1 2/10/2005thrust chamber assembly concept design review components of tca injector chamber nozzle

2/10/2005 Thrust Chamber Assembly Concept Design Review

1

Components of TCA

• Injector

• Chamber

• Nozzle


2

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

• 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. This allows for the propellant tanks to empty at the same rate

• 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


3

• With the information available we run the NASA thermochemistry code to obtain some useful data:– Chamber Temp (Tc) = 6341 R

– C* = 6044 ft/s

– Exit pressure (pe) = 5.66 psi

– Exit velocity (ve) = 9627.8 ft/s

– Cfvac = 1.593

– Specific heat ratio γ = 1.1398

– Molecular weight = 21.313

– Ispvac = 327.6 s

Design Process


4

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 ,,

effSLF

SL g

CcIsp

0

,*

SLIsp

Fm

From NASA code

Design parameters

From NASA code

Design parameter

Design Process


5

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 choose a contraction ratio of 2 to help with combustion stability

0

*

gp

mcA

ct

Design Process


6

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

tcconv

RRL

tan

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


7

• Injector design pressure loss is 70 psi.

• We use .2*Pc = 60 psi for the drop across the orifices

• Area for injection is found with the pressure drop from the manifold to the chamber with:

id

ii

pgC

mA

02

Cd is discharge coefficient = .80

Design Process

We need to select hole sizes based on drill bits that can be purchased. By selecting the number of orifices that we want we can find the hole sizes that we need. Going back we can find the new mass flows and actual O/F.


8

Design Requirements – Chamber

• L* = 42.5 in• Should withstand heat flux for burn time• Should withstand any transient pressure• Should not be overly complicated (Cheap to build)• Cannot use regenerative cooling because of lack

of pressure budget• Use ablative liner and film cooling or O/F bias.• Convergence ratio of 2• Need to be able to flange onto injector


9

Design Specs - Chamber

• Chamber Diameter = 3.69 in

• Length of chamber = 20.73 in

• Length of converging section ≈ .64 in

• Diameter of throat = 2.61 in


10

Current Chamber Design

• Put drawing here


11

Design Requirements – Nozzle

• Expansion ratio = 8• 75% bell to assist in weight reduction• Manufacturing must be taken into

consideration– Conical nozzle used to be cheaper to

manufacture

– CNC manufacturing has reduced cost of bell nozzle

• UncooledNASA Dryden


12

Design Specs - Nozzle

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

• 75% Bell– Lower Weight– Better Performance

Bell Nozzle on Pump-Fed LRE


13

Current Nozzle Design


14

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 = 8.45 lbm/s• Ox flow rate = 5.87 lbm/s• Fuel Flow rate = 2.58 lbm/s


15

O-F-O Impinging Injector

• Injector provides for propellant mixing by impinging jets. Two oxidizer jets impinge on one fuel jet.

O OF

Fan


16

O-F-O Injector

• Well known design process• Better performance compared to pintle• Allows for O/F biasing against wall and

film cooling• Propellants are well suited for this option

– SG propylene = .5– SG LOX = 1.14– O/F = 2.27


17

Injector Sizing

• 18 – triplets • 18 film cooling elements• Oversize outboard oxidizer element to

ensure jets stay away from the wall• Impingement point length/ diameter of

orifice should be ~ 5• Bore length/diameter of orifice should be >

3.5 to ensure Cd = .80• Manifolds – 10*area of orifices they feed


18

Injector Concept


19

Injector Concept


20

Injector Concept

• Put the picture here


21

Injector Performance Analysis

• With these sizes:

Stream Lengths

68.3

57.6

40.6

,

,

f

f

ino

o

outo

o

D

l

D

l

D

l

43.6

26.6

80.3

,

,

ino

o

outo

o

f

f

D

L

D

L

D

L

Bore Lengths


22

Injector Lengths


23

Manifold Sizes


24

Manifold Sizes


25

Manifold Sizes

• Aox,in = .08165 in2

• Aox,out = .01437 in2

• Afuel = .01452 in2

• Afilm = .00226 in2

• Flow Area/ Injection Area• Oxin = 2.296• Oxout = 7.738• Fuel = 9.043• Film = 33.186


26

Injector performance

• Velocities:– Ox = 88.35 ft/s– Fuel = 120.45 ft/s

• Momenta– Ox_out = 222 lb-in/s2

– Ox_in = 210 lb-in/s2

– Fuel = 224 lb-in/s2

0.9911 : 1.0000 : 0.9375


27

Injector Fill Times

Volumes

Vox = 7.13508 in3

Vf = .26875 in3

Volumetric flows

Qox = 142 in3/s

Qf = 118.5 in3/s

Fill times

tox= .05 sec

tfuel= .002 sec


28

Combustion Stability

Stable

Unstable


29

Current Concept Summary

• Injector: O-F-O Injector

• Chamber: Ablative Lining

• Nozzle: 80% Bell

Picture here


30

Numbers SummaryO/F 2.2729

Pc 300 psi

F 2200 lbf

ε 8

Tc 6340 R

c* 6044 ft/s

Pe 5.66 psi

ve 9628 ft/s

Cf)vac 1.593

γ 1.1398

MW 21.313 lb/lbmole

Ivac 327.6 s

Iopt 299.2 s

L* 42.5 in

εc 2

Pc/Pa 20.41

Cf 1.44207

η 0.95

Isp 257.35 s

c 8280 ft/s


31

Numbers

mdot 8.4519 lb/s

mo 5.8695 lb/s

mf 2.5824 lb/s

Dc 3.692 in Chamber

Ac 10.706 in^2

Dt 2.6107 in Throat

At 5.353 in^2

De 7.3841 in Exit

Ae 42.82 in^2

Lc 20.72 in Length of chamber


32

Numbers

ΔP 60 psi injector pressure drop

ρ lox 0.0412 lb/in^3 density lox

ρ fuel 0.0218 lb/in^3

Cd 0.8 discharge coeff

Dfilm 0.031 in film orifices

Dox,out 0.0781 in outside orifices

Dox,in 0.076 in inside orifices

Dfuel 0.0785 in fuel orifices

Aox 0.167887 in^2

Afuel 0.100703 in^2

Vox 1060 in/s 88.33333 ft/s

Vfuel 1445 in/s 120.4167 ft/s

mom_ox_o 222 lb-in/s

mom_ox_in 210 lb-in/s

mom_f 224 lb-in/s


33

Numbers

ac 49259 in/s 4104.917 ft/s

f1t 7821 Hz

f2t 12974 Hz

f3t 17845 Hz

f1r 16276 Hz

f2r 29800 Hz

Df/Vf 5.40E-05 s


34

Adiabatic Flame Temperature vs. O/F Ratio

3000

3500

4000

4500

5000

5500

6000

6500

7000

0 0.5 1 1.5 2 2.5 3 3.5

Phi

T_a

diab

atic

- R


35

5000

5200

5400

5600

5800

6000

6200

0 0.5 1 1.5 2 2.5 3 3.5

Phi

C_s

tar

- ft/

s

Cstar vs. O/F Ratio


36

250

260

270

280

290

300

310

320

330

340

0 0.5 1 1.5 2 2.5 3 3.5

Phi

I_va

c -

sec

Ivac vs. O/F Ratio


37

1.5

1.51

1.52

1.53

1.54

1.55

1.56

1.57

1.58

1.59

1.6

0 0.5 1 1.5 2 2.5 3 3.5

Phi

Thr

ust C

oeffi

cien

t

Thrust Coefficient vs. O/F Ratio

1 2/10/2005thrust chamber assembly concept design review components of tca injector chamber nozzle

Documents

volume design process

psi chamber pressure

s design process slide

combustion chamber

nasa code design parameters

injector slide

chamber temp t c

mixture ratio