The two-step filter has been combined with a modified Sage-Husa time-varying measurement noise statistical estimator, which is able to estimate the covariance of measurement noise on line, to generate an adaptive two-step filter. In many practical applications such as the bearings-only guidance, some model parameters and the process noise covariance are also unknown a priori. Based on the adaptive two-step filter, we utilize multiple models in the first-step filtering as well as in the time update of the second-step filtering to handle the uncertainties of model parameters and process noise covariance. In each timestep of the multiple model filtering, probabilistic weights punishing the estimates of first-step state from different models, and their associated covariance matrices are acquired according to Bayes’ rule. The weighted sum of the estimates of first-step state and that of the associated covariance matrices are extracted as the ultimate estimate and covariance of the first-step state, and are used as measurement information for the measurement update of the second-step state. Thus there is still only one iteration process and no apparent enhancement of computation burden. A motion tracking sliding-mode guidance law is presented for missiles with non-negligible delays in actual acceleration. This guidance law guarantees guidance accuracy and is able to enhance observability in bearings-only tracking. In bearings-only cases, the multiple model adaptive two-step filter is applied to the motion tracking sliding-mode guidance law, supplying relative range, relative velocity, and target acceleration information. In simulation experiments satisfactory filtering and guidance results are obtained, even if the filter runs into unknown target maneuvers and unknown time-varying measurement noise covariance, and the guidance law has to deal with a large time lag in acceleration.

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An efficient way to systematically reproduce inflow profiles across a compressible, turbulent boundary layer for the mean flow variables, as well as the turbulent quantities in one-equation and two-equation turbulence-models, from given external flow conditions and one boundary layer parameter, is described. The reproduced profiles are checked against both experimental results and numerical solution of boundary layer equations. Theoretical analysis shows that a new form of density-weighted velocity, rather the Van Driest density-weighted velocity, obeys the linear-law at the viscous sublayer in a compressible turbulent boundary layer. Especially for non-adiabatic wall at hypersonic Mach numbers, where there are large density gradients, these two kinds of density-weighted velocity could differ considerably. Therefore, the new form of density-weighted velocity proposed in this paper should be employed for the viscous sublayer. It is also shown that power-law fitting for the streamwise velocity gives an unacceptable profile in the viscous sublayer. An efficient way to specify the normal velocity profile is also proposed and tested. The reproduced normal velocity at the boundary layer edge is found to agree remarkably well with the numerical solution of boundary layer equations.

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The paper is focused on an equilibrium analysis of the HOPE-X with six degrees of freedom nonlinear equations of motion and nonlinear aerodynamic data. The analysis is made primarily by using the continuation method, and the problem formulation for the method is described. Results are compared with those of two types of simplified analyses to suggest that rigorous analyses are necessary for control system development. Stability analyses of the obtained equilibrium branches are then undertaken to pose the problem that the vehicle has no attractor in a subsonic range under investigation. Future works are finally referred to with regard to the continuation method and the speed range of the vehicle to be pursued further.

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The third-generation Korean sounding rocket, KSR-III, was successfully launched on Nov. 28, 2002 from the western coastal area of the Korean Peninsula. The previous KSR-I and -II series had employed a solid propulsion system whereas the KSR-III utilizes Korea’s first liquid propulsion system. The prime objective of the mission was to evaluate the performance of the liquid propulsion system. The onboard electronics system of the KSR-III is an enhanced version of the ones used in the previous series. The system has extended data channels and has adopted a distributed data processing system using the RS-485 bus network. The rocket has various sensors to measure physical characteristics such as temperature, pressure, strain, acceleration, etc., and has scientific instruments including an ozone detector and two magnetometers. The flight data is transmitted to the ground station in real time by the onboard telemetry system. The onboard electronics system of the KSR-III mainly consists of telemetry, an RF subsystem, a tele-command system, a power supply system, and scientific payloads. Herein, we present an overview of the enhanced electronics system of the KSR-III and representative flight test results analyses including scientific data are discussed.

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The effects of flare control on the aerodynamic characteristics, performance, and stability of a cylindrical body under laminar and turbulent boundary layer conditions have been studied experimentally and computationally. The experimental study has been carried out in a hypersonic gun tunnel at a Mach number of 8.2 and a Reynolds number of 158,100, based on the cylinder diameter, at flare angles 0, 5, 10, 15, 20, 25 and 30 degrees and at pitch angles of −12 to 12 deg for the 10 deg flare case only. The surface flow was studied using the oil-dot technique. Some information regarding the shock layer was obtained from schlieren pictures. The effects of turbulence on onset of separation were also deduced from pressure measurements over the cylinder and the flare. The forces were measured with a three-component balance equipped with semiconductor strain gauges. The effects of centre of gravity (CG) location on the aerodynamic characteristics and in particular on the C_Mαwere examined. The results under turbulent conditions and zero-incidence were compared with numerical simulations performed using a 3-D time-marching Navier-Stokes code. The magnitude of the separated region, the minimum flare angle required to induce separation, and the effects of small-scale separation are detailed.

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The structures and characteristics of λ-shaped and X-shaped pseudo-shocks in a square duct are investigated through numerical simulations and experiments at Mach 2 and Mach 4, respectively. The experiments were carried out in a pressure-vacuum supersonic wind tunnel with a test cross section of 80×80 mm². Numerical simulations were carried out using the Harten-Yee second-order TVD scheme and the Baldwin-Lomax turbulence model. The Reynolds numbers for the Mach 2 and 4 cases were Re_∞=2.53×10⁷ and Re_∞=2.36×10⁷, respectively, and the flow confinement was δ_∞⁄h=0.35 for both cases. The computational results for the Mach 2 pseudo-shock wave are in good agreement with the experimental results. Based on this agreement, the flow quantities, which are very difficult to obtain experimentally, were analyzed by numerical simulation. Although several differences were found between the computational results and experiments in the case of Mach 4 due to asymmetric characteristics in experiment which could not be reproduced in numerical simulation, the computational results are valuable for understanding this complex asymmetric phenomenon.

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The effects of jet control on the aerodynamic characteristics and performance of a flat-plate body configuration have been studied experimentally. The study has been carried out in a hypersonic gun tunnel at a Mach number of 8.2 and a Reynolds number of 1,900,000, based on the flat-plate length-chord, using both sharp and blunted leading edges, at pitch angles 0, −5 and −10 deg. Air was used as the working gas for both the free stream and the sonic jet. The tests employed schlieren photography to study the overall flow field. Quantitative studies of the effect of the jet have been made by pressure measurements. The results were compared with numerical simulations performed using a 3-D time-marching Navier-Stokes code. Theoretical algorithms have been developed to predict the shape of the separation front and jet penetration height. The parameters, which determine the upstream extent and lateral spreading of the separation front around the transverse jet, and the magnitude of the separated region, are detailed.

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This paper proposes an exact polytopic model for the linear parameter varying (LPV) system of aircraft. The LPV system of the linearized equation of aircraft is represented by a descriptor form which reserves physical features of the equation. A polytopic model, called descriptor polytopic model, is derived through a variable transformation to satisfy conditions for the polytope. Using the obtained descriptor polytopic model, a gain scheduling state feedback law is then designed by means of a linear matrix inequality (LMI) formulation. It is shown in a numerical example of a longitudinal flight control that the proposed descriptor polytopic model of aircraft had no model error and exactly represented the original LPV system without any unnecessary flight region.

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The temperature depression of liquids due to latent heat of vaporization causes vapor pressure depression and suppresses cavity growth. This phenomenon is called “thermodynamic effect of cavitation.” This effect is especially significant in cryogenic fluids such as LOX and LH₂. Due to this effect, the performance of hydraulic equipment for cryogenic fluids, such as turbopumps of rocket engines, is not as bad as predicted. In this paper, the size of the cavity in cryogenic fluid is estimated numerically taking the thermodynamic effect of cavitation into consideration. A cavity is assumed to be a sheet cavity. From the results, the effects of the properties of liquid and Reynolds number on the thermodynamic effect of cavitation are investigated.

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The classical Schwartz-Christoffel transformation can be extended to airfoils by a modern computational approach, which otherwise would be limited to calculating potential flows around simple geometric shapes. A numerical method is proposed to determine the Schwartz-Christoffel transformation from a unit circle to paneled general aviation airfoils. The present calculations agree with the theoretical predictions, and they are also compared to experimental results.

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