• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2012.05a
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• pp.282-289
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• 2012

The IR Signature and radiative base heating from an aircraft plume have been important factors for aircraft survivability in modern battle fields. In order to enhance the aircraft survivability and reduce the base heating, infrared signatures emitted from an aircraft exhaust plume should be determined. In this work, therefore, IR signatures and radiative base heating characteristics are examined in the plume exhausted from the aircraft with operating at altitude of 5km in M=0.9 and 1.6, respectively. As a result, it is found that the particular wavelength IR signature has high spectral characteristics because of $H_2O$ and $CO_2$ gases in the plume, and the radiative heat flux coming into the base plane increases with higher Mach number and shorter distance.

• Journal of the Korean Society of Propulsion Engineers
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• v.9 no.3
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• pp.85-91
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• 2005

Radiative heating of a liquid rocket base plane due to plume emission is numerically investigated. Calculation of flow and temperature fields around rocket nozzle precedes and thereby realistic plume shape and temperature distribution inside the plume are obtained. Based on the calculated temperature field, radiative transfer equation is solved by discrete ordinate method. With the sample rocket plume, the averaged radiative heat flux reaching the base plane is calculated about 5 kw/m$^{2}$ at the flight altitude of 10.9 km. This value is small compared with radiative heat flux caused by constant-temperature (1500 K) plume emission, but it is not negligibly small. At higher. altitude (29.8km), view factor between the base plane and the exhaust plume is increased due to the increased expansion angle of the plume. Nevertheless, the radiative heating disappears since the base plane is heated to high temperature (above 1000 K due to convective heat transfer.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.42 no.8
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• pp.640-647
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• 2014

Infrared signature of aircraft exhaust plume is the critical factor for aircraft survivability. To improve the military aircraft survivability, the accurate prediction of infrared signature for the propulsion system is needed. The numerical analysis of thermal fluid field for nozzle inflow, free stream flow, and plume region is conducted by using the in-house code. Weighted Sum of Gray Gases Model based on Narrow Band with regrouping is adopted to calculate the spectral infrared signature emitted from aircraft exhaust plume. The accuracy and reliability of the developed code are validated in the one-dimensional band model. It is found that the infrared radiant intensity is relatively more strong in the plume through the analysis, the results show the different characteristic of the spectral infrared signature along the temperature, the partial pressure, and the species distribution. The continuous spectral radiant intensity is shown near the nozzle exit due to the emission from the nozzle wall.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2017.05a
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• pp.596-597
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• 2017

In this study, the measurements of infrared (IR) signature were carried out using a small scale engine with the variation of the engine performance and target positions in the exhaust plume. The operating conditions of the engine were kept constant for each test, and the measured positions were sapced at refular intervals from the nozzle exit. The measured IR signature was calibrated by using a blackbody. The results of infrared signature measurements are shown in three bands for analysis of spectral characteristics. As the engine performance decreased and the distance from the nozzle exit increased, the IR signature decreased and the level of decrease varied according to the bands.

• Journal of the Korean Society of Propulsion Engineers
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• v.9 no.1
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• pp.35-45
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• 2005

Numerical study is carried out to investigate the effects of chemistry and thermal radiation on the rocket plume flow field at various altitudes. Navier-Stokes equations for compressible flows were solved by a fully-implicit TVD code based on the finite volume method. An infinitely fast chemistry module for hydrocarbon mixture with detailed thermo-chemical properties and a thermal radiation module for optically thick media were incorporated with the fluid dynamics code. The plume flow fields of a kerosene-fueled rocket flying at Mach number zero at sea-level, 1.16 at altitude of 5.06 km and 2.90 at 17.34 km were numerically analyzed. Results showed the plume structures at different altitude conditions with the effects of chemistry and radiation. It is understood that the excess temperature by the chemical reactions in the exhaust gas may not be ignored in the view point of propulsion performance and thermal protection of the rocket base, especially at higher altitude conditions.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.857-862
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• 2011

3 dimensional turbine exhaust gas flow was simplified to an axisymmetrical flow and calculated with detailed chemistry models. GRI 35 species-217 reaction step model and simplified 11 species 15 reaction model was applied to the secondary reaction of the turbine exhaust gas and compared. All the model captured the secondary combustion on the base region, and the temperature was 600K higher than that without turbine exhaust gas. This means the local temperature of the base can be higher in the case of real 3 dimensional flow. The simplified model show the similar results to the GRI detailed chemistry model although the former affected the engine plume structure slightly.

• 한국전산유체공학회:학술대회논문집
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• 1999.11a
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• pp.71-76
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• 1999

The base flow regions of a three-body sounding rocket containing multiple exhaust plumes were numerically investigated in three dimensions for a free stream Mach number of 2.7 at flight altitude 18.5 km. The flowfields were calculated using the full compressible Navier-Stokes equations with an one-equation turbulence model of Baldwin-Earth. The present calculations were executed based upon a chemically frozen, single perfect gas model assumption. Due to the symmetry of the three-body rocket of each single nozzle, only one fourth of the computational domain was considered for the analysis. The results indicate that a babe heating effect is not considerable due to the small expansion of the plumes. In the base, however, a low speed recirculating flow dominates the region.

• Journal of computational fluids engineering
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• v.18 no.2
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• pp.66-71
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• 2013

In this paper, the supersonic flows around space launch vehicles have been numerically simulated by using a 3-D RANS flow solver. The focus of the study was made for investigating plume-induced flow separation(PIFS). For this purpose, a vertex-centered finite-volume method was utilized in conjunction with 2nd-order Roe's FDS to discretize the inviscid fluxes. The viscous fluxes were computed based on central differencing. The Spalart-Allmaras model was employed for the closure of turbulence. The Gauss-Seidel iteration was used for time integration. To validate the flow solver, calculation was made for the 0.04 scale model of the Saturn-5 launch vehicle at the supersonic flow condition without exhaust plume, and the predicted results were compared with the experimental data. Good agreements were obtained between the present results and the experiment for the surface pressure coefficient and the Mach number distribution inside the boundary layer. Additional calculations were made for the real scale of the Saturn-5 configuration with exhaust plume. The flow characteristics were analyzed, and the PIFS distances were validated by comparing with the flight data. The KSLV-1 is also simulated at the several altitude conditions. In case of the KSLV-1, PIFS was not observed at all conditions, and it is expected that PIFS is affected by the nozzle position.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.45 no.9
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• pp.766-774
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• 2017

When a plume flow exhausted from a lunar lander descent engine impinges on the lunar surface, regolith particles on the lunar surface will be dispersed due to a plume-surface interaction. If the dispersed particles collide with the lunar lander, some adverse effects such as a performance degradation can be caused. Thus, this study tried to predict the plume flow behaviors using the CFD methods. A nozzle inside region was analyzed by a continuum flow model based on the Navier-Stokes equations while the plume behaviors of the outside nozzle was performed by comparing and analyzing the individual results using the continuum flow model and the DSMC method. As a result, it was possible to establish an optimum procedure of the plume analysis for the lunar lander descent engine in the vacuum condition. In the future, it is expected to utilize the present results for the development of the Korean lunar lander.

• 한국전산유체공학회:학술대회논문집
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• 2004.03a
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• pp.146-153
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• 2004

The Numerical study has been carried out to investigate the effects of chemical reaction and thermal radiation on the rocket plume flow-field at various altitudes. The theoretical formulation is based on the Navier-Stokes equations for compressible flows along with the infinitely fast chemistry and thermal radiation. The governing equations were solved by a finite volume fully-implicit TVD(Total Variation Diminishing) code which uses Roe's approximate Riemann solver and MUSCL(Monotone Upstream-centered Schemes for Conservation Laws) scheme. LU-SGS (Lower Upper Symmetric Gauss Seidel) method is used for the implicit solution strategy. An equilibrium chemistry module for hydrocarbon mixture with detailed thermo-chemical properties and a thermal radiation module for optically thin media were incorporated with the fluid dynamics code. In this study, kerosene-fueled rocket was assumed operating at O/F ratio of 2.34 with a nozzle expansion ratio of 6.14. Flight conditions considered were Mach number zero at ground level, Mach number 1.16 at altitude 5.06km and Mach number 2.9 at altitude 17.34km. Numerical results gave the understandings on the detailed plume structures at different altitude conditions. The diffusive effect of the thermal radiation on temperature field and the effect of chemical recombination during the expansion process could be also understood. By comparing the results from frozen flow and infinitely fast chemistry assumptions, the excess temperature of the exhaust gas resulting from the chemical recombination seems to be significant and cannot be neglected in the view point of performance, thermal protection and flow physics.