Se-Yoon Oh

Magnus and Spin-Damping Measurements of a Spinning Projectile Using Design of Experiments

The purpose of this research is to efficiently characterize the dynamic Magnus effect and spin-damping data of a spinning projectile or missile over the required experimental region using statistically designed experiments for wind-tunnel tests. In the present work, the wind-tunnel tests for theMagnus effect and spin-dampingmeasurements were conducted on a 155-mm spin-stabilized projectile model. The tests used techniques involving the design of the experiments and response surfacemodeling. TheMagnus effect and spin-damping datawere investigated for various Mach numbers and angle-of-attack combinations. The validity of the wind-tunnel measurement techniques was evaluated by comparing themwith the previous test results on the same configuration. The experimental results show that regressionmodels were found to successfully estimate theMagnus effect and spin-damping needed to investigate aerodynamic characteristics of a spinning projectile over the required experimental design space.

Optimal Aft End Distorted Fin Model Using Response Surface Method

In wind-tunnel testing of bomb ormissile-type models, aft end geometry distortion may become necessary when a model is tested by using a sting support system.When the diameter of themodel base is increased, the exposed tail fin area is reduced to result in aerodynamic data alteration. In the present work, various tests are conducted on anMK82 bomb model with base diameter and exposed tail fin area variations. The tests use techniques involving design of experiments and response surface modeling. The effects of the model base geometry modifications on the aerodynamic characteristics are investigated. Experimental design optimization is then performed on the basis of three design factors: the base diameter, the extended tail fin area, and the freestream velocity. The experimental results show that the altered aerodynamic characteristics due to aft end model distortion can be minimized by changing the tail fin geometry on the basis of the regression model.

An Experimental Study of Fuselage Drag and Stability Characteristics of a Helicopter Configuration

This paper describes the test carried out on an experimental study of fuselage drag and stability characteristics of a helicopter configuration and the test techniques developed for the testing and the lessons learned in the Agency for Defense Development Low Speed Wind Tunnel(ADD-LSWT). The main objective of this test is to determine the drag and stability characteristics of helicopter configurations according to the various configuration changes. The fuselage model with a highly modular structure is a representation of 1:8 scale of the external contour of the conceptual design helicopter configuration with rotating main rotor hub including blade stubs capable of rotating up to 500 rpm. The test results are compared with the available similar data and fair to good agreement is obtained.

An Experimental Study on Magnus Characteristics of a Spinning Projectile at High Speed Region

The purpose of this research is to determine the dynamic Magnus effect data of a spinning projectile in wind-tunnel testing. In the present work, the high-speed wind-tunnel tests for the Magnus effect measurements were conducted on a 155-mm spin-stabilized projectile model in the Agency for Defense Development`s Tri-Sonic Wind Tunnel at spin rates about 12,000 rpm. The test Mach numbers ranged from 0.7 to 2.0, and the angles of attack ranged from -4 to +10 deg. The validity of the wind-tunnel measurement techniques was evaluated by comparing them with the previous test results on the same configuration. The experimental results show that fair to good agreement is obtained with resonable accuracy.

Development of Magnus Effect Measurement Technique for Spinning Projectile

The Magnus effect measurement apparatus was designed and built for spinning wind tunnel model which would simulate the rotation of projectiles. Prior to the high speed test, the ground functional test and the low speed test were carried out in the Agency for Defense Development`s Low Speed Wind Tunnel(ADD-LSWT) at spin rates from about 6,000 to 10,000 rpm. Magnus force and moment were measured on the spinning projectile model at velocity of 100 m/s. It was shown that the Magnus force and moment were linear function of spin parameter. The test results were compared with Magnus test run on the same configuration in the Arnold Engineering Development Center`s Propulsion Tunnel 4T(AEDC-4T).

Response surface smoothing for wind tunnel testing based on design of experiment with subspace partitioning

This paper presents a new approach in resolving discontinuities between subspaces when the subspace partitioning method is applied to measure aerodynamic coefficients. The discontinuities between response surfaces in each subspace are hard to resolve in an analytical method, but it can be smoothened by overlapping between subspaces and integrating with proper interpolation. The traditional One Factor at a Time (OFAT) method is used for gathering aerodynamic force and moment coefficient data before applying the Design of Experiments (DOE) method for the purpose of later validation of the suggested method. In each subspace, the Central Composite Design (CCD) is used to reduce experimental points. The Separated Response Surface Model compared to the Integrated Response Surface Model. Quality of fit of the Developed Integrated Model calculated with OFAT data for validation. The test model base is the lambda wing configuration of the UCAV 1303. The study was carried out at the low speed wind tunnel of the Agency for Defense Development (ADD-LSWT).

Experimental Investigations of Systematic Errors in Wind Tunnel Testing Using Design of Experiments

The variation of systematic bias errors in the wind tunnel testing has been studied. A Design of Experiments(DOE) approach to an experimental study of fuselage drag and stability characteristics of a helicopter configuration was applied. When forces and moments measured in one time block differ significantly from measurements made in another time block under assumption that sample observations can be expected to yield same results within permissible measuring errors. The practical implication of this paper is that the systematic error can not be assumed not to exist. The those error reduction could be achieved through the process of randomization, blocking, and replication of the data points.

Wind-Tunnel Testing of an Inertial Particle Separator Inlet Using Design of Experiments

Wind-tunnel tests were conducted to determine the optimum bypass duct area of the inertial particle separator for the aircraft inlet duct. When the area of the bypass duct for the air inlet is excessively increased, the amount of the dumped air to overboard is increased to result in the inlet performance degradation. In the present work, various tests are conducted on an inertial particle separator of the nacelle inlet model with bypass area variations. The tests utilize techniques involving design of experiments and response surface modeling. The effects of the bypass area variation on the bypass ratio of the inertial particle separator system and the performance of the inlet system are investigated. Experimental design optimization is then performed on the basis of two design factors: the bypass area and the engine air flow condition. The experimental results show that the bypass duct area to meet the bypass ratio requirement provided by the aircraft manufacturer can be determined to yield maximum performance on the basis of the regression model.

An Experimental Study of Test Section Velocity Calibration for Low-Speed Wind Tunnel

The purpose of this research is to determine the calibration constants of the wind speed measurement systems required to calculate the wind tunnel velocity in the test section. In the present work, the aerodynamic calibration tests of the test section were conducted in the Agency for Defense Development`s Low-Speed Wind Tunnel. The test speed ranged from 10 to 100 m/s with a reference pitot-static pressure probe. The validity of the calibration results was evaluated by comparing them with the previous calibration constants. The calibration results show that fair to good agreement is obtained with resonable accuracy.

An Experiment Study on Sideslip Angle Effect of Lambda Wing Configuration

ABSTRACT An experimental study on aerodynamic coefficients of a lambda w ing configuration was performed at the low speed wind tunnel of Agency for Defense Development. The main purpose of this study was to investigate the effects of sidesli p angle on various aerodynamic coefficients. In the case of 0 ° sideslip angle, nose-up pitching moment rapidly increases at a specific angle of attack. This unstable pitching moment characteristic is referred to as pitch break or pitch up. As the sideslip angle increases, the pitch break is found to be generated at a higher angle of attack. Rolling mome nt is found to show similar behavior pattern to ‘pitch break’ style with angle of attack at non-zero sideslip angles. This trend gets severer at greater sideslip angles. Yawing moment also shows substantial variation of the slope and the unstable directional stability with sideslip angles at higher angles of attack. These characteristics of the three moments clearly implies the difficulty of the flight control which requires efficient contr ol augmentation system.