Overall Pressure Ratio

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## History of overall pressure ratios

## Advantages of high overall pressure ratios

## Disadvantages of high overall pressure ratios

## Examples

## Differences from other similar terms

## Compression ratio versus overall pressure ratio

## See also

## References

This article uses material from the Wikipedia page available here. It is released under the Creative Commons Attribution-Share-Alike License 3.0.

Overall Pressure Ratio

In aeronautical engineering, **overall pressure ratio**, or **overall compression ratio**, is the ratio of the stagnation pressure as measured at the front and rear of the compressor of a gas turbine engine. The terms *compression ratio* and *pressure ratio* are used interchangeably.^{[1]}**Overall compression ratio** also means the *overall cycle pressure ratio* which includes intake ram.^{[2]}

Early jet engines had limited pressure ratios due to construction inaccuracies of the compressors and various material limits. For instance, the Junkers Jumo 004 from World War II had an overall pressure ratio 3.14:1. The immediate post-war Snecma Atar improved this marginally to 5.2:1. Improvements in materials, compressor blades, and especially the introduction of multi-spool engines with several different rotational speeds, led to the much higher pressure ratios common today.

Modern civilian engines generally operate between 40 and 55:1. The highest in-service is the General Electric GEnx-1B/75 with an OPR of 58 at the end of the climb to cruise altitude (Top of Climb) and 47 for takeoff at sea level.^{[3]}

Generally speaking, a higher overall pressure ratio implies higher efficiency, but the engine will usually weigh more, so there is a compromise. A high overall pressure ratio permits a larger area ratio nozzle to be fitted on the jet engine. This means that more of the heat energy is converted to jet speed, and energetic efficiency improves. This is reflected in improvements in the engine's specific fuel consumption.

The GE Catalyst has a 16:1 OPR and its thermal efficiency is 40%, the 32:1 Pratt & Whitney GTF has a thermal efficiency of 50% and the 58:1 GEnx has a thermal efficiency of 58%.^{[4]}

One of the primary limiting factors on pressure ratio in modern designs is that the air heats up as it is compressed. As the air travels through the compressor stages it can reach temperatures that pose a material failure risk for the compressor blades. This is especially true for the last compressor stage, and the outlet temperature from this stage is a common figure of merit for engine designs.

Military engines are often forced to work under conditions that maximize the heating load. For instance, the General Dynamics F-111 Aardvark was required to operate at speeds of Mach 1.1 at sea level. As a side-effect of these wide operating conditions, and generally older technology in most cases, military engines typically have lower overall pressure ratios. The Pratt & Whitney TF30 used on the F-111 had a pressure ratio of about 20:1, while newer engines like the General Electric F110 and Pratt & Whitney F135 have improved this to about 30:1.

An additional concern is weight. A higher compression ratio implies a heavier engine, which in turn costs fuel to carry around. Thus, for a particular construction technology and set of flight plans an optimal overall pressure ratio can be determined.

Engine | Overall pressure ratio | Major applications |
---|---|---|

General Electric GE9X | 60:1 | 777X |

Rolls-Royce Trent XWB | 52:1 | A350 XWB |

General Electric GE90 | 42:1 | 777 |

General Electric CF6 | 30.5:1 | 747, 767, A300, MD-11, C-5 |

General Electric F110 | 30:1 | F-14, F-15, F-16 |

Pratt & Whitney TF30 | 20:1 | F-14, F-111 |

Rolls-Royce/Snecma Olympus 593 | 15.5:1/80:1 Supersonic.^{[5]} |
Concorde |

The term should not be confused with the more familiar term compression ratio applied to reciprocating engines. Compression ratio is a ratio of volumes. In the case of the Otto cycle reciprocating engine, the maximum expansion of the charge is limited by the mechanical movement of the pistons (or rotor), and so the compression can be measured by simply comparing the volume of the cylinder with the piston at the top and bottom of its motion. The same is not true of the "open ended" gas turbine, where operational and structural considerations are the limiting factors. Nevertheless, the two terms are similar in that they both offer a quick way of determining overall efficiency relative to other engines of the same class.

The broadly equivalent measure of rocket engine efficiency is chamber pressure/exit pressure, and this ratio can be over 2000 for the Space Shuttle Main Engine.

Compression ratio and overall pressure ratio are interrelated as follows (what follows is valid only for static pressures, not overall pressure ratios, as compression while slowing down flow without doing work is not changing stagnation pressure at all, and in supersonic flow with shock waves stagnation pressure will be less than 1 with > 1 compression ratio):

Compression ratio | 2:1 | 3:1 | 5:1 | 10:1 | 15:1 | 20:1 | 25:1 | 35:1 |
---|---|---|---|---|---|---|---|---|

Pressure ratio | 2.64:1 | 4.66:1 | 9.52:1 | 25.12:1 | 44.31:1 | 66.29:1 | 90.60:1 | 145.11:1 |

The reason for this difference is that compression ratio is defined via the volume reduction:

- ,

while pressure ratio is defined as the pressure increase:

- .

In calculating the pressure ratio, we assume that an adiabatic compression is carried out (i.e. that no heat energy is supplied to the gas being compressed, and that any temperature rise is solely due to the compression). We also assume that air is a perfect gas. With these two assumptions, we can define the relationship between change of volume and change of pressure as follows:

where is the ratio of specific heats (air: approximately 1.4). The values in the table above are derived using this formula. Note that in reality the ratio of specific heats changes with temperature and that significant deviations from adiabatic behavior will occur.

**^**"The aircraft Gas Turbine Engine and its operation" P&W Oper.Instr.200, United Technologies Pratt & whitney December, 1982, p.49**^**http://www.ulb.tu-darmstadt.de/tocs/210525592.pdf p.695**^**Bjorn Fehrm (October 28, 2016). "Bjorn's Corner: Turbofan engine challenges, Part 1".*Leeham News*.**^**Bjorn Fehrm (June 14, 2019). "Bjorn's Corner: Why hybrid cars work and hybrid airliners have challenges".*Leeham News*.**^**Concorde: story of a supersonic pioneer By Kenneth Owen

This article uses material from the Wikipedia page available here. It is released under the Creative Commons Attribution-Share-Alike License 3.0.

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