• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.47 no.8
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• pp.574-581
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• 2019

This paper investigates the effects of thrust misalignments and offsets of the lateral center of gravity of a space launch vehicle on its guidance performance. Sensitivity analysis and Monte Carlo simulations are applied to analyze their effects by computing changes in orbit injection errors when including the error sources. To compensate their effects, the attitude controller including an integrator additionally and the Steering Misalignment Correction (SMC) routine of the Saturn V are considered, and then Monte Carlo simulations are performed to evaluate their performances.

• Transactions of the Korean Society of Mechanical Engineers A
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• v.36 no.11
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• pp.1399-1404
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• 2012

Because satellites operate in space, it is impossible to repair them when they malfunction. Therefore, to ensure the normal function of the payload used in the satellites, accurate assembly and installation of parts are crucial. To prevent abnormal functioning in the extreme environments during launch and in space, it is essential to test changes at the parts and system levels by performing alignment measurement before and after the launch environment test and the space environment test. Recently, noncontact three-dimensional precision machinery for medium- and large-sized parts has been developed. One of these is the theodolite measurement system, which is widely used in the aerospace industry. This study measures the angle of the dual thruster module that is used to control the attitude of KOMPSAT by using a theodolite, and alignment measurement and a reliability analysis are performed.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2007.04a
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• pp.281-285
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• 2007

The flow characteristics in the thruster should be analyzed considering its geometry and the pressure ratio to estimate its performance and etc. This paper suggests the computational result of an axisymmetric real nozzle for the altitude control of a satellite to find out the application limit that the assumption of continuum mechanics holds. The steady non-reacted compressible flow field in the unstructured grid system is computed and analyzed with varying the environmental pressure (or the degree of vacuum) under the fixed pressure ratio in a real thruster of which the area ratio of exit to throat is 56. The assumption of the continuum mechanics is not approved when the environmental pressure is reduced less than $10^{-3}$ atm.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.39 no.12
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• pp.1097-1106
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• 2011

Unsteady numerical analysis was performed to simulate air-launched missiles from a complete helicopter in hover by using an unstructured overset mesh flow solver coupled with a module of six degree-of-freedom motion of equations. The unsteady computations have been performed to obtain flow fields around the complete helicopter including main rotor, tail rotor, and fuselage equipped with multiple missiles, and six-DOF simulation has been performed to predict the behavior of the air-launched missile. The effects of the launching position and the missile thrust on the trajectory of the missile were investigated as well as the aerodynamic interference of the air-launched missile under the unsteady downwash produced by main rotor.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.31 no.5
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• pp.1-9
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• 2003

Detailed flow of a shrouded tail rotor in hover is studied by using a compressible inviscid flow solver on unstructured meshes. The numerical method is based on a cell-centered finite-volume discretization and an implicit Gauss-Seidel time integration. Numerical simulation is made for a single blade attached to the center body and guide by the duct by imposing a periodic boundary condition between adjacent rotor blades. The results show that the performance of an isolated rotor without shroud compares well with experiment. In case of a shrouded rotor, correction of the collective pitch angle is made such that the overall performance matches with experiment to account for the uncertainties of the experimental model configuration. Details of the flow field compare well with the experiment confirming the validity of the present method.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.33 no.1
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• pp.1-10
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• 2005

An aerodynamic shape optimization technique has been developed for helicopter rotor blades in hover based on a continuous adjoint method on unstructured meshes. The Euler flow solver and the continuous adjoint sensitivity analysis were formulated on the rotating frame of reference for hovering rotor blades. In order to handle the repeated evaluation of the design cycle efficiently, the flow and adjoint solvers were parallelized using a domain decomposition strategy. A solution-adaptive mesh refinement technique was adopted for the accurate capturing of the tip vortex. Applications were made for the aerodynamic shape optimization of Caradonna-Tung rotor blades and UH60 rotor blades in hover. The results showed that the present method is an effective tool to determine optimum aerodynamic shapes of rotor blades requiring less torque while maintaining the desired thrust level.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.30 no.7
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• pp.137-143
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• 2002

This paper concerns the spin-stabilization for the second stage of KSR-III(Korea Sounding Rocket-III). The error sources in the second stage are defined, and then the effect on attitude error of KSR-III is analyzed quantitatively and qualitatively. Based on the analysis, some conditions for optimal spinning frequency and optimal steady-spinning duration are suggested. Among several solutions, an optimal value can be determined in the point of minimum impact-point error. An approach to a sub-optimal solution is also suggested. Application to the KSR-III shows that a proper spinning frequency and a steady-spin duration to give the smallest impact-point error can be effectively determined.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.46 no.5
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• pp.419-426
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• 2018

For an orbit transfer in a space exploration mission, a solid or liquid rocket booster is included at the last stage of the launch vehicle. During the orbit transfer, thrust misalignment can cause a severe orbit error. Three axis attitude control or spin stabilization can be implemented to minimize the error. Spin stabilization technique has advantages in structural simplicity and lightness. One of ways to apply the spin stabilization to the payload is to include a spin table system in the launch vehicle. In this paper, effect of the spin table system on separation dynamics of the payload is analyzed. Simple model of the spin table to mimic basic functions is designed and simulation environment is established with the model. Effect of the spin table is tested by evaluating separation dynamics of a payload with and without the spin table. Analysis on tolerance effect of separation spring constant on separation dynamics of a payload is conducted.

• Aerospace Engineering and Technology
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• v.6 no.1
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• pp.146-156
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• 2007

In this paper, the TVC actuation test and analysis results for a flexible seal kick motor nozzle are presented. A real-sized test model of KSLV-I kick motor system is applied to water pressurizing TVC tests which investigate the property changes in TVC nozzle expansion and TVC actuation performance against chamber pressure changes. The equipments which are required for TVC actuation tests are briefly explained. The TVC actuation tests are firstly accomplished in static mode, which reveals TVC error characteristics including thrust misalignment, control accuracy, and TVC stroke increase, etc. The properties in frequency domain is given via dynamic tests. These results may play an important role in enhancing the TVC control performance of KSLV-I.

• Aerospace Engineering and Technology
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• v.3 no.1
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• pp.170-178
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• 2004

This paper deals with six degree-of-freedom HILS(hardware-in-the-loop-simulation) of KSR-III rocket using a TAFMS(three axis flight motion simulator). This TAFMS HILS test is accomplished before main HILS tests in order to verify the control stability in the presence of TAFMS dynamic effects. The TAFMS HILS test includes initial attitude holding tests for INS initial alignment procedures, timer synchronization tests with an auxiliary lift-off signal, real-time calibration tests using an external thermal recorder, open-loop TAFMS operating tests, and final closed-loop TAFMS HILS tests using the TAFMS attitude measurements as inputs to the closed control loop. The HILS tests are accomplished for several flight conditions composed with nominal flight condition, TWD effect added condition, slosh modes and/or bending modes existing condition, and windy condition, etc.