## Overall pressure ratio – wikipedia electricity lesson plans 8th grade

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] Advantages of high overall __pressure ratios__ [ edit ]

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. Disadvantages of high overall pressure ratios [ edit ]

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. Examples [ edit ] Engine

The contribution of intake to the overall __pressure ratio__, defined as the ratio of the stagnation pressure as measured at the front and rear of the intake is negative in supersonic flight, and approximately non-existent in subsonic flight. Therefore with the definition given in the first sentence in this article, the overall effective pressure ratio of Concorde taking intake into account is less than without it. In other words less than 15.5:1 as in the table. Information from the source given above obviously uses a different definition for overall __pressure ratio__, comparing static pressures instead of stagnation pressures, and is misleading given the definition in this article as it is.

Compression ratio and overall pressure ratio would be interrelated as follows, if the first sentence in this article would be incorrect, given as it is, what follows is valid only for static pressures, but not overall **pressure ratios**, as compresson 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

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: P 1 V 1 γ = P 2 V 2 γ ⇒ P 2 P 1 = ( V 1 V 2 ) γ {\displaystyle P_{1}V_{1}^{\gamma }=P_{2}V_{2}^{\gamma }\Rightarrow {\frac {P_{2}}{P_{1}}}=\left({\frac {V_{1}}{V_{2}}}\right)^{\gamma }}

where γ {\displaystyle \gamma } 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. See also [ edit ]