# California State University, Los Angeles - PowerPoint by 64JB6CO

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```									 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

FA  m / m  Fuel  AirRatio
f      a                              m  mass  of  Fuel
f

AF  ma / mf  Air  FuelRatio                    m  mass  of  Air
a

  ( FA) ACT /(FA)STOICH  ( AF)STOICH /( AF) ACT

 1                                  Running lean: Oxygen in exhaust
 1                                   Running rich: CO and fuel in exhaust

 1                                             Stoichiometric

General equation of combusting hydrocarbon fuel, excess air remaining after CO2
and H2O are formed

         y       79               y             y               79    
C x H y   x        O2     N 2   xCO2  H 2O  ( x  ) (1   )O2     N2 
         4       21               2             4               21    
Geometry and Boundary Conditions

Interior
Pressure Outlet

COMBUSTION CHAMBER DATA
Wall
•   Fuel: n-Heptane (Gas and Liquid)
•   Oxidizer: Air (%79 N2 - %21 O2)
•   Vertical Chamber                         Interior
•   Parallel Injections for Fuel and Air
(Study Velocity Inlet)                    Chamber Wall
•   Range   between 0.55-0.95                    300 mm

Diameter 70 mm
CONTOURS OF STATIC TEMPERATURE (K) & MASS
FRACTION OF CO2 FOR GAS HEPTANE

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

Static Temperature                     Mass 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


CP             a1  a2T  a3T 2  a4T 3  a5T 4
R
H           a1  a2 T 2  a3 T 2 3  a4 T 3 4  a5 T 4 5  a6 T
RT

S         a1 ln T  a2T  a3 T 2 2  a4 T 3 3  a5 T 4 4  a6
R

•     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
Fluent Results were Compared to:

•   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
Results
2D Subsonic Validation Successful – Spalart-Allmaras Turbulence Model was found to
produce the most accurate results.

2D Supersonic            Successful – Inviscid Solver was found to produce the most
Validation               accurate results.

3D Supersonic            Successful - Inviscid Solver was found to produce the most
Validation               accurate results.

3D Hypersonic            Successful - Inviscid Solver was found to produce the most
Validation               accurate results.
3D Dimensional Supersonic Validation of Winged Cone GHV At Mach 4

A Lift Coefficient Comparison between NASA's, CFD Team's and Fluent's Computational
Results for the Winged Cone GHV at Mach 4
0.2

0.15

0.1
Lift Coefficient

0.05

0
0          2               4                 6              8                 10   12

-0.05
Angle of Attack (Degrees)

NASA Report      S-A Solver       CFD Team       Inviscid Solver
3D Dimensional Supersonic Validation of Winged Cone GHV At Mach 4

0.07        A Drag Coefficient Com parison betw een NASA's, CFD Team 's and Fluent's Com putational Results for the
Wing Cone GHV at Mach 4

0.06

0.05
Drag Coefficient

0.04

0.03

0.02

0.01

0
0                  2                   4                   6                     8                 10          12
Angle of Attack (Degrees)

S-A Solver     CFD Team        Inviscid Solver       NASA Report
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)

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

Altitude
Speed
   Analyze NASA Langley, CIAM,
Angle of attack          Altitude
NAL, etc. scramjet test data for                                             Speed
performance curves, altitude, fuel
consumption, speed, flight angle                            Input engine
of attack, emissions, etc.                                   parameters
   Compare test data to Fluent
Model
   Create engine analysis
Performance
methodology for use as a design                         look-up
code)
   Output engine data will provide
results for CFD team                                    Output engine
parameters

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.
•   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.
Shock Wave Diagram with Ramp Injector

Combustor Duct

G (gap length) = 3 in
X/G = 16
G (gap length) = 3 in                         X/G=16
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 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
10.3 deg Ramp Angle: Mach 2 Airflow With No 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

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