the limits of tolerable effective dihedral determined in the present investigation, of course, cannot be applied generally to all airplanes. It is believed that an investigation should be conducted with control over other stability parameters, such as directional stability and directional damping, as well as control over effective dihedral. With such additional control, it would be possible to vary the characteristics of the airplane motion (period, damping,", "timestamp": "2026-07-22T04:46:00.290097+00:00"}
{"citation_id": "19930082245", "source_url": "https://ntrs.nasa.gov/api/citations/19930082245/downloads/19930082245.pdf", "page_number": 27, "total_pages": 66, "image_filename": "19930082245_p27.jpg", "text": "```markdown\n26\n\n8\n6\n4\n2\n0\n-2\n-4\n-6\n-8\n-10\n-12\n-14\n.1 .2 .3 .4 .5 .6 .7 .8 .9\nMach number, M\n\nSection angle of attack, $\\alpha$, deg\n\n$\\delta_a$\n(deg)\n-12\n-6\n-4\n-2\n0\n2\n4\n-12\n-18\n\n(b) $c_n = -0.2$.\nFigure 6.—Continued.\n\n.16\n.12\n.08\n.04\n0\n-.04\n-.08\n-.12\n-.16\n-.20\n-.24\n-.28\n.1 .2 .3 .4 .5 .6 .7 .8 .9\nMach number, M\n\nSection pitching moment coefficient, $c_m$\n\n$\\delta_a$\n(deg)\n-12\n-6\n-4\n2\n0\n2\n4\n12\n18\n\nNACA\nNACA TN NO. 1596\n```", "timestamp": "2026-07-22T04:46:00.569742+00:00"}
{"citation_id": "19930082447", "source_url": "https://ntrs.nasa.gov/api/citations/19930082447/downloads/19930082447.pdf", "page_number": 21, "total_pages": 24, "image_filename": "19930082447_p21.jpg", "text": "```markdown\nNACA TN No. 1775\n\n10\n\nTime coefficient, $C_t$\n\nTheoretical,\nreference 1\n$\\circ \\square \\triangle$ Experimental\n$\\circ \\square \\triangle$ Consistency runs\n\nAt rebound\n\nAt maximum\ndraft\n\n1.0\n\nAt maximum\nacceleration\n\n.1\n\n.1 1.0 10\nApproach parameter, $\\kappa$\n\nNACA\n\nFigure 5.—Variation of time coefficient with approach\nparameter.\n\n19\n```", "timestamp": "2026-07-22T04:46:02.390848+00:00"}
{"citation_id": "19930082614", "source_url": "https://ntrs.nasa.gov/api/citations/19930082614/downloads/19930082614.pdf", "page_number": 3, "total_pages": 36, "image_filename": "19930082614_p3.jpg", "text": "```markdown\nTECH LIBRARY KAFB, NM\n0065356\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nTECHNICAL NOTE 1939\n\nTHE EFFECTS OF AERODYNAMIC BRAKES UPON THE\nSPEED CHARACTERISTICS OF AIRPLANES\n\nBy Jack D. Stephenson\n\nSUMMARY\n\nA study has been made of the factors influencing the performance\nof aerodynamic brakes. The requirements which must be met in order for\nthe brakes to provide the necessary control over the forward speed are\ndiscussed for various flight conditions under which they may be used.\nEquations relating the speed and altitude are presented for several\ncases in which certain simplifying assumptions are made. For these\ncases, formulas and graphs in the report furnish a means of quickly\ncomputing the longitudinal speed variations, the dive angles, and the\nrates of descent for airplanes having known drag characteristics. For\nthe cases to which the simplifying assumptions do not apply, it is\nindicated that a satisfactory solution, which takes into account all\nof the possible variables (such as atmospheric density, drag coeffi-\ncient, and flight-path angle) can be obtained by a step-by-step method\nof calculation. Graphs are presented to reduce the time required for\nstep-by-step calculations. Example calculations, which show each step\nin detail, illustrate the use of the graphs and formulas.\n\nThe increases in drag coefficient that are characteristic of\nseveral types of wing and fuselage aerodynamic brakes, which have been\ntested in wind tunnels or in flight, are summarized in the report.\nThe effect of Mach number on the drag coefficient and the effect of\npartial brake deflection are included where such data are available.\n\nINTRODUCTION\n\nThe continued improvement in the aerodynamic design of high-speed\nairplanes has brought the normal operating speeds near to their maximum\nsafe speeds. As a result of low drag and of high engine output,\nairplanes may under certain circumstances accelerate to speeds at which\ndangerous compressibility effects or structural loadings arise. The\ntrend toward higher wing loadings has added considerably to the possi-\nbility of attaining dangerous speeds, especially at the high operating\naltitudes which are characteristic of many modern airplanes.\n```", "timestamp": "2026-07-22T04:46:02.573622+00:00"}
{"citation_id": "19930085838", "source_url": "https://ntrs.nasa.gov/api/citations/19930085838/downloads/19930085838.pdf", "page_number": 107, "total_pages": 118, "image_filename": "19930085838_p107.jpg", "text": "NACA RM No. L9B23\n105\n\nMaximum section lift coefficient, $c_{l_{max}}$\n\nAirfoil\n$\\circ$ 0.154c thick\n$\\square$ .154c thick (true-contour aileron, reference 3)\n$\\diamond$ .176c thick\n\nFlap deflection, $\\delta_f$, deg\n\n[Figure: NACA logo]\n\nFigure 15.- Maximum lift characteristics of the NACA 7-series-type airfoils with double slotted flap, straight-sided Frise aileron, and flip. $R = 6 \\times 10^6$ (approx.); $\\delta_a = 0^\\circ$; $\\delta_l = 0^\\circ$; $c_b = 0.408c_a$.", "timestamp": "2026-07-22T04:46:09.946097+00:00"}
{"citation_id": "19930082485", "source_url": "https://ntrs.nasa.gov/api/citations/19930082485/downloads/19930082485.pdf", "page_number": 20, "total_pages": 62, "image_filename": "19930082485_p20.jpg", "text": "NACA TN No. 1810\n19\n\n$$F(t_2) = F(t_m - \\frac{y}{2}) \\tag{32}$$\n\nThese functions may be expanded in a Taylor's series.\n\n$$F(t_1) - F(t_2) = 2 \\left[ \\left(\\frac{y}{2}\\right) F'(t_m) + \\left(\\frac{y}{2}\\right)^3 \\frac{F'''(t_m)}{3!} + \\dots + \\left(\\frac{y}{2}\\right)^n \\frac{F^n(t_m)}{n!} \\right] \\tag{33}$$\n\nwhere $n$ is an odd positive integer and\n$$F'(t_m) = \\exp(t_m^2)$$\n$$F'''(t_m) = (2 + 4 t_m^2) \\exp(t_m^2)$$\n\nBy use of the first five terms of the Taylor's series,\n\n$$J = 1 + \\frac{y^2}{12} (1 + 2 t_m^2) + \\frac{y^4}{480} (3 + 12 t_m^2 + 4 t_m^4) \\tag{34}$$\n\nNow, let\n\n$$a = t_1 y = \\frac{n_o C_1}{2} \\tag{35}$$\n\n$$b = \\frac{t_1}{y} = - \\frac{C_1}{\\Delta C} \\tag{36}$$\n\nThen,\n\n$$t_m^2 = \\frac{a}{4b} (2b - 1)^2$$\n\nand equation (34) becomes\n\n$$J = 1 + \\frac{a}{12b} \\left[ 1 + \\frac{a}{2b} (2b - 1)^2 \\right]$$\n$$+ \\frac{1}{480} \\left(\\frac{a}{b}\\right)^2 \\left[ 3 + 5 \\left(\\frac{a}{b}\\right) (2b-1)^2 + \\frac{1}{4} \\left(\\frac{a}{b}\\right)^2 (2b - 1)^4 \\right] \\tag{37}$$\n\nand by the same method but using seven terms of the Taylor's series", "timestamp": "2026-07-22T04:46:11.363184+00:00"}
{"citation_id": "19930082496", "source_url": "https://ntrs.nasa.gov/api/citations/19930082496/downloads/19930082496.pdf", "page_number": 11, "total_pages": 50, "image_filename": "19930082496_p11.jpg", "text": "10\nNACA TN No. 1836\n\ncommonly used. These alloys have a short-time tensile strength as great as 37,800 pounds per square inch for alloy 422-19 at 1800° F as indicated in reference 9. In general, the short-time tensile strength for high-temperature alloys is about 33,000 pounds per square inch at 1800° F. Reference 1 indicates that National Bureau of Standards Body 4811 has a short-time tensile strength of 19,000 pounds per square inch at 1800° F. At 2200° F, there are few short-time tensile data for the metal alloys or for ceramics with which the ceramal tensile strengths of 8900 and 13,200 pounds per square inch compared. This value can be considered good at temperatures approaching the melting points of metal alloys. Figure 8 shows specimen 3D5 after failure and illustrates a fracture typical of a brittle material.\n\nWhen densities of 5.5 (reference 2), 3.0 (reference 1), and 8.3 (reference 9) grams per cubic centimeter were used for the ceramal, the National Bureau of Standards Body 4811, and alloy 422-19, respectively, which yielded better than average strength-to-weight ratios for the ceramic and the metal, the strength-to-weight ratio of the ceramal at 1800° F is 0.95 times that of the ceramic and 1.3 times that of the alloy. Inasmuch as centrifugal forces are largely responsible for stresses in a turbine blade, the high strength-to-weight ratio of the ceramal compared with metals is of importance.\n\nThermal-Shock Evaluation\n\nThe results obtained in the thermal-shock evaluation are presented in table II. The ceramal survived 25 cycles at 2400° F, whereas a zircon ceramic survived one cycle at 1800° F. Data for the ceramic are included to enable a comparative evaluation using the same apparatus. Zircon is regarded as having good thermal-shock properties (reference 10). Titanium carbide was investigated to allow comparison of the thermal-shock characteristics of titanium carbide with titanium carbide plus cobalt. The addition of cobalt tends to improve the thermal-shock characteristics of the ceramic. The thermal-shock resistance of this ceramic is very good; specimen 3A7 survived 14 cycles at 2400° F and specimen 3A10 survived 21 cycles at 2400° F.\n\nDuring the thermal-shock investigation, severity of oxidation of the ceramal specimens increased as evaluation temperatures progressed from 1800° to 2400° F. Upon completion of 25 cycles at 1800° F, a relatively thin, tight film was noted. Upon completion of the 25 cycles at 2400° F, the film had become scaly and more", "timestamp": "2026-07-22T04:46:12.537685+00:00"}
{"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 1, "total_pages": 78, "image_filename": "19930082618_p1.jpg", "text": "NACA\nTN\n1945\na1\n\nNACA TN 1945\n8381\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\n\nTECHNICAL NOTE 1945\n\nLOAN COPY: RET\nAFWL TECHNICAL\nKIRTLAND AFB, NM\n\nTECH LIBRARY KAFB, NM\n0065341\n\nAERODYNAMIC CHARACTERISTICS OF 15 NACA AIRFOIL\nSECTIONS AT SEVEN REYNOLDS NUMBERS\nFROM $0.7 \\times 10^6$ TO $9.0 \\times 10^6$\n\nBy Laurence K. Loftin, Jr. and Hamilton A. Smith\n\nLangley Aeronautical Laboratory\nLangley Air Force Base, Va.\n\n[Figure: NACA logo]\n\nWashington\nOctober 1949\n\nAFMDC\nTECHNICAL LIBRARY\nAFL 2811\n\n219.98/41", "timestamp": "2026-07-22T04:46:15.183265+00:00"}
{"citation_id": "19930082498", "source_url": "https://ntrs.nasa.gov/api/citations/19930082498/downloads/19930082498.pdf", "page_number": 10, "total_pages": 49, "image_filename": "19930082498_p10.jpg", "text": "NACA TN No. 1838\n\nresults in an additional sharp noise increase of 7 decibels (muffler 37, table II). These characteristics indicate that the noise reduction is principally due to the low-frequency chambers, so for this engine a small number of large low-frequency chambers would be much more effective than a large number of small high-frequency chambers. This result agrees with the design equations of reference 1, which state that the low-frequency cut-off of a muffler of this type is inversely proportional to the square root of the chamber volume.\n\nSeveral arrangements utilizing two resonant chambers were investigated and some had excellent muffling capabilities. Included in this group was the muffler used in the demonstration of light-airplane noise reduction (muffler 43, table II, fig. 1, and reference 1). In the process of this investigation one of the mufflers was received and tested with the central baffle spot-welded at only four points on the circumference. The baffle was later seam-welded to eliminate any leakage between the two chambers and the muffler was retested. The sound-pressure level for the 1650-rpm condition is reduced more than 10 decibels by this means, this reduction indicating that there should be no leakage between the chambers. The back pressures of these straight-through mufflers with a $2\\frac{3}{4}$-inch-diameter central tube are, in general, lower than those for the original engine installation at 1650 rpm and 2000 rpm but are higher at 2790 rpm.\n\nThe muffler attenuation curves of reference 1 show that in order to obtain sufficient noise reduction the cut-off frequency must be chosen somewhat below the lowest frequency for which attenuation is desired. An allowance must also be made in the design of a muffler for the accuracy with which the exhaust-gas temperature, and hence the speed of sound, is known and for the fact that this temperature is not constant. The effect of choosing the design cut-off frequency too close to the engine fundamental frequency, at which high attenuation is required, was demonstrated by the tests of muffler 49, which was designed for a low-frequency cut-off about 7 percent lower than the fundamental frequency at 2790 rpm. The performance of this muffler was poor, inasmuch as the sound-pressure level of the engine fundamental was 97 decibels at 2790 rpm with this muffler, which is only 4 decibels below that with the wye alone (configuration 4, table II). Note that at 2000 rpm the fundamental, which is lower than the design cut-off frequency, is not attenuated by muffler 49.\n\nAn attempt was made to shape a resonant-chamber muffler to have the largest volume possible in the available space within the engine cowling (fig. 2(b)). This particular muffler (muffler 57, table II) was of little value as a noise reducer because the flat sides vibrated with the exhaust pulses and radiated sound. This muffler was modified by connecting the resonant chambers (fig. 2(c) and muffler 58, table II) with the hope that such an alteration might have a beneficial effect similar to that obtained by joining the two exhaust pipes ahead of the", "timestamp": "2026-07-22T04:46:19.323221+00:00"}
{"citation_id": "19930082487", "source_url": "https://ntrs.nasa.gov/api/citations/19930082487/downloads/19930082487.pdf", "page_number": 20, "total_pages": 33, "image_filename": "19930082487_p20.jpg", "text": "18\nNACA TN No. 1813\n\nPressure coefficient, P\n$$M, 1.0$$\n$$\\frac{PM_o=0}{\\sqrt{1-M_o^2}}$$\nx/c\n0.20\n.30\n.40\n.60\n.10\n.05\n.90\n0\n-.4\n-.8\n.3 .4 .5 .6 .7 .8 .9\nMach number, $M_o$\n$M_{cr}$ $M_d$ $M_s$\n\n(b) Lower surface.\nFigure 3.- Concluded.\nNACA", "timestamp": "2026-07-22T04:46:22.929209+00:00"}
{"citation_id": "19930093773", "source_url": "https://ntrs.nasa.gov/api/citations/19930093773/downloads/19930093773.pdf", "page_number": 24, "total_pages": 47, "image_filename": "19930093773_p24.jpg", "text": "NACA RM E9G09\n23\n\nNet thrust, $F_n$, lb\n\nAltitude\n(ft)\nO 5,000\n□ 15,000\n◇ 25,000\n△ 35,000\n▽ 45,000\n◁ 50,000\n\nEngine speed, N, rpm\n\n(a) Net thrust.\n\nFigure 4. - Effect of altitude on variation of engine performance with\nengine speed at flight Mach number of 0.21.", "timestamp": "2026-07-22T04:46:28.447580+00:00"}
{"citation_id": "19930085965", "source_url": "https://ntrs.nasa.gov/api/citations/19930085965/downloads/19930085965.pdf", "page_number": 61, "total_pages": 67, "image_filename": "19930085965_p61.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T04:46:29.261583+00:00"}
{"citation_id": "19930082450", "source_url": "https://ntrs.nasa.gov/api/citations/19930082450/downloads/19930082450.pdf", "page_number": 15, "total_pages": 37, "image_filename": "19930082450_p15.jpg", "text": "14\nNACA TN No. 1778\n\nTABLE 4.- Z-PLANE PROPERTIES\n$$\n\\left[ \\frac{v_{\\infty}}{u_{\\infty}} = 0.79; \\frac{v_{\\infty}}{u_{\\infty}} = -9.8; \\frac{v_{\\infty}}{u_{\\infty}} = 0.4; \\frac{v_{\\infty}}{u_{\\infty}} = 3; \\frac{v_{\\infty}}{u_{\\infty}} = 4; \\frac{v_{\\infty}}{u_{\\infty}} = 1.933; \\frac{v_{\\infty}}{u_{\\infty}} = 12.3 \\right]\n$$\n\n| $\\frac{v_{\\infty}}{u_{\\infty}}$ | 20 | 21 | 22 | 23 | 24 | 25 | 26 | 27 | 28 | 29 | 30 | 31 | 32 |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| 25 | 1.858 | 1.903 | 1.928 | 1.963 | 1.998 | 2.033 | 2.068 | 2.103 | 2.135 | 2.172 | 2.207 | 2.242 | 2.277 |\n| 26 | 1.825 | 1.856 | 1.892 | 1.926 | 1.959 | 1.993 | 2.027 | 2.060 | 2.094 | 2.127 | 2.161 | 2.195 | 2.228 |\n| 27 | 1.794 | 1.837 | 1.858 | 1.901 | 1.934 | 1.966 | 1.999 | 2.031 | 2.064 | 2.096 | 2.128 | 2.161 | 2.193 |\n| 28 | 1.766 | 1.797 | 1.828 | 1.860 | 1.891 | 1.922 | 1.953 | 1.984 | 2.016 | 2.047 | 2.078 | 2.109 | 2.140 |\n| 29 | 1.740 | 1.770 | 1.800 | 1.830 | 1.860 | 1.890 | 1.920 | 1.950 | 1.980 | 2.010 | 2.040 | 2.070 | 2.100 |\n| 30 | 1.715 | 1.744 | 1.773 | 1.802 | 1.831 | 1.861 | 1.890 | 1.919 | 1.948 | 1.977 | 2.006 | 2.035 | 2.064 |\n| 31 | 1.692 | 1.720 | 1.748 | 1.776 | 1.804 | 1.832 | 1.860 | 1.888 | 1.916 | 1.944 | 1.972 | 2.000 | 2.028 |\n| 32 | 1.670 | 1.698 | 1.725 | 1.752 | 1.779 | 1.807 | 1.834 | 1.861 | 1.889 | 1.916 | 1.943 | 1.971 | 1.998 |\n| 33 | 1.650 | 1.676 | 1.703 | 1.729 | 1.756 | 1.782 | 1.809 | 1.835 | 1.862 | 1.888 | 1.914 | 1.940 | 1.966 |\n| 34 | 1.631 | 1.657 | 1.682 | 1.708 | 1.733 | 1.759 | 1.785 | 1.811 | 1.836 | 1.862 | 1.888 | 1.913 | 1.939 |\n| 35 | 1.613 | 1.638 | 1.663 | 1.688 | 1.713 | 1.738 | 1.763 | 1.788 | 1.812 | 1.837 | 1.862 | 1.887 | 1.912 |\n| 36 | 1.596 | 1.620 | 1.645 | 1.669 | 1.694 | 1.718 | 1.743 | 1.767 | 1.792 | 1.816 | 1.840 | 1.864 | 1.888 |\n| 37 | 1.580 | 1.603 | 1.627 | 1.650 | 1.674 | 1.698 | 1.721 | 1.745 | 1.769 | 1.792 | 1.816 | 1.840 | 1.863 |\n| 38 | 1.564 | 1.587 | 1.610 | 1.633 | 1.656 | 1.679 | 1.702 | 1.725 | 1.748 | 1.771 | 1.794 | 1.817 | 1.840 |\n| 39 | 1.550 | 1.572 | 1.595 | 1.617 | 1.640 | 1.662 | 1.684 | 1.707 | 1.729 | 1.752 | 1.774 | 1.797 | 1.819 |\n| 40 | 1.536 | 1.558 | 1.580 | 1.602 | 1.624 | 1.646 | 1.667 | 1.689 | 1.711 | 1.733 | 1.755 | 1.777 | 1.799 |\n| 41 | 1.523 | 1.545 | 1.567 | 1.588 | 1.609 | 1.631 | 1.652 | 1.673 | 1.694 | 1.715 | 1.736 | 1.757 | 1.778 |\n| 42 | 1.511 | 1.532 | 1.553 | 1.574 | 1.595 | 1.616 | 1.637 | 1.658 | 1.679 | 1.700 | 1.720 | 1.741 | 1.761 |\n| 43 | 1.499 | 1.520 | 1.541 | 1.562 | 1.582 | 1.603 | 1.624 | 1.644 | 1.665 | 1.685 | 1.706 | 1.726 | 1.746 |\n| 44 | 1.488 | 1.509 | 1.529 | 1.550 | 1.570 | 1.591 | 1.611 | 1.631 | 1.652 | 1.672 | 1.692 | 1.712 | 1.732 |\n| 45 | 1.477 | 1.498 | 1.518 | 1.538 | 1.558 | 1.578 | 1.598 | 1.618 | 1.638 | 1.658 | 1.678 | 1.698 | 1.718 |\n| 46 | 1.467 | 1.487 | 1.507 | 1.527 | 1.547 | 1.567 | 1.587 | 1.607 | 1.626 | 1.646 | 1.666 | 1.685 | 1.705 |\n| 47 | 1.457 | 1.477 | 1.497 | 1.517 | 1.536 | 1.556 | 1.576 | 1.595 | 1.615 | 1.634 | 1.654 | 1.673 | 1.692 |\n| 48 | 1.447 | 1.467 | 1.487 | 1.506 | 1.526 | 1.545 | 1.565 | 1.584 | 1.604 | 1.623 | 1.642 | 1.661 | 1.680 |\n| 49 | 1.438 | 1.458 | 1.477 | 1.496 | 1.516 | 1.535 | 1.554 | 1.573 | 1.592 | 1.611 | 1.630 | 1.649 | 1.668 |\n| 50 | 1.429 | 1.448 | 1.467 | 1.486 | 1.505 | 1.524 | 1.543 | 1.562 | 1.581 | 1.600 | 1.619 | 1.638 | 1.657 |\n| 51 | 1.420 | 1.439 | 1.458 | 1.477 | 1.495 | 1.514 | 1.533 | 1.551 | 1.570 | 1.589 | 1.607 | 1.626 | 1.644 |\n| 52 | 1.412 | 1.430 | 1.449 | 1.467 | 1.486 | 1.504 | 1.523 | 1.541 | 1.560 | 1.578 | 1.596 | 1.614 | 1.632 |\n| 53 | 1.403 | 1.422 | 1.440 | 1.458 | 1.477 | 1.495 | 1.513 | 1.531 | 1.549 | 1.567 | 1.585 | 1.603 | 1.621 |\n| 54 | 1.395 | 1.413 | 1.431 | 1.449 | 1.467 | 1.485 | 1.503 | 1.521 | 1.539 | 1.557 | 1.575 | 1.593 | 1.610 |\n| 55 | 1.387 | 1.405 | 1.423 | 1.441 | 1.458 | 1.476 | 1.494 | 1.512 | 1.530 | 1.547 | 1.565 | 1.582 | 1.600 |\n| 56 | 1.379 | 1.397 | 1.415 | 1.432 | 1.450 | 1.468 | 1.485 | 1.503 | 1.520 | 1.538 | 1.555 | 1.572 | 1.590 |\n| 57 | 1.371 | 1.389 | 1.406 | 1.424 | 1.441 | 1.459 | 1.476 | 1.494 | 1.511 | 1.528 | 1.545 | 1.562 | 1.580 |\n| 58 | 1.363 | 1.381 | 1.398 | 1.415 | 1.433 | 1.450 | 1.467 | 1.484 | 1.501 | 1.518 | 1.535 | 1.552 | 1.569 |\n| 59 | 1.356 | 1.373 | 1.390 | 1.407 | 1.424 | 1.441 | 1.458 | 1.475 | 1.492 | 1.509 | 1.526 | 1.543 | 1.559 |\n| 60 | 1.348 | 1.365 | 1.382 | 1.399 | 1.416 | 1.433 | 1.450 | 1.466 | 1.483 | 1.500 | 1.516 | 1.533 | 1.549 |\n| 61 | 1.341 | 1.358 | 1.374 | 1.391 | 1.408 | 1.424 | 1.441 | 1.457 | 1.474 | 1.490 | 1.506 | 1.523 | 1.539 |\n| 62 | 1.333 | 1.350 | 1.367 | 1.383 | 1.400 | 1.416 | 1.432 | 1.448 | 1.465 | 1.481 | 1.497 | 1.513 | 1.529 |\n| 63 | 1.326 | 1.342 | 1.359 | 1.375 | 1.391 | 1.407 | 1.423 | 1.439 | 1.455 | 1.471 | 1.487 | 1.503 | 1.519 |\n| 64 | 1.", "timestamp": "2026-07-22T04:46:29.613152+00:00"}
{"citation_id": "19930082476", "source_url": "https://ntrs.nasa.gov/api/citations/19930082476/downloads/19930082476.pdf", "page_number": 21, "total_pages": 41, "image_filename": "19930082476_p21.jpg", "text": "NACA TN No. 1801\n19\n\nCHART 4.- SPIN CHARACTERISTICS OF MODEL WITH INCREASED RELATIVE DENSITY\n(LINKED RUDDER AND AILERON CONTROLS)\n$$\\left[ \\frac{I_x - I_y}{mb^2} = -18 \\times 10^{-4}; \\mu = 10.39 \\text{ (loading 4 in table II and point 4 in fig. 4); right erect spins} \\right]$$\n\n| | Wheel setting | | | |\n| :--- | :---: | :---: | :---: | :---: |\n| | **Left** | **0** | **$\\frac{1}{4}$** | **$\\frac{1}{2}$** | **Right** | **Full** |\n| **Elevator setting, degrees** | | | | | | |\n| **Up** | | | | | | |\n| 30 | | a<br>16 3D<br>25 10D<br>202 0.55 | | | | |\n| 20 | | No spin | | a<br>22 5D<br>29 12D<br>212 0.57 | a<br>22 0<br>30 7D<br>189 0.75 | |\n| 13 | | No spin | b | a<br>23 70<br>207 0.67 | a, c<br>20 6U<br>32 5D<br>215 | |\n| 5 | | | | c | | |\n| **Down** | | | | | | |\n| 0 | | No spin | | No spin | b | |\n| 12 | | No spin | | No spin | b | |\n\n$^a$Oscillatory spin, range of values or average value given.\n$^b$Steep spiral.\n$^c$Wandering spin.\n\nModel values converted to corresponding full-scale values.\nU inner wing up\nD inner wing down\n\n| $\\alpha$ (deg) | $\\beta$ (deg) |\n| :---: | :---: |\n| V (fps) | $\\omega$ (rps) |\n\nNACA", "timestamp": "2026-07-22T04:46:30.374198+00:00"}
{"citation_id": "19930082245", "source_url": "https://ntrs.nasa.gov/api/citations/19930082245/downloads/19930082245.pdf", "page_number": 28, "total_pages": 66, "image_filename": "19930082245_p28.jpg", "text": "NACA TN NO. 1596\n\n8\n6\n4\n2\n0\n-2\n-4\n-6\n-8\n-10\n-12\n-14\n.1 .2 .3 .4 .5 .6 .7 .8 .9\nMach number, M\n\nSection angle of attack, $\\alpha$, deg\n\n$\\delta_a$\n(deg)\n-12\n-6\n-4\n-2\n0\n2\n4\n12\n18\n30\n\n.16\n.12\n.08\n.04\n0\n-.04\n-.08\n-.12\n-.16\n-.20\n-.24\n-.28\n.1 .2 .3 .4 .5 .6 .7 .8 .9\nMach number, M\n\nSection pitching-moment coefficient, $C_m$\n\n$\\delta_a$\n(deg)\n-12\n-6\n-4\n-2\n0\n2\n4\n12\n18\n30\n\n[Figure: NACA logo]\n\n(c) $c_n = 0$.\nFigure 6.—Continued.\n\n27", "timestamp": "2026-07-22T04:46:35.870392+00:00"}
{"citation_id": "19930082617", "source_url": "https://ntrs.nasa.gov/api/citations/19930082617/downloads/19930082617.pdf", "page_number": 3, "total_pages": 58, "image_filename": "19930082617_p3.jpg", "text": "2\nNACA TN 1962\n\nAll the cylinders previously tested, however, had cutouts of the same length except for three side-cutout specimens. Little information was obtained with regard to the effect of the length of the cutout upon the stress concentration. Furthermore, in all specimens tested, the axial length of the general-instability bulge was either longer than or equal to the length of the cutout. It was conjectured that with longer cutouts the bulge might be shorter than the cutout, and in some cases several waves might even develop.\n\nTherefore it was felt that cylinders with varying lengths of cutout should be tested so as to supply the missing information. It is the purpose of the present report to describe tests which were run on cylinders with long cutouts. As long side cutouts are seldom met in practice, it was thought more advisable to construct cylinders with series of side cutouts. Thus the windows in the fuselage of a transport airplane were simulated.\n\nThe investigation described in this report was carried out at PIBAL under the sponsorship and with the financial assistance of the National Advisory Committee for Aeronautics. Acknowledgment is due to Mr. Sebastian V. Nardo who took part in the initial stages of the experimental work.\n\nTEST SPECIMENS, RIG, AND PROCEDURE\n\nFive cylinders with bottom cutouts and three with symmetrical side cutouts were tested. They were numbered consecutively from 72 to 79. Their characteristics are given in table I and shown in figure 1.\n\nIn all specimens the diameter was 20 inches, the sheet was 0.012 inch in thickness, and the longitudinal reinforcement consisted of sixteen 3/8-inch-square 24S-T aluminum-alloy stringers equally spaced along the circumference on the inside of the sheet. The number of rings varied between 14 and 22 and the lengths of the cylinders between 57.9 and 61.78 inches except for cylinder 77 which had 29 rings and was 77.10 inches long. The rings were either square or rectangular in cross section and were equally spaced on the outside of the sheet. The ring spacing was either 2.57 or 3.86 inches.\n\nRings and stringers were attached to the sheet covering by means of 1/8-inch Al7S-T aluminum-alloy round-head rivets. The rivet spacing was 0.643 inch on the stringers and approximately 1 inch on the rings. The rings and stringers were fastened to each other at their intersections by 1/8-inch steel machine screws except in cylinders 75 and 77 where 3/16-inch steel machine screws were used. Cylinder 77 had heat-treated Allen head bolts at the ring-stringer intersections in the vicinity of the cutout.", "timestamp": "2026-07-22T04:46:38.725000+00:00"}
{"citation_id": "19930085842", "source_url": "https://ntrs.nasa.gov/api/citations/19930085842/downloads/19930085842.pdf", "page_number": 95, "total_pages": 104, "image_filename": "19930085842_p95.jpg", "text": "NACA RM L9C29\n91\n\nPropeller advance-diameter ratio, V/nD\n8\n6\n4\n\nV/nD\n\nResultant drag coefficient, CDR\n6\n4\n2\n0\n-2\n\nCDR\n\nLift coefficient, CL\n2.2\n2.0\n1.8\n1.6\n1.4\n1.2\n\nCL\n\n0.1 0.2 0.3 0.4 0.5\nTorque coefficient, Qc\n\nNATIONAL ADVISORY\nCOMMITTEE FOR AERONAUTICS\n\n(e) $\\alpha_u = 35^\\circ$.\nFigure 44.— Continued.", "timestamp": "2026-07-22T04:46:46.244167+00:00"}
{"citation_id": "19930082614", "source_url": "https://ntrs.nasa.gov/api/citations/19930082614/downloads/19930082614.pdf", "page_number": 4, "total_pages": 36, "image_filename": "19930082614_p4.jpg", "text": "```markdown\n2\nNACA TN 1939\n\nOne means of controlling speed is the use of aerodynamic brakes.\nThis report is concerned with the problem of relating the drag increases\ndue to aerodynamic brakes to the control of forward speed. Problems of\nbuffeting, of changes in stability, of aerodynamic loads, and of changes\nin trim, which may arise when a particular brake is used on an airplane,\nare not considered. If a satisfactory brake is to be selected for an\nairplane, however, the possible occurrence of such phenomena must be\ninvestigated for the airplane and brake combination.\n\nThe increases in drag that result from the use of air brakes have\nbeen measured in wind-tunnel and flight tests for a large variety of\nbrakes. In this report a summary of such drag data is presented. The\neffect of increases in the drag coefficient upon the speed variation\ncalculated for a hypothetical airplane is illustrated. A procedure\nfor calculating the speed of an airplane at any point in arbitrarily\nspecified maneuvers is presented and discussed.\n\nSYMBOLS\n\n$C_{D_n}$ net drag coefficient $\\left(\\frac{D_n}{qS}\\right)$\n\n$\\Delta C_D$ increase in drag coefficient due to the aerodynamic\nbrake $\\left(\\frac{\\text{brake drag}}{qS}\\right)$\n\n$\\Delta C_{D_B}$ drag coefficient due to the aerodynamic brake referred\nto the area of the brake $\\left(\\frac{\\text{brake drag}}{qS_B}\\right)$\n\n$C_L$ lift coefficient $\\left(\\frac{\\text{lift}}{qS}\\right)$\n\n$D_n$ algebraic sum of aerodynamic forces parallel to the direction\nof flight, positive toward the rear, pounds\n\n$K$ deceleration factor $\\left(\\frac{1}{2} \\frac{\\rho g}{W/S} C_{D_n}\\right)$, per foot\n\n$M$ Mach number\n\n$S$ wing area, square feet\n\n$S_B$ maximum projected area of the aerodynamic brake, including\nslots, gaps, and perforations, square feet\n```", "timestamp": "2026-07-22T04:46:47.475320+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 13, "total_pages": 99, "image_filename": "19930082511_p13.jpg", "text": "NACA TN No. 1826\n\nTwo-dimensional closed-open-closed tunnel.- The setup for the two-dimensional closed-open-closed tunnel with a vortex on the center line (fig. 8(a)) is an obvious modification of the corresponding setup for the closed-open tunnel. The same would be true for the off-center vortex except for the necessity of satisfying the condition that the velocities in the upstream and downstream closed regions be equal. Thus, the setup of figure 8(b) provides an upstream perturbation velocity but no downstream perturbation velocity, and cannot, therefore, solve the problem completely. An additional flow, found by the setup of figure 8(c) must be included. An electrode is here located at both the upstream and downstream ends, and the potentials relative to the free boundary are so adjusted that the entrance-lip condition is satisfied. It is apparent that in order to satisfy this condition the downstream current flow will be much greater than the upstream current flow; that is, the downstream perturbation velocity for a contracting or expanding jet is much greater than the upstream perturbation velocity. Because of this difference, a suitable amount of the flow of figure 8(c) may be added to that of figure 8(b) to produce equal upstream and downstream perturbation velocities.\n\nDisplacement of the free surfaces.- The current density normal to the surface of a metal plate representing a free surface is proportional to $\\frac{\\partial \\phi}{\\partial y}$ and corresponds to the local vertical perturbation velocity. The total vertical displacement at a point on the free surface is then given by $\\int \\frac{\\partial \\phi}{\\partial y} dx$ integrated from the entrance lip to the point. In particular, the integral along the entire lower free surface of a closed-open-closed tunnel represents the displacement at the exit lip and it may be measured by means of an ammeter in the line that goes to the lower metal plate.\n\nIf the pressure on the lower free surface can adjust itself so that the displacement at the downstream end is zero, the perturbation velocity at the lower surface will be different from that at the upper surface. If the perturbation velocity on the upper surface is taken as zero, that on the lower surface will be negative, so that the potential on the lower surface must drop uniformly from entrance to exit. Such a variation could be accomplished if the lower surface were represented by a number of short metal strips instead of a single plate.\n\nThree-dimensional closed-open and closed-open-closed tunnels.- The analogies for the three-dimensional tunnels are obvious modifications of those for the two-dimensional tunnels. The free boundary may not be simulated by a single metal cylinder because, as was previously noted, different elements of the free boundary do not have the same potential, although the potential is constant along each element. The free boundary must thus be simulated by a number of longitudinal metal strips, insulated from each other, with a feeler electrode immediately in front of each. When the lifting element lies in the horizontal plane of", "timestamp": "2026-07-22T04:46:50.516793+00:00"}
{"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 2, "total_pages": 78, "image_filename": "19930082618_p2.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T04:46:53.731155+00:00"}
{"citation_id": "19930092013", "source_url": "https://ntrs.nasa.gov/api/citations/19930092013/downloads/19930092013.pdf", "page_number": 20, "total_pages": 21, "image_filename": "19930092013_p20.jpg", "text": "```markdown\n16\nREPORT 948—NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nresponse, spiral divergence) which seem to be important to the pilots over a much wider range than is possible at present. The formulation of more generally applicable conclusions should thereby be made possible.\n\nCONCLUSIONS\n\nA flight investigation to evaluate a device for varying effective dihedral and to determine the tolerable (safe for normal fighter operation) limits of effective dihedral at landing-approach, cruising, and high speeds for a conventional fighter airplane resulted in the following conclusions:\n\n1. The device permits large changes in stick-fixed and stick-free dihedral effect to be made readily in flight. The apparatus exhibited a small amount of lag during dynamic maneuvers which, although perceptible to the pilots, was not considered by them to cause any significant change in the feel of the airplane from that of an airplane with similar dihedral.\n\n2. An effective dihedral as high as $28.4^\\circ$ did not cause the airplane to exhibit intolerable stability and control characteristics at landing-approach speed, even though it caused rolling-velocity reversals in rudder-fixed aileron rolls and even though the airplane was oscillatorily unstable. It appears that this was because the period was long, the rolling velocities experienced in rough air were low, and the rudder was very effective in producing roll.\n\n3. The maximum tolerable effective dihedral at cruising and high speeds was indicated to be about $22^\\circ$. With higher values of dihedral the large and poorly damped rolling motions caused by rough air made the lateral oscillations difficult to control.\n\n4. The minimum tolerable effective dihedral at landing-approach speed was indicated from pilots' opinions formed during flights at altitude to be about $-7^\\circ$ for field landings. With more negative values the adverse rolling response to rudder control (left roll with right rudder) was considered to be dangerously high for an approach.\n\n5. The minimum tolerable effective dihedral at cruising and high speeds was indicated to be about $-5^\\circ$. With more negative values, the rolling response to gusts and the adverse rolling response to rudder control were so rapid that, in rough air, the pilot had to be constantly on the controls, a situation which was considered dangerous from the standpoint of fatigue for flights of normal duration.\n\nREFERENCES\n\n1. Gilruth, R. R.: Requirements for Satisfactory Flying Qualities of Airplanes. NACA Rep. 755, 1943.\n2. Anon.: Flying Qualities of Piloted Airplanes. Spec. No. 1815-B, U. S. Air Force, June 1, 1948.\n3. Anon.: Specification for Stability and Control Characteristics of Piloted Airplanes. Spec. No. SR-119B, Bur. Aero., Navy Dept., June 1, 1948.\n4. Kauffman, William M., Smith, Allan, Liddell, Charles J., Jr., and Cooper, George E.: Flight Tests of an Apparatus for Varying Dihedral Effect in Flight. NACA TN 1788, 1948.\n5. Liddell, Charles J., Jr., Van Dyke, Rudolph, D., Jr., and Heinle, Donovan R.: A Flight Determination of the Tolerable Range of Effective Dihedral on a Conventional Fighter Airplane. NACA TN 1936, 1949.\n6. Pearson, Henry A., and Jones, Robert T.: Theoretical Stability and Control Characteristics of Wings with Various Amount of Taper and Twist. NACA Rep. 635, 1938.\n7. Harper, Charles W., and Jones, Arthur L.: A Comparison of the Lateral Motions Calculated for Tailless and Conventional Airplanes. NACA TN 1154, 1947.\n8. Sternfield, Leonard: Effect of Product of Inertia on Lateral Stability. NACA TN 1193, 1947.\n\nTABLE 1.—VALUES OF POSSIBLE CRITERIA FOR LIMITATION OF POSITIVE EFFECTIVE DIHEDRAL AS MEASURED ON THE TEST AIRPLANE\n\n| $\\Gamma_e$ (deg) | $P$ (sec) | $T_{\\frac{1}{2}}$ (sec) | $\\left| \\frac{\\dot{p}}{\\beta} \\right|$ (per sec) | $\\left| \\frac{\\phi}{\\beta} \\right|$ | Pilot opinion |\n| :--- | :--- | :--- | :--- | :--- | :--- |\n| **Landing-approach condition** | | | | | |\n| 28.4 | 3.6 | (Unstable, $T_2$=38) | 3.8 | 2.3 | Tolerable |\n| 22.7 | 3.9 | 11.5 | 3.2 | 2.0 | Good |\n| 14.2 | 4.4 | 5.2 | 2.1 | 1.4 | Good |\n| 5.3 | 5.2 | 2.4 | 0.5 | 0.4 | Good |\n| **Cruising condition** | | | | | |\n| 24.4 | 3.0 | 8.3 | 11.2 | 5.4 | Intolerable |\n| 18.2 | 3.3 | 5.2 | 9.1 | 4.7 | Tolerable |\n| 12.9 | 3.6 | 3.5 | 5.8 | 3.3 | Good |\n| 6.2 | 4.0 | 2.6 | 2.3 | 1.5 | Good |\n| **High-speed condition** | | | | | |\n| 28.4 | 2.3 | 7.0 | 15.5 | 5.7 | Intolerable |\n| 18.2 | 2.5 | 3.5 | 11.5 | 4.2 | Tolerable |\n| 12.9 | 2.6 | 2.5 | 7.5 | 2.7 | Good |\n| 6.2 | 2.8 | 2.1 | 4.6 | 1.7 | Good |\n\nAMES AERONAUTICAL LABORATORY,\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS,\nMOFFETT FIELD, CALIF., Aug. 24, 1949.\n\nU. S. GOVERNMENT PRINTING OFFICE: 1950\n```", "timestamp": "2026-07-22T04:46:54.304024+00:00"}
{"citation_id": "19930082585", "source_url": "https://ntrs.nasa.gov/api/citations/19930082585/downloads/19930082585.pdf", "page_number": 5, "total_pages": 30, "image_filename": "19930082585_p5.jpg", "text": "4\nNACA TN 1907\n\nRotor Aerodynamic Characteristics\n\n| | |\n| :--- | :--- |\n| T | rotor thrust, pounds |\n| Q | rotor torque, pound-feet |\n| $C_Q$ | torque coefficient $\\left( C_Q = \\frac{Q}{\\pi \\rho R^3 (\\Omega R)^2} \\right)$ |\n\nRotor-Blade Motion\n\n| | |\n| :--- | :--- |\n| $\\beta$ | flapping angle, radians unless otherwise noted |\n\nMiscellaneous\n\n| | |\n| :--- | :--- |\n| k | decay constant in assumed variation of induced velocity with time |\n| $a_1, a_2, A_1, A_2$ | constants in expression for $\\theta$ as a function of time; $A_1$ and $A_2$ in radians unless otherwise noted |\n| $b_1, b_2, b_3$ | coefficients in differential equation for $\\beta(t)$ |\n| $B_1, B_2, B_3, B_4, B_5, B_6, B_7$ | coefficients in complete solution for $\\beta(t)$, radians unless otherwise noted |\n| $m_1, \\alpha, \\omega$ | decay constants in expression for $\\beta(t)$ |\n\nSubscripts:\n\n| | |\n| :--- | :--- |\n| 0 | initial value, for hovering |\n| f | final value, for steady autorotative vertical descent |\n| a | average value $\\left( ( )_a = \\frac{1}{2} [( )_0 + ( )_f] \\right)$ |\n| $\\Delta t$ | value after time interval of $\\Delta t$ seconds |", "timestamp": "2026-07-22T04:46:56.023571+00:00"}
{"citation_id": "19930082485", "source_url": "https://ntrs.nasa.gov/api/citations/19930082485/downloads/19930082485.pdf", "page_number": 21, "total_pages": 62, "image_filename": "19930082485_p21.jpg", "text": "```markdown\n20\nNACA TN No. 1810\n\n$$\nK = 1 + \\frac{a}{4b} \\left\\{ \\left[ \\frac{3}{2} \\left( \\frac{a}{b} \\right) (2b - 1)^2 \\right] + 1 \\right\\} + \\frac{9}{160} \\left( \\frac{a}{b} \\right)^2 \\left\\{ \\frac{4}{3} \\left[ \\left( \\frac{a}{b} \\right) (2b - 1)^2 \\right] \\right.\n$$\n$$\n\\left. + 3 \\left( \\frac{a}{b} \\right) (2b - 1)^2 + 1 \\right\\} + \\frac{9}{4480} \\left( \\frac{a}{b} \\right)^3 \\left\\{ \\frac{9}{8} \\left[ \\left( \\frac{a}{b} \\right) (2b - 1)^2 \\right]^3 \\right.\n$$\n$$\n\\left. + \\frac{90}{8} \\left[ \\frac{a}{b} (2b - 1)^2 \\right]^2 + \\frac{90}{4} \\left( \\frac{a}{b} \\right) (2b - 1)^2 + 5 \\right\\} \\quad (38)\n$$\n\nEquations (37) and (38) apply for negative or positive values of b, that is, for $C_1 > C_2$ or $C_1 < C_2$ but not for $C_1 = C_2$. If $C_1 = C_2$, the expressions for J and K are\n\n$$\nJ = \\frac{\\sinh a}{a}\n$$\n\n$$\nK = \\frac{\\sinh 3a}{3a}\n$$\n\nThese equations can easily be verified by integrating equation 1 with the curvature C constant and substituting in equations (22) and (23).\n\nCharts of J and K/J are shown in figures 15 and 16, respectively, as a function of\n\n$$\na = \\frac{C_1 n_o}{2}\n$$\n\nand\n\n$$\nb = \\frac{-C_1}{AC}\n$$\n\nEvaluation of channel-surface velocities. - The relation between the velocity at the point of the center of the channel and the velocities along the surface of the channel (figs. 17 and 18) can be derived, inasmuch as from equation (21)\n\n1026\n```", "timestamp": "2026-07-22T04:46:57.825229+00:00"}
{"citation_id": "19930082487", "source_url": "https://ntrs.nasa.gov/api/citations/19930082487/downloads/19930082487.pdf", "page_number": 21, "total_pages": 33, "image_filename": "19930082487_p21.jpg", "text": "NACA TN No. 1813\n\nPressure coefficient, P\n\n-2.0\n-1.6\n-1.2\n-.8\n-.4\n0\n\nx/c\nM, 1.0 1.4\nCrest\nx/c\n0.025\n.100\n.230\n.380\nMcr Md Ms\n.4 .6 .8\n\nx/c\nM, 1.0 1.4\n0.075\n.200\n.325\n.475\nMcr Md Ms\n.4 .6 .8\n\nx/c\nM, 0.8 1.0 1.4\n0.050\n.150\n.265\n.415\nMcr Md Ms\n.4 .6 .8\n\nx/c\n* 0.265\n+ .415\nUnpublished\nGerman data\n\nFree-stream Mach number, M₀\n\n(a) α, -4°; lower surface. (b) α, -2°; upper surface. (c) α, 0°; upper surface.\n\nFigure 4.— Pressure coefficient at various chordwise locations as a function of free-stream Mach number for NACA 23015 airfoil section.\n\nNACA\n\n19", "timestamp": "2026-07-22T04:47:00.554587+00:00"}