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    COOLED FUEL INJECTOR SYSTEM FOR A GAS TURBINE ENGINE
    4.
    发明申请
    COOLED FUEL INJECTOR SYSTEM FOR A GAS TURBINE ENGINE 审中-公开
    用于燃气涡轮发动机的冷却燃料喷射器系统

    公开(公告)号：US20160273453A1

    公开(公告)日：2016-09-22

    申请号：US15029503

    申请日：2014-10-31

    Applicant: UNITED TECHNOLOGIES CORPORATION

    Inventor： John D. Frish ,  Alexander W. Williams ,  Sean D. Bradshaw

    IPC: F02C7/14 ,  F23R3/28 ,  F02C7/22

    CPC classification number: F02C7/14 ,  F02C7/222 ,  F05D2220/32 ,  F05D2260/231 ,  F23R3/283 ,  Y02T50/675

    Abstract: A cooling system for a fuel injector system of a gas turbine engine has a heat exchanger for cooling a portion of diffuser case air and then routing the cooled diffuser case air through a sleeve that surrounds a fuel injector conduit located in at least the diffuser case plenum for minimizing fuel heat-up rates in the conduit. By minimizing fuel temperatures within the injector conduit, coking accumulation is thereby eliminated or reduced.

    Abstract translation: 用于燃气涡轮发动机的燃料喷射器系统的冷却系统具有热交换器，用于冷却扩散器壳体空气的一部分，然后将冷却的扩散器壳体空气穿过围绕位于至少扩散器壳体增压室中的燃料喷射器管道的套筒 以最小化管道中的燃料加热速率。 通过最小化喷射器导管内的燃料温度，由此消除或减少焦化积聚。

    Combustor panels having angled rail
    9.
    发明授权
    Combustor panels having angled rail 有权

    公开(公告)号：US10260750B2

    公开(公告)日：2019-04-16

    申请号：US14982642

    申请日：2015-12-29

    Applicant: United Technologies Corporation

    Inventor： John S. Tu ,  James B. Hoke ,  Jonathan Jeffery Eastwood ,  Kevin Joseph Low ,  Sean D. Bradshaw

    IPC: F23R3/00 ,  F23R3/26 ,  F23R3/06

    Abstract: A combustor of a gas turbine engine including a combustor shell having an interior surface defining a combustion chamber, a first panel mounted to the interior surface at a first position, the first panel having a first surface and a first rail extending from the first surface toward the combustor shell, the first rail configured at a first angle relative to the first surface, and a second panel mounted to the interior surface at a second position axially adjacent to the first panel, the second panel having a second surface and a second rail extending from the second surface toward the combustor shell, the second rail configured at a second angle relative to the second surface. The first and second rails are proximal to each other and define a circumferential gap there between and at least one of the first or second angles is an acute angle.

    Turbine blade having film cooling hole arrangement
    10.
    发明授权
    Turbine blade having film cooling hole arrangement 有权

    公开(公告)号：US10060268B2

    公开(公告)日：2018-08-28

    申请号：US14965329

    申请日：2015-12-10

    Applicant: United Technologies Corporation

    Inventor： Sean D. Bradshaw

    IPC: F01D5/18 ,  F01D5/20

    CPC classification number: F01D5/186 ,  F01D5/20 ,  F05D2240/303 ,  F05D2240/307 ,  F05D2250/74 ,  F05D2260/202 ,  Y02T50/672 ,  Y02T50/673 ,  Y02T50/676

    Abstract: A turbine blade includes a platform that has a platform leading edge and trailing edge joined by two platform circumferential sides. An airfoil extends radially outwardly from the platform to a free tip end. The airfoil includes an airfoil leading edge and trailing edge joined by opposed pressure and suction sides. A root extends radially inwardly from the platform. The platform and the airfoil include film cooling holes that have external breakout points that are located in substantial conformance with the Cartesian coordinates set forth in Table 1. The Cartesian coordinates provided by an axial coordinate, a circumferential coordinate, and a radial coordinate, relative to a zero-coordinate, and the cooling holes have a diametrical surface tolerance relative to the specified coordinates of 0.200 inches (5.0 mm).

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