Buckets:
| {"citation_id": "19930082487", "source_url": "https://ntrs.nasa.gov/api/citations/19930082487/downloads/19930082487.pdf", "page_number": 24, "total_pages": 33, "image_filename": "19930082487_p24.jpg", "text": "22\nNACA TN No. 1813\n\n<!-- Image (101, 100, 852, 858) -->\n\n(a) Airfoil section, NACA 4415.\n\n(b) Airfoil section, NACA 0015.\n\nFigure 6.—Variation of critical, drag-divergence, and shock-stall Mach numbers with angle of attack for several NACA airfoil sections.", "timestamp": "2026-07-22T04:49:19.628710+00:00"} | |
| {"citation_id": "19930082485", "source_url": "https://ntrs.nasa.gov/api/citations/19930082485/downloads/19930082485.pdf", "page_number": 24, "total_pages": 62, "image_filename": "19930082485_p24.jpg", "text": "NACA TN No. 1810\n\n$$\n\\left( \\frac{\\mu n_0}{\\tau} \\right)_1 = \\sqrt{\\frac{\\gamma - 1}{\\gamma + 1}} \\left[ 1 - \\frac{\\gamma - 1}{\\gamma + 1} \\left( \\frac{V}{V_{cr}} \\right)_1^2 \\right]^{\\frac{1}{\\gamma - 1}} \\left( \\frac{V}{V_{cr}} \\right)_1 \\cos \\alpha_1 \\tag{41}\n$$\n\n$$\n\\left( \\frac{\\mu n_0}{\\tau} \\right)_e = \\sqrt{\\frac{\\gamma - 1}{\\gamma + 1}} \\left[ 1 - \\frac{\\gamma - 1}{\\gamma + 1} \\left( \\frac{V}{V_{cr}} \\right)_e^2 \\right]^{\\frac{1}{\\gamma - 1}} \\left( \\frac{V}{V_{cr}} \\right)_e \\cos \\alpha_e \\tag{42}\n$$\n\nWith the flow network of velocity potentials and streamlines drawn and the value of $\\mu$ known for each velocity potential line, the surface velocity may be computed. The values of $\\left( \\frac{V}{V_{cr}} \\right)_1$ and $\\left( \\frac{V}{V_{cr}} \\right)_2$ are plotted against blade-surface length $S_1$ and $S_2$, respectively. (See fig. 22.) The velocity potential $\\phi$ is then determined by graphically integrating the velocity-distribution profile.\n\n$$\n\\phi = \\int V \\, ds \\tag{43}\n$$\n\nThe velocity potential is plotted against blade-surface length $S$ (fig. 22). Inasmuch as the absolute value of $\\phi$ is not important, the value of $\\phi_4$, the velocity-potential line connecting $S_1$ and $S_2$ at the rear stagnation point on $S_2$, is set equal to zero on both $S_1$ and $S_2$. This plot provides a convenient method for relocating the velocity-potential lines, inasmuch as the value of $S_1$ and $S_2$ for constant values of $\\phi$ are easily found.\n\nThe circulation $\\Gamma$ around the airfoil is specified by the vector diagram and is expressed as follows:\n\n$$\n\\begin{aligned}\n\\Gamma &= \\tau \\left[ (V_u)_1 - (V_u)_e \\right] \\\\\n&= \\tau (V_u)_1 - \\tau (V_u)_e \\\\\n&= \\Delta \\Gamma_1 - \\Delta \\Gamma_e \\tag{44}\n\\end{aligned}\n$$", "timestamp": "2026-07-22T04:49:24.643229+00:00"} | |
| {"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 5, "total_pages": 78, "image_filename": "19930082618_p5.jpg", "text": "NACA TN 1945\n3\n\nsufficient amount of data is also included to show the effects of\nleading-edge roughness and split flaps upon the characteristics of the\nairfoils.\n\nSYMBOLS\n\n| | |\n| :--- | :--- |\n| $c_d$ | section drag coefficient |\n| $c_l$ | section lift coefficient |\n| $c_{l_{max}}$ | maximum section lift coefficient |\n| $c_{l_1}$ | section design lift coefficient |\n| $c_{mac}$ | section pitching-moment coefficient about aerodynamic center |\n| $c_{mc/4}$ | section pitching-moment coefficient about quarter-chord point |\n| $\\alpha_0$ | section angle of attack |\n| $\\alpha_{l_0}$ | section angle of zero lift |\n| $dc_l/d\\alpha_0$ | section lift-curve slope |\n| R | Reynolds number |\n| c | airfoil chord |\n| x | distance along chord |\n| y | distance perpendicular to chord |\n\nAIRFOILS\n\nThe airfoils investigated consisted of 10 NACA 6-series sections\nand 5 NACA 4- and 5-digit-series sections. The airfoils were selected\nto show the effect upon the resultant aerodynamic characteristics of\nsystematic variations in thickness, camber, and thickness distribution.", "timestamp": "2026-07-22T04:49:26.635639+00:00"} | |
| {"citation_id": "19930082498", "source_url": "https://ntrs.nasa.gov/api/citations/19930082498/downloads/19930082498.pdf", "page_number": 14, "total_pages": 49, "image_filename": "19930082498_p14.jpg", "text": "NACA TN No. 1838\n\nTABLE I.- ENGINE AND PROPELLER NOISE\n\n| Configuration | Engine speed (rpm) | Engine fundamental frequency, F (cps) | Propeller fundamental frequency, F' (cps) | Sound-pressure level (db) |\n|---------------|--------------------|----------------------------------------|------------------------------------------|----------------------------|\n| | | | | Over-all | F' | F | 2F' | 1.5F | 2F = 3F' | 2.5F | 3F | 4F = 6F' |\n| 1 Original engine installation (See fig. 3) | 2000 | 100 | 66.5 | 98 | 92 | 97 | 80 | 76 | 80 | (a) | 67 | (a) |\n| 2 Exhaust manifold removed | 2000 | 100 | 66.5 | 100 | 90 | 97 | 82 | 78 | 87 | 75 | 68 | 67 |\n\na The sound-pressure level was below the range of the analyzer.\n\n[Figure: NACA logo]\n\n13", "timestamp": "2026-07-22T04:49:29.827958+00:00"} | |
| {"citation_id": "19930082245", "source_url": "https://ntrs.nasa.gov/api/citations/19930082245/downloads/19930082245.pdf", "page_number": 32, "total_pages": 66, "image_filename": "19930082245_p32.jpg", "text": "```markdown\nNACA TN No. 1596\n\n16\n14\n$\\delta_a$\n(deg)\n12\n-12\n10\n-6\n-4\n-2\n0\n2\n4\n8\n6\n4\n2\n0\n-2\n-4\n-6\n.1 .2 .3 .4 .5 .6 .7 .8 .9\nMach number, M\n\nSection angle of attack, $\\alpha$, deg\n\n.16\n$\\delta_a$\n(deg)\n.12\n-12\n.08\n.04\n-6\n-4\n-2\n0\n2\n4\n0\n-.04\n-.08\n-.12\n12\n-.16\n-.20\n18\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(g) $C_n=0.7$.\nFigure 6 .-Continued.\n\n[Figure: NACA logo]\n\n31\n```", "timestamp": "2026-07-22T04:49:35.856565+00:00"} | |
| {"citation_id": "19930085842", "source_url": "https://ntrs.nasa.gov/api/citations/19930085842/downloads/19930085842.pdf", "page_number": 99, "total_pages": 104, "image_filename": "19930085842_p99.jpg", "text": "NACA RM L9C29\n95\n\n[Figure: A multi-panel graph plotting various aerodynamic coefficients against the lift coefficient $C_L$. The graph contains four sub-plots arranged vertically. The x-axis for all plots is labeled \"Lift coefficient, $C_L$\" and ranges from -6 to 4. The top plot shows $V/nD$ on the y-axis (0 to 6). The second plot shows $\\alpha, deg$ on the y-axis (40 to 42). The third plot shows $Q_c$ on the y-axis (0.02 to 0.10). The bottom plot shows $C_{D_R}$ on the y-axis (-6 to 4). A logo for the \"NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\" is present in the bottom right corner of the graph area.]\n\n(a) $\\alpha_u = 42^\\circ$.\n\nFigure 47.- Variation of $C_{D_R}$, $Q_c$, $\\alpha$, and $V/nD$ with $C_L$ for propeller operation at $\\beta = 11.5^\\circ$. Basic model configuration; all control surfaces neutral.", "timestamp": "2026-07-22T04:49:36.053576+00:00"} | |
| {"citation_id": "19930085838", "source_url": "https://ntrs.nasa.gov/api/citations/19930085838/downloads/19930085838.pdf", "page_number": 111, "total_pages": 118, "image_filename": "19930085838_p111.jpg", "text": "NACA RM No. L9B23\n109\n\nIncrement of section angle of attack, $\\Delta\\alpha_o$, deg\n$\\delta_1$ (deg)\nPlain symbols, $c_b = 0.351c_a$\nFlagged symbols, $c_b = 0.408c_a$\nAileron deflection, $\\delta_a$, deg\n(b) 0.177c thick airfoil; $\\delta_f = 25^\\circ$; $c_l = 1.1$.\n\nIncrement of section angle of attack, $\\Delta\\alpha_o$, deg\n$\\delta_1$ (deg)\nPlain symbols, $c_b = 0.351c_a$\nFlagged symbols, $c_b = 0.408c_a$\nAileron deflection, $\\delta_a$, deg\n(c) 0.177c thick airfoil; $\\delta_f = 40^\\circ$; $c_l = 2.0$.\nFigure 16.- Continued.", "timestamp": "2026-07-22T04:49:38.260708+00:00"} | |
| {"citation_id": "19930082496", "source_url": "https://ntrs.nasa.gov/api/citations/19930082496/downloads/19930082496.pdf", "page_number": 15, "total_pages": 50, "image_filename": "19930082496_p15.jpg", "text": "14\nNACA TN No. 1836\n\nSUMMARY OF RESULTS\n\nThe investigations to determine characteristics of a carbide-type ceramal, strength at elevated temperatures, resistance to thermal shock, and performance characteristics when formed into a blade shape and operated under quasi-service conditions, gave the following results:\n\n1. The short-time tensile strengths of 33,200 pounds per square inch at $1800^\\circ$ F and as high as 13,200 pounds per square inch at $2200^\\circ$ F of the particular carbide ceramal compared favorably with similar data for ceramics and currently used alloys. On a strength-to-weight basis, this ceramal was, in general, superior to alloys and ceramics.\n\n2. The thermal-shock resistance of the ceramal was excellent compared with zircon ceramic and good compared with titanium carbide ceramic. The ceramal survived 25 thermal-shock cycles at $1800^\\circ$ F, 25 cycles at $2000^\\circ$ F, 25 cycles at $2200^\\circ$ F, and 25 cycles at $2400^\\circ$ F, whereas the zircon ceramic survived 1 cycle at $1800^\\circ$ F, and the titanium carbide ceramic survived 21 cycles at $2400^\\circ$ F.\n\n3. The quasi-service turbine-blade evaluation using a sample of three ceramal blades and a sample of 12 metal blades resulted in the following data:\n\n (a) At 9 hours and 42 minutes, three ceramal blades had survived whereas six of the metal blades remained; at this time, one ceramal blade was accidentally broken.\n\n (b) At 12 hours and 13 minutes, two of the three ceramal blades remained, whereas a total of only four metal blades remained. At this time, a second ceramal blade had been destroyed because of disk-dovetail failure and the third blade failed in service.\n\n4. During operation, a film consisting of two layers formed on the ceramal blades. The outer layer was mainly titanium dioxide $\\text{TiO}_2$ (rutile). The inner layer consisted mainly of cobalt titanate $\\text{CoTiO}_3$. No significant change occurred in the base material.\n\nCONCLUDING REMARKS\n\nThe investigation yielded data from which it was indicated that:", "timestamp": "2026-07-22T04:49:41.620667+00:00"} | |
| {"citation_id": "19930093773", "source_url": "https://ntrs.nasa.gov/api/citations/19930093773/downloads/19930093773.pdf", "page_number": 28, "total_pages": 47, "image_filename": "19930093773_p28.jpg", "text": "NACA RM E9G09\n27\n\n1159\n\n[Figure: A graph plotting Fuel-air ratio, f/a against Engine speed, N, rpm. The y-axis ranges from 0 to .020. The x-axis ranges from 2 to 8 x 10^3. The graph contains multiple curves with data points marked by different symbols corresponding to altitudes. A NACA logo is present in the bottom right corner of the plot area.]\n\nAltitude\n(ft)\nO 5,000\n□ 15,000\n◊ 25,000\n△ 35,000\n▽ 45,000\n◁ 50,000\n\n(e) Fuel-air ratio.\nFigure 4. - Continued. Effect of altitude on variation of engine performance with engine speed at flight Mach number of 0.21.", "timestamp": "2026-07-22T04:49:49.579540+00:00"} | |
| {"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 16, "total_pages": 99, "image_filename": "19930082511_p16.jpg", "text": "14\nNACA TN No. 1826\n\nTwo-dimensional closed-open-closed tunnel.- The value of $v$ at the downstream end of the free boundary is given by the total flow of current into the metal plate and, in general, is not zero. At the lip of the closed exit, however, $v$ must be suddenly reduced to zero in order for the flow to follow the solid boundary; hence, a short electrode must be added at the exit lip, and as much current must be forced out of it as flows into the long electrode that represents the open boundary; that is, the integral of $\\frac{\\partial u}{\\partial y} dx$ along the free boundary must be canceled at the exit lip. The setup (fig. 10(a)) therefore shows a voltage source to supply this current and means for measuring and equalizing the current flow into adjacent electrodes. If these additional short electrodes are omitted, the setup will correspond to a tunnel the exit section of which has been alined with the deflected jet (fig. 10(b)) because the condition that $\\frac{\\partial u}{\\partial y} = 0$ on the closed exit boundary would merely permit $v$ to remain at the value it had at the end of the free boundary.\n\nFor the off-center position of the lifting surface, a similar setup is used and, as before, probes in the regions far upstream and far downstream are used to determine the potential $u$ in these regions relative to the potential of the free boundary. Since these potentials far upstream and far downstream will not be equal, an additional perturbation field must be provided such that the sum of the two fields will have the same potential in the two regions. This additional perturbation field, which corresponds to a contraction or expansion of the jet, is provided by the setup shown in figure 10(c). It is clear from this figure that the downstream perturbation potential is much greater than the upstream perturbation potential; this result corresponds to that indicated in the velocity-potential analogy.\n\nThe condition in which the pressure on the lower free surface is higher than that on the upper free surface is easily represented by applying a voltage difference between the two surfaces. (See fig. 10(d).) The corresponding displacement of the lower surface, however, is not so readily obtained. The vertical velocity at every point is $\\int \\frac{\\partial u}{\\partial y} dx$, so that the displacement at each point is $\\iint \\frac{\\partial u}{\\partial y} dx dx$. In order to accomplish this integration $\\frac{\\partial u}{\\partial y}$ must be determined at points along the boundary, perhaps by breaking the long plate into a number of short pieces and determining the current flowing into each.\n\nThree-dimensional closed-open and closed-open-closed tunnels.- The analogies for the three-dimensional tunnels are again obvious modifications of those for the two-dimensional tunnels. For the closed-open analogy, the free boundary may be represented by a single cylinder of metal (fig. 11(a)). For the closed-open-closed analogy, the free boundary must be represented by a number of separate strips (fig. 11(b)) in order that the total current into each strip may be measured and an equal", "timestamp": "2026-07-22T04:49:50.736695+00:00"} | |
| {"citation_id": "19930082585", "source_url": "https://ntrs.nasa.gov/api/citations/19930082585/downloads/19930082585.pdf", "page_number": 8, "total_pages": 30, "image_filename": "19930082585_p8.jpg", "text": "NACA TN 1907\n\nand the equation for the transient flapping motion is\n\n$$\nI_1 \\left( \\beta \\Omega_a^2 + \\ddot{\\beta} \\right) = \\frac{\\rho a c R^4 \\Omega_a^2}{2} \\left( \\frac{V - v}{3 \\Omega_a R} - \\frac{\\dot{\\beta}}{4 \\Omega_a} + \\frac{\\theta}{4} \\right)\n\\tag{4}\n$$\n\nSolving equation (4) for $V$, differentiating for $\\dot{V}$, and substituting in equation (3) gives\n\n$$\n\\dddot{\\beta} + b_1 \\ddot{\\beta} + b_2 \\dot{\\beta} + b_3 \\beta = b_4 + \\frac{\\rho a c R^3 \\Omega_a}{6 I_1} \\left( -v + \\frac{3}{4} \\Omega_a R \\dot{\\theta} + \\frac{\\rho a b c R^3 \\Omega_a^2}{4 \\dot{W}/g} \\theta \\right)\n\\tag{5}\n$$\n\nwhere\n\n$$\nb_1 = \\frac{\\rho a c R^3 \\Omega_a}{2 I_1} \\left( \\frac{R}{4} + \\frac{b I_1}{2 R \\dot{W}/g} \\right)\n$$\n\n$$\nb_2 = \\frac{\\rho a c R^3 \\Omega_a}{6 I_1} \\left( \\Omega_a^2 + \\frac{\\rho a b c R^3 \\Omega_a}{4 \\dot{W}/g} \\right)\n$$\n\n$$\nb_3 = \\frac{\\rho a b c R^2 \\Omega_a^3}{4 \\dot{W}/g}\n$$\n\n$$\nb_4 = \\frac{g \\rho a c R^3 \\Omega_a}{6 I_1}\n$$\n\nThe transient solution of equation (5) is of the form\n\n$$\n\\beta = B_1 e^{m_1 t} + e^{\\alpha t} \\left( B_2 \\cos \\omega t + B_3 \\sin \\omega t \\right)\n\\tag{6}\n$$", "timestamp": "2026-07-22T04:49:51.483338+00:00"} | |
| {"citation_id": "19930082914", "source_url": "https://ntrs.nasa.gov/api/citations/19930082914/downloads/19930082914.pdf", "page_number": 2, "total_pages": 66, "image_filename": "19930082914_p2.jpg", "text": "NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nTECHNICAL NOTE NO. 1857\n\nINVESTIGATION WITH AN INTERFEROMETER OF THE TURBULENT MIXING OF A FREE SUPERSONIC JET\n\nBy Paul B. Gooderum, George P. Wood, and Maurice J. Brevoort\n\nSUMMARY\n\nThe free turbulent mixing of a supersonic jet of Mach number 1.6 has been experimentally investigated. An interferometer, of which a description is given, was used for the investigation. Density and velocity distributions through the mixing zone have been obtained. It was found that there was similarity in distribution at the cross sections investigated and that, in the subsonic portion of the mixing zone, the velocity distribution fitted the theoretical distribution for incompressible flow. It was found that the rates of spread of the mixing zone both into the jet and into the ambient air were less than those of subsonic jets.\n\nINTRODUCTION\n\nA considerable amount of work, both theoretical and experimental, has been done in the past on free jets. Most of the theoretical work has been based on Prandtl's concept of the \"mixing length.\" Tollmien (reference 1) and Görtler (reference 2) treated the turbulent mixing of incompressible jets. Abramovich (reference 3) published a theory of the free subsonic jet in which effects of compressibility were included. Experimental results on velocity distribution in the mixing region and rate of spread of the mixing region both into the jet and into the surrounding air have been published by various investigators (references 4 to 8). Most of the experimental work has dealt with constant-density cases; that is, the density of the jet was the same as that of the ambient air with which it mixed, and the velocity of the jet was quite small. The present paper reports results of an experimental investigation of the mixing of a supersonic jet. In this jet the Mach number was 1.6 and the density was about one and one half times the density of the ambient air. Measurements were made with an interferometer of the density variation across the mixing zone in the region near the nozzle. From the density variations the velocity variations have been calculated. A full description of the interferometer is given, together with discussions of the technique of adjusting the interferometer, the theory of fringe formation, and the method of evaluating interferograms.", "timestamp": "2026-07-22T04:49:53.923281+00:00"} | |
| {"citation_id": "19930082542", "source_url": "https://ntrs.nasa.gov/api/citations/19930082542/downloads/19930082542.pdf", "page_number": 9, "total_pages": 53, "image_filename": "19930082542_p9.jpg", "text": "```markdown\n8\nNACA TN No. 1867\n\nas could be determined, a temperature of $2050^\\circ$ or $2100^\\circ$ F was necessary to obtain complete solution of all constituents except the apparently insoluble columbium carbides. Partial grain growth started at $1950^\\circ$ F. Higher temperatures produced uniform grains gradually increasing in size with increasing solution-treating temperature. The grain size was not excessive, however, even after solution-treating at $2300^\\circ$ F.\n\nAging temperature.— (See fig. 7.) Aging treatments at 24 hours had relatively little effect on properties at room temperature. The major effect of aging on the results of rupture tests was to increase ductility. There was a maximum in 100-hour rupture strengths on aging at $1350^\\circ$ to $1400^\\circ$ F although the specific effects varied depending on the prior treatment and rupture time considered. Aging after solution-treating widened the difference between the strengths for rupture in 100 and 1000 hours in most cases. In general, better properties were obtained by aging as the solution temperature increased. The changes in properties seem small in comparison with the changes in microstructure shown in the typical examples in figure 8.\n\nAging time.— (See fig. 9.) Aging at $1400^\\circ$ F after a $2050^\\circ$ F solution treatment improved the rupture strength and ductility for 100 hours at $1200^\\circ$ F but did not improve the strength at 1000 hours. A period of 8 hours at $1400^\\circ$ F produced nearly the maximum effect.\n\nAging at $1400^\\circ$ F after a $2050^\\circ$ F solution treatment produced rupture strengths at 100 hours about equal to those similarly aged after a solution treatment at $2200^\\circ$ F. The ductility in the rupture test was much higher after a $2050^\\circ$ F treatment. The material solution-treated at $2050^\\circ$ F had somewhat lower strengths at 1000 hours than those treated at $2200^\\circ$ F when aged at $1400^\\circ$ F.\n\nApparently aging at $1350^\\circ$ F for 24 hours was slightly more beneficial to 100-hour rupture strength than aging at $1400^\\circ$ F. The 1000-hour strength of the material solution-treated at $2050^\\circ$ F was also improved but not when treated at $2200^\\circ$ F. Aging at $1350^\\circ$ F for 50 hours did not change the properties appreciably over those of material aged for 50 hours at $1400^\\circ$ F.\n\nRoom-temperature properties were not appreciably changed by aging at the indicated times.\n\nTreatment prior to hot-cold-work.— (See fig. 10.) Various solution treatments prior to 15 percent hot-cold-work had very little effect on properties at room temperature. The hot-rolled stock, however, became harder and stronger than the solution-treated stocks.\n\nThe minimum rupture strength occurred when the stock was treated at $1800^\\circ$ F prior to hot-cold-working. Stress for rupture in 1000 hours was independent of solution temperature above $1950^\\circ$ F and higher for the solution-treated material than for the hot-rolled material. The 100-hour strengths were at a maximum when solution-treated at $1950^\\circ$\n```", "timestamp": "2026-07-22T04:49:54.364317+00:00"} | |
| {"citation_id": "19930082712", "source_url": "https://ntrs.nasa.gov/api/citations/19930082712/downloads/19930082712.pdf", "page_number": 3, "total_pages": 14, "image_filename": "19930082712_p3.jpg", "text": "NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nTECHNICAL NOTE 1998\n\nAERODYNAMIC CHARACTERISTICS OF THE\n\nNACA 8-H-12 AIRFOIL SECTION AT SIX REYNOLDS NUMBERS\n\nFROM $1.8 \\times 10^6$ TO $11.0 \\times 10^6$\n\nBy Raymond F. Schaefer and Hamilton A. Smith\n\nSUMMARY\n\nAn investigation has been conducted in the Langley two-dimensional low-turbulence pressure tunnel to determine the aerodynamic characteristics of the NACA 8-H-12 airfoil section at four Reynolds numbers from $3.0 \\times 10^6$ to $11.0 \\times 10^6$. The section lift, drag, and pitching-moment characteristics are presented for both the smooth and rough surface condition at these four Reynolds numbers, together with previously published results for the same section at Reynolds numbers of $1.8 \\times 10^6$ and $2.6 \\times 10^6$. Some of the more important aerodynamic characteristics of the NACA 8-H-12 airfoil are compared with those of two sections commonly used in rotor-blade design, the NACA 0012 and NACA 23012.\n\nThe data indicate that no unusual scale effects on lift, drag, and pitching moment are present for the smooth NACA 8-H-12 airfoil within the range of Reynolds number from $1.8 \\times 10^6$ to $11.0 \\times 10^6$. In general, this is also true for the airfoil with leading-edge roughness.\n\nThe maximum lift coefficient of the smooth NACA 8-H-12 airfoil is lower than those for the NACA 23012 and NACA 0012 sections over the range of Reynolds number tested. Leading-edge roughness on the NACA 8-H-12 airfoil, however, has a less detrimental effect on the maximum lift coefficient than it does on the other two airfoils. The value of the drag coefficient at the design lift coefficient is lower than that for either the NACA 0012 or NACA 23012 section.\n\nINTRODUCTION\n\nSeveral low-drag airfoil sections have been derived solely for use in rotor blades of rotating-wing aircraft. References 1, 2, and 3", "timestamp": "2026-07-22T04:49:56.423454+00:00"} | |
| {"citation_id": "19930082614", "source_url": "https://ntrs.nasa.gov/api/citations/19930082614/downloads/19930082614.pdf", "page_number": 8, "total_pages": 36, "image_filename": "19930082614_p8.jpg", "text": "```markdown\n6\nNACA TN 1939\n\nreport through a study of equations of longitudinal motion of airplanes,\nof the validity of various initial assumptions, and of the accuracy of\napproximate calculations.\n\nAlgebraic Analysis\n\nThe deceleration of an airplane with air brakes depends upon a\nnumber of factors, such as variations of flight-path angle, of alti-\ntude, and of speed, as well as upon the characteristics of the brakes\nthemselves. In order to separate the effects of the brakes from the\nother effects, the problem is simplified by assuming that one or more\nof these factors is constant. The effects of various air-brake\ndesigns and locations on the variation of speed with time may then be\nreadily determined. The following sections present analyses employing\nsuch simplifications.\n\nLevel-flight deceleration.- In level flight, one form of the\nequation of longitudinal motion of an airplane is\n\n$$ \\frac{dV}{dt} = - \\frac{g}{W} D_n \\quad (1) $$\n\nwhere $D_n$ is the algebraic sum of the aerodynamic forces acting\nparallel to the flight direction, positive being taken toward the rear.\nThe value of $D_n$ is affected by any changes in the drag of the air-\nplane, which may result, for example, from variations in angle of\nattack, variations of Mach number, or changes in the setting of the\nair brakes. Any factors that cause the thrust to change, such as\nvariations in engine output, variations in propeller efficiency, and\neffects of velocity upon propeller or engine thrust also influence $D_n$.\nIt is convenient to express $D_n$ in the form of a force coefficient\n\n$$ C_{D_n} = \\frac{D_n}{\\frac{1}{2}\\rho V^2 S} $$\n\nSubstituting in equation (1),\n\n$$ \\frac{dV}{dt} = - C_{D_n} \\frac{\\rho}{2} \\frac{gV^2}{W/S} $$\n\nIf $C_{D_n}$ is constant, this expression may be integrated to give\n\n$$ V = \\frac{1}{Kt + (1/V_0)} \\quad (2) $$\n```", "timestamp": "2026-07-22T04:49:58.212606+00:00"} | |
| {"citation_id": "19930082487", "source_url": "https://ntrs.nasa.gov/api/citations/19930082487/downloads/19930082487.pdf", "page_number": 25, "total_pages": 33, "image_filename": "19930082487_p25.jpg", "text": "NACA TN No. 1813\n23\n\n<!-- Image (142, 105, 887, 457) -->\n\n(c) Airfoil section, NACA 65$_2$-215, a=0.5.\n\n<!-- Image (142, 517, 887, 870) -->\n\n(d) Airfoil section, NACA 66,2-215, a=0.6.\n\nFigure 6.-Concluded.", "timestamp": "2026-07-22T04:49:58.415130+00:00"} | |
| {"citation_id": "19930082476", "source_url": "https://ntrs.nasa.gov/api/citations/19930082476/downloads/19930082476.pdf", "page_number": 23, "total_pages": 41, "image_filename": "19930082476_p23.jpg", "text": "CHART 6.- EFFECT OF INDIVIDUAL AILERON DEFLECTIONS ON THE SPIN CHARACTERISTICS OF MODEL (RUDDERS AND AILERONS UNLINKED)\n[Right erect spins; elevator set to 13° up, rudders set to neutral, ailerons set as indicated]\n\nA. Loading 1 ($\\frac{I_x - I_y}{mb^2} = -3 \\times 10^{-4}$; $\\mu = 5.04$; loading 1 in table II and point 1 in fig. 4)\nB. Loading 2 ($\\frac{I_x - I_y}{mb^2} = -49 \\times 10^{-4}$; $\\mu = 5.89$; loading 1 in table II and point 1 in fig. 4)\nC. Loading 4 ($\\frac{I_x - I_y}{mb^2} = -18 \\times 10^{-4}$; $\\mu = 10.39$; loading 4 in table II and point 4 in fig. 4)\n\n| Right aileron up setting, degrees | Left aileron down setting, degrees |\n| :--- | :--- |\n| **51½** | **No spin** |\n| **21½** | **a**<br>15 2U<br>25 9D<br>161 0.69 |\n| **0** | **No spin** |\n| | **No spin** |\n| | **No spin** |\n| | 0 |\n| | 5 |\n| | 21½ |\n\n| Right aileron up setting, degrees | Left aileron down setting, degrees |\n| :--- | :--- |\n| **51½** | **No spin** |\n| **21½** | **a**<br>15 11D<br>22 14D<br>151 0.59 |\n| **0** | **No spin** |\n| | **No spin** |\n| | **No spin** |\n| | 0 |\n| | 5 |\n| | 21½ |\n\n| Right aileron up setting, degrees | Left aileron down setting, degrees |\n| :--- | :--- |\n| **51½** | **b** |\n| **30** | **a**<br>22 2D<br>32 11D<br>205 0.57 |\n| **21½** | **23 8D**<br>215 0.65 |\n| **10** | **No spin** |\n| **0** | **No spin** |\n| | **No spin** |\n| | **No spin** |\n| | 0 |\n| | 21½ |\n| | 30 |\n\naOscillatory spin, range of values or average value given.\nbSteep spiral.\n\nNACA\nModel values converted to corresponding full-scale values.\nU Inner wing up\nD Inner wing down\n\n| $\\alpha$ (deg) | $g$ (deg) |\n| :--- | :--- |\n| $V$ (fps) | $\\Omega$ (rps) |\n\nNACA TN No. 1801\n21", "timestamp": "2026-07-22T04:49:59.419198+00:00"} | |
| {"citation_id": "19930085965", "source_url": "https://ntrs.nasa.gov/api/citations/19930085965/downloads/19930085965.pdf", "page_number": 65, "total_pages": 67, "image_filename": "19930085965_p65.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T04:50:00.382199+00:00"} | |
| {"citation_id": "19930085842", "source_url": "https://ntrs.nasa.gov/api/citations/19930085842/downloads/19930085842.pdf", "page_number": 100, "total_pages": 104, "image_filename": "19930085842_p100.jpg", "text": "96\nNACA RM L9C29\n\n[Figure: A graph with four subplots plotted on a grid. The x-axis is labeled \"Lift coefficient, $C_L$\" with values from 0 to 4. The y-axes are labeled as follows from top to bottom: \"$\\alpha$, deg\" (46 to 48), \"$Q_c$\" (0.2 to 1.0), \"$C_{D\\pi}$\" (-4 to 6), and \"$\\psi_{HD}$\" (2 to 6). The plot contains four curves labeled $\\alpha$, $Q_c$, $C_{D\\pi}$, and $\\psi_{HD}$. A box in the lower right corner of the grid reads \"NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\".]\n\n(b) $\\alpha_u = 48^\\circ$.\nFigure 47.- Continued.", "timestamp": "2026-07-22T04:50:11.578854+00:00"} | |
| {"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 6, "total_pages": 78, "image_filename": "19930082618_p6.jpg", "text": "4\nNACA TN 1945\n\nThe 10 NACA 6-series airfoils can be grouped as follows to show the systematic variation in design parameters:\n\n| Thickness variation | Camber variation | Thickness-distribution variation |\n| :--- | :--- | :--- |\n| | NACA airfoil | |\n| $64_1$-409 | $64_1$-012 | $63_2$-415 |\n| $64_1$-412 | $64_1$A212 | $64_2$-415 |\n| $64_2$-415 | $64_1$-412 | $65_2$-415 |\n| $64_3$-418 | $64_1$-612 | $66_2$-415 |\n\nThe NACA 64-series thickness form was chosen for the basic investigation of the effects of thickness ratio and camber because, on the basis of the higher Reynolds number results presented in reference 1, this thickness form was believed to represent the best compromise between airfoil lift and drag characteristics in both the smooth and rough surface conditions. The use of an NACA 6A-series thickness form for the investigation of the 12-percent-thick airfoil with 0.2 design lift coefficient was prompted only by the availability of the test model. On the basis of the data presented in reference 3, the use of the slightly modified thickness form in this case would not be expected to alter the validity of the comparison of the more important effects of camber upon the aerodynamic characteristics of the airfoils. Except for the investigation of the effect of variation in amount of camber, the NACA a = 1.0 mean line cambered for a design lift coefficient of 0.4 was used in all cases, since the use of this mean line with 0.4 design lift coefficient generally results in good maximum lift characteristics without causing appreciable increases in the minimum drag or excessive values of the pitching moment (reference 1). Amounts of camber corresponding to design lift coefficients greater than 0.6 were not investigated because previous experience (reference 1) has indicated that such large amounts of camber have an adverse effect upon the drag without causing any marked improvement in maximum lift. Airfoils having thickness ratios not included in the range from 9 to 18 percent of the chord were not investigated because they were not thought to be of very great interest in the design of personal-type airplanes.\n\nThe NACA 4- and 5-digit-series airfoils for which experimental data were obtained are as follows:\n\nNACA 0012\nNACA 4412\nNACA 23012\nNACA 4415\nNACA 23015", "timestamp": "2026-07-22T04:50:12.660023+00:00"} | |
| {"citation_id": "19930082245", "source_url": "https://ntrs.nasa.gov/api/citations/19930082245/downloads/19930082245.pdf", "page_number": 33, "total_pages": 66, "image_filename": "19930082245_p33.jpg", "text": "```markdown\n16\n$\\delta_a$\n(deg)\n-12\n12\n10\n-6\n-4\n-2\n0\n2\n4\n8\n6\n4\n2\n0\n-2\n-4\n-6\n.1 .2 .3 .4 .5 .6 .7 .8 .9\nMach number, M\n\n.16\n$\\delta_a$\n(deg)\n.12\n-12\n.08\n.04\n-6\n-4\n-2\n0\n2\n4\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\nSection pitching-moment coefficient, $c_m$\n\n(h) $c_n = 0.8$.\nFigure 6 .-Concluded.\n\n[Figure: NACA logo]\n\nNACA TN No. 1596\n8\n```", "timestamp": "2026-07-22T04:50:12.873443+00:00"} | |
| {"citation_id": "19930082617", "source_url": "https://ntrs.nasa.gov/api/citations/19930082617/downloads/19930082617.pdf", "page_number": 7, "total_pages": 58, "image_filename": "19930082617_p7.jpg", "text": "6\nNACA TN 1962\n\nCONCLUSIONS\n\nPure bending tests carried out with eight reinforced monocoque\ncylinders substantiated the conjecture that the wave length of the\ngeneral-instability bulge can be smaller than the length of the cutout.\nIt was found that such a situation arises only when the cutout is\nconsiderably longer than most cutouts encountered in actual airplanes.\n\nWhen two specimens were tested which differed only in the length\nof the cutout and the total length of the cylinder, the specimen\nhaving the longer cutout failed at a lower applied load.\n\nPolytechnic Institute of Brooklyn\nBrooklyn, N.Y., July 12, 1948", "timestamp": "2026-07-22T04:50:14.264833+00:00"} | |
| {"citation_id": "19930082447", "source_url": "https://ntrs.nasa.gov/api/citations/19930082447/downloads/19930082447.pdf", "page_number": 24, "total_pages": 24, "image_filename": "19930082447_p24.jpg", "text": "22\n\n$$\n\\begin{array}{c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|", "timestamp": "2026-07-22T04:50:15.914202+00:00"} | |
| {"citation_id": "19930085838", "source_url": "https://ntrs.nasa.gov/api/citations/19930085838/downloads/19930085838.pdf", "page_number": 112, "total_pages": 118, "image_filename": "19930085838_p112.jpg", "text": "110\nNACA RM No. L9B23\n\nIncrement of section angle of attack, $\\Delta\\alpha_o$, deg\n\n$\\delta_1$ (deg)\n$\\circ$ 0\n$\\diamond$ 1.1\n$\\nabla$ 2.2\n$\\triangle$ 3.3\n$\\square$ 4.4\n$\\bullet$ 5.5\n$\\blacksquare$ 10\n$\\blacktriangle$ 15\n$\\blacktriangledown$ 20\n\nPlain symbols, $c_b = 0.351c_a$\nFlagged symbols, $c_b = 0.408c_a$\n\nAileron deflection, $\\delta_a$, deg\n\n(d) 0.154c thick airfoil; $\\delta_f = 25^\\circ$; $c_l = 1.1$.\n\nIncrement of section angle of attack, $\\Delta\\alpha_o$, deg\n\n$\\delta_1$ (deg)\n$\\circ$ 0\n$\\diamond$ 1.1\n$\\nabla$ 2.2\n$\\triangle$ 3.3\n$\\square$ 4.4\n$\\bullet$ 5.5\n$\\blacksquare$ 10\n$\\blacktriangle$ 15\n$\\blacktriangledown$ 20\n$\\times$ 25\n\nPlain symbols, $c_b = 0.351c_a$\nFlagged symbols, $c_b = 0.408c_a$\n\nAileron deflection, $\\delta_a$, deg\n\n(e) 0.154c thick airfoil; $\\delta_f = 40^\\circ$; $c_l = 2.0$.\n\nFigure 18.- Continued.", "timestamp": "2026-07-22T04:50:19.537573+00:00"} | |
| {"citation_id": "19930093773", "source_url": "https://ntrs.nasa.gov/api/citations/19930093773/downloads/19930093773.pdf", "page_number": 29, "total_pages": 47, "image_filename": "19930093773_p29.jpg", "text": "28\nNACA RM E9G09\n\n[Figure: A line graph plotting Exhaust-gas total temperature against Engine speed. The graph includes a legend for Altitude (ft) with symbols: ○ 5,000, □ 15,000, ◇ 25,000, △ 35,000, ▽ 45,000, ◁ 50,000. A dashed line indicates \"Limiting temperature\". The NACA logo is present in the bottom right corner of the plot area.]\n\nExhaust-gas total temperature, $T_7$, °R\n1800\nLimiting temperature\n1600\n1400\n1200\n1000\n800\n2 3 4 5 6 7 8x10³\nEngine speed, N, rpm\n\n(f) Exhaust-gas total temperature.\nFigure 4. - Concluded. Effect of altitude on variation of engine\nperformance with engine speed at flight Mach number of 0.21.\n\n1159", "timestamp": "2026-07-22T04:50:21.876622+00:00"} | |
| {"citation_id": "19930082496", "source_url": "https://ntrs.nasa.gov/api/citations/19930082496/downloads/19930082496.pdf", "page_number": 16, "total_pages": 50, "image_filename": "19930082496_p16.jpg", "text": "NACA TN No. 1836\n15\n\n1. More care should be exercised in the handling of blades made of carbide-type cermals then is customary with blades of alloys.\n\n2. Blades of carbide-type cermals having high thermal conductivities cause the turbine disk rim to run hotter than do metal blades.\n\n3. The scale is tenacious and tends to preserve aerodynamic shapes and to inhibit scaling under conditions of blade operation similar to those reported. For more severe operation, protective coatings against oxidation might be required.\n\n4. The 80-percent titanium carbide plus 20-percent cobalt carbide-type ceramal shows promise for gas-turbine-blade application at relatively high temperatures for short times.\n\nLewis Flight Propulsion Laboratory,\nNational Advisory Committee for Aeronautics,\nCleveland, Ohio, December 6, 1948.\n\nREFERENCES\n\n1. Geller, R. F., and Burdick, M. D.: Progress Report on Strength and Creep of Special Ceramic Bodies in Tension at Elevated Temperatures. NACA ARR No. 6D24, 1946.\n\n2. Anon: Carbides for High Temperature Applications. Materials and Methods, vol. 26, no. 6, Dec. 1947, pp. 85-86.\n\n3. Kopecki, E. S.: Metallurgy. The Iron Age, vol. 161, no. 1, Jan. 1, 1948, pp. 198-207.\n\n4. McKenna, Philip M.: Tantalum Carbide; Its Relation to Other Hard Refractory Compounds. Ind. and Eng. Chem. (Ind. ed.), vol. 28, no. 7, July 1936, pp. 767-772.\n\n5. McKenna, Philip M.: Hard Composition of Matter. U.S. Patent Office No. 2,093,844, Sept. 21, 1937; Method of Producing Hard Compositions of Matter. U. S. Patent Office No. 2,093,845, Sept. 21, 1937; Tungsten Titanium Carbide, WTic2. U. S. Patent Office No. 2,113,353, April 5, 1938; Process of Preparing Tungsten Titanium Carbide. U. S. Patent Office No. 2,113,354, April 5, 1938; Hard Compositions of Matter. U. S. Patent Office No. 2,113,355, April 5, 1938; Process of Making Hard Compositions of Matter. U. S. Patent Office No. 2,113,356, April 5, 1938;", "timestamp": "2026-07-22T04:50:24.681069+00:00"} | |
| {"citation_id": "19930082914", "source_url": "https://ntrs.nasa.gov/api/citations/19930082914/downloads/19930082914.pdf", "page_number": 3, "total_pages": 66, "image_filename": "19930082914_p3.jpg", "text": "2\nNACA TN No. 1857\n\nSYMBOLS\n\na constant\nb fringe spacing; constant\nc constant ($\\sqrt{1/2\\sigma^3}$)\n$c_1$ constant in Snell's law\n$c_2$ constant of integration\n$c_p$ specific heat of air at constant pressure\n$C = \\frac{\\lambda_o}{L(n - 1)}$\nk Gladstone-Dale constant\nl mixing length\nL distance through test section\nn index of refraction of air of density $\\rho$\nn' index of refraction of air of density $\\rho'$\n$p = \\frac{ac_1}{V_o}$\n$q = \\frac{bc_1}{V_o}$\nS(y) dimensionless fringe shift, Y/b, where density is $\\rho'$\nT static temperature of air\n$T_{stag}$ stagnation temperature of air\nu velocity of air\n$u_o$ velocity of air in free stream\nV velocity of light", "timestamp": "2026-07-22T04:50:25.053510+00:00"} | |
| {"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 17, "total_pages": 99, "image_filename": "19930082511_p17.jpg", "text": "NACA TN No. 1826\n\ncurrent forced out of the short strip immediately behind it. Contraction or expansion of the jet is represented as in the two-dimensional case; but the setup that would correspond to different pressures along different strips seems to have no practical significance in the three-dimensional case.\n\nSome form of this acceleration-potential analogy is probably the most convenient for solving problems similar to that of the helicopter in the Langley full-scale tunnel. Simply neglecting the exit, as with a closed-open tunnel, permits the free surface to be represented by a single sheet of metal and eliminates any measurements of current flow to or from the surface. Improved accuracy should be attainable by cutting the sheet into two parts with two short strips at the rear. (See fig. 11(c).) The need for many strips seems unlikely, at least in view of the previously mentioned uncertain definition of the physical flow in the region of the exit.\n\nCorrespondence between velocity-potential and acceleration-potential analogies.- As has already been indicated, the acceleration potential is identical with the x-component of the perturbation velocity and is hence merely the x-derivative of the perturbation-velocity potential. It is of interest to point out the related fact that the acceleration potential analogies are, in a sense, the x-derivatives of the velocity-potential analogies. For example (see fig. 12),\n\n(1) For the velocity-potential analogy, an infinitely long double layer represented a lifting element located at its forward edge. The difference between two such double layers, of which one is shifted slightly relative to the other, is merely the short double layer that was used in the acceleration-potential analogy.\n\n(2) For the velocity-potential analogy, the free boundary consisted of constant-potential strips on which the potentials were so adjusted that the gradient was zero at the leading edge. If each strip is now shifted and subtracted, there remains a long strip, with a short strip at the front and back. Since, in the velocity-potential analogy, the gradient was zero at the entrance lip, the short strip at the front may be neglected. The remainder corresponds to the arrangement used in the acceleration-potential analogy, and the fact that the total current after the subtraction must be zero corresponds to the fact that the total current out of the short strip must be made equal to the total current into the long strip.\n\n(3) When the lifting element was off-center, the velocity-potential analogies required electrodes upstream and downstream, with uniform current flow along the upstream and downstream closed regions. That the subtraction eliminates these current flows corresponds to the fact that no upstream or downstream electrodes are used in the acceleration-potential analogies.", "timestamp": "2026-07-22T04:50:31.820824+00:00"} | |
| {"citation_id": "19930082712", "source_url": "https://ntrs.nasa.gov/api/citations/19930082712/downloads/19930082712.pdf", "page_number": 4, "total_pages": 14, "image_filename": "19930082712_p4.jpg", "text": "2\nNACA TN 1998\n\npresent aerodynamic data for a number of such sections designed to give near-zero pitching moments about the aerodynamic center, low drag over the range of lift coefficient most useful for normal operation, and moderate drag at higher lift coefficients. Because of the present interest in rotors of larger dimensions, it was considered desirable to investigate the aerodynamic characteristics of one of the more promising of these airfoils at Reynolds numbers higher than those at which the former investigations were conducted. The NACA 8-H-12 airfoil was selected on the basis of the generally favorable data given for this airfoil at the lower Reynolds numbers of reference 2. The aerodynamic results for this airfoil, initially tested in the Langley two-dimensional low-turbulence tunnel at Reynolds numbers of $1.8 \\times 10^6$ and $2.6 \\times 10^6$, have therefore been extended to include data for Reynolds numbers of $3.0 \\times 10^6$, $6.0 \\times 10^6$, $9.0 \\times 10^6$, and $11.0 \\times 10^6$ in the present investigation.\n\nThe data given in the present paper were obtained from measurements of the lift, drag, and pitching moments for both smooth and rough surface conditions at the six Reynolds numbers. For comparison, some of the more important aerodynamic parameters of two sections frequently used in rotor blades, the NACA 0012 and NACA 23012, are included. The basic aerodynamic data from which these parameters were taken are given in reference 4.\n\nCOEFFICIENTS AND SYMBOLS\n\n| | |\n| :--- | :--- |\n| $c$ | chord |\n| $c_d$ | section drag coefficient |\n| $c_l$ | section lift coefficient |\n| $c_{l_1}$ | design section lift coefficient |\n| $c_{l_{max}}$ | maximum section lift coefficient |\n| $c_{mac}$ | section pitching-moment coefficient about the aerodynamic center |\n| $c_{mp}$ | section pitching-moment coefficient about the axis on which the airfoil model was pivoted |\n| $dc_l/d\\alpha_o$ | slope of section lift curve per degree |", "timestamp": "2026-07-22T04:50:32.030090+00:00"} | |
| {"citation_id": "19930082487", "source_url": "https://ntrs.nasa.gov/api/citations/19930082487/downloads/19930082487.pdf", "page_number": 26, "total_pages": 33, "image_filename": "19930082487_p26.jpg", "text": "24\nNACA TN No. 1813\n\n$\\alpha$\n(deg)\n$\\circ$ -4\n$\\square$ -2\n$\\diamond$ 0\n$\\triangle$ 2\n$\\nabla$ 4\n$\\triangleright$ 6\n\n$(M_p)_{calc}$\n\nSection drag coefficient, $c_d$\n.12\n.08\n.04\n0\n.6 .7 .8\nFree-stream Mach number, $M_o$\n(a) NACA 64$_1$-206.\n\n.6 .7 .8\n(b) NACA 64$_1$-208.\n\n.12\n.08\n.04\n0\n.6 .7 .8\nFree-stream Mach number, $M_o$\n(c) NACA 64$_1$-210.\n\n.6 .7 .8\n(d) NACA 64$_1$-212.\n\nNACA\n\nFigure 7.- Calculated Mach number for occurrence of sonic velocity at airfoil crest and variation of section drag coefficient with Mach number for a variety of airfoil sections.", "timestamp": "2026-07-22T04:50:35.150521+00:00"} | |
| {"citation_id": "19930082614", "source_url": "https://ntrs.nasa.gov/api/citations/19930082614/downloads/19930082614.pdf", "page_number": 9, "total_pages": 36, "image_filename": "19930082614_p9.jpg", "text": "NACA TN 1939\n\nwhere\n\n$$\nK = C_{Dn} \\frac{\\rho}{2} \\frac{g}{W/S}\n$$\n\nThe quantity $ K $ is a measure of the effects of altitude, wing loading, and net drag coefficient upon the longitudinal acceleration and in this report is referred to as the \"deceleration factor.\" The value of the deceleration factor may be obtained for an airplane from figure 1(a), which takes into account the effects of altitude and of wing loading, and figure 1(b) which includes the effect of the net drag coefficient.\n\nThe velocity variation given by equation (2) has been plotted in figure 2 for four initial speeds: 300, 500, 700, and 900 feet per second. Curves are given for various values of deceleration factor $ K $. Comparison of the slopes of curves for equal values of $ K $ shows a considerable increase in the deceleration as the initial speed increases.\n\nIn deriving equation (2), it was assumed that $ C_{Dn} $ was constant during the time interval considered. For this assumption to be valid, the engine output must be constant and the effects of speed changes on the thrust and drag coefficients must be negligible. If aerodynamic brakes are extended at the start of the time interval, the resulting increase in drag must be so rapid that it can be approximated by an instantaneous increase.\n\nIt is obvious that assuming $ C_{Dn} $ (or $ K $) constant will in some cases lead to large errors. However, the results can represent the actual velocity variations fairly accurately under certain conditions. Below the Mach number of drag divergence the variation of airplane drag coefficient with Mach number is generally small. The effect of a delay due to the time required for full extension of the air brakes could introduce a large discrepancy; however, a rapid rate of extension is an essential characteristic of brakes which are to be used at high speeds.\n\nIt is possible to express variations in $ C_{Dn} $ as a series of instantaneous changes occurring at chosen time intervals during the deceleration. The variation of speed with time is then given by equation (2) with values of $ K $ computed for each interval and values of $ V_0 $ successively determined as the velocity at the end of each preceding interval. The same variation may be obtained from figure 2 by shifting along the time axis portions of the curves having these values of deceleration factor and initial velocity. The curve thus obtained is continuous but changes slope abruptly at the start of each interval.", "timestamp": "2026-07-22T04:50:35.759144+00:00"} | |
| {"citation_id": "19930082542", "source_url": "https://ntrs.nasa.gov/api/citations/19930082542/downloads/19930082542.pdf", "page_number": 10, "total_pages": 53, "image_filename": "19930082542_p10.jpg", "text": "NACA TN No. 1867\n\nto $2050^\\circ$ F; but were lower than for the as-rolled material. Increasing the solution temperature lowered the difference between the stress for rupture in 100 and 1000 hours of hot-cold-worked material.\n\nSolution-treating above $2100^\\circ$ F prior to hot-cold-work resulted in severe stress-concentration brittleness. This was so severe in the material quenched from $2200^\\circ$ F that it was necessary to reduce the diameter of the test specimens in order to avoid fractures in fillets and threads of the specimens.\n\nTemperature of hot-cold-working.— (See fig. 11.) Strength at room temperature fell off as the temperature of hot-cold-working was raised from $1000^\\circ$ to $1800^\\circ$ F. This decrease in resulting strength with increasing temperature of hot-cold-work was most pronounced between $1300^\\circ$ and $1600^\\circ$ F. Ductility at room temperature was not affected by the temperature of working up to $1600^\\circ$ F.\n\nThe temperature of hot-cold-work had practically no effect on the results from rupture tests in the range from $1000^\\circ$ to $1400^\\circ$ F. When worked at higher temperatures the resulting rupture strength was lower and ductility higher.\n\nA solution treatment at $2200^\\circ$ F prior to hot-cold-work resulted in a lower stress for rupture in 100 hours than when the solution treatment was $2050^\\circ$ F. There was practically no difference in the rupture strength at 1000 hours. Elongation in the rupture test was low after both treatments.\n\nIn general, the properties resulting from a reduction of 10 percent at room temperature were not quite so good as when the reduction was at $1200^\\circ$ F.\n\nAmount of hot-cold-work.— (See fig. 12.) Hardness and strength at room temperature were increased and ductility was reduced to a pronounced degree by increasing amounts of hot-cold-work at $1200^\\circ$ F. The indications are, however, that the yield strength may be adversely affected by more than limited amounts of hot-cold-work. Between 10- and 15-percent reduction, depending on the prior treatment, is needed to produce a yield strength of 100,000 psi for 0.02-percent offset.\n\nRupture strengths increased markedly with percent reduction. Both prior treatment and rupture time influenced the relative rupture strengths to some extent. In general, the hot-rolled material had the best strength and ductility although its 1000-hour strength did not increase with percent reduction as much as for the solution-treated materials. Apparently solution treatments in the range from $1950^\\circ$ to $2200^\\circ$ F had relatively little effect on rupture strength, except possibly at the shorter time periods for the material solution-treated at $2200^\\circ$ F and reduced more than 10 percent.", "timestamp": "2026-07-22T04:50:35.952470+00:00"} | |
| {"citation_id": "19930085965", "source_url": "https://ntrs.nasa.gov/api/citations/19930085965/downloads/19930085965.pdf", "page_number": 66, "total_pages": 67, "image_filename": "19930085965_p66.jpg", "text": "NACA RM E9E06\n65\n\n<!-- Image (169, 112, 873, 847) -->\n\nFigure 18. - Calculated unit heat dissipation along inlet-guide-vane chord in icing condition. Blade-surface temperature, 35° F; inlet-air temperature, 0° F; liquid-water content, 1 gram per cubic meter.", "timestamp": "2026-07-22T04:50:36.990963+00:00"} | |
| {"citation_id": "19930082485", "source_url": "https://ntrs.nasa.gov/api/citations/19930082485/downloads/19930082485.pdf", "page_number": 25, "total_pages": 62, "image_filename": "19930082485_p25.jpg", "text": "24\nNACA TN No. 1810\n\nthat is,\n\n$$\n\\Delta \\Gamma_1 = \\tau(V_u)_1\n$$\n\n$$\n\\Delta \\Gamma_e = \\tau(V_u)_e\n$$\n\nThe difference in velocity potential on $S_1$ and that on $S_2$ at the front stagnation point is equal to $\\Delta \\Gamma_1$ and that at the rear stagnation point is equal to $\\Delta \\Gamma_e$ (fig. 22).\n\nInasmuch as the flow is considered irrotational, the line integral of velocity around any closed path in the flow field must equal zero.\n\n$$\n\\int_C V \\, ds = 0 \\tag{45}\n$$\n\nwhere $c$, for convenience, is taken as any area in the channel bounded by two streamlines and two velocity-potential lines. For example,\n\n$$\n\\int_{\\phi_3}^{\\phi_4} V_1 \\, ds_1 + \\int_{\\phi_4}^{\\phi_3} V_2 \\, ds_2 = 0 \\tag{46}\n$$\n\nThe streamlines are so drawn that equal amounts of gas pass between them.\n\n$$\n\\psi = \\int_0^n \\rho V \\, dn = n_0 \\int_0^{\\frac{n}{n_0}} \\rho V \\, d\\left(\\frac{n}{n_0}\\right) \\tag{47}\n$$\n\nThe value of $\\left(\\frac{n}{n_0}\\right)_i$ for $\\psi_i$ ($i = 1, 2, 3, \\dots, j$; where there are $j$ streamlines) is graphically found by plotting $\\rho V$ against $\\frac{n}{n_0}$ (fig. 23) and integrating graphically (fig. 24).", "timestamp": "2026-07-22T04:50:40.859163+00:00"} | |
| {"citation_id": "19930085842", "source_url": "https://ntrs.nasa.gov/api/citations/19930085842/downloads/19930085842.pdf", "page_number": 101, "total_pages": 104, "image_filename": "19930085842_p101.jpg", "text": "NACA RM L9C29\n97\n\n[Figure: A graph with four plots on a grid background. The x-axis is labeled \"Lift coefficient, $C_L$\" with values from 2 to 6. The left y-axis is labeled \"$\\alpha$, deg\" with values 52 and 54, and \"$C_{PR}$\" with values from -6 to 4. The right y-axis has values .4, .2, 0, .12, .08, .04. The plots are labeled $V_{hD}$, $\\alpha$, $Q_c$, and $C_{PR}$. The National Advisory Committee for Aeronautics logo is in the bottom right corner of the graph.]\n\n(c) $\\alpha_u = 54^\\circ$.\n\nFigure 47.- Continued.", "timestamp": "2026-07-22T04:50:43.340557+00:00"} | |
| {"citation_id": "19930082617", "source_url": "https://ntrs.nasa.gov/api/citations/19930082617/downloads/19930082617.pdf", "page_number": 8, "total_pages": 58, "image_filename": "19930082617_p8.jpg", "text": "NACA TN 1962\n7\n\nREFERENCES\n\n1. Hoff, N. J., and Boley, Bruno A.: Stresses in and General Instability of Monocoque Cylinders with Cutouts. I - Experimental Investigation of Cylinders with a Symmetric Cutout Subjected to Pure Bending. NACA TN 1013, 1946.\n\n2. Hoff, N. J., Boley, Bruno A., and Klein, Bertram: Stresses in and General Instability of Monocoque Cylinders with Cutouts. II - Calculation of the Stresses in a Cylinder with a Symmetric Cutout. NACA TN 1014, 1946.\n\n3. Hoff, N. J., Boley, Bruno A., and Klein, Bertram: Stresses in and General Instability of Monocoque Cylinders with Cutouts. III - Calculation of the Buckling Load of Cylinders with Symmetric Cutout Subjected to Pure Bending. NACA TN 1263, 1947.\n\n4. Hoff, N. J., Boley, Bruno A., and Viggiano, Louis R.: Stresses in and General Instability of Monocoque Cylinders with Cutouts. IV - Pure Bending Tests of Cylinders with Side Cutout. NACA TN 1264, 1948.\n\n5. Hoff, N. J., and Klein, Bertram: Stresses in and General Instability of Monocoque Cylinders with Cutouts. V - Calculation of the Stresses in Cylinders with Side Cutout. NACA TN 1435, 1948.\n\n6. Hoff, N. J., Klein, Bertram, and Boley, Bruno A.: Stresses in and General Instability of Monocoque Cylinders with Cutouts. VI - Calculation of the Buckling Load of Cylinders with Side Cutout Subjected to Pure Bending. NACA TN 1436, 1948.", "timestamp": "2026-07-22T04:50:49.666021+00:00"} | |
| {"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 7, "total_pages": 78, "image_filename": "19930082618_p7.jpg", "text": "NACA TN 1945\n\nThese particular airfoils were chosen for investigation because they have been employed quite extensively in the past; hence, a comparison of their merits relative to those of the NACA 6-series sections throughout the range of Reynolds number from $0.7 \\times 10^6$ to $9.0 \\times 10^6$ seemed desirable.\n\nComplete descriptions, including the methods of derivation and theoretical pressure-distribution data, can be found in reference 1 for all the airfoils investigated except the NACA $64_1A212$ section, for which corresponding information is included in reference 3. Ordinates for the 15 airfoils tested are presented in tables I to XV.\n\nAPPARATUS AND TESTS\n\nModels.- The 24-inch-chord models of the airfoil sections tested were constructed of laminated mahogany. The surfaces of the models were lacquered and then sanded with No. 400 carborundum paper.\n\nWind tunnel and test methods.- The experimental investigation was conducted in the Langley two-dimensional low-turbulence tunnel. The test section of this tunnel measures 3 feet by 7.5 feet and, when mounted, the model completely spans the 3-foot dimension. Since this tunnel operates at atmospheric pressure only, the Reynolds number is varied by means of the tunnel airspeed. Lift measurements are usually made in this tunnel by taking the difference of the integrated pressure reaction upon the floor and ceiling of the tunnel (reference 4). Because of the small dynamic pressures involved in the present investigation, however, more accurate measurements of the lift were obtainable with the three-component balance which is part of the equipment of the low-turbulence tunnel. The pitching-moment measurements were also made with the balance.\n\nFor the tests using the balance, the models were supported in the tunnel on trunnions extending through the tunnel walls from the balance frame. A small gap was allowed between the ends of the model and the tunnel walls to insure freedom of movement of the balance. Since air leakage through these gaps was considered as a possible source of error, lift tests were made at Reynolds numbers of $2.0 \\times 10^6$ and $1.5 \\times 10^6$ with the gaps open and then sealed. The measurements for the gaps-sealed condition were made by means of the tunnel floor and ceiling pressure orifices; for the gaps-open tests the balance was used. Results obtained by the two methods agreed to within the experimental error for these Reynolds numbers and would be expected to agree equally well at the lower Reynolds numbers.\n\nSimilar comparative tests have shown, however, that more accurate measurements of the drag are possible with the wake-survey apparatus", "timestamp": "2026-07-22T04:50:51.157140+00:00"} | |
| {"citation_id": "19930085838", "source_url": "https://ntrs.nasa.gov/api/citations/19930085838/downloads/19930085838.pdf", "page_number": 113, "total_pages": 118, "image_filename": "19930085838_p113.jpg", "text": "NACA RM No. L9B23\n111\n\nIncrement of section angle of attack, $\\Delta c_0$, deg\n\n$\\delta_f$ (deg) | Aileron slot | $c_b$\n--- | --- | ---\n0 | Unsealed | $0.408c_a$\n0 | Sealed, not faired | $.408c_a$\n0 | Sealed and faired | $.408c_a$\n-3 | Unsealed | $.351c_a$\n-3 | Sealed | $.351c_a$\n\n(f) $0.154c$ thick airfoil; $\\delta_f = 40^\\circ$; $c_l = 2.0$.\nFigure 18.- Concluded.\n\nIncrement of section angle of attack, $\\Delta c_0$, deg\n\nPlain symbols, $c_b = 0.351c_a$\nFlagged symbols, $c_b = 0.408c_a$\n\n$\\delta_f$ (deg) | $c_l$ | Airfoil\n--- | --- | ---\n25 | 1.1 | $0.154c$ thick\n40 | 2.0 | $.154c$ thick\n25 | 1.1 | $.177c$ thick\n40 | 2.0 | $.177c$ thick\n\nFlap deflection, $\\delta_f$, deg\nNACA\n\nFigure 19.- Lift effectiveness of the flap on the NACA 7-series-type airfoils with straight-sided Frise aileron and double slotted flap. $R = 6 \\times 10^6$ (approx.); $\\delta_a = 0^\\circ$.", "timestamp": "2026-07-22T04:50:51.792018+00:00"} | |
| {"citation_id": "19930082450", "source_url": "https://ntrs.nasa.gov/api/citations/19930082450/downloads/19930082450.pdf", "page_number": 17, "total_pages": 37, "image_filename": "19930082450_p17.jpg", "text": "16\nNACA TN No. 1778\n\nTABLE 5.- 2-PANEL PROPERTIES\n$$ \\left[ \\frac{t_w}{t_B} = 1.00; \\frac{b_w}{t_w} = 8.6; \\frac{t_F}{t_w} = 0.4; \\frac{t_A}{t_w} = 3; \\frac{t_F}{t_w} = 4; \\frac{t_B}{t_B} = 1.95; \\frac{b}{t_w} = 11.7 \\right] $$\n\n<!-- Table (130, 110, 927, 937) -->\n\\begin{tabular}{|c|c|c|c|c|c|c|c|c|c|c|c|c|c|}\n\\hline\n\\multicolumn{2}{|c|}{} & \\multicolumn{12}{c|}{$\\frac{b}{t_w}$} \\\\\n\\cline{3-14}\n\\multicolumn{2}{|c|}{} & 20 & 21 & 22 & 23 & 24 & 25 & 26 & 27 & 28 & 29 & 30 & 31 & 32 \\\\\n\\hline\n\\multirow{12}{*}{$\\frac{t_F}{t_w}$} & 25 & 2.247 & 2.295 & 2.342 & 2.385 & 2.427 & 2.467 & 2.505 & 2.541 & 2.575 & 2.607 & 2.637 & 2.665 & 2.691 \\\\\n& 26 & 2.276 & 2.320 & 2.363 & 2.403 & 2.441 & 2.477 & 2.511 & 2.543 & 2.573 & 2.601 & 2.627 & 2.651 & 2.673 \\\\\n& 27 & 2.293 & 2.230 & 2.372 & 2.408 & 2.442 & 2.474 & 2.504 & 2.532 & 2.558 & 2.582 & 2.604 & 2.624 & 2.642 \\\\\n& 28 & 2.295 & 2.330 & 2.363 & 2.393 & 2.421 & 2.447 & 2.471 & 2.493 & 2.513 & 2.531 & 2.547 & 2.562 & 2.575 \\\\\n& 29 & 2.284 & 2.312 & 2.338 & 2.362 & 2.384 & 2.404 & 2.422 & 2.439 & 2.454 & 2.468 & 2.480 & 2.491 & 2.501 \\\\\n& 30 & 2.262 & 2.284 & 2.304 & 2.322 & 2.338 & 2.353 & 2.366 & 2.378 & 2.389 & 2.399 & 2.408 & 2.416 & 2.423 \\\\\n& 31 & 2.230 & 2.247 & 2.262 & 2.276 & 2.288 & 2.299 & 2.309 & 2.318 & 2.326 & 2.333 & 2.340 & 2.346 & 2.351 \\\\\n& 32 & 2.190 & 2.202 & 2.213 & 2.223 & 2.232 & 2.240 & 2.247 & 2.254 & 2.260 & 2.265 & 2.270 & 2.274 & 2.278 \\\\\n& 33 & 2.143 & 2.151 & 2.158 & 2.165 & 2.171 & 2.176 & 2.181 & 2.185 & 2.189 & 2.193 & 2.196 & 2.199 & 2.202 \\\\\n& 34 & 2.090 & 2.095 & 2.099 & 2.103 & 2.107 & 2.110 & 2.113 & 2.116 & 2.118 & 2.120 & 2.122 & 2.124 & 2.125 \\\\\n& 35 & 2.032 & 2.034 & 2.036 & 2.038 & 2.040 & 2.041 & 2.042 & 2.043 & 2.044 & 2.045 & 2.046 & 2.047 & 2.047 \\\\\n& 36 & 1.970 & 1.970 & 1.970 & 1.970 & 1.970 & 1.970 & 1.970 & 1.970 & 1.970 & 1.970 & 1.970 & 1.970 & 1.970 \\\\\n\\hline\n\\multirow{12}{*}{$\\frac{t_F}{t_w}$} & 25 & 1.803 & 1.850 & 1.895 & 1.938 & 1.979 & 2.018 & 2.055 & 2.090 & 2.123 & 2.154 & 2.183 & 2.210 & 2.236 \\\\\n& 26 & 1.821 & 1.860 & 1.899 & 1.935 & 1.969 & 2.001 & 2.031 & 2.059 & 2.085 & 2.109 & 2.131 & 2.151 & 2.169 \\\\\n& 27 & 1.824 & 1.856 & 1.885 & 1.912 & 1.937 & 1.960 & 1.981 & 2.000 & 2.017 & 2.033 & 2.047 & 2.060 & 2.071 \\\\\n& 28 & 1.815 & 1.841 & 1.864 & 1.885 & 1.904 & 1.921 & 1.937 & 1.951 & 1.964 & 1.975 & 1.985 & 1.994 & 2.002 \\\\\n& 29 & 1.796 & 1.816 & 1.834 & 1.850 & 1.864 & 1.877 & 1.889 & 1.899 & 1.908 & 1.916 & 1.923 & 1.929 & 1.935 \\\\\n& 30 & 1.768 & 1.784 & 1.798 & 1.810 & 1.821 & 1.831 & 1.840 & 1.848 & 1.855 & 1.861 & 1.866 & 1.871 & 1.875 \\\\\n& 31 & 1.732 & 1.744 & 1.754 & 1.763 & 1.771 & 1.778 & 1.784 & 1.790 & 1.795 & 1.799 & 1.803 & 1.806 & 1.809 \\\\\n& 32 & 1.690 & 1.698 & 1.705 & 1.711 & 1.716 & 1.721 & 1.725 & 1.729 & 1.732 & 1.735 & 1.738 & 1.740 & 1.742 \\\\\n& 33 & 1.643 & 1.648 & 1.652 & 1.656 & 1.659 & 1.662 & 1.665 & 1.667 & 1.669 & 1.671 & 1.673 & 1.674 & 1.675 \\\\\n& 34 & 1.592 & 1.594 & 1.596 & 1.598 & 1.599 & 1.600 & 1.601 & 1.602 & 1.603 & 1.603 & 1.604 & 1.604 & 1.605 \\\\\n& 35 & 1.538 & 1.538 & 1.538 & 1.538 & 1.538 & 1.538 & 1.538 & 1.538 & 1.538 & 1.538 & 1.538 & 1.538 & 1.538 \\\\\n& 36 & 1.482 & 1.480 & 1.478 & 1.476 & 1.474 & 1.472 & 1.470 & 1.468 & 1.466 & 1.464 & 1.462 & 1.460 & 1.458 \\\\\n\\hline\n\\multirow{12}{*}{$\\frac{t_F}{t_w}$} & 25 & 6.576 & 7.044 & 7.519 & 8.003 & 8.493 & 8.990 & 9.493 & 10.00 & 10.52 & 11.03 & 11.54 & 12.06 & 12.62 \\\\\n& 26 & 6.473 & 6.895 & 7.308 & 7.708 & 8.097 & 8.476 & 8.846 & 9.208 & 9.563 & 9.910 & 10.25 & 10.58 & 10.90 \\\\\n& 27 & 6.316 & 6.685 & 7.038 & 7.377 & 7.704 & 8.020 & 8.326 & 8.623 & 8.912 & 9.193 & 9.467 & 9.734 & 9.994 \\\\\n& 28 & 6.110 & 6.432 & 6.738 & 7.030 & 7.310 & 7.579 & 7.838 & 8.088 & 8.330 & 8.564 & 8.791 & 9.011 & 9.224 \\\\\n& 29 & 5.860 & 6.138 & 6.402 & 6.654 & 6.895 & 7.126 & 7.348 & 7.562 & 7.768 & 7.967 & 8.159 & 8.345 & 8.525 \\\\\n& 30 & 5.570 & 5.808 & 6.033 & 6.247 & 6.452 & 6.648 & 6.837 & 7.019 & 7.194 & 7.363 & 7.526 & 7.683 & 7.835 \\\\\n& 31 & 5.245 & 5.446 & 5.636 & 5.817 & 5.990 & 6.156 & 6.316 & 6.470 & 6.619 & 6.762 & 6.901 & 7.035 & 7.164 \\\\\n& 32 & 4.890 & 5.057 & 5.215 & 5.365 & 5.509 & 5.647 & 5.780 & 5.908 & 6.032 & 6.151 & 6.266 & 6.377 & 6.484 \\\\\n& 33 & 4.510 & 4.647 & 4.777 & 4.901 & 5.020 & 5.134 & 5.244 & 5.350 & 5.452 & 5.550 & 5.645 & 5.736 & 5.824 \\\\\n& 34 & 4.110 & 4.220 & 4.324 & 4.423 & 4.518 & 4.609 & 4.697 & 4.781 & 4.862 & 4.940 & 5.015 & 5.087 & 5.156 \\\\\n& 35 & 3.690 & 3.775 & 3.855 & 3.931 & 4.004 & 4.074 & 4.141 & 4.205 & 4.267 & 4.326 & 4.383 & 4.437 & 4.489 \\\\\n& 36 & 3.260 & 3.322 & 3.380 & 3.435 & 3.487 & 3.537 & 3.585 & 3.630 & 3.673 & 3.714 & 3.753 & 3.790 & 3.825 \\\\\n\\hline\n\\multirow{12}{*}{$\\frac{t_F}{t_w}$} & 25 & 7.258 & 7.720 & 8.179 & 8.634 & 9.086 & 9.535 & 9.982 & 10.43 & 10.87 & 11.31 & 11.74 & 12.18 & 12.61 \\\\\n& 26 & 7.092 & 7.506 & 7.910 & 8.304 & 8.689 & 9.066 & 9.435 & 9.797 & 10.15 & 10.50 & 10.84 & 11.18 & 11.51 \\\\\n& 27 & 6", "timestamp": "2026-07-22T04:50:52.516900+00:00"} | |
| {"citation_id": "19930085471", "source_url": "https://ntrs.nasa.gov/api/citations/19930085471/downloads/19930085471.pdf", "page_number": 1, "total_pages": 28, "image_filename": "19930085471_p1.jpg", "text": "```markdown\nNACA RM No. L811\n\nUNCLASSIFIED\n~~CONFIDENTIAL~~\nRESTRICTED\n\nNACA-RM-L8J11\nCopy No. 24\nRM No. L811\n\nUNCLASSIFIED\nUNCLASSIFIED\n~~RESTRICTED~~\n\nNACA\nNACA No. 46.9-713153\nElectro. Roll #31, data\n11-1-53\nDate\n114154\nD.E. Newlan\n[signature]\n\nRESEARCH MEMORANDUM\n\nINITIAL EXPERIMENTS ON FLUTTER OF UNSWEPT\nCANTILEVER WINGS AT MACH NUMBER 1.3\n\nBy\nW. J. Tuovila, John E. Baker, and Arthur A. Regier\n\nLangley Aeronautical Laboratory\nLangley Field, Va.\n\nClassification changed to\nUNCLASSIFIED\nAuthority Jan. 13, 1954\nNACA\nResearch Abstract #56\nDate\nFEB 24 1954\nBy\nD.E. Newlan /sa\n\nCLASSIFIED DOCUMENT\nThis document contains classified information\naffecting the National Defense of the United\nStates within the meaning of the Espionage Act,\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 proven\nloyalty and discretion who of necessity must be\ninformed thereof.\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\nWASHINGTON\nJanuary 6, 1949\n\nJAN 11 1949\n\n~~CONFIDENTIAL~~\n~~RESTRICTED~~ UNCLASSIFIED\nUNCLASSIFIED\n\n[Stamp: NACA LIBRARY\nLANGLEY MEMORIAL AERONAUTICAL LABORATORY\nLANGLEY FIELD, VA.]\n```", "timestamp": "2026-07-22T04:50:53.890793+00:00"} | |
| {"citation_id": "19930093773", "source_url": "https://ntrs.nasa.gov/api/citations/19930093773/downloads/19930093773.pdf", "page_number": 30, "total_pages": 47, "image_filename": "19930093773_p30.jpg", "text": "NACA RM E9G09\n29\n\nNet thrust, $F_n$, lb\n\n| Flight Mach number | |\n| :--- | :--- |\n| O | 0.21 |\n| $\\square$ | .53 |\n| $\\diamond$ | .72 |\n| $\\triangle$ | .85 |\n| $\\nabla$ | .97 |\n\n[Figure: Graph plotting Net thrust vs Engine speed with multiple curves corresponding to different Flight Mach numbers]\n\nEngine speed, N, rpm\n\nNACA\n\n(a) Net thrust.\n\nFigure 5. - Effect of flight Mach number on variation of engine performance with engine speed at altitude of 25,000 feet.", "timestamp": "2026-07-22T04:50:54.087363+00:00"} | |
| {"citation_id": "19930082585", "source_url": "https://ntrs.nasa.gov/api/citations/19930082585/downloads/19930082585.pdf", "page_number": 9, "total_pages": 30, "image_filename": "19930082585_p9.jpg", "text": "8\nNACA TN 1907\n\nwhere the roots of\n$$m^3 + b_1m^2 + b_2m + b_3 = 0 \\tag{7}$$\nare\n$$\n\\left.\n\\begin{aligned}\nm &= m_1 \\\\\nm &= \\alpha \\pm j\\omega\n\\end{aligned}\n\\right\\} \\tag{7a}\n$$\nwhere\n$$j = \\sqrt{-1}$$\nThe steady-state solutions have the same form as the forcing functions, and together must satisfy equation (5). Thus the steady-state solutions are:\n$$\\beta = B_4 + B_5e^{-kt} + B_6e^{-a_1t} + B_7e^{-a_2t} \\tag{8}$$\nwhere, by substitution into equation (5) and equating coefficients of identical terms,\n$$B_4 = \\beta_f \\tag{8a}$$\n$$B_5 = \\left( \\frac{\\rho acR^3\\Omega_a}{6I_1} \\right) \\frac{k(v_f - v_0)}{(k^3 - b_1k^2 + b_2k - b_3)} \\tag{8b}$$\n$$B_6 = \\left( \\frac{\\rho acR^3\\Omega_a}{6I_1} \\right) \\left( \\frac{\\frac{3}{4}\\Omega_a RA_1a_1 - \\frac{\\rho abcR^3\\Omega_a^2}{4\\delta M/g}A_1}{a_1^3 - b_1a_1^2 + b_2a_1 - b_3} \\right) \\tag{8c}$$", "timestamp": "2026-07-22T04:50:58.471286+00:00"} | |
| {"citation_id": "19930082496", "source_url": "https://ntrs.nasa.gov/api/citations/19930082496/downloads/19930082496.pdf", "page_number": 17, "total_pages": 50, "image_filename": "19930082496_p17.jpg", "text": "16\nNACA TN No. 1836\n\nHard Compositions of Matter. U. S. Patent Office No. 2,123,574,\nJuly 12, 1938; Hard Compositions of Matter. U. S. Patent Office\nNo. 2,123,575, July 12, 1938; Hard Compositions of Matter.\nU. S. Patent Office No. 2,123,576, July 12, 1938; Carbides of\nTantalum and Like Metals and Method of Producing the Same. U. S.\nPatent Office No. 2,124,509, July 19, 1938.\n\n6. Schmid, R.: Ceramic Materials for Machine Parts Exposed to High\nHeat. Trans. No. F-TS-955-RE, Air Materiel Command, Army Air\nForces, March 1947.\n\n7. Anon.: Test Procedures for Ceramic Materials for Application to\nAircraft Power Plants. The Tensile Test. U. S. Air Force\n(Wright Field), Jan. 1948.\n\n8. Hoffman, Charles A., and Ault, G. Merwin: Application of Statis-\ntical Methods to Study of Gas-Turbine Blade Failures. NACA TN\nNo. 1603, 1948.\n\n9. Grant, Nicholas J., Frederickson, A. F., and Teylor, M. E.: A\nSummary of Heat Resistant Alloys from 1200° F to 1600° F.\nThe Iron Age, vol. 161, no. 12, March 18, 1948, pp. 73-78.\n\n10. Russell, Ralston, Jr.: Zircon Porcelain - A Modern Ceramic.\nWestinghouse Eng., vol. 6, no. 3, May 1946, pp. 90-95.\n\nLIBRARY\nOffice of\nAeronautical Intelligence\nNational Advisory Committee\nfor Aeronautics", "timestamp": "2026-07-22T04:51:02.195406+00:00"} | |
| {"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 18, "total_pages": 99, "image_filename": "19930082511_p18.jpg", "text": "16\nNACA TN No. 1826\n\nTechnical Difficulties\n\nIt should be pointed out that the analogies here described may be rather unwieldy, experimentally. Even for the simplest types of analogies, the literature indicates considerable uncertainty as to the most satisfactory electrolyte and electrode materials, and appreciable difficulty in balancing capacitances (alternating current is generally used in analogies, in order to minimize polarization at the electrodes). In the present analogies, the need for separate current sources that are exactly in phase and the large capacitances that will certainly characterize the vortex and the open-boundary representations should greatly complicate the technique. Perhaps the use of direct current instead of alternating current, with nonpolarizing electrodes (as platinized platinum), would be a more practical approach in this respect. Simultaneously satisfying the entrance-lip condition at a number of points around the inlet (or satisfying the corresponding exit condition for the acceleration-potential analogies) may also turn out to be very difficult.\n\nRÉSUMÉ OF PART I\n\nThe most significant points of the preceding discussion of open wind tunnels and their electrical analogies are as follows:\n\n1. Continuity of velocity at the lip of the entrance cone is a basic characteristic of the flow in an open wind tunnel.\n\n2. Equality of the velocities in the upstream and downstream closed regions is a further condition on the tunnel flow if extraneous longitudinal pressure gradients are to be avoided.\n\n3. The velocity on the free surface need not be the same as the velocity in the closed upstream region. In general, the two velocities are the same only when the lifting element lies in the plane of symmetry of the tunnel.\n\n4. For the two-dimensional open tunnel the velocities on the two free surfaces need not be equal. If the space below the lower free surface is closed off, the pressure on the lower free surface will adjust itself so that the displacement at the exit lip is zero.\n\n5. Considerable uncertainty exists with regard to conditions at the exit or the mathematical equivalents of these conditions. Correspondingly, certain compromises in complying with the idealized downstream conditions may be justified in a determination of boundary interference.\n\n6. In any analysis that neglects the closed entrance and exit regions, the condition of zero upstream induced flow must be retained.", "timestamp": "2026-07-22T04:51:07.486019+00:00"} | |
| {"citation_id": "19930082498", "source_url": "https://ntrs.nasa.gov/api/citations/19930082498/downloads/19930082498.pdf", "page_number": 15, "total_pages": 49, "image_filename": "19930082498_p15.jpg", "text": "```markdown\nTABLE II.- RESULTS OF MUFFLER INVESTIGATION MADE WITH PROPELLER REMOVED\n\n| Muffler configuration | Engine speed (rpm) | Fundamental firing frequency, F (cps) | Sound-pressure level (db) | | | | | | | | | | | | | | | Other sounds | | | Back pressure |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| (a) | | | Over-all | 0.2F | 1.0F | 1.2F | 2.0F | 2.2F | 3.0F | 3.2F | 4.0F | 4.2F | 5.0F | 5.2F | 6.0F | 6.2F | 7.0F | cps | db | cps | db | |\n| 1 Original exhaust manifold with tapered exit cones (see fig. 3) | 1690<br>2000<br>2790 | 82.5<br>100.0<br>139.5 | 97.0<br>98.0<br>97.5 | 58<br>68<br>77 | 97<br>98<br>97 | 81<br>87<br>87 | 81<br>81<br>81 | 70<br>79<br>75 | 70<br>68<br>70 | 69<br>67<br>73 | 65<br>65<br>72 | 63<br>65<br>67 | 63<br>63<br>66 | 58<br>63<br>63 | 58<br>62<br>63 | 57<br>60<br>63 | 55<br>57<br>63 | 57<br>57<br>63 | ----<br>----<br>---- | ----<br>----<br>---- | ----<br>----<br>---- | Medium |\n| 2 Original exhaust manifold with exit cones removed (see fig. 4) | 1690<br>2000<br>2790 | 82.5<br>100.0<br>139.5 | 97.5<br>99.0<br>100.0 | 57<br>65<br>75 | 97<br>98<br>98 | 81<br>75<br>90 | 86<br>86<br>88 | 70<br>75<br>87 | 70<br>78<br>73 | 68<br>72<br>70 | 66<br>63<br>67 | 63<br>63<br>66 | 57<br>63<br>65 | 58<br>63<br>65 | 55<br>60<br>65 | 55<br>57<br>63 | 55<br>57<br>63 | ----<br>----<br>---- | ----<br>----<br>---- | ----<br>----<br>---- | Low |\n| 3 Exhaust manifold removed | 1690<br>2000<br>2790 | 82.5<br>100.0<br>139.5 | 98.0<br>98.5<br>102.0 | 65<br>67<br>65 | 98<br>97<br>99 | 90<br>86<br>88 | 86<br>86<br>86 | 70<br>73<br>75 | 72<br>73<br>72 | 77<br>78<br>71 | 72<br>71<br>72 | 75<br>71<br>72 | 65<br>65<br>73 | 65<br>71<br>73 | 63<br>71<br>73 | 64<br>71<br>73 | 65<br>70<br>73 | 619<br>130<br>---- | 63<br>66<br>---- | ----<br>----<br>---- | ----<br>----<br>---- | ---- |\n| 4 Open vye | 1690<br>2000<br>2790 | 82.5<br>100.0<br>139.5 | 91.5<br>78.0<br>99.0 | 20<br>61<br>55 | 63<br>67<br>91 | 63<br>74<br>67 | 89<br>74<br>88 | 67<br>67<br>87 | 75<br>70<br>70 | 71<br>60<br>70 | 75<br>58<br>79 | 75<br>49<br>81 | 65<br>59<br>79 | 65<br>59<br>81 | 57<br>54<br>75 | 55<br>54<br>57 | 55<br>54<br>57 | 260<br>790<br>375<br>230 | 77<br>146<br>88<br>72 | 74<br>275<br>320<br>47 | 63<br>84<br>84<br>71 | Low |\n| 5 Open vye with 126-inch exhaust pipe | 1690<br>2000<br>2790 | 82.5<br>100.0<br>139.5 | 93.0<br>94.5<br>104.5 | (b)<br>(b)<br>(b) | 93<br>93<br>104 | 65<br>67<br>(b) | 81<br>88<br>95 | 65<br>70<br>90 | 75<br>73<br>95 | 68<br>68<br>89 | 65<br>78<br>84 | 65<br>72<br>72 | 65<br>65<br>77 | 65<br>65<br>72 | (b)<br>(b)<br>(b) | (b)<br>(b)<br>(b) | (b)<br>(b)<br>(b) | 790<br>130 | 70<br>1230 | ----<br>84 | ----<br>---- | Medium |\n| 6 Open vye with 90° upturned elbow | 1690<br>2000<br>2790 | 82.5<br>100.0<br>139.5 | 86.0<br>88.0<br>97.5 | 45<br>50<br>60 | 78<br>86<br>96 | 61<br>63<br>73 | 84<br>84<br>91 | 65<br>70<br>73 | 75<br>84<br>73 | 66<br>69<br>75 | 73<br>63<br>75 | 67<br>63<br>75 | 61<br>63<br>79 | 57<br>58<br>80 | 55<br>57<br>72 | 55<br>57<br>72 | 53<br>57<br>72 | 275<br>465<br>---- | 72<br>80<br>---- | 750<br>800<br>---- | 69<br>61<br>---- | Medium |\n| 7 Two commercial airplane mufflers (see fig. 2(a)) | 2000 | 100.0 | 95.0 | 77 | 90 | 89 | 86 | 86 | 79 | 81 | 65 | 65 | 70 | 65 | 68 | 68 | (b) | ---- | ---- | ---- | ---- | ---- |\n| 8 One commercial airplane muffler-heater and one muffler (see fig. 2(a)) | 1690<br>2000<br>2790 | 82.5<br>100.0<br>139.5 | 89.0<br>95.0<br>94.0 | 77<br>86<br>89 | 89<br>86<br>90 | 87<br>95<br>90 | 77<br>95<br>87 | 71<br>81<br>75 | 73<br>87<br>66 | 67<br>76<br>68 | 65<br>70<br>71 | 63<br>67<br>65 | 65<br>68<br>68 | 61<br>68<br>63 | 55<br>60<br>60 | 58<br>60<br>60 | 58<br>60<br>60 | 619<br>790<br>---- | 59<br>62<br>---- | ----<br>----<br>---- | ----<br>----<br>---- | Low |\n| 9 Two commercial airplane mufflers (see fig. 2(a)) | 1690<br>2000<br>2790 | 82.5<br>100.0<br>139.5 | 92.0<br>96.0<br>98.0 | 76<br>84<br>89 | 91<br>84<br>93 | 91<br>95<br>92 | 73<br>81<br>93 | 74<br>88<br>72 | 68<br>86<br>72 | 69<br>86<br>68 | 63<br>71<br>73 | 62<br>71<br>74 | 65<br>66<br>73 | 59<br>69<br>72 | 55<br>69<br>67 | 55<br>69<br>67 | 55<br>69<br>70 | 619<br>790<br>---- | 59<br>62<br>---- | 860<br>860<br>---- | 59<br>61<br>---- | Low |\n| 10 Two commercial airplane muffler-heaters (see fig. 2(a)) | 1690<br>2000<br>2790 | 82.5<br>100.0<br>139.5 | 89.0<br>93.0<br>95.5 | 81<br>70<br>78 | 86<br>87<br>81 | 87<br>91<br>88 | 81<br>79<br>79 | 77<br>88<br>70 | 72<br>88<br>62 | 69<br>79<br>60 | 67<br>69<br>63 | 67<br>69<br>65 | 67<br>69<br>65 | 67<br>69<br>67 | 55<br>68<br>65 | 55<br>68<br>65 | 55<br>68<br>65 | 619<br>790<br>---- | 59<br>62<br>---- | 860<br>860<br>---- | 59<br>61<br>---- | Low |\n| 11 [Figure: Flow diagram showing dimensions 5, 10.4, 5.5, 26, 14.1] <br> 2 mufflers, one on each bank of cylinders (see fig. 1) | 2000 | 100.0 | 92.0 | 78 | 91 | 84 | (b) | 81 | (b) | 67 | (b) | (b) | 70 | (b) | (b) | (b) | (b) | ---- | ---- | ---- | ---- | Low |\n| 12 1 muffler, same as 11 | 2000 | 100.0 | 89.0 | 55 | 77 | 60 | 81 | 60 | 81 | 59 | 58 | 55 | 61 | 52 | 55 | (b) | (b) | ---- | ---- | ---- | ---- | High |\n\n$^a$In the sketches of the configurations, all dimensions are in inches and all cross sections are circular, except where otherwise indicated.\nThe sound-pressure level was below the range of the analyzer.\n$^b$Power off.\n\nNACA\nNACA TN No. 1838\n74\n```", "timestamp": "2026-07-22T04:51:08.413694+00:00"} | |
| {"citation_id": "19930082914", "source_url": "https://ntrs.nasa.gov/api/citations/19930082914/downloads/19930082914.pdf", "page_number": 4, "total_pages": 66, "image_filename": "19930082914_p4.jpg", "text": "NACA TN No. 1857\n3\n\n$V_o$ velocity of light in vacuum\n$V'$ velocity of light in air of density $\\rho'$\n$x$ coordinate\n$X$ retardation of light beam that causes fringe shift $Y$\n$y$ coordinate\n$Y$ fringe shift\n$z$ coordinate\n$\\lambda$ wavelength of light\n$\\lambda_o$ wavelength of light in vacuum\n$\\lambda'$ wavelength of light in air of density $\\rho'$\n$\\xi = p + qy$\n$\\rho, \\rho', \\rho_{atm}$ density of air\n$\\sigma$ scale factor determined by comparing experimental and theoretical velocity distributions\n$\\phi$ angle of incidence of light ray\n\nAPPARATUS\n\nJet\n\nThe open jet is operated by the air from a 500-cubic-foot storage tank in which the initial pressure is about 10 pounds per square inch and from which a pipe leads to a supersonic nozzle. The exit end of the nozzle is open to the atmosphere. In the pipe that runs from the tank to the nozzle is a hand-operated gate valve, an air-operated quick-opening valve, and a pressure regulator. The purpose of the pressure regulator is to reduce the pressure of the air going into the nozzle to a value that will result in the pressure of the air, as it leaves the nozzle, being atmospheric. The result is that no strong shock waves or expansion waves originate at the rim of the nozzle to adjust the jet pressure to atmospheric.\n\nThe results reported herein were obtained with a nozzle that gave a Mach number of 1.6. The nozzle was of the two-dimensional type and was constructed of steel. It had an exit opening of 3 by 3 inches. The", "timestamp": "2026-07-22T04:51:09.482139+00:00"} | |
| {"citation_id": "19930082712", "source_url": "https://ntrs.nasa.gov/api/citations/19930082712/downloads/19930082712.pdf", "page_number": 5, "total_pages": 14, "image_filename": "19930082712_p5.jpg", "text": "NACA TN 1998\n\nR Reynolds number \nx distance along chord from leading edge \ny distance perpendicular to chord \n$\\alpha_0$ section angle of attack, degrees \n$\\alpha_{l_0}$ section angle of zero lift, degrees \n\nMODEL AND TESTS\n\nThe 24-inch-chord model of the NACA 8-H-12 airfoil was constructed of laminated mahogany. For tests in the smooth condition, the surfaces of the model were lacquered and then sanded in a chordwise direction with No. 400 carborundum paper until aerodynamically smooth. For tests with standard roughness, carborundum grains of 0.011-inch diameter were applied over a surface length of 0.08c to each surface measured from the airfoil leading edge. The grains were sparsely spread to cover from 5 to 10 percent of the area. The model completely spanned the smaller dimension of the 3- by $7\\frac{1}{2}$-foot rectangular test section of the Langley two-dimensional low-turbulence pressure tunnel. The model was pivoted at 0.25c in the chordwise direction and, because of the strength requirements of this particular model, at a vertical distance of $1/2$ inch above the chord line. The gaps between the tunnel walls and the ends of the model were sealed to prevent air leakage. Ordinates for the NACA 8-H-12 airfoil section are given in table I.\n\nThe tests consisted of measurements of the lift, drag, and pitching-moment coefficients at Reynolds numbers of $3.0 \\times 10^6$, $6.0 \\times 10^6$, $9.0 \\times 10^6$, and $11.0 \\times 10^6$. Lift was obtained from the resultant of the integrated pressure distributions along the floor and ceiling of the tunnel test section. Drag was obtained by means of the wake-survey method, and pitching moments were measured with a torque balance. For variations in Reynolds number, the density of the air within the tunnel was changed over a pressure range of 2 to 10 atmospheres. The maximum Mach number attained during the tests was less than 0.13, therefore the results may be considered to be relatively free of compressibility effects. Detailed information on the tunnel and its operation can be found in reference 5.", "timestamp": "2026-07-22T04:51:16.029998+00:00"} | |
| {"citation_id": "19930085965", "source_url": "https://ntrs.nasa.gov/api/citations/19930085965/downloads/19930085965.pdf", "page_number": 67, "total_pages": 67, "image_filename": "19930085965_p67.jpg", "text": "66\n\n$\\frac{25}{32}''$\n\nA\n\n$\\frac{25}{32}''$\n\nA\n\n$4\\frac{1}{2}''$\n\ns/C\n\n0.25\n\n0.125''\n\n.50\n\n.75\n\n.080''\n\n.050''\n\n62°\n\nSection A-A \n(mean section)\n\nNACA\n\nFigure 19. — Physical dimensions of inlet guide vanes of typical axial-flow compressor.\n\nNACA-Langley - 8-24-49 - 350\n\nNACA RM E9E06", "timestamp": "2026-07-22T04:51:17.009650+00:00"} | |
| {"citation_id": "19930085842", "source_url": "https://ntrs.nasa.gov/api/citations/19930085842/downloads/19930085842.pdf", "page_number": 102, "total_pages": 104, "image_filename": "19930085842_p102.jpg", "text": "98\nNACA RM L9C29\n\n[Figure: A graph with four subplots plotted on a grid. The x-axis is labeled \"Lift coefficient, $C_L$\" with values 2, 3, 4, 5, 6. The left y-axis is labeled \"$\\alpha$, deg\" with values 58, 60. The right y-axis has multiple scales: top scale labeled \"$V_{MD}$\" with values 0, 2, 4; middle scale labeled \"$Q_C$\" with values 0.4, 0.8, 1.2; bottom scale labeled \"$C_{DR}$\" with values -0.4, -0.2, 0, 0.2, 0.4. The top subplot shows data points for $V_{MD}$. The second subplot shows data points for $\\alpha$. The third subplot shows data points for $Q_C$. The bottom subplot shows data points for $C_{DR}$. A box in the bottom right corner of the graph contains the text \"NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\".]\n\n(a) $\\alpha_u = 60^\\circ$.\nFigure 47.- Continued.", "timestamp": "2026-07-22T04:51:20.480874+00:00"} | |
| {"citation_id": "19930082485", "source_url": "https://ntrs.nasa.gov/api/citations/19930082485/downloads/19930082485.pdf", "page_number": 26, "total_pages": 62, "image_filename": "19930082485_p26.jpg", "text": "NACA TN No. 1810\n\nThen,\n\n$\\psi_{1} = \\frac{i-1}{j-1} W$\n\nthat is,\n\n$\\psi_{1} = 0, \\; \\psi_{2} = \\frac{W}{4}, \\; \\psi_{3} = \\frac{2}{4} W \\quad \\text{for} \\quad j = 5$\n\nThese relations are used to relocate the potential lines and the streamlines. The velocities are recomputed and the flow network again checked. Satisfactory accuracy is usually obtained in two adjustments of the flow network.\n\nThe accuracy of the network from $\\phi_{4}$ and the rear stagnation point and from $\\phi_{8}$ and the front stagnation point cannot be checked, inasmuch as the accuracy of the assumed streamline bounding the channel in these regions cannot be checked by these methods. If the network for uniform flow some distance ahead of and behind the cascade is, however, smoothly blended with the network between the blade surfaces, the surface velocities from $\\phi_{4}$ to the rear stagnation point and from $\\phi_{8}$ to the front stagnation point can be approximated with reasonable accuracy.\n\nThe first approximation of the blade shape may be altered by increasing the camber or changing the blade thickness or otherwise altering the blade shape to satisfy equation (44) and to establish a desirable velocity distribution. For example, it is desirable to have the maximum velocity on the blade surface no more than 15 percent greater than the leaving velocity, as specified by the vector diagram, inasmuch as too great a reduction in velocity in the region of the trailing edge on the suction surface is likely to result in flow separation. Accelerating flow over as much of the blade surface as possible in order to minimize boundary-layer growth is also desirable.\n\nBlade sections are developed for a number of radii and faired to form a complete blade. Stator blades may be designed with a linear taper, but the operating stress in a rotor blade for a given tip speed will be reduced if a parabolic taper is used.", "timestamp": "2026-07-22T04:51:21.029122+00:00"} | |
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