By Richard Puster and Marco Egoavil, NASA Langley, Hampton, VA

The 8-Foot High Temperature Tunnel (8-Ft HTT) at NASA Langley is a combustion heated, hypersonic wind tunnel that has been used for many years to simulate supersonic flight conditions at altitudes in excess of 90,000 feet. The test section, 8 feet in diameter and 12 feet long, accommodates large, air breathing, hypersonic propulsion systems, such as the National Aerospace Plane (NASP) concept engine. It is also used to test structural and thermal protection system components, such as the exterior tiles used for the Space Shuttle. The high enthalpy environment is generated by a methane-air flame in a pressurized combustion chamber that is expanded through a hypersonic nozzle at the chamber exit. For air breathing propulsion tests, additional oxygen is added so that the molar concentration of oxygen is equal to that of air.

The spray bar, the original fuel injection system on the 8-Ft HTT

Researchers at NASA Langley have used FLUENT for many years to study several aspects of the flow inside the 8-Ft HTT. The fuel injector, the heart of the combustion chamber, was among the first components to be analyzed. The original fuel injector consisted of fifteen concentric rings, each of which was perforated with a number of small holes for spraying fuel into the combustion space. Called the spray bar, it was feared that the flame temperature near the rings would increase in the oxygen-rich environment of the chamber, such that melting of the spray bar material might occur. To prevent this from happening, tabs were installed at the sides of each ring that led to the formation of small side vortices. These acted to stabilize the high velocity fuel injection sites with lifted flames. The flame produced by the spray bar was first studied using a 2D cross-sectional model in FLUENT 3.02. In one simulation, the transient shutdown of the combustor was modeled, as the fuel velocity from the spray bar was gradually reduced. The FLUENT results indicated that the lower fuel velocities caused the flame to attach to the rings during the shutdown period. As a result of the simulation, engineers developed a more rapid shutdown procedure, and monitored the spray bar environment more closely than before during this process.

Temperature contours show how the pilot burner at the top of the new fuel injector array ignites the methane in a simulation of the upstream half of the combustor; the flow is from left to right

The original spray bar was subsequently retired, and replaced with a completely revised fuel injection and ignition system. The new system, which is still in use today, consists of an array of airfoil fuel injectors, an ignitor booster system, and a perforated plate positioned upstream of the fuel injectors to critically damp the pressure oscillations inside the combustor. As the name suggests, the airfoil fuel injectors have a cross-section in the shape of an airfoil, with holes at the blunt trailing edge for the spraying of fuel into the air that has passed over the airfoil. Following their success on the spray bar, tabs have been positioned at the sides of the widest portion of the airfoils to help stabilize the flame. The fuel injectors produce high speed methane jets, which are ignited by a pilot burner positioned at the top of the injector array. The ignitor burns methane and produces hot (3000 R), rich reaction products (H 2 , CO, and H 2 O), which ignite quickly in the nearby methane jets, producing a stable, high temperature zone. The high temperature zone then acts to ignite the remaining bulk of methane emanating from the array, giving rise to a large, stable methane-air flame downstream in the combustor. This ignition process is one of the fastest known for a hydro-carbon, and has been used in other super and hypersonic applications. The initial concept for this system was constructed in a pilot-scale, 9-Inch High Temperature Tunnel, and validated using FLUENT 5. The FLUENT predictions of axial temperature profiles for this ignition scenario are in good agreement with experimental data

To model the flow in the nozzle region, the exit flow from the combustor is required as a boundary condition. Due to the length of the combustor, engineers have solved for the reacting flow using two successive simulations in FLUENT 5. The first includes the industry Simulations at NASA Langley action of the fuel injectors and ignitor, extending to the halfway point in the combustor. The second simulation takes the exit profiles for all variables from the first simulation and uses them as inlet boundary conditions for the second. The exit flow from the second half can then be used for simulations of the nozzle. The nozzle itself is aircooled by tiny jets of transpiration air injected through small holes in the wall. To simulate this cooling method, eight mass flow inlets were positioned along the nozzle wall to correspond to the eight coolant circuits used in the actual HTT. The simulations verified suspicions that excessive injection flow rates (overcooling) would lead to separation upstream of the nozzle throat. Based on these results, it was determined that the coolant flow rates needed to be adjusted very carefully to adequately cool the nozzle and yield the desired conditions in the test region.

In yet another project, simulations of the combustor were used to study the spacing of the inlets used to deliver oxygen to the chamber. The oxygen is transported to the combustor area in the liquid state, in piping that runs behind an insulating liner that surrounds the flame. When gaseous oxygen is introduced to the hot combustion gases at the injector site, it is cold by comparison, and can cause a significant buoyancy effect. This drives the hottest gases to the top if the chamber and can lead to overheating on the upper combustor wall. FLUENT 5 was used to test the uniformity of the temperature distribution as a function of the cold oxidant bypass, or spacing between the fuel injectors and liner. The results showed that by reducing the spacing to the smallest value tested, better mixing could be achieved, leading to more uniform temperature distributions throughout the combustion chamber.

Editor's Note: Richard Puster and Marco Egoavil have worked for a combined total of 55 years at NASA Langley. While many FLUENT simulations have been run at NASA Langley during their tenure, the organization also provided the funding for the development of FLUENT/ UNS, the first segregated FLUENT solver to work on an unstructured grid, and a precursor to FLUENT 5. Puster and Egoavil are both planning to retire at the end of this calendar year.

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