379
11.1 General Concept and Components
20, four stages suffice in the turbine part (enthalpy drop is about 980 kJ/kg). With
a compressor pressure ratio 40, five stages are needed (verify with the formulae of
Sect. 11.2).
The interior parts of the combustion chamber and the first stages of the turbine
are exposed to the hottest combustion gas. These parts therefore need cooling and
shielding from the hot gas. Heat shielding is done with a thermal barrier coating
(TBC). This is a layer of 100–400 μm thickness of ceramic material, formed by a
vapour deposition process of a mixture of zirconium and yttrium. It can withstand
very high temperatures and has an extremely low heat conduction coefficient (in the
order of 1 W/mK). This layer creates a temperature difference of 100–300 °C. The
ceramic layer is attached to a metal bond layer of about 100 μm thickness, formed
by an alloy of nickel, cobalt, chromium, aluminium and yttrium, which itself covers
the blade metal. The intermediate layer, or bond coat, serves as an elastic layer between the blade metal and the ceramic top coat. In between the bond coat and the top
coat, a small oxide layer grows due to the high temperature and this oxide has a role
in binding the two coats to each other. The white surface material of the blades of
the first three stages in Fig. 11.6 is the ceramic layer. The metal materials for combustors, turbine blades and vanes are alloys of nickel, chromium and cobalt (nickelbase super-alloys; melting temperature about 1350 °C). The blades of the first stage
are typically cast with temperature controlled cooling such that a blade forms a
single crystal. This very expensive technique is necessary for obtaining maximum
strength. The blades of the second stage are subjected to lower temperatures. These
blades are typically cast with a less delicate cooling technique in a directionally
solidified form, which means that the blade is composed of several crystals, but all
in the longitudinal direction. Blades of non-cooled stages may be cast conventionally, leading to a large number of crystals with borders in all directions.
Cooling of the first stages of the turbine part of a gas turbine (stator and rotor
parts of stages 1 and 2 and the stator of stage 3 of the SGT6-5000F in Fig. 11.2)
is normally realised by air bleeds on the compressor. This air is led to the interior
of the hollow stator vanes and rotor blades of the turbine, typically by external
conduits in a land-based gas turbine and by internal perforations in an aero gas turbine. For rotor cooling, the air has to be brought into perforations of the rotor discs
through slip ring sealed chambers. Three cooling principles are often combined.
Figure 11.7 (left) sketches sections through a rotor blade. First, there is convection
cooling by the air flowing through the internal passages. Heat transfer is enhanced
by ribs in the broader passages and by pins in the narrower ones. A large fraction of
the air is exhausted at the tip of the blade to provide cooling of the casing. A small
fraction is exhausted at the trailing edge of the blade. By perforations between the
front convection cooling channel and a channel at the leading edge, jets are formed
impinging at the internal side of the trailing edge. This way of cooling is called
impingement cooling. Impinging jets perpendicular to the surface realise a much
higher heat transfer than flow aligned to the surface, as with convection cooling.
This more intense cooling is very advantageous at the leading edge, where the heat
load is the largest. Film cooling may be added. The air is led to the blade surface by
holes, typically with an angle of about 30° to the surface, forming a protective film.
11.1 General Concept and Components
20, four stages suffice in the turbine part (enthalpy drop is about 980 kJ/kg). With
a compressor pressure ratio 40, five stages are needed (verify with the formulae of
Sect. 11.2).
The interior parts of the combustion chamber and the first stages of the turbine
are exposed to the hottest combustion gas. These parts therefore need cooling and
shielding from the hot gas. Heat shielding is done with a thermal barrier coating
(TBC). This is a layer of 100–400 μm thickness of ceramic material, formed by a
vapour deposition process of a mixture of zirconium and yttrium. It can withstand
very high temperatures and has an extremely low heat conduction coefficient (in the
order of 1 W/mK). This layer creates a temperature difference of 100–300 °C. The
ceramic layer is attached to a metal bond layer of about 100 μm thickness, formed
by an alloy of nickel, cobalt, chromium, aluminium and yttrium, which itself covers
the blade metal. The intermediate layer, or bond coat, serves as an elastic layer between the blade metal and the ceramic top coat. In between the bond coat and the top
coat, a small oxide layer grows due to the high temperature and this oxide has a role
in binding the two coats to each other. The white surface material of the blades of
the first three stages in Fig. 11.6 is the ceramic layer. The metal materials for combustors, turbine blades and vanes are alloys of nickel, chromium and cobalt (nickelbase super-alloys; melting temperature about 1350 °C). The blades of the first stage
are typically cast with temperature controlled cooling such that a blade forms a
single crystal. This very expensive technique is necessary for obtaining maximum
strength. The blades of the second stage are subjected to lower temperatures. These
blades are typically cast with a less delicate cooling technique in a directionally
solidified form, which means that the blade is composed of several crystals, but all
in the longitudinal direction. Blades of non-cooled stages may be cast conventionally, leading to a large number of crystals with borders in all directions.
Cooling of the first stages of the turbine part of a gas turbine (stator and rotor
parts of stages 1 and 2 and the stator of stage 3 of the SGT6-5000F in Fig. 11.2)
is normally realised by air bleeds on the compressor. This air is led to the interior
of the hollow stator vanes and rotor blades of the turbine, typically by external
conduits in a land-based gas turbine and by internal perforations in an aero gas turbine. For rotor cooling, the air has to be brought into perforations of the rotor discs
through slip ring sealed chambers. Three cooling principles are often combined.
Figure 11.7 (left) sketches sections through a rotor blade. First, there is convection
cooling by the air flowing through the internal passages. Heat transfer is enhanced
by ribs in the broader passages and by pins in the narrower ones. A large fraction of
the air is exhausted at the tip of the blade to provide cooling of the casing. A small
fraction is exhausted at the trailing edge of the blade. By perforations between the
front convection cooling channel and a channel at the leading edge, jets are formed
impinging at the internal side of the trailing edge. This way of cooling is called
impingement cooling. Impinging jets perpendicular to the surface realise a much
higher heat transfer than flow aligned to the surface, as with convection cooling.
This more intense cooling is very advantageous at the leading edge, where the heat
load is the largest. Film cooling may be added. The air is led to the blade surface by
holes, typically with an angle of about 30° to the surface, forming a protective film.