Editing Lockheed SR-71 Blackbird (section)

====Engine and nacelle====
{{main|Pratt & Whitney J58}}
The engine was an extensively re-designed version of the J58-P2, an existing supersonic engine which had run 700 development hours in support of proposals to power various aircraft for the US Navy. Only the compressor and turbine aerodynamics were retained. New design requirements for cruise at Mach 3.2 included:
* operating with very high ram temperature air entering the compressor, at {{Convert|800|F|C|sigfig=2}}
* a continuous turbine temperature capability {{Convert|450|F-change}} hotter than previous experience ([[Pratt & Whitney J75]])
* continuous use of maximum afterburning
* the use of new, more expensive, materials and fluids required to withstand unprecedented high temperatures

The engine was an afterburning turbojet for take-off and transonic flight (bleed bypass closed) and a low bypass augmented turbofan for supersonic acceleration (bleed bypass open). It approximated a ramjet during high speed supersonic cruise (with a pressure loss, compressor to exhaust, of 80% which was typical of a ramjet). It was a low bypass turbofan for subsonic loiter (bleed bypass open).<ref>https://authors.library.caltech.edu/records/6s4e6-b2j60, AE107_SR-71_Case_Study_321-450, p. 27.</ref><ref>A-12 Utility Flight Manual, 15 September 1965, changed 15 June 1968, 'Start Bleed And Bypass Valve Actuation', Figure 1-7</ref>

Analysis of the J58-P2 supersonic performance<ref name="roadrunnersinternationale.com">Bob Abernethy. https://www.roadrunnersinternationale.com/pw_tales.htm, 'More Never Told Tales of Pratt & Whitney'.</ref> showed the high compressor inlet temperature would have caused stalling, choking and blade breakages in the compressor as a result of operating at low corrected speeds on the compressor map. These problems were resolved by Pratt & Whitney engineer Robert Abernethy and are explained in his patent, "Recover Bleed Air Turbojet".<ref>{{cite web | url=https://patents.google.com/patent/US3344606A/en | title=Recover bleed air turbojet }}</ref> His solution was to 1) incorporate six air-bleed tubes, prominent on the outside of the engine, to transfer 20% of the compressor air to the afterburner, and 2) to modify the inlet guide vanes with a 2-position, trailing edge flap. The compressor bleed enabled the compressor to operate more efficiently and with the resulting increase in engine airflow matched the inlet design flow with an installed thrust increase of 47%.<ref name="roadrunnersinternationale.com"/><ref>William Brown. J58/SR-71 Propulsion Integration, attachment to CIA-RDP90B001170R000100050008-1, Fig. 3, 'Inlet and engine airflow match'.</ref> A continuous turbine temperature of {{Convert|2000|F|C}} was enabled with air-cooled first stage turbine vane and blades. Continuous operation of maximum afterburning was enabled by passing relatively cool air from the compressor along the inner surface of the duct and nozzle. Ceramic thermal barrier coatings were also used.

The secondary airflow through the nacelle comes from the cowl boundary layer bleed system which is oversized (flows more than boundary layer) to give a high enough pressure recovery to support the ejector pumping action. Additional air comes from the rear bypass doors and, for low speed operation with negligible inlet ram, from suck-in doors by the compressor case.

<gallery widths="220px" heights="165px" mode="packed" class="center" caption="Engine/nacelle">
File:Pratt & Whitney J58-JT11D-20K turbojet engine, 1962 - Lockheed SR-71A Blackbird, 1966 - Evergreen Aviation & Space Museum - McMinnville, Oregon - DSC01037.jpg|View of J58 engine shows some features required for flight at Mach 3.2: titanium inlet guide vanes and first stage compressor blades for lighter weight at high ram temperatures, transonic first stage compressor blades and low hub/tip ratio compressor entry, both scaled from the bigger Mach-3 J91 engine compressor, 2-position flaps on the inlet guide vanes and three of the six bypass tubes.
File:Pratt & Whitney J58 18.jpg|The afterburner was rated for continuous operation at {{cvt|3200|F|C|-2}} made possible with ceramic coatings (colored white) on duct liner and flame holders<ref>https://ntrs.nasa.gov/citations/20090018047, "History Of Thermal Barrier Coatings For Gas Turbine Engines: Emphasising NASA's Role from 1942 to 1990".</ref> and compressor bleed air cooling the duct and nozzle (above Mach 2.1 when the bleed was flowing). The nozzle is fully open, the maximum afterburning position. The main purpose of the variable nozzle area was to control engine operation which it did in conjunction with varying heat release in the afterburner.
File:Lockheed SR-71A Blackbird at Evergreen Aviation & Space Museum (6586637067).jpg|The inlet at left was depressed when the engine ran at high power settings with inadequate inlet ram (stationary and low flight speeds). The lower-than-ambient pressure in the inlet brought in extra air through the spike bleed and forward bypass louvers shown on the inlet external surface. Adequate secondary cooling air came in through the suck-in doors (shown open on the hinged nacelle).
</gallery>