• Proceeding of EDISON Challenge
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• 2014.03a
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• pp.544-549
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• 2014

임계마하수보다 큰 자유흐름 마하수에서는 충격파의 발생으로 인해 급격한 항력증가가 발생하므로, 임계마하수 증가는 고속 공기역학에서 중요한 분야로 다뤄지고 있다. Whitcomb R. T.에 의해 천음속영역에서 순항할 수 있는 초임계익형이 개발되었으나, 충격파 제어 기법들에 대한 실험적인 검증은 형상 제작의 어려움으로 인해 한계를 지닌다. 따라서 본 논문에서는 2D_Comp-2.1_P와 Prandtl-Glauert 압축성 보정식을 이용하여 NACA0012와 RAE2822의 임계마하수를 해석하고, 충격파 제어 장치 중 하나인 Bump를 RAE2822에 설치하여 임계마하수를 향상시키기 위한 연구를 수행하였다. 연구 결과 충격파를 압축파로 분산시켜 충격파의 강도를 약화시키고, 양항비의 4.7% 증가를 확인하였다. 따라서 Bump를 설계한 RAE2822가 기본 익형보다 높은 천음속 조건에서 효율적인 공력특성을 가지는 것을 확인하였다.

• 한국가시화정보학회:학술대회논문집
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• 2002.11a
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• pp.109-112
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• 2002

The supersonic jet discharging from a petal nozzle is known to enhance mixing effect with the surrounding gas because it produces strong longitudinal vortices due to the velocity difference from both the major and minor axes of petal nozzle. In the present study, the supersonic free jet discharging from the petal nozzle is investigated experimentally. The nozzles used are 4, 6, and 8 lobed petal nozzles with a design Mach number of 1.7, and the flow fields are compared with a circular nozzle with the same design Mach number. The pitot impact pressures are measured using a fine pilot probe. The flow fields are visualized using a Schlieren optical method. The results show that the petal nozzle has more increased supersonic length compared with the circular jet.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.48 no.11
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• pp.849-859
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• 2020

The semi-empirical equation and commercial computational tool were used to predict the base drag of a guided missile with free-stream Mach numbers and chamber pressures, and the results were generally agree each other. Differences in flow characteristics and base drags were observed with over/under expansion conditions by the nozzle. Under the over-expansion condition, the base pressure decreased as the expansion fan was generated at upper region of the base, and base pressure decreased further with increasing free-stream Mach number as the expansion becomes strong. Under the under-expansion conditions, a shock wave was generated around the base by the influence of the nozzle flow, which increased the base pressure, and the effect increased as the chamber pressure increased. Under the same chamber pressure condition, as the free-stream Mach number increases, the characteristic that the base pressure decreases as the shock wave generated at the base moves downstream was observed.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.32 no.8
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• pp.64-71
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• 2004

A computational study on the supersonic flow around the lateral jet controlled missile has been performed. For this purpose a three dimensional Navier-Stokes computer code(AADL3D) has been developed and case studies have been performed by comparing the normal force coefficient and the moment coefficient of a missile body for different jet flow conditions including jet pressure and jet Mach number. The results show different behavior of normal force and moment variation according to jet pressure variation and jet Mach number variation. From the detailed flow field analyses, it is verified that most of the normal force loss and the pitching moment generation are taken place at the low-pressure region behind the jet nozzle. Furthermore, it is shown that the pitching moment can be efficiently reduced by obtaining the lateral thrust through higher jet Mach number rather than through high jet pressure.

• Journal of the Korean Society of Propulsion Engineers
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• v.24 no.3
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• pp.31-40
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• 2020

The structure and dynamics of multiple shock waves are studied numerically using a finite volume solver for a model with nozzle exit Mach number of 1.75. At first, the shock variation based on images were analyzed using a Matlab program then later to the wall static pressure variation. The amplitude and frequency variation for multiple shock waves are analyzed. The cross-correlation between the shock location suggests that the first and the second shocks are well correlated while the other shocks show a phase lag in the oscillation characteristics. The rms values of pressure fluctuations are maximum at the shock locations while the other parts in the flow exhibit a lower value os standard deviation.

• Journal of the Korean Society of Visualization
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• v.18 no.2
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• pp.51-58
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• 2020

Research on shock structures of supersonic jet through visualization experiments in low-pressure environment have not been actively conducted. Therefore, in this study, shock waves and supersonic jets were analyzed and compared by numerical analysis and Schlieren technique at low-pressure. Schlieren technique is commonly used to visualize the shock waves generated by density gradient as interferometric methods. Pressure ratio of entrance and ambient was set around 4 to observe moderate under-expanded jet. For validation of experimental and numerical results, the shock structure and frequency were compared. In the case of ST and C nozzle, the results were shown that the difference of shock cell distance was within 10%. The Mach number gradually decreased due to energy reduction, and the error rate was within 7%. D nozzle was not fitted to be observing the shock structure. Because the interface between rarefaction fan and supersonic jet was ambiguous and oscillating phemenoma occurred at end of jet, the supersonic jet in low ambient pressure was observed and analyzed.

• Journal of the Korean Society of Propulsion Engineers
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• v.5 no.3
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• pp.25-32
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• 2001

A passive control method of the interaction between a weak normal shock-wave and a turbulent boundary-layer was simulated using two-dimensional Navier-Stokes computations. The inflow Mach number just upstream of the normal shock wave was 1.33. A porous plate wall having a cavity underneath was used to control the shock-wave/turbulent boundary-layer interaction. The flows through the porous holes and inside the cavity were investigated to get a better understanding of the flow physics involved in this kind of passive control method. The present computations were validated by some recent wind tunnel tests. The results showed that downstream of the rear leg of the $\lambda$-shock wave the main stream inflows into the cavity, but upstream of the rear leg of the $\lambda$-shock wave the flow proceeds from the cavity toward to the main stream. The flow through the porous holes did not choke fur the present shock/boundary layer interaction.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.38 no.9
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• pp.856-861
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• 2010

In this work, experiments on hybrid sounding rocket were conducted to investigate the aerodynamic characteristics and analyze longitudinal static stability. Tests were performed on 1/10 scale models of sounding rocket through Mach number ranging from 1.75 to 2.5 and for angle of attack from $0^{\circ}$ to $6^{\circ}$. Aerodynamic forces and moments were measured by means of a 4 component internal balance. With measured forces and moments, static stability characteristics of rocket were calculated. Tests were made for three models with different length to determine the effect of body length. The visualization of shock waves was carried out by Schlieren optical system to observe variations of shock waves with Mach number and angle of attack.

• Transactions of the Korean Society of Mechanical Engineers B
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• v.38 no.3
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• pp.219-225
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• 2014

In this study, to find the characteristics of the oscillation of a terminating shock wave in a transonic airfoil flow with non-equilibrium condensation, a NACA00-12,14,15 airfoil flow with non-equilibrium condensation is investigated through numerical analysis of TVD scheme. Transonic free stream Mach number of 0.81-0.90 with the variation of stagnation relative humidity and airfoil thickness is tested. For the free stream Mach number 0.87 and attack angle of ${\alpha}=0^{\circ}$, the increase in stagnation relative humidity attenuates the strength of the terminating shock wave and inactivates the oscillation of the terminating shock wave. For the case of $M_{\infty}=0.87$ and ${\phi}_0=60%$, the decreasing rate in the frequency of the shock oscillation caused by non-equilibrium condensation to that of ${\phi}_0=30%$ amounts to 5%. Also, as the stagnation relative humidity gets larger, the maximum coefficient of drag and the difference between the maximum and minimum in $C_D$ become smaller. On the other hand, as the thickness of the airfoil gets larger, the supersonic bubble size becomes bigger and the oscillation of the shock wave becomes higher.

• Proceedings of the KSME Conference
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• 2002.08a
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• pp.739-742
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• 2002

The supersonic, dual, coaxial jet impinging upon a vertical flat plate has recently been applied to a variety of industrial manufacturing processes, since it has several advantages over a conventional supersonic impinging jet. In the present study, experimentation is carried out to investigate the effects of the impinging angle of the annular flow and the design Mach number on the flow field formed over the vertical flat plate. A convergent-divergent nozzle is used to obtain the inner jet flow, its design Mach number being changed between $1.0\;and\;2.0$. The outer annular nozzle has a constant area of the Mach number of 1.0, and its impinging angle of $0^{\circ}\;and\;20^{\circ}$. The primary jet pressure ratio is changed in the range from 6.0 to 10.0 and for the annular flow, the assistant jet pressure ratio is changed from 1.0 to 4.0. The distance between the dual, coaxial nozzle and flat plate is also changed. Detailed pressure measurements are conducted along the axis of the jet and on the flat plate as well. The impinging coaxial Jet flows are visualized using the Schlieren and Shadow optical methods. The results show that the flow field on the plate is not strongly dependent only on the primary and assistant pressure ratios but also the impinging angle of the annular nozzle.