The Ramjet Cycle - AA283 Aircraft and Rocket Propulsion

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AA283Aircraft and Rocket Propulsion

Chapter 3 - The Ramjet Cycle

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French Nord 1500 - turbo-ramjet, Mach 2.16 X-15 scramjet mockup

P-61 ramjet test SR-71(M-21 version) with D-21 ramjet drone

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Mass Flow

(3.1)

(3.2)

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M

M

Rayleigh relations for Tt and Pt

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We construct the flow through a ramjet engine beginning with supersonic flow in a straight constant area tube. Viscous Friction on the walls of the tube is neglected.

Add an inlet convergence and divergence

3.1 Ramjet flow field

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Introduce a burner and add heat to the flow.

Mass balance across the burner (neglect added fuel).

(3.5)

(3.4)

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(3.6)

(3.7)

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Add enough heat to unstart the internal flow.

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The stagnation pressure ratio across the engine is the same after the unstart.

(3.8)

(3.9)

This ratio is the result of losses across the inlet shock and across the burner. Across the shock

The Mach number at station 3 is 0.475. The added heat chokes the flow

(3.10)

The product of (3.9) and (3.10) is equal to (3.8)

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Now look at the thrust generated by this engine.

The net thrust is zero.

(3.11)

(3.12)

(3.13)

(3.14)

(3.15)

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Adding more heat generates some thrust.

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(3.16)

(3.17)

(3.18)

(3.19)

(3.20)

The thrust comes from the imbalance of pressure forces on the inlet.

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3.2 The role of the nozzle

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Equate mass flows at the inlet and nozzle throats - neglect fuel flow.

Stagnation pressure loss due to heat addition is proportional to the Mach number squared.

Neglect the stagnation pressure loss across the burner.

Adding the convergent nozzle increases the thrust.

(3.21)

(3.22)

(3.23)

(3.24)

(3.25)

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The thrust comes from the imbalance of forces on the inlet and nozzle.

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3.3 The ideal ramjet cycle

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The nozzle is fully expanded.

The exit Mach number equals the flight Mach number.

Additional thrust is generated by the diverging section of the nozzle.

(3.26)

(3.27)

(3.28)

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Effect of nozzle exit Mach number on thrust

Rearrange to express the thrust in terms of the nozzle exit Mach number.

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Fuel/air ratio.

(3.29)

(3.30)

Effect of fuel/air ratio on thrust.

(3.31)

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Specific impulse.

Overall efficiency.

Propulsive efficiency.

(3.32)

(3.33)

(3.34)

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Thermal efficiency.

For the ideal ramjet.

(3.35)

(3.36)

Recall the Brayton efficiency.

(3.37)

= 0.643

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T0 = 216K Tt 4 = 2000K A3 / A1.5 = 8 A4 / Ae = 3

Assume static pressure outside ramjet exit nozzle has recovered to free stream value

3.7 Ramjet with unstarted inlet

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State I

T0 = 216K Tt 4 = 2000K A1 / A1.5 = 8 A4 / Ae = 3

(3.62)

(3.63)

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State II - Increase inlet throat area until throat unchokes.

T0 = 216K Tt 4 = 2000K A1 / A1.5 = 5.747 A4 / Ae = 3

(3.72)

(3.73)

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State III - Remove the inlet throat altogether.

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State IV - Open the nozzle exit fully.

(3.80)

(3.81)

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State V - Reduce the burner temperature until the shock is very close to station 1.

(3.88)

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State VI - Reduce the burner temperature slightly to establish supersonic flow up to the burner.

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(3.38)

3.4 Optimization of the ideal ramjet cycle

(3.39)

Maximize thrust with respect to Mach number.

Express thrust in terms of component parameters.

Fuel shut-off

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Differentiate.

(3.40)

= 0(3.41)

(3.42)

Mach number for maximum thrust, f<<1.

neglect.

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Specific impulse.

(3.43)

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3.5 The non-ideal ramjet

Non-ideal effects include stagnation pressure losses due to:

Inlet shocks and viscous frictionBurner heat addition and frictionNozzle shocks and viscous friction

and stagnation temperature losses due to:

Heat conduction and radiation to wallsIncomplete combustion

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3.6 Ramjet control

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Mass balance.

(3.44)

(3.45)

(3.46)

The nozzle is choked.

Thrust

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(3.47)

(3.47)

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Scramjets

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3.3 The ideal ramjet cycle

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The problem of high flight Mach numbers

Thrust formula for a fully expanded nozzle

Thrust and drag coefficients

(3.89)

(3.90)

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Thrust coefficient for the ideal ramjet, Me = M0

Across the burner

(3.92)

(3.91)

f =CpTte −CpTt0hf −CpTte

=

TteTt0

−1

τ f

1+ γ −12

⎛⎝⎜

⎞⎠⎟ M

20

− TteTt0

Assume constant Cp and Cv. Rearrange (3.92) to express the fuel-air ratio as

Recall

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Complete (stoichiometric) combustion of JP-4

Complete (stoichiometric) combustion of hydrogen

Heating values

(3.93)

(3.94)

(3.95)

(3.96)

(3.97)

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Thrust coefficient at constant fuel-air ratio

(3.98)

(3.99)

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The solution?

Stanford PSAAP Center – AST Review Oct 20-21, 2008

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Scramjet flight envelope

Mach Number

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Stagnation pressure compared to dynamic pressure

(3.102)

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NASA Roadmap

X-43A- Integrated Vehicle Demonstration- Scramjet Engine- Short Duration Flight (Heat Sink Materials)

Dual Mode Scramjet- Actively Cooled Structure for long duration flight

Turbine Based Combined Cycle Rig

X-51A

Durable Combustor Rig

HIFiRE

Flight Experimentation

Long Duration

Combined Cycle

Stanford PSAAP Center – AST Review Oct 20-21, 2008

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HyShot - U. Queensland

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X-51A Waverider

2006 GDE-2 Test

2007 X-51 X1 Test

2008 X-51 X2 Test

2009 X-51 Flight Test

Stanford PSAAP Center – AST Review Oct 20-21, 2008

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Aug 14, 2012 failed because of a control fin failure after 16 sec.

Conditions for ignition appear to have been achieved but there was no ignition.

May 2010 flight lasted 200 sec and reached Mach 5 telemetry lost toward the end of flight.

2004 flight failed due to shock overpressure.

One X-51A left.

May 1, 2013 successful 6 minute flight with 210 seconds at Mach 5.

X-51A Flight history

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A Mach 6 concept from the Lockheed Skunk Works possibly operational 2030

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Problems

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