California State University MFDC Lab. Combustion- Propulsion Team Students: Amir Massoudi – Justin Rencher Andrew Clark – Uche Ofoma Professor: Darrell Guillaume Feb. 10, 2004

OBJECTIVES  Improve combustor performance in Ramjet and Scramjet engines by optimizing air-fuel mixing to reduce pollutant formation and to increase engine efficiency Validate the CFD Software called “Fluent” Verify that it can accurately predict the products of combustion Verify that is can accurately predict energy output Verify that it produces CFD result that are consistent with STARS Model both Ramjet and Scramjet engines Modify fuel injection locations Alter the fuel-air ratios Modify the combustor geometry  Develop a Scramjet engine model that accurately predicts engine thrust given parameters such as angle of attack, speed, and altitude Seek out all published data on Scramjet engines Develop a Fluent model of a Scramjet Compare model performance to published data Run model under a variety of conditions to develop a look-up table to be used with the testbed.

AMIR MASSOUDI (Graduate Student California State Univ. L.A) OBJECTIVE Build 2D combustion chamber model with numerical software Make a physical combustor based on results produced by computer models Compare results from two models

Equivalence Ratio & Equation of Burning Hydrocarbon Fuel Running lean: Oxygen in exhaust Running rich: CO and fuel in exhaust Stoichiometric General equation of combusting hydrocarbon fuel, excess air remaining after CO 2 and H 2 O are formed

Geometry and Boundary Conditions Diameter 70 mm Wall Chamber Wall 300 mm Interior Pressure Outlet COMBUSTION CHAMBER DATA Fuel: n-Heptane (Gas and Liquid) Oxidizer: Air (%79 N 2 - %21 O 2 ) Vertical Chamber Parallel Injections for Fuel and Air (Study Velocity Inlet) Range between

CONTOURS OF STATIC TEMPERATURE (K) & MASS FRACTION OF CO2 FOR GAS HEPTANE Static TemperatureMass Fraction CO2

CONTOURS OF STATIC TEMPRATURE (K) & MASS FRACTION OF CO2 FOR LIQUID HEPTANE Static TemperatureMass Fraction CO2

ANDREW CLARK (Intern from Univ. of Manchester England) Objectives Find a non-technical method of creating a thermodynamic database for Fluent. This would allow the usage of liquid aviation fuels which are not currently contained in Fluent’s original thermodynamic database. Validate Fluent as a CFD code by comparing lift and drag coefficients obtained in Fluent with coefficients obtained experimentally and coefficients obtained with STARS.

Thermodynamic Database Summary of Fluent’s Thermodynamic Database: Contains NASA thermodynamic polynomials Thermodynamic polynomials are used to find thermodynamic and thermochemical properties of species within a temperature range Thermodynamic database used primarily for combustion and propulsion. Fluent uses a modified CHEMKIN II Format database Database was created in MS Access and mail-merged to MS Word Reasons to Construct Database: Update the current fuel types found in the Fluent database. More up-to-date polynomials can be used, most of Fluent’s data is from the 1980’s where the source database is updated monthly Be able to utilize new fuel types.

Thermodynamic Database NASA Thermodynamic polynomials have the form Completed thermodynamic tables for three fuels (n-Heptane gas, n-Heptane liquid, Jet A liquid) Used data from Caltech Fluent has different format for Polynomial Coefficients Converted polynomial coefficients from source format to Fluent format

Validating Fluent 2D Subsonic validation using JavaFoil (panel method) to produce theoretical data for a NACA 4415 airfoil 2D Supersonic validation using linearised supersonic airfoil theory for a diamond airfoil 3D Subsonic validation using STARS data supplied by CFD team for Titan (a NASA award winning student design) 3D Supersonic and Hypersonic validation using NASA’s report for a Winged-Cone GHV and the CFD team’s results from a CFD research code called STARS Fluent Results were Compared to:

Results 2D Subsonic Validation Successful – Spalart-Allmaras Turbulence Model was found to produce the most accurate results. 2D Supersonic Validation Successful – Inviscid Solver was found to produce the most accurate results. 3D Supersonic Validation Successful - Inviscid Solver was found to produce the most accurate results. 3D Hypersonic Validation Successful - Inviscid Solver was found to produce the most accurate results.

3D Dimensional Supersonic Validation of Winged Cone GHV At Mach 4

Uche Ofoma (Graduate Student California State Univ. L.A) Objective Seek out all published results on Scramjet engines Results  Many tests have been performed at NASA Langley  Results are classified so we cannot get them

Other Engine Performance Data (Tunnel)  Results from 2001 CIAM tunnel tests  Gaseous hydrogen used as fuel  Mach 6 flow velocity  Approx. 75 kg thrust measured

Other Engine Performance Data (Tunnel)  Published data from NASA/CIAM Hypersonic Flying Laboratory (Feb. 1998)

Other Engine Performance Data (Tunnel)  Japan’s NAL Kaduka Space Propulsion Laboratory scramjet engine tests at Mach 4, 6 and 8  Tests similar to NASA Langley’s  Net thrust of 500 N produced

Engine Model  Analyze NASA Langley, CIAM, NAL, etc. scramjet test data for performance curves, altitude, fuel consumption, speed, flight angle of attack, emissions, etc.  Compare test data to Fluent Model  Create engine analysis methodology for use as a design tool (spreadsheet or program code)  Output engine data will provide results for CFD team Altitude Speed Pitch Change

Justin Rencher (Undergraduate Student California State Univ. L.A) Objectives Accurately simulate supersonic combustion of an appropriate fuel in a two dimensional scramjet using the CFD software, Fluent. Approach Build geometry and cases based upon existing research and results, applying known methods and accepted approaches to the Fluent CFD environment. Results Building supersonic combusting ramjet simulations within Fluent that actually converge has proven to be quite difficult. Information on how to create such simulations is scarce and sometimes classified. Observations made at the recent AIAA conference in Reno have shown us that we are on the right track.

Description of Current Task The geometry for this particular model is based on published data from the NASA Langley Scramjet Test Complex The focus of these CFD cases is primarily on the behavior of fluid flow and combustion characteristics as they are affected by what are known as ramp injectors. These ramp injectors are utilized to enhance fuel/air mixing so that the combustor length can be reduced. A ramp angle of 10.3 deg was used in published data. The following slides show results of 10.3, 12.3, and 8.3 deg angles as determined by Fluent CFD Software. Results of airflow and combustion for each model are pictured. A Mach 2 airflow is used and combustion is carried out with gaseous n-heptane.

X/G=16G (gap length) = 3 in Shock Wave Diagram with Ramp Injector Combustor Duct X/G = 16 G (gap length) = 3 in

8.3 deg Ramp Angle: Mach 2 Airflow and Combustion The two top slides are air flow only, displaying contours of mach number for the 8.3 degree ramp injectors The slide to the left displays contours of static temperature for air flow with combustion

10.3 deg Ramp Angle: Mach 2 Airflow With No Combustion 10.3 deg Ramp Angle: Mach 2 Airflow and Combustion The two top slides are air flow only, displaying contours of mach number for the 10.3 degree ramp injectors The slide to the left displays contours of static temperature for air flow with combustion

12.3 deg Ramp Angle: Mach 2 Airflow and Combustion The two top slides are air flow only, displaying contours of mach number for the 12.3 degree ramp injectors The slide to the left displays contours of static temperature for air flow with combustion