Numerical Studies on the Performance of Scramjet Combustor with Alternating Wedge-Shaped Strut Injector

2017 ◽  
Vol 34 (1) ◽  
Author(s):  
Gautam Choubey ◽  
K. M. Pandey
Keyword(s):  
Numerical Analysis ◽  
Flow Structure ◽  
Fuel Injection ◽  
Angle Of Attack ◽  
Supersonic Combustion ◽  
Scramjet Combustor ◽  
Numerical Studies ◽  
Scramjet Engine

AbstractNumerical analysis of the supersonic combustion and flow structure through a scramjet engine at Mach 7 with alternating wedge fuel injection and with three angle of attack (

scholarly journals The Effect of Fuel Injection Location on Supersonic Hydrogen Combustion in a Cavity-Based Model Scramjet Combustor

Energies ◽  
2020 ◽  
Vol 13 (1) ◽  
pp. 193 ◽  
Author(s):  
Eunju Jeong ◽  
Sean O’Byrne ◽  
In-Seuck Jeung ◽  
A. F. P. Houwing
Keyword(s):  
Equivalence Ratio ◽  
Fuel Injection ◽  
Flow Direction ◽  
Flame Structure ◽  
Flow Characteristics ◽  
Supersonic Combustion ◽  
Reacting Flow ◽  
Total Enthalpy ◽  
Scramjet Combustor ◽  
Injection Location

Supersonic combustion experiments were performed using three different hydrogen fuel-injection configurations in a cavity-based model scramjet combustor with various global fuel–air equivalence ratios. The configurations tested were angled injection at 15° to the flow direction upstream of the cavity, parallel injection from the front step, and upstream injection from the rear ramp. Planar laser-induced fluorescence of the hydroxyl radical and time-resolved pressure measurements were used to investigate the flow characteristics. Angled injection generated a weak bow shock in front of the injector and recirculation zone to maintain the combustion as the equivalence ratio increased. Parallel and upstream injections both showed similar flame structure over the cavity at low equivalence ratio. Upstream injection enhanced the fuel diffusion and enabled ignition with a shorter delay length than with parallel injection. The presence of a flame near the cavity was determined while varying the fuel injection location, the equivalence ratio, and total enthalpy of the air flow. The flame characteristics agreed with the correlation plot for the stable flame limit of non-premixed conditions. The pressure increase in the cavity for reacting flow compared to non-reacting flow was almost identical for all three configurations. More than 300 mm downstream of the duct entrance, averaged pressure ratios at low global equivalence ratio were similar for all three injection configurations.

scholarly journals Direct Measurements of Skin Friction in Supersonic Combustion Flow Fields

1992 ◽  
Author(s):  
K. M. Chadwick ◽  
D. J. Deturris ◽  
J. A. Schetz
Keyword(s):  
Skin Friction ◽  
Fuel Injection ◽  
Force Measurement ◽  
Tangential Force ◽  
Transverse Component ◽  
Supersonic Combustion ◽  
Quartz Tube ◽  
Hydrogen Fuel ◽  
Sensor Head ◽  
Scramjet Combustor

An experimental investigation was conducted to measure skin friction along the chamber walls of supersonic combustors. A direct force measurement device was used to simultaneously measure an axial and transverse component of the small tangential shear force passing over a non-intrusive floating element. This measurement was made possible with a sensitive piezoresistive deflection sensing unit. The floating head is mounted to a stiff cantilever beam arrangement with deflection due to the flow on the order of 0.00254 mm (0.0001 in). This allowed the instrument to be a non-nulling type. A second gauge was designed with active cooling of the floating sensor head to eliminate non-uniform temperature effects between the sensor head and the surrounding wall. The key to this device is the use of a quartz tube cantilever with piezoresistive strain gages bonded directly to its surface. A symmetric fluid flow was developed inside the quartz tube to provide cooling to the backside of the floating head. Tests showed that this flow did not influence the tangential force measurement. Measurements were made in three separate combustor test facilities. Tests at NASA Langley Research Center consisted of a Mach 3.0 vitiated air flow with hydrogen fuel injection at Pt = 500 psia (3446 kPa) and Tt = 3000 R (1667 K). Two separate sets of tests were conducted at the General Applied Science Laboratory (GASL) in a scramjet combustor model with hydrogen fuel injection in vitiated air at Mach = 3.3, Pt = 800 psia (5510 kPa), and Tt = 4000 R (2222 K). Skin friction coefficients between 0.001–0.005 were measured dependent on the facility and measurement location. Analysis of the measurement uncertainties indicate an accuracy to within ±10–15% of the streamwise component.

Direct Measurements of Skin Friction in Supersonic Combustion Flow Fields

1993 ◽  
Vol 115 (3) ◽  
pp. 507-514 ◽  
Author(s):  
K. M. Chadwick ◽  
D. J. DeTurris ◽  
J. A. Schetz
Keyword(s):  
Skin Friction ◽  
Fuel Injection ◽  
Force Measurement ◽  
Tangential Force ◽  
Transverse Component ◽  
Supersonic Combustion ◽  
Quartz Tube ◽  
Hydrogen Fuel ◽  
Sensor Head ◽  
Scramjet Combustor

An experimental investigation was conducted to measure skin friction along the chamber walls of supersonic combustors. A direct force measurement device was used to measure simultaneously an axial and a transverse component of the small tangential shear force passing over a nonintrusive floating element. This measurement was made possible with a sensitive piezoresistive deflection sensing unit. The floating head is mounted to a stiff cantilever beam arrangement with deflection due to the flow on the order of 0.00254 mm (0.0001 in). This allowed the instrument to be a nonnulling type. A second gage was designed with active cooling of the floating sensor head to eliminate nonuniform temperature effects between the sensor head and the surrounding wall. The key to this device is the use of a quartz tube cantilever with piezoresistive strain gages bonded directly to its surface. A symmetric fluid flow was developed inside the quartz tube to provide cooling to the backside of the floating head. Tests showed that this flow did not influence the tangential force measurement. Measurements were made in three separate combustor test facilities. Tests at NASA Langley Research center consisted of a Mach 3.0 vitiated air flow with hydrogen fuel injection at Pt = 500 psia (3466 kPa) and Tt = 3000 R (1667 K). Two separate sets of tests were conducted at the General Applied Science Laboratory (GASL) in a scramjet combustor model with hydrogen fuel injection in vitiated air at Mach = 3.3, Pt = 800 psia (5510 kPa), and Tt = 4000 R (2222 K). Skin friction coefficients between 0.001–0.005 were measured dependent on the facility and measurement location. Analysis of the measurement uncertainties indicate an accuracy to within ± 10–15 percent of the streamwise component.

Investigation of Performance of Scramjet Combustion Used Kerosene with Clean Air

2012 ◽  
Vol 601 ◽  
pp. 294-298
Author(s):  
Wei Qiang Li ◽  
Fei Teng Luo
Keyword(s):  
Supersonic Combustion ◽  
Experimental Result ◽  
Scramjet Combustor ◽  
Clean Air ◽  
Cavity Design ◽  
Total Temperature ◽  
Scramjet Engine ◽  
Ignition Characteristics ◽  
Flight Regimes ◽  
Inflow Conditions

The aaviation kerosene is a practical candidate fuel for scramjet engine in the flight regimes of Mach number less than 8, in considerations of safety, economy and portable. In the present paper, the ignition characteristics of a scramjet combustor model fueled by aviation kerosene was experimentally investigated, using the resistance heating direct-connected supersonic combustion facility of Northwestern Polytechnical University. The inflow conditions at the direct-connected combustor entrance were specified as: the nominal Mach of 2.0, the total temperature of 870～930K and the total pressure of about 770kPa. Three combustor performance of difference ER were compared applied CFD simulated. The experimental result suggested that: the cavity design is a key problem for kerosene successful ignition is the combustor; That chooses appropriate gas-oil ratio could increased performance of combustor and kept isolator interfered by gas.

CFD Analysis of Hypersonic Combustion of H2-Fueled Scramjet Combustor with Cavity Based Fuel Injector at Flight Mach 6

2014 ◽  
Vol 656 ◽  
pp. 53-63 ◽  
Author(s):  
Krishna Murari Pandey ◽  
Sukanta Roga
Keyword(s):  
Flow Structure ◽  
Combustion Process ◽  
Wave Pattern ◽  
Combustion Efficiency ◽  
Fuel Injector ◽  
Recirculation Flow ◽  
Scramjet Combustor ◽  
Supersonic Mixing ◽  
Stable Flame ◽  
Scramjet Engine

This paper presents a numerical analysis of the inlet-combustor interaction and flow structure through a scramjet engine at a flight Mach 6 with cavity based injection. Fuel is injected at supersonic speed of Mach 2 through a cavity based injector. These numerical simulations are aimed to study the flow structure, supersonic mixing and combustion for cavity based injection. For the reacting cases, the shock wave pattern is modified which is due to the strong heat release during combustion process. The shock structure and combustion phenomenon are not only affected by the geometry but also by the flight Mach number and the trajectory. The inlet-combustor interaction is studied with a fix location of cavity based injection. Cavity is of interest because recirculation flow in cavity would provide a stable flame holding while enhancing the rate of mixing or combustion. The cavity effect is discussed from a view point of mixing and combustion efficiency.

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