scholarly journals Investigations on the Influence of the In-Stream Pylon and Strut on the Performance of a Scramjet Combustor

2014 ◽  
Vol 2014 ◽  
pp. 1-10 ◽  
Author(s):  
Hao Ouyang ◽  
Weidong Liu ◽  
Mingbo Sun
Keyword(s):  
Mach Number ◽  
Pressure Loss ◽  
Total Pressure ◽  
Wall Pressure ◽  
Total Pressure Loss ◽  
Research Results ◽  
Scramjet Combustor ◽  
Total Temperature

The influence of the in-stream pylon and strut on the performance of scramjet combustor was experimentally and numerically investigated. The experiments were conducted with a direct-connect supersonic model combustor equipped with multiple cavities. The entrance parameter of combustor corresponds to scramjet flight Mach number 4.0 with a total temperature of 947 K. The research results show that, compared with the scramjet combustor without pylon and strut, the wall pressure and the thrust of the scramjet increase due to the improvement of mixing and combustion effect due to the pylon and strut. The total pressure loss caused by the strut is considerable whereas pylon influence is slight.

Numerical investigations on the hydrogen jet pressure variations in a strut based scramjet combustor

2021 ◽  
Vol ahead-of-print (ahead-of-print) ◽  
Author(s):  
Jeyakumar Suppandipillai ◽  
Jayaraman Kandasamy ◽  
R. Sivakumar ◽  
Mehmet Karaca ◽  
Karthik K.
Keyword(s):  
Shock Wave ◽  
Supersonic Flow ◽  
Pressure Loss ◽  
Total Pressure ◽  
Injection Pressure ◽  
Total Pressure Loss ◽  
Normal Shock ◽  
Content Type ◽  
Normal Shock Wave ◽  
Scramjet Combustor

Purpose This paper aims to study the influences of hydrogen jet pressure on flow features of a strut-based injector in a scramjet combustor under-reacting cases are numerically investigated in this study. Design/methodology/approach The numerical analysis is carried out using Reynolds Averaged Navier Stokes (RANS) equations with the Shear Stress Transport k-ω turbulence model in contention to comprehend the flow physics during scramjet combustion. The three major parameters such as the shock wave pattern, wall pressures and static temperature across the combustor are validated with the reported experiments. The results comply with the range, indicating the adopted simulation method can be extended for other investigations as well. The supersonic flow characteristics are determined based on the flow properties, combustion efficiency and total pressure loss. Findings The results revealed that the augmentation of hydrogen jet pressure via variation in flame features increases the static pressure in the vicinity of the strut and destabilize the normal shock wave position. Indeed, the pressure of the mainstream flow drives the shock wave toward the upstream direction. The study perceived that once the hydrogen jet pressure is reached 4 bar, the incoming flow attains a subsonic state due to the movement of normal shock wave ahead of the strut. It is noticed that the increase in hydrogen jet pressure in the supersonic flow field improves the jet penetration rate in the lateral direction of the flow and also increases the total pressure loss as compared with the baseline injection pressure condition. Practical implications The outcome of this research provides the influence of fuel injection pressure variations in the supersonic combustion phenomenon of hypersonic vehicles. Originality/value This paper substantiates the effect of increasing hydrogen jet pressure in the reacting supersonic airstream on the performance of a scramjet combustor.

Aerodynamic Performance of Suction-Side Gill Region Film Cooling

2010 ◽  
Vol 132 (3) ◽  
Author(s):  
Justin Chappell ◽  
Phil Ligrani ◽  
Sri Sreekanth ◽  
Terry Lucas ◽  
Edward Vlasic
Keyword(s):  
Mach Number ◽  
Film Cooling ◽  
Pressure Loss ◽  
Total Pressure ◽  
Suction Side ◽  
Total Pressure Loss ◽  
Pressure Loss Coefficient ◽  
Blowing Ratio ◽  
Exit Mach Number ◽  
Total Pressure Loss Coefficient

The performance of suction-side gill region film cooling is investigated using the University of Utah transonic wind tunnel and a simulated turbine vane in a two-dimensional cascade. The effects of film cooling hole orientation, shape, and number of rows, and their resulting effects on the aerodynamic losses, are considered for four different hole configurations: round axial (RA), shaped axial (SA), round radial (RR), and round compound (RC). The mainstream Reynolds number based on axial chord is 500,000, exit Mach number is 0.35, and the tests are conducted using the first row of holes, or both rows of holes at blowing ratios of 0.6 and 1.2. Carbon dioxide is used as the injectant to achieve density ratios of 1.77–1.99 similar to values present in operating gas turbine engines. Presented are the local distributions of total pressure loss coefficient, local normalized exit Mach number, and local normalized exit kinetic energy. Integrated aerodynamic losses (IAL) increase anywhere from 4% to 45% compared with a smooth blade with no film injection. The performance of each hole type depends on the airfoil configuration, film cooling configuration, mainstream flow Mach number, number of rows of holes, density ratio, and blowing ratio, but the general trend is an increase in IAL as either the blowing ratio or the number of rows of holes increase. In general, the largest total pressure loss coefficient Cp magnitudes and the largest IAL are generally present at any particular wake location for the RR or SA configurations, regardless of the film cooling blowing ratio and number of holes. The SA holes also generally produce the highest local peak Cp magnitudes. IAL magnitudes are generally lowest with the RA hole configuration. A one-dimensional mixing loss correlation for normalized IAL values is also presented, which matches most of the both rows data for RA, SA, RR, and RC hole configurations. The equation also provides good representation of the RA, RC, and RR first row data sets.

The Design, Analysis and Test of a Unique Gas Turbine Scroll for the F-35, Joint Strike Fighter

2006 ◽  
Author(s):  
Hakan Aksoy ◽  
Stony W. Kujala ◽  
Craig W. McKeever ◽  
Ly D. Nguyen
Keyword(s):  
Mach Number ◽  
Pressure Loss ◽  
Total Pressure ◽  
Total Pressure Loss ◽  
Flow Distortion ◽  
Cfd Simulations ◽  
Flow Angle ◽  
Auxiliary Power Unit ◽  
Auxiliary Power ◽  
Stringent Performance

The design of the APU (Auxiliary Power Unit) for the F-35 JSF (Joint Strike Fighter) focused on minimizing size and weight while meeting stringent performance goals. To help realize that goal, a unique turbine scroll was designed. The scroll design delivers air from the combustor to the turbine inlet with minimal loss and flow distortion while minimizing design space. CFD (Computational Fluid Dynamics) results of scroll total pressure loss and exit peripheral distribution of total pressure, Mach number, and flow angle are presented. Rig tests were utilized for measuring and validating the computed total pressure and Mach number distributions around the periphery of the scroll exit. Comparisons of the CFD simulations and test data indicate strong correlation in values of average total pressure loss, local total pressure loss and Mach number around the exit periphery.

scholarly journals A Compact Diffuser System for Annular Combustors

1983 ◽  
Author(s):  
R. C. Adkins ◽  
J. O. Yost
Keyword(s):  
High Pressure ◽  
Mach Number ◽  
Pressure Loss ◽  
Total Pressure ◽  
Aerodynamic Performance ◽  
Total Pressure Loss ◽  
Pressure Recovery ◽  
Analytical Relationship ◽  
Exit Mach Number ◽  
Aerodynamic Simulation

Airflow tests have been conducted on an aerodynamic simulation of a combustor with pre-diffuser of compact configuration. The inlet Mach number throughout the tests was 0.35. The configuration was successful because of the attainment of a high pressure recovery, (Cp = 0.80), coupled with an exceptionally low total pressure loss (λ = 0.04). A useful analytical relationship is derived between the aerodynamic performance of combustor, compressor exit Mach number and diffuser performance.

Multipoint design optimization for a controlled diffusion airfoil compressor cascade

2020 ◽  
Vol 234 (11) ◽  
pp. 2143-2159
Author(s):  
Song Huang ◽  
Chengwu Yang ◽  
Ge Han ◽  
Shengfeng Zhao ◽  
Xingen Lu
Keyword(s):  
Mach Number ◽  
Turbulence Intensity ◽  
Pressure Loss ◽  
Total Pressure ◽  
Separation Bubble ◽  
Total Pressure Loss ◽  
Laminar Separation Bubble ◽  
Laminar Separation ◽  
Controlled Diffusion ◽  
Wide Range

Under high altitude and low Reynolds number conditions, the aerodynamic performance of compressor cascades deteriorates drastically. In this paper, an optimally designed system combining class-shape-transformation method, S1 surface flow solver and whale optimization algorithm was established to achieve for a controlled diffusion airfoil, called MANGHH. The aim of this work is to improve our understanding of the loss mechanism for the original cascade and optimal cascade under different inflow conditions. The study shows that the total pressure loss of the optimal cascade at an angle of attack of −4°, 0°, and 6° decreases by 55.9%, 16.1%, and 16.3%, respectively, compared with the original controlled diffusion airfoil. The range of the available low loss incidence improves significantly. At different incidences, the optimal cascade moves the blade loading forward compared with that of the original controlled diffusion airfoil while reducing the growth rate of the boundary layer thickness, eliminating a wide range of flow separations. The optimal cascade reduces the total pressure loss mainly by reducing trailing edge mixing loss compared with that of the original controlled diffusion airfoil. Under different inlet Mach number conditions, a laminar separation bubble appears on the suction surface of the original controlled diffusion airfoil. As the inlet Mach number increases, the position of the laminar separation bubble moves slightly upstream, while the length and depth of the laminar separation bubble increase. Fortunately, the total pressure loss of the optimal cascade decreases significantly compared with that of the original controlled diffusion airfoil. Under different incoming turbulence intensity conditions, the total pressure loss of the optimal cascade is always lower than that of the original controlled diffusion airfoil. As the incoming turbulence intensity increases, the total pressure loss of the original controlled diffusion airfoil decreases first and then increases. However, the total pressure loss of the optimal cascade increases with increasing incoming turbulence intensity due to the improvement of the turbulence dissipation capacity.

Effects of pylon geometry on mixing enhancement in a scramjet pylon-cavity flameholder

2020 ◽  
Vol 124 (1278) ◽  
pp. 1262-1280
Author(s):  
A. Oamjee ◽  
R. Sadanandan
Keyword(s):  
Pressure Loss ◽  
Total Pressure ◽  
Supersonic Combustion ◽  
Total Pressure Loss ◽  
Absolute Pressure ◽  
Mixing Efficiency ◽  
Mixing Enhancement ◽  
Scramjet Combustor ◽  
Fuel Jet ◽  
Jet Penetration

ABSTRACTNumerical investigation of the effect of pylon geometry within a pylon-cavity aided Supersonic Combustion Ramjet (SCRAMJET) combustor on mixing enhancement, flame-holding capability, fuel jet penetration and total pressure loss are conducted in the current study. RANS equations for compressed real gas are solved by coupled, implicit, second-order upwind solver. A two-equation SST model is used for turbulence modelling. Validation of the computational model is performed with the help of experimental data collected using surface pressure taps, Schlieren flow visualisation and particle image velocimetry (PIV). The study uses four distinct pylon geometry cases, which include the baseline geometry. Sonic injection of hydrogen fuel through a 1mm diameter hole at 2mm downstream of the pylon rear face along the axis of the test section floor is performed for every case. A crossflow of Mach number 2.2 at four bar absolute pressure and standard atmospheric temperature is maintained. A comparative study of mixing efficiency, total pressure loss, fuel jet penetration and fuel plume area fraction for the different cases evaluate the mixing performance. The simulations show that the Pylon 2 case gives a significant improvement in the performance parameters compared to the other geometries. It is observed that mixing efficiency and fuel jet penetration capability of the system are highly dependent on the streamwise vortex within the flameholder.

Some Preliminary Results for Conical Diffusers with high Subsonic entry mach Numbers

1966 ◽  
Vol 8 (4) ◽  
pp. 384-391 ◽  
Author(s):  
J. L. Livesey ◽  
T. Hugh
Keyword(s):  
Mach Number ◽  
Pipe Flow ◽  
Pressure Loss ◽  
Total Pressure ◽  
Static Pressure ◽  
Loss Coefficient ◽  
Total Pressure Loss ◽  
Preliminary Results ◽  
Mach Numbers ◽  
High Mach

Some preliminary results are presented for the variation of total pressure loss coefficient with entry Mach number for conical diffusers. Included angles of 5, 8 and 12°, area ratios from 2 to 16 and entry lengths from 0 to 65 diameters are considered. In these initial tests the junction between the parallel entry pipe and the cone is sharp. The variation of Mach number and total pressure loss in the entry pipe is predicted simply using measured friction factors and typical velocity profiles. These calculations give the entry conditions to the diffusers in the absence of the diffusers. The effect of the junction of entry pipe and diffuser on the entry pipe flow is thus correctly attributed to the diffuser in the assessment of its total pressure loss. In addition a correctly specified mean entry Mach number is obtained. These considerations enable the losses in diffuser-pipe conjunctions of high Mach number to be logically analysed from separate pipe flow and diffuser data. The results obtained are unusual and indicate that previous assessments of diffuser performance at high Mach numbers have been confused by incorrect consideration of the effect of the diffuser entry pipe conjunction and incorrect specification of the effective mean entry Mach number. It is concluded that further experimental results are needed for developing compressible flows in constant area ducts in order that the present preliminary results may be made more precise. The momentum equation analysis in terms of suitable mean values is presented briefly. Previous diffuser results for low subsonic entry Mach numbers are considered briefly in comparison. Serious errors are shown to be present in these early results. Typically the errors originate in the definition of loss coefficient and static pressure efficiency and the use of the mass derived mean concept. In some cases the quoted static pressure efficiencies imply a decrease in entropy for the diffuser flow at low area ratios.

scholarly journals The Combustion Characteristics of Double Ramps in a Strut-Based Scramjet Combustor

Energies ◽  
2021 ◽  
Vol 14 (4) ◽  
pp. 831
Author(s):  
A. Antony Athithan ◽  
S. Jeyakumar ◽  
Norbert Sczygiol ◽  
Mariusz Urbanski ◽  
A. Hariharasudan
Keyword(s):  
Pressure Loss ◽  
Total Pressure ◽  
Computational Study ◽  
Flame Temperature ◽  
Total Pressure Loss ◽  
Reacting Flow ◽  
Hydrogen Distribution ◽  
Spatial Direction ◽  
Shock Interactions ◽  
Scramjet Combustor

This paper focuses on the influence of ramp locations upstream of a strut-based scramjet combustor under reacting flow conditions that are numerically investigated. In contrast, a computational study is adopted using Reynolds Averaged Navier Stokes (RANS) equations with the Shear Stress Transport (SST) k-ω turbulence model. The numerical results of the Deutsches Zentrum für Luft- und Raumfahrt or German Aerospace Centre (DLR) scramjet model are validated with the reported experimental values that show compliance within the range, indicating that the adopted simulation method can be extended for other investigations as well. The performance of the ramps in the strut-based scramjet combustor is analyzed based on parameters such as wall pressures, combustion efficiency and total pressure loss at various axial locations of the combustor. From the numerical shadowgraph, more shock interactions are observed upstream of the strut injection region for the ramp cases, which decelerates the flow downstream, and additional shock reflections with less intensity are also noticed when compared with the DLR scramjet model. The shock reflection due to the ramps enhances the hydrogen distribution in the spatial direction. The ignition delay is noticed for ramp combustors due to the deceleration of flow compared to the baseline strut only scramjet combustor. However, a higher flame temperature is observed with the ramp combustor. Because more shock interactions arise from the ramps, a marginal increase in the total pressure loss is observed for ramp combustors when compared to the baseline model.

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