Plume Induced Flow Separation Research Articles

Numerical analysis on thermal environment of reusable launch vehicle during supersonic retropropulsion

Developing reusable launch vehicles (RLV) is an effective way to significantly reduce space transportation costs. However, in the return stage of the vehicle, especially the supersonic retropropulsion (SRP), the first stage is partially submerged in the hot exhaust plume and faces severe convective heating at the base plate and sidewalls. As a rule of thumb, the layout of nozzles at the rocket base significantly affects the heat load. In order to evaluate the influence of heat load under different nozzle configurations during SRP (33 km–22 km), the thermal environment of RLV was studied through numerical investigation for three typical nozzle configurations. Using Reynolds-averaged Navier–Stokes equations (RANS)-based computational methods and the k−ω SST turbulence model, the flowfield is characterized at different trajectory points during retropropulsion. Results showed that the high heat flux is mainly concentrated in the baseplate's center area and the sidewall's near base area during SRP. In general, activating engines in the radial position is beneficial for the thermal protection system of a reusable launch vehicle while the thrust is identical. Base heating primarily results from high-temperature counterflows due to plume interactions, while sidewall heating is predominantly a result of high-temperature recirculation zones generated by plume-induced flow separation. During the supersonic retropropulsion phase, the 3-nozzle-180°, a non-closed nozzle configuration, has significantly lower heat loads on the baseplate and the sidewall than that of the 4-nozzle-90° and the 3-nozzle-120°. In addition, reducing the number of engines is also an effective solution under the same conditions of thrust. Studying the layout of the working engine without changing the original layout has important guiding significance for improving the reusability and safety factor of the RLV.

Computational investigation of axi-symmetric plume induced flow separation

The flow over a rocket may result in plume induced flow separation (PIFS) depending on flight conditions and nozzle-to-ambient pressure ratios. In the current study, the flow structure over a generic rocket model has been investigated to determine the conditions that would results in PIFS over the rocket body. The rocket used in the current study has a tangent-ogive nose with a length of 25.4 cm (nose length/diameter ∼ 4.0), followed by a 57.2-cm long cylindrical rocket body (overall length/diameter ∼ 13.0). A 45° conical nozzle with an exit Mach number of 2.7 (nozzle expansion ratio, ε ∼1.786) was used to simulate the plume. The total temperature and the total pressure in the combustion chamber were kept as 3589 K (6000 °F), and 5.86 MPa ab (850 psia). The flow conditions investigated including flight Mach numbers from 0.7 to 2.0, at five different altitudes; 0, 3.05, 6.10, 9.14 and 12.2 km (0, 10,000, 20,000, 30,000 and 40,000 feet). Based on the validation study using various eddy-viscosity turbulence models, the k-ε RNG turbulence model was selected for the parametric study. PIFS is not observed at Mach numbers less than 1.0 up to altitudes including 9.1 km (30,000 ft), and occurs only between Mach 1.0 and 2.0. The Mach number range where PIFS is observed increases with increasing altitude. However at 12.2 km (40,000 ft) PIFS is observed at every Mach number considered, including subsonic range.

Effect of freestream-jet plume interaction on aerodynamic coefficients with different flared aft-bodies

Experiments were carried out on a blunt-cone-cylinder geometry with various flared aft body configurations to understand the effect freestream (M=5)-jet plume (M=2.65) interaction on aerodynamic coefficients. Supersonic jet plume of varying pressure ratios (Pj/P∞=0–130) were simulated and the model angle of attack (α) is varied from 0 to 4°. The complex freestream-jet plume interaction was captured through Schlieren flow visualization while the forces and moments are measured through conventional strain gauge balance. The introduction of supersonic jet plume resulted in reduction of aerodynamic coefficients in all the cases. The change in aerodynamic coefficients depends primarily on the extent of Plume Induced Flow Separation (PIFS) due to freestream-jet plume interaction and the aft body configurations. A reduction of 37–46% in axial force coefficient (CA), 20–23% in normal force coefficient (CN) and 31–72% in pitching moment coefficient (Cm) were observed due to freestream-jet plume interaction. In cylindrical aft body, separation due to freestream-jet plume interaction was found to be dominant whereas it is less pronounced in flared aft body configurations. Among the tested configurations 14and16° flared aft body configurations show more stability compared to cylindrical and other flared aft body configurations under jet plume-on and off conditions.

Computational framework for Launch, Ascent, and Vehicle Aerodynamics (LAVA)

The Launch Ascent and Vehicle Aerodynamics (LAVA) framework, developed at NASA Ames Research Center, is introduced. This technology originated from addressing some of the key challenges that were present during the re-design of the launch infrastructure at Kennedy Space Center. The solver combines Computational Fluid Dynamics (CFD) capabilities with auxiliary modules, such as Conjugate Heat Transfer (CHT) and Computational Aero-Acoustics (CAA). LAVA is designed to be grid agnostic, i.e., it can handle block-structured Cartesian, generalized curvilinear overset and unstructured polyhedral grids either as stand-alone mode or by coupling different grid types through an overset interface. A description of the spatial discretizations utilized for each grid type, along with the available explicit and implicit time-stepping schemes, is provided. The overset grid coupling procedure for Cartesian and unstructured mesh types, as well as the CHT and CAA capabilities is discussed in some detail. Several NASA mission related applications are also presented to demonstrate the capabilities for large-scale applications, such as pressure, thermal and acoustic analyses of the geometrically complex launch environment, steady and unsteady ascent aerodynamics, plume-induced flow separation analyses of heavy lift launch vehicles and aeroacoustic applications. In addition, two validation cases related to NASA aeronautics applications are presented: the 1st AIAA Sonic Boom Prediction Workshop test cases and a computational study of slat noise.

Skylon Aerospace Plane and Its Aerodynamics and Plumes

The Skylon concept incorporates the highly innovative synergetic air-breathing rocket engine concept that has the potential to revolutionize the mode of propulsion for transportation of medium-weight payloads to low Earth orbits. An independent partial assessment is provided of the technical viability of the Skylon concept. Pressure lift and drag coefficients derived from Euler simulations for unpowered flight compare very well and fairly well, respectively, with those from engineering methods. The engineering-method coefficients for powered flight are increasingly less acceptable as the freestream Mach number is increased beyond 8.5 because these methods did not account for the increasing favorable (in terms of pressure forces) effect of underexpanded rocket engine plumes on the aft fuselage. At Mach numbers greater than 8.5, the thermal environment around the aft fuselage is a known unknown: a potential design and/or performance risk issue. The adverse effects of shock waves on the aft fuselage and plume-induced flow separation are other potential risks. A preliminary design of Skylon requires the judicious use of a combination of engineering methods, advanced methods based on required physics or analytical fidelity, test data, and independent assessments. The demonstration of a synergetic air-breathing rocket-engine-powered experimental aerospace plane calls for the second revival of the Aerospace Plane Program.

우주발사체의 플룸에 따른 유동박리 현상에 대한 수치적 연구

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.

USM3D Simulations of Saturn V Plume-Induced Flow Separation

Covers advancements in spacecraft and tactical and strategic missile systems, including subsystem design and application, mission design and analysis, materials and structures, developments in space sciences, space processing and manufacturing, space operations, and applications of space technologies to other fields.

USM3D Simulations of Saturn V Plume-Induced Flow Separation

Covers advancements in spacecraft and tactical and strategic missile systems, including subsystem design and application, mission design and analysis, materials and structures, developments in space sciences, space processing and manufacturing, space operations, and applications of space technologies to other fields.

Investigation on the Characteristics of Plume-Induced Flow Separation and Wall Heat Transfer

Plume induced flow separation phenomena and the convective heat flux to the base surface were investigated according to the base configurations. The flight condition of a vehicle was the flight Mach number of 5 and the plume Mach number of 3. A twodimensional axisymmetric Navier-Stokes solver containing a k-�& SST turbulence model was used to solve the launch vehicle configuration in which a plume was incorporated. The results acquired from plume induced flow separation computations were compared to results acquired through experimentation as a means to validate the codes utilized in these investigations. Three after-body configurations were simulated in order to investigate the effects experienced by the launch. The three configurations examined in this study were designated as: (1) boat-tail, (2) straight, and (3) flare. The boat tail after body configuration caused plume induced flow separation, while the flare configuration suppressed plume induced flow separation. Additionally, the flare type base reduced the convective heat flux to the base surface. The maximum heat flux occurred around the nozzle exit due to the plume’s rapid expansion.

Analysis of Linear Aerospike Plume-Induced X-33 Base-Heating Environment

Covers advancements in spacecraft and tactical and strategic missile systems, including subsystem design and application, mission design and analysis, materials and structures, developments in space sciences, space processing and manufacturing, space operations, and applications of space technologies to other fields.

Numerical study of plume-induced flow separation

Laminar, Mach 26 flow past a blunt-nosed cone with a spherical after-body experiencing axial mass injection from its base was numerically studied in two calculations. Mass injection rate for both calculations was 0·15 lb/sec and Reynolds numbers relative to base diameter were 735 and 4150. Numerical computations were performed with a code (AFTON 2A) which solves the time-dependent Navier-Stokes equations for systems with axial symmetry and mass injection. Steady flow in both cases was approached asymptotically with time from initial flow fields derived from simple inviscid theory. At the lower Reynolds number, the cone and after-body boundary layers remained attached, and the interaction between the incident airstream and the injected plume gases is similar to that of two impinging supersonic streams. By contrast, at the higher Reynolds number the cone and after-body boundary layers both separate, giving rise to a double-vortex pattern and a complex system of shock waves. AFTON 2A results were compared with a theoretical inviscid model of a mass source in a uniform hypersonic stream. In the AFTON 2A calculations, the plume radius is predicted to be much larger, with an increased plume volume caused by heat transfer across the dividing streamline. Although absolute magnitudes of plume-shock and dividing-streamline radii differ from inviscid prediction, ratios of these radii are consistent with the inviscid scaling law; i.e. radii are inversely proportional to the fourth root of the plume drag coefficient. The two calculations also demonstrated a structural difference in the plume inner core. At the lower Reynolds number the inner core was nearly isentropic. By contrast, an amount of heat sufficient to destroy isentropy penetrates the inner core of the plume at the higher Reynolds number.

Flow separation due to jet pluming.

D missile ascent rapidly decreasing ambient pressure causes the rocket exhaust plume diameter to equal and exceed the vehicle length. At sufficiently high altitudes, the plume interacts with the vehicle flowfield and produces a boundary-layer separation that can occur well forward of the vehicle base. Because this phenomenon significantly influences vehicle stability, heating, and stage separation, the conditions at which plume-induced flow separations can occur are naturally of interest. In a series of wind-tunnel experiments, large regions of plume-induced flow separation have been demonstrated by Falanga et al and by Vick et al using cold-flow jet simulation techniques. Since these available data indicate a considerable degree of configuration dependence, they cannot be used readily to predict flow-separation characteristics for arbitrary missile shapes. Furthermore, the adequacy of the jet simulation techniques used to simulate vehicle-plume interaction has not been demonstrated. Additional wind-tunnel tests now have been conducted in order to determine the plume-induced flow-separation pattern to be expected on several typical missile configurations. Furthermore, a flight program using one of the windtunnel configurations has served to check the validity of the jet simulation techniques. The results of these tests are reported in this note.