1. Choking Due To Friction The theory here predicts that for adiabatic frictional flow in a constant area duct, no matter what the inlet Mach number M1
is, the flow downstream tends toward the sonic point. There is a certain duct length L*(M1) for which the exit Mach number will be exactly unity. But what if the actual duct length L is greater than the predict-ed “maximum” length L* ? Then the flow condition must change and there are two classifications. Subsonic Inlet If L ˃ L*(M1), the flow slows down until an inlet Mach number M2 is reached such that L = L*(M2). The exit flow is sonic and the mass flow has been reduced by frict- ional choking. Further increases in duct length will continue to decrease the inlet Mach number M and mass flux.
Prof. Dr. MOHSEN OSMAN

2. Supersonic Inlet From Table we see that friction has a very large effect on supersonic duct flow. Even an infinite inlet Mach number will be reduced to sonic conditions in only 41
diameters M = M = ∞ Fig. (4.3) Behavior of duct flow with a nominal supersonic inlet condition M = 3.0 (a) L/D ≤ 26, flow is supersonic throughout duct; (b) L/D = 40 ˃L*/D; normal shock at M = 2.0 with subsonic flow then accelerating to sonic exit point.
Prof. Dr. MOHSEN OSMAN

3. (c) L/D = 53, shock must now occur at M = 2
(c) L/D = 53, shock must now occur at M = 2.5; (d) L/D ˃ 63, flow must be entirely subsonic and choked at exit Some typical numerical values are shown in Fig. (4.3), assuming
an inlet M = 3.0 and For this condition L* = 26 diam. If L is increased beyond 26 D, the flow will not choke but a normal shock will form at just the right place for the subsequent subsonic frictional flow to become sonic exactly at the exit. Figure (4.3) shows two examples, for As length increases, the required shock moves upstream until, for Fig.(4.3), the shock is at the inlet for Further increase in L causes the shock to move upstream of the inlet into the supersonic nozzle feeding the duct Yet the mass flux is still the same as for the very short duct, because presumably the feed nozzle still has a sonic throat.
Prof. Dr. MOHSEN OSMAN

4. Eventually, a very long duct will cause the feed-nozzle throat to become choked, thus reducing the duct mass flux. Thus supersonic friction changes the flow pattern if L ˃L* but does
not choke the flow until L is much larger than L* Example Air enters a 3-cm-diameter duct at and V1 = 100 m/s. The friction factor is Compute (a) the maximum duct length for this condition, (b) the mass flux if the duct length is 15 m, and (c) the reduced mass flux if L = 30 m Required: D = 3 cm & f = (a) Lmax for given condition (b) for L = 15 m Po =200 kPa (c) for L = 30 m To =500 K
V1 = 100 m/s
Prof. Dr. MOHSEN OSMAN

5. Solution (a) Apply energy equation between stagnation and inlet states CP To = CP T1 + ½ V Interpolating into Fanno-Line Table across, from M=M1=0.225 we get
Prof. Dr. MOHSEN OSMAN

6. (b) The given L=15 m is less than Lmax , and so is not choked and the mass flux follows from inlet conditions;
Prof. Dr. MOHSEN OSMAN

7. (c) Since L = 30 m is greater than Lmax , the duct must choke
back until L = Lmax , corresponding to a lower inlet M1 Interpolation in Fanno-Line Table we find that this value of 20 corresponds to M1, choked = (23% less) From Isentropic Table :
Prof. Dr. MOHSEN OSMAN

8. Isothermal Flow with Friction The adiabatic frictional-flow assumption is appropriate to high-speed flow in short ducts. For flow in long ducts, e.g.,
natural-gas pipelines, the gas state more closely approximates an isothermal flow. The analysis is the same except that the iso-energetic energy equation (4.1c) is replaced by the simple relation. T = constant dT = 0.0 (4.11) Again it is possible to write all property changes in terms of the Mach number. Integration of the Mach number versus friction relation yields . which is the isothermal analog of equation (4.7) for adiabatic flow.
Prof. Dr. MOHSEN OSMAN

9. This friction relation has the interesting result that Lmax becomes zero not at the sonic point but at a If the tube length L is greater than Lmax from Eq. (4.12), a sub-sonic flow will choke back to a smaller M1 and mass flux and a supersonic flow will experience a normal-shock adjustment to Fig. (4.3) The exit isothermal choked flow is not sonic, and so the use of the asterisk (*) is approximate Let represent properties at the choking point L=Lmax . Then the isothermal analysis leads to the following Mach number relations for the flow properties:
Prof. Dr. MOHSEN OSMAN

10. An interesting by-product of the isothermal analysis is an exact relation between pressure drop and duct mass flux. In contrast, it is practically impossible in adiabatic flow to eliminate the
Mach number between eqns. (4.7) and (4.9a) to get a direct relation between pressure and friction Consequently, the common problem of predicting adiabatic duct mass flux for a given pressure drop can only be solved by a laborious Mach-number-iteration procedure For isothermal flow, we can substitute in Eq.(4.1b) to obtain But from Eqs.(4.1a) and (4.2) with dT = 0.0, we have so that Eq. (4.14) becomes
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11. Since G2RT is constant for isothermal flow, these are exact differentials and can be integrated from giving: Thus we have an explicit expression for isothermal mass flux as a function of the pressure drop in the duct. If you don’t think eqn (4.16) is beautiful, try solving the same problem for adiabatic flow with the Mach-number relations or Table Eqn (4.16) has one flow: with the Mach number eliminated, it cannot recognize the choking phenomenon. Therefore one should check the physical realism of a solution using Eq(4.16) by computing the exit Mach number M2 to ensure that it is not greater than for a subsonic inlet flow.
Prof. Dr. MOHSEN OSMAN

12. Example 4.4 Air enters a pipe of 1 in diameter at subsonic velocity and P1 = 30 psia, T1 = 550oR.
If the pipe is 10 ft long, and the exit pressure P2 =20 psia, compute the mass flux for (a) isothermal flow and (b) adiabatic flow Solution f = A) Isothermal Flow D = 1inch Lpipe = 10 inches P2 =20 psia P1 = 30 psia T2 =T1 = 550oR T1 = 550oR For isothermal flow Eq. (4.16) applies. Compute
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13. Accordingly,
G = 1.7 lbf .s/ft3 = 1.7 slugs/(s.ft2 ) (Note: mass unit = slug = (weight unit / g) ) Check the Mach number at inlet and exit.
Prof. Dr. MOHSEN OSMAN

14. Then &
Since these are well below choking, the solution is accurate. B) Adiabatic Flow
Lmax1 – Lmax2 = 10
f = 0.025
P1 = 30 psia P2 = 20 psia
T1 = 550oR
For adiabatic flow, the only solution you can think of is to guess M1 Compute read M2 , finally compute
which is the ratio of the given pressures
Prof. Dr. MOHSEN OSMAN

15. If so, you have found the correct solution
If so, you have found the correct solution. If not try another M1 and iterate to find the proper solution. The isothermal relation provides a guess for The last value is close enough : M1 M2
P1 / P2
0.30
– 3 =
2.2993
0.4005
1.344
0.32
– 3 =
1.4467
0.4604
1.454
0.33
– 3 =
1.0821
0.4985
1.531
0.326
– 3 =
1.2235
0.4823
1.498
Prof. Dr. MOHSEN OSMAN

16. Example 4. 5 At the inlet of 0. 1 m diameter duct, the initial Mach no
Example At the inlet of 0.1 m diameter duct, the initial Mach no. is 0.2 and the static pressure and temperature are 200 kPa and 300K
respectively. If the friction factor is 0.02, find the Mach number and properties after the flow has traversed 50 m of duct Note:
M1 = 0.2, P1 = 200 kPa
T1 = 300 K
(1)________________________(2) (1)_______________(2) P*,T*
f = L = 50 m D = 0.1 m f = L = 50 m D = 0.1m
_________________________ ______________
(a) Problem Statement
(b) Sonic position indicated
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17. From Isentropic Table From Fanno line Table @ M1 = 0
From Isentropic Table From Fanno line M1 = Interpolating into Fanno Line Table across from
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18. Example A converging-diverging nozzle / constant-area duct system is connected to a reservoir containing air with a stagnation
pressure of 700 kPa. The exit to throat area ratio of the nozzle is 3; the constant area duct is 6 m long and 0.3 m in diameter with a friction coefficient of Determine the range of back pressures for which a normal shock wave will stand in the duct SOLUTION Using Isentropic Table for we get: .Me = 2.64, If a normal shock wave will stand at inlet to the duct; use N.S. Table for Mx = Me = 2.64, we get kPa Using Fanno Line Table for M = My = , we get:
Prof. Dr. MOHSEN OSMAN

19. Flow parameters If a normal shock wave will stand at exit of the duct; use Fanno Line Table for M1 = Me = 2.64, we get: Using N.S. Table for Mx = M2 = , we get: kPa ˂ Pb ˂ kPa
Prof. Dr. MOHSEN OSMAN