• Journal of Advanced Navigation Technology
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• v.26 no.1
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• pp.9-14
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• 2022

In this paper we investigate the advantages when the GPS Antenna is installed on the rear missile body. In conventional design the GPS antenna locates on the front part of missile. However it causes degraded GPS positioning performance since the missile body blocks the GPS signals. This paper proposes the GPS array antenna design which locates on the rear part of missile body and has the tilted antenna patches to achieve the maximum area of receiving GPS signals. We simulate LOS region of receiving signals and conducted anechoic chamber test to define the effective signal receiving region. And we conduct field test and flight test to check out the enhancement of signal receiving area.

• Transactions of the Korean Society for Noise and Vibration Engineering
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• v.15 no.7 s.100
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• pp.813-820
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• 2005

A deployable missile control fin has some structural nonlinearities because of the worn or loose hinges and the manufacturing tolerance. The structural nonlinearity cannot be eliminated completely, and exerts significant effects on the static and dynamic characteristics of the control fin. Thus, It is important to establish the accurate deployable missile control fin model. In the present study, the nonlinear dynamic model of 4he deployable missile control fin is developed using a substructure synthesis method. The deployable missile control fin can be subdivided Into two substructures represented by linear dynamic models and a nonlinear hinge with structural nonlinearities. The nonlinear hinge model is established by using a system identification method, and the substructure modes are improved using the Frequency Response Method. A substructure synthesis method Is expanded to couple the substructure models and the nonlinear hinge model, and the nonlinear dynamic model of the fin is developed. Finally, the established nonlinear dynamic model of the deployable missile control fin is verified by dynamic tests. The established model is In good agreement with test results, showing that the present approach is useful in aeroelastic stability analyses such as time-domain nonlinear flutter analysis.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.33 no.3
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• pp.26-31
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• 2005

An efficient estimation of the aerodynamic characteristics for missile configurations is essential in the preliminary stage of a missile design. In this study, a Missile DATCOM family code based on the semi-empirical method was utilized for this purpose. In order to check the accuracy and reliability of the code several test cases have been considered: subsonic flow with high angles of attack and supersonic flow with moderate angles of attack. It turned out that the code in general provides prediction in qualitative agreement with the experimental data and results by other works. Finally, the code was applied to a more complicated missile configuration with canard and freely spinning tail fin.

• Proceedings of the KIPE Conference
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• 2010.07a
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• pp.520-521
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• 2010

본 논문은 군사용 미사일의 조립 후 성능 검증을 위한 Missile Assembly Test System(MATS)용 IPMSM의 위치 제어 방법을 구현하였다. 이를 위하여 슬라이딩 모드 제어기와 PI 제어기를 Membership 함수를 사용하여 혼합한 개선된 위치제어 알고리즘을 제안하였다. 오버슈트가 없고, 외란에 대하여 강인하게 운전 궤적을 추정하는 특성이 뛰어난 슬라이딩 모드 제어의 장점과, 빠른 제어 응답성과 정상상태에서 리플이 적은 PI제어를 적절하게 혼합하였고, 이를 시뮬레이션을 통하여 IPMSM의 위치제어 성능을 검증하였다.

• Convergence Security Journal
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• v.20 no.4
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• pp.3-12
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• 2020

Although the related regulations or guidelines are not specified in the defense weapon system R&D process, R&D authorities frequently encounter problems that require rational decision-making. If the rational process is not applied in the matter of alternative choice, the project could be disrupted, which can result in longer project periods or more resource provision. In particular, a variety of decision-making methods are needed for test&evaluation of missile R&D. The issue of selecting a test site is one of the representative decision-making problems. If it is needed to determine the priority of multiple sites, Delphi Method and Analytic Hierarchy Process(AHP) will be applied. However, if the input of cost is to be considered, Data Envelopment Analysis(DEA) is more valuable to solve the problem. This paper proposes a solution to handle quantitatively various decision-making problems that can occur in missile flight test, and shows how DEA is applied through a simulated case study of selecting a foreign test site.

• Proceedings of the KSME Conference
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• 2002.08a
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• pp.87-90
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• 2002

Computational study has been undertaken to investigate the aerodynamic influence of side jet on a supersonic missile and to find a similarity condition between the flight condition and the wind tunnel testing. Tasks were performed to validate the existing Raytheon test body with side jet, to simulate the flow inside the supersonic wind tunnel, and to compare the flow fields between the missile in free flight and that in the wind tunnel. Then sub-scale model of body-tail configuration was analyzed to estimate the influence of the side jet on the missile components. It Is found that the influence of side Jet is not as significant on the tail region as on the body surface and a simple algebraic formula for aerodynamic coefficients accounting for the side jet as a point force may be cautiously utilized in setting up control logic.

• Proceedings of the Korean Society for Noise and Vibration Engineering Conference
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• 2004.05a
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• pp.629-634
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• 2004

Modal parameters of the total missile structure including a hinge mechanism are estimated by the experimental modal analysis. The free-free boundary condition is simulated by hanging the missile structure with a wire rope, and the missile structure is excited by the random vibration technique. Test results are used to verify the FE analysis, the 1-D FE model is modified by 3-D model at the hinge part. Consequently, the modal parameters of the missile structure are estimated preciously.

• Journal of Korean Society for Quality Management
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• v.48 no.1
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• pp.215-225
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• 2020

Purpose: The purpose of this study was to show current states of domestic Precision Guided Missile(PGM) by analyzing Live-fire test results using general methods to get the Confidence Levels. Methods: Live-fire test results were used to get Confidence Levels of PGM. The Confidence Levels were derived by two general methods. The first method was Binomial distribution and second was convergence of Hypergeometric distribution and Bayes' rule. Results: The results of this study are as follows; The more Live-fire tests of PGM are performed, the higher Confidence Level of PGM will be estimated. And the number of Live-fire tests are related to a unit price of PGM. This results means that the increase of live-fire test, which is useful data for preparation and evaluation of Development Tests / Operation Tests for PGMs, is only way to enhance the Confidence Levels of each PGMs. Conclusion: This study shows the relationship between the Live-fire tests and Confidence Levels of PGMs and it will be used on Live-fire Test & Evaluation of PGMs for reference.

• Journal of the Korean Society of Propulsion Engineers
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• v.4 no.3
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• pp.19-28
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• 2000

In general the data such as thrust, pressure, temperature and combustion time are measured in developing the propulsion system of solid rocket motor through static firing test. But in the case of personal surface to air missile there are required a severe safety specifications in order to eliminate gunner hazard from the exhaust plume of motors. The safety requirements lead to the design of separation device and safety igniter device. The dynamic firing test for the designed two devices should be conducted under the flight environmental conditions to verify the requirements compliance. In this study the technique for dynamic firing test of propulsion system of personal surface to air missile is proposed and the method to design the dynamic test bench is also studied.

• 제어로봇시스템학회:학술대회논문집
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• 1990.10a
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• pp.475-480
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• 1990

A design of aeroload simulator is proposed which is used for the dynamic test of guided missile servoactuation system. Since the simulator is linked to the servoactuation system, it is a two-input one-output system. The mathematical model of the aeroload simulator is derived and the root-locus method is applied to design compensator to improve the stability, response speed and accuracy of the system. Computer simulations are done to verify the goodness of the system design.