Open access peer-reviewed chapter

# Effect of Stagnation Temperature on Supersonic Flow Parameters with Application for Air in Nozzles

By Toufik Zebbiche

Submitted: November 4th 2010Reviewed: March 7th 2011Published: November 2nd 2011

DOI: 10.5772/19633

## 1. Introduction

The obtained results of a supersonic perfect gas flow presented in (Anderson, 1982, 1988& Ryhming, 1984), are valid under some assumptions. One of the assumptions is that the gas is regarded as a calorically perfect, i. e., the specific heats C P is constant and does not depend on the temperature, which is not valid in the real case when the temperature increases (Zebbiche & Youbi, 2005b, 2006, Zebbiche, 2010a, 2010b). The aim of this research is to develop a mathematical model of the gas flow by adding the variation effect of C P and γ with the temperature. In this case, the gas is named by calorically imperfect gas or gas at high temperature. There are tables for air (Peterson & Hill, 1965) for example) that contain the values of C P and γ versus the temperature in interval 55 K to 3550 K. We carried out a polynomial interpolation of these values in order to find an analytical form for the function C P (T).

The presented mathematical relations are valid in the general case independently of the interpolation form and the substance, but the results are illustrated by a polynomial interpolation of the 9th degree. The obtained mathematical relations are in the form of nonlinear algebraic equations, and so analytical integration was impossible. Thus, our interest is directed towards to the determination of numerical solutions. The dichotomy method for the solution of the nonlinear algebraic equations is used; the Simpson’s algorithm (Démidovitch & Maron, 1987& Zebbiche & Youbi, 2006, Zebbiche, 2010a, 2010b) for numerical integration of the found functions is applied. The integrated functions have high gradients of the interval extremity, where the Simpson’s algorithm requires a very high discretization to have a suitable precision. The solution of this problem is made by introduction of a condensation procedure in order to refine the points at the place where there is high gradient. The Robert’s condensation formula presented in (Fletcher, 1988) was chosen. The application for the air in the supersonic field is limited by the threshold of the molecules dissociation. The comparison is made with the calorically perfect gas model.

The problem encounters in the aeronautical experiments where the use of the nozzle designed on the basis of the perfect gas assumption, degrades the performances. If during the experiment measurements are carried out it will be found that measured parameters are differed from the calculated, especially for the high stagnation temperature. Several reasons are responsible for this deviation. Our flow is regarded as perfect, permanent and non-rotational. The gas is regarded as calorically imperfect and thermally perfect. The theory of perfect gas does not take account of this temperature.

To determine the application limits of the perfect gas model, the error given by this model is compared with our results.

## 2. Mathematical formulation

The development is based on the use of the conservation equations in differential form. We assume that the state equation of perfect gas (P=ρRT) remains valid, with R=287.102 J/(kg K). For the adiabatic flow, the temperature and the density of a perfect gas are related by the following differential equation (Moran, 2007& Oosthuisen & Carscallen, 1997& Zuker & Bilbarz, 2002, Zebbiche, 2010a, 2010b).

CPγdTRTρdρ=0E1

Using relationship between C P and γ [C P =γR/(γ-1)], the equation (1) can be written at the following form:

dρρ=dTT[γ(T)1]E2

The integration of the relation (2) gives the adiabatic equation of a perfect gas at high temperature.

The sound velocity is (Ryhming, 1984),

a2=(dPdρ)entropy=constantE3

The differentiation of the state equation of a perfect gas gives:

dPdρ=ρRdTdρ+RTE4

Substituting the relationship (2) in the equation (4), we obtain after transformation:

a2(T)=γ(T)RTE5

Equation (5) proves that the relation of speed of sound of perfect gas remains always valid for the model at high temperature, but it is necessary to take into account the variation of the ratio γ(T).

The equation of the energy conservation in differential form (Anderson, 1988& Moran, 2007) is written as:

CPdT+VdV=0E6

The integration between the stagnation state (V 0 ≈ 0, T 0 ) and supersonic state (V, T) gives:

V2=2H(T)E7

Where

H(T)=TT0CP(T)dTE8

Dividing the equation (6) by V 2 and substituting the relation (7) in the obtained result, we obtain:

dVV=CP(T)2H(T)dTE9

Dividing the relation (7) by the sound velocity, we obtain an expression connecting the Mach number with the enthalpy and the temperature:

M(T)=2H(T)a(T)E10

The relation (10) shows the variation of the Mach number with the temperature for calorically imperfect gas.

The momentum equation in differential form can be written as (Moran, 2007, Peterson & Hill1, 1965, & Oosthuisen & Carscallen, 1997):

VdV+dPρ=0E11

Using the expression (3), the relationship (10), can be written as:

dρρ=Fρ(T)dTE12

Where

Fρ(T)=CP(T)a2(T)E13

The density ratio relative to the temperature T 0 can be obtained by integration of the function (13) between the stagnation state (ρ 0 ,T 0 ) and the concerned supersonic state (ρ,T):

ρρ0=Exp(TT0Fρ(T)dT)E14

The pressure ratio is obtained by using the relation of the perfect gas state:

PP0=(ρρ0)(TT0)E15

The mass conservation equation is written as (Anderson, 1988& Moran, 2007)

ρVA=constantE16

The taking logarithm and then differentiating of relation (16), and also using of the relations (9) and (12), one can receive the following equation:

dAA=FA(T)dTE17

Where

FA(T)=CP(T)[1a2(T)12H(T)]E18

The integration of equation (17) between the critical state (A * , T * ) and the supersonic state (A, T) gives the cross-section areas ratio: *

AA*=Exp(TT*FA(T)dT)E19

To find parameters ρ and A, the integrals of functions F ρ (T) and F A (T) should be found. As the analytical procedure is impossible, our interest is directed towards the numerical calculation. All parameters M, ρ and A depend on the temperature.

The critical mass flow rate (Moran, 2007, Zebbiche & Youbi, 2005a, 2005b) can be written in non-dimensional form:

mA*ρ0a0=.A(ρρ0)(aa0)Mcos(θ)dAA*E20

As the mass flow rate through the throat is constant, we can calculate it at the throat. In this section, we have ρ=ρ * , a=a * , M=1, θ=0 and A=A * . Therefore, the relation (20) is reduced to:

m˙A*ρ0a0=(ρ*ρ0)(a*a0)E21

The determination of the velocity sound ratio is done by the relation (5). Thus,

aa0=[γ(T)γ(T0)]1/2[TT0]1/2E22

The parameters T, P, ρ and A for the perfect gas are connected explicitly with the Mach number, which is the basic variable for that model. For our model, the basic variable is the temperature because of the implicit equation (10) connecting M and T, where the reverse analytical expression does not exist.

## 3. Calculation procedure

In the first case, one presents the table of variation of CP and γ versus the temperature for air (Peterson & Hill, 1965, Zebbiche 2010a, 2010b). The values are presented in the table 1.

 T (K) CP(J/(KgK) γ(T) T (K) CP (J/(Kg K) γ(T) T (K) CP J/(Kg K) γ(T) 55.538 1001.104 1.402 833.316 1107.192 1.350 2111.094 1256.813 1.296 . . . 888.872 1119.078 1.345 2222.205 1263.410 1.294 222.205 1001.101 1.402 944.427 1131.314 1.340 2333.316 1270.097 1.292 277.761 1002.885 1.401 999.983 1141.365 1.336 2444.427 1273.476 1.291 305.538 1004.675 1.400 1055.538 1151.658 1.332 2555.538 1276.877 1.290 333.316 1006.473 1.399 1111.094 1162.202 1.328 2666.650 1283.751 1.288 361.094 1008.281 1.398 1166.650 1170.280 1.325 2777.761 1287.224 1.287 388.872 1011.923 1.396 1222.205 1178.509 1.322 2888.872 1290.721 1.286 416.650 1015.603 1.394 1277.761 1186.893 1.319 2999.983 1294.242 1.285 444.427 1019.320 1.392 1333.316 1192.570 1.317 3111.094 1297.789 1.284 499.983 1028.781 1.387 1444.427 1204.142 1.313 3222.205 1301.360 1.283 555.538 1054.563 1.374 1555.538 1216.014 1.309 3333.316 1304.957 1.282 611.094 1054.563 1.370 1666.650 1225.121 1.306 3444.427 1304.957 1.282 666.650 1067.077 1.368 1777.761 1234.409 1.303 3555.538 1308.580 1.281 722.205 1080.005 1.362 1888.872 1243.883 1.300 777.761 1093.370 1.356 1999.983 1250.305 1.298

### Table 1.

Variation of C P (T) and γ(T) versus the temperature for air.

For a perfect gas, the γ and C P values are equal to γ=1.402 and C P =1001.28932 J/(kgK) (Oosthuisen & Carscallen, 1997, Moran, 2007& Zuker & Bilbarz, 2002).. The interpolation of the C P values according to the temperature is presented by relation (23) in the form of Horner scheme to minimize the mathematical operations number (Zebbiche, 2010a, 2010b):

CP(T)=a1+T(a2+T(a3+T(a4+T(a5+T(a6+T(a7+T(a8+T(a9+T(a10)))))))))E23

The interpolation (a i i=1, 2, …, 10) of constants are illustrated in table 2.

 I ai I ai 1 1001.1058 6 3.069773 10-12 2 0.04066128 7 -1.350935 10-15 3 -0.000633769 8 3.472262 10-19 4 2.747475 10-6 9 -4.846753 10-23 5 -4.033845 10-9 10 2.841187 10-27

### Table 2.

Coefficients of the polynomial C P (T).

A relationship (23) gives undulated dependence for temperature approximately low thanT¯=240K. So for this field, the table value (Peterson & Hill, 1965), was taken

C¯P=Cp(T¯)=1001.15868J/(kgK)E24

Thus:

forTT¯, we have CP(T)=C¯PforT>T¯, relation (23) is used.

The selected interpolation gives an error less than ε=10 -3 between the table and interpolated values.

Once the interpolation is made, we determine the function H(T) of the relation (8), by integrating the function C P (T) in the interval [T, T 0 ]. Then, H(T) is a function with a parameter T 0 and it is defined when T≤T 0 .

Substituting the relation (23) in (8) and writing the integration results in the form of Horner scheme, the following expression for enthalpy is obtained

H(T)=H0-[c1+T(c2+T(c3+T(c4+T(c5+Tc6+T(c7+T(c8+T(c9+T(c10)))))))))]E25

Where

H0=T0(c1+T0(c2+T0(c3+T0(c4+T0(c5+T0(c6+T0(c7+T0(c8+T0(c9+T0(c10))))))))))E26

and

ci=aii(i=1,2,3,...,10)E27 Figure 1.Variation of function F ρ (T) in the interval [T S ,T 0 ] versusT 0 .

Taking into account the correction made to the function C P (T), the function H(T) has the following form:

ForT0<T¯,

H(T)=C¯P(T0T)E28
ForT0>T¯,we have two cases:

ifT>T¯:H(T)=relation(24)E29
ifTT¯:H(T)=C¯P(T¯T)+H(T¯)E30

The determination of the ratios (14) and (19) require the numerical integration of F ρ (T) and F A (T) in the intervals [T, T 0 ] and [T, T * ] respectively. We carried out preliminary calculation of these functions (Figs. 1, 2) to see their variations and to choice the integration method. Figure 2.Variation of the function F A (T) in the interval [T S ,T * ] versus T 0

Due to high gradient at the left extremity of the interval, the integration with a constant step requires a very small step. The tracing of the functions is selected for T 0 =500 K (low temperature) and M S =6.00 (extreme supersonic) for a good representation in these ends. In this case, we obtain T * =418.34 K and T S =61.07 K. the two functions presents a very large derivative at temperature T S .

A Condensation of nodes is then necessary in the vicinity of T S for the two functions. The goal of this condensation is to calculate the value of integral with a high precision in a reduced time by minimizing the nodes number. The Simpson’s integration method (Démidovitch & Maron, 1987& Zebbiche & Youbi, 2006) was chosen. The chosen condensation function has the following form (Zebbiche & Youbi, 2005a):

si=b1zi+(1b1)[1tanh[b2(1zi)]tanh(b2)]E31

Where

zi=i1N11iNE32

Obtained s i values, enable to find the value of T i in nodes i:

Ti=si(TDTG)+TGE33

The temperature T D is equal to T 0 for F ρ (T), and equal to T * for F A (T). The temperature T G is equal to T * for the critical parameter, and equal to T S for the supersonic parameter. Taking a value b 1 near zero (b 1 =0.1, for example) and b 2 =2.0, it can condense the nodes towards left edge T S of the interval, see figure 3.

The stagnation state is given by M=0. Then, the critical parameters correspond to M=1.00, for example at the throat of a supersonic nozzle, summarize by:

When M=1.00 we have T=T * . These conditions in the relation (10), we obtain:

2H(T*)a2(T*)=0E34

The resolution of equation (29) is made by the use of the dichotomy algorithm (Démidovitch & Maron, 1987& Zebbiche & Youbi, 2006), with T * <T 0 . It can choose the interval [T 1 ,T 2 ] containing T * by T 1 =0 K and T 2 =T 0 . The value T * can be given with a precision ε if the interval of subdivision number K is satisfied by the following condition:

K=1.4426Log(T0ε)+1E35

If ε=10-8 is taken, the number K cannot exceed 39. Consequently, the temperature ratio T * /T 0 can be calculated.

Taking T=T * and ρ=ρ * in the relation (14) and integrating the function F ρ (T) by using the Simpson’s formula with condensation of nodes towards the left end, the critical density ratio is obtained.

The critical ratios of the pressures and the sound velocity can be calculated by using the relations (15) and (22) respectively, by replacing T=T * , ρ=ρ * , P=P * and a=a * ,

### 3.2. Parameters for a supersonic Mach number

For a given supersonic cross-section, the parameters ρ=ρ S , P=P S , A=A S , and T=T S can be determined according to the Mach number M=M S . Replacing T=T S and M=M S in relation (10) gives

2H(TS)MS2a2(TS)=0E36

The determination of T S of equation (31) is done always by the dichotomy algorithm, excepting T S <T * . We can take the interval [T 1 ,T 2 ] containing T S , by (T 1 =0 K, and T 2 =T *.

Replacing T=T S and ρ=ρ S in relation (14) and integrating the function F ρ (T) by using the Simpson’s method with condensation of nodes towards the left end, the density ratio can be obtained.

The ratios of pressures, speed of sound and the sections corresponding to M=M S can be calculated respectively by using the relations (15), (22) and (19) by replacing T=T S , ρ=ρ S , P=P S , a=a S and A=A S .

The integration results of the ratios ρ * / ρ 0 , ρ S 0 and A S /A * primarily depend on the values of N, b 1 and b 2 .

### 3.3. Supersonic nozzle conception

For supersonic nozzle application, it is necessary to determine the thrust coefficient. For nozzles giving a uniform and parallel flow at the exit section, the thrust coefficient is (Peterson & Hill, 1965& Zebbiche, Youbi, 2005b)

CF=FP0A*E37

Where

F=mVE=mMEaEE38

The introduction of relations (21), (22) into (32) gives as the following relation:

CF=γ(T0)ME(aEa0)(ρ*ρ*)(a*a0)E39

The design of the nozzle is made on the basis of its application. For rockets and missiles applications, the design is made to obtain nozzles having largest possible exit Mach number, which gives largest thrust coefficient, and smallest possible length, which give smallest possible mass of structure.

For the application of blowers, we make the design on the basis to obtain the smallest possible temperature at the exit section, to not to destroy the measuring instruments, and to save the ambient conditions. Another condition requested is to have possible largest ray of the exit section for the site of instruments. Between the two possibilities of construction, we prefer the first one.

### 3.4. Error of perfect gas model

The mathematical perfect gas model is developed on the basis to regarding the specific heat C P and ratio γ as constants, which gives acceptable results for low temperature. According to this study, we can notice a difference on the given results between the perfect gas model and developed here model.The error given by the PG model compared to our HT model can be calculated for each parameter. Then, for each value (T 0 , M), the ε error can be evaluated by the following relationship:

εy(T0,M)=|1yPG(T0,M)yHT(T0,M)|×100E40

The letter y in the expression (35) can represent all above-mentioned parameters. As a rule for the aerodynamic applications, the error should be lower than 5%.

## 4. Application

The design of a supersonic propulsion nozzle can be considered as example. The use of the obtained dimensioned nozzle shape based on the application of the PG model given a supersonic uniform Mach number M S at the exit section of rockets, degrades the desired performances (exit Mach number, pressure force), especially if the temperature T 0 of the combustion chamber is higher. We recall here that the form of the nozzle structure does not change, except the thermodynamic behaviour of the air which changes with T 0 . Two situations can be presented.

The first situation presented is that, if we wants to preserve the same variation of the Mach number throughout the nozzle, and consequently, the same exit Mach number M E , is necessary to determine by the application of our model, the ray of each section and in particular the ray of the exit section, which will give the same variation of the Mach number, and consequently another shape of the nozzle will be obtained.

MS(HT)=MS(PG)E41
MS(PG)=2H[TS(HT)]a[TS(HT)]E42
ASA*(HT)=eTS(HT)T*FA(T)dT>ASA*(PG)E43

The relation (36) indicates that the Mach number of the PG model is preserved for each section in our calculation. Initially, we determine the temperature at each section; witch presents the solution of equation (37). To determine the ratio of the sections, we use the relation (38). The ratio of the section obtained by our model will be superior that that determined by the PG model as present equation (38). Then the shape of the nozzle obtained by PG model is included in the nozzle obtained by our model. The temperature T 0 presented in equation (38) is that correspond to the temperature T 0 for our model.

The second situation consists to preserving the shape of the nozzle dimensioned on the basis of PG model for the aeronautical applications considered the HT model.

ASA*(HT)=ASA*(PG)E44
MS(HT)<MS(PG)E45

The relation (39) presents this situation. In this case, the nozzle will deliver a Mach number lower than desired, as shows the relation (40). The correction of the Mach number for HT model is initially made by the determination of the temperature T S as solution of equation (38), then determine the exit Mach number as solution of relation (37). The resolution of equation (38) is done by combining the dichotomy method with Simpson’s algorithm.

Figures 4 and 5 respectively represent the variation of specific heat C P (T) and the ratio γ(T) of the air versus the temperature up to 3550 K for HT and PG models. The graphs at high temperature are presented by using the polynomial interpolation (23). We can say that at low temperature until approximately 240 K, the gas can be regarded as calorically perfect, because of the invariance of specific heat C P (T) and the ratio γ(T). But if T 0 increases, we can see the difference between these values and it influences on the thermodynamic parameters of the flow. Figure 4.Variation of the specific heat for constant pressure versus stagnation temperature T 0 . Figure 5.Variation of the specific heats ratio versus T 0 .

### 5.1. Results for the critical parameters

Figures 6, 7 and 8 represent the variation of the critical thermodynamic ratios versus T 0 . It can be seen that with enhancement T 0 , the critical parameters vary, and this variation becomes considerable for high values of T 0 unlike to the PG model, where they do not depend on T 0.. For example, the value of the temperature ratio given by the HT model is always higher than the value given by the PG model. The ratios are determined by the choice of N=300000, b 1 =0.1 and b 2 =2.0 to have a precision better than ε=10 -5. The obtained numerical values of the critical parameters are presented in the table 3. Figure 6.Variation of T * /T 0 versus T 0 . Figure 7.Variation of ρ * /ρ 0 versus T 0 . Figure 8.Variation of P * /P 0 versus T 0 .

Figure 9 shows that mass flow rate through the critical cross section given by the perfect gas theory is lower than it is at the HT model, especially for values of T 0 . Figure 9.Variation of the non-dimensional critical mass flow rate with T 0 .

Figure 10 presents the variation of the critical sound velocity ratio versus T 0 . The influence of the T 0 on this parameter can be found.

 T*/T0 P*/P0 ρ*/ρ 0 a*/a 0 m/A* ρ 0 a 0 PG (γ=1.402) 0.8326 0.5279 0.6340 0.9124 0.5785 T0=298.15 K 0.8328 0.5279 0.6339 0.9131 0.5788 T0=500 K 0.8366 0.5293 0.6326 0.9171 0.5802 T0=1000 K 0.8535 0.5369 0.6291 0.9280 0.5838 T0=2000 K 0.8689 0.5448 0.6270 0.9343 0.5858 T0=2500 K 0.8722 0.5466 0.6266 0.9355 0.5862 T0=3000 K 0.8743 0.5475 0.6263 0.9365 0.5865 T0=3500 K 0.8758 0.5484 0.6262 0.9366 0.5865

### Table 3.

Numerical values of the critical parameters at high temperature.

### 5.2. Results for the supersonic parameters

Figures 11, 12 and 13 presents the variation of the supersonic flow parameters in a cross-section versus Mach number for T 0 =1000 K, 2000 K and 3000 K, including the case of perfect gas for γ=1.402. When M=1, we can obtain the values of the critical ratios. If we take into account the variation of C P (T), the temperature T 0 influences on the value of the thermodynamic and geometrical parameters of flow unlike the PG model.

The curve 4 of figure 11 is under the curves of the HT model, which indicates that the perfect gas model cool the flow compared to the real thermodynamic behaviour of the gas, and consequently, it influences on the dimensionless parameters of a nozzle. At low temperature and Mach number, the theory of perfect gas gives acceptable results. The obtained numerical values of the supersonic flow parameters, the cross section area ratio and sound velocity ratio are presented respectively if the tables 4, 5, 6, 7 and 8.

 T/T0 M=2.00 M=3.00 M=4.00 M=5.00 M=6.00 PG (γ=1.402) 0.5543 0.3560 0.2371 0.1659 0.1214 T0=298.15 K 0.5544 0.3560 0.2372 0.1659 0.1214 T0=500 K 0.5577 0.3581 0.2386 0.1669 0.1221 T0=1000 K 0.5810 0.3731 0.2481 0.1736 0.1269 T0=1500 K 0.6031 0.3911 0.2594 0.1810 0.1323 T0=2000 K 0.6163 0.4058 0.2694 0.1873 0.1366 T0=2500 K 0.6245 0.4162 0.2778 0.1928 0.1403 T0=3000 K 0.6301 0.4233 0.2848 0.1977 0.1473 T0=3500 K 0.6340 0.4285 0.2901 0.2018 0.1462

### Table 4.

Numerical values of the temperature ratio at high temperature

 ρ/ρ0 M=2.00 M=3.00 M=4.00 M=5.00 M=6.00 PG (γ=1.402) 0.2304 0.0765 0.0278 0.0114 0.0052 T0=298.15 K 0.2304 0.0765 0.0278 0.0114 0.0052 T0=500 K 0.2283 0.0758 0.0276 0.0113 0.0052 T0=1000 K 0.2181 0.0696 0.0250 0.0103 0.0047 T0=1500 K 0.2116 0.0636 0.0220 0.0089 0.0041 T0=2000 K 0.2087 0.0601 0.0197 0.0077 0.0035 T0=2500 K 0.2069 0.0581 0.0182 0.0069 0.0030 T0=3000 K 0.2057 0.0569 0.0173 0.0063 0.0027 T0=3500 K 0.2049 0.0560 0.0166 0.0058 0.0024

### Table 5.

Numerical values of the density ratio at high temperature

 P/P0 M=2.00 M=3.00 M=4.00 M=5.00 M=6.00 PG (γ=1.402) 0.1277 0.0272 0.0066 0.0019 0.0006 T0=298.15 K 0.1277 0.0272 0.0066 0.0019 0.0006 T0=500 K 0.1273 0.0271 0.0065 0.0018 0.0006 T0=1000 K 0.1267 0.0259 0.0062 0.0017 0.0006 T0=1500 K 0.1276 0.0248 0.0057 0.0016 0.0005 T0=2000 K 0.1286 0.0244 0.0053 0.0014 0.0004 T0=2500 K 0.1292 0.0242 0.0050 0.0013 0.0004 T0=3000 K 0.1296 0.0240 0.0049 0.0004 0.0003 T0=3500 K 0.1299 0.0240 0.0048 0.0011 0.0003

### Table 6.

Numerical values of the Pressure ratio at high temperature.

 A/A* M=2.00 M=3.00 M=4.00 M=5.00 M=6.00 PG (γ=1.402) 1.6859 4.2200 10.6470 24.7491 52.4769 T0=298.15 K 1.6859 4.2195 10.6444 24.7401 52.4516 T0=500 K 1.6916 4.2373 10.6895 24.8447 52.6735 T0=1000 K 1.7295 4.4739 11.3996 26.5019 56.1887 T0=1500 K 1.7582 4.7822 12.6397 29.7769 63.2133 T0=2000 K 1.7711 4.9930 13.8617 33.5860 72.0795 T0=2500 K 1.7795 5.1217 14.8227 37.2104 81.2941 T0=3000 K 1.7851 5.2091 15.5040 40.3844 90.4168 T0=3500 K 1.7889 5.2727 16.0098 43.0001 98.7953

### Table 7.

Numerical Values of the cross section area ratio at high temperature.

Figure 14 represent the variation of the critical cross-section area section ratio versus Mach number at high temperature. For low values of Mach number and T 0 , the four curves fuses and start to be differs when M>2.00. We can see that the curves 3 and 4 are almost superposed for any value of T 0 . This result shows that the PG model can be used for T 0 <1000 K.

Figure 15 presents the variation of the sound velocity ratio versus Mach number at high temperature. T 0 value influences on this parameter.

Figure 16 shows the variation of the thrust coefficient versus exit Mach number for various values of T 0 . It can be seen the effect of T 0 on this parameter. We can found that all the four curves are almost confounded when M E <2.00 approximately. After this value, the curves begin to separates progressively. The numerical values of the thrust coefficient are presented in the table 9. Figure 14.Variation of the critical cross-section area ratio versus Mach number.
 a/a0 M=2.00 M=3.00 M=4.00 M=5.00 M=6.00 PG (γ=1.402) 0.7445 0.5966 0.4870 0.4074 0.3484 T0=298.15 K 0.7450 0.5970 0.4873 0.4076 0.3486 T0=500 K 0.7510 0.6019 0.4913 0.4110 0.3515 T0=1000 K 0.7739 0.6245 0.5103 0.4268 0.3651 T0=1500 K 0.7862 0.6408 0.5254 0.4398 0.3762 T0=2000 K 0.7923 0.6501 0.5354 0.4489 0.3841 T0=2500 K 0.7959 0.6556 0.5420 0.4553 0.3898 T0=3000 K 0.7985 0.6595 0.5465 0.4600 0.3942 T0=3500 K 0.7998 0.6618 0.5495 0.4632 0.3973

### Table 8.

Numerical values of the sound velocity ratio at high temperature. Figure 15.Variation of the ratio of the velocity sound versus Mach number.
 CF M=2.00 M=3.00 M=4.00 M=5.00 M=6.00 PG (γ=1.402) 1.2078 1.4519 1.5802 1.6523 1.6959 T0=298.15 K 1.2078 1.4518 1.5800 1.6521 1.6957 T0=500 K 1.2076 1.4519 1.5802 1.6523 1.6958 T0=1000 K 1.2072 1.4613 1.5919 1.6646 1.7085 T0=1500 K 1.2062 1.4748 1.6123 1.6871 1.7317 T0=2000 K 1.2048 1.4832 1.6288 1.7069 1.7527 T0=2500 K 1.2042 1.4879 1.6401 1.7221 1.7694 T0=3000 K 1.2038 1.4912 1.6479 1.7337 1.7828 T0=3500 K 1.2033 1.4936 1.6533 1.7422 1.7932

### Table 9.

Numerical values of the thrust coefficient at high temperature

### 5.3. Results for the error given by the perfect gas model

Figure 17 presents the relative error of the thermodynamic and geometrical parameters between the PG and the HT models for several T 0 values.

It can be seen that the error depends on the values of T 0 and M. For example, if T 0 =2000 K and M=3.00, the use of the PG model will give a relative error equal to ε=14.27 % for the temperatures ratio, ε=27.30 % for the density ratio, error ε=15.48 % for the critical sections ratio and ε=2.11 % for the thrust coefficient. For lower values of M and T 0 , the error ε is weak. The curve 3 in the figure 17 is under the error 5% independently of the Mach number, which is interpreted by the use potential of the PG model when T 0 <1000 K.

We can deduce for the error given by the thrust coefficient that it is equal to ε=0.0 %, if M E =2.00 approximately independently of T 0 . There is no intersection of the three curves in the same time. When M E =2.00. Figure 17.Variation of the relative error given by supersonic parameters of PG versus Mach number.

### 5.4. Results for the supersonic nozzle application

Figure 18 presents the variation of the Mach number through the nozzle for T 0 =1000 K, 2000 K and 3000 K, including the case of perfect gas presented by curve 4. The example is selected for M S =3.00 for the PG model. If T 0 is taken into account, we will see a fall in Mach number of the dimensioned nozzle in comparison with the PG model. The more is the temperature T 0 , the more it is this fall. Consequently, the thermodynamics parameters force to design the nozzle with different dimensions than it is predicted by use the PG model. It should be noticed that the difference becomes considerable if the value T 0 exceeds 1000 K.

Figure 19 present the correction of the Mach number of nozzle giving exit Mach number M S , dimensioned on the basis of the PG model for various values of T 0 .

One can see that the curves confound until Mach number M S =2.0 for the whole range of T 0 . From this value, the difference between the three curves 1, 2 and 3, start to increase. The curves 3 and 4 are almost confounded whatever the Mach number if the value of T 0 is lower than 1000 K. For example, if the nozzle delivers a Mach number M S =3.00 at the exit section, on the assumption of the PG model, the HT model gives Mach number equal to M S =2.93, 2.84 and 2.81 for T 0 =1000 K, 2000 K and 3000 K respectively. The numerical values of the correction of the exit Mach number of the nozzle are presented in the table 10. Figure 18.Effect of stagnation temperature on the variation of the Mach number through the nozzle.
 MS (PG γ=1.402) 1.5 2 3 4 5 6 MS (T0=298.15 K) 1.4995 1.9995 2.9995 3.9993 4.9989 5.9985 MS (T0=500 K) 1.4977 1.9959 2.9956 3.9955 4.9951 5.9947 MS (T0=1000 K) 1.4879 1.9705 2.9398 3.9237 4.9145 5.904 MS (T0=1500 K) 1.483 1.9534 2.8777 3.8147 4.7727 5.7411 MS (T0=2000 K) 1.4807 1.9463 2.8432 3.7293 4.6372 5.5675 MS (T0=2500 K) 1.4792 1.9417 2.8245 3.6765 4.536 5.4209 MS (T0=3000 K) 1.4785 1.9388 2.8121 3.6454 4.4676 5.3066 MS (T0=3500 K) 1.4778 1.9368 2.8035 3.6241 4.4216 5.2237

### Table 10.

Correction of the exit Mach number of the nozzle.

Figure 20 presents the supersonic nozzles shapes delivering a same variation of the Mach number throughout the nozzle and consequently given the same exit Mach number M S =3.00. The variation of the Mach number through these 4 nozzles is illustrated on curve 4 of figure 18. The three other curves 1, 2, and, 3 of figure 15 are obtained with the HT model use for T 0 =3000 K, 2000 K and 1000 K respectively. The curve 4 of figure 20 is the same as it is in the figure 13a, and it is calculated with the PG model use. The nozzle that is calculated according to the PG model provides less cross-section area in comparison with the HT model. Figure 19.Correction of the Mach number at High Temperature of a nozzle dimensioned on the perfect gas model. Figure 20.Shapes of nozzles at high temperature corresponding to same Mach number variation througout the nozzle and given M S =3.00 at the exit.

## 6. Conclusion

From this study, we can quote the following points:

If we accept an error lower than 5%, we can study a supersonic flow using a perfect gas relations, if the stagnation temperature T 0 is lower than 1000 K for any value of Mach number, or when the Mach number is lower than 2.0 for any value of T 0 up to approximately 3000 K.

The PG model is represented by an explicit and simple relations, and do not request a high time to make calculation, unlike the proposed model, which requires the resolution of a nonlinear algebraic equations, and integration of two complex analytical functions. It takes more time for calculation and for data processing.

The basic variable for our model is the temperature and for the PG model is the Mach number because of a nonlinear implicit equation connecting the parameters T and M.

The relations presented in this study are valid for any interpolation chosen for the function C P (T). The essential one is that the selected interpolation gives small error.

We can choose another substance instead of the air. The relations remain valid, except that it is necessary to have the table of variation of CP and γ according to the temperature and to make a suitable interpolation.

The cross section area ratio presented by the relation (19) can be used as a source of comparison for verification of the dimensions calculation of various supersonic nozzles. It provides a uniform and parallel flow at the exit section by the method of characteristics and the Prandtl Meyer function (Zebbiche & Youbi, 2005a, 2005b, Zebbiche, 2007, Zebbiche, 2010a& Zebbiche, 2010b). The thermodynamic ratios can be used to determine the design parameters of the various shapes of nozzles under the basis of the HT model.

We can obtain the relations of a perfect gas starting from the relations of our model by annulling all constants of interpolation except the first. In this case, the PG model becomes a particular case of our model.

## Acknowledgments

The author acknowledges Djamel, Khaoula, Abdelghani Amine, Ritadj Zebbiche and Fettoum Mebrek for granting time to prepare this manuscript.

## How to cite and reference

### Cite this chapter Copy to clipboard

Toufik Zebbiche (November 2nd 2011). Effect of Stagnation Temperature on Supersonic Flow Parameters with Application for Air in Nozzles, Thermodynamics - Interaction Studies - Solids, Liquids and Gases, Juan Carlos Moreno-Pirajan, IntechOpen, DOI: 10.5772/19633. Available from:

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