A rotor disk for a gas turbine engine is disclosed and formed to enable operation at high rotational speeds in a high temperature environment. The rotor disk is formed to include a bore, a live rim diameter and an outer diameter related to each other according to defined relationships.

A gas turbine engine of an aircraft includes: an engine core having a turbine including a lowest pressure rotor stage, a turbine diameter, a fan including a plurality of fan blades extending from a hub, an annular fan face at a leading edge of the fan defining a fan tip radius at the fan face; wherein a downstream blockage ratio is defined as: $\frac{\begin{array}{c}\mathrm{the}\mathrm{turbine}\mathrm{diameter}\mathrm{at}\mathrm{an}\mathrm{axial}\mathrm{location}\\ \mathrm{of}\mathrm{the}\mathrm{lowest}\mathrm{pressure}\mathrm{rotor}\mathrm{stage}\end{array}}{a\mathrm{distance}f\mathrm{rom}}$

A gas turbine engine of an aircraft includes: an engine core having a turbine including a lowest pressure rotor stage, a turbine diameter, a fan including a plurality of fan blades extending from a hub, an annular fan face at a leading edge of the fan; wherein a downstream blockage ratio of: $\frac{\begin{array}{c}\mathrm{the}\mathrm{turbine}\mathrm{diameter}\mathrm{at}\mathrm{an}\mathrm{axial}\\ \mathrm{location}\mathrm{of}\mathrm{the}\mathrm{lowest}\mathrm{pressure}\mathrm{rotor}\mathrm{stage}\end{array}}{a\mathrm{distance}\mathrm{from}a\mathrm{ground}}$

A gas turbine engine of an aircraft includes: an engine core having a turbine including a lowest pressure rotor stage, a turbine diameter, a fan including a plurality of fan blades extending from a hub, an annular fan face at a leading edge of the fan; wherein a downstream blockage ratio is:

[00001]$\frac{\begin{array}{c}\mathrm{the}.Math..Math.\mathrm{turbine}.Math..Math.\mathrm{diameter}.Math..Math.\mathrm{at}.Math..Math.\mathrm{an}.Math..Math.\mathrm{axial}.Math..Math.\mathrm{location}\\ \mathrm{of}.Math..Math.\mathrm{the}.Math..Math.\mathrm{lowest}.Math..Math.\mathrm{pressure}.Math..Math.\mathrm{rotor}.Math..Math.\mathrm{stage}\end{array}}{\mathrm{ground}.Math..Math.\mathrm{plane}.Math..Math.\mathrm{to}.Math..Math.\mathrm{wing}.Math..Math.\mathrm{distance}}$

and a quasi-non-dimensional mass flow rate Q defined as:

[00002]$Q=W.Math.\frac{\sqrt{T.Math.0}}{P.Math..Math.0.Math.A_{\mathrm{flow}}}$

where: W is mass flow rate through the fan in Kg/s; T0 is average stagnati

A gas turbine engine of an aircraft includes: an engine core having a turbine including a lowest pressure rotor stage, a turbine diameter, a fan including a plurality of fan blades extending from a hub, an annular fan face at a leading edge of the fan; wherein a downstream blockage ratio is:

[00001]$\frac{\begin{array}{c}\mathrm{the}.Math..Math.\mathrm{turbine}.Math..Math.\mathrm{diameter}.Math..Math.\mathrm{at}.Math..Math.\mathrm{an}.Math..Math.\mathrm{axial}\\ \mathrm{location}.Math..Math.\mathrm{of}.Math..Math.\mathrm{the}.Math..Math.\mathrm{lowest}.Math..Math.\mathrm{pressure}.Math..Math.\mathrm{rotor}.Math..Math.\mathrm{stage}\end{array}}{\mathrm{ground}.Math..Math.\mathrm{plane}.Math..Math.\mathrm{to}.Math..Math.\mathrm{wing}.Math..Math.\mathrm{distance}}$

and a quasi-non-dimensional mass flow rate Q defined as:

[00002]

A gas turbine engine of an aircraft includes: an engine core having a turbine including a lowest pressure rotor stage, a turbine diameter, a fan including a plurality of fan blades extending from a hub, an annular fan face at a leading edge of the fan; wherein a downstream blockage ratio is:

An aircraft gas turbine engine includes a compressor system including a low pressure compressor coupled to a low pressure shaft, and a high pressure compressor coupled to a high pressure shaft. An inner core casing is provided radially inwardly of compressor blades of the compressor system. A fan is coupled to the low pressure shaft via a reduction gearbox. The high pressure compressor includes an outer core casing arrangement provided radially outwardly of compressor blades of the high pressure compressor, the inner and outer core casing arrangements defining a core working gas flow path therebetween; and the outer core casing arrangement includes a first and a second outer core casing spaced radially outward from the first outer core casing. At an axial plane of an inlet to the high pressure compressor, the second outer core casing has a radius greater than 0.25 times a radius of the fan.

The present invention belongs to the technical field of engine modeling, and provides a methodology for modeling a starting process of a micro gas turbine engine, comprising the following steps: modeling micro gas turbine engine speed; modeling a relationship between micro gas turbine engine performance parameters and speed; and performing error analysis. Since most existing micro gas turbine engine modeling methodologies are methodologies using pure mechanism or machine learning, it is difficult to accurately describe the starting process of a micro gas turbine engine, and machine learning requires a lot of test data. In engineering practice, pure mechanisms commonly used at present are complicated in operation, low in efficiency, and low in modeling accuracy. The invention provides a methodology for modeling a starting process of a micro gas turbine engine based on a combination of a mechanism and an identification method, which makes up for the deficiencies of the prior art. The present invention is simple in operation and high in accuracy, and can implement modeling of the whole micro gas turbine engine. The methodology has certain extensibility and can be extended to other fields.

A rotor disk for a gas turbine engine is disclosed and formed to enable operation at high rotational speeds in a high temperature environment. The rotor disk is formed to include a bore, a live rim diameter and an outer diameter related to each other according to defined relationships.