propulsion systems design · 2020. 10. 29. · propulsion systems design . enae 483/788d -...

Propulsion Systems Design ENAE 483/788D - Principles of Space Systems Design

U N I V E R S I T Y O FMARYLAND

Propulsion Systems Design• Lecture #18 - October 29, 2020 • Rocket engine basics • Survey of the technologies • Propellant feed systems • Propulsion systems design

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© 2020 David L. Akin - All rights reserved http://spacecraft.ssl.umd.edu

http://spacecraft.ssl.umd.edu



Grading Rubric for Team Project #2• Overall design (interior/exterior views, dimensioned 3-

view, functionality in mission application) – 10 pts • Habitability design (crew accommodations, usage

allocation, galley, waste management, windows) - 10 pts

• Life support systems design (trade studies, installation in CAD design, provision of resupply) – 10 pts

• Accommodation of visiting vehicles (docking ports, hatches, cargo interfaces) – 10 pts

• EVA accommodations (airlock(s)/suitports, suit donning/doffing/recharging stations, dust remediation) – 10 pts

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Grading Rubric for Term Project #2• Logistics management (volume allocation, crew

access, ability to restock) – 10 pts • Concept of operations (How will the habitat be

used? “day in the life”… how well does it fit the mission?) – 10 pts

• Presentation quality (“telling the story”, engaging the reader, adhering to principles from class) - 10 pts

• CAD quality (details, fidelity, complexity) - 20 pts • “Above and beyond” (extra effort outside of

nominal expectations) - 10 pts extra credit

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Propulsion Taxonomy

Ion

MPD

Non-ThermalThermal

Non-ChemicalChemical

Monopropellants Bipropellants

Pressure-Fed Pump-Fed

LiquidsSolids Hybrids Air-Breathing

Solar Sail

Laser Sail

Microwave Sail

MagnetoPlasma

ED Tether

Nuclear

Electrical

Solar

Beamed

Cold Gas

Mass Expulsion Non-Mass Expulsion

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Thermal Rocket Exhaust Velocity• Exhaust velocity is

where

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ve =2γ

γ − 1ℜToM̄

1 − ( pepo )γ − 1

γ

M̄ ≡ average molecular weight of exhaust

ℜ ≡ universal gas constant = 8314.3JoulesmoleoK

γ ≡ ratio of specific heats ≈ 1.2



Ideal Thermal Rocket Exhaust Velocity• Ideal exhaust velocity is

• This corresponds to an ideally expanded nozzle • All thermal energy converted to kinetic energy of

exhaust • Only a function of temperature and molecular

weight!

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ve,ideal =2γ

γ − 1ℜToM̄



Thermal Rocket Performance• Thrust is

• Effective exhaust velocity

• Expansion ratio

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T = ·mve + (pe − pamb)Ae

T = ·mc ⟹ c = ve + (pe − pamb)Ae·m (Isp = cgo )

AtAe

= ( γ + 12 )1

γ − 1

( pepo )1γ γ + 1

γ − 11 − ( pepo )

γ − 1γ



Nozzle Design• Pressure ratio p0/pe=100 (1470 psi-->14.7 psi)

Ae/At=11.9 • Pressure ratio p0/pe=1000 (1470 psi-->1.47 psi)

Ae/At=71.6 • Difference between sea level and vacuum Ve

• Isp,vacuum=455 sec --> Isp,sl=397 sec

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ve1ve2

=p

γ − 1γ

o − pγ − 1

γe1

pγ − 1

γo − p

γ − 1γ

e2



Solid Rocket Motor

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From G. P. Sutton, Rocket Propulsion Elements (5th ed.) John Wiley and Sons, 1986



Solid Propellant Combustion Characteristics


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Solid Grain Configurations


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Short-Grain Solid Configurations


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Advanced Grain ConfigurationsFrom G. P. Sutton, Rocket Propulsion Elements (5th ed.) John Wiley and Sons, 1986

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Liquid Rocket Engine

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Liquid Propellant Feed Systems

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Space Shuttle OMS Engine


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Turbopump Fed Liquid Rocket Engine


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Sample Pump-fed Engine Cycles


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Gas Generator Engine Schematic

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SpaceX Merlin 1D Engines

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Falcon 9 Octoweb Engine Mount

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Staged-Combustion Engine Schematic

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RD-180 Engine(s) (Atlas V)

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SSME Powerhead Configuration

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SSME Engine Cycle


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Liquid Rocket Engine Cutaway

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H-1 Engine Injector Plate

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Injector Concepts


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TR-201 Engine (LM Descent/Delta)

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Solid Rocket Nozzle (Heat-Sink)


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Ablative Nozzle Schematic


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Active Chamber Cooling Schematic


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Boundary Layer Cooling Approaches


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Hybrid Rocket Schematic


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Hybrid Rocket Combustion


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Thrust Vector Control Approaches


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Reaction Control Systems• Thruster control of vehicle attitude and translation • “Bang-bang” control algorithms • Design goals:

– Minimize coupling (pure forces for translation; pure moments for rotation)except for pure entry vehicles

– Minimize duty cycle (use propellant as sparingly as possible)

– Meet requirements for maximum rotational and linear accelerations

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Single-Axis Equations of Motion

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⌧ = I ✓̈

⌧

It = ✓̇ + C1

at t = 0, ✓̇ = ✓̇o =)⌧

It = ✓̇ � ✓̇o

at t = 0, ✓ = ✓o =)1

2

⌧

It2 + ✓̇ot = ✓ � ✓o

1

2

⌧

It2 + ✓̇ot = ✓ + C2

1

2

⇣✓̇2 � ✓̇o

2⌘=

⌧

I(✓ � ✓o)



Attitude Trajectories in the Phase Plane

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!3.00%

!2.00%

!1.00%

0.00%

1.00%

2.00%

3.00%

!20.00% !10.00% 0.00% 10.00% 20.00% 30.00% 40.00%

tau/I=0%

!0.001%

!0.002%

!0.003%

!0.004%



Gemini Entry Reaction Control System

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Apollo Reaction Control System Thrusters

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RCS Quad

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Apollo CSM RCS Assembly

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Lunar Module Reaction Control System

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LM RCS Quad

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Viking Aeroshell RCS Thruster

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Viking RCS Thruster Schematic

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Space Shuttle Primary RCS Engine


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Monopropellant Engine Design


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Cold Gas Thruster Exhaust Velocity

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Ve =

vuut 2�� 1



Cold-gas Propellant Performance


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Total Impulse• Total impulse It is the total thrust-time product for

the propulsion system, with units

• To assess cold-gas systems, we can examine total impulse per unit volume of propellant storage

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It = Tt = ṁvet

t =⇢V

ṁ

It = ⇢V ve

ItV

= ⇢ve



Performance of Cold-Gas Systems

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Self-Pressurizing Propellants (CO2)

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Self-Pressurizing Propellants (N2O)

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Density1300 kg/m3

Density625 kg/m3



N2O Performance Augmentation• Nominal cold-gas exhaust velocity ~600 m/sec • N2O dissociates in the presence of a heated

catalyst engine temperature ~1300°C exhaust velocity ~1800 m/sec

• NOFB (Nitrous Oxide Fuel Blend) - store premixed N2O/hydrocarbon mixture exhaust velocity >3000 m/sec

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2N2O �! 2N2 +O2



Pressurization System Analysis

Pg0, Vg

PL, VL

Pgf, Vg

PL, VL

Adiabatic Expansion of Pressurizing Gas

Initial Final

€

pg,0Vgγ = pg, fVg

γ + plVlγ

Known quantities:

Pg,0=Initial gas pressure

Pg,f=Final gas pressure

PL=Operating pressure of propellant tank(s)

VL=Volume of propellant tank(s)

Solve for gas volume Vg

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Boost Module Propellant Tanks• Gross mass 23,000 kg

– Inert mass 2300 kg – Propellant mass 20,700 kg – Mixture ratio N2O4/A50 = 1.8 (by mass)

• N2O4 tank – Mass = 13,310 kg – Density = 1450 kg/m3 – Volume = 9.177 m3 --> rsphere=1.299 m

• Aerozine 50 tank – Mass = 7390 kg – Density = 900 kg/m3 – Volume = 8.214 m3 --> rsphere=1.252 m

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Boost Module Main Propulsion• Total propellant volume VL = 17.39 m3 • Assume engine pressure p0 = 250 psi • Tank pressure pL = 1.25*p0 = 312 psi • Final GHe pressure pg,f = 75 psi + pL = 388 psi • Initial GHe pressure pg,0 = 4500 psi • Conversion factor 1 psi = 6892 Pa • Ratio of specific heats for He = 1.67

• Vg = 3.713 m3 • Ideal gas: T=300°K --> ρ=49.7 kg/m3 (4500 psi = 31.04 MPa) MHe=185.1 kg

€

4500 psi( )Vg1.67 = 388 psi( )Vg

1.67 + 312 psi( ) 17.39 m 3( )1.67

€

ρHe =pg,0M ℜT0

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Autogenous Pressurization• Use gaseous propellants to pressurize tanks with

liquid propellants • Heat exchanger to gasify and warm propellants,

then route back into ullage volume • Eliminates need for pressurized gases for ullage

and high-pressure storage bottles (e.g., Falcon 9 failures)

• Issue: start-up transient

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Nuclear Thermal Rockets• Heat propellants by passing

through nuclear reactor • Isp limited by temperature

limits on reactor elements (~900 sec for H2 propellant)

• Mass impacts of reactor, shielding

• High thrust system

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Electrostatic Ion Thruster

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Hall-Effect Thruster

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Ion Engine Schematic

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Ion Engines Existing/In Development• NSTAR (NASA Solar Technology Application

Readiness) – DS1 and Dawn – 30 cm, 2.3 kW, 92 mN, 3120 sec

• NEXT (NASA Evolutionary Xenon Thruster) – Available 2019 – 6.9 kW power, 236 mN thrust, Isp 4190 sec

• HiPEP (High Power Electric Propulsion) – TRL 3 – 670 mN, 39.3 kW, 7 mg/sec prop, Isp 9620 sec

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Variable Specific Impulse Magnetoplasma

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By Source (WP:NFCC#4), Fair use, https://en.wikipedia.org/w/index.php?curid=35831241



VASIMR Engine Concept

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VASIMR Operating Specifications• Optimum operating point

– Isp 5000 sec – Thrust 5.7 N – Power 200 kW

• Can be derated for higher thrust at lower Isp • Compare to ion engines at equivalent power

– 87 NSTAR thrusters: 8 N, 3120 sec Isp – 29 NEXT thrusters: 6.8 N, 4190 sec Isp – 5 HiPEP thrusters: 3.4 N, 9620 sec Isp

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Solar Sails• Sunlight reflecting off sail

produces momentum transfer

• At 1 AU, P=1394 W/m2 • c=3x108 m/sec • T=9x10-6 N/m2

€

T = 2 ˙ m V = 2 ˙ m c

€

E = mc2 ⇒ m = Ec 2

⇒ ˙ m = Et

1c2

=Pc 2

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Propulsion TaxonomyMass Expulsion Non-Mass Expulsion

Non-ThermalThermal

Non-ChemicalChemical Solar Sail

Laser Sail

Microwave Sail

MagnetoPlasma

Monopropellants Bipropellants

LiquidsSolids Hybrids

Pressure-Fed Pump-Fed

Ion

MPDNuclear

Electrical

Solar

Air-Breathing ED Tether

Beamed

Cold Gas

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propulsion systems design · 2020. 10. 29. · propulsion systems design . enae 483/788d -...

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