9.E.3."Công nghệ" khác như sau:

Note 9.E.3. includes technology applicable to aero, aero-derivative, marine or industrial gas turbine engines.
a. "Technology" "required" for the "development" or "production" of any of the following gas turbine engine components or systems:
1. Turbine blades, vanes or "tip shrouds", made from directionally-solidified or single-crystal alloys and having (in the 001 Miller Index Direction) a stress-rupture life exceeding 400 hours at 1,273 K (1,000°C) at a stress of 200 MPa, based on the average property values;
Technical Note
For the purposes of 9.E.3.a.1., stress-rupture life testing is typically conducted on a test specimen.
a. 2. Combustors having any of the following:
a. 'Thermally decoupled liners' designed to operate at 'combustor exit temperature' exceeding 1,883K (1,610°C);
b. Non-metallic liners;
c. Non-metallic shells;
d. Liners designed to operate at 'combustor exit temperature' exceeding 1,883K (1,610°C) and having holes that meet the parameters specified in 9.E.3.c.; or
e. 'Pressure gain combustion';
Note The "required" "technology" for holes in 9.E.3.a.2. is limited to the derivation of the geometry and location of the holes.
Technical Notes
1. For the purposes of 9.E.3.a.2.a., 'thermally decoupled liners' are liners that feature at least a support structure designed to carry mechanical loads and a combustion facing structure designed to protect the support structure from the heat of combustion. The combustion facing structure and support structure have independent thermal displacement (mechanical displacement due to thermal load) with respect to one another, i.e., they are thermally decoupled.
2. For the purposes of 9.E.3.a.2.a. and 9.E.3.a.2.d., 'combustor exit temperature' is the bulk average gas path total (stagnation) temperature between the combustor exit plane and the leading edge of the turbine inlet guide vane (i.e., measured at engine station T40 as defined in SAE ARP 755A) when the engine is running in a "steady state mode" of operation at the certificated maximum continuous operating temperature.
3. For the purposes of 9.E.3.a.2.e., 'pressure gain combustion' refers to a combustor where the bulk average stagnation pressure at the combustor outlet is greater than the bulk average stagnation pressure at the combustor inlet due primarily to the combustion process, when the engine is running in a "steady state mode" of operation.
N.B. For "technology" "required" for manufacturing cooling holes, see 9.E.3.c.
a. 3. Components that are any of the following:
a. Manufactured from organic "composite" materials designed to operate above 588 K (315°C);
b. Manufactured from any of the following:
1. Metal "matrix" "composites" reinforced by any of the following:
a. Materials specified in 1.C.7.;
b. "Fibrous or filamentary materials" specified in 1.C.10.; or
c. Aluminides specified in 1.C.2.a.; or
2. Ceramic "matrix" "composites" specified in 1.C.7.; or
a. 3. c. Stators, vanes, blades, tip seals (shrouds), rotating blings, rotating blisks, or 'splitter ducts', that are all of the following:
1. Not specified in 9.E.3.a.3.a.;
2. Designed for compressors or fans; and
3. Manufactured from material specified in 1.C.10.e. with resins specified in 1.C.8.;
Technical Note
For the purposes of 9.E.3.a.3.c., a 'splitter duct' performs the initial separation of the air-mass flow between the bypass and core sections of the engine.
a. 4. Uncooled turbine blades, vanes or "tip shrouds", designed to operate at a 'gas path temperature' of 1,373 K (1,100°C) or more;
a. 5. Cooled turbine blades, vanes, "tip shrouds" other than those described in 9.E.3.a.1., designed to operate at a 'gas path temperature' of 1,693 K (1,420°C) or more;
Technical Note
For the purposes of 9.E.3.a.4. and 9.E.3.a.5., 'gas path temperature' is the bulk average gas path total (stagnation) temperature at the leading edge plane of the turbine component when the engine is running in a "steady state mode" of operation at the certificated or specified maximum continuous operating temperature.
a. 6. Aerofoil-to-disk blade combinations using solid-state joining;
7. Not used since 2018
8. 'Damage tolerant' rotor components using powder metallurgy materials specified in 1.C.2.b.; or
Technical Note
For the purposes of 9.E.3.a.8., 'damage tolerant' components are designed using methodology and substantiation to predict and limit crack growth.
9. Not used since 2009
N.B. For "FADEC systems", see 9.E.3.h.
10. Not used since 2010
N.B. For adjustable flow path geometry, see 9.E.3.i.
a. 11. 'Fan blades' having all of the following:
a. 20% or more of the total volume being one or more closed cavities containing vacuum or gas only; and
b. One or more closed cavities having a volume of 5 cm³ or larger;
Technical Note
For the purposes of 9.E.3.a.11., a 'fan blade' is the aerofoil portion of the rotating stage or stages, which provide both compressor and bypass flow in a gas turbine engine.
b. "Technology" "required" for the "development" or "production" of any of the following:
1. Wind tunnel aero-models equipped with non-intrusive sensors capable of transmitting data from the sensors to the data acquisition system; or
b. 2. "Composite" propeller blades or prop fans, capable of absorbing more than 2,000 kW at flight speeds exceeding Mach 0.55;
c. "Technology" "required" for manufacturing cooling holes in gas turbine engine components incorporating any of the "technologies" specified in 9.E.3.a.1., 9.E.3.a.2. or 9.E.3.a.5., and having any of the following:
1. Having all of the following:
a. Minimum 'cross-sectional area' less than 0.45 mm²;
b. 'Hole shape ratio' greater than 4.52; and
c. 'Incidence angle' equal to or less than 25°; or
2. Having all of the following:
a. Minimum 'cross-sectional area' less than 0.12 mm²;
b. 'Hole shape ratio' greater than 5.65; and
c. 'Incidence angle' more than 25°;
Note 9.E.3.c. does not apply to "technology" for manufacturing constant radius cylindrical holes that are straight through and enter and exit on the external surfaces of the component.
Technical Notes
For the purposes of 9.E.3.c.:
1. The 'cross-sectional area' is the area of the hole in the plane perpendicular to the hole axis.
2. 'Hole shape ratio' is the nominal length of the axis of the hole divided by the square root of its minimum 'cross-sectional area'.
3. 'Incidence angle' is the acute angle measured between the plane tangential to the aerofoil surface and the hole axis at the point where the hole axis enters the aerofoil surface.
4. Methods for manufacturing holes include "laser" beam machining, water jet machining, Electro-Chemical Machining (ECM) or Electrical Discharge Machining (EDM).
d. "Technology" "required" for the "development" or "production" of helicopter power transfer systems or tilt rotor or tilt wing "aircraft" power transfer systems;
e. "Technology" for the "development" or "production" of reciprocating diesel engine ground vehicle propulsion systems having all of the following:
1. 'Box volume' of 1.2 m³ or less;
2. An overall power output of more than 750 kW based on 80/1269/EEC, ISO 2534 or national equivalents; and
3. Power density of more than 700 kW/m³ of 'box volume';
Technical Note
For the purposes of 9.E.3.e., 'box volume' is the product of three perpendicular dimensions measured in the following way:
Length: The length of the crankshaft from front flange to flywheel face;
Width: The largest of any of the following:
a. The outside dimension from valve cover to valve cover;
b. The dimensions of the outside edges of the cylinder heads; or
c. The diameter of the flywheel housing;
Height: The largest of any of the following:
a. The dimension of the crankshaft centre-line to the top plane of the valve cover (or cylinder head) plus twice the stroke; or
b. The diameter of the flywheel housing.
f. "Technology" "required" for the "production" of specially designed components for "high output diesel engines", as follows:
1. "Technology" "required" for the "production" of engine systems having all of the following components employing ceramics materials specified in 1.C.7.:
a. Cylinder liners;
b. Pistons;
c. Cylinder heads; and
d. One or more other components (including exhaust ports, turbochargers, valve guides, valve assemblies or insulated fuel injectors);
f. 2. "Technology" "required" for the "production" of turbocharger systems with single-stage compressors and having all of the following:
a. Operating at pressure ratios of 4:1 or higher;
b. Mass flow in the range from 30 to 130 kg per minute; and
c. Variable flow area capability within the compressor or turbine sections;
f. 3. "Technology" "required" for the "production" of fuel injection systems with a specially designed multifuel (e.g., diesel or jet fuel) capability covering a viscosity range from diesel fuel (2.5 cSt at 310.8 K (37.8°C)) down to gasoline fuel (0.5 cSt at 310.8 K (37.8°C)) and having all of the following:
a. Injection amount in excess of 230 mm³ per injection per cylinder; and
b. Electronic control features specially designed for switching governor characteristics automatically depending on fuel property to provide the same torque characteristics by using the appropriate sensors;
g. "Technology" "required" for the "development" or "production" of "high output diesel engines" for solid, gas phase or liquid film (or combinations thereof) cylinder wall lubrication and permitting operation to temperatures exceeding 723 K (450°C), measured on the cylinder wall at the top limit of travel of the top ring of the piston;
h. "Technology" for gas turbine engine Full Authority Digital Engine Control "(FADEC) systems" as follows:
1. "Development" "technology" for deriving the functional requirements for the components necessary for the "FADEC system" to regulate engine thrust or shaft power (e.g., feedback sensor time constants and accuracies, fuel valve slew rate);
2. "Development" or "production" "technology" for control and diagnostic components unique to the "FADEC system" and used to regulate engine thrust or shaft power;
3. "Development" "technology" for the control law algorithms, including "source code", unique to the "FADEC system" and used to regulate engine thrust or shaft power;
Note 9.E.3.h. does not apply to technical data related to engine-"aircraft" integration required by civil aviation authorities of one or more Wassenaar Arrangement Participating States to be published for general airline use (e.g., installation manuals, operating instructions, instructions for continued airworthiness) or interface functions (e.g., input/output processing, airframe thrust or shaft power demand).
i. "Technology" for gas turbine engine adjustable flow path systems designed to maintain engine stability for gas generator turbines, fan or power turbines, or propelling nozzles, as follows:
1. "Development" "technology" for deriving the functional requirements for the components that maintain engine stability;
2. "Development" or "production" "technology" for components unique to the adjustable flow path system and that maintain engine stability;
3. "Development" "technology" for the control law algorithms, including "source code", unique to the adjustable flow path system and that maintain engine stability;
Note 9.E.3.i. does not apply to "technology" for any of the following:
a. Inlet guide vanes;
b. Variable pitch fans or prop-fans;
c. Variable compressor vanes;
d. Compressor bleed valves; or
e. Adjustable flow path geometry for reverse thrust.
j. "Technology" "required" for the "development" of wing-folding systems designed for fixed-wing "aircraft" powered by gas turbine engines;
N.B. For "technology" "required" for the "development" of wing-folding systems designed for fixed-wing "aircraft" specified in ML10., see ML22.
k. "Technology", not specified in 9.E.3.a., 9.E.3.h., or 9.E.3.i., "required" for the "development" of any of the following components or systems, specially designed for aero gas turbine engines to enable "aircraft" to cruise at Mach 1 or greater for more than 30 minutes:
1. Propulsion inlet systems;
2. Propulsion exhaust systems;
3. 'Reheat systems';
4. 'Active thermal management systems' to condition fluids used to lubricate or cool 'engine rotor supports';
5. Oil-free 'engine rotor supports'; or
6. Systems to remove heat from 'compression system' core gas path flow.
Technical Notes
For the purposes of 9.E.3.k.:
1. Propulsion inlet systems include core flow pre-coolers.
2. 'Reheat systems' provide additional thrust by combusting fuel in exhaust and/or bypass flow downstream of the last turbomachinery stage. 'Reheat systems' are also referred to as afterburners.
3. 'Active thermal management systems' employ methods other than passive oil-to-air cooling or oil-to-fuel cooling, such as vapour cycle systems.
4. 'Compression system' is any stage or combination of stages between the engine inlet face and the combustor that increases gas path pressure through mechanical work.
5. An 'engine rotor support' is the bearing supporting the main engine shaft that drives the compression system or turbine rotors.
N.B.1. For engine control technology, see 9.E.3.h.
N.B.2. For adjustable flow path systems technology, see 9.E.3.i.
[1] ^(*) The Russian Federation and Ukraine view this list as a reference list drawn up to help in the selection of dual-use goods which could contribute to the indigenous development, production or enhancement of conventional munitions capabilities.