AIAA-2004-3514 · Hydrogen, Methane, Propane, Octane, RP-1/Kerosene Easily add new fuel, oxidizers, and product species by supplying table of H, Rho, and S versus T & P. Staged-Combustion,

Copyright 2004 SpaceWorks Engineering, Inc. (SEI) All rights reserved.Engineering Today, Enabling Tomorrow Page 1

www.sei.aeroSpaceWorks Engineering, Inc. (SEI)

SpaceWorks Engineering, Inc. (SEI)

AIAA-2004-3514REDTOP-2: Rocket Engine Design Tool Featuring Engine Performance, Weight, Cost and Reliability

Director of Hypersonics:Dr. John E. Bradford

Senior Futurist:Mr. A.C. Charania

Director of Advanced Concepts:Dr. Brad St. Germain

AIAA Joint Propulsion Conference and ExhibitFt. Lauderdale, FloridaJuly, 2004

Including:- 2nd, 3rd, and 4th generation single-stage and two- stage Reusable Launch Vehicle (RLV) designs (rocket , airbreather, combined-cycle)- Human Exploration and Development of Space (HEDS) infrastructures including Space Solar Power (SSP)- In-space transfer vehicles and upper stages and or bital maneuvering vehicles- Lunar and Mars transfer vehicles and landers for h uman exploration missions- In-space transportation nodes and propellant depot s

Concepts and Architectures



Image sources: SpaceWorks Engineering, Inc. (SEI), Space Systems Design Lab (SSDL) / Georgia Institute of Technology

MotivationEnable improved prediction of liquid rocket engine performance (thrust, Isp) at the conceptual and preliminary stages of design

Capture impacts on engine performance and weight du e to design variable selections like mixture ratio, chamber pressure, and nozzle ar ea ratio

Establish main engine component (pumps, turbines, v alves) details like size and weight to provide better POD for detailed studies

Propagate propulsion-level effects (e.g. preburner m ixture ratio) to vehicle-level parameters like GLOW, dry weight, and length



REDTOP-2 Overview



REDTOP-2



SpaceWorks Engineering, Inc. (SEI) introduces the Rocket Engine Design Tool for Optimal

Performance (REDTOP)-2, an analysis code for the propulsion expert conducting conceptual

and preliminary rocket engine design studies.

REDTOP-2 is capable of performing a steady-state engine power balance for a variety of

cycles, predicting engine weight on a component basis, computing numerous engine cost

metrics, and estimating the reliability of the engine. REDTOP-2 allows for parametric engine

design and sizing which include designing for a required thrust level (at a specified ambient

condition), sizing at a specified total mass flow-rate, or designing for a specific throat area.

This package is currently available for purchase through individual licenses. The full

product suite includes self-installing executable, documentation with case study examples,

and selected online support.

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Capabilities Summary

Oxygen, Hydrogen-Peroxide, Water, Nitrogen

Hydrogen, Methane, Propane, Octane, RP-1/Kerosene

Easily add new fuel, oxidizers, and product species by supplying table of H, Rho, and S versus T & P.

Staged-Combustion, Gas Generator, Expander, Split-Expander, and Tap-Off

Fuel and/or Oxidizer-Rich Preburners/GGs or Catalyst Pack

Dual or Single Preburners with Series or Parallel Turbine Flow

Sizes engine at maximum operating condition to determine weight, then analyzes at throttled

engine setting for performance assessment.

Detailed weight predictions for chamber(s), nozzle(s), valves, low and high pressure pumps/turbines,

controllers, accessories, nacelle, etc.

Multiple cost model options to determine engine DDT&E, first unit cost (TFU), and costs over the

production run with learning curve effects.

Top-down reliability modeling approach based on design characteristics. Outputs include overall

safety, overall reliability, true/false cuts, premature shutdown occurrences, etc.

Built-in Oxidizer Propellant Options

Built-in Fuel Options

Generic Equilibrium Model

Cycle Options

Throttled Engine Analysis

Weight Breakdown Statement

Cost Modeling

Reliability Model



Input Parameters

Define engine cycle, configuration, propellant type, and accessories

Establish sizing method: thrust, flowrate, throat area

Define top-level characteristics: chamber pressure, mixture ratio, expansion ratio, throttle

Define cycle-specific parameters (defaults provided):

preburner mixture ratiopump and turbine efficiencies

valve pressure drops

nozzle shape, length factor, half-angle, equilibrium-to-frozen flow fraction

heat transfer characteristics (nozzle regen.?, reference Qflux and pressure, efficiency)

Complete reliability model questionnaire

Define costing scenario and environment

inflation rate, programs fees, learning curve, manufacturing methods, etc.



(12)

(2)

(4)

(20-40)

(6)

(15-40)

Quantity



Sample Dual-Preburner, SC Cycle Flowpath



REDTOP-2 Engine Component Models

Chamber- Used for main combustion chamber, preburner(s), and gas-generator- Supports multiple inflow streams- Determine equilibrium flow conditions in chamber an d at throat- Computes chamber geometry (length, diameter, throat area, etc.) and weight (case and injector)

Nozzle- Either Conical or Bell-shaped geometries with a fro zen-to-equilibrium flow fraction for kinetics effec t- Bell nozzle use Rao-method for detailed contour sha ping- Weight based on cooling method (radiative, regenera tive, ablative) and pressure

Pump- Power balance optimizer (CPS algorithm) specifies d ischarge pressure(s)- Shaft speed based on propellant cavitation and stage -specific speed limits- Number stages, impeller inlet/exit diameter, and im peller tip speed calculations also performed

Turbine- Sized from either specified pressure ratio or requi red horsepower, given shaft speed

Valve- Used for inlet valves, main valves, preburner/GG val ves, and bypass valves- User specifies pressure drop- Will compute weight for cases with and without flui d flows

Flow Splitter and Flow Divider- Used for flow splits (e.g. to parallel turbines) an d flow mergers (e.g. coolant flows)

Heat Exchanger- Supports either heat addition or removal from flow- User specified heat fluxes (BTU/in 2-s) and empirical scaling law coefficients based on reference pressure

Injector- Simple pressure drop analysis (no weight estimation )- Used for MCC, preburners, GG entering flows

Results and Output Parameters

PerformanceThrust and Isp at sea-level, vacuum, and ambient conditions

-frozen, equilibrium, and effective resultsPower Balance

Pump shaft speeds, horsepower, discharge pressures, NPSH, etc.Turbine pressure ratios and shaft speedsOpen-cycle flowrates (GG or tapoff)

WeightMain Chamber(s) and Injector(s)Nozzle(s)TurbomachineryPreburner(s)/GG Chamber and InjectorInlet, Main, Bypass, Control, and Preburner/GG ValvesOpen-Cycle Nozzles/Turbine Exhaust DuctsController/AvionicsAccessories

ReliabilityOverall Safety - catastrophic failureOverall Reliability - mission success

CostDesign, Development, Testing, Evaluation (DDT&E)Theoretical First Unit (TFU)Unit Costs (average or per build)

Component DetailsResults for each component with flow properties, areas, composition, weights, etc.



User-Interfaces



ModelCenter© Collaborative Environment“Phoenix Integration allows manufacturing companies to integrate and automate numerous software tools, rem ote locations, and different computing platforms into a cohesive environment for systems design…

…Our client software and back-end server software p roducts help you build an integrated process for your engineerin g design team.”

Phoenix Integration Inc.http://www.phoenix-int.comImage Source: Phoenix Integration Inc.

http://www.phoenix-int.com/products/index.html





REDTOP-2 User-Interface Options:Engine Definition



REDTOP-2 User-Interface Options:Nozzle Specifications

Cycle-Specific Nozzle Sub-Tab



REDTOP-2 User-Interface Options

Detailed Component Results – ASCII Files

Cost Model Results

WBS, Component Summary, and Reliability Results Fil es

Top-Level Performance and Power Balance Results

Test Cases





SSME Comparison

Staged Combustion cycle with fuel-rich dual-preburners100% regeneratively cooled nozzleREDTOP-2 sized engine at 109% throttle with tanked propellant flowrate of 1,129 lbm/sChamber pressure of 2,995 psi at 100% throttle setting

77.5 : 1Nozzle Area Ratio

5.0 %MFV delta-P

16.3 %MOV delta-P

78 %HPFTP Turbine Eta

78 %HPOTP Turbine Eta

73 %HPFTP Eta

0.68Ox-Side Preburner Mixture Ratio

Actual ValuesParameter

2,995 @ 100%Chamber Pressure (psi)

80 %HPOTP-S2 Eta

6.0Tank Supplied Mixture Ratio

67 %HPOTP Eta

0.976Fuel-Side Preburner Mixture Ratio

LOX/LH2Propellants

*Reference: Manski, “Cycles for Earth-to-Orbit Propulsion”, Journal Propulsion and Power



SSME Modeling Results

13.7813.02Overall Length (ft)

59.92 : 159.81 : 1SLS T/W

ActualREDTOP-2

Value

Parameter

73.05 : 1

7,0006,988Weight (lbs)

73.13 : 1Vacuum T/W

8,1017,970HPOTP-S2 Pressure Out (psi)

4,7904,681HPOTP Pressure Out (psi)

7,0546,940HPFTP Pressure Out (psi)

1.541.516HPOTP Turbine Pressure Ratio

1.5771.602HPFTP Turbine Pressure Ratio

44.0244.01Nozzle Exit Area (ft2)

0.5679

370.76

417,944

453.36

511,052

REDTOP-2

Value

0.568

371.9

419,404

453.5

512,350

Comparison Value**

SLS Isp (s)

SLS Thrust (lbs)

Parameter

Throat Area (ft2)

Vacuum Isp (s)

Vacuum Thrust (lbs)

Performance and Power Balance Weight and Geometry

**Reference: Manski, “Cycles for Earth-to-Orbit Propulsion”, Journal Propulsion and Power



RL10-3-3A Engine Comparison

53 : 152.2 : 1Vacuum T/W

ActualREDTOP-2Output Parameter

5.8

310

444.4

37.1

316Weight (lbs)

443.85Vacuum Isp (s)

5.6Length (ft)

37.2Tank Supplied Flowrate (lbm/s)

475Chamber Pressure (psi)

5.5Mixture Ratio

ValueParameter

61 : 1Nozzle Area Ratio

16,500Vacuum Thrust (lbs)

LOX/LH2Propellants

Pratt and Whitney engine used on Centaur upperstage of Atlas and Titan launch vehiclesExpander-CycleActual engine uses a single-shaft/turbine configuration with gear-box to drive pumpsREDTOP-2 flowpath utilizes non-geared, dual-shaft/turbine configuration with parallel flows

Engine Specifications to REDTOP-2

Results



J-2S Engine Comparison

69.7 : 173.39 : 1Vacuum T/W

ActualREDTOP-2

Value

Parameter

9.7

3,800

436

607.8

3,611Weight (lbs)

436.1Vacuum Isp (s)

10.5Length (ft)

607.6Tank Supplied Flowrate (lbm/s)

1,200Chamber Pressure (psi)

5.5Mixture Ratio

ValueParameter

40 : 1Nozzle Area Ratio


LOX/LH2Propellants

Tap-off cycle derived from J-2 engine used to power upperstages of Saturn VSlightly higher thrust class than J-2Improved performance and simpler flowpath over J-2Some flowpath differences between actual engine and flowpathin REDTOP-2User specified tap-off gas temperature limit; film cooling flow required to obtain this temperature then determined by REDTOP-2REDTOP-2 sized tap-off gas flowrate required to meet pump power requirements subject to a max. turbine pressure ratio


Results



RS-68 Engine Comparison

17.110.3Length (ft)

14,76113,620Weight (lbs)

357359.1SLS Isp (s)

ActualREDTOP-2Output Parameter

44.4 : 1

50.9 : 1

409

1,836

55.1 : 1Vacuum T/W

410.4Vaccum Isp (s)

48.2 : 1SLS T/W

1,772Tank Supplied Flowrate (lbm/s)

1,420Chamber Pressure (psi)

6.0Mixture Ratio

ValueInput Parameter

21.5 : 1Nozzle Area Ratio


LOX/LH2Propellants

Expendable, low-cost gas generator engine

Built by Boeing/Rocketdyne for Delta-IV EELV

Parallel flow turbines

Regeneratively cooled-chamber with ablative nozzle

Some flowpath differences between actual engine and flowpath in REDTOP-2


Results

Sensitivity Studies





Fuel-Rich, Single-Preburner Staged-Combustion Engine

340.3

388.6

316.7

377.2

281.5

359.4

SLS Isp (s)

69.9

76.9

70.6

79.2

68.1

80.6

Vacuum T/W

456.8

446.5

456.6

446.9

456.4

446.5

Vacuum Isp(s)

16.1100

50

100

50

100

50

Expansion Ratio

17.4

12.92,500

Overall Length (ft)Chamber Pressure (psi)

11.93,000

19.2

14.22,000

Designed a 600Klbs thrust-class LOX/LH2 staged-combustion engine

Fuel rich preburner (o/f=0.7) with series flow turbines (oxidizer-to-fuel)

100% regenerative cooled nozzle and chamber

Examined mixture ratios of 6:1 and 7:1

Performed sweeps of chamber pressure and expansion ratio at eachmixture ratio

Mixture Ratio 6:1 -- Results

Case Studies



Airbreathing TSTO RLV Concept





Concept of Study: TSTO Airbreathing RLVIOC: 2020

• First stage has combination propulsion system with TBCC, DMSJ, and multiple rocket engines• Second stage has single tail-rocket

*Representative Concept Illustration



A/B TSTO RLV Closure Model in ModelCenter© Environment

Approximately 10-15 iterations

Single Iteration Run Time:20 min.

Platforms: (1)Dual 1.8Ghz G5 on Mac OS X, (2) 3.0GHz Dell PIV PC, (1) SGI Octane Workstation

Closure Process

REDTOP-2, First Stage Engines

REDTOP-2, Second Stage Engines

ARWB SSTO Vehicle





Design Study: All-Rocket Winged Body (ARWB) Reusable Launch Vehicle (RLV)

Engines: 5 Advanced Single-Preburner Staged Combustion Engines (Pc 3,500 psi, mixture ratio 6.0)Propellants: NBP LOX and NBP LH2T/We: ~70

Propulsion

Payload: 15k lbs. (100 nmi. @ 28.5 degrees inclination from KSC), Cargo delivery or passenger delivery and returnReference Mission

Single-Stage-To-Orbit (SSTO) Vertical Take-Off Horizontal Landing (VTHL) Earth-To-Orbit (ETO) Reusable Launch Vehicle (RLV); commercial focus with initial flight capable in 2020, technology freeze date of 2015

Concept

CharacteristicsItem

29 ft78 ft

143 ft

LH2 TankLOX Tank

Payload Bay (15 ft dia. x 25 ft)

Main LOX/LH2Engines (5)

He Pressurant Spheres (4)Aft OMS/RCSTanks (LOX/Ethanol/H2O2)

Forward RCS Tanks(H2O2)

OMS Engines (2)



ARWB Reference Engine Specifications

60.616Average Cost ($M)

VALUE / DESCRIPTIONPARAMETER

39.386First Unit Cost ($M)

37.228Flight Unit Cost ($M)

99.998011Overall Safety

99.942512Overall Reliability

MR=0.85, T=1718.0 RPreburner

1,534.8Total Development Cost ($M)

60:1Expansion Ratio

80% Bell with 15o Half-angle,

100% Regenerative

Nozzle

11.9Engine Length (ft)

Exit Area (ft2)

Tank Mixture Ratio

Pc (psi)

Configuration

Cycle

Fuel/Oxidizer

33.259

6.0

3,500 psi

Single-Preburner

Parallel Turbines

Staged-Combustion

LH2 / LOX

ProbWorks

ProbWorks©, from Pi Blue Software (www.piblue.com) is a new suite of uncertainty and sensitivity analysis tools for use with Phoenix Integration’s ModelCenter © collaborative design environment.

This suite consists of four tools to help employ uncertainty analysis techniques , each implemented as a Java-based component which can function on any platform running Phoenix Integratio n’s ModelCenter ©

and Analysis Server ©.





REDTOP-2 (C++ on PC)

Approximately 5 iterations from Deterministic solution

Relative convergence criteria of 0.001 on mass ratio and engine T/W

Total Run Time: 1.75 hours

POST trajectory model (Fortran, PC)

Weight and Sizing Model (Excel)

Probabilistic Closure Model in ModelCenter© Environment

Closure Script (VB)

Design Variables (Pc, T/W takeoff, etc.)

Closure Process

DPOMD (Java on PC)



ARWB Concept Closure Results

11.811.91Engine Length (ft)

62.77

453.46

481,938

7.76

184,651

1,830,510

64.07T/W)sls

Vacuum Isp (s)

Engine Vacuum Thrust (lbs, each)

Mass Ratio

Dry Weight (lbs)

Gross Weight (lbs)

453.73

445,708

7.61

166,671

1,631,900

• Mission: Deliver 15Klbs to LEO (100 nmi. circular) a t 28.5o inclination from KSC Spaceport• T/W at takeoff of 1.15• Engine Mixture Ratio of 6:1• Nozzle Expansion Ratio of 60:1

PROBABILISTICDETERMINISTIC

Probabilistic vehicle closure resulted in ~11% incr ease in vehicle gross and dry weights

Conclusions and Future Work



Summary and Conclusions



A new design tool for use in the conceptual and preliminary stages of liquid rocket engine

design has been created. This tool, called REDTOP-2, has demonstrated its ability to model

existing engines designs with a degree of accuracy well suited for the early phase of vehicle

and propulsion design studies.

REDTOP-2 has been utilized in a number of design studies including:

- engine sensitivity analysis to chamber pressure, mixture ratio, and nozzle area ratio

- closure of an airbreathing TSTO RLV in a collaborative environment

-closure of an all-rocket SSTO RLV with propulsion uncertainty assessment

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Future WorkIncorporate detailed heat-transfer model for chambe r and nozzle regenerative cooling

Continue to add additional cycles, including monopr opellant and tripropellant configurations

Refine specific component analysis and improve exec ution speeds

Implement a bottom-up reliability model based on in dividual component/hardware items, safety factors, design fe atures, and flowpath details





SpaceWorks Engineering, Inc. (SEI)

Contact InformationBusiness Address:SpaceWorks Engineering, Inc. (SEI)1200 Ashwood ParkwaySuite 506Atlanta, GA 30338 U.S.A.

Phone: 770-379-8000Fax: 770-379-8001

Internet:WWW: www.sei.aeroE-mail: [email protected]

President / CEO: Dr. John R. OldsPhone: 770-379-8002E-mail: [email protected]

Director of Hypersonics: Dr. John E. BradfordPhone: 770-379-8007E-mail: [email protected]

Director of Advanced Concepts: Dr. Brad St. GermainPhone: 770-379-8010E-mail: [email protected]

Project Engineer: Mr. Matthew GrahamPhone: 770-379-8009E-mail: [email protected]

Project Engineer: Mr. Jon WallacePhone: 770-379-8008E-mail: [email protected]

Senior Futurist: Mr. A.C. CharaniaPhone: 770-379-8006E-mail: [email protected]

Documents

AIAA-2004-3514 · Hydrogen, Methane, Propane, Octane, RP-1/Kerosene Easily add new fuel, oxidizers, and product species by supplying table of H, Rho, and S versus T & P. Staged-Combustion,