1. A method of mounting a drain valve in a drain opening in an aircraft skin, the method comprising:
applying a conductive coating to the aircraft skin adjacent the drain opening;
inserting a drain valve stem of the drain valve into the drain opening; and
securing the drain valve to the aircraft skin such that the conductive coating extends radially beyond an outer radial dimension of a drain valve flange of the drain valve.
2. The method of claim 1, further comprising applying a dielectric sealant to at least one of the aircraft skin and the drain valve flange of the drain valve.
3. The method of claim 1, further comprising attaching a conductive retaining bracket to the drain valve flange and the aircraft skin such that the drain valve flange is sandwiched between the conductive retaining bracket and the aircraft skin.
4. The method of claim 1, wherein the step of applying a conductive coating further comprises applying the conductive coating to an outer surface of the aircraft skin.
5. The method of claim 2, wherein the step of applying a dielectric sealant comprises applying the dielectric sealant at a location such that it is interposed between the aircraft skin and the drain valve flange.
6. A kit for installing at least one drain valve for a fuel tank in for an aircraft, the kit comprising:
a drain valve including a stem and a flange, wherein the flange is adapted to rest against a surface of the skin of an associated aircraft; and
a chemical conversion coating adapted to be applied to the skin for directing electric charge from a lightning strike away from the drain valve; and
a dielectric material adapted to be applied to at least one of the flange and the skin.
7. The kit of claim 6, wherein the dielectric material comprises a synthetic rubber sealant.
8. A method of mounting a drain valve in a drain valve opening in an aircraft skin, the method comprising:
applying a conductive coating material to the aircraft skin;
mounting the drain valve in a drain valve opening in the aircraft skin such that the conductive coating extends beyond a periphery of the drain valve.
9. The method of claim 8, wherein the step of applying a conductive coating comprises applying the coating circumferentially continuously around the drain valve opening.
10. The method of claim 8, further comprising applying a nonconductive sealant to at least one of the drain valve and the aircraft skin.
11. The method of claim 10 wherein the step of applying a nonconductive sealant further comprises applying the nonconductive sealant such that when the drain valve is mounted in the drain valve opening the nonconductive sealant is interposed between a surface of the aircraft skin and a portion of the drain valve.
12. The method of claim 11, wherein the nonconductive sealant is interposed between an outer surface of the aircraft skin and a drain valve flange of the drain valve.
13. The method of claim 8, wherein the step of applying a conductive coating material further comprises applying the conductive coating material to an outer surface of the aircraft skin.
14. In combination, an aircraft skin and a drain valve assembly mounted to the aircraft skin, the drain valve assembly comprising:
a drain valve including a stem and an integral flange; and
a conductive retaining ring attached to the aircraft skin and sandwiching at least a portion of the flange to the aircraft skin.
15. An assembly comprising:
an aircraft skin including a drain opening;
a drain valve including a drain valve flange, a portion of the drain valve being received in the drain opening; and
a conductive coating disposed on a surface of the aircraft skin adjacent the drain opening, the conductive coating extending beyond an outer radial dimension of the drain valve flange.
16. The assembly of claim 15, wherein the conductive coating comprises a chemical conversion coating.
17. The assembly of claim 15, wherein the drain valve flange is disposed adjacent an outer surface of aircraft skin.
18. The assembly of claim 15, further comprising a dielectric material disposed between the aircraft skin and the drain valve flange.
19. The assembly of claim 18, wherein the dielectric material is disposed between an outer surface of the aircraft skin and the drain valve flange.
20. The assembly of claim 15, wherein the conductive coating is disposed on an outer surface of the aircraft skin.
The claims below are in addition to those above.
All refrences to claim(s) which appear below refer to the numbering after this setence.
1. A system for conducting a working fluid and for receiving input heat from a heat source and for providing work out, the system comprising,
a capillary device for receiving the input heat from the heat source and phase changing and separating the working fluid from a liquid state to the first vapor state at a first vapor pressure,
a superheater for receiving the input heat and heating the working fluid in the first vapor state to a second vapor state, the working fluid in the first vapor state is heated into the second vapor state at a second vapor pressure at a superheated temperature in staggered amounts,
a turbine for converting thermal energy of the working fluid in the second vapor state into mechanical energy as the work out while converting the working fluid from the second vapor state to a third vapor state at the third vapor pressure, and
a condenser for phase changing the working fluid in the third vapor state into a liquid state while rejecting waste heat.
2. The system of claim 1 wherein,
the first vapor state is a saturated vapor state,
the second vapor state is a superheated vapor state, and
the third vapor state is a saturated vapor state.
3. The system of claim 1 wherein,
the capillary device is a loop heat pipe.
4. The system of claim 1 wherein,
the capillary device is a capillary pumped loop.
5. The system of claim 1 wherein,
the system is a power generator, and
the turbine converts energy by extracting thermal energy from the working fluid to produce output power as work out.
6. The system of claim 1 wherein,
the first vapor state has a first vapor pressure and a first temperature,
the second vapor state has a second vapor pressure and a second temperature,
the third vapor state has a third vapor pressure at a third temperature,
the first temperature is lower than the second temperature,
the second temperature is higher than the third temperature, and
the first vapor pressure is higher than the third vapor pressure.
7. The system of claim 1, wherein,
heat is radiated from the condenser.
8. The system of claim 1 wherein the superheater comprises,
stages respective comprising:
continuous control valves for receiving the working fluid from the capillary device;
heating chambers for superheating the staggered amounts of the working fluid; and
control valves for ejecting the staggered amounts of the working fluid, wherein each of the control valves are activated at staggered times for ejected the staggered amounts of the working fluid at staggered times in pulses.
9. The system of claim 1 wherein the superheater comprises,
stages respective comprising: continuous control valves for receiving the working fluid from the capillary device; heating chambers for superheating the staggered amounts of the working fluid; and control valves for ejecting the staggered amounts of the working fluid, wherein each of the control valves are activated at staggered times for ejected the staggered amounts of the working fluid at staggered times in pulses, and
a controller for controlling the control valves for providing the ejection of the staggered amounts of the working fluid at staggered times in pulse, the controller for controlling the control valve when temperatures of the working fluid in the heating changer are at predetermined values.
10. The system of claim 1 further comprising,
a vapor accumulator disposed between the capillary device and the superheater for continuously accumulating the working fluid from the capillary device and for dispensing the working fluid in incremental amounts respectively into the stages of the superheater for dampening pressure oscillations entering superheater.
11. The system of claim 1 further comprising,
a liquid pump for pressurizing the working fluid in the liquid state, and
a preheater for heating the working fluid in the liquid state.
12. The system of claim 1 wherein,
the superheater heat source and the capillary heat source are the same heat source.
13. The system of claim 1 wherein,
the superheater heat source and the capillary heat source and the preheater heat source are all separate heat sources where the heat source for the superheater is at a temperature higher than temperatures of the input heat from the capillary heat source and the preheater heat source.
14. The generator of claim 1 wherein,
the superheater heat source is selected from the group consisting of a radioisotope heat source, an active nuclear heat source, a solar heat source, and a waste heat source.
15. The generator of claim 1 wherein,
the capillary heat source is solar energy.
16. The generator of claim 1 wherein,
the capillary heat source and preheater heat source receive heat dissipated from spacecraft electronics.
17. The generator of claim 1 wherein,
the system is for powering a spacecraft,
the superheater heat source and the capillary heat source are selected from the group consisting of a radioisotope power system, or spacecraft electronics or solar radiation, and
the condenser radiates heat out for rejection of waste heat into outer space.
18. The generator of claim 1 wherein,
the heat input to the superheater comprises a thermal energy storage material.
19. The system of claim 1 further comprising,
a liquid pump for pressurizing the working fluid in the liquid state into a pressurized liquid state, the liquid pump being coupled to the condenser, and
a preheater for heating the working fluid in the pressurized liquid state into a heated pressurized liquid state, the preheater being coupled to the capillary device.
20. The system of claim 19 wherein,
the superheater heat source and the capillary heat source and the preheater heat source are the same heat source.