16.50 Lecture 7: Modeling of Rocket Nozzles, Effects of Nozzle Area Ratio

16.50 Lecture 7 Subject: Modeling of rocket nozzles; effects of nozzle area ratio. In the last lecture we saw how the throat area of the nozzle controls the mass flow rate. Now we will explore the effects of the shape of the nozzle downstream of the throat. The Mach number and hence velocity at any point in the nozzle is determined by the ratio of the area of the stream tube to the area of the throat (the area ratio), so if we assume that a) The nozzle flows full, i.e. the streamtube shape matches the shape of the nozzle b) The flow is supersonic to the point x downstream of the throat, then we can find M(x) from A(x)/At and it in turn determines p(x), T(x) and u(x). As we shall see, it is the pressure that determines whether the nozzle flows full, so it is convenient to relate the exit velocity directly to the pressure. We do this from the Energy Equation, u2 Tc ! T + 2c p u 2 = 2c p (Tc ! T ) = 2c pTc [1 ! T ] Tc So at any point in the nozzle where s=sc " !1 p " 12 ) ] pc and this is independent of whether the nozzle is full at this pressure, because the area is not referred to. In particular, if we apply this at the end of the nozzle where the pressure is pe, the velocity at that point is u = 2c pTc [1! ( " !1 pe " 12 ue = 2c pTc [1! ( ) ] pc You know from your previous work that if the nozzle flows full the thrust of the rocket can be written . F = m ue + Ae ( pe ! po ) substituting the expression for ue we then have the following expression for F in terms of the pressure ratio pe/p0. " !1 pe " 12 F = m! 2c pTc [1! ( ) ] + Ae ( pe ! po ) pc The mass flow rate is given in terms of c* by . pA m! c t c* . 1 and for ideal gases, we have the estimate of c* given by " +1 c = * RTc 2 2(" #1) ; != "( ) ! " +1 F Substituting these into the expression for F and non-dimensionalizing it: ! +1 ! "1 p A # p " po &( F 2 ! 2 #% 2 & ! "1 = [1 " ( e ) ! ] + e % e ! " 1 $ ! + 1' pc At $ pc ' pc At where the area ratio is itself related to the pressure ratio through continuity: " +1 Ae ! t ut " #1 2 2(" #1) 1 = = ( ) " #1 At ! e ue 2 " +1 p p ( e )1/ " 1# ( e ) " pc pc More generally, we define a Thrust Coefficient by F ! pc At cF and the above expression then gives us an estimate of cF for ideal gases. The dependence of cF on nozzle area ratio and pressure ratio is conventionally summarized as in the figure below, which is drawn for γ=1.2 and 1.3. Notice: -γ is replaced by k in these graphs -pc is replaced by p1 -pe is replaced by p2 -pa, or p0, the ambient pressure, is replaced by p3 2 2.0 k = 1.20 1.8 = 8 p 1lp 3 Line of optimum thrust coefficient P2=P3 p1 lp 3 100 1.6 33.3 2.246 lue= um va Maxim 1000 500 333 = 20 0 50. 0 20 1.4 CF .0 10.0 1.2 Region of flow separation for conical and bell shaped nozzles 5. 0 1.0 3. 3 2. 5 0.8 2. 0 0.6 1 2 4 6 Area ratio 8 10 20 40 60 80 100 ε = A2 lAt Thrust coefficient CF versus nozzle area ratio for k = 1.20. Image by MIT OpenCourseWare. After Sutton and Biblarz 2001. 2.0 k = 1.30 Line of optimum thrust coefficient P2=P3 p1lp3 = 100 50 33.3 1.4 CF 1000 500 p1 lp 3 = 333 200 1.5 20. 0 10. 0 1.2 Region of flow separation for conical and bell shaped nozzles 5.0 1.0 3. 3 2. 5 0.8 0.6 4 lue=1.96 Maximum va 8 1.8 2.0 1 2 4 6 8 Area ratio 10 20 40 60 80 100 ε = A2 lAt Thrust coefficient CF versus nozzle area ratio for k = 1.30. Image by MIT OpenCourseWare. After Sutton and Biblarz 2001. 3 The thrust is also expressible in terms of the effective exhaust velocity c, as we discussed in Lecture 1. Then we have pA F ! m! c = c t = pc At cF c* and we see that c = c * cF The "Characteristic Velocity" c* depends mainly on propellant properties. The Thrust Coefficient cF depends on propellant properties through γ (or equivalent), but mainly on the pressure ratio and nozzle geometry. So by the definitions of cF and c* we have managed to separate the effects of propellant properties and the effects of nozzle geometry into the two factors of c. This separation, though demonstrated only for ideal gases, holds also for the more general situation of complex chemically reacting propellants. Effects of Non-ideal Expansion So far we have assumed in the discussion of cF that the flow fills the nozzle and is supersonic to the exit. If this is the case, the thrust is given by the above expression for cF . But in fact the flow can be somewhat more complex than this, depending the ratio pc/p0 compared to that which leads to ideal expansion. This behavior is summarized in the figure below. Kerrebrock, Jack L. (1992). Aircraft Engines and Gas Turbines (2nd Edition). MIT Press, © Massachusetts Institute of Technology. Used with permission. Here An is the throat area, Ae the exit area. The ideally expanded situation is at the upper right, and it is the condition at which thrust is maximum for a given external pressure p0. This is easier to visualize than to prove analytically: if the divergent nozzle were extended a little by adding a section at its exit, this section would see an internal pressure lower than p0, and would therefore generate suction, or negative thrust. If on the contrary, the nozzle were shortened a little, one would lose the positive thrust that was being produced by the removed portion. Returning to consideration of a given nozzle, if p0 is lowered below its ideal matching pressure (for example by the rocket ascending in the atmosphere) the nozzle becomes underexpanded, as at the lower right. In this case the flow fills the nozzle and our formula for F works fine. 4 If on the other hand p0 is larger than corresponds to ideal expansion, the situation can become more complex. For pressure ratios p0/ pe< 2 to 2.5, the nozzle is likely to remain full, and again the formula holds. But for larger pressure ratios the oblique shocks that form at the exit of the nozzle are strong enough to separate the boundary layer, and the point of separation moves into the nozzle so that its effective area decreases, as shown at the upper left. In this case the separation occurs approximately at a pressure ps such that ps/p0 = ½ to 1/(2.5). Again we can use the formula for thrust by replacing pe by ps and Ae by As, the area at which the pressure is ps. So now . F = m us + As (ps ! po ) With this understanding, we can write the thrust coefficient for the separated nozzle as ! +1 F = pc At p 2 ! 2 %# 2 & ! "1 [1 " ( s ) ! " 1 $ ! + 1' pc ! "1 ! ]+ As %# ps " po &( At $ pc ' where now it is understood that ps is the pressure at which the separation occurs and As is the corresponding area. 5