Example:	Takeoff thrust at the standard ambient temperature of 59°F will suffer ∼20% drop at an ambient temperature of 99°F!
Solution:	Water or water–alcohol injection in the compressor

• Heat absorbed to evaporate the water leads to a lower Δs, higher pressure and density
• Increased mass flow rate is not shown, since the h–s diagram is per unit mass
• Compressor pressure ratio increases as the corrected shaft speed increases with lower temperature
• Latent heat of evaporation of water is higher than alcohol, therefore it takes more alcohol for the same cooling effect.

At any engine speed, airflow rate varies directly with density at the compressor face, therefore p_t2

Normal shock inlets are suitable for Mach number
External–compression inlets are suitable for
Mixed–compression inlets are suitable for
Internal–compression inlets
External–compression inlets exhibit		“Buzz” instability in subcritical mode
All supersonic inlets with internal throat exhibit		Starting Problem/Unstart
Limit on the mean throat Mach number
Limit nacelle drag rise at transonic speeds by		“Slimline”, supercritical nacelle design
Inlet total pressure recovery at cruise

Interblade row spacing is typically one quarter of average stage axial chord length, namely,

Reynolds number based on chord
Modern transonic fan (stage) pressure ratio
Modern transonic fan aspect ratio
Compressor pressure ratio, per spool

Circumferential groove in the casing
Cubic boron nitride (CBN) tip coating effectively reduces tip clearance
This is the creep-rupture limit temperature of nickel-based superalloys, namely:
This limits the compressor pressure ratio to about 45–50.
In a conventional turbofan engine, axial Mach number at the engine face has a design value of about 0.5.
The current trend is in increasing Mz2 to ∼0.6.
Rotating stall cell spins at the rotor angular speed, in the opposite direction, in the relative (i.e., rotor) frame of reference.

Cruise TSFC drops dramatically with BPR The TSFC numbers are “ballpark” values.

		Burner efficiency drops with altitude (since pressure drops)
Typical burner efficiency at takeoff
Typical burner efficiency at cruise
Flow deceleration in the burner prediffuser
Typical combustor total pressure ratio

Since η_b drops with altitude, fuel-related parameters (e.g., f, TSFC) do not exactly follow the corrected parameters defined for TSFC_c

Stoichiometric fuel-to-air ratio for typical hydrocarbon or JP fuels
Fuel-to-air ratio in primary burner
Hydrogen versus hydrocarbon or JP fuel
Combustor liner cooling effectiveness
		TSFC doubles, while thrust increases by 50–80% Low-frequency oscillation: “Rumble”, f ∼ 50–100 Hz High-frequency instability: “Screech”, f ∼ 1–5 kHz

Gas path temperature in nozzle
Gas path temperature in rotor frame of reference
Blade service temperature
Adiabatic wall temperature
Coolant temperature at the rotor blade root (preswirl)
Coolant mass fraction in the first nozzle blade row
Coolant mass fraction in the first rotor
Thermal barrier coating (TBC) offers protection

In a turbine component, for example, blades, the blade life is reduced by 50% for each 10 K rise in material temperature

First nozzle, per spool is choked
All rotor exits in relative frame unchoked
Turbine Exit Mach number

Turbine efficiency loss per percentage increase in tip clearance is between 1.5 and 3%.

Critical Nozzle Pressure Ratio, (NPR)crit
Gross thrust is maximized when perfectly expanded
A subsonic jet is always perfectly expanded
Penalty of convergent nozzle versus C–D nozzle
Benefit of mixing the fan and core flows in TF engines
Thrust reverser effectiveness in separate flow TF
Thrust reverser effectiveness in mixed flow TF
Chevron nozzle jet noise reduction benefit
Chevron nozzle penalty on cruise gross thrust
Throat area scheduled with afterburner
NPR grows exponentially with flight Mach number, for example, in hypersonic regime we have