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Lockheed L-1011 TriStar

The Lockheed L-1011 TriStar (pronounced "El-ten-eleven") is an American medium-to-long-range, wide-body trijet airliner built by the Lockheed Corporation. It was the third wide-body airliner to enter commercial operations, after the Boeing 747 and the McDonnell Douglas DC-10. The airliner has a seating capacity of up to 400 passengers and a range of over 4000 nmi. Its trijet configuration has three Rolls-Royce RB211 engines with one engine under each wing, along with a third engine center-mounted with an S-duct air inlet embedded in the tail and the upper fuselage. The aircraft has an autoland capability, an automated descent control system, and available lower deck galley and lounge facilities.

The L-1011 TriStar was produced in two fuselage lengths. The original L-1011-1 first flew in November 1970 and entered service with Eastern Air Lines in 1972. The shortened, longer range L-1011-500 first flew in 1978 and entered service with British Airways a year later. The original-length TriStar was also produced as the high gross weight L-1011-100, the up-rated engine L-1011-200, and the further upgraded L-1011-250. Post-production conversions for the L-1011-1 with increased takeoff weights included the L-1011-50 and L-1011-150.

The L-1011 TriStar's sales were hampered by two years of delays due to developmental and financial problems at Rolls-Royce, the sole manufacturer of the aircraft's engines. Between 1968 and 1984, Lockheed manufactured a total of 250 TriStars, assembled at the Lockheed plant located at the Palmdale Regional Airport in southern California north of Los Angeles. After L-1011 production ended, Lockheed withdrew from the commercial aircraft business due to its below-target sales. As of 2024, only one L-1011 is in service, as Stargazer.

Origins
In the 1960s, American Airlines approached Lockheed and competitor Douglas (later McDonnell Douglas) with the need for an airliner that could carry 250 passengers on transcontinental routes. Lockheed had not produced civilian airliners since 1961 with the L-188 Electra. In the 1950s the Electra was designed for turboprop propulsion, which Lockheed had successfully used on the C-130 Hercules military transport. Even after the Electra overcame vibration problems that caused several crashes early in its career, the market for large airliners would soon shift over to jet airliners such as the Boeing 707 and Douglas DC-8. Lockheed won contracts for jet military transports with the C-141 StarLifter, and pioneered very large jet transports with the large C-5 Galaxy with its high-bypass turbofan engines. Boeing lost the military contract, but its private-venture 747 captured what would become a much larger civilian airliner market for wide-body airliners.

Having experienced difficulties with some of its military programs, Lockheed was eager to re-enter the civilian market with a smaller wide-body jet, and its response was the L-1011 TriStar. Douglas Aircraft answered American Airlines with the DC-10, which had a similar three-engine configuration and dimensions. Despite their similarities, the L-1011 and DC-10's engineering approach differed greatly. McDonnell, who had recently taken over Douglas Aircraft, directed DC-10 development on a "very firm budget, and cost overruns were unacceptable – even at the expense of safety", and the conservative approach meant reusing Douglas DC-8 technology. By contrast, Lockheed would "take the most advanced technology of the day and when that technology was lacking, Lockheed created it" for the L-1011 in order to give it lower noise emissions, improved reliability, and higher efficiency over first-generation jet airliners. The TriStar name was selected in a Lockheed employee naming contest for the airliner. The advanced technology that went into the TriStar resulted in a high purchase price.

The TriStar's design featured a twin-aisle interior with a maximum of 400 passengers and a three-engine layout. The TriStar was originally conceived as a "jumbo twin", but a three-engine design was ultimately chosen to give the aircraft enough thrust to take off from existing runways. Also, before the establishment of Extended Operations standards by the FAA in the 1980s, commercial jets with only two engines were not allowed to fly more than 30 minutes away from an airport, making trans-oceanic flights impossible. The main visible difference between the TriStar and its similar trijet competitor, the McDonnell Douglas DC-10, is the central tail engine configuration: the DC-10's engine is mounted above the fuselage for simplicity of design and more economical construction, while the TriStar's engine is mounted to the rear fuselage and fed through an S-duct (similar to the Boeing 727) for reduced drag and improved stability. Lockheed engineers were able to maintain straight-through engine performance by limiting the curve of the S-duct to less than a quarter of the radius of the engine intake diameter. The S-duct design also reduced the total empty aircraft weight. The research undertaken during the design of the L-1011 indicated that losses of using an S-duct were more than compensated for by the above savings. A further major difference between the L-1011 and the DC-10 was Lockheed's selection of the Rolls-Royce RB211 as the only engine for the L-1011. As originally designed, the RB211 turbofan was an advanced three-spool design with a carbon fiber fan, which would have better efficiency and power-to-weight ratio than any competing engine like the General Electric CF6 that powered the DC-10. In theory, the triple spool would produce the same or more power as existing double spool engines while having a smaller cross section that would reduce drag.

American Airlines opted for the Douglas DC-10, although it showed considerable interest in the L-1011. American intended to convince Douglas to lower its price for the DC-10, which it did. Without the support of American, the TriStar was launched on orders from TWA and Eastern Air Lines.

Although the TriStar's design schedule closely followed that of its competitor, McDonnell Douglas beat Lockheed to market by a year due to delays in powerplant development. In February 1971, after massive development costs associated with the RB211, Rolls-Royce went into receivership. This halted L-1011 final assembly and Lockheed investigated the possibility of a US engine supplier. However the engineering was finalized at that stage in the TriStar's development and its S-duct, which was designed to fit the smaller cross-section of the triple spool RB-211 engine that would have reduced drag, was too small in diameter to accommodate an existing double spool engine. One option presented was potential outsource of RB-211 production to Canadian manufacturer Orenda Engines.

The British government agreed to approve a large state subsidy to restart Rolls-Royce operations on condition the U.S. government guarantee the bank loans Lockheed needed to complete the L-1011 project. Despite some opposition, not least from the then Governor of California, Ronald Reagan, the U.S. government provided these guarantees. For the rest of the RB211 project, Rolls-Royce remained a government-owned company.

Production

The TriStar's internal Lockheed model number is L-093. The TriStar was manufactured in Lockheed facilities in Burbank and Palmdale, California.

The prototype L1011 first flew on November 16, 1970. The L-1011 was certified on April 14, 1972, with the first airliner delivered to Eastern Air Lines on April 26, 1972. In 1972, its unit cost was US$20 million (~$ in ). To further publicize the new aircraft, an L-1011 was taken on a world tour during 1972 by famed Lockheed test pilot Tony LeVier. In a demonstration by test pilots LeVier and Charles Hall, 115 crew members, employees, and reporters embarked on the TriStar for a 4-hour, 13-minute flight from Palmdale to Dulles Airport "with the TriStar's AFCS [Automatic Flight Control System] feature engaged from takeoff roll to landing", and Lockheed touted it as "a groundbreaking moment: the first cross-country flight without the need for human hands on the controls".

Lockheed discovered fairly early on that the TriStar suffered from higher than estimated structural weight, engine weight, and specific fuel consumption. To rectify this problem and to meet performance guarantees, Lockheed developed a structural kit that allowed maximum takeoff weight (MTOW) to be increased on production aircraft from 409000 to 430000 lb. However, the weight problems affected the weight and desirability of early production L-1011-1 aircraft, known as Group 1 (serial numbers 1002 through to 1012).

Group 1 aircraft have an Operating empty weight (OEW) of 252700 lb, about 12700 lb higher than later aircraft, while Group 2 aircraft (serial numbers 1013 through 1051) have an OEW of 247000 lb, some 4700 lb lower. These aircraft, in general, also have different center of gravity envelopes with the forward center of gravity limit on the early aircraft being more restrictive at higher gross weights. Groups 1 and 2 aircraft (serial numbers 1002 to 1051) are upgradeable only to -50 or -150 specifications, although the Group 1 aircraft (up to serial number 1012) still maintain their operating disadvantages. All L-1011-1 aircraft from serial number 1052 onwards are Group 3 aircraft and are fully upgradeable to all variants up to -250 specification.

Costs at Rolls-Royce were controlled and its efforts largely went into the original TriStar engines, which needed considerable modifications between the L-1011's first flight and service entry. The competition, notably General Electric, was very quick to develop its CF6 engine with more thrust, which meant that a heavier "intercontinental" DC-10-30 could be more quickly brought to market. The flexibility afforded to potential customers by a long-range DC-10 put the L-1011 at a serious disadvantage. Rolls-Royce went on to develop the high-thrust RB211-524 for the L-1011-200 and -500, but this took many years.

The resultant delay in Lockheed and Rolls-Royce offering a high gross variant with a longer range, coupled with the TriStar's delayed introduction, meant that only 250 TriStars were sold compared to some 400 DC-10s. Lockheed needed to sell 500 airliners to break even, but in 1981, the company announced production would end with the delivery of the 250th and last L-1011 on order in 1984.

The TriStar's failure to achieve profitability caused Lockheed to withdraw from the civilian aircraft business. The TriStar's rivalry with the DC-10 has been seen as a "case study in what can happen when two manufacturers attempt to split a market that simply could not support both aircraft". Lockheed lacked the resources to follow up with several proposals based on the TriStar wing and airframe, including a wide-body twinjet and a stretched quad-jet (one of the quadjet proposals consisting of two underwing engines and two rear fuselage-mounted engines). McDonnell Douglas was also financially weakened and could only develop the MD-11, a refinement of the DC-10, instead of an all-new design to challenge the next generation of twinjets like the Boeing 777.

Engines
The RB211 and its features, despite the delays in its development, provided the L-1011 with then-unmatched fuel economy and noise levels.

Reversers and nozzle
During development the RB211 engines on the L-1011 had thrust-spoilers – target-type hot-stream reverse buckets – alongside the cold-bypass reversers. Despite capturing 25% of the total engine thrust, aerodynamic interference with the flaps diminished the braking effect of the flaps, so the thrust-spoilers were removed; instead an 11-degree afterbody was installed, which improved the specific range by 1.5%. Further improvements led to a 15-degree afterbody, enabling the L-1011 "to beat its predicted specific air range at 0.85 Mach by between 3.5 and 5.5 percent, the exact figure depending on cruise weight."

All-flying tail
Instead of the trimmable horizontal stabilizer (THS) found on most jetliners of its time, the L-1011 incorporated an all-flying tail – a stabilator. The aft portion had a geared (anti-servo) elevator that was linked to and moved with the stabilator, changing the stabilator's camber and improving the overall control surface effectiveness. Lockheed's main drive away from a THS was "[eliminating] mis-trim and runaway trim problems that have contributed to a number of accidents in the past." The fact that the elevators are not moved directly led to the failure in recognizing the jamming (trailing edge up) of the left elevator aboard Delta Air Lines Flight 1080 in 1977.

Fuel system
The L-1011-1 has four wing tanks; each inboard tank feeds the respective wing engine, and the two outboard tanks feed the tail engine via a flow equalizer. The additional center tank of the long-range variants, which is two halves, is located between the wing halves. Each center tank half acts as additional capacity to its adjacent inboard wing tank; filling it by way of ejector pumps. When the center tank is filled, it is used to crossfeed all three engines in flight (by way of the ejector pumps and crossfeed valves) until the center tank is empty and the remaining tanks are equalized.

Landing gear
The nose landing gear had two attachment points forward and aft, allowing a short-enough tug to push or pull the plane from directly underneath, a feature to allow operations where there wasn't enough forward space at some airports, which was more common at the time.

Electrical system and avionics
The L-1011 was the first jetliner to have an integrated drive generator (IDG).

The FMS on the L-1011, certified by the FAA in September 1977, offered many features that have since become common. The features were aimed at greatly reducing crew workload and improving fuel efficiency. Of those are a Mach/IAS cruise control, an automatic Rough Air Mode that detects turbulence and adjusts the engine power setting accordingly, and a descent mode that figures out the optimum location to start a descent by "back computing" from a preselected point, allowing "on-altitude and on-speed" arrival.

The L-1011 also featured a highly advanced autopilot system and was the first widebody to receive FAA certification for Cat-IIIc autolanding, which approved the TriStar for completely blind landings performed by the aircraft's autopilot in zero-visibility weather. The L-1011 used an inertial navigation system to navigate; this included aligning the navigation system by entering current coordinates of longitude and latitude.

It also had a unique direct lift control (DLC) system, which allowed for smooth approaches when landing, without having to use significant pitch changes while on the approach path. DLC helps maintain the aircraft on the glideslope during final approach by automatically deploying spoiler panels on the wings. Thus, rather than maintaining the descent by adjusting pitch, DLC helps control the descent while maintaining a more consistent pitch angle, using four redundant hydraulic systems. The production also used a unique "autoclave" system for bonding fuselage panels together; this made the L-1011 extremely resistant to corrosion.

Other components and systems
The APU was capable of operating up to 30,000 feet; its two square-shaped inlet doors are situated on the bottom fuselage on the aircraft's centerline towards the rear of the plane.

Compared to the typical three-system of its era, the L-1011 had four independent 3,000-psi hydraulic systems, A through D.

Commercial

TWA heralded the TriStar as one of the safest aircraft in the world in promotional literature in the 1980s when concern over the safety record of the McDonnell Douglas DC-10, flown by rival airlines, was at its peak. The L-1011 has been involved in five fatal accidents, only one of which was due to a problem with the aircraft.

Delta Air Lines was the type's largest customer. Delta retired its TriStars in 2001 to replace them with the Boeing 767-400ER. Cathay Pacific eventually became the largest non-U.S. operator of the type by acquiring many of the Eastern Air Lines examples when Eastern went bankrupt, operating as many as 21 aircraft. Cathay Pacific retired its L-1011s in October 1996 and replaced the type with the Airbus A330-300. TWA withdrew its last TriStar from service in 1997.

To secure the Japanese market, Lockheed secretly bribed several members of the Japanese government to subsidize All Nippon Airways' purchase of L-1011s; this caused a significant scandal when the bribes were uncovered. The discovered scale to what has become known as the Lockheed bribery scandal led to the arrest of Japanese Prime Minister Kakuei Tanaka, as well as several other officials. Within Lockheed, board chairman Daniel Haughton and vice chairman and president Carl Kotchian resigned their posts on February 13, 1976. Tanaka was eventually tried and found guilty of violating foreign exchange control laws but was not charged with bribery, a more serious criminal offense. Crucially for Lockheed, the fallout from the scandal included the loss of a contract worth over $1 billion.

The Soviet Union at that time lacked a widebody airliner. Development of its own Ilyushin Il-86 was delayed; consequently, in the mid-1970s, the Soviets started negotiations to buy 30 TriStars and license-produce up to 100 a year. The talks collapsed as US President Jimmy Carter made human rights an important consideration in US foreign policy. The TriStar was also listed by the Coordinating Committee as embodying advanced technology forbidden for sale to potential enemies, which presented a serious obstacle to the export deal.

The last three airlines to operate the L-1011 in scheduled service were Brussels Airlines (codeshare with Hewa Bora Airways), Thai Sky Airlines, and Lloyd Aereo Boliviano, with final flights in August 2007, February 2008, and May 2008, respectively. In later years the L-1011 has been used by smaller start-up carriers, particularly in Africa and Asia. These operators mainly do their business in the ad hoc charter and wet leasing businesses. ATA Airlines (formerly known as American Trans Air) fleet included over 19 TriStars, but operations dwindled to only three L-1011-500s before the company's shutdown in April 2008.

Military
The TriStar has also been used as a military tanker and passenger/cargo aircraft. The British Royal Air Force had nine aircraft of four variants. The aircraft were six ex-British Airways and three Pan Am L-1011-500s. All of the aircraft served with No. 216 Squadron, and were based at RAF Brize Norton. The TriStar was replaced in RAF service by the Airbus A330 MRTT under the Future Strategic Tanker Aircraft program. 216 Squadron was officially disbanded on March 20, 2014, and flew its last sorties with the TriStar on March 24, 2014.

Other

In the early 1990s, Orbital Sciences began to use a converted L-1011-100 named Stargazer to launch Pegasus rockets into orbit around Earth. This venture effectively rendered the small Scout rocket obsolete. This aircraft was also used in support of the X-34 and X-43 programs. NASA performed aerodynamic research on Orbital Sciences' L-1011 in 1995. In 2014, three L-1011s in the world were airworthy. As of 2019, Stargazer is the only active L-1011.

Variants
The earlier versions of the L-1011, such as the -1, -100, and -150 can be distinguished from the later models by the design of the middle engine nacelles. The earlier version nacelle has a round intake, whereas the later models have a small vertical fin between the bottom of the middle engine intake and the top of the fuselage.

The two L-1011 aircraft delivered to Pacific Southwest Airlines were configured with internal airstair doors that led into an entry hall in what was normally the forward lower baggage hold. This was to allow operations from airfields that did not have terminal buildings with jet bridges. These two aircraft were later in service with Aeroperú and Worldways Canada.

L-1011-1

The L-1011-1 (FAA certification L-1011-385-1) was the first production model of the L-1011, designed for short- and medium-range flights. This variant served as the basis for subsequent variants. This type was purchased by Air Canada, ANA, Cathay Pacific, Eastern, and other operators with regional trunk routes requiring a widebody aircraft. Pacific Southwest Airlines purchased two L-1011-1 models with lower deck seating. This variant was also one of the few widebodies to have the option for a full-height built-in airstair.

The L-1011-1 was first delivered to Eastern Air Lines on April 5, 1972. A total of 160 L-1011-1 TriStars were built before production ended in 1983, although the majority of these, 119 or 75% of the total, were completed during a four-year period from 1972 to 1975. Most sales of the L-1011-1 were to US operators, with just three airlines, Delta, Eastern, and TWA, taking delivery of 110 combined. A further two aircraft were placed with a fourth US airline, Pacific Southwest Airlines.

L-1011-100

The L-1011-100 (FAA certification L-1011-385-1-14) was the second production model of the L-1011 and first flew in 1975 and featured a new center fuel tank and higher gross weights that increased the aircraft's range by nearly 930 mi. Launch orders for the L-1011-100 were placed by Saudia and Cathay Pacific, for two each, in May 1974. Deliveries began in June 1975.

The variant was also purchased by several airlines with longer-range routes, such as TWA, Air Canada, and BEA (which merged with BOAC to form British Airways). The first two L-1011-100s (serial numbers 1110 and 1116) were delivered new to Saudia with the same fuel capacity as the L-1011-1 (FAA certification L-1011-385-1-14); these were later upgraded to L-1011-200 specification.

L-1011-50
The L-1011-50 was an upgraded version of the L-1011-1 with an increase in maximum takeoff weight from 430,000 lb to either 440,000 lb or 450,000 lb. Fuel capacity was not increased. The -50 was available only as a conversion package for the L-1011-1 and was never built new.

L-1011-150
The L-1011-150 was a development of the L-1011-1 with its maximum takeoff weight increased to 470000 lb. It was available only as a conversion for the L-1011-1. The -150 involves the conversion of Group 1 and Group 2 L-1011-1 aircraft to an MTOW of 470000 lb, an increase of 40000 lb, about 10%, from the L-1011-1, giving the aircraft a slightly better range than the -50, but without the additional center-section fuel tank, less than the L-1011-100 aircraft. The first aircraft was converted by MBB at Lemwerder in Germany during the winter of 1988–89 and was subsequently handed over to Air Transat of Canada on May 11, 1989.

L-1011-200

The L-1011-200 (FAA certification L-1011-385-1-15), the third production model of the L-1011, was introduced in 1976. Although otherwise similar to the -100, the -200 uses Rolls-Royce RB.211-524B engines to improve its performance in hot and high-altitude conditions. Gulf Air used -200 models to replace its earlier-generation Vickers VC10 fleet.

Other than the engines, the basic TriStar -200 is identical to the -100, with center-section fuel, having a MTOW of 466000 lb, and fuel capacity of 26400 usgal as the -100. An increase of gross weight to 474000 lb is possible, with the heavier aircraft offered by Lockheed as -200I or -200(Improved). Saudi Arabian Airlines (Saudia) was a launch customer for the -200 series and operated a sizable fleet until 1998. A total of 24 L-1011-200 aircraft were built new, with the first delivered to Saudia on May 28, 1977. Like other TriStar improvements, a conversion program has also been offered.

L-1011-250
The L-1011-250 was an upgrade developed for late-model L-1011-1 aircraft and all L-1011-100 and L-1011-200 aircraft. The more powerful engines, lengthened wing, active-load-control ailerons and other systems that had been developed for the L-1011-500 were adapted into the baseline model. The changes resulted in increases in maximum takeoff weight to 510000 lb and fuel capacity from 23,600 US gal (89,335 L) to 31,632 US gal (119,735 L). This variant also used the upgraded RB211-524B4I engine, which could be easily retrofitted to the existing RB211-524B powerplants of the L-1011-200, but it required a re-engining on the L-1011-1 and L-1011-100, which used the original RB211-22B. The conversion allowed the L-1011 to match the performance of the long-range McDonnell Douglas DC-10-30. Although it could be applied to all L-1011 models, the upgrade was only undertaken by Delta on six late-model L-1011-1 aircraft.

L-1011-500

The L-1011-500 (FAA certification L-1011-385-3) was the last L-1011 variant to enter production. It was a longer-range variant first flight-tested in 1978. Its fuselage length was shortened by 14 ft and MTOW increased to allow higher fuel loads. More powerful RB.211-524 engines, increased wingspan, active-load-control ailerons and other improved systems were features introduced by Lockheed to exploit newly available technologies in the late 1970s. The -500 variant was popular among international operators and formed a significant portion of the L-1011 fleet of Delta and British Airways. However, it entered service seven years after the similar DC-10-30 entered service.

The TriStar 500 first flew on October 16, 1978, with the first delivery to British Airways on April 27, 1979. It entered service with British Airways on May 7, 1979, flying between London and Abu Dhabi. The last L-1011 produced was a TriStar 500, operated by the Las Vegas Sands.

Dimensions
The TriStar 500 has an overall length of 164 ft and wingspan increased to 164 ft (early TriStar versions originally had the TriStar 1 wing with a span of 155 ft).

Flying surfaces
Lockheed developed some aerodynamic improvements for the TriStar 500 which included a modified wing-to-body fairing, a fillet below the central intake, extended wingtips, and "active ailerons" or active control system (ACS). The new fairing reduced drag, while the fillet reduced noise in the rear cabin. The wingtip extensions increased aspect ratio, thus reducing induced drag, but resulted in increased bending. The ACS developed to solve this, provided gust alleviation, improving ride during flight, reduced fuel burn, and increased fatigue life.

Earlier TriStar 500s were delivered with the standard wing; these were later retrofitted with ailerons and extended wingtips. Pan Am was the first customer to order the -500 with the extended wingtips and active ailerons. Aircraft serial number 1176, the first for Pan Am, was the first TriStar 500 to be fitted with them as standard.

Powerplant

The TriStar 500 is equipped with the more powerful RB211-524B engines. Initially rated at 50000 lbf thrust each, the higher-thrust 53000 lbf -524B4 Improved (also referred to as the -524B4I) later became available, which also offered improved fuel efficiency.

Performance
Originally certified with an MTOW of 496000 lb, an increased MTOW of 504000 lb was later certified in 1979, and all earlier production aircraft were certified at this weight. A further increase, to 510000 lb, is also available, and most TriStar 500s are thought to have had this increase. Standard fuel capacity is 31600 USgal, giving the TriStar 500 a range of about 5200 nmi with 246 passengers and baggage.

Cabin
The TriStar 500's maximum passenger capacity is 315, although no aircraft were operated with that number of seats. A typical two-class layout might include 21 First Class and 229 Economy Class for a maximum of 250 passengers. More spacious three-class layouts used on longer routes include 233 with 12 First Class, 32 Business Class, and 189 Economy Class with Delta Air Lines. The aircraft is equipped with six exits, two fewer than the long-body TriStars, thus reducing the exit limit maximum.

Operators
One L-1011 TriStar is in service as of 2024 is the Stargazer air-launch mothership, operated by Northrop Grumman Space Systems.

Accidents and incidents

As of December 2011, the L-1011 has been involved in 35 aviation occurrences, including 10 hull-losses, with 540 fatalities. Of the four pioneering widebody aircraft (Boeing 747, McDonnell Douglas DC-10, L-1011, and Airbus A300/A310 family), the Lockheed L-1011 had comparatively few accidents and a better safety record than its competitors.
 * On December 29, 1972, Eastern Air Lines Flight 401, an L-1011, crashed in the Florida Everglades as a result of the flight crew's failure to monitor the flight instruments during a malfunction of the landing gear position indicator system. The crash resulted in 101 fatalities, and was the subject of two TV movies, Crash and The Ghost of Flight 401. It was also the subject of a Mayday episode.
 * On April 12, 1977, Delta Air Lines Flight 1080, on takeoff from San Diego, had a left stabilizer jammed undetected in the full trailing-edge-up position. This failure resulted in a large noseup and rolling moment that almost exceeded the capability of the flight controls. The airplane was just about to stall in the clouds when Captain Jack McMahan, with unusual insight, reduced power on the wing engines and began using the throttles to supplement the remaining flight controls, using differential and collective engine thrust. Cabin crew moved all the passengers forward in the cabin to redistribute weight and help get the nose down. Steve Heidt, the flight engineer, said, "It probably didn't help much, but in that situation, we figured every little bit would help." All the way from San Diego to Los Angeles, the aircraft flew with its pitch controlled by differential thrust between tail and wing engines, while the left roll tendency was compensated by wing differential thrust, and made a successful emergency landing in Los Angeles. According to an incident analysis by Warren VanderBurgh, comprehensive crew training played a critical role in control recovery.
 * On August 19, 1980, a fire destroyed the L-1011-200 used for Saudia Flight 163 on the ground after the pilots made an emergency landing at Riyadh's former International Airport due to fire in the rear of the aircraft. Delays in initiating the evacuation of the aircraft led to the deaths of all 301 occupants.
 * On December 23, 1980, Saudia Flight 162, an L-1011, had a tire explode and penetrate the passenger cabin whilst in flight. The aircraft lost cabin pressure and two passengers were ejected through a hole in the cabin floor. The aircraft was later repaired and returned to service.
 * On September 22, 1981, Eastern Air Lines Flight 935 departed Newark, New Jersey, and suffered an uncontained failure of its number two (tail) engine at 14500 ft, while en route to San Juan, Puerto Rico. The fragments from that engine damaged three of its four hydraulic systems resulting in fluid loss in them. The rudder pedals also jammed. The fragments struck but did not puncture the lines for the other hydraulic system; the captain was able to safely land the aircraft at John F. Kennedy International Airport, with some limited use of the outboard spoilers, the inboard ailerons and the horizontal stabilizer, plus differential engine power of the remaining two engines. There were no injuries. The L-1011 having four hydraulic systems (instead of three like the DC-10) allowed for a safe landing.
 * On May 5, 1983, Eastern Air Lines Flight 855, L-1011 registration N334EA, while flying from Miami to Nassau, shut down the number 2 engine due to low oil pressure and began a return to Miami. Both of the remaining engines later failed. Without power, Flight 855 descended from 13,000 to 4000 ft before the number 2 engine was restarted and the aircraft landed in Miami without injuries. Incorrect engine maintenance had led to the loss of oil on all three engines.
 * On April 5, 1984, a Saudia Lockheed L-1011 TriStar on final approach to Damascus from Jeddah was hijacked by a Syrian national. The hijacker demanded to be taken to Istanbul, Turkey, but changed his mind and requested to go to Stockholm. After landing in Istanbul to refuel, the pilot pushed the hijacker out the emergency exit whereupon he was arrested.

 * On May 27, 1985, British Airtours Flight 101, registration G-BBAI, from Palma Airport, Mallorca, Spain overran the runway on landing at Leeds Bradford International Airport, West Yorkshire, United Kingdom. 12 of the 412 people on board suffered minor injuries when exiting down steep rear ramps. The aircraft was severely damaged.
 * On August 2, 1985, Delta Air Lines Flight 191, an L-1011, crashed while approaching Dallas/Fort Worth International Airport in microburst conditions. The crash killed eight of 11 crew members and 128 of the 152 passengers on board, as well as one person on the ground.
 * On October 18, 1985, a Jordanian Airlines L-1011 experienced an inflight fire at 24,000 feet while on approach to Singapore. The fire burned through the rear pressure bulkhead, causing explosive depressurization of the cabin. The air rushing out of the cabin extinguished the fire, saving the aircraft. All 118 passengers and crew survived. The aircraft was later repaired and placed back into service.
 * On May 3, 1986, Air Lanka Flight 512 (now SriLankan Airlines), an L-1011, was destroyed on the ground in Colombo, Sri Lanka, after a bomb exploded in the rear cargo hold, severing the tail and killing 21 people.
 * On June 28, 1991, LTU International L-1011 registration D-AERI, suffered an interior fire during maintenance in a hangar at Düsseldorf Airport. Four engineers escaped injury, but the aircraft was declared a total loss.
 * On July 30, 1992, TWA Flight 843, an L-1011, had its takeoff aborted by the captain after liftoff from John F. Kennedy International Airport, in response to a false stall warning. The aircraft landed too hard, breaking a wing spar and starting a fire. All 292 passengers and crew evacuated safely, with only 10 minor injuries. The airliner was destroyed by fire.
 * On August 23, 1995, Delta Air Lines Flight 157, an L-1011 TriStar 1, suffered a rapid decompression after the pressure bulkhead failed. The flight crew initiated an emergency descent to 14,000 feet, and the plane landed safely at Los Angeles International Airport with no deaths or injuries to the 226 passengers or 10 crew. However, the aircraft was substantially damaged and later written off.

Aircraft on display

 * N1011 – L-1011-1 on display at the Delta Flight Museum in Atlanta, Georgia. This is the forward upper fuselage of the prototype aircraft and is painted in Delta colors.
 * C-FTNA – L-1011-1 on display at the Lyon–Saint-Exupéry Airport in Lyon, France.  After Air Transat flight TSC906 was damaged in a hailstorm, the plane returned to Lyon and was written off. It is still used today for emergency training.
 * N31019 – L-1011-50 on display at the National Airline History Museum in Kansas City, Missouri, at the Charles B. Wheeler Downtown Airport.
 * TT-DWE – L-1011-100 on display at the Emirates National Auto Museum in Abu Dhabi, United Arab Emirates. This aircraft was originally delivered to British Airways.
 * HZ-AHP – L-1011-200 in Saudia livery on display as a gate guardian at the Royal Saudi Air Force Museum in Riyadh, Saudi Arabia.
 * 9Y-TGN – L-1011-500 on display at the Chaguaramas Military History and Aerospace Museum in Chaguaramas, Trinidad. This aircraft was previously operated by BWIA West Indies Airways.
 * N910TE – L-1011-1 on display with TriStar Experience at Kansas City International Airport. It was previously in storage in Tucson after being retired by the Flying Hospital group as P4-MED in favor of an MD-10. N910TE is the only L-1011-1 with working RB211-22b engines.
 * HS-AXE (Thai Sky Airlines), formerly N718DA (Delta) L-1011-1 was converted to a bar/restaurant located in Bangkok, Thailand.
 * 9Q-CHC (Hewa Bora Airways), MSN 1209 formerly C-GAGI (Air Canada) and N767DA (Delta) L-1011-385-3 TriStar 500 moved to Parc de la Vallée de la Nsele near Kinshasa, Democratic Republic of the Congo where it will be converted to a bar/restaurant.
 * CS-TMP (Luzair) MSN 1248 formerly JY-AGJ (Royal Jordanian) L-1011-385-3 TriStar 500 moved to Underwater Military Museum Dive Site, Aqaba, Jordan on 26 August 2019 for use as a tourist dive site.
 * N102CK (Gee Bee Airways), MSN 1198, in Kitty Hawk basic colours, formerly G-BHBM (British Airways) L-1011 TriStar 200F, stored at the apron of Kavala Airport LGKV/KVA - Greece, since November 2004.
 * N388LS, the former Las Vegas Sands L-1011-500, was damaged beyond repair in 2013 by floods while it was parked at Bangkok Airport. In 2018, the aircraft's body was disassembled and reassembled for display at Chic Chic Market, Nong Khai.

Popular culture

 * The band El Ten Eleven, a Los Angeles post-rock duo, derives its name from the L-1011.

title: Lockheed L-1011 TriStar


text:
Turbofan

A turbofan or fanjet is a type of airbreathing jet engine that is widely used in aircraft propulsion. The word "turbofan" is a combination of the preceding generation engine technology of the turbojet, and a reference to the additional fan stage added. It consists of a gas turbine engine which achieves mechanical energy from combustion, and a ducted fan that uses the mechanical energy from the gas turbine to force air rearwards. Thus, whereas all the air taken in by a turbojet passes through the combustion chamber and turbines, in a turbofan some of that air bypasses these components. A turbofan thus can be thought of as a turbojet being used to drive a ducted fan, with both of these contributing to the thrust.

The ratio of the mass-flow of air bypassing the engine core to the mass-flow of air passing through the core is referred to as the bypass ratio. The engine produces thrust through a combination of these two portions working together; engines that use more jet thrust relative to fan thrust are known as low-bypass turbofans, conversely those that have considerably more fan thrust than jet thrust are known as high-bypass. Most commercial aviation jet engines in use are of the high-bypass type, and most modern fighter engines are low-bypass. Afterburners are used on low-bypass turbofan engines with bypass and core mixing before the afterburner.

Modern turbofans have either a large single-stage fan or a smaller fan with several stages. An early configuration combined a low-pressure turbine and fan in a single rear-mounted unit.

Principles
The turbofan was invented to improve the fuel consumption of the turbojet. It achieves this by pushing more air, thus increasing the mass and lowering the speed of the propelling jet compared to that of the turbojet. This is done mechanically by adding a ducted fan rather than using viscous forces by adding an ejector, as first envisaged by Whittle.

Frank Whittle envisioned flight speeds of 500 mph in his March 1936 UK patent 471,368 "Improvements relating to the propulsion of aircraft", in which he describes the principles behind the turbofan, although not called as such at that time. While the turbojet uses the gas from its thermodynamic cycle as its propelling jet, for aircraft speeds below 500 mph there are two penalties to this design which are addressed by the turbofan.

Firstly, energy is wasted as the propelling jet is going much faster rearwards than the aircraft is going forwards, leaving a very fast wake. This wake contains kinetic energy that reflects the fuel used to produce it, rather than the fuel used to move the aircraft forwards. A turbofan harvests that wasted velocity and uses it to power a ducted fan that blows air in bypass channels around the rest of the turbine. This reduces the speed of the propelling jet while pushing more air, and thus more mass.

The other penalty is that combustion is less efficient at lower speeds. Any action to reduce the fuel consumption of the engine by increasing its pressure ratio or turbine temperature to achieve better combustion causes a corresponding increase in pressure and temperature in the exhaust duct which in turn cause a higher gas speed from the propelling nozzle (and higher KE and wasted fuel). Although the engine would use less fuel to produce a pound of thrust, more fuel is wasted in the faster propelling jet. In other words, the independence of thermal and propulsive efficiencies, as exists with the piston engine/propeller combination which preceded the turbojet, is lost. In contrast, Roth considers regaining this independence the single most important feature of the turbofan which allows specific thrust to be chosen independently of the gas generator cycle.

The working substance of the thermodynamic cycle is the only mass accelerated to produce thrust in a turbojet which is a serious limitation (high fuel consumption) for aircraft speeds below supersonic. For subsonic flight speeds the speed of the propelling jet has to be reduced because there is a price to be paid in producing the thrust. The energy required to accelerate the gas inside the engine (increase in kinetic energy) is expended in two ways, by producing a change in momentum ( i.e. a force), and a wake which is an unavoidable consequence of producing thrust by an airbreathing engine (or propeller). The wake velocity, and fuel burned to produce it, can be reduced and the required thrust still maintained by increasing the mass accelerated. A turbofan does this by transferring energy available inside the engine, from the gas generator, to a ducted fan which produces a second, additional mass of accelerated air.

The transfer of energy from the core to bypass air results in lower pressure and temperature gas entering the core nozzle (lower exhaust velocity) and fan-produced temperature and pressure entering the fan nozzle. The amount of energy transferred depends on how much pressure rise the fan is designed to produce (fan pressure ratio). The best energy exchange (lowest fuel consumption) between the two flows, and how the jet velocities compare, depends on how efficiently the transfer takes place which depends on the losses in the fan-turbine and fan.

The fan flow has lower exhaust velocity, giving much more thrust per unit energy (lower specific thrust). Both airstreams contribute to the gross thrust of the engine. The additional air for the bypass stream increases the ram drag in the air intake stream-tube, but there is still a significant increase in net thrust. The overall effective exhaust velocity of the two exhaust jets can be made closer to a normal subsonic aircraft's flight speed and gets closer to the ideal Froude efficiency. A turbofan accelerates a larger mass of air more slowly, compared to a turbojet which accelerates a smaller amount more quickly, which is a less efficient way to generate the same thrust (see the efficiency section below).

The ratio of the mass-flow of air bypassing the engine core compared to the mass-flow of air passing through the core is referred to as the bypass ratio. Engines with more jet thrust relative to fan thrust are known as low-bypass turbofans, those that have considerably more fan thrust than jet thrust are known as high-bypass. Most commercial aviation jet engines in use are high-bypass, and most modern fighter engines are low-bypass. Afterburners are used on low-bypass turbofans on combat aircraft.

Bypass ratio
The bypass ratio (BPR) of a turbofan engine is the ratio between the mass flow rate of the bypass stream to the mass flow rate entering the core. A bypass ratio of 6, for example, means that 6 times more air passes through the bypass duct than the amount that passes through the combustion chamber.

Turbofan engines are usually described in terms of BPR, which together with overall pressure ratio, turbine inlet temperature and fan pressure ratio are important design parameters. In addition BPR is quoted for turboprop and unducted fan installations because their high propulsive efficiency gives them the overall efficiency characteristics of very high bypass turbofans. This allows them to be shown together with turbofans on plots which show trends of reducing specific fuel consumption (SFC) with increasing BPR. BPR can also be quoted for lift fan installations where the fan airflow is remote from the engine and doesn't flow past the engine core.

Considering a constant core (i.e. fixed pressure ratio and turbine inlet temperature), core and bypass jet velocities equal and a particular flight condition (i.e. Mach number and altitude) the fuel consumption per lb of thrust (sfc) decreases with increase in BPR. At the same time gross and net thrusts increase, but by different amounts. There is considerable potential for reducing fuel consumption for the same core cycle by increasing BPR.This is achieved because of the reduction in pounds of thrust per lb/sec of airflow (specific thrust) and the resultant reduction in lost kinetic energy in the jets (increase in propulsive efficiency).

If all the gas power from a gas turbine is converted to kinetic energy in a propelling nozzle, the aircraft is best suited to high supersonic speeds. If it is all transferred to a separate big mass of air with low kinetic energy, the aircraft is best suited to zero speed (hovering). For speeds in between, the gas power is shared between a separate airstream and the gas turbine's own nozzle flow in a proportion which gives the aircraft performance required. The trade off between mass flow and velocity is also seen with propellers and helicopter rotors by comparing disc loading and power loading. For example, the same helicopter weight can be supported by a high power engine and small diameter rotor or, for less fuel, a lower power engine and bigger rotor with lower velocity through the rotor.

Bypass usually refers to transferring gas power from a gas turbine to a bypass stream of air to reduce fuel consumption and jet noise. Alternatively, there may be a requirement for an afterburning engine where the sole requirement for bypass is to provide cooling air. This sets the lower limit for BPR and these engines have been called "leaky" or continuous bleed turbojets (General Electric YJ-101 BPR 0.25) and low BPR turbojets (Pratt & Whitney PW1120). Low BPR (0.2) has also been used to provide surge margin as well as afterburner cooling for the Pratt & Whitney J58.

Efficiency

Propeller engines are most efficient for low speeds, turbojet engines for high speeds, and turbofan engines between the two. Turbofans are the most efficient engines in the range of speeds from about 500 to 1000 km/h, the speed at which most commercial aircraft operate.

In a turbojet (zero-bypass) engine, the high temperature and high pressure exhaust gas is accelerated when it undergoes expansion through a propelling nozzle and produces all the thrust. The compressor absorbs the mechanical power produced by the turbine. In a bypass design, extra turbines drive a ducted fan that accelerates air rearward from the front of the engine. In a high-bypass design, the ducted fan and nozzle produce most of the thrust. Turbofans are closely related to turboprops in principle because both transfer some of the gas turbine's gas power, using extra machinery, to a bypass stream leaving less for the hot nozzle to convert to kinetic energy. Turbofans represent an intermediate stage between turbojets, which derive all their thrust from exhaust gases, and turbo-props which derive minimal thrust from exhaust gases (typically 10% or less). Extracting shaft power and transferring it to a bypass stream introduces extra losses which are more than made up by the improved propulsive efficiency. The turboprop at its best flight speed gives significant fuel savings over a turbojet even though an extra turbine, a gearbox and a propeller are added to the turbojet's low-loss propelling nozzle. The turbofan has additional losses from its greater number of compressor stages/blades, fan and bypass duct.

Froude, or propulsive, efficiency can be defined as:

$$\eta_f = \frac {2}{1 + \frac {V_j}{V_a}}$$

where:

 * $Vj$ = thrust equivalent jet velocity
 * $Va$ = aircraft velocity

Thrust
While a turbojet engine uses all of the engine's output to produce thrust in the form of a hot high-velocity exhaust gas jet, a turbofan's cool low-velocity bypass air yields between 30% and 70% of the total thrust produced by a turbofan system.

The thrust (FN) generated by a turbofan depends on the effective exhaust velocity of the total exhaust, as with any jet engine, but because two exhaust jets are present the thrust equation can be expanded as:

$$F_N = \dot{m}_e v_{he} - \dot{m}_o v_o + BPR\, (\dot{m}_c) v_f$$

where:

 * $• me$ = the mass rate of hot combustion exhaust flow from the core engine
 * $• mo$ = the mass rate of total air flow entering the turbofan = $• mc + • mf$
 * $• mc$ = the mass rate of intake air that flows to the core engine
 * $• mf$ = the mass rate of intake air that bypasses the core engine
 * $vf$ = the velocity of the air flow bypassed around the core engine
 * $vhe$ = the velocity of the hot exhaust gas from the core engine
 * $vo$ = the velocity of the total air intake = the true airspeed of the aircraft
 * $BPR$ = bypass ratio

Nozzles
The cold duct and core duct's nozzle systems are relatively complex due to the use of two separate exhaust flows. In high bypass engines, the fan is situated in a short duct near the front of the engine and typically has a convergent cold nozzle, with the tail of the duct forming a low pressure ratio nozzle that under normal conditions will choke creating supersonic flow patterns around the core. The core nozzle is more conventional, but generates less of the thrust, and depending on design choices, such as noise considerations, may conceivably not choke. In low bypass engines the two flows may combine within the ducts, and share a common nozzle, which can be fitted with afterburner.

Noise
Most of the air flow through a high-bypass turbofan is lower-velocity bypass flow: even when combined with the much-higher-velocity engine exhaust, the average exhaust velocity is considerably lower than in a pure turbojet. Turbojet engine noise is predominately jet noise from the high exhaust velocity. Therefore, turbofan engines are significantly quieter than a pure-jet of the same thrust, and jet noise is no longer the predominant source. Turbofan engine noise propagates both upstream via the inlet and downstream via the primary nozzle and the by-pass duct. Other noise sources are the fan, compressor and turbine.

Fan noise may come from the interaction of the fan-blade wakes with the pressure field of the downstream fan-exit stator vanes. It may be minimized by adequate axial spacing between blade trailing edge and stator entrance. At high engine speeds, as at takeoff, shock waves from the supersonic fan tips, because of their unequal nature, produce noise of a discordant nature known as "buzz saw" noise.

All modern turbofan engines have acoustic liners in the nacelle to damp their noise. They extend as much as possible to cover the largest surface area. The acoustic performance of the engine can be experimentally evaluated by means of ground tests or in dedicated experimental test rigs.

In the aerospace industry, chevrons are the "saw-tooth" patterns on the trailing edges of some jet engine nozzles that are used for noise reduction. The shaped edges smooth the mixing of hot air from the engine core and cooler air flowing through the engine fan, which reduces noise-creating turbulence. Chevrons were developed by GE under a NASA contract. Some notable examples of such designs are Boeing 787 and Boeing 747-8 – on the Rolls-Royce Trent 1000 and General Electric GEnx engines.

History
Early turbojet engines were not very fuel-efficient because their overall pressure ratio and turbine inlet temperature were severely limited by the technology and materials available at the time.

The first turbofan engine, which was only run on a test bed, was the German Daimler-Benz DB 670, designated the 109-007 by the German RLM (Ministry of Aviation), with a first run date of 27 May 1943, after the testing of the turbomachinery using an electric motor, which had been undertaken on 1 April 1943. Development of the engine was abandoned with its problems unsolved, as the war situation worsened for Germany.

Later in 1943, the British ground tested the Metrovick F.3 turbofan, which used the Metrovick F.2 turbojet as a gas generator with the exhaust discharging into a close-coupled aft-fan module comprising a contra-rotating LP turbine system driving two co-axial contra-rotating fans.

Improved materials, and the introduction of twin compressors, such as in the Bristol Olympus, and Pratt & Whitney JT3C engines, increased the overall pressure ratio and thus the thermodynamic efficiency of engines. They also had poor propulsive efficiency, because pure turbojets have a high specific thrust/high velocity exhaust, which is better suited to supersonic flight.

The original low-bypass turbofan engines were designed to improve propulsive efficiency by reducing the exhaust velocity to a value closer to that of the aircraft. The Rolls-Royce Conway, the world's first production turbofan, had a bypass ratio of 0.3, similar to the modern General Electric F404 fighter engine. Civilian turbofan engines of the 1960s, such as the Pratt & Whitney JT8D and the Rolls-Royce Spey, had bypass ratios closer to 1 and were similar to their military equivalents.

The first Soviet airliner powered by turbofan engines was the Tupolev Tu-124 introduced in 1962. It used the Soloviev D-20. 164 aircraft were produced between 1960 and 1965 for Aeroflot and other Eastern Bloc airlines, with some operating until the early 1990s.

The first General Electric turbofan was the aft-fan CJ805-23, based on the CJ805-3 turbojet. It was followed by the aft-fan General Electric CF700 engine, with a 2.0 bypass ratio. This was derived from the General Electric J85/CJ610 turbojet 2,850 lbf to power the larger Rockwell Sabreliner 75/80 model aircraft, as well as the Dassault Falcon 20, with about a 50% increase in thrust to 4,200 lbf. The CF700 was the first small turbofan to be certified by the Federal Aviation Administration (FAA). There were at one time over 400 CF700 aircraft in operation around the world, with an experience base of over 10 million service hours. The CF700 turbofan engine was also used to train Moon-bound astronauts in Project Apollo as the powerplant for the Lunar Landing Research Vehicle.

Low-bypass turbofan

A high-specific-thrust/low-bypass-ratio turbofan normally has a multi-stage fan behind inlet guide vanes, developing a relatively high pressure ratio and, thus, yielding a high (mixed or cold) exhaust velocity. The core airflow needs to be large enough to ensure there is sufficient core power to drive the fan. A smaller core flow/higher bypass ratio cycle can be achieved by raising the inlet temperature of the high-pressure (HP) turbine rotor.

To illustrate one aspect of how a turbofan differs from a turbojet, comparisons can be made at the same airflow (to keep a common intake for example) and the same net thrust (i.e. same specific thrust). A bypass flow can be added only if the turbine inlet temperature is not too high to compensate for the smaller core flow. Future improvements in turbine cooling/material technology can allow higher turbine inlet temperature, which is necessary because of increased cooling air temperature, resulting from an overall pressure ratio increase.

The resulting turbofan, with reasonable efficiencies and duct loss for the added components, would probably operate at a higher nozzle pressure ratio than the turbojet, but with a lower exhaust temperature to retain net thrust. Since the temperature rise across the whole engine (intake to nozzle) would be lower, the (dry power) fuel flow would also be reduced, resulting in a better specific fuel consumption (SFC).

Some low-bypass ratio military turbofans (e.g. F404, JT8D) have variable inlet guide vanes to direct air onto the first fan rotor stage. This improves the fan surge margin (see compressor map).

Afterburning turbofan

Since the 1970s, most jet fighter engines have been low/medium bypass turbofans with a mixed exhaust, afterburner and variable area exit nozzle. An afterburner is a combustor located downstream of the turbine blades and directly upstream of the nozzle, which burns fuel from afterburner-specific fuel injectors. When lit, large volumes of fuel are burnt in the afterburner, raising the temperature of exhaust gases by a significant degree, resulting in a higher exhaust velocity/engine specific thrust. The variable geometry nozzle must open to a larger throat area to accommodate the extra volume and increased flow rate when the afterburner is lit. Afterburning is often designed to give a significant thrust boost for take off, transonic acceleration and combat maneuvers, but is very fuel intensive. Consequently, afterburning can be used only for short portions of a mission.

Unlike in the main engine, where stoichiometric temperatures in the combustor have to be reduced before they reach the turbine, an afterburner at maximum fuelling is designed to produce stoichiometric temperatures at entry to the nozzle, about 2100 K. At a fixed total applied fuel:air ratio, the total fuel flow for a given fan airflow will be the same, regardless of the dry specific thrust of the engine. However, a high specific thrust turbofan will, by definition, have a higher nozzle pressure ratio, resulting in a higher afterburning net thrust and, therefore, a lower afterburning specific fuel consumption (SFC). However, high specific thrust engines have a high dry SFC. The situation is reversed for a medium specific thrust afterburning turbofan: i.e., poor afterburning SFC/good dry SFC. The former engine is suitable for a combat aircraft which must remain in afterburning combat for a fairly long period, but has to fight only fairly close to the airfield (e.g. cross border skirmishes). The latter engine is better for an aircraft that has to fly some distance, or loiter for a long time, before going into combat. However, the pilot can afford to stay in afterburning only for a short period, before aircraft fuel reserves become dangerously low.

The first production afterburning turbofan engine was the Pratt & Whitney TF30, which initially powered the F-111 Aardvark and F-14 Tomcat. Low-bypass military turbofans include the Pratt & Whitney F119, the Eurojet EJ200, the General Electric F110, the Klimov RD-33, and the Saturn AL-31, all of which feature a mixed exhaust, afterburner and variable area propelling nozzle.

High-bypass turbofan

To further improve fuel economy and reduce noise, almost all jet airliners and most military transport aircraft (e.g., the C-17) are powered by low-specific-thrust/high-bypass-ratio turbofans. These engines evolved from the high-specific-thrust/low-bypass-ratio turbofans used in such aircraft in the 1960s. Modern combat aircraft tend to use low-bypass ratio turbofans, and some military transport aircraft use turboprops.

Low specific thrust is achieved by replacing the multi-stage fan with a single-stage unit. Unlike some military engines, modern civil turbofans lack stationary inlet guide vanes in front of the fan rotor. The fan is scaled to achieve the desired net thrust.

The core (or gas generator) of the engine must generate enough power to drive the fan at its rated mass flow and pressure ratio. Improvements in turbine cooling/material technology allow for a higher (HP) turbine rotor inlet temperature, which allows a smaller (and lighter) core, potentially improving the core thermal efficiency. Reducing the core mass flow tends to increase the load on the LP turbine, so this unit may require additional stages to reduce the average stage loading and to maintain LP turbine efficiency. Reducing core flow also increases bypass ratio. Bypass ratios greater than 5:1 are increasingly common; the Pratt & Whitney PW1000G, which entered commercial service in 2016, attains 12.5:1.

Further improvements in core thermal efficiency can be achieved by raising the overall pressure ratio of the core. Improvements in blade aerodynamics can reduce the number of extra compressor stages required, and variable geometry stators enable high-pressure-ratio compressors to work surge-free at all throttle settings.

The first (experimental) high-bypass turbofan engine was the AVCO-Lycoming PLF1A-2, a Honeywell T55 turboshaft-derived engine that was first run in February 1962. The PLF1A-2 had a 40 in geared fan stage, produced a static thrust of 4,320 lb, and had a bypass ratio of 6:1. The General Electric TF39 became the first production model, designed to power the Lockheed C-5 Galaxy military transport aircraft. The civil General Electric CF6 engine used a derived design. Other high-bypass turbofans are the Pratt & Whitney JT9D, the three-shaft Rolls-Royce RB211 and the CFM International CFM56; also the smaller TF34. More recent large high-bypass turbofans include the Pratt & Whitney PW4000, the three-shaft Rolls-Royce Trent, the General Electric GE90/GEnx and the GP7000, produced jointly by GE and P&W. The Pratt & Whitney JT9D engine was the first high bypass ratio jet engine to power a wide-body airliner.

The lower the specific thrust of a turbofan, the lower the mean jet outlet velocity, which in turn translates into a high thrust lapse rate (i.e. decreasing thrust with increasing flight speed). See technical discussion below, item 2. Consequently, an engine sized to propel an aircraft at high subsonic flight speed (e.g., Mach 0.83) generates a relatively high thrust at low flight speed, thus enhancing runway performance. Low specific thrust engines tend to have a high bypass ratio, but this is also a function of the temperature of the turbine system.

The turbofans on twin-engined transport aircraft produce enough take-off thrust to continue a take-off on one engine if the other engine shuts down after a critical point in the take-off run. From that point on the aircraft has less than half the thrust compared to two operating engines because the non-functioning engine is a source of drag. Modern twin engined airliners normally climb very steeply immediately after take-off. If one engine shuts down, the climb-out is much shallower, but sufficient to clear obstacles in the flightpath.

The Soviet Union's engine technology was less advanced than the West's, and its first wide-body aircraft, the Ilyushin Il-86, was powered by low-bypass engines. The Yakovlev Yak-42, a medium-range, rear-engined aircraft seating up to 120 passengers, introduced in 1980, was the first Soviet aircraft to use high-bypass engines.

Turbofan configurations
Turbofan engines come in a variety of engine configurations. For a given engine cycle (i.e., same airflow, bypass ratio, fan pressure ratio, overall pressure ratio and HP turbine rotor inlet temperature), the choice of turbofan configuration has little impact upon the design point performance (e.g., net thrust, SFC), as long as overall component performance is maintained. Off-design performance and stability is, however, affected by engine configuration.

The basic element of a turbofan is a spool, a single combination of fan/compressor, turbine and shaft rotating at a single speed. For a given pressure ratio, the surge margin can be increased by two different design paths:

 * 1) Splitting the compressor into two smaller spools rotating at different speeds, as with the Pratt & Whitney J57; or
 * 2) Making the stator vane pitch adjustable, typically in the front stages, as with the J79.

Most modern western civil turbofans employ a relatively high-pressure-ratio high-pressure (HP) compressor, with many rows of variable stators to control surge margin at low rpm. In the three-spool RB211/Trent the core compression system is split into two, with the IP compressor, which supercharges the HP compressor, being on a different coaxial shaft and driven by a separate (IP) turbine. As the HP compressor has a modest pressure ratio its speed can be reduced surge-free, without employing variable geometry. However, because a shallow IP compressor working line is inevitable, the IPC has one stage of variable geometry on all variants except the −535, which has none.

Single-shaft turbofan
Although far from common, the single-shaft turbofan is probably the simplest configuration, comprising a fan and high-pressure compressor driven by a single turbine unit, all on the same spool. The Snecma M53, which powers Dassault Mirage 2000 fighter aircraft, is an example of a single-shaft turbofan. Despite the simplicity of the turbomachinery configuration, the M53 requires a variable area mixer to facilitate part-throttle operation.

Aft-fan turbofan
One of the earliest turbofans was a derivative of the General Electric J79 turbojet, known as the CJ805-23, which featured an integrated aft fan/low-pressure (LP) turbine unit located in the turbojet exhaust jetpipe. Hot gas from the turbojet turbine exhaust expanded through the LP turbine, the fan blades being a radial extension of the turbine blades. This arrangement introduces an additional gas leakage path compared to a front-fan configuration and was a problem with this engine with higher-pressure turbine gas leaking into the fan airflow. An aft-fan configuration was later used for the General Electric GE36 UDF (propfan) demonstrator of the early 1980s.

In 1971 a concept was put forward by the NASA Lewis Research Center for a supersonic transport engine which operated as an aft-fan turbofan at take-off and subsonic speeds and a turbojet at higher speeds. This would give the low noise and high thrust characteristics of a turbofan at take-off, together with turbofan high propulsive efficiency at subsonic flight speeds. It would have the high propulsive efficiency of a turbojet at supersonic cruise speeds.

Basic two-spool

Many turbofans have at least basic two-spool configuration where the fan is on a separate low pressure (LP) spool, running concentrically with the compressor or high pressure (HP) spool; the LP spool runs at a lower angular velocity, while the HP spool turns faster and its compressor further compresses part of the air for combustion. The BR710 is typical of this configuration. At the smaller thrust sizes, instead of all-axial blading, the HP compressor configuration may be axial-centrifugal (e.g., CFE CFE738), double-centrifugal or even diagonal/centrifugal (e.g. Pratt & Whitney Canada PW600).

Boosted two-spool
Higher overall pressure ratios can be achieved by either raising the HP compressor pressure ratio or adding compressor (non-bypass) stages to the LP spool, between the fan and the HP compressor, to boost the latter. All of the large American turbofans (e.g. General Electric CF6, GE90, GE9X and GEnx plus Pratt & Whitney JT9D and PW4000) use booster stages. The Rolls-Royce BR715 is another example. The high bypass ratios used in modern civil turbofans tend to reduce the relative diameter of the booster stages, reducing their mean tip speed. Consequently, more booster stages are required to develop the necessary pressure rise.

Three-spool
Rolls-Royce chose a three-spool configuration for their large civil turbofans (i.e. the RB211 and Trent families), where the booster stages of a boosted two-spool configuration are separated into an intermediate pressure (IP) spool, driven by its own turbine. The first three-spool engine was the earlier Rolls-Royce RB.203 Trent of 1967.

The Garrett ATF3, powering the Dassault Falcon 20 business jet, has an unusual three spool layout with an aft spool not concentric with the two others.

Ivchenko Design Bureau chose the same configuration as Rolls-Royce for their Lotarev D-36 engine, followed by Lotarev/Progress D-18T and Progress D-436.

The Turbo-Union RB199 military turbofan also has a three-spool configuration, as do the military Kuznetsov NK-25 and NK-321.

Geared fan

As bypass ratio increases, the fan blade tip speed increases relative to the LPT blade speed. This will reduce the LPT blade speed, requiring more turbine stages to extract enough energy to drive the fan. Introducing a (planetary) reduction gearbox, with a suitable gear ratio, between the LP shaft and the fan enables both the fan and LP turbine to operate at their optimum speeds. Examples of this configuration are the long-established Garrett TFE731, the Honeywell ALF 502/507, and the recent Pratt & Whitney PW1000G.

Military turbofans
Most of the configurations discussed above are used in civilian turbofans, while modern military turbofans (e.g., Snecma M88) are usually basic two-spool.

High-pressure turbine
Most civil turbofans use a high-efficiency, 2-stage HP turbine to drive the HP compressor. The CFM International CFM56 uses an alternative approach: a single-stage, high-work unit. While this approach is probably less efficient, there are savings on cooling air, weight and cost.

In the RB211 and Trent 3-spool engine series, the HP compressor pressure ratio is modest so only a single HP turbine stage is required. Modern military turbofans also tend to use a single HP turbine stage and a modest HP compressor.

Low-pressure turbine
Modern civil turbofans have multi-stage LP turbines (anywhere from 3 to 7). The number of stages required depends on the engine cycle bypass ratio and the boost (on boosted two-spools). A geared fan may reduce the number of required LPT stages in some applications. Because of the much lower bypass ratios employed, military turbofans require only one or two LP turbine stages.

Cycle improvements
Consider a mixed turbofan with a fixed bypass ratio and airflow. Increasing the overall pressure ratio of the compression system raises the combustor entry temperature. Therefore, at a fixed fuel flow there is an increase in (HP) turbine rotor inlet temperature. Although the higher temperature rise across the compression system implies a larger temperature drop over the turbine system, the mixed nozzle temperature is unaffected, because the same amount of heat is being added to the system. There is, however, a rise in nozzle pressure, because overall pressure ratio increases faster than the turbine expansion ratio, causing an increase in the hot mixer entry pressure. Consequently, net thrust increases, whilst specific fuel consumption (fuel flow/net thrust) decreases. A similar trend occurs with unmixed turbofans.

Turbofan engines can be made more fuel efficient by raising overall pressure ratio and turbine rotor inlet temperature in unison. However, better turbine materials or improved vane/blade cooling are required to cope with increases in both turbine rotor inlet temperature and compressor delivery temperature. Increasing the latter may require better compressor materials.

The overall pressure ratio can be increased by improving fan (or) LP compressor pressure ratio or HP compressor pressure ratio. If the latter is held constant, the increase in (HP) compressor delivery temperature (from raising overall pressure ratio) implies an increase in HP mechanical speed. However, stressing considerations might limit this parameter, implying, despite an increase in overall pressure ratio, a reduction in HP compressor pressure ratio.

According to simple theory, if the ratio of turbine rotor inlet temperature/(HP) compressor delivery temperature is maintained, the HP turbine throat area can be retained. However, this assumes that cycle improvements are obtained, while retaining the datum (HP) compressor exit flow function (non-dimensional flow). In practice, changes to the non-dimensional speed of the (HP) compressor and cooling bleed extraction would probably make this assumption invalid, making some adjustment to HP turbine throat area unavoidable. This means the HP turbine nozzle guide vanes would have to be different from the original. In all probability, the downstream LP turbine nozzle guide vanes would have to be changed anyway.

Thrust growth
Thrust growth is obtained by increasing core power. There are two basic routes available:
 * 1) hot route: increase HP turbine rotor inlet temperature
 * 2) cold route: increase core mass flow

Both routes require an increase in the combustor fuel flow and, therefore, the heat energy added to the core stream.

The hot route may require changes in turbine blade/vane materials or better blade/vane cooling. The cold route can be obtained by one of the following:

 * 1) adding booster stages to the LP/IP compression
 * 2) adding a zero-stage to the HP compression
 * 3) improving the compression process, without adding stages (e.g. higher fan hub pressure ratio)

all of which increase both overall pressure ratio and core airflow.

Alternatively, the core size can be increased, to raise core airflow, without changing overall pressure ratio. This route is expensive, since a new (upflowed) turbine system (and possibly a larger IP compressor) is also required.

Changes must also be made to the fan to absorb the extra core power. On a civil engine, jet noise considerations mean that any significant increase in take-off thrust must be accompanied by a corresponding increase in fan mass flow (to maintain a T/O specific thrust of about 30 lbf/lb/s).

Technical discussion

 * 1) Specific thrust (net thrust/intake airflow) is an important parameter for turbofans and jet engines in general. Imagine a fan (driven by an appropriately sized electric motor) operating within a pipe, which is connected to a propelling nozzle. It is fairly obvious, the higher the fan pressure ratio (fan discharge pressure/fan inlet pressure), the higher the jet velocity and the corresponding specific thrust. Now imagine we replace this set-up with an equivalent turbofan – same airflow and same fan pressure ratio. Obviously, the core of the turbofan must produce sufficient power to drive the fan via the low-pressure (LP) turbine. If we choose a low (HP) turbine inlet temperature for the gas generator, the core airflow needs to be relatively high to compensate. The corresponding bypass ratio is therefore relatively low. If we raise the turbine inlet temperature, the core airflow can be smaller, thus increasing bypass ratio. Raising turbine inlet temperature tends to increase thermal efficiency and, therefore, improve fuel efficiency.
 * 2) Naturally, as altitude increases, there is a decrease in air density and, therefore, the net thrust of an engine. There is also a flight speed effect, termed thrust lapse rate. Consider the approximate equation for net thrust again:$$F_n = m \cdot (V_{jfe} - V_a).$$ With a high specific thrust (e.g., fighter) engine, the jet velocity is relatively high, so intuitively one can see that increases in flight velocity have less of an impact upon net thrust than a medium specific thrust (e.g., trainer) engine, where the jet velocity is lower. The impact of thrust lapse rate upon a low specific thrust (e.g., civil) engine is even more severe. At high flight speeds, high-specific-thrust engines can pick up net thrust through the ram rise in the intake, but this effect tends to diminish at supersonic speeds because of shock wave losses.
 * 3) Thrust growth on civil turbofans is usually obtained by increasing fan airflow, thus preventing the jet noise becoming too high. However, the larger fan airflow requires more power from the core. This can be achieved by raising the overall pressure ratio (combustor inlet pressure/intake delivery pressure) to induce more airflow into the core and by increasing turbine inlet temperature. Together, these parameters tend to increase core thermal efficiency and improve fuel efficiency.
 * 4) Some high-bypass-ratio civil turbofans use an extremely low area ratio (less than 1.01), convergent-divergent, nozzle on the bypass (or mixed exhaust) stream, to control the fan working line. The nozzle acts as if it has variable geometry. At low flight speeds the nozzle is unchoked (less than a Mach number of unity), so the exhaust gas speeds up as it approaches the throat and then slows down slightly as it reaches the divergent section. Consequently, the nozzle exit area controls the fan match and, being larger than the throat, pulls the fan working line slightly away from surge. At higher flight speeds, the ram rise in the intake increases nozzle pressure ratio to the point where the throat becomes choked (M=1.0). Under these circumstances, the throat area dictates the fan match and, being smaller than the exit, pushes the fan working line slightly towards surge. This is not a problem, since fan surge margin is much better at high flight speeds.
 * 5) The off-design behaviour of turbofans is illustrated under compressor map and turbine map.
 * 6) Because modern civil turbofans operate at low specific thrust, they require only a single fan stage to develop the required fan pressure ratio. The desired overall pressure ratio for the engine cycle is usually achieved by multiple axial stages on the core compression. Rolls-Royce tend to split the core compression into two with an intermediate pressure (IP) supercharging the HP compressor, both units being driven by turbines with a single stage, mounted on separate shafts. Consequently, the HP compressor need develop only a modest pressure ratio (e.g., ~4.5:1). US civil engines use much higher HP compressor pressure ratios (e.g., ~23:1 on the General Electric GE90) and tend to be driven by a two-stage HP turbine. Even so, there are usually a few IP axial stages mounted on the LP shaft, behind the fan, to further supercharge the core compression system. Civil engines have multi-stage LP turbines, the number of stages being determined by the bypass ratio, the amount of IP compression on the LP shaft and the LP turbine blade speed.
 * 7) Because military engines usually have to be able to fly very fast at sea level, the limit on HP compressor delivery temperature is reached at a fairly modest design overall pressure ratio, compared with that of a civil engine. Also the fan pressure ratio is relatively high, to achieve a medium to high specific thrust. Consequently, modern military turbofans usually have only 5 or 6 HP compressor stages and require only a single-stage HP turbine. Low-bypass-ratio military turbofans usually have one LP turbine stage, but higher bypass ratio engines need two stages. In theory, by adding IP compressor stages, a modern military turbofan HP compressor could be used in a civil turbofan derivative, but the core would tend to be too small for high thrust applications.

Aerodynamic modelling
Aerodynamics is a mix of subsonic, transonic and supersonic airflow on a single fan/gas compressor blade in a modern turbofan. The airflow past the blades must be maintained within close angular limits to keep the air flowing against an increasing pressure. Otherwise air will be rejected back out of the intake.

The Full Authority Digital Engine Control (FADEC) needs accurate data for controlling the engine. The critical turbine inlet temperature (TIT) is too harsh an environment, at 1700 °C and 17 bar, for reliable sensors. Therefore, during development of a new engine type a relation is established between a more easily measured temperature like exhaust gas temperature and the TIT. Monitoring the exhaust gas temperature is then used to make sure the engine does not run too hot.

Blade technology
A 100 g turbine blade is subjected to 1700 °C, at 17 bar and a centrifugal force of 40 kN, well above the point of plastic deformation and even above the melting point. Exotic alloys, sophisticated air cooling schemes and special mechanical design are needed to keep the physical stresses within the strength of the material. Rotating seals must withstand harsh conditions for 10 years, 20,000 missions and rotating at 10 to 20,000 rpm.

Fan blades
Fan blades have been growing as jet engines have been getting bigger: each fan blade carries the equivalent of nine double-decker buses and swallows air the equivalent volume of a squash court every second. Advances in computational fluid dynamics (CFD) modelling have permitted complex, 3D curved shapes with very wide chord, keeping the fan capabilities while minimizing the blade count to lower costs. Coincidentally, the bypass ratio grew to achieve higher propulsive efficiency and the fan diameter increased.

Rolls-Royce pioneered the hollow, titanium wide-chord fan blade in the 1980s for aerodynamic efficiency and foreign object damage resistance in the RB211 then for the Trent. GE Aviation introduced carbon fiber composite fan blades on the GE90 in 1995, manufactured since 2017 with a carbon-fiber tape-layer process. GE partner Safran developed a 3D woven technology with Albany Composites for the CFM56 and CFM LEAP engines.

Future progress
Engine cores are shrinking as they operate at higher pressure ratios and become more efficient and smaller compared to the fan as bypass ratios increase. Blade tip clearances are more difficult to maintain at the exit of the high-pressure compressor where blades are 0.5 in high or less; backbone bending further affects clearance control as the core is proportionately longer and thinner and the fan to low-pressure turbine driveshaft space is constrained within the core.

Pratt & Whitney VP technology and environment Alan Epstein argued "Over the history of commercial aviation, we have gone from 20% to 40% [cruise efficiency], and there is a consensus among the engine community that we can probably get to 60%".

Geared turbofans and further fan pressure ratio reductions may continue to improve propulsive efficiency. The second phase of the FAA's Continuous Lower Energy, Emissions and Noise (CLEEN) program is targeting for the late 2020s reductions of 33% fuel burn, 60% emissions and 32 dB EPNdb noise compared with the 2000s state-of-the-art. In summer 2017 at NASA Glenn Research Center in Cleveland, Ohio, Pratt has finished testing a very-low-pressure-ratio fan on a PW1000G, resembling an open rotor with fewer blades than the PW1000G's 20.

The weight and size of the nacelle would be reduced by a short duct inlet, imposing higher aerodynamic turning loads on the blades and leaving less space for soundproofing, but a lower-pressure-ratio fan is slower. UTC Aerospace Systems Aerostructures will have a full-scale ground test in 2019 of its low-drag Integrated Propulsion System with a thrust reverser, improving fuel burn by 1% and with 2.5-3 EPNdB lower noise.

Safran expects to deliver another 10–15% in fuel efficiency through the mid-2020s before reaching an asymptote, and next will have to increase the bypass ratio to 35:1 instead of 11:1 for the CFM LEAP. It is demonstrating a counterrotating open rotor unducted fan (propfan) in Istres, France, under the European Clean Sky technology program. Modeling advances and high specific strength materials may help it succeed where previous attempts failed. When noise levels are within existing standards and similar to the LEAP engine, 15% lower fuel burn will be available and for that Safran is testing its controls, vibration and operation, while airframe integration is still challenging.

For GE Aviation, the energy density of jet fuel still maximises the Breguet range equation and higher pressure ratio cores; lower pressure ratio fans, low-loss inlets and lighter structures can further improve thermal, transfer and propulsive efficiency. Under the U.S. Air Force's Adaptive Engine Transition Program, adaptive thermodynamic cycles will be used for the sixth-generation jet fighter, based on a modified Brayton cycle and Constant volume combustion. Additive manufacturing in the advanced turboprop will reduce weight by 5% and fuel burn by 20%.

Rotating and static ceramic matrix composite (CMC) parts operates 500 °F hotter than metal and are one-third its weight. With $21.9 million from the Air Force Research Laboratory, GE is investing $200 million in a CMC facility in Huntsville, Alabama, in addition to its Asheville, North Carolina site, mass-producing silicon carbide matrix with silicon-carbide fibers in 2018. CMCs will be used ten times more by the mid-2020s: the CFM LEAP requires 18 CMC turbine shrouds per engine and the GE9X will use it in the combustor and for 42 HP turbine nozzles.

Rolls-Royce Plc aim for a 60:1 pressure ratio core for the 2020s Ultrafan and began ground tests of its 100,000 hp gear for 100,000 lbf and 15:1 bypass ratios. Nearly stoichiometric turbine entry temperature approaches the theoretical limit and its impact on emissions has to be balanced with environmental performance goals. Open rotors, lower pressure ratio fans and potentially distributed propulsion offer more room for better propulsive efficiency. Exotic cycles, heat exchangers and pressure gain/constant volume combustion may improve thermodynamic efficiency. Additive manufacturing could be an enabler for intercooler and recuperators. Closer airframe integration and hybrid or electric aircraft can be combined with gas turbines.

Rolls-Royce engines have a 72–82% propulsive efficiency and 42–49% thermal efficiency for a 0.63 - 0.49 lb/lbf/h TSFC at Mach 0.8, and aim for theoretical limits of 95% for open rotor propulsive efficiency and 60% for thermal efficiency with stoichiometric turbine entry temperature and 80:1 overall pressure ratio for a 0.35 lb/lbf/h TSFC

As teething troubles may not show up until several thousand hours, the latest turbofans' technical problems disrupt airlines operations and manufacturers deliveries while production rates rise sharply. Trent 1000 cracked blades grounded almost 50 Boeing 787s and reduced ETOPS to 2.3 hours down from 5.5, costing Rolls-Royce plc almost $950 million. PW1000G knife-edge seal fractures have caused Pratt & Whitney to fall behind in deliveries, leaving about 100 engineless A320neos waiting for their powerplants. The CFM LEAP introduction had been smoother but a ceramic composite HP Turbine coating was prematurely lost, necessitating a new design, causing 60 A320neo engine removals for modification and delaying deliveries by up to six weeks late.

On a widebody, Safran estimates 5–10% of fuel could be saved by reducing power intake for hydraulic systems, while swapping to electrical power could save 30% of weight, as initiated on the Boeing 787, while Rolls-Royce plc hopes for up to 5%.

Manufacturers
The turbofan engine market is dominated by General Electric, Rolls-Royce plc and Pratt & Whitney, in order of market share. General Electric and Safran of France have a joint venture, CFM International. Pratt & Whitney also have a joint venture, International Aero Engines with Japanese Aero Engine Corporation and MTU Aero Engines of Germany, specializing in engines for the Airbus A320 family. Pratt & Whitney and General Electric have a joint venture, Engine Alliance selling a range of engines for aircraft such as the Airbus A380.

For airliners and cargo aircraft, the in-service fleet in 2016 is 60,000 engines and should grow to 103,000 in 2035 with 86,500 deliveries according to Flight Global. A majority will be medium-thrust engines for narrow-body aircraft with 54,000 deliveries, for a fleet growing from 28,500 to 61,000. High-thrust engines for wide-body aircraft, worth 40–45% of the market by value, will grow from 12,700 engines to over 21,000 with 18,500 deliveries. The regional jet engines below 20,000 lb (89 kN) fleet will grow from 7,500 to 9,000 and the fleet of turboprops for airliners will increase from 9,400 to 10,200. The manufacturers market share should be led by CFM with 44% followed by Pratt & Whitney with 29% and then Rolls-Royce and General Electric with 10% each.

Extreme bypass jet engines
In the 1970s, Rolls-Royce/SNECMA tested a M45SD-02 turbofan fitted with variable-pitch fan blades to improve handling at ultralow fan pressure ratios and to provide thrust reverse down to zero aircraft speed. The engine was aimed at ultraquiet STOL aircraft operating from city-centre airports.

In a bid for increased efficiency with speed, a development of the turbofan and turboprop known as a propfan engine was created that had an unducted fan. The fan blades are situated outside of the duct, so that it appears like a turboprop with wide scimitar-like blades. Both General Electric and Pratt & Whitney/Allison demonstrated propfan engines in the 1980s. Excessive cabin noise and relatively cheap jet fuel prevented the engines being put into service. The Progress D-27 propfan, developed in the U.S.S.R., was the only propfan engine equipped on a production aircraft.

Terminology

 * Afterburner: jetpipe equipped for afterburning
 * Augmentor: afterburner for turbofan with burning in hot and cold flows
 * Bypass: that part of the engine as distinct from the core in terms of components and airflow, eg that part of fan blading (fan outer) and stators which pass bypass air, bypass duct, bypass nozzle
 * Bypass ratio: bypass air mass flow /core air mass flow
 * Core: that part of the engine as distinct from the bypass in terms of components and airflow, eg core cowl, core nozzle, core airflow and associated machinery, combustor and fuel system
 * Core power: also known as "available energy" or "gas horsepower". It is used to measure the theoretical (isentropic expansion) shaft work available from a gas generator or core by expanding hot, high pressure gas to ambient pressure. Since the power depends on the pressure and temperature of the gas (and the ambient pressure) a related figure of merit for thrust-producing engines is one which measures the thrust-producing potential from hot, high pressure gas and known as "stream thrust". It is obtained by calculating the velocity obtained with isentropic expansion to atmospheric pressure. The significance of the thrust obtained appears when multiplied by the aircraft velocity to give the thrust work. The thrust work which is potentially available is far less than the gas horsepower due to the increasing waste in the exhaust kinetic energy with increasing pressure and temperature before expansion to atmospheric pressure. The two are related by the propulsive efficiency, a measure of the energy wasted as a result of producing a force (ie thrust) in a fluid by increasing the speed (ie momentum) of the fluid.
 * Dry: engine ratings/ throttle lever positions below afterburning selection
 * EGT: exhaust gas temperature
 * EPR: engine pressure ratio
 * Fan: turbofan LP compressor
 * Fanjet: turbofan or aircraft powered by turbofan (colloquial)
 * Fan pressure ratio: fan outlet total pressure/fan inlet total pressure
 * Flex temp: At reduced take-off weights commercial aircraft can use reduced thrust which increases engine life and reduces maintenance costs. Flex temperature is a higher than actual outside air temperature (OAT) which is input to the engine monitoring computer to achieve the required reduced thrust (also known as "assumed temperature thrust reduction").
 * Gas generator: that part of the engine core which provides the hot, high pressure gas for fan-driving turbines (turbofan), for propelling nozzles (turbojet), for propeller- and rotor-driving turbines (turboprop and turboshaft), for industrial and marine power turbines
 * HP: high-pressure
 * Intake ram drag: Loss in momentum of engine stream tube from freestream to intake entrance, ie amount of energy imparted to air required to accelerate air from a stationary atmosphere to aircraft speed.
 * IEPR: integrated engine pressure ratio
 * IP: intermediate pressure
 * LP: low-pressure
 * Net thrust: nozzle thrust in stationary air (gross thrust) – engine stream tube ram drag (loss in momentum from freestream to intake entrance, ie amount of energy imparted to air required to accelerate air from a stationary atmosphere to aircraft speed). This is the thrust acting on the airframe.
 * Overall pressure ratio: combustor inlet total pressure/intake delivery total pressure
 * Overall efficiency: thermal efficiency * propulsive efficiency
 * Propulsive efficiency: propulsive power/rate of production of propulsive kinetic energy (maximum propulsive efficiency occurs when jet velocity equals flight velocity, which implies zero net thrust!)
 * Specific fuel consumption (SFC): total fuel flow/net thrust (proportional to flight velocity/overall thermal efficiency)
 * Spooling up: increase in RPM (colloquial)
 * Spooling down: decrease in RPM (colloquial)
 * Stage loading: For a turbine, the purpose of which is to produce power, the loading is an indicator of power developed per lb/sec of gas (specific power). A turbine stage turns the gas from an axial direction and speeds it up (in the nozzle guide vanes) to turn the rotor most effectively ( rotor blades must produce high lift), the proviso being that this is done efficiently, ie with acceptable losses. For a compressor stage, the purpose of which is to produce a pressure rise, a diffusion process is used. How much diffusion may be allowed ( and pressure rise obtained) before unacceptable flow separation occurs (ie losses) may be regarded as a loading limit.
 * Static pressure: pressure of the fluid which is associated not with its motion but with its state or, alternatively, pressure due to the random motion of the fluid molecules that would be felt or measured if moving with the flow
 * Specific thrust: net thrust/intake airflow
 * Thermal efficiency: rate of production of propulsive kinetic energy/fuel power
 * Total fuel flow: combustor (plus any afterburner) fuel flow rate (e.g., lb/s or g/s)
 * Total pressure: static pressure plus kinetic energy term
 * Turbine rotor inlet temperature: maximum cycle temperature, ie temperature at which work transfer takes place

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title: Turbofan