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176 records · Page 10
All sourcesNASA NTRSDOE OSTI
NASA NTRS2000

Genetic Algorithm Approaches for Actuator Placement

This research investigated genetic algorithm approaches for smart actuator placement to provide aircraft maneuverability without requiring hinged flaps or other control surfaces. The effort supported goals of the Multidisciplinary Design Optimization focus efforts in NASA's Aircraft au program. This work helped to properly identify various aspects of the genetic algorithm operators and parameters that allow for placement of discrete control actuators/effectors. An improved problem definition, including better definition of the objective function and constraints, resulted from this research effort. The work conducted for this research used a geometrically simple wing model; however, an increasing number of potential actuator placement locations were incorporated to illustrate the ability of the GA to determine promising actuator placement arrangements. This effort's major result is a useful genetic algorithm-based approach to assist in the discrete actuator/effector placement problem.

Crossley, William A.↗
NASA NTRS1998

Mechanical Design of High Lift Systems for High Aspect Ratio Swept Wings

The NASA Ames Research Center is working to develop a methodology for the optimization and design of the high lift system for future subsonic airliners with the involvement of two partners. Aerodynamic analysis methods for two dimensional and three dimensional wing performance with flaps and slats deployed are being developed through a grant with the aeronautical department of the University of California Davis, and a flap and slat mechanism design procedure is being developed through a contract with PKCR, Inc., of Seattle, WA. This report documents the work that has been completed in the contract with PKCR on mechanism design. Flap mechanism designs have been completed for seven (7) different mechanisms with a total of twelve (12) different layouts all for a common single slotted flap configuration. The seven mechanisms are as follows: Simple Hinge, Upside Down/Upright Four Bar Linkage (two layouts), Upside Down Four Bar Linkages (three versions), Airbus A330/340 Link/Track Mechanism, Airbus A320 Link/Track Mechanism (two layouts), Boeing Link/Track Mechanism (two layouts), and Boeing 767 Hinged Beam Four Bar Linkage. In addition, a single layout has been made to investigate the growth potential from a single slotted flap to a vane/main double slotted flap using the Boeing Link/Track Mechanism. All layouts show Fowler motion and gap progression of the flap from stowed to a fully deployed position, and evaluations based on spanwise continuity, fairing size and number, complexity, reliability and maintainability and weight as well as Fowler motion and gap progression are presented. For slat design, the options have been limited to mechanisms for a shallow leading edge slat. Three (3) different layouts are presented for maximum slat angles of 20 deg, 15 deg and 1O deg all mechanized with a rack and pinion drive similar to that on the Boeing 757 airplane. Based on the work of Ljungstroem in Sweden, this type of slat design appears to shift the lift curve so that higher lift is achieved with the deployed slat with no increase in angle of attack. The layouts demonstrate that these slat systems can be designed with no need for slave links, and an experimental test program is outlined to experimentally validate the lift characteristics of the shallow slat.

Rudolph, Peter K. C.↗
NASA NTRS2018

A Discrete and Distributed Steady Blowing Application on a High Reynolds Number Semispan Supercritical Wing Configuration

The FAST-MAC circulation control model was modified to test an array of steady and unsteady actuators at realistic flight Reynolds numbers in the National Transonic Facility at the NASA Langley Research Center. Previous experiments in the FAST-MAC test series used a fullspan tapered slot, and that configuration is used as a baseline for performance and weight flow requirements. The goal of the latest experiment was to reduce the weight flow required to achieve comparable performance established by the baseline FAST-MAC data. Thirty-nine interchangeable actuator cartridges of various designs were mounted into the FAST-MAC model where the exiting jet was directed over a 15% chord simple hinged-flap. These two types of actuators were fabricated using rapid prototype techniques and their design performance was optimized for a transonic cruise configuration having a 0° flap deflection. The steady actuators were found to provide an off-design drag reduction of 5.5%, nearly equaling the drag reduction of the fullspan tapered slot configuration, but with a 69% weight flow reduction. This weight flow savings is similar to the sweeping jet actuators, but with better drag performance.

Jones, Gregory S.↗
NASA NTRS1990

An airfoil for general aviation applications

A new airfoil, the NLF(1)-0115, has been recently designed at the NASA Langley Research Center for use in general-aviation applications. During the development of this airfoil, special emphasis was placed on experiences and observations gleaned from other successful general-aviation airfoils. For example, the flight lift-coefficient range is the same as that of the turbulent-flow NACA 23015 airfoil. Also, although beneficial for reducing drag and having large amounts of lift, the NLF(1)-0115 avoids the use of aft loading which can lead to large stick forces if utilized on portions of the wing having ailerons. Furthermore, not using aft loading eliminates the concern that the high pitching-moment coefficient generated by such airfoils can result in large trim drags if cruise flaps are not employed. The NASA NLF(1)-0115 has a thickness of 15 percent. It is designed primarily for general-aviation aircraft with wing loadings of 718 to 958 N/sq m (15 to 20 lb/sq ft). Low profile drag as a result of laminar flow is obtained over the range from c sub l = 0.1 and R = 9x10(exp 6) (the cruise condition) to c sub l = 0.6 and R = 4 x 10(exp 6) (the climb condition). While this airfoil can be used with flaps, it is designed to achieve c(sub l, max) = 1.5 at R = 2.6 x 10(exp 6) without flaps. The zero-lift pitching moment is held at c sub m sub o = 0.055. The hinge moment for a .20c aileron is fixed at a value equal to that of the NACA 63 sub 2-215 airfoil, c sub h = 0.00216. The loss in c (sub l, max) due to leading edge roughness, rain, or insects at R = 2.6 x 10 (exp 6) is 11 percent as compared with 14 percent for the NACA 23015.

Selig, Michael S.↗
NASA NTRS1971

Full-scale wind-tunnel tests of a small unpowered jet aircraft with a T-tail

The aerodynamic characteristics of a full scale executive type jet transport aircraft with a T-tail were investigated in a 40 x 80 ft (12.2 by 24.4 meter) wind tunnel (subsonic). Static, longitudinal, and lateral stability, and control characteristics were determined at angles of attack from -2 deg to +42 deg. The aircraft wing had 13 deg of sweep and an aspect ratio of 5.02. The aircraft was tested power off with various wing leading- and trailing-edge high lift devices. The basic configuration was tested with and without such components as engine nacelles, wing tip tanks, and empannage. Hinge-moment data were obtained and downwash angles in the horizontal-tail plane location were calculated. The data were obtained at Reynolds numbers of 4.1 million and 8.7 million based on mean aerodynamic chord. The model had static longitudinal stability through initial stall. Severe tail buffet occurred near the angle of attack for maximum lift. Above initial stall the aircraft had pronounced pitch-up, characteristic of T-tail configurations. A stable trim point was possible at angles of attack between 30 deg and 40 deg (depending on c.g. location and flap setting). Hinge-moment data showed no regions with adverse effects on stick force. Comparisons of wind-tunnel data and flight-test are presented.

Soderman, P. T.↗
NASA NTRS1986

Basic studies on delta wing flow modifications by means of apex fences

The effectiveness of apex fences on a 60-deg delta wing at low speeds was experimentally investigated. Resembling highly swept spoilers in appearance, the fences are designed to fold out of the wing apex region upper surface near the leading edges, where they generate a powerful vortex pair. The intense suction of the fence vortices augments lift in the apex region, the resulting positive pitching moment being utilized to trim trailing edge flaps for lift augmentation during approach and landing at relatively low angles of attack. The fences reduce the apex lift at high angles of attack, leading to a desirable nose-down moment. The above projected functions of the apex fence device were validated and quantified through low speed tunnel tests, comprising upper surface pressure surveys on a semispan model and balance measurements on a geometrically similar fully span wing/body configuration. Fence parameters such as area, shape, hinge position and deflection angle were investigated. Typical results are presented indicating the apex fence potential in controlling the longitudinal characteristics of a tail-less delta.

Hoffler, K. D.↗
NASA NTRS1961

An Investigation of Wing and Aileron Loads Due to Deflected Inboard and Outboard Ailerons on a 4-Percent-Thick 30 deg Sweptback Wing at Transonic Speeds

An investigation has been conducted in the Langley 16-foot transonic tunnel to determine the changes in wing loading characteristics due to deflections of a plain faired flap-type inboard aileron, a plain faired flap-type outboard aileron, and a slab-sided thickened trailing edge outboard aileron. The test wing was 4 percent thick and had 30 sweep of the quarter chord, an aspect ratio of 3.0, a taper ratio of 0.2, and NACA 65A004 airfoil sections. The loading characteristics of the deflected ailerons were also investigated. The model was a sting-mounted wing-body combination, and pressure measurements over one wing panel (exposed area) and the ailerons were obtained for angles of attack from 0 to 20 at deflections up to +/- 15 deg for Mach numbers between 0.80 and 1.03. The test Reynolds number based on the wing mean aerodynamic chord was about 7.4 x 10(exp 6). The results of the investigation indicated that positive deflection of the plain faired flap-type inboard aileron caused significant added loading over the wing sections outboard of the aileron at all Mach numbers for model angles of attack from 0 deg or 4 deg up to 12 deg. Positive deflection of the two outboard ailerons (plain faired and slab sided with thickened trailing edge) caused significant added loading over the wing sections inboard of the ailerons for different model angle-of-attack ranges at the several test Mach numbers. The loading shapes over the ailerons were irregular and would be difficult to predict from theoretical considerations in the transonic speed range. The longitudinal and lateral center-of-pressure locations for the ailerons varied only slightly with increasing angle of attack and/or Mach number. Generally, the negative slopes of the variations of aileron hinge-moment coefficient with aileron deflection for all three ailerons varied similarly with Mach number at the test angles of attack.

Whitcomb, Charles F.↗
NASA NTRS1999

Design of Supersonic Transport Flap Systems for Thrust Recovery at Subsonic Speeds

A study of the subsonic aerodynamics of hinged flap systems for supersonic cruise commercial aircraft has been conducted using linear attached-flow theory that has been modified to include an estimate of attainable leading edge thrust and an approximate representation of vortex forces. Comparisons of theoretical predictions with experimental results show that the theory gives a reasonably good and generally conservative estimate of the performance of an efficient flap system and provides a good estimate of the leading and trailing-edge deflection angles necessary for optimum performance. A substantial reduction in the area of the inboard region of the leading edge flap has only a minor effect on the performance and the optimum deflection angles. Changes in the size of the outboard leading-edge flap show that performance is greatest when this flap has a chord equal to approximately 30 percent of the wing chord. A study was also made of the performance of various combinations of individual leading and trailing-edge flaps, and the results show that aerodynamic efficiencies as high as 85 percent of full suction are predicted.

Mann, Michael J.↗
NASA NTRS2013

Enhancements to the FAST-MAC Circulation Control Model and Recent High-Reynolds Number Testing in the National Transonic Facility

A second wind tunnel test of the FAST-MAC circulation control model was recently completed in the National Transonic Facility at the NASA Langley Research Center. The model was equipped with four onboard flow control valves allowing independent control of the circulation control plenums, which were directed over a 15% chord simple-hinged flap. The model was configured for low-speed high-lift testing with flap deflections of 30 and 60 degrees, along with the transonic cruise configuration with zero degree flap deflection. Testing was again conducted over a wide range of Mach numbers up to 0.88, and Reynolds numbers up to 30 million based on the mean chord. The first wind tunnel test had poor transonic force and moment data repeatability at mild cryogenic conditions due to inadequate thermal conditioning of the balance. The second test demonstrated that an improvement to the balance heating system significantly improved the transonic data repeatability, but also indicated further improvements are still needed. The low-speed highlift performance of the model was improved by testing various blowing slot heights, and the circulation control was again demonstrated to be effective in re-attaching the flow over the wing at off-design transonic conditions. A new tailored spanwise blowing technique was also demonstrated to be effective at transonic conditions with the benefit of reduced mass flow requirements.

Milholen, William E., II↗
NASA NTRS2019

Comparative Study of Active Flow Control Strategies for Lift Enhancement of a Simplified High-Lift Configuration

Numerical simulations have been performed for a simplified high-lift (SHL) version of the Common Research Model (CRM) configuration, where the Fowler flaps of the conventional high-lift (CRM-HL) configuration are replaced by a set of simple hinged flaps. These hinged flaps are equipped with integrated modular active flow control (AFC) cartridges on the suction surface, and the resulting geometry is known as the CRM-SHL-AFC configuration. The main objective is to make use of AFC devices on the CRM-SHL-AFC configuration to recover the aerodynamic performance (lift) of the CRM-HL configuration. In the current paper, a Lattice Boltzmann method-based computational fluid dynamics (CFD) code, known as PowerFLOWQ® is used to simulate the entire flow field associated with the CRM-SHL-AFC configuration equipped with several different types of AFC devices. The transonic version of the PowerFLOWQ® code that has been validated for high speed flows is used to accurately simulate the flow field generated by the high-momentum actuators required to mitigate reversed flow regions on the suction surfaces of the main wing and the flap. The numerical solutions predict the expected trends in aerodynamic forces as the actuation levels are increased. More efficient AFC systems and actuator arrangements emerged based on the parametric studies performed prior to a Fall 2018 wind tunnel test. Preliminary comparisons of the numerical solutions for lift and surface pressures are presented here with the experimental data, demonstrating the usefulness of CFD for predicting the flow field and lift characteristics of AFC-enabled high-lift configurations.

Vatsa, Veer N.↗
NASA NTRS2007

Flight Testing an Iced Business Jet for Flight Simulation Model Validation

A flight test of a business jet aircraft with various ice accretions was performed to obtain data to validate flight simulation models developed through wind tunnel tests. Three types of ice accretions were tested: pre-activation roughness, runback shapes that form downstream of the thermal wing ice protection system, and a wing ice protection system failure shape. The high fidelity flight simulation models of this business jet aircraft were validated using a software tool called "Overdrive." Through comparisons of flight-extracted aerodynamic forces and moments to simulation-predicted forces and moments, the simulation models were successfully validated. Only minor adjustments in the simulation database were required to obtain adequate match, signifying the process used to develop the simulation models was successful. The simulation models were implemented in the NASA Ice Contamination Effects Flight Training Device (ICEFTD) to enable company pilots to evaluate flight characteristics of the simulation models. By and large, the pilots confirmed good similarities in the flight characteristics when compared to the real airplane. However, pilots noted pitch up tendencies at stall with the flaps extended that were not representative of the airplane and identified some differences in pilot forces. The elevator hinge moment model and implementation of the control forces on the ICEFTD were identified as a driver in the pitch ups and control force issues, and will be an area for future work.

Ratvasky, Thomas P.↗
NASA NTRSDate not supplied

CFD and Experimental Data Comparisons for Conventional and AFC-Enabled CRM High-Lift Configurations

Numerical simulations have been performed for a simplified high-lift (SHL) version of the Common Research Model (CRM) configuration, where the Fowler flaps of a representative conventional high-lift (CRM-HL) configuration are replaced by a set of simple hinged flaps. These hinged flaps are equipped with integrated modular active flow control (AFC) cartridges on the suction surface of the flap shoulder, and the resulting geometry is known as the CRM-SHL-AFC configuration. The main objective is to make use of AFC devices on the CRM-SHL-Configuration to produce the aerodynamic performance (lift) comparable to that of the CRM-HL configuration over a large angle of attack range encompassing maximum lift conditions. For comparison purposes, computations are also per-formed for the CRM-HL configuration. In the current paper, PowerFLOWR©, a CFD code based on the Lattice Boltzmann method (LBM), is used to simulate the entire flow field associated with the CRM-SHL-AFC configuration equipped with a combination of two different types of AFC actuators. The transonic version of the PowerFLOWR© code that has been validated for high-speed flows is used to simulate the flow field generated by the high-momentum actuators required to mitigate reversed flow regions on the suction surfaces of the main wing and the flap. This study is focused on the AFC systems and actuator arrangements that had emerged based on the parametric studies conducted previously at the nominal landing condition. Comparisons of the numerical solutions for lift and surface pressures are presented here with the experimental data, demonstrating the usefulness of CFD for predicting the flow field and lift characteristics of AFC-enabled high-lift configurations over a broad angle of attack range. The numerical solutions predict the expected trends in aerodynamic forces with angle of attack variation.

Veer N Vatsa↗
NASA NTRS2019

Surface Flow Visualization of the High-Lift Common Research Model

A 10% scale version of the High-Lift Common Research Model (CRM-HL) was tested in the NASA Langley 14- by 22-Foot Subsonic Tunnel (14x22) in support of the NASA Advanced Air Transport Technology (AATT) Project. The CRM-HL experiment included various configurations such as conventional and simple-hinged flaps, with and without engine nacelle/pylon, with and without nacelle chine, different Active Flow Control (AFC) methods (sweeping jets, alternating pulsed jets, and preconditioned boundary layer blowing), and their various parameters. This particular study is focused on the surface flow visualization of the conventional CRM-HL model at landing configuration. The conventional CRM-HL model with the single-slotted Fowler flap system serves as a baseline for the AFC-enabled simplified high-lift configuration as well as a high-lift technology development platform due to its publicly open geometry. Surface flow visualizations were performed using fluorescent minitufts, which were found to be nonintrusive to the aerodynamic performance. Tuft flow visualizations are supplemented with the relevant pressure and force measurements in order to understand the flow characteristics developed on the conventional CRM- HL model. In addition, three dimensional, unsteady, compressible Computational Fluid Dynamic (CFD) simulations were performed for selective cases. The surface streamlines and transverse velocity fluctuations obtained by the CFD simulations are qualitatively compared to the tuft direction and tuft unsteadiness, respectively. Force measurements of the CRM-HL model show performance degradation at higher angles of attack. Surface flow visualizations revealed the performance loss due to the nacelle/pylon wake that grows with angle of attack and eventually promotes flow separation over the inboard wing. This performance loss was successfully recovered by placing a chine on the engine nacelle.

Koklu, Mehti↗
NASA NTRS2014

A Modular Approach to Model Oscillating Control Surfaces Using Navier Stokes Equations

The use of active controls for rotorcraft is becoming more important for modern aerospace configurations. Efforts to reduce the vibrations of helicopter blades with use of active-controls are in progress. Modeling oscillating control surfaces using the linear aerodynamics theory is well established. However, higher-fidelity methods are needed to account for nonlinear effects, such as those that occur in transonic flow. The aeroelastic responses of a wing with an oscillating control surface, computed using the transonic small perturbation (TSP) theory, have been shown to cause important transonic flow effects such as a reversal of control surface effectiveness that occurs as the shock wave crosses the hinge line. In order to account for flow complexities such as blade-vortex interactions of rotor blades higher-fidelity methods based on the Navier-Stokes equations are used. Reference 6 presents a procedure that uses the Navier-Stokes equations with moving-sheared grids and demonstrates up to 8 degrees of control-surface amplitude, using a single grid. Later, this procedure was extended to accommodate larger amplitudes, based on sliding grid zones. The sheared grid method implemented in EulerlNavier-Stokes-based aeroelastic code ENS AERO was successfully applied to active control design by industry. Recently there are several papers that present results for oscillating control surface using Reynolds Averaged Navier-Stokes (RANS) equations. References 9 and 10 report 2-D cases by filling gaps with overset grids. Reference 9 compares integrated forces with the experiment at low oscillating frequencies whereas Ref. 10 reports parametric studies but with no validation. Reference II reports results for a 3D case by modeling the gap region with a deformed grid and compares force results with the experiment only at the mid-span of flap. In Ref. II grid is deformed to match the control surface deflections at the section where the measurements are made. However, there is no indication in Ref. II that the gaps are explicitly modeled as in Ref. 6. Computations using overset grids are reported in Ref. 12 for a case by adding moving control surface to an existing blade but with no validation either with an experiment or another computation.

Navier↗