Buckets:
| {"citation_id": "19930082498", "source_url": "https://ntrs.nasa.gov/api/citations/19930082498/downloads/19930082498.pdf", "page_number": 33, "total_pages": 49, "image_filename": "19930082498_p33.jpg", "text": "Page intentionally left blank\n\nPage intentionally left blank", "timestamp": "2026-07-22T05:08:31.451402+00:00"} | |
| {"citation_id": "19930082614", "source_url": "https://ntrs.nasa.gov/api/citations/19930082614/downloads/19930082614.pdf", "page_number": 25, "total_pages": 36, "image_filename": "19930082614_p25.jpg", "text": "```markdown\nDeceleration factor, K, per foot\nAltitude, h, ft\n\n.001 .002 .004 .01 .02 .04 .06 .10 .20 40x10³ 0 4 8 12 16 20 30 40 50 60x10³ 1.8x10³\n\nNet drag coefficient, C_Dn = 0.01\n0.2\n0.4\n0.6\n0.8\n1.0\n1.2\n1.4\n1.6\n1.8\n2.0\n4.0\n6.0\n8.0\n10.0\n20.0\n40.0\n60.0\n\nWing loading, lb/sq ft\n30\n40\n50\n60\n80\n100\n120\n140\n\nExample:\nWing loading, 50 lb/sq ft\nAltitude, 25,000 ft\nTrue airspeed, 700 ft/sec\nNet drag coefficient, 0.114\nFlight-path angle, -60°\nThe guide lines indicate an\nacceleration factor g/g₀ of 0.27\n\nK/C_Dn\n1.6\n1.4\n1.2\n1.0\n.8\n.6\n.4\n.2\n0\n\ng/g₀ + sin γ\n0\n-.4\n-.8\n-1.2\n-1.6\n-2.0\n-2.4\n-2.8\n\nTrue airspeed, V, ft/sec\nV=200\n300\n400\n500\n600\n700\n800\n900\n1000\n\nγ =\n90°\n70°\n60°\n50°\n40°\n30°\n\nFlight-path angle, γ\n-90°\n-70°\n-60°\n-50°\n-40°\n-30°\n\nLongitudinal acceleration factor, g/g₀\n-2.4 -2.0 -1.6 -1.2 -.8 -.4 0 .4 .8 1.2 1.6\n\nFigure 1- Graphical solution for determining longitudinal acceleration.\n(A larger copy of this figure is enclosed in an envelope at the end of the report.)\n\nNACA TN 1399\n23\n```", "timestamp": "2026-07-22T05:08:35.054078+00:00"} | |
| {"citation_id": "19930085487", "source_url": "https://ntrs.nasa.gov/api/citations/19930085487/downloads/19930085487.pdf", "page_number": 24, "total_pages": 36, "image_filename": "19930085487_p24.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:08:35.478897+00:00"} | |
| {"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 40, "total_pages": 99, "image_filename": "19930082511_p40.jpg", "text": "38\nNACA TN No. 1826\n\nSymmetrical closed-open-closed tunnel.- In figure 18 are shown similar curves for a symmetrical closed-open-closed tunnel of which the length of the open section is 1.5 times the tunnel height. All curves show a sharp reduction in the induced angle as the closed exit is approached and entered; however, for $\\xi_1 < 1$, the closed exit has practically no effect on the induced velocities at and upstream of the vortex. The induced angle at the vortex decreases rapidly as the vortex moves downstream from about $\\xi_1 = 1.0$, and is $\\frac{\\Gamma^1}{4}$ in the plane of the exit ($\\xi_1 = 1.5$).\n\nClosed-open-closed tunnel with one exit lip.- Figure 19 shows results for a tunnel similar to that just discussed except that one exit lip is omitted. The two curves shown are very similar to the corresponding curves for the symmetrical condition. As was pointed out earlier, the horizontal component of the induced velocity on the axis is not zero for this unsymmetrical configuration. Values of this horizontal component have been plotted in figure 20 for the same two vortex locations as in figure 19. The values are seen to be very small in the forward part of the tunnel but become quite large in the neighborhood of the exit lip. The effect is consistent with the concept of the exit lip as a concentration of vortices having total strength equal and opposite to that of the bound vortex and serving thereby to turn the air back to its original direction. The fact that the two curves are practically identical lends further support to this viewpoint.\n\nComparison of the three tunnel types.- In figure 21 are compared the induced-angle curves for $\\xi_1 = 0.5$ and 1.0 for the three tunnel types just discussed. It is seen that the differences are slight up to about $\\xi = 1.0$; beyond this value the curves for the closed-open tunnel continue to rise, while the others descend rapidly. The effect of the closed exit is somewhat larger for the tunnel with two exit lips than for the tunnel with one exit lip. Although the induced angles become slightly negative in the downstream closed region they eventually return to zero.\n\nSymmetrical closed-open-closed tunnel with unequal pressures on the two free surfaces.- By means of equation (24) calculations were made of the induced vertical velocities on the axis of a closed-open-closed tunnel of jet length equal to 1.5 times the tunnel height and having equal and opposite horizontal perturbation velocities ($-u_p$ and $u_p$) on the upper and lower free surfaces, respectively. The results are plotted in figure 22. The curve shows that the vertical velocity component (or the induced angle) has an almost linear variation along the axis, which corresponds to the fairly uniform curvature of the jet that would be expected to result from the pressure difference between the upper and lower surfaces. For this same condition, the integral of the normal velocity along the free surface (equation (25)), which is the downward displacement of the jet boundary at the exit lip, was found to be $3.89u_p$.", "timestamp": "2026-07-22T05:08:35.830252+00:00"} | |
| {"citation_id": "19930082914", "source_url": "https://ntrs.nasa.gov/api/citations/19930082914/downloads/19930082914.pdf", "page_number": 26, "total_pages": 66, "image_filename": "19930082914_p26.jpg", "text": "NACA TN No. 1857\n25\n\nTollmien obtained (reference 1) the theoretical velocity distribution for the mixing region of an incompressible jet. His results, given in table I of reference 1, are shown in figure 16. The vertical coordinate is $u/u_0$, where $u_0$ is the free-stream velocity. The horizontal coordinate is again $\\sigma y/x$. (For this figure only, the position of the x-axis has not been adjusted to coincide with the $\\frac{u}{u_0} = 0.5$ contour but is horizontal, at right angles to the y-axis.) Tollmien's results show that the outside edge of the mixing region of an incompressible jet is given by $\\frac{\\sigma y}{x} = -2.04$. Furthermore, Abramovich found (reference 3) that compressibility had no effect on the value of $\\sigma y/x$ at the outside boundary of the mixing region. He treated theoretically a jet in which the stagnation temperature was the same as that of the ambient air and in which the free-stream velocity was large - up to the velocity of sound. He also treated the low-speed jet in which the stagnation temperature was different from the temperature of the ambient air. He found that compressibility effects arising from the high velocity or from the temperature difference did not affect the value of $\\sigma y/x$ at the outside boundary, but that the value for both cases was -2.04. In the present paper, therefore, this value has been accepted as correct.\n\nTollmien's results also show that the inside edge of the mixing region of an incompressible jet lies along a value $\\frac{\\sigma y}{x} = 0.98$. Abramovich found, however, that the value of $\\sigma y/x$ at the inside boundary should be affected by compressibility considerations. His analysis, however, was not extended to supersonic velocities, and no theoretical value for $\\sigma y/x$ at the inside boundary is available for supersonic flow.\n\nThe quantity $\\sigma$ was introduced in the theory of free jet mixing as a scale factor. (For example, see reference 1.) The theory does not give the value of $\\sigma$, but its value is determined by fitting the experimentally determined velocity distribution to the theoretical velocity distribution. In the present case, the inside boundary of the mixing zone is clearly shown on the interferograms, and the angle that it makes with the free-stream direction can be measured and shown to be $3^\\circ$. Because the theoretical value for $\\sigma y/x$ at this boundary is not known, the value of $\\sigma$ cannot be determined from the measured rate of spread of the mixing zone into the jet. Furthermore, at the outer edge of the mixing zone where $\\frac{\\sigma y}{x} = -2.04$, the density gradient is extremely small. An interferometer is not very sensitive to such small gradients, however, and the outer edge of the mixing zone cannot, therefore, be determined accurately from the interferograms. Shadowgrams that covered an extent of 5 feet along the jet were therefore taken. On the shadowgrams the outside boundary of the mixing zone appeared to lie at approximately $6^\\circ$. This angle would give a value of $\\sigma$ of approximately 20,\n$$ \\sigma \\approx \\frac{-2.04}{-\\tan 6^\\circ} $$", "timestamp": "2026-07-22T05:08:36.355784+00:00"} | |
| {"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 2, "total_pages": 96, "image_filename": "19930085880_p2.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:08:36.459123+00:00"} | |
| {"citation_id": "19930085879", "source_url": "https://ntrs.nasa.gov/api/citations/19930085879/downloads/19930085879.pdf", "page_number": 4, "total_pages": 29, "image_filename": "19930085879_p4.jpg", "text": "```markdown\n2\nNACA RM L9D11\n\ndata indicate that some means will be necessary to stabilize such a jettisoned nose section to prevent linear and centripetal accelerations dangerous to the pilot and show that the addition of suitable fins would accomplish this at low speeds. Rocket-propelled test vehicles designated RM-11A and RM-11B have been constructed and tested to investigate the problems of jettisoning such a fin-stabilized nose section during high-speed power-off flight and to measure the accelerations throughout its flight path. The results of the tests of these two vehicles are covered by the present paper. The nose-section model used was not a model of any particular airplane nose but was fairly representative of a $\\frac{1}{4}$- or $\\frac{1}{5}$-scale-airplane nose section dynamically similar at about 40,000 feet altitude. Since one of the problems of this escape method is to insure that the rear body will not overtake the jettisoned nose, an attempt was made to measure the relative decelerations of the two sections. The models were launched at the Pilotless Aircraft Research Station, Wallops Island, Va.\n\nAPPARATUS AND METHODS\n\nModel\n\nThe RM-11 rocket-propelled test vehicle covered by the present paper was similar to the FR-1-A (reference 4) and so constructed that the nose could be jettisoned at a station 40.5 inches from its tip. The nose section had four stabilizing fins of 32.4-square-inches area each, permanently installed such that the trailing edge was at the separation station and the center of gravity was located 60 percent back of the nose. The results of previous tests of a one-half-scale model of this nose configuration in the Langley 20-foot free-spinning tunnel (reference 3) had indicated that the fin size selected for this center-of-gravity location would give good stability. The center-of-gravity position of the jettisonable nose was made approximately 60 percent back from the tip because designers of jettisonable nose sections for most research airplanes have found that no practical layout will give a center-of-gravity location farther forward. The rear body was designed so that it would be stable after separation. A sketch of the model is shown in figure 1 and a photograph in figure 2. The mass-balance characteristics of the two models used are listed in table I. The model was launched from a near zero length launcher at an angle of 60° using methods described in reference 4.\n\nA sketch of the jettison mechanism is shown in figure 3. A mercury deceleration switch was used to close the firing circuit of the jettison charge and a delay squib of approximately 0.8 second was used to insure complete loss of thrust before ejection. When the jettison charge is fired, the piston cannot move; the jettison cylinder therefore moves forward on the piston, releasing the toggle latches. The cylinder continues to move forward off the piston, carrying the entire nose section with it.\n```", "timestamp": "2026-07-22T05:08:40.314427+00:00"} | |
| {"citation_id": "19930082485", "source_url": "https://ntrs.nasa.gov/api/citations/19930082485/downloads/19930082485.pdf", "page_number": 47, "total_pages": 62, "image_filename": "19930082485_p47.jpg", "text": "46\nNACA TN No. 1810\n\n1026\n\n$$J$$\n\n$$\\frac{C_1}{\\Delta C}$$\n-4.0\n-8.0\n-16.0\n$$\\infty$$\n16.0\n8.0\n4.0\n3.0\n2.0\n1.5\n1.0\n\n$$\\frac{C_1}{\\Delta C}$$\n1.0\n1.5\n2.0\n3.0\n4.0\n8.0\n$$\\infty$$\n-8.0\n-4.0\n\n$$C_1 n_0 / 2$$\n\n[Figure: NACA logo]\n\nFigure 15.- Chart for finding J from channel width and blade curvature. (A 10-by-16-in. print of this figure is attached.)", "timestamp": "2026-07-22T05:08:41.593724+00:00"} | |
| {"citation_id": "19930082245", "source_url": "https://ntrs.nasa.gov/api/citations/19930082245/downloads/19930082245.pdf", "page_number": 51, "total_pages": 66, "image_filename": "19930082245_p51.jpg", "text": "```markdown\n50\n\nAileron section normal-force coefficient, $C_{n_a}$\nAileron section hinge-moment coefficient, $C_h$\n\n<!-- Image (48, 134, 932, 685) -->\n\n(a) $C_n = -0.4$.\n\nFigure 10.- Variation of aileron normal-force and hinge-moment coefficients with Mach number for an unsealed 0.20c plain aileron of beveled-trailing-edge profile mounted on an NACA 66,1-115 airfoil section.\n\nNACA TN No. 1596\n```", "timestamp": "2026-07-22T05:08:47.355198+00:00"} | |
| {"citation_id": "19930085626", "source_url": "https://ntrs.nasa.gov/api/citations/19930085626/downloads/19930085626.pdf", "page_number": 7, "total_pages": 24, "image_filename": "19930085626_p7.jpg", "text": "6\nCONFIDENTIAL\nNACA RM No. L8K23\n\nREFERENCES\n\n1. Sandahl, Carl A.: Free-Flight Investigation at Transonic and Super-\nsonic Speeds of the Rolling Effectiveness of a 42.7° Sweptback Wing\nHaving Partial-Span Ailerons. NACA RM No. L8E25, 1948.\n\n2. Turner, Thomas R., Lockwood, Vernard E., and Vogler, Raymond D.:\nPreliminary Investigation of Various Ailerons on a 42° Sweptback\nWing for Lateral Control at Transonic Speeds.\nNACA RM No. L8D21, 1948.\n\n3. Sandahl, Carl A., and Marino, Alfred A.: Free-Flight Investigation\nof Control Effectiveness of Full-Span 0.2-Chord Plain Ailerons at\nHigh Subsonic, Transonic, and Supersonic Speeds to Determine Some\nEffects of Section Thickness and Wing Sweepback.\nNACA RM No. L7D02, 1947.\n\n4. Sandahl, Carl A.: Free-Flight Investigation of Control Effectiveness\nof Full-Span, 0.2-Chord Plain Ailerons at High Subsonic, Transonic,\nand Supersonic Speeds to Determine Some Effects of Wing Sweepback,\nTaper, Aspect Ratio, and Section Thickness Ratio.\nNACA RM No. L7F30, 1947.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:08:50.157099+00:00"} | |
| {"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 26, "total_pages": 78, "image_filename": "19930082618_p26.jpg", "text": "```markdown\n24\n\n<!-- Image (79, 78, 879, 764) -->\n\n(c) Section drag characteristics and section pitching-moment characteristics about the aerodynamic center of the plain NACA 64-409 airfoil section.\n\nFigure 1.- Concluded.\n\nNACA TN 1945\n```", "timestamp": "2026-07-22T05:08:54.141165+00:00"} | |
| {"citation_id": "19930085471", "source_url": "https://ntrs.nasa.gov/api/citations/19930085471/downloads/19930085471.pdf", "page_number": 25, "total_pages": 28, "image_filename": "19930085471_p25.jpg", "text": "UNCLASSIFIED\nCONFIDENTIAL\n\nWing\ndesignation\nE.A.\n(percent\nchord)\n$r_a^2$\n\nA-1\nB-1\nB-2\nB-3\nB-4\nC-1\nC-2\nD-1\nD-2\nE-1\nE-2\nF-1\nF-2\nG-1\nG-2\n\n40\n40\n50\n\n0.38\n.25\n.25\n\nV/b$\\omega_\\alpha$\n1/k\n\n[Figure: Graph showing experimental data points and theoretical lines for flutter analysis]\n\nNACA\n\nFigure 9.- Comparison of experimental values of V/b$\\omega_\\alpha$ for coupled wing flutter with theoretical values for one-degree-of-freedom torsional flutter. M = 1.3, $\\alpha_\\alpha$ = 0.04.\n\nNACA RM No. L5A11\n23", "timestamp": "2026-07-22T05:08:54.422725+00:00"} | |
| {"citation_id": "19930082496", "source_url": "https://ntrs.nasa.gov/api/citations/19930082496/downloads/19930082496.pdf", "page_number": 37, "total_pages": 50, "image_filename": "19930082496_p37.jpg", "text": "36\n\nPage intentionally left blank\n\nPage intentionally left blank", "timestamp": "2026-07-22T05:08:55.843236+00:00"} | |
| {"citation_id": "19930085519", "source_url": "https://ntrs.nasa.gov/api/citations/19930085519/downloads/19930085519.pdf", "page_number": 22, "total_pages": 46, "image_filename": "19930085519_p22.jpg", "text": "```markdown\nNACA RM No. L58K19\n\nPlane of Symmetry\n\n.20c\n42.00\n\n.25 b/2\n.35 b/2\n.45 b/2\n.55 b/2\n.65 b/2\n.75 b/2\n\n.10 b/2\n.05 b/2\n\nb/2 = 68.25\n\n.20c\n26.25\n\nSlot lip\nFlap leading edge\nActuating arms\n\n0.70 chord line\n\nNACA\n\nFigure 3.- Plug-aileron and slotted-flap locations on the 42° sweptback wing. All dimensions are in inches unless otherwise noted.\n\n21\n```", "timestamp": "2026-07-22T05:08:57.420460+00:00"} | |
| {"citation_id": "19930085544", "source_url": "https://ntrs.nasa.gov/api/citations/19930085544/downloads/19930085544.pdf", "page_number": 10, "total_pages": 33, "image_filename": "19930085544_p10.jpg", "text": "NACA RM No. L8K26\n\nFrom figure 1 it can be seen that the maximum increment in the geometric angle of attack occurs when the quantity $\\sin \\omega t = 1$ ($\\omega t = 90^\\circ$) and at $\\sin \\omega t = -1$ ($\\omega t = 270^\\circ$). The amplitude $\\left(\\alpha_{P_0}\\right)$ of the geometric angle of attack at $\\omega t = 90^\\circ$ is\n\n$$\n\\phi_o - \\phi_{90} = \\tan^{-1} \\frac{J}{\\pi x} \\cos \\alpha_T - \\tan^{-1} \\frac{\\cos \\alpha_T}{\\frac{\\pi x}{J} + \\sin \\alpha_T}\n$$\n\nand the amplitude at $\\omega t = 270^\\circ$ is\n\n$$\n\\phi_{270} - \\phi_o = \\tan^{-1} \\frac{\\cos \\alpha_T}{\\frac{\\pi x}{J} - \\sin \\alpha_T} - \\tan^{-1} \\frac{J}{\\pi x} \\cos \\alpha_T\n$$\n\nThese values have been calculated for several values of $\\frac{J}{\\pi x}$ with the angle $\\alpha_T$ as parameter, and the results are shown plotted in figure 3. In figure 4 are shown results of similar calculations made to determine the value of $\\epsilon$ at $\\omega t = 90^\\circ$ and $\\omega t = 270^\\circ$. Figures 3 and 4 show that in the propeller case the deviation from sinusoidal variations in the resultant velocity and angle of attack is small at thrust-axis angles less than about $6^\\circ$ and values of $\\frac{J}{\\pi x}$ less than approximately 2.\n\nThe flapping motion $h$ is a function of the blade stiffness and will not be considered here. Calculations show that the effect of this motion on the maximum force is in general small but that the lag in the position of the maximum force may become large depending on the frequency of the oscillations.\n\nDropping the $h$ terms, equation (7) reduces to\n\n$$\n\\begin{aligned}\nL = & -\\pi \\rho \\frac{c^2}{4} \\left[ \\dot{W}_{\\omega t} \\dot{\\alpha}_P + \\dot{W}_{\\omega t} (\\alpha + \\alpha_P) \\right] - 2\\pi \\rho \\dot{W}_{\\omega t} \\frac{c}{2} \\left[ W_o \\alpha \\right. \\\\\n& \\left. + (\\dot{W}_{\\omega t} - \\dot{W}_o) \\alpha C(k_{W_{\\omega t}}) + \\left( \\frac{c}{4} \\dot{\\alpha}_P + \\dot{W}_o \\alpha_P \\right) C(k_{\\alpha_P}) \\right. \\\\\n& \\left. + (\\dot{W}_{\\omega t} - \\dot{W}_o) \\alpha_P C(k_{W_{\\omega t}} + k_{\\alpha_P}) \\right]\n\\end{aligned}\n\\tag{8}\n$$", "timestamp": "2026-07-22T05:09:00.736317+00:00"} | |
| {"citation_id": "19930085869", "source_url": "https://ntrs.nasa.gov/api/citations/19930085869/downloads/19930085869.pdf", "page_number": 6, "total_pages": 36, "image_filename": "19930085869_p6.jpg", "text": "4\nCONFIDENTIAL\nNACA RM L9D15\n\nTake-off Stability\n\nIn order to find the trim limits of stability, the towing carriage was held at constant speeds, while the model trim was slowly increased or decreased until the porpoising limit was crossed. The variation of trim with speed for three locations of the center of gravity ($0.20\\bar{c}$, $0.30\\bar{c}$, and $0.40\\bar{c}$) was determined during accelerated runs ($1.0 \\text{ ft/sec}^2$) to take-off with full power and fixed elevators. The range of available center-of-gravity and elevator positions that would permit operation at trims above $2^\\circ$ without porpoising of greater than $2^\\circ$ amplitude was investigated during accelerated runs. A minimum trim of $2^\\circ$ appeared to be a reasonable limit for purposes of evaluation.\n\nLanding Stability\n\nThe landing stability was investigated by trimming the model in the air to the desired landing trim, while the carriage was held at a constant speed slightly above the model flying speed, and then decelerating the carriage at a constant rate of 3 feet per second per second, allowing the model to glide onto the water in simulation of an actual landing. The descent to the water from flight was made from a height of $0.3b$ above the water. This method was used to hold the sinking speeds to a value of approximately 300 feet per minute. After the first contact the rise restriction was removed. Landings were made with the center of gravity located at $0.20\\bar{c}$, $0.30\\bar{c}$, and $0.40\\bar{c}$, using one-quarter thrust.\n\nResistance\n\nThe resistance characteristics were obtained with the wing and tail surfaces removed. Constant speed runs were made with the model fixed in trim. Lift curves (assuming lift to vary as the square of the speed) were calculated for the model from the take-off speeds observed for various trims during the take-off stability tests. The load on the water, applied by dead weights, was determined from the lift curves. The range of trim tested at each speed was the range of stable trim obtained at that speed during the stability tests with the center of gravity located at $0.30\\bar{c}$ except in the speed range from $C_y = 6.0$ to take-off where the maximum trim was arbitrarily limited to $12^\\circ$. The resistance selected was the lowest resistance obtained at each speed.\n\nRESULTS AND DISCUSSION\n\nSpray Characteristics\n\nThe gross load coefficient at which spray entered the propellers with the two spray-strip arrangements shown in figure 4 is plotted\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:09:02.980515+00:00"} | |
| {"citation_id": "19930082614", "source_url": "https://ntrs.nasa.gov/api/citations/19930082614/downloads/19930082614.pdf", "page_number": 26, "total_pages": 36, "image_filename": "19930082614_p26.jpg", "text": "24\nNACA TN 1939\n\nTrue airspeed, V, ft/sec\n900\n800\n700\n600\n500\n400\n300\n200\n100\n0\n\nDeceleration factor\nK, per ft\n0.005 x 10⁻³\n.010 x 10⁻³\n.020 x 10⁻³\n.030 x 10⁻³\n.050 x 10⁻³\n.070 x 10⁻³\n.100 x 10⁻³\n\nTime, t, sec\n0 4 8 12 16 20 24 28\n\n[Figure: NACA logo]\n\nFigure 2.— The effect of deceleration factor K upon the variation of speed with time for several initial speeds. Level flight condition.", "timestamp": "2026-07-22T05:09:14.943175+00:00"} | |
| {"citation_id": "19930085487", "source_url": "https://ntrs.nasa.gov/api/citations/19930085487/downloads/19930085487.pdf", "page_number": 25, "total_pages": 36, "image_filename": "19930085487_p25.jpg", "text": "NACA RM No. E8J22\n\n[Figure: Sand pattern on turbine blades with scale marked \"INCHES 0 1\"]\n\nNACA\nC-14992\n5-20-46\n\nFigure 8. - Sand pattern for first torsional-mode vibration on ninth-stage blade showing distorted node shape. Frequency of vibration, 6160 cycles per second.\n\n23", "timestamp": "2026-07-22T05:09:15.530524+00:00"} | |
| {"citation_id": "19930085536", "source_url": "https://ntrs.nasa.gov/api/citations/19930085536/downloads/19930085536.pdf", "page_number": 19, "total_pages": 20, "image_filename": "19930085536_p19.jpg", "text": "```markdown\nNACA RM No. E8K05\n\nPressure coefficient, $C_p$\n\n| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | |", "timestamp": "2026-07-22T05:09:17.341670+00:00"} | |
| {"citation_id": "19930082498", "source_url": "https://ntrs.nasa.gov/api/citations/19930082498/downloads/19930082498.pdf", "page_number": 34, "total_pages": 49, "image_filename": "19930082498_p34.jpg", "text": "NACA TN No. 1838\n\n[Figure: A metallic, U-shaped mechanical assembly with multiple cylindrical ports or nozzles mounted on two vertical plates connected by a horizontal tube. A ruler marked in inches is visible at the bottom center for scale. In the lower right corner of the image, a label reads “NACA L-52705”.]\n\n(c) Configuration 58.\n\nFigure 2.— Continued.\n\n33", "timestamp": "2026-07-22T05:09:17.755045+00:00"} | |
| {"citation_id": "19930085542", "source_url": "https://ntrs.nasa.gov/api/citations/19930085542/downloads/19930085542.pdf", "page_number": 13, "total_pages": 46, "image_filename": "19930085542_p13.jpg", "text": "NACA RM No. L5L29\n\nHigh directional stability is indicated at high lift coefficients for low aspect ratios in figure 18. Also indicated in figure 18 are very high values of the effective dihedral parameter $C_{l_\\psi}$ for low aspect ratios.\n\nReducing the aspect ratio by cutting portions from the tips parallel to the plane of symmetry results in an increase in directional stability which is in fair agreement with the theory of reference 5 down to an aspect ratio of 2. (See fig. 24.) At an aspect ratio of 1.0 the directional stability is considerably larger than that predicted by reference 5. The increase in the variation of $C_{l_\\psi}$ with $C_L$ with decreasing aspect ratio is considerably larger than is indicated by either reference 2 or reference 5. As the aspect ratio is decreased, $C_{l_p}$ at high lift coefficients, has a tendency to increase. (See fig. 19.) The experimental variation of $C_{Y_p}$ with $C_L$ (fig. 25) shows little change with aspect ratio; whereas the theory of reference 5 indicates a small reduction in this parameter with decreasing aspect ratio and the triangular wing theory of reference 2 indicates a large increase in $\\partial C_{Y_p}/\\partial C_L$ with decreasing aspect ratio. In general, better agreement was obtained with reference 5. The experimental curve of figure 25 indicates a decrease in the variation of $C_{n_p}$ with $C_L$ with decreasing aspect ratio which is in fair agreement with the theory of reference 5. Reducing the aspect ratio results in a decrease in damping in roll $C_{l_p}$. (See fig. 25.) Good agreement is indicated with experiment by the theory of reference 5 while the theory of reference 2 is in fair agreement with experiment only at the lowest test aspect ratio ($A = 1.0$). The theory of reference 4 is in fair agreement at low aspect ratios and excellent agreement at $A = 3$ and $A = 4$.\n\nCONCLUSIONS\n\nTests conducted in the 6-foot-diameter rolling-flow test section of the Langley stability tunnel to determine the effects of a number of geometric variables on the low-speed static-stability and rolling characteristics of triangular wings indicate the following conclusions:\n\n1. In general, variations in profile had small effects on the static and rolling characteristics of triangular wings at low lift coefficients. The greatest effect of profile was in the high-lift-coefficient range. The linear range of the static-stability parameters was decreased (in much the same manner as for untapered swept wings) as the leading-edge sharpness was increased.\n\n2. Several of the characteristics of triangular wings may be estimated with fair accuracy by available swept-wing theory. Available low-aspect-ratio triangular-wing theory was found to be reliable for", "timestamp": "2026-07-22T05:09:19.795661+00:00"} | |
| {"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 41, "total_pages": 99, "image_filename": "19930082511_p41.jpg", "text": "NACA TN No. 1826\n39\n\nFor a vortex $\\Gamma'$ located at $\\xi_1 = 0.5$ in the equal-pressure case, the integral of the normal velocity along the free surface (equation (21)), which is proportional to the upward displacement of the jet boundary at the exit lip, was found to be $1.20\\Gamma'$. Accordingly, zero displacement at the exit, corresponding to the existence of a closed space above or below the jet (reference 2), will result if the flow described in the preceding paragraph is superposed on the equal-pressure flow in such proportion that $3.89u_b = 1.20\\Gamma'$; that is, $\\frac{u_b}{\\Gamma'} = 0.31$. The corresponding effect on the induced angle at $\\xi = \\xi_1$ is found as follows: at $\\xi = \\xi_1 = 0.5$, $\\frac{v}{\\Gamma'}$ for the equal-pressure case (fig. 18) is $0.48$. From figure 22, $\\frac{v}{u_b}$ at $\\xi = 0.5$ for the unequal-pressure case is $-1.44$. Since $0.31 \\times -1.44 = -0.45$, it is seen that, if spillage at the exit lip is prevented, the induced velocity in the region of the vortex is nearly eliminated. A similar comparison of the slopes of the curves in figures 18 and 22 in the neighborhood of $\\xi = 0.5$ shows that the induced curvature in the region of the vortex is also nearly eliminated.\n\nRésumé of numerical results.- The induced angle at the lifting vortex is essentially that for an infinite open jet if the vortex is more than half the tunnel height from the entrance and the exit. The induced angles for case 2 (one fixed exit boundary) are nearly the same as for case 3 (symmetrical exit), so that any failure of the flow to contact the upper exit lip should not appreciably affect the tunnel correction. Finally, for case 3, if enough of the different-pressure flow is added to assure zero displacement of the free boundary at the exit (that is, if spillage at the exit is prevented, as by enclosing the space into which the spillage would normally occur), the induced angle at the vortex may be nearly eliminated.\n\nIII - CIRCULAR TUNNELS\n\nIn part III an outline is given of a general method for calculating the boundary effect in an open circular tunnel of finite jet length. The solution, involving expansions in Bessel functions, is somewhat similar to the solution for a closed circular tunnel (reference 13), but is constructed so that it satisfies the condition of uniformity of pressure over the open boundary and also the condition of continuity of velocity at the entrance lip. Numerical results are given for a lifting element on the tunnel axis.", "timestamp": "2026-07-22T05:09:25.860430+00:00"} | |
| {"citation_id": "19930082485", "source_url": "https://ntrs.nasa.gov/api/citations/19930082485/downloads/19930082485.pdf", "page_number": 48, "total_pages": 62, "image_filename": "19930082485_p48.jpg", "text": "NACA TN No. 1810\n47\n\n<!-- Image (125, 109, 826, 919) -->\n\nFigure 16.- Chart for finding K/J from channel width and blade curvature. (A 10-by-20-in. print of this figure is attached.)", "timestamp": "2026-07-22T05:09:26.465721+00:00"} | |
| {"citation_id": "19930082245", "source_url": "https://ntrs.nasa.gov/api/citations/19930082245/downloads/19930082245.pdf", "page_number": 52, "total_pages": 66, "image_filename": "19930082245_p52.jpg", "text": "```markdown\nNACA TN No. 1596\n\nAileron section normal-force coefficient, $c_{n_a}$\n\n| $\\delta_a$ (deg) |\n|---|\n| 12 |\n| 18 |\n| 6 |\n| 4 |\n| 2 |\n| 0 |\n| -2 |\n| -4 |\n| -6 |\n| -12 |\n\nMach number, M\n\nAileron section hinge-moment coefficient, $c_{h}$\n\n| $\\delta_a$ (deg) |\n|---|\n| -12 |\n| -6 |\n| -4 |\n| -2 |\n| 0 |\n| 2 |\n| 4 |\n| 6 |\n| 12 |\n| 18 |\n\nMach number, M\n\n(b) $c_n = -0.2$.\n\nFigure 10.—Continued.\n\n51\n\n[Figure: NACA logo]\n```", "timestamp": "2026-07-22T05:09:29.213738+00:00"} | |
| {"citation_id": "19930082914", "source_url": "https://ntrs.nasa.gov/api/citations/19930082914/downloads/19930082914.pdf", "page_number": 27, "total_pages": 66, "image_filename": "19930082914_p27.jpg", "text": "26\nNACA TN No. 1857\n\nIt was felt, however, that the shadowgrams might not accurately indicate the true boundary of the mixing zone, and that the most satisfactory method of determining $\\sigma$ would be to choose a value that would give the best fit of the experimentally determined velocity distribution with Tollmien's theoretically determined velocity distribution over the subsonic portion of the mixing region. First, Tollmien's curve, shown in figure 16, was shifted horizontally by 0.39 in order that the 0.5 value of $u/u_0$ would lie at the zero value of $\\sigma y/x$, as shown by the solid curve of figure 17. Then the velocity distributions that were calculated from the measured density distributions, as has been indicated, were plotted, with $\\sigma$ taken as 15 and with the x-axis taken along the $u/u_0 = 0.5$ contour. This value of $\\sigma$ was chosen because it makes the data agree with Tollmien's curve between about 0.2 and 0.6 on the vertical scale. At values of $u/u_0$ smaller than approximately 0.2 the data are probably not very accurate. At values of $u/u_0$ greater than 0.625 the flow is supersonic and it is to be expected that compressibility effects alter the velocity distribution from that of incompressible flow. In fact, figure 17 shows that the inside boundary of the mixing region of the supersonic jet is at $\\frac{\\sigma y}{x} = 1.12$, compared to 1.37 for the incompressible jet. In the supersonic jet, therefore, the mixing region spreads into the free stream at a slower rate than in an incompressible jet.\n\nThe value of $\\sigma$ obtained here can be compared with the values obtained by other investigators for incompressible jets. Measurements in a large jet at Göttingen (see reference 1), in which the free-stream velocity was about 100 feet per second, gave a value for $\\sigma$ of 11.8. Liepmann and Laufer state (reference 8) that Cordes found a value of 11.95 for $\\sigma$. Liepmann and Laufer found that for their jet, which had a free-stream velocity of 59 feet per second, the value of $\\sigma$ was 12.0. These values are to be compared with the results of the present measurements, which give a value of 15 for $\\sigma$ for a jet of Mach number 1.6. Because $\\sigma$ is a measure of the rate of spread of the mixing region, the mixing region of the present supersonic jet spreads less rapidly than that of incompressible jets. The rate of spread of the outside boundary is 12/15 that of incompressible jets. The ratio of the rates of spread of the inside boundary is even less, as is shown by the fact that, on figure 17, the inside boundary lies at a smaller value of $\\sigma y/x$ than the theoretical.\n\nIn the theory of the incompressible jet, it has been customary to assume that the value of the \"mixing length\" is constant across the mixing zone and that the value of the mixing length $l$ is proportional to the distance from the nozzle,\n\n$$l = cx$$", "timestamp": "2026-07-22T05:09:31.644778+00:00"} | |
| {"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 3, "total_pages": 96, "image_filename": "19930085880_p3.jpg", "text": "```markdown\nNACA RM No. L9C03\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nRESEARCH MEMORANDUM\n\nPLANING CHARACTERISTICS OF THREE SURFACES\n\nREPRESENTATIVE OF HYDRO-SKI FORMS\n\nBy Kenneth L. Wadlin and John R. McGehee\n\nSUMMARY\n\nThe planing characteristics, as determined by tank tests, are presented for three surfaces representative of hydro-ski forms. One surface was of rectangular plan form with a flat bottom, the second surface had a rectangular plan form with transversely curved bottom and the third surface had a flat bottom but was triangular in plan form. The range of trims investigated was $4^\\circ$ to $20^\\circ$. The data are presented in the form of plots of load, resistance, trimming moment, and draft against wetted area. Plots of wetted length, wetted area forward of the observed wetted length at the chine, and aerodynamic tare forces are included.\n\nINTRODUCTION\n\nThe use of retractable planing surfaces, called hydro-skis, for supporting jet-propelled water-based airplanes during the high-speed part of their take-offs and landings, was proposed in reference 1. The results of some preliminary tests of models fitted with hydro-skis are presented in references 1 and 2.\n\nHydro-skis are intended to be parts of the airplane which can be extended for landing and take-off. Since the skis come from the fuselage which is generally rounded or the wing which is more or less flat, the skis also will generally have rounded or flat cross sections. Though some data are available on flat rectangular planing surfaces (see references 3 and 4), the range of trims is limited. Practically no data are available on the characteristics of planing surfaces with convex cross sections or plan forms other than rectangles.\n\nInformation as to the effects of plan form and cross-sectional curvature should therefore be an aid in designing hydro-skis and hydro-ski arrangements. Because of this, an investigation was initiated at Langley tank no. 2 to determine the characteristics of planing surfaces of several plan forms and transversely curved bottoms. This paper presents the results of some planing tests of three such surfaces. For convenience in locating the data presented, an index of figures is presented in table 1. Because of current interest of the Air Force and the Bureau of Aeronautics in obtaining data pertaining to hydro-skis, the data in this paper are presented without analysis or discussion to make it available as quickly as possible.\n```", "timestamp": "2026-07-22T05:09:35.729865+00:00"} | |
| {"citation_id": "19930085471", "source_url": "https://ntrs.nasa.gov/api/citations/19930085471/downloads/19930085471.pdf", "page_number": 26, "total_pages": 28, "image_filename": "19930085471_p26.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:09:36.014366+00:00"} | |
| {"citation_id": "19930085879", "source_url": "https://ntrs.nasa.gov/api/citations/19930085879/downloads/19930085879.pdf", "page_number": 5, "total_pages": 29, "image_filename": "19930085879_p5.jpg", "text": "NACA RM L9D11\n\nBecause it was felt that the failure of a release latch in the first test (model A) was due to high bearing forces and/or a twisting moment during power-on flight, bearing plates as shown in figure 4 were installed in model B to absorb these forces and moments so that they would not be applied on the ejection mechanism.\n\nThe nose section was equipped with drag flaps which were driven radially outward by a small electric motor. The flaps started to open approximately 2 seconds after ejection and were completely out 6 seconds after ejection. The flaps are shown retracted and opened in figures 5 and 6, respectively.\n\nInstrumentation\n\nA four-channel telemeter was installed in the nose section to transmit signals from four accelerometers. Three accelerometers to measure longitudinal (along X-axis), transverse (along Y-axis), and normal (along Z-axis of the nose section) accelerations were installed in the jettisonable nose at the locations given in table II. A longitudinal accelerometer installed in the rear body was connected to the telemeter in the nose section by a pull-out plug and about 15 feet of excess cable in order that readings of relative acceleration between the two bodies could be obtained during separation.\n\nA continuous-wave Doppler radar was used to record velocity and an SCR-584 pulse-type radar was used to record trajectory of the models. Atmospheric conditions prevailing at the time of flight were obtained by a radiosonde.\n\nRESULTS\n\nModel A\n\nThe nose section of model A came off prematurely after 4.65 seconds of power-on flight (velocity 615 ft per sec; Mach number 0.54) due to the malfunction of a release latch. The nose yawed to the right, became detached from the rear body, and was struck by the right wing. One nose fin was torn off but the telemeter remained in operation and a record was obtained. Directly after the nose came off, peak accelerations of approximately ±12g occurred about various axes due to contact and interference with the rear body. Two seconds later these oscillations became more or less regular and had peak values of about 1 g to -7g. The telemeter record indicates that the nose section was assuming a helical flight path. A plot of accelerations against time from 7 seconds after launching is shown in figure 7. Since the nose was not forcibly ejected, the automatic switch to the motor which operates the flaps was not turned on.", "timestamp": "2026-07-22T05:09:37.955212+00:00"} | |
| {"citation_id": "19930082496", "source_url": "https://ntrs.nasa.gov/api/citations/19930082496/downloads/19930082496.pdf", "page_number": 38, "total_pages": 50, "image_filename": "19930082496_p38.jpg", "text": "NACA TN No. 1836\n37\n\n[Figure: A circular, dark, textured specimen labeled \"3D14\" above a scale marked \"INCHES\" from 0 to 1. A NACA stamp with \"C.22200\" and \"9-2-48\" is visible in the lower right corner of the image area.]\n\nFigure 9. - Ceramel thermal-shock specimen after completion of 25 cycles at 1800°, 25 at 2000°, 25 at 2200°, and 25 at 2400° F.", "timestamp": "2026-07-22T05:09:40.745360+00:00"} | |
| {"citation_id": "19930085519", "source_url": "https://ntrs.nasa.gov/api/citations/19930085519/downloads/19930085519.pdf", "page_number": 23, "total_pages": 46, "image_filename": "19930085519_p23.jpg", "text": "```markdown\n22\n\n<!-- Image (107, 89, 890, 782) -->\n\nDimension\nA .0300c\nB .0374c\nC .0486c\nD .0450c\nE .0050c\nF .0102c\nG .0093c\nH .0050c\nJ .0102c\nK .0450c\nL .0450c\nM .0605c\nFlap chord .2000c\n\nNACA\n\nFigure 4.- Section dimensions of the slotted flap tested on the 42° sweptback wing.\n\nNACA RM No. L56K19\n```", "timestamp": "2026-07-22T05:09:41.762281+00:00"} | |
| {"citation_id": "19930085544", "source_url": "https://ntrs.nasa.gov/api/citations/19930085544/downloads/19930085544.pdf", "page_number": 11, "total_pages": 33, "image_filename": "19930085544_p11.jpg", "text": "10\nNACA RM No. L8K26\n\nFor the type of motion being considered here it will be convenient to assign the following values to the parameters appearing in equation (8):\n\n$$\n\\alpha_p = -i\\alpha_{P_0} e^{i\\omega t}\n$$\n\n$$\nh = 0\n$$\n\n$$\nW_{\\omega t} = W_0(1 - i\\epsilon e^{i\\omega t})\n$$\n\n$$\n\\epsilon = \\frac{W_{00}}{W_0} - 1\n$$\n\n$$\nC(k) = F + iG\n$$\n\nIn the above expressions the parameter $\\alpha_{P_0}$ is taken as the amplitude of the aerodynamic angle of attack as estimated by steady-state calculations in potential flow. The function $C(k)$ is a complex function of the parameter $k$ (reference 1) and is given by\n\n$$\nC(k) = F(k) + iG(k)\n$$\n\nwhere $F$ and $G$ are obtained from standard Bessel functions of the first and second kinds with argument $k$. The variation of the functions $F$ and $G$ with the parameter $\\frac{1}{k}$ is given in figure 4 and table II of reference 1. In the present case the functions $F$ and $G$ are evaluated as in reference 2 for $k$'s defined as follows:\n\n$$\nk_{W_{\\omega t}} = \\frac{\\omega W_{\\omega t}}{2W_0}\n$$\n\n$$\nk_{\\alpha_p} = \\frac{\\omega \\alpha_p}{2W_0}\n$$\n\n$$\nk_{W_{\\omega t}} + k_{\\alpha_p} = \\frac{(\\omega W_{\\omega t} + \\omega \\alpha_p)c}{2W_0}\n$$", "timestamp": "2026-07-22T05:09:49.237315+00:00"} | |
| {"citation_id": "19930085869", "source_url": "https://ntrs.nasa.gov/api/citations/19930085869/downloads/19930085869.pdf", "page_number": 7, "total_pages": 36, "image_filename": "19930085869_p7.jpg", "text": "NACA RM L9D15 CONFIDENTIAL 5\n\nagainst speed coefficient in figure 6. With the spray strip of the same depth as that used on the unswept models of reference 2, the propeller spray was too heavy to be considered acceptable. This heavy spray resulted from the shorter forebody, shorter spray strip, and the lower trim which allowed spray to flow around the forward end of the spray strip. This undeflected spray was heavier than the light intermittent spray coming from under the strips. Lengthening and deepening the spray strip to give the arrangement shown in figure 4(b) not only reduced the intensity of spray in the propellers at the design gross load but also decreased the ranges of speed and load coefficients over which spray struck the propellers. The propeller spray was considered to be light at the design gross load with the final configuration used. The worst spray condition for the two spray-strip arrangements is shown in figure 7.\n\nIn practice, the spray strips could be retracted in sections. However, unpublished wind-tunnel tests on the unswept model of reference 2 indicate that the drag of such spray strips may not be enough to warrant retraction.\n\nIn spite of the low bow clearance there was no spray over the bow during take-off runs. The tail surfaces were struck by heavy spray from the forebody roach near hump speed over a speed-coefficient range of about 0.3. At high speeds only light spray from the forebody wetted the tail surfaces and the under surface of the inboard wing panels.\n\nTake-Off Stability and Trims\n\nThe trim limits of stability of the swept-hull model are compared with those of the unswept model in figure 8. The differences between the two lower-limit curves are not large, the peak for the swept hull being $10.5^\\circ$. Upper-limit porpoising was obtained for the swept hull during constant speed runs over a short range ($C_V = 6.3$ to $7.7$) near take-off.\n\nIn figure 9, trim is plotted against speed coefficient for various elevator deflections at three locations of the center of gravity. The static trim (approximately $7^\\circ$) was less than that of the unswept model (approximately $10^\\circ$) as a result of the rearward shift of volume. A trim peak of about $12^\\circ$ was reached near a speed coefficient of 3.5. This peak corresponded to that obtained with the twin-boom configuration. (See reference 2.) The second trim peak that occurred in most of the curves was caused by the action of the roach on the tail boom. Some indication of this peak was found with the single-boom configuration of reference 2. At the aft location of the center of gravity, with large up-elevator deflections, trim increased from rest until the model was near take-off. Aerodynamic tests with the model free to trim indicated that the effectiveness of the elevators in trimming the model began to decrease with an increase in elevator deflections greater than $-15^\\circ$. A wide range of trim was obtainable beyond the hump speed for all center-of-gravity locations.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:09:53.859278+00:00"} | |
| {"citation_id": "19930085487", "source_url": "https://ntrs.nasa.gov/api/citations/19930085487/downloads/19930085487.pdf", "page_number": 26, "total_pages": 36, "image_filename": "19930085487_p26.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:09:54.087562+00:00"} | |
| {"citation_id": "19930082498", "source_url": "https://ntrs.nasa.gov/api/citations/19930082498/downloads/19930082498.pdf", "page_number": 35, "total_pages": 49, "image_filename": "19930082498_p35.jpg", "text": "Page intentionally left blank\n\nPage intentionally left blank", "timestamp": "2026-07-22T05:09:58.152526+00:00"} | |
| {"citation_id": "19930085536", "source_url": "https://ntrs.nasa.gov/api/citations/19930085536/downloads/19930085536.pdf", "page_number": 20, "total_pages": 20, "image_filename": "19930085536_p20.jpg", "text": "```markdown\n18\n\nPressure coefficient, $C_p$\n\n| | | | | | | | | | | | | | | | | |\n|---|---|---|---|---|---|---|---|---|---|---|---|---|---|---|---|---|\n| .40 | | | | | | | | | | | | | | | | |\n| | | | | | | | | | | | | | | | | |\n| .32 | | | | | | | | | | | | | | | | |\n| | | | | | | | | | | | | | | | | |\n| .24 | | | | | | | | | | | | | | | | |\n| | | | | | | | | | | | | | | | | |\n| .16 | | | | | | | | | | | | | | | | |\n| | | | | | | | | | | | | | | | | |\n| .08 | | | | | | | | | | | | | | | | |\n| | | | | | | | | | | | | | | | | |\n| 0 | | | | | | | | | | | | | | | | |\n| | | | | | | | | | | | | | | | | |\n| -.08 | | | | | | | | | | | | | | | | |\n| -16 | -12 | -8 | -4 | 0 | 4 | 8 | 12 | 16 | | | | | | | | |\n| | | | | | | | | | | | | | | | | |\n| | -8 | -4 | 0 | 4 | 8 | 12 | 16 | 20 | 24 | | | | | | | |\n\nAngle of yaw, $\\Psi$, deg\nAngle of flow deflection, deg\n\no Experimental data\n--- Linearized theory (equation (9))\n— Linearized theory (equation (8))\n\n[Figure: NACA logo]\n\nFigure 7. - Variation of pressure coefficient with angle of yaw for $\\theta$ of $180^\\circ$.\n\nNACA RM NO. E8K05\n```", "timestamp": "2026-07-22T05:10:02.001174+00:00"} | |
| {"citation_id": "19930082614", "source_url": "https://ntrs.nasa.gov/api/citations/19930082614/downloads/19930082614.pdf", "page_number": 27, "total_pages": 36, "image_filename": "19930082614_p27.jpg", "text": "NACA TN 1939\n\n1200\n1000\n800\n600\n400\n200\n0\nTrue airspeed, V, ft/sec\n\n.001 .002 .004 .01 .02 .04 .06 .1 .2 .4 .6 .8x10⁻³\nDeceleration factor, K, per ft\n\nFlight-path angle, γ\nRate of descent, Vd\n\n-2°\n-4°\n-6°\n-8°\n-10°\n-12°\n-14°\n-16°\n-18°\n-20°\n-22°\n-24°\n-26°\n-28°\n-30°\n-32°\n-34°\n-36°\n-38°\n-40°\n-42°\n-44°\n-46°\n-48°\n-50°\n-52°\n-54°\n-56°\n-58°\n-60°\n-62°\n-64°\n-66°\n-68°\n-70°\n-72°\n-74°\n-76°\n-78°\n-80°\n-82°\n-84°\n-86°\n-88°\n-90°\n\n1000\n2000\n4000\n6000\n8000\n10000\n15000\n20000\n25000\n30000\n35000\n40000\n45000\n50000\n55000\n60000\n65000\n70000\n75000\n80000\n85000\n90000\n95000\n100000\n\nVd = 400 ft/min\n\n[Figure: Graph showing effect of deceleration factor K on true airspeed and flight-path angle for constant-speed descent]\n\nFigure 3.— Effect of deceleration factor, K, upon the vertical speed and the flight-path angle for constant-speed descent.\n\nNACA\n\n25", "timestamp": "2026-07-22T05:10:03.875712+00:00"} | |
| {"citation_id": "19930082485", "source_url": "https://ntrs.nasa.gov/api/citations/19930082485/downloads/19930082485.pdf", "page_number": 49, "total_pages": 62, "image_filename": "19930082485_p49.jpg", "text": "48\nNACA TN No. 1810\n\n$$ \\sqrt{z_1}/\\sqrt{z_m} $$\n\n$$ C_1 n_0/2 $$\n\n$$ \\frac{C_1}{\\Delta C} $$\n\n| | |\n| :--- | :--- |\n| - 4.0 | |\n| - 8.0 | |\n| -16.0 | |\n| $\\infty$ | |\n| 16.0 | |\n| 8.0 | |\n| 4.0 | |\n| 3.0 | |\n| 2.0 | |\n| 1.5 | |\n| 1.0 | |\n\n[Figure: Graph with NACA logo]\n\nFigure 17.- Variation with channel width and blade curvature of ratio of velocities along blade suction surfaces to velocity at channel center. (A 10-by-16-in. print of this figure is attached.)", "timestamp": "2026-07-22T05:10:05.262414+00:00"} | |
| {"citation_id": "19930082245", "source_url": "https://ntrs.nasa.gov/api/citations/19930082245/downloads/19930082245.pdf", "page_number": 53, "total_pages": 66, "image_filename": "19930082245_p53.jpg", "text": "```markdown\n1.4\n1.2\n1.0\n.8\n.6\n.4\n.2\n0\n-.2\n-.4\n-.6\n.1 .2 .3 .4 .5 .6 .7 .8 .9\nMach number, M\nAileron section normal-force coefficient, $C_{n_a}$\n\n$\\delta_a$ (deg)\n30\n18\n12\n6\n4\n2\n0\n-2\n-4\n-6\n-12\n\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\nAileron section hinge-moment coefficient, $C_h$\n\n$\\delta_a$ (deg)\n-12\n-6\n-4\n-2\n0\n2\n4\n6\n12\n18\n30\n\n(c) $C_n = 0$.\nFigure 10.-Continued.\n\nNACA\nNACA TN No. 1596\n58\n```", "timestamp": "2026-07-22T05:10:07.699246+00:00"} | |
| {"citation_id": "19930085542", "source_url": "https://ntrs.nasa.gov/api/citations/19930085542/downloads/19930085542.pdf", "page_number": 14, "total_pages": 46, "image_filename": "19930085542_p14.jpg", "text": "12\nNACA RM No. L8L29\n\ncertain characteristics but for others, particularly the lift-curve slope,\ndamping in roll, and aerodynamic-center position, the agreement was poor\nexcept at the lowest test aspect ratio (A = 1.07).\n\n3. The vertical fins tested provided good directional stability for\nthe wing-fin combinations throughout the lift-coefficient range. The fins\nincreased the damping in roll and decreased the variation of effective\ndihedral parameter with lift coefficient.\n\n4. The series of wings obtained by cutting various portions from\nthe tips of a basic triangular wing generally had good longitudinal and\ndirectional stability but very high effective dihedral. Most of the\ncharacteristics of these wings, at low lift coefficients, could be\npredicted with fair accuracy by means of available swept-wing theory.\n\nLangley Aeronautical Laboratory\nNational Advisory Committee for Aeronautics\nLangley Air Force Base, Va.", "timestamp": "2026-07-22T05:10:07.802947+00:00"} | |
| {"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 27, "total_pages": 78, "image_filename": "19930082618_p27.jpg", "text": "NACA TN 1945\n\nSection lift coefficient, $c_l$\n\nMoment coefficient, $c_{m_{c/4}}$\n\nSection angle of attack, $\\alpha$, deg\n\nR\n0.7 x $10^6$\n1.5\n2.0\n3.0\n4.0\n6.0\n9.0\n\nFlagged symbols denote\nstandard roughness\n\nNACA\n\n(a) Section lift and pitching-moment characteristics of the plain airfoil section.\n\nFigure 2.— Aerodynamic characteristics of the NACA 64$_1$-412 airfoil section, 24-inch chord.\n\n25", "timestamp": "2026-07-22T05:10:09.197597+00:00"} | |
| {"citation_id": "19930085471", "source_url": "https://ntrs.nasa.gov/api/citations/19930085471/downloads/19930085471.pdf", "page_number": 27, "total_pages": 28, "image_filename": "19930085471_p27.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:10:12.234844+00:00"} | |
| {"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 42, "total_pages": 99, "image_filename": "19930082511_p42.jpg", "text": "40\nNACA TN No. 1826\n\nSYMBOLS\n\n$\\xi, \\eta, \\zeta$\nrectangular coordinates in units of the tunnel radius\nwith origin at lifting element (see fig. 23)\n\n$\\xi, \\rho, \\theta$\ncylindrical coordinates (see fig. 23)\n\na, b\n$\\xi$-coordinate of entrance and exit lips, respectively\n\n$\\beta$\nvariable of integration\n\nq\nvariable of integration (see reference 13)\n\n$\\Phi_0$\ndisturbance potential associated with body (or with\nvortex system)\n\n$\\Phi$\ntunnel-induced potential\n\n$\\Phi_C$\ntunnel-induced potential in closed circular tunnel\n\n$\\Phi_A$\nresidual potential $(\\Phi - \\Phi_C)$\n\n$J_m$\nBessel function of the first kind of order m\n\nu\nconstant longitudinal perturbation velocity on free surface\n\n$\\varepsilon_m^{(1)}(\\xi)$\nm$^{th}$ Fourier sine coefficient of $\\left. \\frac{\\partial \\Phi_A}{\\partial \\rho} \\right|_{\\rho=1}$\n\n$\\varepsilon_m^{(2)}(\\xi)$\nm$^{th}$ Fourier cosine coefficient of $\\left. \\frac{\\partial \\Phi_A}{\\partial \\rho} \\right|_{\\rho=1}$\n\n$h_{mn}^{(j)}$\nn$^{th}$ coefficient in series for $\\varepsilon_m^{(j)}(\\xi)$\n\n$y_{sm}$\ns$^{th}$ zero of $J_m'$ (not including the zero at the origin)\n\n$r_m^{(1)}$\nm$^{th}$ Fourier sine coefficient of $- \\left. \\frac{\\partial(\\Phi_0 + \\Phi_C)}{\\partial \\xi} \\right|_{\\rho=1}$\n\n$r_m^{(2)}$\nm$^{th}$ Fourier cosine coefficient of $- \\left. \\frac{\\partial(\\Phi_0 + \\Phi_C)}{\\partial \\xi} \\right|_{\\rho=1}$\n\n$\\delta$\ntunnel-induced velocity parameter $\\left( \\frac{\\pi}{V} \\frac{\\pi R^2 \\rho V^2}{2L} \\right)$", "timestamp": "2026-07-22T05:10:13.557810+00:00"} | |
| {"citation_id": "19930082496", "source_url": "https://ntrs.nasa.gov/api/citations/19930082496/downloads/19930082496.pdf", "page_number": 39, "total_pages": 50, "image_filename": "19930082496_p39.jpg", "text": "38\n\nPage intentionally left blank\n\nPage intentionally left blank", "timestamp": "2026-07-22T05:10:16.540523+00:00"} | |
| {"citation_id": "19930082914", "source_url": "https://ntrs.nasa.gov/api/citations/19930082914/downloads/19930082914.pdf", "page_number": 28, "total_pages": 66, "image_filename": "19930082914_p28.jpg", "text": "NACA TN No. 1857\n\nThe constant of proportionality $c$ is $\\sqrt{1/2\\sigma^3}$. For the subsonic jets, with $\\sigma = 12$, the proportionality constant $c$ has the value of 0.017. For the supersonic jet, with $\\sigma = 15$, the value of $c$ is 0.012. By the use of the hot-wire method, Liepmann and Laufer have shown that the \"mixing length\" is not constant across the mixing zone. If $\\sigma$ is taken, notwithstanding this fact, to be a measure of the amount of turbulence, then the present measurements show that the mixing region of a supersonic jet is less turbulent than the mixing region of a low-speed jet.\n\nFigure 17 shows that the velocity distributions at the various cross sections are similar. The figure also shows that the flow in the mixing zone is turbulent, because the rate of spread is linear, in agreement with theory, as compared to laminar flow, in which the rate of spread is a function of the Reynolds number based on the x-coordinate.\n\nFigures 15 and 17 show density and velocity distributions, respectively, at seven cross sections. These cross sections lie between 2 and $7\\frac{1}{2}$ inches downstream from the nozzle. Interferograms of the region from $7\\frac{1}{2}$ to 10 inches downstream were also taken and density and velocity distributions were obtained. These distributions were very similar to those shown in figures 15 and 17. The distributions so obtained, however, at these distances from the nozzle are not necessarily accurate. At 10 inches, the path length of the light through the mixing regions on the two sides of the jet is about 3 inches. The effect of this on the results was reduced by the method described in a previous section. When the effect is large, however, it is not certain that the method of correcting is at all accurate. For that reason no results are shown in the present paper for distances greater than $7\\frac{1}{2}$ inches from the nozzle.\n\nSUMMARY OF RESULTS\n\nIt has been found that, for the free supersonic jet of Mach number 1.6,\n\n1. Density distributions through the mixing region were similar to each other at the cross sections investigated.\n\n2. Velocity distributions through the mixing region were similar to each other at the cross sections investigated and were similar to Tollmien's theoretical velocity distribution in the subsonic portion of the mixing region.", "timestamp": "2026-07-22T05:10:17.513070+00:00"} | |
| {"citation_id": "19930085519", "source_url": "https://ntrs.nasa.gov/api/citations/19930085519/downloads/19930085519.pdf", "page_number": 24, "total_pages": 46, "image_filename": "19930085519_p24.jpg", "text": "NACA RM No. L8X19\n\nSection A-A\n\n.077$\\frac{b}{2}$\n\n.063\"\n\n.077$\\frac{b}{2}$\n\nFigure 5.- Locations and details of flap-slot flow-control vanes A on the 42° sweptback wing.\n\nNACA\n\n23", "timestamp": "2026-07-22T05:10:22.482549+00:00"} | |
| {"citation_id": "19930085879", "source_url": "https://ntrs.nasa.gov/api/citations/19930085879/downloads/19930085879.pdf", "page_number": 6, "total_pages": 29, "image_filename": "19930085879_p6.jpg", "text": "4\nNACA RM L9D11\n\nModel B\n\nModel B was launched smoothly and good telemeter signals were received. Rocket thrust decreased after 8.4 seconds of flight and deceleration of the model started at 8.6 seconds. The nose was jettisoned at 9.86 seconds (M = 0.87) producing an instantaneous forward acceleration on the nose section of 10.8g and instantaneous transverse and normal accelerations of 2.4g and 0.75g, respectively. After ejection, the drag of the nose section produced a longitudinal deceleration of about 3.25g (maximum), gradually decreasing as the speed decreased. The normal and transverse accelerations after separation were limited to small oscillations of 0g to 1.3g maximum. A plot of accelerations against time during ejection is shown in figure 8.\n\nReadings of the accelerometer in the rear section were obtained for about 0.3 second after separation, making it possible to calculate the drag of the rear body until it was about 5.5 feet behind the nose section. A plot of the longitudinal accelerations of the nose and rear sections immediately following ejection is shown in figure 9. The accelerometer traces are rough because the nose was oscillating. A plot of drag coefficient $C_D$ against separation distance is shown in figure 10. These drag coefficients are based on body cross-section area at the separation station. Since both bodies were free, there is no assurance that the nose was not displaced somewhat laterally or normally from the rear body as they separated.\n\nThe continuous-wave Doppler radar gave a record of the velocities of both sections up to 14.72 seconds. This is presented as a velocity-time plot in figure 11, and separation velocity-time plot in figure 12.\n\nA plot of flight time against altitude as obtained from the pulse-type tracking radar is shown in figure 13. The pulse-type radar read the rear-body position from separation until 25 seconds after launching, both sections from 25 to 26 seconds after launching, and thereafter only the nose section. Some intermediate values were obtained by integrating the velocity-time curve plot of the continuous-wave Doppler record. Also included in figure 13 are the atmospheric conditions at time of flight as recorded by the radiosonde.\n\nWhile there was no instrumentation to record operation of the flap motor, reduction of the drag data to a plot of drag coefficient against velocity (fig. 14) indicates that they operated, but in an irregular manner. An integration of the accelerometer record indicates that the terminal velocity was approximately 500 feet per second at sea level.", "timestamp": "2026-07-22T05:10:22.910239+00:00"} | |
| {"citation_id": "19930085487", "source_url": "https://ntrs.nasa.gov/api/citations/19930085487/downloads/19930085487.pdf", "page_number": 27, "total_pages": 36, "image_filename": "19930085487_p27.jpg", "text": "NACA RM No. E8J22\n25\n\n100,000\n\n651\n\nFrequency, cps\n\n10,000\nInlet guide vane, 56 blades\nFourth bending mode\nSecond torsional mode\nThird bending mode\n\nRear stator row, 22 blades\nFirst torsional mode\nSecond bending mode\n\n1,000\nFourth order, front bearing supports\nFirst bending mode\nSecond order, split compressor case\nFirst order, rotor speed\n\n100\n4\n6\n8\n10\n12\n14\n16x10³\nRotor speed, rpm\n\nNACA\n\n(a) First stage.\nFigure 9. - Critical-speed diagrams for 10 stages of compressor rotor.", "timestamp": "2026-07-22T05:10:30.564790+00:00"} | |
| {"citation_id": "19930082498", "source_url": "https://ntrs.nasa.gov/api/citations/19930082498/downloads/19930082498.pdf", "page_number": 36, "total_pages": 49, "image_filename": "19930082498_p36.jpg", "text": "NACA TN No. 1838\n\n[Figure: Two expansion-chamber mufflers with elliptical cross sections, labeled 63 and 66, lying on a concrete surface. A ruler is placed on each for scale. A NACA stamp with “L-53825” is visible in the lower right corner of the image.]\n\n(d) Expansion-chamber mufflers with elliptical cross section.\n\nFigure 2.— Continued.\n\n35", "timestamp": "2026-07-22T05:10:34.296661+00:00"} | |
| {"citation_id": "19930085626", "source_url": "https://ntrs.nasa.gov/api/citations/19930085626/downloads/19930085626.pdf", "page_number": 8, "total_pages": 24, "image_filename": "19930085626_p8.jpg", "text": "NACA RM No. L8K23 CONFIDENTIAL 7\n\nTABLE I\n\nPHYSICAL CHARACTERISTICS OF TEST VEHICLES WITH ORIGINAL AILERONS\n\nTotal exposed wing area, sq ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 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. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 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. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .", "timestamp": "2026-07-22T05:10:37.321759+00:00"} | |
| {"citation_id": "19930085881", "source_url": "https://ntrs.nasa.gov/api/citations/19930085881/downloads/19930085881.pdf", "page_number": 1, "total_pages": 31, "image_filename": "19930085881_p1.jpg", "text": "NACA RM L9D12\n\nCONFIDENTIAL\n\nCopy\nRM L9D12\n\nNACA\n\nRESEARCH MEMORANDUM\n\nEFFECTS OF SOME AIRFOIL-SECTION VARIATIONS ON WING-\nAILERON ROLLING EFFECTIVENESS AND DRAG AS DETERMINED IN\nFREE FLIGHT AT TRANSONIC AND SUPERSONIC SPEEDS\n\nBy\nCarl A. Sandahl, William M. Bland, Jr., and H. Kurt Strass\n\nLangley Aeronautical Laboratory\nLangley Air Force Base, Va.\n\nCLASSIFIED DOCUMENT\n\nThis document contains information affecting the national defense of the United States within the meaning of the Espionage Act, USC 50:31 and 32. Its transmission or the revelation of its contents in any manner to an unauthorized person is prohibited by law.\n\nCLASSIFICATION CHANGED TO UNCLASSIFIED\nAUTHORITY: NACA RESEARCH ABSTRACT No. 99\nDATE: APRIL 9, 1956\nWHL\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\nWASHINGTON\nJuly 22, 1949\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:10:38.787509+00:00"} | |
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