• Current Industrial and Technological Trends in Aerospace
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• v.7 no.1
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• pp.43-55
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• 2009

Advanced countries in aerospace have been struggle to realize the hypersonic air-breathing system since originating the concept of the hypersonic air-breathing propulsion system during the first half of the twentieth century. At last, NASA's X-43A Hyper-X did successful Mach 10 flight in November 2004. Each countries are running the program to applicate this hypersonic air-breathing propulsion system to SSTO(Single Stage to Orbit) or TSTO(Two Stage to Orbit) vehicle or hypersonic missile system at present. In this paper, we wrote the history and current issues of the hypersonic air-breathing propulsion system and hypersonic flight with the hypersonic air-breathing propulsion system.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.03a
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• pp.283-292
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• 2008

The foremost criterion in the design of a Satellite Launch Vehicle(SLV) is its performance capability to boost the designated payload to the desired mission orbit; it starts from focusing on the SLV configuration to achieve the velocity requirements($}\Delta}V$) for the mission. In this paper we review an analytical approach which is suitable enough for preliminary conceptual design and is used previously to optimize stage configurations for Two Stage to Orbit SLV for Low Earth Orbit(LEO) Missions; we have extended this approach to Three Stage to Orbit SLV and compared different propellant options for the mission. The objective is to minimize the Gross Lift off Weight(GLOW). The primary performance figures of merit were the total inert weight of the SLV and the payload weight that the SLV could lift into LEO, given candidate propulsion systems. The optimization is achieved by configuring the $}\Delta}V$ between stages. A comparison of configurations of single-stage and multi-stage SLVs is made for different propellants. Based upon the optimized stage configurations a comparative performance analysis is made between Liquid and Solid fueled SLV. A 3 degree of freedom trajectory-analysis program is modeled in SIMULINK and used to conduct the performance analysis. Furthermore, a cost analysis is performed on our stage optimized SLVs. The cost estimation relationships(CER) used give us a comparison of development and fabrication costs for the Liquid vs. Solid fueled SLV in man years. The pros and cons of the production, operation ability, performance, responsiveness, logistics, price, shelf life, storage etc of both Solid and Liquid fueled SLVs are discussed. The statistics and data are used from existing or historical(real) SLV stages.

• Structural Engineering and Mechanics
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• v.47 no.2
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• pp.149-166
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• 2013

Flexible behaviors in new aerospace structures can lead to a degradation of their control and guidance system and undesired performance. The objectives of the current work are to analyze the vibration resulting from the propulsion force on a Single Stage to Orbit (SSTO) launch vehicle (LV). This is modeled as a follower force on a free-free Euler-Bernoulli beam consisting of two concentrated masses at the two free ends. Once the effects on the oscillation of the actuators are studied, a solution to reduce these oscillations will also be developed. To pursue this goal, the stability of the beam model is studied using Ritz method. It is determined that the transverse and rotary inertia of the concentrated masses cause a change in the critical follower force. A new dynamic model and an adaptive control system for an SSTO LV have been developed that allow the aerospace structure to run on its maximum bearable propulsion force with the optimum effects on the oscillation of its actuators. Simulation results show that such a control model provides an effective way to reduce the undesirable oscillations of the actuators.

• Journal of the Korean Society of Manufacturing Technology Engineers
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• v.24 no.3
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• pp.320-326
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• 2015

Despite the wide industrial applications of planetary gear trains, the relationship between the design parameters (tooth profile, carrier mass, etc.) and performance (strength, vibration, noise, etc.) remains poorly understood. A significant amount of research has focused on transmission errors, which are measurable performance indicators directly related to the design parameters. Herein, an experimental test rig for a single-stage planetary gear set built using digital angular encoders and gap sensors is described. To study the static and dynamic characteristics of this planetary gear train, the transmission errors and sun gear orbit are analyzed from the data measured under various levels of torque and speed. The transmission errors of the gear train decrease 40% when the speed increases from 30 to 600 rpm with an output torque of 39.2 Nm, and increase 22% when the output torque increases from 19.6 to 39.2 Nm with an input speed of 30 rpm.

• Journal of the Korean Crystal Growth and Crystal Technology
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• v.13 no.3
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• pp.132-138
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• 2003

A stoichiometric mixture for $CdIn_2Te_4$ single crystal was prepared from horizontal electric furnace. The $CdIn_2Te_4$ single crystal was grown in the three-stage vertical electric furnace by using Bridgman method. The $CdIn_2Te_4$ single crystal was evaluated to be tetragonal by the power method. The (001) growth plane of oriented $CdIn_2Te_4$ single crystal was confirmed from back-reflection Laue patterns. The carrier density and mobility of $CdIn_2Te_4$ single crystal measured with Hall effect by van der Pauw method are $8.61\times 1016 \textrm {cm}^{-3}$ and 242 $\textrm{cm}^2$/V.s at 293 K, respectively. The temperature dependence of the energy band gap of the $CdIn_2Te_4$ single crystal obtained from the absorption spectra was well described by the Varshni's relation, $1.4750ev - (7.69\times10^{-3})\; ev/k)\;T^2$/(T + 2147k).The crystal field and the spin-orbit splitting energies for the valence band of the $CdIn_2Te_4$ single crystal have been estimated to be 0.2704 eV and 0.1465 eV, respectively, by means of the photocurrent spectra and the Hopfield quasicubic model. These results indicate that the splitting of the $\Delta$so definitely exists in the $\Gamma_7$ states of the valence band of the $CdIn_2Te_4$ single crystal. The three photocurrent peaks observed at 10 K are ascribed to the $A_{1-} B_{1-}$ and Cl-exciton peaks for n = 1.

• Proceedings of the Korean Institute of Electrical and Electronic Material Engineers Conference
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• 2003.07a
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• pp.347-351
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• 2003

A p-$CdIn_2Te_4$ single crystal has been grown by the Bridgman method without a seed crystal in a tree-stage vertical electric furnace. From photocurrent measurements, it was found that three peaks, A, B, and C, corresponded to an intrinsic transition due to the band-to-band transition from the valence band states ${\Gamma}_7(A),\;{\Gamma}_6(B),\;and\;{\Gamma}_7(C)$ to the conduction band state ${\Gamma}_6$, respectively. Also, the valence band splitting of the $CdIn_2Te_4$ crystal has been confirmed by photocurrent spectroscopy. The crystal field splitting and the spin orbit splitting were obtained to be 0.2360 and 0.1119 eV, respectively. Also, the temperature dependence of the band gap energy of the $CdIn_2Te_4$ crystal has been driven as the following equation of $E_g(T)\;=E_g(0)\;-\;(9.43\;{\times}\;10^{-3})T^2/(2676\;+\;T)$. In this equation, the Eg(0) was estimated to be 1.4750, 1.7110, and 1.8229 eV at the valence band state A, B, and C, respectively. The band gap energy of the p-$CdIn_2Te_4$ at room temperature was determined to be 1.2023 eV.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2004.03a
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• pp.753-762
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• 2004

Conceptual studies of a combined-cycle engine have been conducted. Herein, the results are presented. The engine is composed of ejector-jet, ramjet, scramjet and rocket modes, and will be mounted on the Single-Stage-to-Orbit aerospace plane. Propellants are hydrogen and oxygen. Calculated engine thrust performances and cooling requirement of the engine are presented. Pitching moment of the plane with the engine will be balanced even in the vacuum condition. The experimental results of the inlet and the ejector-jet, ramjet and scramjet modes are presented. The effect of the airframe configuration on the engine performance and the thermal environment in the in-side of the plane are also presented. Through the investigations, possibilities of the combined-cycle engine and the aerospace plane are being made clear now.

• Proceedings of the Korean Institute of Electrical and Electronic Material Engineers Conference
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• 2004.11a
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• pp.179-182
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• 2004

The $p-CdIn_2Te_4$ single crystal was grown in the three-stage vertical electric furnace by using Bridgman method. The quality of the grown crystal has been investigated by the x-ray diffraction and the photoluminescence measurements. From the Photoluminescence spectra of the as-grown $CdIn_2Te_4$ crystal and the various heat-treated crystals, the $(D^{o},X)$ emission was found to be the dominant intensity in the photoluminescence spectrum of the $CdIn_2Te_4:Cd$, while the $(A^{o},X)$ emission completely disappeared in the $CdIn_2Te_4:Cd$. However, the $(A^{o},X)$ emission in the photoluminescence spectrum of the $CdIn_2Te_4:Te$ was the dominant intensity like an as-grown $p-CdIn_2Te_4$ crystal. These results indicated that the $(D^{o},X)$ is associated with $V_{Te}$ acted as donor and that the $(A^{o},X)$ emission is related to $V_{Cd}$ acted as acceptor, respectively. The $p-CdIn_2Te_4$ crystal was found to be obviously converted into the n-type after annealing in the Cd atmosphere. The origin of $(D^{o},\;A^{o})$ emission and its TO Phonon replicas is related to the interaction between donors such as $V_{Te}$ or $Cd_{int}$, and accepters such as $V_{Cd}$ or $Te_{int}$. Also, the In in the $CdIn_2Te_4$ was confirmed not to form the native defects because it existed in the stable form of bonds.

• Journal of the Korean Society of Visualization
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• v.10 no.1
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• pp.47-54
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• 2012

Single Stage To Orbit (SSTO) missions which require its engines to be operated at varying back pressure conditions, use engines operate at high combustion chamber pressures (more than 100bar) with moderate area ratios (AR 70~80). This ensures that the exhaust jet flows full during most part of the operational regimes by optimal expansion at each altitude. Aero-spike nozzle is a kind of altitude adaptation nozzle where requirement of high combustion chamber pressures can be avoided as the flow is adapted to the outside conditions by the virtue of the nozzle configuration. However, the thrust prediction using the conventional thrust equations remains to be a challenge as the nozzle plume shapes vary with the back pressure conditions. In the present work, the performance evaluation of a new aero-spike nozzle is being carried out. Computational studies are carried out to predict the thrust generated by the aero-spike nozzle in varying back pressure conditions which requires the unsteady pressure boundary conditions in the computational domain. Schlieren pictures are taken to validate the computational results. It is found that the flow in the aero-spike nozzle is mainly affected by the base wall pressure variation. The aerospike nozzle exhibits maximum performance in the properly expanded flow regime due to the open wake formation.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.03a
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• pp.330-339
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• 2008

Ground-based interceptors(GBI) comprise a major element of the strategic defense against hostile targets like Intercontinental Ballistic Missiles(ICBM) and reentry vehicles(RV) dispersed from them. An optimum design of the subsystems is required to increase the performance and reliability of these GBI. Propulsion subsystem design and optimization is the motivation for this effort. This paper describes an effort in which an entire GBI missile system, including a multi-stage solid rocket booster, is considered simultaneously in a Genetic Algorithm(GA) performance optimization process. Single goal, constrained optimization is performed. For specified payload and miss distance, time of flight, the most important component in the optimization process is the booster, for its takeoff weight, time of flight, or a combination of the two. The GBI is assumed to be a multistage missile that uses target location data provided by two ground based RF radar sensors and two low earth orbit(LEO) IR sensors. 3Dimensional model is developed for a multistage target with a boost phase acceleration profile that depends on total mass, propellant mass and the specific impulse in the gravity field. The monostatic radar cross section (RCS) data of a three stage ICBM is used. For preliminary design, GBI is assumed to have a fixed initial position from the target launch point and zero launch delay. GBI carries the Kill Vehicle(KV) to an optimal position in space to allow it to complete the intercept. The objective is to design and optimize the propulsion system for the GBI that will fulfill mission requirements and objectives. The KV weight and volume requirements are specified in the problem definition before the optimization is computed. We have considered only continuous design variables, while considering discrete variables as input. Though the number of stages should also be one of the design variables, however, in this paper it is fixed as three. The elite solution from GA is passed on to(Sequential Quadratic Programming) SQP as near optimal guess. The SQP then performs local convergence to identify the minimum mass of the GBI. The performance of the three staged GBI is validated using a ballistic missile intercept scenario modeled in Matlab/SIMULINK.