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All sourcesNASA NTRSDOE OSTI
NASA NTRS1949

Longitudinal Stability and Control Characteristics of a Semispan Wind-Tunnel Model of the XF7U-1 Airplane and a Comparison with Complete-Model Wind-Tunnel Tests and Semispan-Model Wing-Flow Tests

An investigation was conducted on an 0.08-scale semispan model of the Chance Vought XF7U-1 airplane in the Langley high-speed 7- by 10-foot tunnel in the Mach number range from 0.40 to 0.97. The results are compared with those obtained with an 0.08-scale sting-mounted complete model tested in the same tunnel and with an 0.026-scale semispan model tested by the wing-flow method. The lift-curve slopes obtained for the 0.08-scale semispan model and the 0.026-scale wing-flow model were in good agreement but both were generally lower than the values obtained for the sting model. The results of an unpublished investigation have shown that tunnel-wall boundary-layer and strut-leakage effects can came the difference noted between the lift-curve slopes of the sting and the semispan data. Fair agreement was obtained among the data of the three models as regard the variation of pitching-moment coefficients with lift coefficient. The agreement between the complete and the semispan models was more favorable with the vertical fine on, because the wall-boundary-layer and strut leakage effects were less severe. In the Mach number range between 0.94 and 0.97, ailavator-control reversal was indicated in the wing-flow data near zero lift; Whereas, these same trends were indicated in the larger scale semispan data at somewhat higher lift coefficients.

Goodson, Kenneth W.↗
NASA NTRS1991

Dynamic response of a hammerhead launch vehicle wind-tunnel model

A wind-tunnel test of a 1/10th-scale Atlas-Centaur I large payload fairing launch vehicle model was conducted in the NASA Langley Transonic Dynamics Tunnel. The wind tunnel model was an aeroelastically-scaled version of the flight vehicle and was capable of simulating either of the first two bending vibration modes of the full-scale vehicle by a partial mode technique. The primary purpose of the test was to gather data concerning buffet response which could be used to clear the vehicle for flight. Additionally, angle-of-attack studies were conducted and several payload fairing configurations were tested to assess the buffet response and dynamic stability of off-design flight conditions and geometric parameters. No dynamic instabilities were found for any of the configurations tested. The buffet response data for the nominal flight configuration indicate that the unsteady buffet loads represent 5 to 10 percent of the total design load; therefore, the buffet loads are not a large factor affecting the overall vehicle design. Payload fairing length-to-diameter ratio variations were found to have small effects on the buffet response of the model, except in the case of the smallest length-to-diameter models for the second bending mode simulation. The effects of angle of attack on buffet response were found to be small. The model was more sensitive to Mach number changes than to angle of attack. The buffet response results from this wind tunnel test were influenced by the tunnel facility vibration levels. An attempt was made to experimentally reduce the effect of the facility mechanical vibration for the nominal flight configuration by testing with vertical rods used to stiffen the sting support. The first flight of the Atlas-Centaur I vehicle successfully occurred on July 25, 1990, and a comparison of flight measurements with wind tunnel data is presented. The flight data was found to be well within the 3 sigma level of the wind tunnel data.

Cole, Stanley R.↗
NASA NTRS2015

Static Aeroelastic Modeling of a Sub-Scale Wind Tunnel Model with Novel Flap Concept

This paper examines the static aeroelastic modeling of a flexible wind tunnel model equipped with a novel control effector known as the Variable Camber Continuous Trailing Edge Flap (VCCTEF) system. The wind tunnel model is an approximately 10 percent sub-scale version of a wing designed to be equipped on a commerical transport aircraft such as the full-scale NASA Generic Transport Model (GTM). The structure of the model is made highly flexible such that a 10 percent of semi-span wing tip deflection is expected at a design lift condition, and a representation of the VCCTEF concept is incorporated on the model. Static aeroelastic modeling is conducted by using a representative single beam structural finite-element model coupled to a vortex-lattice aerodynamic model. The resulting aeroelastic model of the flexible wind tunnel model is compared against experimental wind tunnel test results from the actual model tested at the University of Washington Aeronautical Laboratory (UWAL). Comparison cases are made using reference and VICON measurements of VCCTEF flap deflections from the wind tunnel test data, and percent errors between the lift curve parameters are presented. The results show up to a maximum percent of error of <10% with regards to lift curve slope values between the aeroelastic model and the UWAL test data. The agreement of lift curve slope values from the aeroelastic model and experimental test results serves as validation for the coupled vortex-lattice finite-element static aeroelastic model.

Aeroelasticity↗
NASA NTRS1996

Modeling the Benchmark Active Control Technology Wind-Tunnel Model for Application to Flutter Suppression

This paper describes the formulation of a model of the dynamic behavior of the Benchmark Active Controls Technology (BACT) wind-tunnel model for application to design and analysis of flutter suppression controllers. The model is formed by combining the equations of motion for the BACT wind-tunnel model with actuator models and a model of wind-tunnel turbulence. The primary focus of this paper is the development of the equations of motion from first principles using Lagrange's equations and the principle of virtual work. A numerical form of the model is generated using values for parameters obtained from both experiment and analysis. A unique aspect of the BACT wind-tunnel model is that it has upper- and lower-surface spoilers for active control. Comparisons with experimental frequency responses and other data show excellent agreement and suggest that simple coefficient-based aerodynamics are sufficient to accurately characterize the aeroelastic response of the BACT wind-tunnel model. The equations of motion developed herein have been used to assist the design and analysis of a number of flutter suppression controllers that have been successfully implemented.

Waszak, Martin R.↗
NASA NTRS1998

Modeling the Benchmark Active Control Technology Wind-Tunnel Model for Active Control Design Applications

This report describes the formulation of a model of the dynamic behavior of the Benchmark Active Controls Technology (BACT) wind tunnel model for active control design and analysis applications. The model is formed by combining the equations of motion for the BACT wind tunnel model with actuator models and a model of wind tunnel turbulence. The primary focus of this report is the development of the equations of motion from first principles by using Lagrange's equations and the principle of virtual work. A numerical form of the model is generated by making use of parameters obtained from both experiment and analysis. Comparisons between experimental and analytical data obtained from the numerical model show excellent agreement and suggest that simple coefficient-based aerodynamics are sufficient to accurately characterize the aeroelastic response of the BACT wind tunnel model. The equations of motion developed herein have been used to aid in the design and analysis of a number of flutter suppression controllers that have been successfully implemented.

Waszak, Martin R.↗
NASA NTRS2014

Static Aeroelastic Scaling and Analysis of a Sub-Scale Flexible Wing Wind Tunnel Model

This paper presents an approach to the development of a scaled wind tunnel model for static aeroelastic similarity with a full-scale wing model. The full-scale aircraft model is based on the NASA Generic Transport Model (GTM) with flexible wing structures referred to as the Elastically Shaped Aircraft Concept (ESAC). The baseline stiffness of the ESAC wing represents a conventionally stiff wing model. Static aeroelastic scaling is conducted on the stiff wing configuration to develop the wind tunnel model, but additional tailoring is also conducted such that the wind tunnel model achieves a 10% wing tip deflection at the wind tunnel test condition. An aeroelastic scaling procedure and analysis is conducted, and a sub-scale flexible wind tunnel model based on the full-scale's undeformed jig-shape is developed. Optimization of the flexible wind tunnel model's undeflected twist along the span, or pre-twist or wash-out, is then conducted for the design test condition. The resulting wind tunnel model is an aeroelastic model designed for the wind tunnel test condition.

Aeroelasticity↗
NASA NTRS1979

Tests of a protective shell passive release mechanism for hypersonic wind-tunnel models

A protective shell mechanism for wind tunnel models was developed and tested. The mechanism is passive in operation, reliable, and imposes no new structural design changes for wind tunnel models. Methods of predicting the release time and the measured loads associated with the release of the shell are given. The mechanism was tested in a series of wind tunnel tests to validate the removal process and measure the pressure loads on the model. The protective shell can be used for wind tunnel models that require a step input of heating and loading such as a thin skin heat transfer model. The mechanism may have other potential applications.

Puster, R. L.↗
NASA NTRS1991

Model order reduction applied to a hot-bench simulation of an aeroelastic wind-tunnel model

Simulations of an aeroelastically scaled wind-tunnel model were developed for hot-bench testing of a digital controller. The digital controller provided active flutter-suppression, rolling-maneuver-load alleviation, and plant estimation. To achieve an acceptable time scale for the hot-bench application, the mathematical model of the wind-tunnel model was reduced from 220 states to approximately 130 states while assuring that the required accuracy was preserved for all combinations of 10 inputs and 56 outputs. The reduction was achieved by focussing on a linear, aeroelastic submodel of the full mathematical model and by applying a method based on the internally balanced realization of a dynamic system. The error-bound properties of the internally balanced realization significantly contribute to its utility in the model reduction process. The reduction method and the results achieved are described.

Buttrill, Carey S.↗
NASA NTRS1983

A common geometric data-base approach for computer-aided manufacturing of wind-tunnel models and theoretical aerodynamic analysis

A more automated process to produce wind tunnel models using existing facilities is discussed. A process was sought to more rapidly determine the aerodynamic characteristics of advanced aircraft configurations. Such aerodynamic characteristics are determined from theoretical analyses and wind tunnel tests of the configurations. Computers are used to perform the theoretical analyses, and a computer aided manufacturing system is used to fabricate the wind tunnel models. In the past a separate set of input data describing the aircraft geometry had to be generated for each process. This process establishes a common data base by enabling the computer aided manufacturing system to use, via a software interface, the geometric input data generated for the theoretical analysis. Thus, only one set of geometric data needs to be generated. Tests reveal that the process can reduce by several weeks the time needed to produce a wind tunnel model component. In addition, this process increases the similarity of the wind tunnel model to the mathematical model used by the theoretical aerodynamic analysis programs. Specifically, the wind tunnel model can be machined to within 0.008 in. of the original mathematical model. However, the software interface is highly complex and cumbersome to operate, making it unsuitable for routine use. The procurement of an independent computer aided design/computer aided manufacturing system with the capability to support both the theoretical analysis and the manufacturing tasks was recommended.

See, M. J.↗
NASA NTRS1990

Analytical aerodynamic model of a high alpha research vehicle wind-tunnel model

A 6 DOF analytical aerodynamic model of a high alpha research vehicle is derived. The derivation is based on wind-tunnel model data valid in the altitude-Mach flight envelope centered at 15,000 ft altitude and 0.6 Mach number with Mach range between 0.3 and 0.9. The analytical models of the aerodynamics coefficients are nonlinear functions of alpha with all control variable and other states fixed. Interpolation is required between the parameterized nonlinear functions. The lift and pitching moment coefficients have unsteady flow parts due to the time range of change of angle-of-attack (alpha dot). The analytical models are plotted and compared with their corresponding wind-tunnel data. Piloted simulated maneuvers of the wind-tunnel model are used to evaluate the analytical model. The maneuvers considered are pitch-ups, 360 degree loaded and unloaded rolls, turn reversals, split S's, and level turns. The evaluation finds that (1) the analytical model is a good representation at Mach 0.6, (2) the longitudinal part is good for the Mach range 0.3 to 0.9, and (3) the lateral part is good for Mach numbers between 0.6 and 0.9. The computer simulations show that the storage requirement of the analytical model is about one tenth that of the wind-tunnel model and it runs twice as fast.

Cao, Jichang↗
NASA NTRS1985

Investigation of Low-temperature Solders for Cryogenic Wind Tunnel Models

The advent of high Reynolds number cryogenic wind tunnels has forced alteration of manufacturing and assembly techniques and eliminated usage of many materials associated with conventional wind tunnel models. One of the techniques affected is soldering. Solder alloys commonly used for wind tunnel models are susceptible to low-temperature embrittlement and phase transformation. The low-temperature performance of several solder alloys is being examined during research and development activities being conducted in support of design and fabrication of cryogenic wind tunnel models. Among the properties examined during these tests are shear strength, surface quality, joint stability, and durability when subjected to dynamic loading. Results of these tests and experiences with recent models are summarized.

Firth, G. C.↗
NASA NTRSDate not supplied

Gust Load Alleviation Control and Gust Estimation for a High Aspect Ratio Wing Wind Tunnel Model

This paper presents the gust load alleviation (GLA) study of the aspect ratio 13.5 Common Research Model (CRM) wind-tunnel model. This study details the design of the GLA controller in preparation for wind-tunnel testing in the Transonic Dynamics Tunnel at NASA Langley Research Center. An aeroservolastic (ASE) model was first reduced using a model reduction method that takes advantage of the sinusoidal steady-state response. Then, the reduced model was used to design an extended-state Kalman filter which estimates the states and the sinusoidal gust input. The GLA control was then derived using the optimal control solution to a multi-objective cost function. The results of the GLA controller indicate a 75.6% reduction in wing-root strain while maintaining robust stability margins. The final paper will include performance under sub-optimal conditions, e.g., sensor noise, uncertainty in gust frequency, etc..

Christopher Forte↗
NASA NTRSDate not supplied

Gust Load Alleviation Control and Gust Estimation for a High Aspect Ratio Wing Wind Tunnel Model

This paper presents the gust load alleviation (GLA) study of the aspect ratio 13.5 Common Research Model (CRM) wind-tunnel model. This study details the design of the GLA controller in preparation for wind-tunnel testing in the Transonic Dynamics Tunnel at NASA Langley Research Center. An aeroservolastic (ASE) model is first reduced using a model reduction method that takes advantage of the sinusoidal steady-state response. Then, the reduced model is used to design an extended-state Kalman filter which estimates the states and the sinusoidal gust input. The GLA control is then derived using the optimal control solution to a multi-objective cost function. The results of the GLA controller indicate a 69.07% reduction in wing-root strain without sensor noise and 68.45% reduction with sensor noise, while maintaining robust stability margins. An adaptive GLA controller is developed for uncertain gust frequency and shows a 71.03% reduction in wing root strain compared to the non-adaptive control reduction of just 39.19%.

Christopher Forte↗