Flight Prop 1

Energy Input

Method of operation

Medium (Fluid)

Guarantee Thrust/Power in Design and Operation

Safety

Reliablity

Certification Requirements

Input and Service Costs

Maintenance Cost

Initial and Amortization Costs

Unscheduled Cancellation Expenses and Repair Costs

Further Costs…

1. Turbofan

2. Piston Engine (Propeller)

3. Piston Engine

4. Turbofan

5. Piston engine

1. No

2. Yes

3. No

1. No, only rougly 25-30% leaves as shaft power

• Higher thermal efficiency at small size propulsion units

• Lower reduction of efficiency towards part-speed operation

• Fast dynamic performance response with little need for control systems

• Lower price, synergies with automotive engines, choices of modular units available

• Larger range of rotational speed

• Using exhaust turbocharger similar flight altitudes can be reached

• Operation possible with fixed-geometry propeller

• Much higher power per unit and per weight

• More compact design, lower frontal drag, more flexible integration

• Continuous process, higher reliability, lower overhaul intervals

• Low level of vibration due to continuous operation

• Lower oil consumption (lubrication only required for bearings and gears)

• More flexible in terms of fuel quality

• No need for water-cooling

• More simple adaption of thrust (through scaling or thermodyn. cycle mod)

• Small turboprops cant be enhanced to much further in terms of efficiency

• Large turboprops can (Large size of compressor and turbine does allow further thermodyn. cycle enhancements (OPR and TET were and will be increased))

• Piston Engine up to 300 kW are highly efficient

1. Turbofan

2. Turboprop

3. Turboprop

4. Both

• Electric Propulsion e.g. Distributed Propulsion

• Geared Turbofan

• Unducted Fan

• Piston engines

• Turbojet engines

• Turboprop engines

• Turbofan engines

• …

Generation of thrust FN

• at low fuel consumption

• at small size and low weight

Carnot Cycle

n_th = 1-Tmin/Tmax

• Rotary Engine

• Star Engine

• V-Engine

• In-Line Engine

• Boxer Engine

Special cases:

• Isochoric 𝑣 = 𝑐𝑜𝑛𝑠𝑡. : 𝑛 → ∞

• Isobaric 𝑝 = 𝑐𝑜𝑛𝑠𝑡. : 𝑛 = 0

• Isothermal 𝑇 = 𝑐𝑜𝑛𝑠𝑡. : 𝑛 = 1

• Isentropic 𝑠 = 𝑐𝑜𝑛𝑠𝑡. : 𝑛 = 𝜅

1. False

2. True

3. True

4. False

1. True

2. False

3. False

4. True

• Constant massflow through all components in the cycle

• Constant gas properties and gas mix across whole cycle (only heat transfer, not mass change in combustor), and constant specific heat capacity across the whole cycle (perfect gas)

• Adiabatic compression (C) and expansion (E), therefore, no cooling etc.

• No losses in components , therefore isentropic compression and expansion (no dissipation), isobaric combustion (CC)

Change of State:

0 -> t3 isentropic Compression:

𝑑𝑠 = 0, 𝑑𝑝 > 0

t3 -> t4 isobaric Heat Addition:

𝑑𝑠 > 0, 𝑑𝑝 = 0

t4 -> 9 isentropic Expansion:

𝑑𝑠 = 0, 𝑑𝑝 < 0

9 -> 0 isobaric Heat Removal:

𝑑𝑠 < 0, 𝑑𝑝 = 0

1. True

2. False

3. False

4. True

Optima shift up and to the right

Top of climb

In aviation a flame-out refers to the run-down of a jet engine caused by the extinction of the flame in the combustion chamber.

Caused by:

• fuel starvation

• compressor stall

• insufficient oxygen (at high altitudes)

• foreign object damage

• etc.

  
  

-> Restart the engine

• Windmill (ram air, needs sufficent height/speed)

• starter assisted

• APU delivers bleed air to the pneumatic starter

• That start powers a gearbox which in turn rotates the turboshaft

• If one engine is running, you can use the bleed air from this instead of the APU

N.B.: There is also the ASU Auxiliary Starter Unit (Ground support equipment)

Ram effect occurs when engine has some forward speed. As a consequence of stagnation,

some kinematic term adds to static quantities ( c2/2 )

• Turbine Entry Temperature

• Overall Pressure ratio

• Specific Thrust goes down with the Mach No. When AB is enabled it gets shifted to higher Mach numbers

• Specific Fuel Consumption goes up with the Mach No. When Ab is enabled, it gets shifted to higher Mach numbers & up

1. False

2. True (Since this gets in to the core)

3. False (Not at standstill)

• Geometric Scaling to Increase Air Mass Flow

  • required space for integration goes up

  • any synergy from common parts vanishes

• Increasing TET to Increase Net Work (Jet Velocity)

  • if not combined with countermeasures this leads to reduced life

  • leads to a throttling effect in station 4 (higher working line in compressor)

• Increasing Pressure Ratio to Increase Mass Flow (at same HP Turbine Capacity)

  • Only change in compressor

  • Ensure that it still delivers the same reduced exit mass flow

  • May result in a reduction of specific net work of the cycle, however increase in thrust because of the higher massflow

• Increasing Bypass Flow (Fan Diameter) to Increase Air Mass Flow

  • Increase fan diameter -> LP turbine needs to be changed

• Integration of an Afterburner to Increase Exit Momentum

  • Significantly increase the jet exit velocity c9

  • Core engine and afterburner can only be operated safely and stable when the exit nozzle is a variable area nozzle

  • Risk if core compressor surge

  • Increases the length and weight of the engine

• Injection of Water or Methanol to Increase Energetic Input and Jet Mass Flow

  • You can burn even more fuel without exceeding TET since the water has a cooling effect

-> High Bypass ratios -> Less acceleration of the massflow leads to high eta_p -> Leads to high massflow requierements to achieve the thrust

Inner / thermal efficiency → rates the conversion from heat to net power

evaluation of the jet engine process

Outer / propulsive efficiency → rates the conversion of jet power to propulsive

power

Overall efficiency → takes into account all losses, that occur during the

conversion of fuel energy to thrust performance

Overall efficiency = propulsive efficiency * thermal efficiency

1. True

2. False

1. False

• Recuperator is installer after LPT to recupe the heat of the exhaust gases

• These are then used to heat up the gases entering the CC

• -> Less delta_T has to be achieved by the CC -> Less fuel burn

• But for this to work, Tt_5 has to be significantly higher than T_t3

Optima shift to lower Total Pressure ratios!"

The inter-cooler cools the air inbetween compression components

-> reduces the work input required for compression

-> the amount of heat transfer achievable in the recuperator increases with decreasing compressor delivery air temperature

An engine has to be designed for a given flight mission and flight profile; however a number of additional factors have to be taken into account

• Thrust requirements depending on the flight profile

• Life cycle requirement (TET level depends on type of aircraft application)

• Fuel consumption (range / payload / weighting across mission)

• Engine weight and volume, fan diameter (drag)

• Development and production cost, common parts across fleet

• Technology maturation (to allow low risk engine design)

• High-confidence preliminary design and engine performance modeling

• Easy and minimum maintenance

• Employee training

Turbojet (Single Spool)

Turbojet (Twin Spool)

Turbojet (Single Spool) with Afterburner

Turbofan Low Bypass (Twin Spool)

Turbofan High Bypass (Twin Spool)

Turbofan High Bypass (Three Spool)

• Pro: Optimum rotational speed for each module (core/ mid pressure/ low pressure)

• Con: High level of complexity, high weight

Geared Turbofan High Bypass (Twin Spool)

Turboshaft (Twin Spool)

Turboprop (Twin Spool)

• Pro: High BPR possible (up to 100)

• Con: High level of noise through propeller

− Micro gas turbines up to 100 kW

− small gas turbines (mostly aero derivatives) up to 40 MW

− heavy duty gas turbines for permanent load up to 340 MW

No ram effect

Almost full expansion in turbine

• Friction losses in bearings

• Friction losses in sealings

• Losses in gear boxes

• Heat transfer (losses) within lubrication system

Bearings for the main GT shaft:

Operating Temperature ~ 300°C

High DN factor of <2.5e6 (diameter x rot. speed)

High reliability, long life cycle with high stress

Material Science Improved TET quite a bit

Advanced Air Cooling drove the biggest improvement in recent years

In a combined cycle in stationary GT steam can be used to cool Turbines

• Simple and robust arrangement

• All rotating components mechanically arranged on one single shaft

• Close coupling between power and rotational speed, suitable for constant speed and constant power

• Large power required for starting, but quick starting (within <45 min) capability given

• Net power off-take usually on compressor end of the machine (better access on single-spool arrangement)

• Used for permanent load

• More complex design (2 shafts)

• Usually on small GTs

• Rotating components mechanically de-coupled on 2 shafts

• High flexibility in rotational speed, also suitable for constant speed and flexible power

• Quick starting of the gas generator possible

• Special prevention required for over-speed on the „isolated“ GG shaft

• Single shaft gas turbines are fixed to generate electricity at frequencies of 50 Hz

-> Therefore, power can not be controlled using rotational speed

• Compressor Variable Stator Vanes (change of the inlet massflow)

  -> reduction of pressure ratio and the exhaust temperature goes up when maintaining TET

  -> thermal efficiency reduces

• Adaption of Turbine Entry Temperature (changing fuel flow)

  -> De-throttles the compressor

  -> Pressure ratio (hence power) drops

  -> Thermal efficiency reduces

1. False (no ram effect at all, c_0 = 0)

2. False (Mostly Air cooling Technology)

3. False (You expand as much as possible, but leave some delta_T to make exhaust gas leave the GT, and not otherwise not possible)

1. False (otherway around)

2. True (Compressor exit temperature goes down -> Higher delta_T between compressor exit temperature & exhaust gas temperature)

3. True

• Large power density ( in particular when applying in high MW range)

• Good torque characteristics of the power turbine

• Flexible fuel types

• Little vibration (low noise)

• High thermal efficiency, low specific consumption

• High level of reliability

• High specific power output

• Low air pollutant emissions

• Reduction of development time

• Low production costs

• Simple assembly

• Quick start/shut down ability

• Long intervals between overhauls

• Simple inspection and short maintenance work/overhauls

• Wide operating ranges

Open cycle (with continuous combustion at high pressure)

Closed cycle (with continuous combustion at ambient pressure)

Open cycle using recuperation

Closed cycle using recuperation

• Recuperation wins at low pressure ratios

• There, recuperation goes beyond conventional cycle thermal efficiencies

• For the same pressure ratio, inter-cooling reduces compression work (total enthalpy rise)

• The optimum cycle pressure ratio for an inter-cooled GT is much higher than for a conventional GT … thermal cycle efficiency increases but structural complexity increases

• Hence, the duty of the compression system is more difficult due to adverse effects at higher PR on component efficiency and

stability of the compressor

• Extreme OPR may take the rear-end geometry of the compressor (core size this is called) to a critically low physical size, which is difficult to manage aerodynamically – again, reduction of efficiency and stability

• Extraction of air, taking it through inter-cooler and bringing it back to the main flow path of the GT is related to losses through diffusion, friction, bending, and secondary flow effects in the volute

• Extracting and returning air to the GT flow path causes flow distorsion at the HPC entry

• Higher the inlet temperatures, reduce the compressor reduced speeds and the pressure ratio

• The TET stays the same (limited)

-> Higher ambient temperatures, cause lower relative power

-> Higher ambient temperatures, cause lower relative inlet mass flow

-> Higher ambient temperatures, cause lower relative efficiency

-> Higher ambient temperatures, cause higher exhaust temperatures

Combined Cycle Power Plant

Gas Turbine + Steam Turbine

Exhaust gases of the gas turbine are used to heat up water -> Steam turbine

-> Very high efficiencies achievable ~60%

Subsonic Flight Intake

• rounded edges, relatively short arrangement

• Low static pressure rise

• Low total pressure loss

• Causes distorsion to fan and compressor at high incidence or cross-wind (moderate aerodynamic interaction of intake and compression components)

Supersonic Flight Intake

• sharp edges, long arrangement, potentially variable

• High static pressure rise

• High total pressure loss

• causes distorsion to fan and compressor at any flight condition (strong aerodynamic interaction between intake and compression components)

• Can develop intake buzz, i.e. violent fluctuations in pressure and related shock system

Variable Arrangement:

• Allows adaption to varying flight Mach number and core engine massflow

• Manages poor subsonic inlet flow in this geometry

• Manages shock configuration at supersonic inlet flow

• Removes boundary layers to reduce compression system distorsion

• Do compression work and pressure ratio

• Can have variable stator vanes to vary its flow capacity and stage matching

• Feeds the engine secondary air system, providing cooling air as well as sealing air mostly to locations in the hot section of the engine: combustor and turbine(s)

• Low Mach no. entry flow

• Fuel injection

• Fuel/air mixing

• Multi-zone combustion to achieve low emissions

• Increase total temperature via intensive heat input to flow

• Do expansion work and pressure ratio

• Receive cooling air from compressor

• Drive compressor and auxiliary units of the engine; power balance on common shaft

SAS main purpose:

• Lead compressed air from different sources to bearings (sealing air) and to the rear of the engine to cool hot parts of combustor and turbine

• Allow effective cooling with a minimum of compressed air

• Cool life-critical turbine components: blades and vanes in gaspath and disks underneath the gaspath

In turbomachinery the state of flow is given by two components of the velocity vector:

1. the axial velocity defining how much mass flow passes the rotor

2. the blade speed defining how fast the rotor goes

-> Reduced massflow

-> Reduced Spool Speed

• „Reduced mass flow“ is a measure for the level of Mach number in a flow channel

• If that channel has a narrowest area (throat), the highest values will occur there

• When Mach number reaches a value of 1 in the throat, this state is termed „choking“, and effectively means, that reduced mass flow is fixed (can not be increased further)

1. True

2. False (Mach no. similtude)

3. False

Very often, aerodynamic speed and reduced mass flow are considered using reference values for pressure and temperature

-> The advantage of this is, that dot(𝑚)𝑐𝑜𝑟𝑟 has the unit of a mass flow and 𝑛𝑐𝑜𝑟𝑟 has the unit of a rotational speed

-> When Mach no. = 1 is reached in the throat area of a turbomachine blade row, dot(𝑚𝑟𝑒𝑑) (and dot(𝑚𝑐𝑜𝑟𝑟)) can not be further increased. The blade row is choked.

• The turbine is designed to work at choked flow condition, hence the reduced massflow at turbine entry (= compressor exit!) is fixed as long as the combustor gives a constant deltaTt.

• For the compressor, the turbine acts as a primary throttling device in the „engine system“, and therefore dictates the level of pressure (pressure ratio) produced in the compressor

• The compressor operating line (OL) is determined by the flow capacity (max. choking reduced massflow)

  of the turbine! When the turbine is designed to a higher flow capacity the compressor OL would drop

• Throat area defines turbine flow capacity

• Effective throat area is defined by pure geometric area minus additional blockage from aerofoil boundary layers and cooling flows released into the flow path

• Throat area is a major measure to set / correct operating line level

1. False

2. True (Throttling)

3. False

Increasing combustor fuel flow at given compressor inlet conditions (𝑻𝒕𝟐, 𝒑𝒕𝟐) causes:

→ a sudden increase in turbine entry temperature 𝑇𝑡4 (due to higher heat release in combustor)

→ a sudden reduction of HP compressor exit reduced flow, which moves the compressor

operating point futher up on the speed line (implying a rapid rise in compressor back pressure

𝑝𝑡3 and consequently stronger throttling of the compressor)

→ an accelleration to higher rotational speed due to an excess of turbine power

→ a transient excursion of the working line „up“ until a new steady-state power balance has

established between compressor and turbine, leading to stabilisation in a higher operating

point on the steady-state operating line ( SS Working Line) in the compressor map.

Reducing combustor fuel flow at given compressor inlet conditions (𝑻𝒕𝟐, 𝒑𝒕𝟐) causes:

→ a sudden reduction of turbine entry temperature 𝑇𝑡4 (due to lower heat release in combustor)

→ a sudden increase of HP compressor exit reduced flow, which moves the compressor

operating point futher down on the speed line (implying a rapid drop in compressor back

pressure 𝑝𝑡3 and consequently relaxed throttling of the compressor)

→ A decelleration to lower rotational speed due to a deficit of turbine power

→ a transient excursion of the working line „down“ until a new steady-state power balance has

established between compressor and turbine, leading to stabilisation in a lower operating

point on the steady-state operating line ( SS Working Line) in the compressor map.

IPC IS OPPOSITE

Plain Nozzle

-> low mixing rate

Ejector Nozzle

-> High mixing rate

Large Bypass Ratio Engines

Separated nozzles

-> Independant setting of Pt levels at fan exit and LPT exit!

Common nozzle

-> Requires matching of Pt levels at fan exit and LPT exit!

Afterburner Engine with Variable Nozzle

• Used for afterburner supersonic flight engines

• Variable mechanism to control throat area and thrust vector

• External arrangement

• Strong impact of nozzle on fan and core throttling

Requirements:

• Compact design → reduced number of stages

• Efficient work input → increased efficiency

• Safe operation → enlarged op range and stability margin

• Low cost → scaling and development of existing designs

higher specific work per stage ->euler equation:

u ↑ → transonic stage design:

• structural complications

• shock losses and stall margin issues

• Stage matching difficult

𝒄_𝒖𝟐 − 𝒄_𝒖𝟏 ↑ → high flow deflection (high profile camber):

• secondary losses dominate

• stall margin issues

1. True

2. False

3. False

Front split casing

Rear Split casing

Entry Casing

Rotor Drum + Blades

Engine mount

Exit casing

Axial Root

• when single blade weight is high

• need to be combined with stators having inner shrouds!

• Blades with extended roots are held in place using retaining plates

Circumferential Roots

• When single blade weight is low

• When blade count is high and some sealing is required at the root to prevent reverse leakage

• Can be combined with stators having either free inner tip or an inner shroud

No roots (BLISK)

• Integral design of blades and disk (Bladed Disk = BLISK)

• Milling from solid or Attachment of readily prepared blades to the disk by linear friction welding

• Advantages: low weight, low manufacturing cost/time, good structural stress distribution, low assembly effort, no root leakages

• Challenges: manage vibrational behavior at low damping level, repair of damage

Outlet Guide Vane (OGV)

➢ high flow turning, many aerofoils, potentially 3D stacking,

tightly spaced, conventional platform design difficult,

often designed as integral part with diffuser

➢ integral ring, cast in one piece, all aerofoils,

some post-casting machining, high roughness, low precision

(high manufacturing and build tolerances) … usually acceptable on this aerofoil

1. Aerodynamic design on a number of stream sections

2. Transformation of the surface into a CAD-volume model

3. Transformation of the CAD model into a meshed model for FEM analysis

4. Stress and deformation analysis, plus translation of the aerodynamically designed “hot” geometry to a “cold” geometry, which can be manufactured

1. True

2. False

3. False

• Rule to prevent flow separation during diffusion

-> Rotor: w2 > 0.7 * w1

-> Stator: c3 > 0.7*c2

• DH is a measure for the aerodynamic loading of boundary layers in the passage, primarily on the hub and casing

• Defines deceleration ratio between inlet and exit of the blade row

• The degree of reaction 𝜌ℎ defines the ratio of in the rotor static and overall stage static enthalpy rise

• For 𝜌ℎ = 0,5 vectors of the velocity diagrams at rotor inlet and stator inlet get equal when being „mirrored about the axial direction

• For 𝜌ℎ = 1 the inlet and exit velocities of the stator are identical, which implies there is no static pressure rise across the stator

phi_1 = cm_1/u_ 2

Throttling = „deformation“ of vector diagram giving increased rotor inlet flow angle 𝜷𝟏 by reduction of 𝒄𝒎 / 𝒖

𝝍𝒉𝒕 = ∆𝒉𝒕 / 𝒖𝟐^2

Ideal Characteristic in 𝜑−𝜓−Diagram:

Real Characteristic:

Contributing losses:

• Boundary layer (BL) friction losses

• Profile wake losses

• Shock losses and shock-BL interaction losses

• Secondary flow losses (tip leakage etc.)

• Mixing losses

The risk of profile boundary layer separation strongly depends on the surface loading distribution. The maximum velocity wmax is much higher than the inlet velocity w1 . The difference of wmax and the exit velocity w2 can be used to quantify the boundary layer loading on the suction surface of the aerofoil

• When a fixed scaling factor is applied to all geometric dimensions of a compressor, that is called „photographic scaling“.

• Simple but it changes relation of area and volume

• Blade speed for example would be maintained with this scaling approach

• In practise, there are further aspects of the design, which prevent full similitude when doing photographic scaling. As a consequence of these, the aerodynamic loading the boundary layers effectively are exposed to, will vary due to the following features of the geometry, which do not scale, when the compressor is scaled:

• Surface roughness on blades, hub, casing >> surface friction

• Running clearance on rotors and stators >> tip leakage and tip losses

• Stress/vibration motivated spacing/pitching of parts >> parts life

• Tolerances on gas-washed and non-gas-washed surfaces and dimensions of compressor parts >> losses, fit, and function

• Performance cycle (pressure ratio requirement; low PR should be achieved with low speed)

• Aerodynamics (Mach no., loss, stability; all these get non-linearly worse when tip speed goes up)

• Structural strength (centrifugal forces, vibrations, containment; high tip speeds cause high weight of drum and casing)

• Noise (noise emission regulations)

Aerodynamic similitude -> certain proportionality of physical parameters

• Ideally, conditions should also satisfy kinematic similitude of the flow, which requires to maintain both, Mach number und Reynolds number:

-> Reynolds number

-> Mach number

• In many cases this can not be achieved and only the more important of these two, depending on case, is held constant when scaling:

Re: incompressible or low-density flow >> in engine: LPT blades

Ma: compressible flow (gas dynamics) >> in engine: fan blades

All parameters relative to blade speed u are the same for all stages

-> Due to compressibility density of the fluid rises. Therefore, the gaspath area (height) has to be contracted in order to maintain 𝜑3 = 𝜑1 across all stages

Usually is independent of rotational speed of the compressor

But isnt if:

• Reynolds number strongly changes (this is the strongest effect)

• Compressibility effects become strong

• Further changes to the flow occur when Mach number approaches or exceeding the value of 1, and shocks occur in the flow field

Stagewise psi-phi plotting is essential to understand the stage matching of compressors

Multistage map plotting is essential to understand the overall performance of compressors

The geometric area is reduced to the effective area by boundary layer blockage

• 2D profile boundary layers

• 3D endwall boundary layer

• For Ma1 < Ma1,crit the impact on flow turning and loss is small in range up to approx. 0,7

• For Ma1 > Ma1,crit the impact on flow turning and loss is significant

At sufficiently high Mach number the supersonic patch is reaching the pressure surface of the adjacent profile. Now, the choke limit is reached

1. True (Endwall Blocking)

2. True

3. False (Critical Mach No is the incoming Mach number that causes the first supersonic flow on the airfoil)

Radial clearances (Rotor blade tip clearance,Cantilevered stator hub clearance,…)

Shroud cavities

Air off-takes

Loss generation caused by boundary layer friction and mixing

Shocks occur at high Mach no.

Occurance of a 3-dimensional complex overall flow field

• Due to historical reasons flow features are divided into separate individual phenomena. In reality, however, they often occur in combination

• Often caused by viscous effects

• The most important phenomena are the passage vortex occurring on fixed blade ends and the tip leakage vortex occurring on free blade ends

s/h is an important characteristic design parameter!

At the stall/surge line at least in part of the blading of the compressor boundary layer separation occurs

-> Strong oscillation of mass flow and huge pressure fluctuation

-> Surge is always triggered by the onset of rotating stall!

-> Global System Instability!

Pre-requisite for the occurance of surge are large mass flow and large volumes

Frequency ~ 3-10Hz

Inception either due to

-> Blade stall (low s/h gap sizes)

-> Tip stall (high s/h gap sizes)

Local instability!

Frequency ~ 50-100Hz

Unsteady phenomenon, but average compressor mass flow remains constant, locally it varies

1. True

2. False

3. False (Tip stall!)

➢ transient excursions

➢ bleeds and VSV scheduling

➢ thermal behavior / tip clearances

➢ manufacturing and build tolerances

➢ sand / FOD erosion

➢ inlet flow distortion

➢ deterioration

➢ any uncertainties

Loss of thrust

Backlash of flames into the compressor, causing thermal failure

Severe mechanical / thermal stress within the engine

Variable Stator Vanes (geometric adjustment)

→ change of velocity triangles

→ change of stage characteristics

→ change of overall characteristics

Each Vane is connected with a lever to a common ring that moves all vanes (Thermal Mismatch!)

Bleed Valves (releasing air from an inter-stage location)

→ increase of inlet mass flow while exit mass flow remains constant

Multiple spools (de-coupling of spool speeds)

→ individual design speeds

→ avoiding to high a PR on one shaft

→ individual lower PR compressor can be handled with sufficient efficiency and stability

▪ Located in the back of the engine

▪ „Hot End“

▪ Satisfy thermodynamic cycle requirements across flight envelope (altitude and Ma number)

▪ Significant LPT Reynolds number effects

▪ Working with extreme temperatures and pressure

▪ High aerodynamic loading at highest efficiencies

▪ Complex cooling systems

▪ Complex parts life requirements

▪ Continued strong demand for new high- temperature materials

Hot gas:

• Temperatures of up to 1600°C featuring strong local variation of 20% (Due to CC)

• High degree of erosion, corrosion, oxidation

Cooling air:

• Temperatures of approx. 600°C (not too cold because of the thermal stresses)

• Contamination with dust from the environment and from compressor abrasive liners

• Corrosion caused by sulphur content

Structural aspects:

• Steady and unsteady thermal strain

• Aero-elastic vibration (especially on high-AR blades)

• Unsteady load from rotor running-in

• Engine vibrations induced on stator vanes

• Clamping loads between aerofoil platforms

• Relative axial shift between rotor and stator

Radial Compressor: High Work coefficients but lower Flow coefficients (compared to axial comp.)

Axial Compressor: High Flow coefficients but lower work coefficients (compared to radial comp.)

Axial Turbine: Very high negative work coefficient and high flow coefficient (also higher efficiency than compressors)

Choking appears at critical area (Ma = 1)

Prandtl Meyer Expansion, expands the flow even further, accelerates and deflects the flow

-> According to the Euler equation increased turbine work output results from this

When reduced mass flow is increased choking appears (Ma = 1)

-> No further increase in reduced mass flow possible

-> Power output can still be increased however, since PM expasion causes further acceleration/deflection of the flow

Incidence Variation doesn’t affect choked turbin much, outlet stays roughly the same

1. False (Acceleration of the flow)

2. False (Operating Conditions, Performance…)

3. False (PM Expansion)

HPT (Nozzle Guide Vanes):

• Film Cooling

• Way thicker

LPT Vanes:

• No film cooling

• Thinner airfoils

HPT Rotor Blades:

• Either with or without shroud (with more efficient)

• Film cooling

• Fir Tree Root

LPT Rotor Blades:

• Either with or without shroud

• No film cooling

• Fir Tree Root

• Creep

• Oxidation, Corrosion

• Thermal Overload

An increase of blade metal temperature by 15°C is equivalent to a reduction of creep strength of 50%

Conventionally cast turbine blade

• Good mechanical properties in all directions

• Equi-axed crystal structure

Directionally solidified turbine blade

• Improved mechanical properties in longitudinal axis

• Columnar crystal structure

Single crystal turbine blade

• Excellent mechanical properties in longitudinal axis

• Improved heat resistance

1. True

1. False (Opposite is true)

2. True (But should be avoided as much as possible)

Convection Cooling

Impingement cooling

Thermal barrier

Local film cooling

Full coverage film cooling

Transpiration cooling

Typically:

1. Cooling air/core stream interaction

2. Rotor/stator interaction

3. Secondary flows

4. Tip clearance flow + control

5. Cooling system efficiency

6. Noise reduction

7. Sealing arrangements

8. Materials

• Pressure and temperature are simultaneously at their individual peak values in the engine

• To ensure safe operability there are high life requirements to be applied to combustion chambers

Efficient combustion:

▪ proper mixing

▪ stable burning

▪ cool hot combustion products down to level adequate for HP turbine parts

Axial-flow combustor

Reverse-flow combustor

▪ Complicated 3D and distorted flow field

▪ Complex process of mixing hot burnt and cold air

▪ Difficult thermal management of combustor and turbine casing

Sufficient margin against lean extinction needs to be established at all operating conditions

Re-light:

1. False

2. False

3. True

Aerodynamic Loss:

frictional loss in diffuser, frictional and mixing loss in combustion chamber

Thermodynamic Loss:

any deviation from the idealized case of inviscid flow in a duct with perfect heat addition

Based on momentum conservation, in a duct of constant cross-sectional area a reduction of density (hence flow accelleration) can only be achieved if pressure drops simultaneously. This corresponds to a change of state along a straight line in the p-v-diagram (Rayleigh-Curve)

-> Aero engine combustion chambers it is avoided to get even close to thislimiting flow regime

Flow Velocity:

Turbomachinery throughflow velocities > 150 m/sec

However: Burning velocity of kerosene in air is varying in a narrow range: 5-9 m/sec

Air-Fuel-Ratio:

Required range of air-fuel-ratios based on engine cycle energy balance is high within range of: AFR = 35 - 150

Stable combustion range of Air-fuel-ratio is very low: AFR = 4 - 30

Primary zone:

• Combustion is almost stoichiometric

• Flame stabilization in re-circulation area

• Fuel mixture preparation (spraying, vaporization)

Secondary zone:

• Air for post-combustion

• Decrease of combustion temperature

• Adjustment of temperature profile

Mixing (dilution) zone:

• Largest fraction of air is injected

• Temperature decrease

• Adjustment of temperature profile

• Rising temperatures at combustor inlet (Tt3) and exit (Tt4)

+ Cooling effect reduced by higher temperature cooling air Tt3

+ Turbine cooling air demand increased due to higher turbine entry temperature Tt4

• Requirement for low-emission advanced combustion

+ High air demand in central combustion zone in order to achieve lean combustion already in primary zone

→ less air for cooling parts and adjust TET profile

+ much larger portion of overall air flow enters combustion zone directly (lean burn)

+ more efficient cooling of flame tube (and turbine) is required

Advantages:

+ Simple and short design, as mixing and vaporisation is achieved before entry to primary combustion zone; low supply pressure required

Disadvantages:

+ Late stable operation → starting problems

+ Poor operation during transient (accel, decel)

Decel: Tube may be over-heated due to reduced cooling provided by low fuel mass flow

Accel: incomplete vaporisation due to very high cooling induced by high fuel mass flow

Nowadays:

(Conventional): RQL ..… Rich burn - Quick Quench - Lean Burn

(New tech.): LDI ….... Lean Direct Injection

1. False

2. False

3. True

OTDF → Life / Stress of HPT Nozzle Guide Vanes (NGVs) (Static components)

RTDF → Life / Stress of HPT Stage 1 Rotor Blades

Due to incomplete combustion and dissociation at high temperatures (along with Co2 & H2O):

• carbon monoxide CO

• nitrogen oxide NOx

• unburned hydro carbons UHC

• black carbon particles (soot)

Problem:

the highest amount of energy is released at stoichiometric mixture ratios

→ high temperatures → high amount of NOx is formed

Objective:

lower temperature of combustion / reduce residence time of gas in regions of high temperature

But:

reducing the turbine entry temperature leads to a drop in thermal efficiency of the process, therefore should be maintained/increased

Present research:

• the regulation of the mixing process

• as well as the combustion sequence

Non-premixed (RQL) arrangements:

Approach on NOx Reduction:

• high degree of fuel atomization when injected

• short residence time at stoichiometric high-temperature conditions

→ quick addition of a high amount of air mass-flow

→ quick change to lean mixture

-> -> “Rich–Quench–Lean”

Premixed(LPP) arrangements:

• fuel and air are mixed without combustion to form a lean mixture

• after this the combustion takes place

Problems of this concept:

• flame stabilization in a lean mixture

• avoiding flashbacks into the mixing zone and self-ignition

→ difficult to realize, especially in jet engines due to high variety of inlet conditions

Staged combustion chambers:

• two different conventional combustion zones are used for low and high power engine operation

• proper re-light ability is only granted in one of the two combustion zones

• low emissions and high degree of performance are achieved by the other combustion zone (potentially by pre-mixed arrangements)

Relative importance of combustion noise is increasingly growing

Two main mechanisms of combustion noise generation:

• Direct noise generated during heat release and by acoustic waves propagating to the engine outlet

• Indirect noise generated by acceleration of entropy/hot-spot and vorticity waves through turbine blades

FAR = 1/AFR

FAR = mfuel/m_air

Equivalence Raio = FAR/FAR_st

Jet fuel:

𝐹𝐴𝑅_𝑠𝑡 = 0,068

𝝓 < 𝟏: Lean Combustion

𝝓 > 𝟏: Rich Combustion

BPR = m19/m9

Pressure ratio

• The degree of reaction 𝜌ℎ defines the ratio of in the rotor static and overall stage static enthalpy rise:

• For 𝜌ℎ = 0,5 vectors of the velocity diagrams at rotor inlet and stator inlet get equal when being „mirrored about the axial direction

• For 𝜌ℎ = 1 the inlet and exit velocities of the stator are identical, which implies there is no static pressure rise across the stator