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)



   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

                                                                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
    tool (spreadsheet or program                            tables
    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.
                     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.
                 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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