1 2/10/2005Thrust Chamber Assembly Concept Design Review Liquid Bi-Propellant Thruster Concept...

2/10/2005 Thrust Chamber Assembly Concept Design Review

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Liquid Bi-Propellant ThrusterConcept Design Review

Adam PenderJason WennerbergArun Padmanabhan

Josh Revenaugh


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Outline

• Schedule and Milestones• Mission• System Requirements• Requirements Compliance Matrix• Design Process• Concepts Considered• Chosen Concept Details• Application to SLV Project• Conclusion


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Schedule Milestones

• Concept Design Review– Now.

• Preliminary Design Review– March 4

• Critical Design Review– April 7

• Hardware Completed – May 5


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• Develop thruster for Prospector-7– Cal State University-Long Beach /

Garvey Aerospace flight demonstration rocket.

– Prototype first stage for Nano-Launch Vehicle

– 246 lbm propellant– 550 lb GLOW– T/W = 4– Payload TBD– Max Alt: 30,000-50,000 ft.

(Burnout at 10,000 ft)

Mission / Application

~22 ft

26 in

Prospector-6


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Requirements

• Thrust: 2200 lbf• Chamber Pressure: 300 psi• Burn Time: 20 seconds• Nozzle Designed for Sea Level to 10,000 ft. Operation• O/F = 2.27 (propellants deplete at same rate)• Thruster Mass < 15 lbm• Injector Pressure drop of 70 psi• C* efficiency of 95%• Interface

– AN fittings (-8 if possible)– 10+ inch plate with bolt pattern communicated to CSULB

• Development Static Tests Performed at Purdue


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Pre-Launch Test Requirements

• Engine-only Static Test– Verify Performance

• Rocket Static Test– Verify Engine/Rocket Integration– 2-4 Seconds if engine to be reused

• Static tests may be performed by CSULB at Mojave


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Requirements Compliance

• Interface: Document Fittings and Flange Bolt Pattern

• Nozzle: Design with Area Ratio for Max Efficiency from 0-10,000 ft.

• Thrust, Burn Time, C* Efficiency, O/F : Static Tests

• Thruster Mass: Scale


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Design Process


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Components of TCA

• Injector

• Chamber

• Nozzle


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• The first step in the design was to pick propellants– LOX – propene chosen for several reasons

• Performance is increased over LOX - RP-1

• Customer has experience and access

• Allow for partial self pressurization of propellant tanks

• The mixture ratio is specified by CSULB based on the ratio that will give the best operability = 2.27

• A chamber pressure must be chosen– 300 psi was chosen by the customer.

• Current tanks can handle 450 psi 300 psi chamber pressure after losses

• Cooling by passive means is possible (No dump or regenerative cooling required)

Design Process


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• With the information available we run the NASA thermochemistry code to obtain some useful data:– Chamber Temp (Tc) = 6374 R

– C* = 6073 ft/s

– Exit pressure (pe) = 5.7 psi

– Exit velocity (ve) = 9673.5 ft/s

– Cfvac = 1.593

– Specific heat ratio γ = 1.1396

– Molecular weight = 21.232

– Ispvac = 329.3 s

Design Process


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With this data we can continue with the design of the engine. We would like to use the equation that relates mass flow rate to force and Isp so first we need Cf at sea level, and then Isp at sea level and then finally mass flow rate through the engine.

c

evacFSLF p

pCC ,,

0

,*

g

CcIsp SLF

SL

SLIsp

Fm

From NASA code

Design parameters

From NASA code

Design parameter

Design Process


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We know our O/F ratio so we can then split the mass flow into fuel and oxidizer:

1r

mm f

1r

rmmo

Where r is the mixture ratio

The throat area is found with:

We want the diameter of the chamber to be at least twice the diameter of the throat to help with combustion stability

0

*

gp

mcA

ct

Design Process


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We use the design parameter L* to find the size of the combustion chamber. We used an L* of 42.5 in because it has worked successfully in the past with RP-1.

*LAV tc This is the volume needed

c

cttconv

RRL

tan

)1(sec5.1)1(

Length of converging section with θc the converging half angle

)(3

22tctcconvconv RRRRLV

Volume that the converging section makes

c

cylcylconvccyl A

VLVVV

Use a cylinder to make the rest of the volume

Design Process


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• Injector design pressure loss is 70 psi.

• Area for injection is given by:

id

ii

pgC

mA

02

Cd is discharge coefficient = .65 for pintle or .72 for straight elements

Design Process


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Design Requirements – Chamber

• L* = 42.5 in• Should withstand heat flux for burn time• Should not be overly complicated (Cheap to

build)• Cannot use regenerative cooling because of

lack of pressure budget• Use ablative liner or thermal barrier coating and

O/F biasing for cooling• Convergence ratio of 3.5-4


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Design Specs - Chamber

• Chamber Diameter = 5 in

• Length of chamber = 9.32 in

• Length of converging section = 3.63 in

• Diameter of throat = 2.61 in


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Design Requirements – Nozzle

• Expansion ratio ~ 8• 80% bell is desirable• Manufacturing must be taken into

consideration– Conical nozzle used to be cheaper to

manufacture

– CNC manufacturing has reduced cost of bell nozzle

• UncooledNASA Dryden


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Design Specs - Nozzle

• Length of nozzle – 8.91 in (15° cone)– 7.13 in (80% bell)

• Preferred Option: 80% Bell– Lower Weight– Better Performance

Bell Nozzle on Pump-Fed LRE


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Design Requirements - Injector

• By far the most complicated part of design• ΔP = 70 psi• Shouldn’t melt or scorch• Provide combustion stability• No inter-propellant seals• Total flow rate = 8.51 lbm/s• Ox flow rate = 6.21 lbm/s• Fuel Flow rate = 2.3 lbm/s


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Pintle Option

• Easy to manufacture• Simpler design• Increases combustion stability

– Fuel orifices = .0292 in

– Number of fuel elements = 152

– Diameter of pintle = 1.56 in

– Diameter of oxidizer orifice = 1.64 in


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Pintle Injector: Manufacturing

Pros• Few parts to

manufacture

Cons• Little room for error in

manufacturing (Tilted pintle would effect LOX injection)


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Flat Face Injector

• More traditional option/better known• Better performance• O/F biasing against wall• Harder to manufacture (hole sizes are small

and numerous)– Fuel orifice size = .0319 in– Number of elements = 128– Ox orifice size = .0320 in– Number of elements = 256


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Flat Face Injector 1

Pros• Can be manufactured

with minimal brazing (Hand brazing or sweating only)

Cons• Many parts to

manufacture• Extra structural

supports needed?


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Flat Face Injector 2

Pros• Few parts to

manufacture• Solid (no extra

structural supports)

Cons• Outsourcing for

brazing (time & $)


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Injector Selection Criteria

• Wants:– Large knowledge base– Low-cost design– Easily manufactured– Meet performance requirements– Stable combustion


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Materials Compatibility

• LOX is highly corrosive, need resistive materials

• High Resistance– Copper, Nickel, Nickel alloys, and copper-nickel

alloys

• Medium Resistance– Stainless steels, aluminum alloys

• Low Resistance– Carbon steel, iron


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Materials

• Monel (Copper-Nickel Alloys)– High flame resistance (up to 10,000 psi)– High strength to weight ratio

• Inconel (Nickel Alloys)– High flame resistance (up to 10,000 psi)– High strength


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Materials

• Copper– High flame resistance (up to 8000 psi)– Cheap

• Stainless Steel– Medium flame resistance (304 up to 1000 psi,

316 up to 500 psi)– Good structural properties– Cheap


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Preliminary Material Choices

• Copper or Stainless Steel for Lox components – Budget Constraints– Manufacturability

• Nickel plating for injector face heat management if necessary


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Lox Seals

• All seals should be “Hard Seals”– Metal-on-metal contact

• Pure Teflon® – One of a few materials that can be used for

sealing components together if a hard seal is not possible

– Must be many layers for a robust seal to allow for thermal contraction


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Current Concept Summary

• Injector: Flat Face Injector

• Chamber: Ablative Lining

• Nozzle: 80% Bell (Conical Shown)

2” 9.44” 3.61” 8.95”

20°15°

Injector Combustion Chamber

2.62”5” = 8


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Applicability to SLV

• Engine could be used on a Purdue-designed rocket.

• Engine could be modified or scaled to suit the specific needs of the project.

• Design process could be followed for a new engine.


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Near Future Steps

• Injector Iterations

• Heat Transfer / Thermal Analysis

• Material Selection

• Structural Analysis

• Interface Design

• Nozzle Design

• Failure Mode Analysis


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¿Questions?


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Appendix

• Documentation Control

• Additional Design Guidelines


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Documentation Control

• Documentation Control – Anyone can post a file they deem useful, using the file format.– The Designers have sole access to editing any drawing files.

• File Format:– component_mmdd_ver.ext

• File Deletion– Old files are to be moved to the "OLD" directory for at least one week

before they are deleted. – Only the Designers or the Project Manager can delete any file, which

shall occur at regular intervals to clear out any unnecessary files.


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Prospector 7 is a modified version of Prospector 5

Thrust: 2,200 lbf; GLOW: 550 lbm

6 fully loaded tanks of each propellant: 246 lbm

Burnout altitude: 10,000 ft; Coast: 30,000 – 40,000ft

• Engine Design Guidelines:

• @ O/F: 2.27 Propylene and LOX will deplete at about the same rate

• Mass: 15 lbm

• Burntime: 20 seconds

• Injector: flat head designs due to higher performance potential and easier control of temperatures

• Tank Pressure: 450 psi; Injector Pressure Loss: 70 psi

• Chamber Design: Ablative chambers, with some including graphite throat inserts

Silica tape from Cotronics and an epoxy from a NASA TPS

• L*: 1 meter

• C*eff: 0.95 (if possible)

Specified Design Guidelines


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Design Objectives

• Design a liquid propellant thrust chamber assembly to be launched as a flight demonstration engine for CalState Long Beach.

• Design with the constraint that hardware must be designed and built using Purdue facilities and personnel

• Cost??

• Adam – feel free to add stuff here


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Design Objectives Level 2

• Thrust ~ 2200 lbf (Sea Level)

• Chamber Pressure ~ 300 psi

• Propellants: LOX-propylene

• Mixture ratio (O/F) = 2.27

• Burn time ~ 15s

• Engine mass ~ 15 lbm

1 2/10/2005Thrust Chamber Assembly Concept Design Review Liquid Bi-Propellant Thruster Concept...

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