Rocket Engines Turbo Machinery

Hans Mårtensson, VKI-RTO March 2007 Rocket engines: turbomachinery,

Rocket engines: Turbomachinery

Hans MårtenssonSonny AnderssonStefan Trollheden

Staffan Brodin


Where is the meat ?

• No factor of 5-10 on Isp…• Systems are proven in operation

– Flying commercial load to orbit– Have flown to the moon

• ”Industrial” objectives optimization– Cost/Resource consumption per payload delivery– Safety– Operability– Time lines


Contents

• Introduction to how the turbopumps relateto engine type

• Example of a first pass preliminary design of turbopumps for a rocket engine

• Discussion on detailed design issues, and how modern tools affect design


Cycle objectives

• Deliver a maximum of thrust using a minimum of fuel– A rockets may take 10 tons to orbit at a 1000 ton take-off weight,

the absolute majority is fuel.

• The thrust is dependent on pressure ratio to increase jet velocity and allow high nozzle area ratio– Higher pressures give better Isp but more complex machines

jetn

jet

nn

vmF

vmg

F

mgFIsp

&

&&

=

⋅=

ityexit veloc nozzle isflow mass propellant is

constantn gravitatio theis thrust theis


Turbomachine design goals

• Objective 1: Deliver the reactants to the thrust chamber at specified pressure.

• Optimize the efficiency• Lower weight• Allow robust operation• Allow stable manufacturing process• Minimize cost


3 principal engine types

• GG entry pressure is above chamber pressure, sufficient to overcome ”plumbing”, and injection losses

• Stage combustion and expander exit pressures are above chamber pressure, leadingto high pressures in the turbine.

• The expander is simplified by not needing pre-burner, but adapted to hydrogen due to cooling properties. Also very sensitive to turbomachery efficiency due to limitedheating capacity.

NASA SP-8107


Range of applicability of different cycles

• Ranges of applicability of different cycles

1000 psia=70bar


Examples of rocket enginesFlowPressure

ratioTurbine inletpressure

Engine

high flowturbine

low pressureratio

200 barVinciRL-10

Expander

high flowlow pressureratio

400 barSSMERD-170

Stagecombustion

low flowhigh pressureratio

100 barVulcainF1,Atlas,Titan

Gas generator


• Thrust weight ratio– Vulcain 2 T/W=60 – Modern fighter engine T/W=10

• Note fuel flow 320 kg/s to payload ~10 tons -> requirement for rapid start in seconds

– Fighter engines go to full power in minutes (in the extreme)

(Source: fichier technique available at www.snecma.com)

SNECMA Vulcain2 engineGas generator cycle


VINCI LH2 turbopump(SNECMA/VOLVO)

• 2 stage pump + inducer• 1 stage turbine with low blade height and no OGV

– Note thick turbine disk indicating high tip speed– 240 K at turbine entry– No exit guide vane from turbine

Low blade height in turbine

Real gas effects

2.8 MW D=120mm


Vulcain 2 LOX turbo pump (AVIO/VAC)

• Single stage pump w. inducer• 2 stage overhung turbine• Separation of L2-Hot gas


SNECMA VINCI engineexpander cycle

• Turbines in series use H2 heated at combustor jacket• Chamber equilibrium temp. ~3500K• No pre-burner needed


Fuels• The most important difference is density

– Density – pump head and power: example at 100 bar

• Other important properties– Lubricating and cooling properties– Soot/Particles – erosion, wear, clogging

• Hydrocarbons (Methane, Kerosene…)– Material compatibility,

• H2 embrittlement, • O2 Oxidization

Density * [kg/m3]

Pump head at 100 bar [m]

LH2 75 13600 RP-1 (Kerosene) 810 1260 LO2 1200 849


Sample Turbine gas compositions

• The LH2/LO2 is most benign although with concerns for hydrogen embrittlement

• High partial pressures O2 maycause oxiization of materials

• Soot/graphite carbon for fuel richRP-1 may clog/erode the turbine

RP-1/LO2 Fuel leanMole fractions at eq. O/F=37, 1000K, 100bar*CO2 0.06007H2O 0.05857*O2 0.88137

RP-1/LO2 Fuel richO/F=0.13MOLE FRACTIONS

100bar 500barCH4 0.29388 0.33322*CO 0.00736 0.00600*CO2 0.00610 0.00664C2H6 0.00012 0.00040*H2 0.12417 0.07918H2O 0.07177 0.07663C(gr) 0.49660 0.49792

LH2/LO2 Fuel RichMole fractions at eq. O/F=1.023, 100bar*H2 0.87110 H2O 0.12890


Cavitation-NPSH• Erosion and loss of pump head may result• NPSH – Net Positive Suction Head

– Measures Pump inlet pressure margin to vaporpressure

– The necessary NPSH depends on purity of the liquidand inducer

• Cavitation is a dimensioning criterion on the pump– Tank pressure, separate boosters, inducer design

and design speed are controlling devices– The trade goes to weight and complexity of the entire

system


Cycle effects - turbine design

• Each level propagates a requirement to the lower system level.

• The requirement must be made based on estimates of what is possible– Technology demonstrators– Results of earlier experience

Cycle Turbopump Turbine Components

Manufacturingprocesses

Efficiency

Vibration control

Loading

Isp- what system ?

Merging all data


Development issues• Concept/preliminary design – assess potential

– Trade Cost-Risk-Performance (TRL-5/6 )– With engine designer

• Draw on experience• Anticipate technology advances

• Design – Realize solution– Analytical predictability– Draw on experience

• Verification (Analytical – Test)– Satisfaction of spec– Few hardwares

• Margin and experience, introduce robust design concepts


First pass design

• Consider a GG cyclePropellant mass flow rate 300 kg/s at O/F=6, Pchamber 100 bar, Pinlet 1 bar

257.1 kg/s12042.9 kg/s132 bar1 bar

Mass flow LO2

PexitLO2

MassflowLH2

PexitLH2Pinlet

Pump delivery data


Efficiency potential and overall sizing

• Yellow dot indicatessolutiond

Lower rim speed

ns-ds diagram with from NASA SP-8109

Specific speed

( ) 43

gH

Qn ins

Ω=

Specific diameter

( )in

s QgHDd

41

=

Tip speed limitation forces lower efficiency potential ifsingle stage H2 is used

Technology development may off-set this


First pass candidate solutionsfor the pump

(No NPSH concession made, No Bearing DN or rotordynamics consideration)

• Select best solution within experience base for product development

• Or push technology by selecting aggressive single stage solution

102000.846912472600LH2 2-stage

117000.746283106140LH2 1-stage reduced ns,ds

102000.8665104122100LH2 1-stage

35000.815813222980LO2

Power neededP[kW]

Efficiency potentialη[-]

Tip speedUtip[m/s]

Tip diameterD[mm]

Shaft speedN[RPM]


Turbine• Deliver power to pump at given speed

– Generally over op. range• For a GG cycle the exit pressure is low

– Trade design pressure ratio with thrust that can be achieved at exitpressure. Minimum is what is needed not to recirculate hot gas in engine bay.

• Use 1000K inlet temperature as upper limit for uncooled blades and vanes– All current machines are uncooled

• Optimize with respect to– Weight– Efficiency– Cost– Industrial issues

• Quality, Safety• Commonality, Experience, Tooling

– Structural


First glance

• The LO2 needs to deliver 1/3 of the LH2 power, savingson cost to be done at the LO2 hence pick 1-stage

• The LH2 will be based on 2 stage estimates

Green LH2

Red LO2

382.6103.5229801000100LO2

555.271010.2726001000100LH2

Efficiency*[-]

Massflow*[kg/s]

Pout[bar]

Power[MW]

Shaft speed[RPM]

Ttin[K]

Ptin[bar]

( )

−=−=

−γγ 1

000000 1220

PP

TCTisTCC outpp =2655m/s -> LOX U/C0 ~0.12, LH2~0.2

U/C0 U/C00.8 0.8


Picking out the stages in ns-dsdiagram

Turbine Balje

Figure 11 ns-ds diagram for turbines with selected points for the LO2 red and LH2 green options

2 stg 1 stg LO2 1 stg 2 stg

LH2

• Use diagram again to start off iteration


Define a turbine layout• Select and work in pitchline mode

– Using simple loss model + experience– Evaluate limits

Tip clearances critical due to blade height

Axial clearance-Chord-Length-weight

Blade count – loading - OGV-cost

Flow areas and angles – LE/TE thickness - sensitivity to tolerances

Reaction-Axial thrust-Efficiency

Off-design trades


Early structural allowances• Select INCO718

– Stage 2 runs a bit coolerthan stage 1

( ) 2222 Ω⋅−= IO rrAN π

”Specific root stress”


2 stage LH2 Layout view

AN2 on stg 2 is marginal at bestOptions: Select new material, Lowerspeed or push the limit by gooddetailed design

Angles and areas within bounds avoiding later problems with tolerances

Margin on efficiency


Ready to start detailed design• Trends

– Aspect ratio: classical h/c>2 in order to limit sec loss. Now we have allowed 0.8.

– Blade loading Zweifel nmber up >1, 0.8 typicalnumber in litterature. Saves cost.

– Blisk designs avoid expensive firtree and dead rim


Now start 3D CFD

3D multi-stage is standard

Is unsteady important to performance ?

Modest changes to the geometry is possible

Pick up major defects of the design


Secondary flow system• Ingress

– from main gas path may change the state of the cavity

• Axial thrust balance– Bearing stiffness and life vary with axial

load. Total axial load must not changesign

– Governed by reaction this varies with op. point.

• Heat loading– Hot start followed by adaption to steady

state at diffeent rates causes strain– Shut down processes causes a cold

shock– Hot geometry causes shift in

performance - clearances• Purge

– Oxygen pumps may inverts purgepushing hot gas through cavity

– LH2 is extremely cold with very small sealing gaps


Blade and vane active cooling• 25% efficiency gives ~1300K Tin

– Save 30% of the flow used to drive the turbine• 50% efficiency could allow ~2000K

– But thermal strains will increase (30K-2000K)• Hotter and heavier but definitely possible on the

H2 turbine– Cooling with Oxygen is off for fuel rich pre-burne, so

we need plumbing mixing in hydrogen for the LOX turbine.

– Complexity and expense in blades will make this difficult


Design optimisation• Push CFD and FE upstream allowing them to

affect earlier decisions– Adapt to new design points rapidly– Automated area ruling variation is key to matching

At 10% reactionlayout, optimised with nozzle


Robust design multi-point• Changes to the cycle and operability requires

off-design operation, often difficult in earlydesign

Worst case off design point

Refence design point

Solutions with almost flat efficiency over the envelope can be found


Robust design• Impact on cost in manufacturing• Low scatter in preformance parameters• Avoids traditional inherited conservatism

± 0.24.8 mmblade thickness

± 0.212.5 mmblade height

± 0.150.55 mmfillet radius

± 21.786 °stagger angle

± 0.211.3 mmchord length

± 0.10.157 mmtrailing edge radius

± 0.10.209 mmleading edge radius

tolerancedimensionparameter


Mechanics• Many problems in engine development relate to turbopumps

– "History of Liquid Rocket Engine Development in the United States1955-1980"

• Transients– Thick walled components designed for burst (pressure vessels,disk …)

have large thermal inertia.

Order of magnitude comparison of cycles for turbomachines

1-10 hours4-10Rocket

1000-10000 hours1000-10000Fighter engine

30000-100000 hours10000-50000Commercial jet aero engine

1E5-1E6Power generation gas turbine / Steam turbine

Operating timeNumber of cycles


Vibration

• A large share of late hard and expensivedevelopments problems, all programs– Forced response, flutter,rotor dynamic or

unidentified• Testing

– Performed in scale rig, TP rig and at enginelevel.

– Red lines monitored on each individual


Unsteady flow forcing

• Animation The unsteady force canbe >30% of the steadyloading

In our example ~100N oscillating per blade


Blade Vibration Forced response

• Avoid known and 1st order crossingsimperative

• Higher pressures require greaterdamping– Test

Campbell diagram Disk modes

0

1000

2000

3000

4000

5000

6000

7000

800 1000 1200 1400 1600 1800

RPS

Hz

1 EO2 EO3 EO4 EO1ND2ND3ND4ND 60 % 80%

Campbell diagram Blade modes

10000

15000

20000

25000

30000

35000

40000

800 1000 1200 1400 1600 1800

RPS

Hz

12 EO (S2-S1)19 EO (IS1)31 EO (S2)38 EO (2S1)57 EO (S1)60 EO (R1)1F1T1E2F2T2E 60 % 80%


Verify HCF margin

• In the Haigh diagram HCF limits the low AN2 (P/A) will pay off again by allowing higher vibration stress amplitude

load

Amplitudestress

Haighdiagram


Flutter - Aeroelastic instability

-0.5

0.5

1.5

2.5

3.5

0.0 0.2 0.4 0.6 0.8 1.0 1.2

Reduced frequency

Aer

odyn

amic

am

plifi

catio

n fa

ctor

AxialTangentialTorsion

Critical reduced frequencies

Stable =>

<=Unstable

relVfChordk ⋅⋅

=πIn our example:

LO2 1F k∼0.08 -> unstable

LH2 1F k∼0.2 -> marginally stable

1F, 1T probably subcritical

NEED TO ENSURE SUFFICIENT DAMPING


Machine hardware• Manufacture

– Looks easier on drawing !

10 mm !


Conclusion

• The turbomachines are key componentsthat will keep improving gradually.

• New automated methods come into useimprove the designs

• Analytical verification

Rocket Engines Turbo Machinery

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