Design Review Agenda - PDF Free Download

Design Review Agenda 1) Introduction, Motivation, and Previous Work a. Previous Work and Accomplishments i. Platform Launches ii. Successful Test Firings 2) More In-Depth Design Overview of the Existing System a. Hybrid Rocket Definition b. Rocket Chamber Components i. Injector plate ii. Pre-Combustion Chamber iii. Igniter iv. Fuel Grain v. Post-Combustion Chamber vi. Nozzle c. Feed System 3) Organization and Team Assignments 4) P07105 Steel Rocket Team Objectives a. Assisting other METEOR Teams where necessary i. Discuss the impact that our team has on everyone elses b. Improving Pre-Existing Design and Set-Up where necessary i. Manufacturing of Materials ii. Testing Procedures iii. Experimental Design iv. Data Acquisition c. Testing and Optimization of overall Hybrid Rocket design i. Nozzle Geometry ii. Fuel Grain iii. Feed System and Flow Rate iv. Other system components d. Provide Detailed Specifications to P07109 i. Fuel Grain Dimensions ii. Nozzle Geometry iii. Feed system pressures iv. combustion chamber lengths v. Injector plate dimensions e. Test specifications given to P07109 i. Manufacture the necessary components ii. Carryout testing of design iii. Verify that the system outputs are those that were desired f. Adjust specifications according to testing results (if necessary) g. Initiate Preliminary design of larger stages of rocket system 5) Theory and Equations a. Specific Impulse b. Nozzle Design 1

c. O/F Ratio d. Regression Rate Analysis e. Feed System Losses?? f. Fuel Grain Additives and cross-section changes?? 6) Summary of Testing 12/9/06 & 12/10/06 7) Design Improvements/Modifications a. Feed System b. Nozzle Calculations c. Combustion Chambers d. Fuel Grain e. Manufacturing of Materials f. Design of Experiments 8) Goals for Next Testing a. Experimental Design 9) Timeline for Spring 07 a. Larger Chamber 10) Financial Budget a. Current BOM b. Projected BOM 2

Previous Work and Accomplishments Platform Launches - platform reached altitude of approximately 30,000 feet twice - third flight reached a height of 70,000ft, 70% of it s final goal of 100,000 ft - All returned safely to the ground - platforms take live video feed and keep contact with the teams on the ground during entire flight Hybrid Rocket Static Test Firings - static testing: the rocket is not allowed to accelerate - Senior Design Team P06006: 2 successful firings - Fall 2006 ~10 Successful Firings - third flight reached a height of 70,000ft, 70% of it s final goal of 100,000 ft - Data Acquired: pressure, temperature and thrust data - All tests 100% safe More In-Depth Design Overview of the Existing System Hybrid Rocket Definition - Hybrid rocket engines are classified as those that utilize a liquid oxidizer and solid propellant to achieve thrust - Current Materials include Hydroxyl Terminated Poly-Butadiene (HTPB) as the solid fuel and liquid Nitrous Oxide as the oxidizer - Not the most efficient - Considered to be one of the safer combinations Rocket Chamber Components Hydroxyl-Terminated Polybutadiene (HTPB) Fuel Grain Chamber Wall Snap Ring Injector Plate Garolite Pre & Post Combustion Chambers 2-D Nozzle 3

Team Member Assignments I. Ignition System Ryan Kuhns Guion Lucas Joe D Amato -Leader -Leader II. III. IV. Manufacturing of Materials Joe D Amato -Leader Ray Mulato -Leader Guion Lucas Ryan Kuhns Kent Etienne Data Analysis & Experimental Design Joe D Amato -Co-Leader: Data Analysis Joel Baillargeon Guion Lucas -Co-Leader: Experimental Design Ryan Kuhns Propulsion Group Leader: Ray Mulato o Nozzle Design Kent Etienne Joe D Amato o Feed System Joel Baillargeon Joe D Amato o Fuel Grain/ Pre-Post Combustion Guion Lucas Kent Etienne Joel Baillargeon o Injector Plate Ryan Kuhns 7

Goals/Objectives of P07105 Steel Rocket Optimization Theoretical Isp for HTPB & NOX 320 s Properties Efficiency of propulsion system Ratio of thrust to weight Change in momentum per unit mass of propellant Affected by combustion temp, chamber pressure, exit pressure, and mass flow rate I sp F v = I m & g = e sp g0. 8

December Testing Purpose of Testing Vary Nozzle Geometry to see the effects on thrust Conceivably measure mass flow rate of system Weigh nitrous oxide tank and fuel grains prior and after each test Come up with approximate O/F ratios Introduce Team to current design 9

8 Deg Half Ang = Short 700.0000 600.0000 500.0000 400.0000 Pressure (psi) 300.0000 200.0000 Tank Pre-Comb Pre-Inj 100.0000 0.0000 0.0000 2.0000 4.0000 6.0000 8.0000 10.0000 12.0000 14.0000 16.0000-100.0000 Time (s) 8 Deg Half Ang - Short 80.0000 70.0000 60.0000 50.0000 Thrust (lbs) 40.0000 30.0000 Thrust 20.0000 10.0000 0.0000 0.0000 2.0000 4.0000 6.0000 8.0000 10.0000 12.0000 14.0000 16.0000-10.0000 Time (s) 13

8 Deg Half Ang - Long 700.0000 600.0000 500.0000 400.0000 Pressure (psi) 300.0000 200.0000 Tank Pr Inj Pre Pre-Comb Pre 100.0000 0.0000 0.0000 2.0000 4.0000 6.0000 8.0000 10.0000 12.0000-100.0000 Time (s) 8 Deg Half Ang - Long 90.0000 80.0000 70.0000 60.0000 Thrust (lbs) 50.0000 40.0000 30.0000 Thrust 20.0000 10.0000 0.0000 0.0000 2.0000 4.0000 6.0000 8.0000 10.0000 12.0000-10.0000 Time (s) 14

11 deg Half - Long 600.0000 500.0000 400.0000 Pressure (psi) 300.0000 200.0000 Tank Pre-Comb Pre-Inj 100.0000 0.0000 0.0000 2.0000 4.0000 6.0000 8.0000 10.0000 12.0000 14.0000 16.0000-100.0000 Time (s) 11 deg Half - Long 90.0000 80.0000 70.0000 60.0000 Thrust (lb) 50.0000 40.0000 30.0000 20.0000 10.0000 0.0000 0.0000 2.0000 4.0000 6.0000 8.0000 10.0000 12.0000 14.0000 16.0000 Time (s) 15

Project Specifications 17

Manufacturing of Materials 18

IGNITER MOLD 20

P06006 Successful Fire Comparison Design of Experiments Test #1 4 hole injector Needle valve in place Burn time = 17 sec Short chamber (15 ) Conical intake nozzle 1 pre and post combustion chamber Test #2 9 hole injector No needle valve Burn time = 15 sec Long chamber (24 ) Gradual intake nozzle 1 pre-combustion chamber and 2 postcombustion chamber In an effort to ensure efficient testing our team researched Design of Experiment techniques. We have decided to go with a full factorial design. The advantage of a full factorial design is that it uses statistics to eliminate some tests. Let s say we have a system with two factors to test the effect of. Each factor in turn has two levels we want to test at. Test the effects of these factors one at a time would yield six tests. Two would act as a baseline, and four would be at the respective high ends of the factors. Statistically we could achieve the same power of the aforementioned test, with two fewer tests through factorials. High High Factor B 5,6 O.V.A.T. Factor B 2 4 Factorial 1,2 3,4 Factor A Low High Low 1 3 Factor A High As shown in the graphic above, by combining testing of the two factors we can eliminate two tests. We will use the Analysis of Variance in Minitab to analyze our results and determine the highest contributing factor. Our next test will include these factors: pressure (350-450 psi, 50), fuel grain length (11 and18), post-combustion chamber length (1.5-2.0,.5). 22

Feed System Improvements The Goals of adding additional feed system equipment: Acquire a reliable value for mass flowrate in order to calculate a number of parameters Regression rate Test Chamber Pressure Oxidizer-Fuel Ratio Oxidizer Mass Velocity Feedback Pressure Regulator Constant supply pressure leads to overall experimental control Capability to vary supply pressure Optimization of nozzle for a supply pressure Gas Tank Heating Blanket Allows for a controlled internal tank temperature Temperature control leads to internal pressure control 23

Nitrogen Regulator 24

Nitrous Oxide 25

N2O Regulator 27

Mass Flow Measurements Our team researched different technologies to obtain mass flow. The preferred choice was a coriolis mass flow meter. This type of meter vibrates the tube of which the fluid passes. The meter then monitors changes in frequency, phase shift, and amplitude to determine the density of the fluid. Unfortunately, this option proved too costly as a quoted mass flow meter from Invensis Systems was $7,731.80. An alternative method had to be found. As suggested by our mentor Dr. Kozak, we looked into using an orifice plate from Lambda Squared. The orifice plate will create a measurable pressure drop within the fluid. And assuming that we are in the compressed liquid region, we would be able to measure pressures and temperature to determine density with a P-V diagram. Here is the link into the orifice plate: http://www.lambdasquare.com/oripac/5300_tech.htm Pugh Method Analysis 31

Nozzle Design The Rocket Nozzle is regarded as the most difficult portion of the design. It is also the most important, having up to a 30% effect on the thrust capabilities of the engine. Prior to optimization of the nozzle, certain temperatures and pressures within the system are needed. 11 degree Half angle optimized nozzle Nozzle design is based (in our case at least) around an assumed chamber pressure and temperature that we came upon through some calculations and educated assumptions. To calculate our Mach number at the exit of the nozzle (where the diverging section of the nozzle ends, we used the following calculation: γ 1 2 2 P γ 01 M e = 1 γ 1 Pe Where P 01 is your combustion chamber pressure, P e is the pressure at the exit of the nozzle, and γ is your ratio of specific heats. We assumed an ideal case, where exit pressure is equal to ambient pressure at a given altitude. To find nozzle throat area, we needed to approximate some numbers based on experimental data from previous tests. The most critical of these numbers, our gas constant, R, was calculated using a NASA program who s original purpose was to use the Method of Characteristics to design a 3-dimensional nozzle. Based on this R value, as 32

33 well as an experimental mass flow we acquired from our last round of testing, we used the following equation to re-design the nozzle: 2 1 * * = γ M T Rbar P m A throat t t We then calculate the Nozzle Area Ratio using the following equation, assuming a throat Mach number equal to one for choked flow through the nozzle. ( ) ( ) 1 1 2 2 2 1 1 2 1 1 + + + = γ γ γ γ t e e t t e M M M M A A Using combinations of a few equations above, we can find our final parameter for the nozzle, the exit area. ( ) 1 2 1 2 2 1 2 1 1 + + + = γ γ γ γ e e t e M M A A Using all of the above parameters, we can develop a 2-dimensional nozzle based on the area ratio of the nozzle. Achieving the correct area ratio will allow for optimal expansion of the nozzle plume. This proper expansion is crucial because optimal expansion will result in the greatest thrust (not under-expanded or over-expanded).

Variables for Next Testing Variables Levels NOX Pressure (Pre-Inj.) Length of Fuel Grain Post-Comb Chamber 3 levels ( of 50 psi.) 2 levels (11 18 ) 2 levels (1.5 and 2.0 ) Goals for Next Testing - Test the effects that the post-combustion chamber lengths have on thrust and efficiency - Vary the fuel grain length for efficiency and mass flow purposes - Control the pressure of the Nitrous Oxide right before injector late - Optimize the conditions that the Nozzle is designed around - Acquire more data regarding regression rates of the fuel grain - Successfully obtain a pressure and temperature measurement of the Nitrous Oxide during flow so the team can determine the state it is in - Successfully Implement the planned out experimental design Test # Pressure (psi) Fuel Grain Post-Comb Length (in) Chamber Length (in) 1 500 11 1.50 2 450 18 2.00 3 400 11 1.50 4 400 18 2.00 5 450 11 1.50 6 500 18 2.00 7 400 11 2.00 8 400 18 1.50 9 450 11 2.00 10 450 18 1.50 11 500 11 2.00 12 500 18 1.50 Planned Testing Procedure 36

Regression Rate Analysis & Larger Chamber Given the need for a 50 second burn at 100 lbs of thrust, our goal is to use regression rate analysis and data to estimate a chamber that will meet the needs of a launch able rocket with these specifications. Regression rate is basically how fast the fuel grain burns outward. The following two formulas were utilized to estimate the regression rates of the HTPB fuel grains during the December testing. One is an approximation that is based on the radius that was lost and the burn time. The other is dependent on the mass loss. r r & app = t loss burn r& mass = ( mi mo ) ρ π L HTPB fuel t r burn 2 o _ inner r o _ inner From the results below you will see that we calculated an average regression rate of about 0.6 mm/s (0.023 in/s). In calculating new dimensions, an assumption was made to increase the regression rate because of a couple of reasons. Firstly, we will hopefully be producing a higher amount of thrust (about 100 lbs) as compared to the last set of tests (only 70 lbs or so). Secondly, as the inner diameter is burned away, more surface area will be available to burn, ultimately increasing the average burn rate and thus the regression rate. So, as far as the new estimates go, a regression rate of 0.889 mm/s (0.035 in/s). If this number is multiplied by the burn time the result will be the amount of radius that there is to burn. This gives approximately 0.0445 m (1.75 in) to work with, thus increasing the outer radius of the fuel grain to 0.06 m (2.34 in). These numbers were given to the Structures team but we have to remember that they are preliminary estimates. The next round of testing will give the optimization team a better idea regarding regression rate analysis with the extra data that is obtained. Thus, after the next round of testing, a new approximation will be made regarding the outer dimension of the fuel grain and all other components of the system. After this is done, 07105 will be need to acquire all the necessary parts to complete testing on the new sized chambers to verify that the assumptions made were correct and that it is OK for the Structures team to base their design on those dimensions. 37

Current BOM 39

Future Costs (Incomplete) 41