PDF/ocr_dataset_1041.jsonl

{"citation_id": "19930085972", "source_url": "https://ntrs.nasa.gov/api/citations/19930085972/downloads/19930085972.pdf", "page_number": 8, "total_pages": 46, "image_filename": "19930085972_p8.jpg", "text": "6\nNACA RM L9B18\n\nEffect of vertical location of the horizontal tail.- Decreasing the height of the horizontal tail above the wing from the basic position to the alternate position (fig. 1) showed little effect on $n_p$ for the configurations investigated (figs. 12(a), 12(b), and 14). The stability was slightly lower at the higher lift coefficients with the tail in the alternate position owing mainly to a less favorable downwash gradient.\n\nModifications to Model with $15^\\circ$ Sweepback\n\nInasmuch as the configuration with $15^\\circ$ sweepback possessed favorable stability characteristics at the stall, this configuration was adopted as the basic low-speed arrangement and various modifications were investigated in an attempt to reduce the large ($0.56c'$) basic shift in neutral point accompanying the reduction in sweep from $45^\\circ$ to $15^\\circ$.\n\nEffect of cutout profile.- The effects of various cutout arrangements are presented in figure 12(a). From stability considerations, the faired cutout appeared to be superior to the unfaired cutouts and this arrangement was used for the majority of tests with cutouts.\n\nEffect of external airfoil flaps.- In an effort to compensate for the forward movement of the wing aerodynamic center caused by the cutout and at the same time introduce a field of upwash in the vicinity of the cutout, tests were made of a configuration employing essentially full-span external flaps (figs. 4 and 5). The results obtained for the various arrangements tested are presented in figures 7(c), 7(d), 12(c), and 12(d). A comparison of these results with those for the configuration without flaps (figs. 7(a), 7(b), 12(a), and 12(b)) indicate that the flap caused an additional rearward movement of the neutral point of only about $0.02c'$ ($\\Lambda = 0^\\circ$). The flap arrangement which was used in this investigation was not a particularly effective one, however, as is indicated by the rather low lift and pitching-moment increments produced by the flap. (Compare figs. 7(a) with 7(c).) It is possible that, with a well-designed extensible-slotted-flap arrangement, the rearward neutral-point movement resulting from the deflected flap for the configuration with the wing cutout would be considerably increased.\n\nSharp leading edge and wing vane.- Several sharp-leading-edge sections and wing vanes mounted on the inboard sections of the wing panel were investigated in an attempt to reduce the lift on this portion of the wing and thereby increase the stability by reducing the downwash gradient. None of these modifications changed the stability characteristics appreciably. Typical results obtained with the sharp leading edge and wing vane are presented in figures 12(e) and 12(f).", "timestamp": "2026-07-22T05:34:36.155848+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 67, "total_pages": 99, "image_filename": "19930082511_p67.jpg", "text": "NACA TN No. 1826\n\nThe second term in the braces of equation (C1) is not a one-term Fourier series. Its $n^{\\text{th}}$ Fourier coefficient is given by a constant times\n\n$$\n\\int_{0}^{2\\pi} \\frac{\\xi \\sin \\theta \\sin n\\theta}{\\sigma(\\xi^2 + \\sin^2\\theta)} \\left(1 - \\frac{1}{\\sqrt{1 + \\xi^2\\sigma^2}}\\right) d\\theta\n$$\n\nInserting this expression in the inner integral of equation (8) of reference 13, and reversing the order of integration gives\n\n$$\n\\int_{0}^{2\\pi} \\sin \\theta \\sin n\\theta \\int_{-\\infty}^{\\infty} \\frac{\\beta}{\\sigma(\\beta^2 + \\sin^2\\theta)} \\left(1 - \\frac{1}{\\sqrt{1 + \\beta^2\\sigma^2}}\\right) \\cos q(\\beta - \\xi) \\, d\\beta \\, d\\theta\n$$\n\nAfter substitution of $p = \\beta\\sigma$, the limit of the inner integral becomes\n\n$$\n\\lim_{\\sigma \\to 0} \\frac{1}{\\sigma} \\int_{-\\infty}^{\\infty} \\frac{p}{p^2 + \\sigma^2\\sin^2\\theta} \\left(1 - \\frac{1}{\\sqrt{1 + p^2}}\\right) \\cos q\\left(\\frac{p}{\\sigma} - \\xi\\right) dp\n$$\n\nIntegration by parts and elimination of terms in $\\sigma^2$ reduces this expression to\n\n$$\n\\lim_{\\sigma \\to 0} \\frac{1}{q} \\int_{-\\infty}^{\\infty} \\left[\\frac{1}{p^2} - \\frac{2p^2 + 1}{p^2(1 + p^2)^{3/2}}\\right] \\sin q\\left(\\frac{p}{\\sigma} - \\xi\\right) dp\n$$\n\nwhich is zero, by the Riemann-Lebesgue lemma.\n\nFinally, then, for the unit doublet $\\left(\\frac{\\Gamma\\sigma}{4\\pi} = 1\\right)$\n\n$$\n\\phi_2 = \\frac{2 \\sin \\theta}{\\pi} \\int_{0}^{\\infty} \\frac{J_1(iq)}{iqJ_1'(iq)} \\left[qK_0(q) + K_1(q)\\right] \\sin q\\xi \\, dq\n$$\n\n$$\n\\left.\\frac{\\partial\\phi_2}{\\partial\\xi}\\right|_{\\rho=1} = \\frac{2 \\sin \\theta}{\\pi} \\int_{0}^{\\infty} \\frac{J_1(iq)}{iJ_1'(iq)} \\left[qK_0(q) + K_1(q)\\right] \\cos q\\xi \\, dq\n$$", "timestamp": "2026-07-22T05:34:43.943387+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 36, "total_pages": 96, "image_filename": "19930085880_p36.jpg", "text": "34\nNACA RM No. L9C03\n\nWetted area\n(sq ft)\n0\n\nLift, lb\n.35\n.30\n.25\n.20\n.15\n.10\n.05\n0\n\nSpeed, fps\n0 10 15 20 25 30 35\n\n.05\n.10\n.15\n.20\n.25\n.30\n\n[Figure: A graph plotting Lift (lb) against Speed (fps) for various Wetted areas (sq ft). A small diagram of a triangle is shown in the upper left corner of the graph area. The NACA logo is in the lower right corner of the graph area.]\n\n(d) $\\tau = 20^\\circ$.\n\nFigure 13.- Concluded.", "timestamp": "2026-07-22T05:34:44.972467+00:00"}
{"citation_id": "19930085542", "source_url": "https://ntrs.nasa.gov/api/citations/19930085542/downloads/19930085542.pdf", "page_number": 38, "total_pages": 46, "image_filename": "19930085542_p38.jpg", "text": "36\nNACA RM No. L8L29\n\nModel No. A\n7 4.0\n8 3.0\n9 2.0\n10 1.0\n\nAngle of attack, $\\alpha$, deg\nLongitudinal-force coefficient, $C_X$\nPitching-moment coefficient, $C_m$\nLift coefficient, $C_L$\n\n[Figure: Graph showing aerodynamic characteristics of modified triangular wings with varying aspect ratios. The graph plots Angle of attack ($\\alpha$), Longitudinal-force coefficient ($C_X$), and Pitching-moment coefficient ($C_m$) against Lift coefficient ($C_L$). Data points for Model Nos. 7, 8, 9, and 10 are shown with different symbols.]\n\nNACA\n\nFigure 17.-- Effect of aspect ratio on aerodynamic characteristics of modified triangular wings. $\\Lambda_{0}/4 = 36.9^{\\circ}$.", "timestamp": "2026-07-22T05:34:48.006522+00:00"}
{"citation_id": "19930085889", "source_url": "https://ntrs.nasa.gov/api/citations/19930085889/downloads/19930085889.pdf", "page_number": 32, "total_pages": 37, "image_filename": "19930085889_p32.jpg", "text": "```markdown\nNACA RM L9F14\n31\n\nCONFIDENTIAL\n\n<!-- Image (239, 117, 808, 874) -->\n\nFigure 13.- Variation of $C_{Y_p}$, $C_{n_p}$, and $C_{l_p}$ with lift coefficient for the wings tested with fuselage.\n```", "timestamp": "2026-07-22T05:34:49.865104+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 9, "total_pages": 114, "image_filename": "19930086061_p9.jpg", "text": "NACA RM L9J07\n5\n\n$\\overline{x}_{\\overline{c}/4}$\ndistance from local center of pressure to $\\overline{c}/4$ in percent\nof local chord, positive when $\\overline{c}/4$ is behind\n$$\n\\left( \\frac{\\int_{0}^{1.0} P \\left( 0.25 - \\frac{x}{c} \\right) d \\left( \\frac{x}{c} \\right)}{c_n} + \\frac{x_{\\overline{c}/4}}{c} \\right)\n$$\n\n$x_{\\overline{c}/4}$\ndistance along chord from $c/4$ to $\\overline{c}/4$, feet\n\ny\ndistance along span from root chord, positive direction\nto the right, feet\n\nMODELS\n\nThe geometric characteristics and principal dimensions of the three\nlow-aspect-ratio pointed wings with $60^\\circ$ sweptback leading edge and varying\ntrailing-edge sweep are given in figure 2. The wings, designated\nhereinafter respectively as wings 1, 2, and 3, had $30^\\circ$, $0^\\circ$, and $-30^\\circ$\ntrailing-edge sweep. All of the wings had 10-percent-thick biconvex\nsections parallel to the plane of symmetry. The aspect ratios\nwere 3.46, 2.31, and 1.73, and the angles of sweep of the quarter-chord\nline were $55.2^\\circ$, $52.4^\\circ$, and $49.1^\\circ$ for wings 1, 2, and 3, respectively.\nA close-up photograph of wing 1 is shown as figure 3(a) and a photograph\nshowing wing 3 mounted in the entrance cone is given as figure 3(b).\n\nWings 2 and 3 were made of $\\frac{3}{32}$-inch sheet brass attached with flush\nrivets to a rigid steel inner structure. Wing 1 was cast of a tin-\nbismuth alloy with a steel insert for added strength. Approximately\n200 orifices were located on the left semispan of each wing at 7 stations,\nhereinafter designated as stations 1, 2, 3, 4, 5, 6, and 7, which were\nlocated at 0, 16.7, 33.3, 50.0, 66.7, 83.3, and 91.6 percent of the\nsemispan from the plane of symmetry, respectively. The chordwise\nlocation of the orifices on each wing is given in table I. The wing\nsupport sting, which served as a conduit for the pressure tubes, was set\noff center on the right semispan and was faired smoothly into the bottom\nsurface near the trailing edge, leaving the upper surface clear of any\nprotuberance.", "timestamp": "2026-07-22T05:34:52.438041+00:00"}
{"citation_id": "19930085912", "source_url": "https://ntrs.nasa.gov/api/citations/19930085912/downloads/19930085912.pdf", "page_number": 20, "total_pages": 36, "image_filename": "19930085912_p20.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:34:54.773906+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 3, "total_pages": 98, "image_filename": "19930086073_p3.jpg", "text": "```markdown\nNACA RM A9H04\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nRESEARCH MEMORANDUM\n\nAN INVESTIGATION AT LOW SPEED OF A LARGE-SCALE TRIANGULAR\nWING OF ASPECT RATIO TWO.- III. CHARACTERISTICS OF\nWING WITH BODY AND VERTICAL TAIL\n\nBy Adrien E. Anderson\n\nSUMMARY\n\nAn investigation has been made to determine the aerodynamic\ncharacteristics in sideslip of a triangular wing of aspect ratio 2.04\nin combination with a body of fineness ratio 12.5 and a vertical tail\nsurface. The airfoil section was a modified symmetrical double wedge\nwith a maximum thickness of 4.76 percent. Force and moment data were\nobtained at several angles of sideslip for various deflections of\nconstant-chord split flaps, semispan split-flap-type ailerons, and a\nconstant-chord rudder. The Reynolds number, as based on the mean aero-\ndynamic chord, was approximately $15.4 \\times 10^6$ and the Mach number 0.13.\n\nThe results of this investigation show that the body combined with\nthe triangular plan-form wing caused no sizable changes in the lift\ncharacteristics of the wing and caused only a 1-percent decrease in the\nstatic margin. Flap lift and pitching-moment effectiveness decreased\nproportional to the decrease in flap area caused by the addition of the\nbody. The wing with body and vertical tail exhibited positive dihedral\neffect throughout the lift range. Directional stability, however,\ndecreased with increasing lift and the model became directionally\nunstable at high lift coefficients. In contrast, rudder effectiveness\nremained nearly constant throughout the lift range. The contribution\nof the vertical tail to the directional stability and the rudder yawing\neffectiveness could be predicted with reasonable accuracy at zero wing\nlift.\n\nINTRODUCTION\n\nA general study of triangular-plan-form wings has been undertaken\nin the Ames 40- by 80-foot wind tunnel to determine their character-\nistics at low speed and large scale. The study of such a plan form\nhaving a symmetrical double-wedge airfoil section was reported in\nreference 1. An investigation into the effects on the longitudinal\ncharacteristics of airfoil-section modifications was carried out and\nreported in reference 2. This report, the third of the series, contains\n```", "timestamp": "2026-07-22T05:34:55.324395+00:00"}
{"citation_id": "19930085962", "source_url": "https://ntrs.nasa.gov/api/citations/19930085962/downloads/19930085962.pdf", "page_number": 16, "total_pages": 51, "image_filename": "19930085962_p16.jpg", "text": "NACA RM A9E05\nCONFIDENTIAL\n\n1.2\n1.0\n.8\n.6\n.4\n.2\n0\n-.2\n-.4\n-.6\n-.8\n\nLift coefficient, $C_L$\n\n$\\delta_e$\n(deg)\n$\\circ$ 0\n$\\square$ 2\n$\\diamond$ 4\n$\\triangle$ 6\n$\\nabla$ 10\n$\\triangleright$ 20\n$\\triangleleft$ 30\n\n-16 -12 -8 -4 0 4 8 12 16\nAngle of attack, $\\alpha$, deg\n\n.12 .08 .04 0 -.04 -.08 -.12 -.16 -.20 -.24 -.28\nPitching-moment coefficient, $C_m$\n\n(a) $C_L$ vs $\\alpha$, $C_L$ vs $C_m$.\n\nFigure 4. — The effect of elevator deflection on the aerodynamic characteristics of the tail at a Mach number of 0.50.\n\nCONFIDENTIAL\n15", "timestamp": "2026-07-22T05:34:56.964932+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 68, "total_pages": 99, "image_filename": "19930082511_p68.jpg", "text": "66\nNACA TN No. 1826\n\nThe potential $\\phi_0$ of a unit doublet line along the axis is\n\n$$\n\\phi_0 = \\frac{\\xi}{\\eta^2 + \\xi^2} \\left( \\frac{\\xi}{\\sqrt{\\xi^2 + \\eta^2 + \\xi^2}} + 1 \\right) = \\frac{\\sin \\theta}{\\rho} \\left( \\frac{\\xi}{\\sqrt{\\xi^2 + \\rho^2}} + 1 \\right)\n$$\n\nwhence\n\n$$\n\\left. \\frac{\\partial \\phi_0}{\\partial \\xi} \\right|_{\\rho=1} = \\frac{\\sin \\theta}{(\\xi^2 + 1)^{3/2}}\n$$\n\nThe flow of the usual reflection vortices for wings of finite span reduces, as the span becomes arbitrarily small, to a uniform upflow in the finite section of the tunnel and therefore contributes nothing to the longitudinal velocity.\n\nThe coefficient of $\\sin \\theta$ in $-\\left. \\frac{\\partial(\\phi_0 + \\phi_0)}{\\partial \\xi} \\right|_{\\rho=1}$ is thus seen to be the expression given in equation (33).", "timestamp": "2026-07-22T05:35:08.426182+00:00"}
{"citation_id": "19930085889", "source_url": "https://ntrs.nasa.gov/api/citations/19930085889/downloads/19930085889.pdf", "page_number": 33, "total_pages": 37, "image_filename": "19930085889_p33.jpg", "text": "32\nNACA RM L9F14\n\nCONFIDENTIAL\n—○— Measured\n- - - - Calculated, reference 2\n\n[Figure: Three plots of $C_{np}$ vs $C_L$ for different sweepback angles]\n\nSweepback, $\\Lambda$, deg\n3.6\n32.6\n46.7\n\nLift coefficient, $C_L$\n\nCONFIDENTIAL\nNACA\n\nFigure 14.- Comparison of the variation with lift coefficient of the values of $C_{np}$ obtained experimentally with those calculated by the method of reference 2.", "timestamp": "2026-07-22T05:35:09.360542+00:00"}
{"citation_id": "19930085964", "source_url": "https://ntrs.nasa.gov/api/citations/19930085964/downloads/19930085964.pdf", "page_number": 15, "total_pages": 18, "image_filename": "19930085964_p15.jpg", "text": "14\nNACA RM E9G25\n\nTrailing\nedge\nExciting frequency, cps\nLeading\nedge\n\n(a) 950\n(b) 950\n(c) 1220\n(d) 1790\n(e) 2170\n(f) 2580\n\n(g) 3090\n(h) 3170\n(i) 3410\n(j) 4110\n(k) 4730\n(l) 5230\n\n(m) 5610\n(n) 6400\n(o) 6930\n(p) 7290\n(q) 7630\n(r) 8020\n\n(s) 8150\n(t) 8760\n(u) 9810\nNACA\n\nFigure 6. - Vibrational modes of stiffened hollow blade B3. (Solid lines represent node lines on concave side of blade; dashed lines represent node lines on convex side. Exciting frequency in cycles per second is shown below each nodal pattern.)\n\n1171\n40-1575", "timestamp": "2026-07-22T05:35:09.674094+00:00"}
{"citation_id": "19930085934", "source_url": "https://ntrs.nasa.gov/api/citations/19930085934/downloads/19930085934.pdf", "page_number": 16, "total_pages": 23, "image_filename": "19930085934_p16.jpg", "text": "NACA RM E9G12\n\nEnthalpy of vapor per pound of dry air at outlet\n\n$h_{s,a,2} = (0.0601)(1179) = 70.86 \\text{ (Btu/lb)}$\n\nEnthalpy of dry air\n\n$h_{d,2} \\text{ at } 264^\\circ \\text{ F} = 77.85 \\text{ (Btu/lb)}$\n\nEnthalpy of mixture per pound of dry air at outlet\n\n$h_{m,a,2} = h_{s,a,2} + h_{d,2}$\n\n$= 70.86 + 77.85$\n\n$= 148.7 \\text{ (Btu/lb)}$\n\n$h_{m,a,1} = 45.25 \\text{ (Btu/lb)} \\text{ from section III}$\n\n(3) Enthalpy change equivalent to actual work. -\n\n$\\Delta h_{m,a} = h_{m,a,2} - h_{m,a,1}$\n\n$= 148.7 - 45.25$\n\n$= 103.5 \\text{ (Btu/lb)}$\n\nV. Adiabatic Efficiency\n\n$\\eta_{ad} = \\frac{\\text{isentropic } \\Delta h_{m,a}}{\\text{actual } \\Delta h_{m,a}}$\n\n$= \\frac{40.93}{103.5} = 0.395$\n\nAPPLICATION OF METHODS\n\nIsentropic outlet temperature, isentropic enthalpy change, and adiabatic efficiency computed for various pressure ratios are shown in figures 1, 2, and 3, respectively. Water-air ratios from 0 to 0.06 are also given. Inlet conditions of pressure, 14 inches of mercury absolute; temperature, $77^\\circ \\text{ F}$; and specific humidity, 0 are", "timestamp": "2026-07-22T05:35:10.658169+00:00"}
{"citation_id": "19930085972", "source_url": "https://ntrs.nasa.gov/api/citations/19930085972/downloads/19930085972.pdf", "page_number": 9, "total_pages": 46, "image_filename": "19930085972_p9.jpg", "text": "NACA RM L9B18\n7\n\nDesign Considerations\n\nThe variation of the static longitudinal stability characteristics with sweepback and model configuration is summarized in figure 15. It is evident that a combination of cutouts and flaps can aid in minimizing the forward neutral-point movement as the sweep angle is decreased to 15° but that translation of the wing is required to compensate for the greater portion of the neutral-point movement. For the sweep range investigated it would be necessary to translate the wing rearward roughly 0.5c! ($\\Lambda = 0^\\circ$) as the sweep angle is decreased to 15° in order to maintain a constant location of the neutral point. It is probable that a sweep angle greater than 15° (but less than 30°) would be satisfactory from low-speed stability considerations. If, for a particular design, the extent of longitudinal wing translation is limited, the maximum sweep angle for which adequate low-speed stability characteristics are attainable should be determined from wind-tunnel experiments.\n\nAlthough the incorporation of a wing capable of translation as well as rotation affords formidable structural problems, the potential aerodynamic rewards incident to their solution are significant. The ability to adjust the sweep angle in flight not only makes it possible to utilize the most efficient sweep angles for high speed and cruising performance but assures stability in the landing configuration without recourse to wing slots or other stall-control devices. The more efficient moderately swept higher-aspect-ratio wing used for the landing condition can also be equipped with conventional high-lift devices and thus provide minimum landing speeds. The wing sweep angle could be adjusted in flight for optimum cruising configuration, and for the highest sweep angles the wing can be translated to compensate partly for the stability changes usually encountered at the higher Mach numbers with swept wings.\n\nIt appears from low-speed stability data that the cutout formed at the wing-fuselage juncture as the wing is rotated forward is beneficial. If high speeds are contemplated with intermediate sweep angles, however, model tests at higher Mach numbers will be required to evaluate the effect of these cutouts.\n\nThe amount of wing translation required is also dependent on the mass distribution of the airplane and the location of the wing pivot point. The weight of the wings, the location of wing fuel tanks as well as fuel tanks in other parts of the airplane, and the plan for emptying the fuel tanks in flight must be considered in evaluating the stability of the airplane. In the case of a variable-sweepback flying wing, for example, the center of gravity would move almost as much as the aerodynamic center of the wing. A study of the unlimited configurations that could utilize center-of-gravity movements created by expendable fuel and moving structural elements is, however, beyond the scope of this paper.", "timestamp": "2026-07-22T05:35:14.810679+00:00"}
{"citation_id": "19930085542", "source_url": "https://ntrs.nasa.gov/api/citations/19930085542/downloads/19930085542.pdf", "page_number": 39, "total_pages": 46, "image_filename": "19930085542_p39.jpg", "text": "NACA RM No. L8L29\n37\n\n$$C_{Y\\psi}$$\n.008\n.004\n0\n-.004\n\n$$C_{n\\psi}$$\n0\n-.002\n-.004\n-.006\n\nModel No. A\n7 4.0\n8 3.0\n9 2.0\n10 1.0\n\n$$C_{z\\psi}$$\n.012\n.010\n.008\n.006\n.004\n.002\n0\n-.002\n\n-2 0 .2 .4 .6 .8 1.0 1.2 1.4 1.6\nLift coefficient, $C_L$\n\n[Figure: NACA logo]\n\nFigure 18.- Effect of aspect ratio of modified triangular wings of NACA 0012 profile on $C_{Y\\psi}$, $C_{n\\psi}$, and $C_{z\\psi}$. $\\Lambda_{c/4} = 36.9^\\circ$.", "timestamp": "2026-07-22T05:35:15.108478+00:00"}
{"citation_id": "19930085912", "source_url": "https://ntrs.nasa.gov/api/citations/19930085912/downloads/19930085912.pdf", "page_number": 21, "total_pages": 36, "image_filename": "19930085912_p21.jpg", "text": "NACA RM No. E9C16\n\n[Figure: Photograph of model installation in test section of icing research tunnel.]\n\nFigure 1. - Photograph of model installation in test section of icing research tunnel.\n\nNACA\nC-21432\n5-13-48\n\n19", "timestamp": "2026-07-22T05:35:16.650118+00:00"}
{"citation_id": "19930085906", "source_url": "https://ntrs.nasa.gov/api/citations/19930085906/downloads/19930085906.pdf", "page_number": 22, "total_pages": 23, "image_filename": "19930085906_p22.jpg", "text": "NACA RM E9F20 CONFIDENTIAL 21\n\nFinal fuel temperature, °F\n\nDistance between preheater and flame holder, in.\n\nFigure 9. - Effect of preheater position on range of final fuel temperature.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:35:20.772684+00:00"}
{"citation_id": "19930085889", "source_url": "https://ntrs.nasa.gov/api/citations/19930085889/downloads/19930085889.pdf", "page_number": 34, "total_pages": 37, "image_filename": "19930085889_p34.jpg", "text": "NACA RM L9F14\n33\n\nCONFIDENTIAL\n\n$C_{lp}$\n.1\n0\n-.1\n-.2\n-.3\n-.4\n\nSweepback, $\\Lambda$, deg\n3.6\n\no Stability tunnel, M=0.13\n□ 7x10 tunnel, M=0.4\n\n$C_{lp}$\n0\n-.1\n-.2\n-.3\n-.4\n\n32.6\n\n$C_{lp}$\n0\n-.1\n-.2\n-.3\n-.4\n\n46.7\n\nNACA\nCONFIDENTIAL\n\n-2 0 .2 .4 .6 .8 1.0 1.2\nLift coefficient, $C_L$\n\nFigure 15.- Comparison of the values of $C_{lp}$ obtained by free rotation of the models in the Langley 7- by 10-foot tunnel with those obtained by the rolling-flow method in the Langley stability tunnel.", "timestamp": "2026-07-22T05:35:33.796959+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 10, "total_pages": 114, "image_filename": "19930086061_p10.jpg", "text": "6\nNACA RM L9J07\n\nT E S T S\n\nThe over-all arrangement of the testing apparatus as located just inside the entrance cone of the Langley full-scale tunnel is sketched in figure 1 and is shown in part by the photograph of figure 3(b). The wings assumed a wide range of positions in the air stream with varying pitch and yaw because the pitch and yaw axes were located 5.3 feet behind the wing apexes at $\\psi = 0^\\circ$ and $\\alpha = 0^\\circ$.\n\nThe air-stream angularity and the distribution of q in the entrance cone were surveyed with a six-prong yaw-pitch head in a vertical plane located 7 inches behind the apexes of the wings at $\\alpha = 0^\\circ$ and $\\psi = 0^\\circ$. The survey was made in 1-foot vertical increments from 5 feet to 10 feet above the tunnel floor, and in $\\frac{1}{2}$-foot horizontal increments through a distance of 4 feet on each side of the wing center lines.\n\nOrifice pressures over the left semispan were recorded through an extensive angle-of-attack range from $-10^\\circ$ to well through the stall angle for yaw angles of $0^\\circ$, $\\pm 2^\\circ$, $\\pm 4^\\circ$, $\\pm 6^\\circ$, $\\pm 8^\\circ$, $\\pm 10^\\circ$, $\\pm 15^\\circ$, $\\pm 20^\\circ$, $\\pm 25^\\circ$, $\\pm 30^\\circ$, and $\\pm 35^\\circ$. All wings were tested at an approximate airspeed of 55 miles per hour or a Mach number of 0.07 and a Reynolds number of $0.57 \\times 10^6$ for wing 1, $0.85 \\times 10^6$ for wing 2, and $1.14 \\times 10^6$ for wing 3. Wing 2 was also tested for the zero-yaw condition at an airspeed of approximately 95 miles per hour corresponding to a Reynolds number of $1.42 \\times 10^6$ in order to obtain an indication of the scale effect. Surface-tuft studies were made on the three wings at several angles of attack for yaw angles of $0^\\circ$, $10^\\circ$, $20^\\circ$, and $35^\\circ$ and tuft-probing studies were made on wing 2 at zero yaw. Extensive smoke studies were made on each wing at yaw angles of $0^\\circ$ and $20^\\circ$ to observe the vortex flow.\n\nR E D U C T I O N O F D A T A\n\nAIR-STREAM FLOW ANALYSIS\n\nResults of the entrance-cone survey show that the $q_1/q$ ratio (ratio of dynamic pressure in the surveyed plane to the reference dynamic pressure at the pitot tube used throughout the tests (fig. 1)) over the region occupied by the left semispan of the wings varied throughout the yaw and angle-of-attack range from about 0.87 to 0.90 (fig. 4). These ratios were low primarily because the reference pitot tube was in a relatively high velocity field; however, the over-all variation in dynamic pressure was of about the same magnitude as reported in reference 9 for", "timestamp": "2026-07-22T05:35:36.272862+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 37, "total_pages": 96, "image_filename": "19930085880_p37.jpg", "text": "NACA RM No. L9C03\n35\n\nLoad, lb\nWetted area, sq ft\nSpeed (fps)\n30\n25\n20\n15\n10\n\n(a) $\\tau = 4^\\circ$.\n\nFigure 14.- Variation of load with wetted area. Model 250A.", "timestamp": "2026-07-22T05:35:36.878630+00:00"}
{"citation_id": "19930085934", "source_url": "https://ntrs.nasa.gov/api/citations/19930085934/downloads/19930085934.pdf", "page_number": 17, "total_pages": 23, "image_filename": "19930085934_p17.jpg", "text": "```markdown\n16\nNACA RM E9G12\n\nused in each case. For any change in inlet conditions, additional\ncurves must be constructed. The slip factor in figure 3 is defined\nas the ratio of the rotational velocity of the fluid particle at\nthe outlet to tangential velocity of the impeller tip.\n\nIn figure 1, the curves for all water-air ratios closely fol-\nlow the same path until the vapor in the mixture passes into the\nsuperheat state, at which point the curve breaks sharply upward.\nAs the water-air ratio increases, the breakaway occurs at increas-\ningly higher pressure ratios. This effect is reflected in fig-\nure 2, which shows that, for any given isentropic enthalpy change,\nthe change in pressure ratio increases at a decreasing rate as\nwater-air ratio is increased. For a given adiabatic efficiency in\nfigure 3, the rate of increase in pressure ratio decreases as water-\nair ratio is increased. A water-air ratio greater than 0.05 is\nrelatively ineffective for pressure ratios of less than 8 for these\ncompressor-inlet conditions. In view of the large difference in\nflow conditions between dry and wet compression, the design of a\ncompressor that will have good efficiencies for both wet and dry\ncompression may be difficult.\n\nFor the inlet conditions used herein, figures 1 and 2 can be\nused as a source of data in computing efficiencies of actual com-\npressors and in making analyses of the turbojet cycle. For the\ncompressor used in a cycle analysis with a given pressure ratio\nand an assumed efficiency, the isentropic enthalpy change can be\ntaken from figure 2 and the actual enthalpy change computed. The\nactual outlet temperature that will give the actual enthalpy change\ncan be found by trial-and-error methods based on use of steam and\nair tables.\n\nSUMMARY OF RESULTS\n\nFrom the method developed for determining centrifugal-flow-\ncompressor performance with water injection, it has been shown that\nfor any given isentropic enthalpy change or adiabatic efficiency,\nthe rate of increase in compressor pressure ratio decreased as the\nwater-air ratio was increased. For compressor-inlet conditions of\npressure, 14 inches of mercury absolute; temperature, $77^\\circ$ F; and\nspecific humidity, 0, water-air ratios greater than 0.05 were rel-\natively ineffective for pressure ratios of less than 8.\n```", "timestamp": "2026-07-22T05:35:39.307548+00:00"}
{"citation_id": "19930085542", "source_url": "https://ntrs.nasa.gov/api/citations/19930085542/downloads/19930085542.pdf", "page_number": 40, "total_pages": 46, "image_filename": "19930085542_p40.jpg", "text": "38\nNACA RM No. L8L29\n\n$$C_{Yp}$$\n.4\n0\n-.4\n\n| Model No. | A |\n| :--- | :--- |\n| 7 | 4.0 |\n| 8 | 3.0 |\n| 9 | 2.0 |\n| 10 | 1.0 |\n\n$$C_{np}$$\n.2\n0\n-.2\n\n$$C_{lp}$$\n.2\n0\n-.2\n-.4\n-.6\n\n[NACA logo]\n\n-.2 0 .2 .4 .6 .8 1.0\nLift coefficient, $C_L$\n\nFigure 19.- Effect of aspect ratio of modified triangular wings of NACA 0012 profile on $C_{Yp}$, $C_{np}$, and $C_{lp}$. $\\Lambda_{c/4} = 36.9^\\circ$.", "timestamp": "2026-07-22T05:35:39.963644+00:00"}
{"citation_id": "19930085881", "source_url": "https://ntrs.nasa.gov/api/citations/19930085881/downloads/19930085881.pdf", "page_number": 24, "total_pages": 31, "image_filename": "19930085881_p24.jpg", "text": "22\nNACA RM L9D12\n\nCONFIDENTIAL\n\n.12\n.08\n$C_D$\n.04\n0\n\n.12\n.08\n$pb/2V$\n.04\n0\n-.04\n\nModel $\\delta_a$ (deg) $i_w$ (deg)\n— 117c 5.0 0\n--- 117d 5.1 -0.08\n\n.6 .8 1.0 1.2 1.4 1.6 1.8 2.0\nM\n\n[Figure: NACA logo]\n\n(d) 9-percent-thick circular-arc airfoil section\nFigure 5.- Concluded.\nCONFIDENTIAL", "timestamp": "2026-07-22T05:35:42.997720+00:00"}
{"citation_id": "19930085962", "source_url": "https://ntrs.nasa.gov/api/citations/19930085962/downloads/19930085962.pdf", "page_number": 17, "total_pages": 51, "image_filename": "19930085962_p17.jpg", "text": "16\nCONFIDENTIAL\nNACA RM A9E05\n\n<!-- Image (109, 169, 871, 792) -->\n\nFigure 4. — Concluded.\n(b) $C_L$ vs $C_D$.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:35:45.856390+00:00"}
{"citation_id": "19930085906", "source_url": "https://ntrs.nasa.gov/api/citations/19930085906/downloads/19930085906.pdf", "page_number": 23, "total_pages": 23, "image_filename": "19930085906_p23.jpg", "text": "22\nCONFIDENTIAL\nNACA RM E9F20\n\nHeat absorbed by fuel, Btu/lb\n140\n130\n120\n110\n100\n90\n80\n70\n0\n4\n8\n12\nDistance between preheater and flame holder, in.\n\n[Figure: A graph showing a shaded region bounded by two lines. The x-axis is labeled \"Distance between preheater and flame holder, in.\" and ranges from 0 to 12. The y-axis is labeled \"Heat absorbed by fuel, Btu/lb\" and ranges from 70 to 140. The shaded region is defined by data points connected by lines. The upper boundary connects points approximately at (0, 84), (4, 107), (8, 124), and (12, 132). The lower boundary connects points approximately at (0, 78), (4, 82), (8, 99), and (12, 109).]\n\nNACA\n\nFigure 10. - Effect of preheater position on heat absorbed by fuel.\n\nCONFIDENTIAL\nNACA-Langley - 9-1-49 - 350", "timestamp": "2026-07-22T05:35:46.050852+00:00"}
{"citation_id": "19930085972", "source_url": "https://ntrs.nasa.gov/api/citations/19930085972/downloads/19930085972.pdf", "page_number": 10, "total_pages": 46, "image_filename": "19930085972_p10.jpg", "text": "8\nNACA RM L5B18\n\nCONCLUSIONS\n\nThe results of a low-speed wind-tunnel investigation of a complete model having a variable-sweep wing which was tested at $45^\\circ$, $30^\\circ$, $15^\\circ$, and $0^\\circ$ sweepback indicated the following conclusions:\n\n1. Stability at the stall was obtained for the configuration with $15^\\circ$ sweepback without recourse to stall-control devices.\n\n2. The shift in neutral point as the sweep was varied from $45^\\circ$ to $15^\\circ$ was decreased from 56 percent of the chord ($c'$ at $\\Lambda = 0^\\circ$) in the original case to 47 percent by the most effective combination of the modifications tested.\n\n3. It seems unlikely that satisfactory stability for all flight conditions can be achieved with a variable-sweep wing without recourse to relative translation between the wing and the center of gravity of the airplane.\n\nLangley Aeronautical Laboratory\nNational Advisory Committee for Aeronautics\nLangley Air Force Base, Va.\n\nREFERENCES\n\n1. Spearman, M. Leroy, and Comisarow, Paul: An Investigation of the Low-Speed Static Stability Characteristics of Complete Models Having Sweptback and Sweptforward Wings. NACA RM No. L8H31, 1948.\n\n2. Gillis, Clarence L., Polhamus, Edward C., and Gray, Joseph L., Jr.: Charts for Determining Jet-Boundary Corrections for Complete Models in 7- by 10-Foot Closed Rectangular Wind Tunnels. NACA ARR No. L5G31, 1945.\n\n3. Thom, A.: Blockage Corrections in a Closed High-Speed Tunnel. R. & M. No. 2033, British A.R.C., 1943.", "timestamp": "2026-07-22T05:35:47.294913+00:00"}
{"citation_id": "19930085977", "source_url": "https://ntrs.nasa.gov/api/citations/19930085977/downloads/19930085977.pdf", "page_number": 7, "total_pages": 33, "image_filename": "19930085977_p7.jpg", "text": "6 CONFIDENTIAL NACA RM L9H22\n\nfloating angles measured are a measure of the angle-of-zero pitching moment about the tail pivot axis rather than the angle-of-zero lift. It has been estimated, however, that for this tail arrangement a downwash gradient as large as $2^\\circ$ across the span of the tail will result in an error of less than $0.2^\\circ$ in the measured downwash angle.\n\nThe total-pressure readings were obtained at constant angles of attack through the Mach number range without an end plate on the model to eliminate end-plate wakes and with the support-strut gap sealed with a rubber sponge seal to minimize any strut-leakage effects. The static-pressure values used in computing the dynamic-pressure ratios were obtained by use of a static probe with no model in position.\n\nRESULTS AND DISCUSSION\n\nA table of the figures presenting the results follows:\n\n| | Figure |\n| :--- | :--- |\n| Wing-alone force data . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 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. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .", "timestamp": "2026-07-22T05:35:50.374228+00:00"}
{"citation_id": "19930085914", "source_url": "https://ntrs.nasa.gov/api/citations/19930085914/downloads/19930085914.pdf", "page_number": 20, "total_pages": 42, "image_filename": "19930085914_p20.jpg", "text": "```markdown\nNACA RM A9D25\n\nLift coefficient, $C_L$\nDrag coefficient, $C_D$\n\n| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | |", "timestamp": "2026-07-22T05:35:53.407446+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 4, "total_pages": 98, "image_filename": "19930086073_p4.jpg", "text": "2\nNACA RM A9H04\n\nthe results of the investigation into the effects of sideslip on the\ncharacteristics of the wing alone, the wing plus body, and the wing plus\nbody and vertical tail.\n\nNOTATION\n\nThe standard NACA coefficients and symbols used within this report\nare defined as follows and in figure 1:\n\n| | |\n| :--- | :--- |\n| A | aspect ratio $\\left(\\frac{b^2}{S}\\right)$ |\n| b | wing span, feet |\n| c | wing chord, measured parallel to air stream, feet |\n| $\\bar{c}$ | mean aerodynamic chord, measured parallel to air stream $\\left(\\frac{\\int_0^{b/2} c^2 dy}{\\int_0^{b/2} c \\ dy}\\right)$, feet |\n| $C_L$ | lift coefficient $\\left(\\frac{\\text{lift}}{qS}\\right)$ |\n| $C_D$ | drag coefficient $\\left(\\frac{\\text{drag}}{qS}\\right)$<br>(Drag, as used herein, is defined as the component of the resultant force acting along the X axis, fig. 1.) |\n| $C_{D_T}$ | increment of drag coefficient due to wind-tunnel-wall interference |\n| $C_Y$ | side-force coefficient $\\left(\\frac{\\text{side force}}{qS}\\right)$<br>(Side force, as used herein, is defined as the component of the resultant force acting along the Y axis, fig. 1.) |\n| $C_m$ | pitching-moment coefficient $\\left(\\frac{\\text{pitching moment}}{qS\\bar{c}}\\right)$ |\n| $C_l$ | rolling-moment coefficient $\\left(\\frac{\\text{rolling moment}}{qSb}\\right)$ |\n| $C_n$ | yawing-moment coefficient $\\left(\\frac{\\text{yawing moment}}{qSb}\\right)$ |", "timestamp": "2026-07-22T05:35:53.843690+00:00"}
{"citation_id": "19930085964", "source_url": "https://ntrs.nasa.gov/api/citations/19930085964/downloads/19930085964.pdf", "page_number": 16, "total_pages": 18, "image_filename": "19930085964_p16.jpg", "text": "NACA RM E9G25\n15\n\nTrailing\nedge\nExciting frequency, cps\nLeading\nedge\n\n(a) 880\n(b) 955\n(c) 1210\n(d) 1490\n(e) 1860\n\n(f) 2260\n(g) 2580\n(h) 3130\n(i) 3400\n(j) 4330\n\n(k) 4880\n(l) 5220\n(m) 5690\n(n) 6790\n(o) 7000\n\n(p) 7360\n(q) 7810\n(r) 8210\n(s) 8590\n(t) 9750\n\nFigure 7. - Vibrational modes of stiffened hollow blade $B_4$. (Solid lines represent node lines on concave side of blade; dashed lines represent node lines on convex side. Exciting frequency in cycles per second is shown below each nodal pattern.)", "timestamp": "2026-07-22T05:35:54.558227+00:00"}
{"citation_id": "19930085889", "source_url": "https://ntrs.nasa.gov/api/citations/19930085889/downloads/19930085889.pdf", "page_number": 35, "total_pages": 37, "image_filename": "19930085889_p35.jpg", "text": "34\nNACA RM L9F14\n\nCONFIDENTIAL\nStability tunnel, M = .13\n7x10 tunnel, M = .40\nWeissinger, $q_0 = 2\\pi$, M = 0\nWeissinger, $q_0 = 2\\pi$, M = .40\n\n$C_{l_p}$\n0\n-.1\n-.2\n-.3\n-.4\n-.5\n\n0 10 20 30 40 50\nAngle of sweepback, $\\Lambda$, deg\n\nCONFIDENTIAL\nNACA\n\nFigure 16.— Comparison of the variation with sweep of the values of $C_{l_p}$ obtained by free rotation of the models in the Langley 7- by 10-foot tunnel and by the rolling-flow method of the Langley stability tunnel with those calculated by the Weissinger method. $C_L = 0$.", "timestamp": "2026-07-22T05:35:56.771411+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 38, "total_pages": 96, "image_filename": "19930085880_p38.jpg", "text": "36\nNACA RM No. L9C03\n\nLoad, lb\nSpeed (fps)\nWetted area, sq ft\n\n(b) $\\tau = 8^\\circ$.\nFigure 14.- Continued.", "timestamp": "2026-07-22T05:35:58.634420+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 69, "total_pages": 99, "image_filename": "19930082511_p69.jpg", "text": "NACA TN No. 1826\n67\n\nREFERENCES\n\n1. Prandtl, L.: Theory of Lifting Surfaces. Part II. NACA TN No. 10, 1920.\n\n2. Theodorsen, Theodore, and Silverstein, Abe: Experimental Verification of the Theory of Wind-Tunnel Boundary Interference. NACA Rep. No. 478, 1934.\n\n3. Weinig, F.: Der Strahleinfluss bei offenen Windkanälen. Luftfahrtforschung, Bd. 13, Nr. 7, July 20, 1936, pp. 210-213.\n\n4. Poggi, L.: Sulla variazione da apportarsi ai risultati delle esperienze eseguite al tunnel aerodinamico su di un modello alare. L'Aerotecnica, vol. XI, fasc. 4, April 1931, pp. 424-445.\n\n5. Küchemann, Dietrich, and Vandrey, Friedrich: Über den Einfluss der Düse (oder des Auffangtrichters) auf Widerstandsmessungen im Freistrahl. Z.f.a.M.M., Bd. 21, Nr. 1, Feb. 1941, pp. 17-31.\n\n6. Vandrey, F.: Der Düseneinfluss auf die Windkanalkorrekturen bei ebener Strömung. Jahrb. 1942 der deutschen Luftfahrtforschung, R. Oldenbourg (Munich), pp. I 786-I 793.\n\n7. Toussaint, A.: Experimental Methods - Wind Tunnels. Influence of the Dimensions of the Air Stream. Vol. III of Aerodynamic Theory, div. I, part 1, ch. III, W. F. Durand, ed., Julius Springer (Berlin), 1935, p. 304.\n\n8. Theodorsen, Theodore: The Theory of Wind-Tunnel Wall Interference. NACA Rep. No. 410, 1931.\n\n9. Schliestett, George Van: Experimental Verification of Theodorsen's Theoretical Jet-Boundary Correction Factors. NACA TN No. 506, 1934.\n\n10. Kellogg, Oliver Dimon: Foundations of Potential Theory. Julius Springer (Berlin), 1929, p. 373.\n\n11. Whittaker, E. T., and Watson, G. N.: A Course of Modern Analysis. Fourth ed., Cambridge Univ. Press (London), 1927. (Reprinted 1940.)\n\n12. Madelung, Erwin: Die Mathematischen Hilfsmittel des Physikers (Mathematical Tools for the Physicist). Dover Publications, 1943, p. 16.\n\n13. Eisenstadt, Bertram J.: Boundary-Induced Upwash for Yawed and Swept-Back Wings in Closed Circular Wind Tunnels. NACA TN No. 1265, 1947.", "timestamp": "2026-07-22T05:36:01.855721+00:00"}
{"citation_id": "19930085934", "source_url": "https://ntrs.nasa.gov/api/citations/19930085934/downloads/19930085934.pdf", "page_number": 18, "total_pages": 23, "image_filename": "19930085934_p18.jpg", "text": "NACA RM E9G12\n17\n\nCONCLUSION\n\nOn the basis of the large difference in flow conditions between wet and dry compression, the design of a compressor with good efficiency for both wet and dry compression may be difficult.\n\nLewis Flight Propulsion Laboratory,\nNational Advisory Committee for Aeronautics,\nCleveland, Ohio.\n\nREFERENCES\n\n1. Keenan, Joseph H., and Keyes, Frederick G.: Thermodynamic Properties of Steam. John Wiley & Sons, Inc., 1936.\n\n2. Keenan, Joseph H., and Kaye, Joseph: Thermodynamic Properties of Air. John Wiley & Sons, Inc., 1945.\n\n3. Anon.: The Refrigerating Data Book. Am. Soc. Refrigerating Eng. (New York), 5th ed., 1942, p. 428.", "timestamp": "2026-07-22T05:36:02.292476+00:00"}
{"citation_id": "19930085542", "source_url": "https://ntrs.nasa.gov/api/citations/19930085542/downloads/19930085542.pdf", "page_number": 41, "total_pages": 46, "image_filename": "19930085542_p41.jpg", "text": "NACA RM No. L8L29\n39\n\n<!-- Image (138, 96, 876, 810) -->\n\nFigure 20.- Variation of aerodynamic center $C_{l_{max}}$ and $C_{L\\alpha}$ with aspect ratio for wings of triangular plan form. Profile, NACA 0012; $C_L = 0$.", "timestamp": "2026-07-22T05:36:04.444313+00:00"}
{"citation_id": "19930086076", "source_url": "https://ntrs.nasa.gov/api/citations/19930086076/downloads/19930086076.pdf", "page_number": 1, "total_pages": 50, "image_filename": "19930086076_p1.jpg", "text": "NACA RM E9F09\n\nNACA\n\nRESEARCH MEMORANDUM\n\nDESIGN FACTORS FOR 4- BY 8-INCH RAM-JET COMBUSTOR\n\nBy Donald W. Male and Adolph J. Cervenka\n\nLewis Flight Propulsion Laboratory\nCleveland, Ohio\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\n\nWASHINGTON\nAugust 11, 1949", "timestamp": "2026-07-22T05:36:10.829842+00:00"}
{"citation_id": "19930085962", "source_url": "https://ntrs.nasa.gov/api/citations/19930085962/downloads/19930085962.pdf", "page_number": 18, "total_pages": 51, "image_filename": "19930085962_p18.jpg", "text": "NACA RM A59E05\nCONFIDENTIAL\n\n1.2\n1.0\n.8\n.6\n.4\n.2\n0\n-.2\n-.4\n-.6\n-.8\n-1.6 -1.2 -.8 -.4 0 4 8 12 16\nAngle of attack, $\\alpha$, deg\n\n$\\delta_e$\n(deg)\n$\\circ$ 0\n$\\square$ 2\n$\\diamond$ 4\n$\\triangle$ 6\n$\\nabla$ 10\n$\\blacktriangledown$ 20\n$\\blacktriangle$ 30\n\n.12 .08 .04 0 -.04 -.08 -.12 -.16 -.20 -.24 -.28\nPitching-moment coefficient, $C_m$\n\n(a) $C_L$ vs $\\alpha$, $C_L$ vs $C_m$.\n\nFigure 5. — The effect of elevator deflection on the aerodynamic characteristics of the tail at a Mach number of 0.70.\n\nCONFIDENTIAL\n17", "timestamp": "2026-07-22T05:36:11.391942+00:00"}
{"citation_id": "19930085881", "source_url": "https://ntrs.nasa.gov/api/citations/19930085881/downloads/19930085881.pdf", "page_number": 25, "total_pages": 31, "image_filename": "19930085881_p25.jpg", "text": "NACA RM L9D12\n23\n\nCONFIDENTIAL\n\n.12\n.08\n$C_D$ .04\n0\n\n.16\n.12\n$pb/2V$ .08\n.04\n0\n\nModel $\\delta_0 (deg)$ $i_w (deg)$\n53c 3.5 -0.02\n53e 5.3 -0.20\n53f 5.3 0\n53m 3.7 -0.04\n53n 3.8 -0.19\n\n.6 .8 1.0 1.2 1.4 1.6 1.8 2.0\nM\nNACA\n\n(a) NACA 65A009 airfoil section.\nFigure 6.- Experimental results. $\\Lambda = 45^\\circ$.\nCONFIDENTIAL", "timestamp": "2026-07-22T05:36:11.578098+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 11, "total_pages": 114, "image_filename": "19930086061_p11.jpg", "text": "```markdown\nNACA RM L9J07\n7\n\nthe test section of the Langley full-scale tunnel. The pitch angularity\nof the air stream in the region of the left semispan did not vary\nmaterially throughout the angle-of-attack range (fig. 5(a)) although it\ndid vary from about $1^\\circ$ at $\\psi = 35^\\circ$ to $3^\\circ$ at $\\psi = -35^\\circ$. The air-stream\nyaw angularity varied about $1.5^\\circ$ in the area occupied by the left semi-\nspan (fig. 5(b)). The air-stream pitch angle and the local dynamic\npressure fluctuated noticeably in the lower region of the survey plane.\n\nThe extent to which the indicated asymmetric air flow influenced the\nwing-pressure data cannot be ascertained reliably. The survey must be\nconsidered only as an indication of general effect for the survey was\ntaken in just one plane located 7 inches behind the apexes of the wings\nwhen $\\psi = 0^\\circ$ and $\\alpha = 0^\\circ$, or approximately 0.2c ahead of the mean\naerodynamic chord of wing 2. A comparison of the pressure distributions\nalong the centrally located station 1 at equal positive and negative\nyaw angles might be expected to give an indication of the magnitude of\nthe flow irregularity, especially since station 1 was located on a ridge\n(section A-A of fig. 2(b)) where the local pressures were sensitive to\ncross-flow velocity components. Because, however, these pressures also\nwere very sensitive to minor construction irregularities along the ridge,\nespecially to slight asymmetries in the location of the orifices along\nthe ridge, such a procedure was not considered trustworthy.\n\nCORRECTIONS TO DATA\n\nA constant stream-angle correction, determined by the zero-lift\ncondition at zero yaw, was used throughout the yaw range. Corrections\nfor support-sting interference and for tunnel-boundary effects were\nassumed to be negligible for the tests. At $\\pm 20^\\circ$ and $\\pm 35^\\circ$ yaw, however,\nthe pressure data of the 80-percent and 90-percent chord orifices of\nstation 1 on the bottom surface were not used because of noticeable\nsupport-sting interference.\n\nThe contribution of chord force to the lift and pitching-moment\ncoefficients was considered to be small enough to neglect. A represen-\ntative calculation made for wing 2 at zero yaw and at $24.1^\\circ$ angle of\nattack showed that the greatest increment in the section lift coefficient\ndue to chord force was 2.4 percent at station 2, while the over-all wing-\nlift-coefficient increment was only 1.1 percent.\n\nThe pressure data for all yaw angles were plotted and analyzed, but\nonly the results of representative yaw angles are presented in this\npaper.\n```", "timestamp": "2026-07-22T05:36:12.725071+00:00"}
{"citation_id": "19930085964", "source_url": "https://ntrs.nasa.gov/api/citations/19930085964/downloads/19930085964.pdf", "page_number": 17, "total_pages": 18, "image_filename": "19930085964_p17.jpg", "text": "Exciting frequency, cps\n\nTrailing edge\nLeading edge\n\n(a) 2050\n(b) 2160\n(c) 3625\n(d) 4020\n(e) 6325\n\n(f) 8500\n(g) 8800\n(h) 9450\n(i) 9625\n(j) 10,700\n\nFigure 8. - Vibrational modes of solid blade C. Frequency of fundamental bending mode, 1270 cycles per second. (Exciting frequency in cycles per second is shown below each nodal pattern.)\n\nNACA RM E9B25\n\n117", "timestamp": "2026-07-22T05:36:17.487023+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 39, "total_pages": 96, "image_filename": "19930085880_p39.jpg", "text": "NACA RM No. I9003\n37\n\n[Figure: A line graph plotting Load (lb) against Wetted area (sq ft). The Y-axis ranges from 0 to 32. The X-axis ranges from 0 to .35. The graph contains five curves representing different speeds (10, 15, 20, 25, 30 fps), distinguished by markers (circles, squares, diamonds, triangles). A legend box is present in the upper left corner.]\n\n(c) $\\tau = 12^\\circ$.\nFigure 14.- Continued.", "timestamp": "2026-07-22T05:36:21.074270+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 5, "total_pages": 98, "image_filename": "19930086073_p5.jpg", "text": "NACA RM A9H04\n3\n\n$C_{l\\beta}$ rate of change of rolling-moment coefficient with sideslip, per degree\n\n$C_{n\\beta}$ rate of change of yawing-moment coefficient with sideslip, per degree\n\n$C_{n\\beta_t}$ rate of change with sideslip of yawing-moment coefficient contributed by the vertical tail, per degree\n\n$C_{Y\\beta}$ rate of change of side-force coefficient with sideslip, per degree\n\n$C_{N\\alpha_t}$ rate of change of tail normal-force coefficient with tail angle of attack, per degree\n\nF ratio of exposed rudder area to total rudder area\n\nl tail length, feet\n\nL/D lift-drag ratio\n\nq free-stream dynamic pressure, pounds per square foot\n\n$q_t$ dynamic pressure at tail surface, pounds per square foot\n\nS wing area, square feet\n\n$S_t$ vertical tail area to the body center line, square feet\n\nV free-stream velocity, feet per second\n\n$V_v$ velocity component at tail due to separation vortices, feet per second\n\n$V_R$ resultant velocity at tail, feet per second\n\ny spanwise distance, outboard from wing center line, feet\n\n$\\alpha$ free-stream angle of attack, degrees\n\n$\\alpha_T$ increment of angle of attack due to wind-tunnel-wall interference, degrees\n\n$\\alpha_t$ angle of attack of vertical tail surface, degrees\n\n$\\beta$ angle of sideslip, degrees\n\n$\\delta_a$ split-flap-type aileron deflection, measured perpendicular to hinge line, degrees\n(Subscripts L and R designate left and right aileron, respectively.)", "timestamp": "2026-07-22T05:36:22.077994+00:00"}
{"citation_id": "19930086076", "source_url": "https://ntrs.nasa.gov/api/citations/19930086076/downloads/19930086076.pdf", "page_number": 2, "total_pages": 50, "image_filename": "19930086076_p2.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:36:24.410728+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 70, "total_pages": 99, "image_filename": "19930082511_p70.jpg", "text": "68\nNACA TN No. 1826\n\n14. Watson, G. N.: A Treatise on the Theory of Bessel Functions.\n    Second ed., The Macmillan Co., 1944.\n\n15. Gray, Andrew, Mathews, G. B., and MacRobert, T. M.: A Treatise\n    on Bessel Functions and Their Applications to Physics. Second\n    ed., Macmillan and Co., Ltd., 1931.\n\n16. Jahnke, Eugene, and Emde, Fritz: Tables of Functions with Formulae\n    and Curves. Rev. ed., Dover Publications (New York), 1943.\n\n17. Milne-Thomson, L. M.: The Calculus of Finite Differences.\n    Macmillan and Co., Ltd., 1933.", "timestamp": "2026-07-22T05:36:25.697956+00:00"}
{"citation_id": "19930085889", "source_url": "https://ntrs.nasa.gov/api/citations/19930085889/downloads/19930085889.pdf", "page_number": 36, "total_pages": 37, "image_filename": "19930085889_p36.jpg", "text": "NACA RM L9F14\n35\n\nCONFIDENTIAL\n\n$C_{Yp}$\n.2\n0\n-.2\n\n$C_{np}$\n.1\n0\n-.1\n\n$C_{lp}$\n0\n-.1\n-.2\n-.3\n-.4\n\n[Figure: Graph showing variation of $C_{Yp}$, $C_{np}$, and $C_{lp}$ with lift coefficient. Data points are plotted for two configurations: circles for \"46.7° wing alone\" and squares for \"46.7° wing + fuselage\". The x-axis is labeled \"Lift coefficient, $C_L$\" ranging from -2 to 12. The y-axes are labeled $C_{Yp}$, $C_{np}$, and $C_{lp}$ with respective scales. A NACA logo is present in the bottom right of the graph area. \"CONFIDENTIAL\" is stamped at the top and bottom of the graph area.]\n\n46.7° wing alone\n46.7° wing + fuselage\n\nLift coefficient, $C_L$\n-2 0 .2 .4 .6 .8 1.0 1.2\n\nCONFIDENTIAL\n\nFigure 17.- Variation of $C_{Yp}$, $C_{np}$, and $C_{lp}$ with lift coefficient for the 46.7° sweptback wing tested alone and for the 46.7° sweptback wing tested with fuselage.", "timestamp": "2026-07-22T05:36:25.901327+00:00"}
{"citation_id": "19930085934", "source_url": "https://ntrs.nasa.gov/api/citations/19930085934/downloads/19930085934.pdf", "page_number": 19, "total_pages": 23, "image_filename": "19930085934_p19.jpg", "text": "18\nNACA RM E9G12\n\n[Figure: A graph plotting Isentropic outlet temperature, $t_{1,2}$, °F against Pressure ratio, $P_2/P_1$. The y-axis ranges from 60 to 340. The x-axis ranges from 1.5 to 5.0. The graph contains multiple curves representing different Water-air ratios, w/a. The curves are labeled 0, 0.02, 0.03, 0.04, 0.05, and 0.06. A NACA logo is present in the bottom right corner of the plot area.]\n\n(a) Pressure ratio, 1.5 to 5.0.\n\nFigure 1. - Variation of outlet temperature with pressure ratio for isentropic compression. Compressor-inlet conditions: pressure, 14 inches mercury absolute; temperature, 77° F; specific humidity, 0.", "timestamp": "2026-07-22T05:36:29.135483+00:00"}
{"citation_id": "19930085542", "source_url": "https://ntrs.nasa.gov/api/citations/19930085542/downloads/19930085542.pdf", "page_number": 42, "total_pages": 46, "image_filename": "19930085542_p42.jpg", "text": "40\nNACA RM No. L8L29\n\n$$\n\\frac{\\partial C_{n\\psi}}{\\partial C_L^2}\n$$\n0\n-.004\n-.008\n-.012\n\n$$\n\\frac{\\partial C_{l\\psi}}{\\partial C_L}\n$$\n.016\n.012\n.008\n.004\n0\n0 1 2 3 4\nAspect ratio, A\n\n— Experimental\n--- Reference 2\n—·— Reference 5 (Calculated)\n\n[Figure: NACA logo]\n\nFigure 21.— Variation of $\\partial C_{n\\psi}/\\partial C_L^2$ and $\\partial C_{l\\psi}/\\partial C_L$ with aspect ratio for wings of triangular plan form. Profile, NACA 0012; $C_L = 0$.", "timestamp": "2026-07-22T05:36:29.343750+00:00"}
{"citation_id": "19930085962", "source_url": "https://ntrs.nasa.gov/api/citations/19930085962/downloads/19930085962.pdf", "page_number": 19, "total_pages": 51, "image_filename": "19930085962_p19.jpg", "text": "18\nCONFIDENTIAL\nNACA RM A9E05\n\nLift coefficient, $C_L$\nDrag coefficient, $C_D$\n\n| $\\delta_e$ (deg) | |\n| :--- | :--- |\n| 0 | $\\circ$ |\n| 2 | $\\square$ |\n| 4 | $\\diamond$ |\n| 6 | $\\triangle$ |\n| 10 | $\\nabla$ |\n| 20 | $\\triangleright$ |\n| <30 | $\\triangleleft$ |\n\n[NACA logo]\n\n(b) $C_L$ vs $C_D$.\n\nFigure 5.—Concluded.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:36:30.372506+00:00"}
{"citation_id": "19930085977", "source_url": "https://ntrs.nasa.gov/api/citations/19930085977/downloads/19930085977.pdf", "page_number": 8, "total_pages": 33, "image_filename": "19930085977_p8.jpg", "text": "NACA RM L9H22 CONFIDENTIAL 7\n\nDrag rise at zero lift began at a Mach number of about 0.87 for both the wing-alone and wing-fuselage configurations.\n\nThe lateral center of pressure for the wing alone was located at 42 percent of the semispan at a Mach number of 0.60 at lift coefficients below 0.5. The same lateral center-of-pressure location was obtained at low speed and high Reynolds numbers in the Langley two-dimensional low-turbulence tunnel for a geometrically similar model. The lateral center of pressure gradually moved outboard as the subsonic speeds increased and was located at about 44.5 percent of the semispan at M = 0.93. Between M = 0.93 and 1.05 there was a fairly abrupt inboard movement of $y_{cp}$ to 41 percent of the semispan and this value remained about constant up to M = 1.18. The addition of the fuselage moved $y_{cp}$ inboard from 1 to 2 percent of the semispan through the Mach number range.\n\nPitching-Moment Characteristics\n\nNear the zero lift coefficient the wing-alone aerodynamic center was located at about 24 percent of the mean aerodynamic chord\n\n$$\n\\left( \\frac{\\partial C_m}{\\partial C_L} \\right)_M = 0.01\n$$\n\nat low Mach numbers. The aerodynamic center moved forward about 3 percent of the mean aerodynamic chord as the Mach number was increased to 0.84. In the speed range between M = 0.84 and 1.03 the aerodynamic center moved back to about 37 percent mean aerodynamic chord and thereafter remained about constant up to M = 1.18. The addition of the fuselage moved the aerodynamic center forward about 7 percent mean aerodynamic chord at the lower Mach numbers and from 4 to 5 percent forward at Mach numbers above unity. By using the theoretical methods of reference 6, it was estimated that the fuselage would move the wing-alone aerodynamic center forward about 6 percent mean aerodynamic chord at low subsonic speeds.\n\nDownwash and Dynamic-Pressure Surveys\n\nThe downwash gradient $d\\epsilon/d\\alpha$ near zero lift for the wing alone was a maximum slightly above the chord plane extended throughout the Mach number range. (See fig. 12.) The variation of $d\\epsilon/d\\alpha$ with Mach number for tail positions of 0 and 30 percent of the semispan above and below the chord line extended was quite similar to the lift-curve-slope variation with Mach number in that a double peaking was present at about the same Mach numbers. (See fig. 15.) Between the peak values of downwash gradient which occurred at M = 0.90 and 1.02, a rather rapid variation of $d\\epsilon/d\\alpha$ with Mach number is\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:36:30.962759+00:00"}
{"citation_id": "19930085964", "source_url": "https://ntrs.nasa.gov/api/citations/19930085964/downloads/19930085964.pdf", "page_number": 18, "total_pages": 18, "image_filename": "19930085964_p18.jpg", "text": "NACA RM E9G25\n17\n\nNodes\nExtreme positions of\nblade-wall movement\n\nFigure 9. - Cross-sectional view showing\nbreathing effect in basic hollow blade.\n\nNACA\nFigure 10. - Section of basic hollow turbine blade\nstrengthened by integral fins (reference 1).\n\nNACA-Langley - 10-3-49 - 250", "timestamp": "2026-07-22T05:36:34.951167+00:00"}