RS2 SESSION PAPERS SESSION 7 JOYNER 7001C

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AIAA Responsive Space 2004

April 22, 2004

Highly Operable Propulsion System

Approaches and Propulsion Technologies

for Operationally Responsive Space

Systems

Russell Joyner

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6/4/02

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Presentation Outline

•Introduction

•Responsive Space, Historically Speaking

•“Spirally Develop” with A Focus

•Ground Rules for Study: Responsive Small Launch Vehicle

•Analysis Process

•Results – TSTO RSLV “Spiral 0-1”

•TSTO RSLV “Spiral 0-1” – Geometry Comparison

•“Spiral Development” from TSTO RSLV to HTO-RSLV

•Horizontal Take-Off (HTO) RSLV Concept Trades

•HTO RSLV Concept Comparison to Legacy Systems

•Boil Off Issues for Cryogenics - Impact of Integrated Thermal

Management Unit (ITMU)

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Introduction

• AFSPC 001-01/02; Operationally Responsive Spacelift (ORS) and Prompt Global Strike

Mission Needs Statement Decomposition

• “.. capability to rapidly put spacecraft into orbit”

• “.. maneuver spacecraft to any point in earth-centered space”

• “.. logistically support them on orbit or return them to earth”

• “.. strike globally and rapidly high value difficult to defeat targets in a single or

multi-theater environment”

• Operationally Responsive Spacelift Needs Architectures that Support an Over-arching

Vision That Can Evolve

• “Spiral Development”, Merging of Technical Capability and Budget Realities

• A “Spiral Development” Approach for ORS Needs A Roadmap that Includes the

‘Present” and “the Possible..Technologies on the Shelf or at High Readiness”

• An Approach for Creating the “Roadmap” from an “Operationally Responsive”

Propulsion and Propellants Point of View

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Responsive Space, Historically Speaking

Use of Cryogenics for Propellants Was Successful Because of Focused Process and Mission

Images Courtesy: Strategic Missile Website

Titan I

Jupiter

Thor

•

Time to Launch <20-minutes

•

Total Propellant Loading in

15-minutes After Launch

Commit Was Issued

~220,000 lbs. ~105,000 lbs.

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20mm Cann on

Sidewinder AA M

Sidewinder AAM

Pilot

APG-6 5 Rada

r

•

Original F-16 was designed for

an important, but limited role as

only an air-to-air fighter aircraft

•

But Evolved to Be More

Multi-Mission Capable

Data: ONE Team Payload & Sensors Presentation Jan 2002

“Spirally Develop” With A Focus

Visionary (But Focused) Approach Needed Early to Meet Full Operational Responsiveness Needs

+

•

A Responsive Small

Launch Vehicle Could

“Spirally” Evolve Into a

Highly Responsive

Launch Architecture

•

A Total Systems

Architecture Vision Is

Needed

A Horizontal Take-off

Type RSLV Carrier?

?

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Ground Rules for Study: Responsive Small Launch Vehicle (RSLV)

•

Notional Two Stage To Orbit (TSTO) RSLV As Baseline Concept (e.g. Similar to

Current TSTO Approaches Coming On-line)

•

Orbit Notional Mission: 1,700 pounds to 100/28.5

•

LOX/Kerosene Propulsion and Propellant as Baseline

– Boost and Upper Stage Performance Per Optimum ISP Nozzle Area Ratio and Max Diameter Per Stage Diameter, O/F, and 2 Combustion Chamber Designs

• Low Pressure, < 500 Psia; Higher Pressure, 750-900 Psia

– Pressure Fed for < 500, Gas Generator and Expander Cycles for 750-900

•

Start With LOX/Methane and 98% Hydrogen Peroxide(HTP)/Solid Fuel Hybrid

Evaluated for Upper Stages and Booster Propulsion

– Take LOX Operability As “workable” Per Historical Systems and Current Experience • Look at Methane (Tboil (K) 112) ... versus (Tboil (K) 90 for LOX)

– +15 Seconds ISP increase over Kerosene, O/F 3.5 versus 2.7 Gives Average Bulk Density Difference ~20% Which Trades With Lower Required Propellant Fraction

• Look at HTP/Solid Hybrid To See How The Performance Differences Vary So System Cost Attributes Could be Investigated

•

Look At General Thermal Storage Impact for “Sized” Vehicle Propellant Loads

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Analysis Process

• Start With Concepts Based on “Available

Hardware”, Investigate “Spiral Development

Elements”

• Define Notional TSTO (2-stage) Responsive

Small Launch Vehicles: LOX/Kerosene

Propellants

• Fly-off with POST (Trajectory Code) to

100nm/28.5 Nominal Mission, “Re-size” to Meet

1,700 pound Payload (Performance for Systems

Flying 1,000 to Higher, Polar Orbits

• Evaluate Alternative Engines/Propellants As

“Spiral Evolutions” to Base Notional Concept

PROPULSION PERFORMANCE

FLIGHT PERFORMANCE

ANALYSIS (POST)

MISSION-CONCEPT DEFINITION

MASS PROPERTIES

(Sizing)

SUMMARIZE RESULTS AND VALIDATE WITH DATA BASE

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Results – TSTO RSLV “Spiral 0-1”

0 20000 40000 60000 80000 100000 120000 140000 160000

LPLOX/RP Bst + U/S HTP Hybrid Bst + LOX/CH4 U/S

HP LOX/RP Bst + U/S HP LOX/RP Bst + LOX/CH4 U/S

LOX/CH4 Bst + U/S

G ro s s W t. (L b s )

+37% P/L

-20% +126%

+12%

Use HP O2/RP Bst,

Add CH4 U/S

TSTO RSLV Payload Sizing Trends

0 50000 100000 150000 200000 250000

0 500 1000 1500 2000 2500 3000 3500 4000

Payload(LEO/28.5deg) (lbs) G ro ss W t. (l b s)

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TSTO RSLV “Spiral 0-1” – Geometry Comparison

64 ft

75 ft

58 ft

Payload(lb) 1,700

1,700

2,300

1,700

GLOW(lb)

64k

144k

52k

68k

72k

Empty(lb)

3.7k

18k

4.7k

4.2k

11k

• Objective: Achieve Greater Responsiveness with

“Core” and Evolve Via “Spiral Development” to be

Fully Responsive With Technology Insertion via

Upgraded Stage Propulsion and ITMU Usage

_{Images Courtesy: Strategic Missile Website}

Design the “Spiral

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Level 1

“Spiral”

Level 0

“Spiral”

Level 2

“Spiral”

Options

High Pressure All LOX/Kerosene

High Pressure LOX/Kerosene Boost

LOX/Methane U/S

High Pressure Hybrid Boost or S/O

LOX/Methane U/S

Horizontal T/O & Hybrid Boost or S/O

LOX/Methane U/S

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Horizontal Take-Off (HTO) RSLV Concept Trades

0 50000 100000 150000 200000 250000 300000 350000 400000 HTO Bst+Hybrid Bst+LOX/CH4 U/S

HTO Bst+HP Bst LOX/RP+LOX/CH4 U/S

HTO Bst+LOX/CH4 Bst+LOX/CH4 U/S

HTO Bst+LP LOX/RP Bst+LP LOX/RP U/S

G ro s s W t. (L b s )

For TO t/w 0.5 2 x 51k TSLI

For TO t/w 0.5 2 x 40k TSLI

For TO t/w 0.5 4 x 28k TSLI For TO t/w 0.5

2 x 31k TSLI

B-58 Hustler w/Ext. POD

163,000 lb

Other A/C TOGW(lb)

For Comparison

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USAF/General Dynamics B-58 “Hustler”

TOGW(lb)

163,000

Payload(lb) < 40,000

Empty(lb)

56,000

Mach

_cruise

2.2

Sref (ft^2)

1,550

Length(ft) 97, b_span 56 ft

Runway Field Length < 7,900 ft, T/W ~0.3

TOGW(lb)

160,000

Payload(lb) <70,000(LEO 3,000)

Empty(lb)

74,000

Mach

_max

3.5

Sref(ft^2)

1,600

Length(ft) < 100, b_span 65 ft

Runway Field Length

< 5,500 ft, T/W ~0.5

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Boil Off Issues for Cryogenics - Impact of Integrated Thermal Management Unit (ITMU)

-20.0 40.0 60.0 80.0 100.0 120.0 140.0 160.0

0 20 40 60 80 100 120 140 160

Hold Time(hours) NO TOP OFF

B o il-O ff ( lb s/ H r)

HP LOX/KERO Both(Stg1&2) HP LOX/CH4 (Stg2)

HP LOX/KERO+LOX/CH4 Both(Stg1&2) LP LOX/KERO LOX Both (Stg1&2)

1-inch Insulation (Foam/MLI/Shields) Tambient 70degF 0 10 20 30 40 50 60 70 LP All LOX/Kero HP All LOX/Kero HP LOX/Kero+ LOX/Meth HTP + LOX/Meth P o w er R eq u ir ed ( kW e)

Images Courtesy: NASA GRC

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Summary of Observations

•

Responsive Spacelift Must Be Approached With A Careful “Spiral

Development” Approach

– A RSLV system must also keep trading off how the system meets affordable cost criteria, obtains high reliability and low maintenance, and has the performance to deliver a wide range of payload that could go as low as 100 pounds or as high as 12,000 pounds to LEO

• Most likely not done by a single launch vehicle design due to the affordability trade-offs but by some combination of stages that builds off the base design without compromising the “Demonstrated Responsiveness”

•

To Meet Global Reach and Rapid Spacelift Mission Needs, Systems Must

Respond in Minutes Like Current Military Aircraft

•

The Goal Should be; “Spirally Develop” Systems Using Evolved Propulsion

Technologies With A Strong Focus on Operability Within a Military Mission

Environment (e.g. F119, RL10)

– Evolve them to formulate a reliable, Operationally Responsive Spacelift and on-orbit architecture

– Evolve in innovative use of air-breathing propulsion, employment of soft-cryogenic fuels and oxidizers, low cost hybrid motors, and an integrated vehicle-engine health management system to create higher levels of

operational responsiveness

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20Klb Methane Expander Engine Concept

LEGEND Pressure PSIA Temperature Deg R Mass Flow Rate Lbm/Sec Flange, Orifice ,

Main Turbine Inlet

Main Turbine Exit

990 800 13.0 591 742 13.0

Total Area Ratio

70 : 1

Vacuum Thrust 22,000 lbf Vacuum Isp 353.2 sec 25 200 13.8 CH4 In 1549 216 13.8 1010 800 13.8 591 742 13.0

Turbine bypass=0.8 lb/s 500 6244 62.3 43 175 48.4 O2 In 880 179 48.4 OFC FSV OIV FIV

Regen Area Ratio

10 : 1

Radiation-cooled skirt

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~20Klb Thrust CH4/O2

Systems Integration RL10 CH4 History

Leverage current RL10 hardware - O2 turbo pump and fuel turbo pump - Fuel and oxidizer inlet valves

- Main fuel and main oxidizer valves - Thrust control valve

- Cool down valves

- Pneumatic control approach

Minimum

modifications to existing injector

Use existing 40Klb test chamber

Insert new TCA technology

Regeneratively or radiatively cooled nozzle

20Klb Methane Expander Engine Concept Attributes

Alter Gear Ratio Between Fuel and Oxidizer Pumps

Low Risk CH

₄

Expander Demo

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The Symbiotic Hybrid As Part Of an RLV

Key Features

:

No Additional Turbomachinery

Low Risk Pressurization Flow

Uses LOX tank pressurant from

vehicle main engines

Affordability Thrust Augmentation

Modular development

Benign environments Low complexity

Low cost fuel canisters Expendable Hybrid Fuel Canister (Pc ~ 1000 psia)

Vehicle Interface Pressurant flow from

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Notional HRC Characteristics For Hybrid Motor

Total length, feet 65 Avg thrust, lb 593,300 Vac Isp, sec 291 Total impulse, lb-sec 5.26E+07 Avg pressure 1200 Burn Ttme, sec 100

Weight:

Nose fairing ₃₁₀

Ignition system ₉₅

Injector ₁₇₅

Combustor case _6,415

Nozzle assembly ₈₀₀

Aft skirt/attach structures _1,050

Misc. ₅₀₀

Separation system ₂₅₀

Total Inerts _9,595

Fuel 67,240

Total HRC 76,835

Simplified HRC characteristics derived from

HPDP program (P&W team member)

Leverage HPDP

Technology

Fixed Nozzle ‘Baseline’

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Firebolt/HAST Hybrid Propulsion System

First Production Hybrid with Flight Maturity--Demonstrated Throttling Capability

Gross vehicle weight, lbs Total weight, CTA lb Fuel weight, lb Mass fraction

Max thrust, nominal, lb Total impulse, lb-sec Propellant

Throttle range Temperature limits, °F Status: Completed advanced technology Quantity produced 1231 268 153 0.53 1,200 156,000 Irfna/Butyl Rubber/Plexiglas 10:1 -45 to +82

50 1,600 1,400 1,200 1,000 800 600 400 200 0 T h ru st , lb 4.2 3.6 3.0 2.4 1.8 1.2 0.6 0 F lo w , l b /s ec 20 16 12 8 4 0 -4 -8 -12

0 40 80 120 160 200 240 280 Time, sec V o lt ag e, V D C Altitude thrust OTV command OTV feedback Oxidizer flow 12969

13.0 in. 13.1 in.

56.8 in.