Results for “Lift to Drag Ratio” — Page 18

NASA NTRS2007

Mars Science Laboratory: Entry, Descent, and Landing System Performance

In 2010, the Mars Science Laboratory (MSL) mission will pioneer the next generation of robotic Entry, Descent, and Landing (EDL) systems, by delivering the largest and most capable rover to date to the surface of Mars. To do so, MSL will fly a guided lifting entry at a lift-to-drag ratio in excess of that ever flown at Mars, deploy the largest parachute ever at Mars, and perform a novel Sky Crane maneuver. Through improved altitude capability, increased latitude coverage, and more accurate payload delivery, MSL is allowing the science community to consider the exploration of previously inaccessible regions of the planet. The MSL EDL system is a new EDL architecture based on Viking heritage technologies and designed to meet the challenges of landing increasing massive payloads on Mars. In accordance with level-1 requirements, the MSL EDL system is being designed to land an 850 kg rover to altitudes as high as 1 km above the Mars Orbiter Laser Altimeter defined areoid within 10 km of the desired landing site. Accordingly, MSL will enter the largest entry mass, fly the largest 70 degree sphere-cone aeroshell, generate the largest hypersonic lift-to-drag ratio, and deploy the largest Disk-Gap-Band supersonic parachute of any previous mission to Mars. Major EDL events include a hypersonic guided entry, supersonic parachute deploy and inflation, subsonic heatshield jettison, terminal descent sensor acquisition, powered descent initiation, sky crane terminal descent, rover touchdown detection, and descent stage flyaway. Key performance metrics, derived from level-1 requirements and tracked by the EDL design team to indicate performance capability and timeline margins, include altitude and range at parachute deploy, time on radar, and propellant use. The MSL EDL system, which will continue to develop over the next three years, will enable a notable extension in the advancement of Mars surface science by delivering more science capability than ever before to the surface of Mars. This paper describes the current MSL EDL system performance as predicted by end-to-end EDL simulations, highlights the sensitivity of this baseline performance to several key environmental assumptions, and discusses some of the challenges faced in delivering such an unprecedented rover payload to the surface of Mars.

Way, David W.↗

NASA NTRS2006

Mars Science Laboratory: Entry, Descent, and Landing System Performance

In 2010, the Mars Science Laboratory (MSL) mission will pioneer the next generation of robotic Entry, Descent, and Landing (EDL) systems, by delivering the largest and most capable rover to date to the surface of Mars. To do so, MSL will fly a guided lifting entry at a lift-to-drag ratio in excess of that ever flown at Mars, deploy the largest parachute ever at Mars, and perform a novel Sky Crane maneuver. Through improved altitude capability, increased latitude coverage, and more accurate payload delivery, MSL is allowing the science community to consider the exploration of previously inaccessible regions of the planet. The MSL EDL system is a new EDL architecture based on Viking heritage technologies and designed to meet the challenges of landing increasing massive payloads on Mars. In accordance with level-1 requirements, the MSL EDL system is being designed to land an 850 kg rover to altitudes as high as 1 km above the Mars Orbiter Laser Altimeter defined areoid within 10 km of the desired landing site. Accordingly, MSL will enter the largest entry mass, fly the largest 70 degree sphere-cone aeroshell, generate the largest hypersonic lift-to-drag ratio, and deploy the largest Disk-Gap-Band supersonic parachute of any previous mission to Mars. Major EDL events include a hypersonic guided entry, supersonic parachute deploy and inflation, subsonic heatshield jettison, terminal descent sensor acquisition, powered descent initiation, sky crane terminal descent, rover touchdown detection, and descent stage flyaway. Key performance metrics, derived from level-1 requirements and tracked by the EDL design team to indicate performance capability and timeline margins, include altitude and range at parachute deploy, time on radar, and propellant use. The MSL EDL system, which will continue to develop over the next three years, will enable a notable extension in the advancement of Mars surface science by delivering more science capability than ever before to the surface of Mars. This paper describes the current MSL EDL system performance as predicted by end-to-end EDL simulations, highlights the sensitivity of this baseline performance to several key environmental assumptions, and discusses some of the challenges faced in delivering such an unprecedented rover payload to the surface of Mars.

Way, David W.↗

NASA NTRS1959

An Experimental Investigation of the Effect of a Canard Control on the Lift, Drag, and Pitching Moment of an Aspect-Ratio 2.0 Triangular Wing Incorporating a Form of Conical Camber

The results of an experimental investigation to determine the effect of a canard control on the lift, drag, and pitching-moment characteristics of an aspect-ratio-2.0 triangular wing incorporating a form of conical camber are presented. The canard had a triangular plan form of aspect ratio 2.0 and was mounted in the extended chord plane of the wing. The ratio of the area of the exposed canard panels to the total wing area was 6.9 percent, and the ratio of the total areas was 12.9 percent. Data were obtained at Mach numbers from 0.70 to 2.22 through an angle-of-attack range from -6 deg to +18 deg with the canard on, and with the canard off. To provide a basis for comparison, the canard was also tested with a symmetrical wing having the same plan form, aspect ratio, and thickness distribution as the cambered wing. The results of the investigation showed that at the high subsonic speeds the gain in maximum lift-drag ratio achieved by camber was considerably reduced by the addition of a canard. At the supersonic speeds, the addition of the canard did not change the effect of camber on the maximum lift-drag ratios.

Menees, Gene P.↗

NASA NTRS1999

Demonstration of Automatically-Generated Adjoint Code for Use in Aerodynamic Shape Optimization

Gradient-based optimization requires accurate derivatives of the objective function and constraints. These gradients may have previously been obtained by manual differentiation of analysis codes, symbolic manipulators, finite-difference approximations, or existing automatic differentiation (AD) tools such as ADIFOR (Automatic Differentiation in FORTRAN). Each of these methods has certain deficiencies, particularly when applied to complex, coupled analyses with many design variables. Recently, a new AD tool called ADJIFOR (Automatic Adjoint Generation in FORTRAN), based upon ADIFOR, was developed and demonstrated. Whereas ADIFOR implements forward-mode (direct) differentiation throughout an analysis program to obtain exact derivatives via the chain rule of calculus, ADJIFOR implements the reverse-mode counterpart of the chain rule to obtain exact adjoint form derivatives from FORTRAN code. Automatically-generated adjoint versions of the widely-used CFL3D computational fluid dynamics (CFD) code and an algebraic wing grid generation code were obtained with just a few hours processing time using the ADJIFOR tool. The codes were verified for accuracy and were shown to compute the exact gradient of the wing lift-to-drag ratio, with respect to any number of shape parameters, in about the time required for 7 to 20 function evaluations. The codes have now been executed on various computers with typical memory and disk space for problems with up to 129 x 65 x 33 grid points, and for hundreds to thousands of independent variables. These adjoint codes are now used in a gradient-based aerodynamic shape optimization problem for a swept, tapered wing. For each design iteration, the optimization package constructs an approximate, linear optimization problem, based upon the current objective function, constraints, and gradient values. The optimizer subroutines are called within a design loop employing the approximate linear problem until an optimum shape is found, the design loop limit is reached, or no further design improvement is possible due to active design variable bounds and/or constraints. The resulting shape parameters are then used by the grid generation code to define a new wing surface and computational grid. The lift-to-drag ratio and its gradient are computed for the new design by the automatically-generated adjoint codes. Several optimization iterations may be required to find an optimum wing shape. Results from two sample cases will be discussed. The reader should note that this work primarily represents a demonstration of use of automatically- generated adjoint code within an aerodynamic shape optimization. As such, little significance is placed upon the actual optimization results, relative to the method for obtaining the results.

Green, Lawrence↗

NASA NTRS2022

Comparison of Computational Predictions of the Mach 0.80 Transonic Truss-Braced Wing Configuration with Experimental Data

The NASA Langley and Ames Research Centers have teamed together for comparisons of computational predictions of the Boeing Mach 0.80Transonic Truss-Braced Wing (TTBW) configuration with a high-speed experimental dataset. The Mach 0.80 TTBW vehicle is a high wing, high aspect ratio configuration, designed for a high lift-to-drag ratio. System studies have predicted significant fuel burn and emissions benefits with the TTBW technology moving toward meeting NASA Subsonic Transport Systems-Level-Metrics. A 4.5% scale Mach 0.80 design TTBW model was recently tested at the NASA Ames Research Center 11-by 11-Foot Transonic Wind Tunnel(11-Ft TWT),providing a valuable dataset to validate computational tools and investigate best practices as risk reduction efforts continue for the development of the advanced TTBW vehicle. The NASA Computational Fluid Dynamics (CFD) team has computed free-air flow solutions on the Mach 0.80 design flight configuration and two wind tunnel configuration variants using the USM3D and LAVA flow solvers. Accurate modeling of the configuration tested in the wind tunnel environment is critical to validating the CFD tools, thus the team has included the internal cavity region and sting in their modeling of the configuration, similar to that tested in the 11-Ft TWT. Overall, the CFD simulations compared well and show similar trends as the corrected experimental data for lift and drag polars. The CFD predicted lift curve is shifted in angle of attack from what was observed in the experiment. The shift in lift also was seen in the pitching moment comparison plots. CFD solutions were computed at constant CL test point values and showed overall very good agreement when comparing constant spanwise cuts of pressure coefficient data on the wing and strut with experimental data. CFD cavity corrections were also investigated using the 11-Ft TWT cavity correction method, similar to that used to correct the wind tunnel data. Results showed some improvement in pitching moment coefficient predictions, and an increase in drag, shifting the data to the right in drag polars, further from the experimental data at lower lift conditions, good agreement near the design CL, and slight improvement at the higher lift conditions.

Transonic Truss-Braced Wing↗

NASA NTRS1991

In-flight simulation at the NASA Dryden Flight Research Facility

An account of the in-flight simulation at the Dryden Flight Research Facility is presented and various aircraft and tests performed are discussed. In-flight simulation has been utilized for a variety of flying quality investigations including low-lift-to-drag ratio approach characteristics for vehicles like the X-15, the lifting bodies, and the Space Shuttle. Consideration is given to the effects of time delays on controllability of aircraft with digital flight-control systems, flight control systems for such aircraft as the X-29, and the causes and cures of pilot-induced oscillation in diverse aircraft.

Shafer, Mary F.↗

NASA NTRS1992

In-flight simulation studies at the NASA Dryden Flight Research Facility

Since the late 1950's, the National Aeronautics and Space Administration's Dryden Flight Research Facility has found in-flight simulation to be an invaluable tool. In-flight simulation has been used to address a wide variety of flying qualities questions, including low-lift-to-drag ratio approach characteristics for vehicles like the X-15, the lifting bodies, and the Space Shuttle; the effects of time delays on controllability of aircraft with digital flight-control systems, the causes and cures of pilot-induced oscillation in a variety of aircraft, and flight-control systems for such diverse aircraft as the X-15 and the X-29. In-flight simulation has also been used to anticipate problems and to avoid them and to solve problems once they appear. Presented here is an account of the in-flight simulation at the Dryden Flight Research Facility and some discussion. An extensive bibliography is included.

Shafer, Mary F.↗

NASA NTRS1994

In-flight simulation studies at the NASA Dryden Flight Research Facility

Since the late 1950's the National Aeronautics and Space Administration's Dryden Flight Research Facility has found in-flight simulation to be an invaluable tool. In-flight simulation has been used to address a wide variety of flying qualities questions, including low lift-to-drag ratio approach characteristics for vehicles like the X-15, the lifting bodies, and the space shuttle; the effects of time delays on controllability of aircraft with digital flight control systems; the causes and cures of pilot-induced oscillation in a variety of aircraft; and flight control systems for such diverse aircraft as the X-15 and the X-29. In-flight simulation has also been used to anticipate problems, avoid them, and solve problems once they appear. This paper presents an account of the in-flight simulation at the Dryden Flight Research Facility and some discussion. An extensive bibliography is included.

Shafer, Mary F.↗

NASA NTRS1986

How different a modern SST would be

The characteristics of a proposed SST are described. The proposed aircraft is to have two engines and an arrow-wing design, a passenger capacity of 250, and attain speeds of Mach 2.7. The low fineness ratio, low-aspect ratio wing planform, different engine nacelles location, and improved lift-to-drag ratio of the aircraft contribute to attaining an economical supersonic cruise. The proposed structural design and materials for the aircraft are examined; the material and design are to be applicable to high Mach and high temperature. Changes in the SST propulsion system, the nozzle designs, and landing and takeoff procedures to improve the operation of the aircraft are discussed.

Driver, C.↗

NASA NTRS2020

Optimization of Low Reynolds Number Airfoils for Martian Rotor Applications Using an Evolutionary Algorithm

The Mars Helicopter (MH) will be flying on the NASA Mars 2020 rover mission scheduled to launch in July of 2020. Research is being performed at the Jet Propulsion Laboratory (JPL) and NASA Ames Research Center to extend the current capabilities and develop the Mars Science Helicopter (MSH) as the next possible step for Martian rotorcraft. The low atmospheric density and the relatively small-scale rotors result in very low chord-based Reynolds number flows over the rotor airfoils. The low Reynolds number regime results in rapid performance degradation for conventional airfoils due to laminar separation without reattachment. Unconventional airfoil shapes with sharp leading edges are explored and optimized for aerodynamic performance at representative Reynolds-Mach combinations for a concept rotor. Sharp leading edges initiate immediate flow separation, and the occurrence of large-scale vortex shedding is found to contribute to the relative performance increase of the optimized airfoils, compared to conventional airfoil shapes. The oscillations are shown to occur independent from laminar-turbulent transition and therefore result in sustainable performance at lower Reynolds numbers. Comparisons are presented to conventional airfoil shapes and peak lift-to-drag ratio increases between 17% and 41% are observed for similar section lift.

Koning, Witold J.↗

NASA NTRSDate not supplied

Comparative Performance of Entry Simulations using Guided Human-Scale Landers at Mars

There has been ongoing interest in human exploration of Mars since the 1970s. Most studies to date have investigated moderate lift-to-drag ratio vehicles to investigate this problem. This paper considers the development and cross-simulation comparison of a slender-bodied, moderate lift-to-drag human-scale lander vehicle for Mars. The vehicle employs the Fully Numerical Predictor-corrector Entry Guidance to target a powered descent state, subsequently leveraging Fractional Polynomial Powered Descent Guidance laws to reach the desired touchdown target. Differences in simulation environment as well as nominal trajectory results are shared.

Evan Roelke↗

NASA NTRSDate not supplied

Comparative Performance of Entry Simulations using Guided Human-Scale Landers at Mars

There has been ongoing interest in human exploration of Mars since the 1970s. Most studies to date have investigated moderate lift-to-drag ratio vehicles to investigate this problem. This paper considers the development and cross-simulation comparison of a slender-bodied, moderate lift-to-drag human-scale lander vehicle for Mars. The vehicle employs the Fully Numerical Predictor-corrector Entry Guidance to target a powered descent state, subsequently leveraging Fractional Polynomial Powered Descent Guidance laws to reach the desired touchdown target. Differences in simulation environment as well as nominal trajectory results are shared.

Evan Roelke↗

NASA NTRS1992

The Triton: Design concepts and methods

During the design of the C & P Aerospace Triton, a few problems were encountered that necessitated changes in the configuration. After the initial concept phase, the aspect ratio was increased from 7 to 7.6 to produce a greater lift to drag ratio (L/D = 13) which satisfied the horsepower requirements (118 hp using the Lycoming O-235 engine). The initial concept had a wing planform area of 134 sq. ft. Detailed wing sizing analysis enlarged the planform area to 150 sq. ft., without changing its layout or location. The most significant changes, however, were made just prior to inboard profile design. The fuselage external diameter was reduced from 54 to 50 inches to reduce drag to meet the desired cruise speed of 120 knots. Also, the nose was extended 6 inches to accommodate landing gear placement. Without the extension, the nosewheel received an unacceptable percentage (25 percent) of the landing weight. The final change in the configuration was made in accordance with the stability and control analysis. In order to reduce the static margin from 20 to 13 percent, the horizontal tail area was reduced from 32.02 to 25.0 sq. ft. The Triton meets all the specifications set forth in the design criteria. If time permitted another iteration of the calculations, two significant changes would be made. The vertical stabilizer area would be reduced to decrease the aircraft lateral stability slope since the current value was too high in relation to the directional stability slope. Also, the aileron size would be decreased to reduce the roll rate below the current 106 deg/second. Doing so would allow greater flap area (increasing CL(sub max)) and thus reduce the overall wing area. C & P would also recalculate the horsepower and drag values to further validate the 120 knot cruising speed.

Meholic, Greg↗

NASA NTRS1961

An Investigation of the Influence of Body Size and Indentation Asymmetry of the Effectiveness of Body Indentation in Combination with a Cambered Wing

An investigation has been made of a 450 sweptback cambered wing in combination with an unindented body and a body symmetrically indented with respect to its axes designed for a Mach number of 1.2. The ratio of body frontal area to wing planform area was 0.08 for these wing-body combinations. In order to determine the influence of body size on the effectiveness of indentation, the test data have been compared with previously obtained data for similar configurations having a ratio of body frontal area to wing planform area of 0.04. Also, in order to investigate the relative effectiveness of indentation asymmetry, a specially indented body designed to account for the wing camber and also designed for a Mach number of 1.2 has been included in these tests. The investigation was conducted in the Langley 8-Foot Tunnels Branch at Mach numbers from 0.80 to 1.43 and a Reynolds number of approximately 1.85 x 10(exp 6), based on a mean aerodynamic chord length of 5.955 inches. The data indicate that the configurations with larger ratio of body frontal area to wing planform area had smaller reductions in zero-lift wave drag associated with body indentation than the configurations with smaller ratio of body frontal area to wing planform area. The 0.08-area-ratio configurations also had correspondingly smaller increases in the values of maximum lift-drag ratio than the 0.04-area-ratio configurations. The consideration of wing camber in the body indentation design resulted in a 35.5-percent reduction in zero-lift wave drag, compared with a 21.5-percent reduction associated with the symmetrical indentation, but had a negligible effect on the values of maximum lift-drag ratio.

Patterson, James C., Jr.↗

NASA NTRS1980

Generic aerocapture atmospheric entry study, volume 1

An atmospheric entry study to fine a generic aerocapture vehicle capable of missions to Mars, Saturn, and Uranus is reported. A single external geometry was developed through atmospheric entry simulations. Aerocapture is a system design concept which uses an aerodynamically controlled atmospheric entry to provide the necessary velocity depletion to capture payloads into planetary orbit. Design concepts are presented which provide the control accuracy required while giving thermal protection for the mission payload. The system design concepts consist of the following elements: (1) an extendable biconic aerodynamic configuration with lift to drag ratio between 1.0 and 2.0; (2) roll control system concepts to control aerodynamic lift and disturbance torques; (3) aeroshell design concepts capable of meeting dynamic pressure loads during aerocapture; and (4) entry thermal protection system design concepts to meet thermodynamic loads during aerocapture.

Source record↗

NASA NTRS1991

Experiences with optimizing airfoil shapes for maximum lift over drag

The goal was to find airfoil shapes which maximize the ratio of lift over drag for given flow conditions. For a fixed Mach number, Reynolds number, and angle of attack, the lift and drag depend only on the airfoil shape. This then becomes a problem in optimization: find the shape which leads to a maximum value of lift over drag. The optimization was carried out using a self contained computer code for finding the minimum of a function subject to constraints. To find the lift and drag for each airfoil shape, a flow solution has to be obtained. This was done using a two dimensional Navier-Stokes code.

Doria, Michael L.↗

NASA NTRS1986

An experimental study of a three lifting surface configuration

The aerodynamic ramifications of utilizing three lifting surfaces as opposed to the conventional or canard lifting configurations have been studied on a theoretical basis by previous researchers. This paper presents an experimental investigation of various configuration modifications for an unyawed typical business jet at a Reynolds number of 1.3 million. The three surface has better lift and high-lift drag characteristics than either the canard or tail-aft configurations, but the cruise drag is more. The induced drag at cruise is highest for the canard and lowest for the tail-aft configuration. The pitching moment characteristics are somewhat between those of the canard and tail-aft configurations. A decrease in gap adversely affects the pitching moment characteristics. A smaller stagger leads to better aerodynamic and stability characteristics. A decrease in span of the forward wing gives better cruise drag and longitudinal stability characteristics, but has adverse effects on high-lift drag. A variation in the incidence angles of either or both the forward and aft wings changes the zero-lift moments of the configuration, while marginally affecting overall lift and drag. At cruise, the lift to drag ratio is highest for the conventional and lowest for the three surface. For high lift conditions, the order is reversed.

Ostowari, C.↗

NASA NTRS1994

A Two Element Laminar Flow Airfoil Optimized for Cruise

Numerical and experimental results are presented for a new two-element, fixed-geometry natural laminar flow airfoil optimized for cruise Reynolds numbers on the order of three million. The airfoil design consists of a primary element and an independent secondary element with a primary to secondary chord ratio of three to one. The airfoil was designed to improve the cruise lift-to-drag ratio while maintaining an appropriate landing capability when compared to conventional airfoils. The airfoil was numerically developed utilizing the NASA Langley Multi-Component Airfoil Analysis computer code running on a personal computer. Numerical results show a nearly 11.75 percent decrease in overall wing drag with no increase in stall speed at sailplane cruise conditions when compared to a wing based on an efficient single element airfoil. Section surface pressure, wake survey, transition location, and flow visualization results were obtained in the Texas A&M University Low Speed Wind Tunnel. Comparisons between the numerical and experimental data, the effects of the relative position and angle of the two elements, and Reynolds number variations from 8 x 10(exp 5) to 3 x 10(exp 6) for the optimum geometry case are presented.

Steen, Gregory Glen↗