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Niklas Wingborg FOI, Energetic materials

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Chemical Rockets Performance and propellants. Niklas Wingborg FOI, Energetic materials. Principle of rocket engines. Combustion chamber Nozzle. Throat Exit. Principle of rocket engines. De Laval nozzle.

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Chemical Rockets

Performance and propellants

Niklas Wingborg

FOI, Energetic materials

principle of rocket engines
Principle of rocket engines

Combustion chamber Nozzle

Throat Exit

principle of rocket engines1
Principle of rocket engines

De Laval nozzle

M<1 M=1 M>1

Tc Tt<Tc Te<Tt

gustav de laval 1845 1913
Gustav de Laval, 1845-1913

1883 AB Separator → Alfa Laval

1893 AB de Lavals Ångturbin → Stal-Laval AB →ALSTOM Sverige AB

rocket propellant classification
Rocket propellant classification

Fuel + oxidizer  gas + energy

  • Propellant = fuel + oxidizer
  • Liquid propellants
    • Bipropellant (storable, non-storable, hypergol)
    • Monopropellant
  • Solid propellants
propulsion systems in the ariane 5
Propulsion systems in the Ariane 5

Upper stage with storable propellants and Aestus engine

Solid propellant booster

Cryogenic main core stage

Vulcain engine

propellant performnace
Propellant performnace
  • Propellant content: up to 90%
  • Not unusual with 50%
  • The performance of the propellant very important
  • Propellant figure of merit: Specific impulse, Isp
  • Isp unit: Ns/kg, m/s or s
specific impulse i sp
Specific impulse, Isp
  • Optimum mixture oxidizer/fuel  high Tc
  • High heat of formation, ΔHf  high Tc
  • High hydrogen content  low M

CO2 44 g/mol

CO 28 g/mol

N2 28 g/mol

H2O 18 g/mol

H2 2 g/mol

calculation of specific impulse
Calculation of specific impulse
  • Nozzle/chamber
    • Pressure in combustion chamber, pc
    • Nozzle expansion
        • Pressure ratio: pc/pe
        • Area ratio: Ae/At
        • Chemical equilibrium or frozen equilibrium
  • Propellant
    • Chemical composition of fuel and oxidizer
    • Heat of formation of fuel and oxidizer
    • Mixing ratio fuel/oxidizer
thermochemical computation
Thermochemical computation
  • Computer programs for calculation of thermochemical equilibrium and Isp
  • NASA CEA (chemical equilibrium with applications)
    • NASA Reference Publication 1311 (June 1996)
  • Equation of state: ideal
  • Chemical equilibrium  minimizing ΔG = ΔH-TΔS
  • CEA can be obtained for free
    • http://www.grc.nasa.gov/WWW/CEAWeb/
    • http://www.openchannelsoftware.com/projects/CEA
common liquid rocket propellants
Common liquid rocket propellants
  • Oxidizers
    • Liquid oxygen, O2
    • Dinitrogen tetroxide, N2O4
    • Nitric acid, HNO3
    • Hydrogen Peroxide, H2O2
  • Fuels
    • Liquid hydrogen, H2
    • Hydrazine, N2H4
    • Monomethylhydrazine
    • Methane
    • Unsymetrical dimethylhydrazine
    • Kerosene
    • Ethanol
liquid oxygen lox o 2
Liquid oxygen (LOX), O2
  • Non storable oxidizer
  • Nontoxic
  • Mp= -219oC, Bp = -183oC
  • Used in combination with H2, kerosene, ethanol
  • Density = 1.14 g/cm3
dinitrogen tetroxide nto n 2 o 4
Dinitrogen tetroxide (NTO), N2O4
  • Widely used storable oxidizer
  • Different percentages (1-3%) of nitric oxide, NO, added as stress corrosion inhibitor (MON-1 and MON-3)
  • MON-1 and MON-3 are used more often than pure NTO
  • Bp= 21°C, Mp=-11°C, dens=1.43 g/cm3
dinitrogen tetroxide nto n 2 o 41
Dinitrogen tetroxide (NTO), N2O4
  • Safety concerns
  • Concern about reactivity of MON with titanium alloys, ignition by friction on freshly formed surfaces (e.g., pyrovalves).
  • History of accidents
  • Toxicity of vapor clouds in case of launch mishaps
  • State governments impose restrictions on transportation of NTO/MON
  • Space agencies have considered manufacturing NTO (and other toxic fuels) at the launch site to alleviate transportation restrictions
liquid hydrogen h 2
Liquid hydrogen, H2
  • Non storable cryogenic fuel, Mp= -259oC, Bp = -253oC
  • Used in combination with LOX
  • Density = 0.07 g/cm3 bulky fuel tank
  • Material problems  brittle at low temperature
  • Air / H2 explosive
hydrazine n 2 h 4
Hydrazine, N2H4
  • Can be used as a bipropellant fuel and as a monopropellant
  • Thermally unstable and cannot be used as a regenerative coolant in bipropellant engines
  • As a fuel, it is hypergolic with many oxidizers
  • Positive enthalpy of formation (+50.434 kJ/mol =+12.05 kcal/mol, liquid at 298 K)
  • Bp= 114°C, Mp=+2°C, dens= 1.00 g/cm3
hydrazine n 2 h 41
Hydrazine, N2H4
  • Hydrazine toxicity concerns
  • Acute toxicity: short-term exposure
  • Chronic toxicity: long-term exposure
  • Volatile
  • Carcinogen
monomethylhydrazine mmh h 3 c nh nh 2
Monomethylhydrazine (MMH), H3C-NH-NH2
  • Frequently used storable, hypergolic bipropellant fuel for satellites and upper stages
  • Can be used as a regenerative coolant in bipropellant engines
  • Low freezing point (-52°C)
  • Density = 0.87
  • Concern about toxicity of vapors (more volatile than hydrazine itself), Bp= +88°C
amsat p3 d launch campaign kourou
AMSAT P3-D Launch Campaign Kourou

MMH filling operation

N2O4 filling operation

http://www.amsat-dl.org/launch

aestus ariane 5 upper stage engine
Aestus: Ariane 5 upper stage engine
  • Fuel: MMH
  • Oxidizer: N2O4
  • MMH regenerative cooling
  • Multiple re-ignition capability
  • Thrust: 3 tons
  • Engine mass: 120 kg
  • Length: 2183 mm
rocket engine design
Rocket engine design

Injector

Chamber

At

Lc

Ae

injector face
Injector face

Mass flow and mixing  diameter of chamber

characteristic velocity c
Characteristic velocity, c*
  • Depends on the properties of the propellant
  • Unit: m/s (but it is not a velocity)
  • Independent of pressure (as long the reactions don\'t change)
  • CEA  c*
  • c*-efficency; ratio between calc. and measured c*
rocket engine design summary
Rocket engine design: summary
  • Propellant
  • Thrust, pressure and Ae/At
  • CEA  Isp, c*
  • Massflow
  • c* and massflow  At Ae
  • Injector and massflow  Ac
  • Propellant  Lc
solid rocket motors1
Solid rocket motors

Igniter

Nozzle

Case with propellant

solid propellants
Solid propellants
  • Solid mixture of oxidizer and fuel
  • Oxidizer: Ammonium perchlorate (AP), NH4ClO4
  • Rubber binder matrix: HTPB
  • Fuel: Aluminium powder
  • Burns on the surface
  • Burn time determined by the smallest dimension
solid propellant geometry
Solid propellant geometry
  • The case is protected by the propellant
  • Shape of combustion channel  pre-programmed pressure and thrust profile
combustion of solid propellants
Combustion of solid propellants

Piece of solid propellant: 10x20x50 mm

combustion of solid propellants1
Combustion of solid propellants

Small pices of propellants

combustion of solid propellants2
Combustion of solid propellants
  • Small pieces burn fast
  • The combustion proceeds perpendicular to the surface
  • Gas generation proportional to burning surface and burning rate, r
combustion of solid propellants3
Combustion of solid propellants
  • r measured at different pressures
  • a and n calculated

In this case at atmospheric pressure

combustion of solid propellants4
Combustion of solid propellants

n must be < 1, preferably 0.5 or lower

combustion of solid propellants5
Combustion of solid propellants
  • r is altered by the initial temperature. A warm propellant burn faster

T2 > T1

T2

T1

pressure

time

solid propellant mechanical properties
Solid propellant mechanical properties
  • Cracks in the propellant  > Ab > pc
  • Might lead to failure
  • Good mechanical properties is important
  • Must be elastic
  • Tg < minimum service temperature
  • Good bonding to case important
  • Debonding  > Ab > pc
manufacturing composite solid propellants
Manufacturing composite solid propellants
  • Liquid rubber (HTPB), AP and Al are mixed under vacuum
  • When properly mixed a curing agent is added
  • Continued mixing
  • Cast in mould to obtain desired shape
  • Cured at elevated temperatures
  • Mould = rocket motor
  • Machining
  • Final charge X-rayed to detect cracks, voids etc
manufacturing composite solid propellants1

~80%

Isp (Ns/kg)

% AP

Manufacturing composite solid propellants
  • Not possible to obtain maximum theoretical Isp
  • Isp limited by viscosity
  • AP particle size: bimodal or trimodal
composite solid propellants
Composite solid propellants
  • Large amount of smoke is formed
  • AP  HCL  hydrochloric acid
  • Shuttle  ~600 tons conc. hydrochloric acid
  • Ariane-5  ~300 tons conc. hydrochloric acid
current trends1
Current trends
  • Green solid propellants to replace AP (ADN, AN, HNF)
  • Green cryogenic solid propellants
  • Green oxidizers (N2O, H2O2)
  • Hypergolic rocket fuels to replace hydrazine and MMH
  • Green monopropellants to replace hydrazine
  • Exotic molecules, HEDM (N4, N8)
why is smoke a concern
Why is smoke a concern?

NC-baserat

AP/Al/HTPB

ammonium dinitramide adn
Ammonium dinitramide, ADN
  • Solid white salt
  • Intended for solid propellants
  • No chlorine content
  • Minimum smoke
  • High performance
  • Very soluble in water (80% at RT)
  • Synthesis developed at FOI
  • Produced on license by EURENCO Bofors in Sweden

NH4·N(NO2)2

solid propellant testing at foi
Solid propellant testing at FOI

Testing of missiles for the Swedish defense

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