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

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_Pis 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_Pand 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_Pand γversus 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 9^th 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).

$$\frac{C_P}{\gamma }dT-\frac{RT}{\rho }d\rho =0$$E1

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

$$\frac{d\rho }{\rho }=\frac{dT}{T\text{[}\gamma (T)-1\text{]}}$$E2

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

The sound velocity is (Ryhming, 1984),

$$a^2={\left(\frac{dP}{d\rho }\right)}_{entropy=cons\tan t}$$E3

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

$$\frac{dP}{d\rho }=\rho R\frac{dT}{d\rho }+RT$$E4

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

$$a^2\text{(}T\text{)}=\gamma \text{(}T\text{)}RT$$E5

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:

$$C_{P}dT+VdV=0$$E6

The integration between the stagnation state (V₀≈ 0, T₀) and supersonic state (V, T) gives:

$$V^2=2H\text{(}T\text{)}$$E7

Where

$$H(T)={\displaystyle {\int }_T^{T_0}{C}_P\left(T\right)dT}$$E8

Dividing the equation (6) by V²and substituting the relation (7) in the obtained result, we obtain:

$$\frac{dV}{V}=-\frac{C_P\text{(}T\text{)}}{2H\text{(}T\text{)}}dT$$E9

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

$$M(T)=\frac{\sqrt{2H\text{(}T\text{)}}}{a\text{(}T\text{)}}$$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):

$$VdV+\frac{dP}{\rho }=0$$E11

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

$$\frac{d\rho }{\rho }=F_{\rho }\text{(}T\text{)}dT$$E12

Where

$$F_{\rho }\text{(}T\text{)}=\frac{C_P\text{(}T\text{)}}{a^2\text{(}T\text{)}}$$E13

The density ratio relative to the temperature T₀can be obtained by integration of the function (13) between the stagnation state (ρ₀,T₀) and the concerned supersonic state (ρ,T):

$$\frac{\rho }{{\rho }_0}=Exp\left(-{\displaystyle {\int }_T^{T_0}{F}_{\rho }\left(T\right)dT}\right)$$E14

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

$$\frac{P}{P_0}=\left(\frac{\rho }{{\rho }_0}\right)\left(\frac{T}{T_0}\right)$$E15

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

$$\rho VA=cons\tan t$$E16

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

$$\frac{dA}{A}=F_A\text{(}T\text{)}dT$$E17

Where

$$F_A\text{(}T\text{)}=C_P\text{(}T\text{)}\left[\frac{1}{a^2\text{(}T\text{)}}-\frac{1}{2H\text{(}T\text{)}}\right]$$E18

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

$$\frac{A}{A_{*}}=Exp\left({\int }_T^{T_{*}}{F}_A\left(T\right)dT\right)$$E19

To find parameters ρand 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:

$$\frac{m}{A_{*}{\rho }_0{a}_0}={\displaystyle {\int }_{.A}\left(\frac{\rho }{{\rho }_0}\right)\left(\frac{a}{a_0}\right)M\cos \left(\theta \right)}\frac{dA}{A_{*}}$$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=0and A=A_*.Therefore, the relation (20) is reduced to:

$$\frac{\dot{m}}{A_{*}{\rho }_0{a}_0}=\left(\frac{{\rho }_{*}}{{\rho }_0}\right)\left(\frac{a_{*}}{a_0}\right)$$E21

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

$$\frac{a}{a_0}={\left[\frac{\gamma \text{(}T\text{)}}{\gamma \text{(}{T}_0\text{)}}\right]}^{1/2}{\left[\frac{T}{T_0}\right]}^{1/2}$$E22

The parameters T, P, ρand Afor 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 Mand T, where the reverse analytical expression does not exist.

3. Calculation procedure

In the first case, one presents the table of variation of C_P 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 γand C_Pvalues 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_Pvalues according to the temperature is presented by relation (23) in the form of Horner scheme to minimize the mathematical operations number (Zebbiche, 2010a, 2010b):

$$C_P\text{(}T\text{)}=a_1+T\text{(}{a}_2+T\text{(}{a}_3+T\text{(}{a}_4+T\text{(}{a}_5+T\text{(}{a}_6+T\text{(}{a}_7+T\text{(}{a}_8+T\text{(}{a}_9+T\text{(}{a}_{10}\text{)))))))))}$$E23

The interpolation (a_ii=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 than$$\text{<mover accent="true"><mi>T</mi><mo>¯</mo></mover>}=\text{240K}$$. So for this field, the table value (Peterson & Hill, 1965), was taken

$${\overline{C}}_P=C_p\left(\overline{T}\right)=1001.15868\text{J/(kgK})$$E24

Thus:

for$$T\le \overline{T}$$, we have $$C_P\text{(}T\text{)}={\overline{C}}_P$$for$$T>\overline{T}$$, relation (23) is used.

The selected interpolation gives an error less than ε=10^-3between 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, T₀].Then, H(T) is a function with a parameter T₀and it is defined when T≤T₀.

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

$$\begin{array}{l}H\text{(}T\text{)}=H_0-\\ \text{[}{c}_1+T\text{(}{c}_2+T\text{(}{c}_3+T\text{(}{c}_4+T\text{(}{c}_5+T{c}_6+T\text{(}{c}_7+T\text{(}{c}_8+T\text{(}{c}_9+T\text{(}{c}_{10}\text{)))))))))]}\end{array}$$E25

Where

$$\begin{array}{l}{H}_0=T_0\text{(}{c}_1+T_0\text{(}{c}_2+T_0\text{(}{c}_3+T_0\text{(}{c}_4+T_0\text{(}{c}_5+T_0\text{(}{c}_6+\\ T_0\text{(}{c}_7+T_0\text{(}{c}_8+T_0\text{(}{c}_9+T_0\text{(}{c}_{10}\text{))))))))))}\end{array}$$E26

and

$$c_i=\frac{a_i}{i}(i=\mathrm{1,}\mathrm{2,}\mathrm{3,}\mathrm{...,}10)$$E27

Figure 1.

Variation of functionF_ρ(T)in the interval [T_S,T₀] versusT₀.

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

For$$T_0<\overline{T}$$,

$$H\text{(}T\text{)}={\overline{C}}_P\left(T_0-T\right)$$E28

For$$T_0>\overline{T}$$,we have two cases:

$$\text{if}T>\overline{T}\text{:}H\text{(}T\text{)}=\text{relation(24)}$$E29

$$\text{if}T\le \overline{T}\text{:}H\text{(}T\text{)}={\overline{C}}_P\text{(}\overline{T}-T\text{)}+H\text{(}\overline{T}\text{)}$$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, T₀]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 functionF_A(T)in the interval [T_S,T_*] versusT₀

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 T₀=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_Sfor 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):

$$s_i=b_1{z}_i+\text{(}1-b_1\text{)}\left[1-\frac{\tanh \left[b_2\cdot \text{(}1-z_i\text{)}\right]}{\tanh \text{(}{b}_2\text{)}}\right]$$E31

Where

$$z_i=\frac{i-1}{N-1}1\le i\le N$$E32

Obtained s_ivalues, enable to find the value of T_iin nodes i:

$$T_i=s_i\text{(}{T}_D-T_G\text{)}+T_G$$E33

The temperature T_Dis equal to T₀for F_ρ(T), and equal to T_*for F_A(T).The temperature T_Gis equal to T_*for the critical parameter, and equal to T_Sfor the supersonic parameter. Taking a value b ₁near zero (b ₁=0.1, for example) and b ₂=2.0, it can condense the nodes towards left edge T_Sof the interval, see figure 3.

Figure 3.

Presentation of the condensation of nodes

3.1. Critical parameters

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:

$$2H(T_{*})-a^2(T_{*})=0$$E34

The resolution of equation (29) is made by the use of the dichotomy algorithm (Démidovitch & Maron, 1987& Zebbiche & Youbi, 2006), with T_*<T₀. It can choose the interval [T₁,T₂]containing T_*by T₁=0 Kand T₂=T₀. The value T_*can be given with a precision εif the interval of subdivision number Kis satisfied by the following condition:

$$K=1.4426Log\left(\frac{T_0}{\epsilon }\right)+1$$E35

If ε=10^-8 is taken, the number Kcannot exceed 39. Consequently, the temperature ratio T_*/T₀can 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_Scan be determined according to the Mach number M=M_S. Replacing T=T_Sand M=M_Sin relation (10) gives

$$2H\text{(}{T}_S\text{)}-M_S^2{a}^2\text{(}{T}_S\text{)}=0$$E36

The determination of T_Sof equation (31) is done always by the dichotomy algorithm, excepting T_S<T_*.We can take the interval [T₁,T₂]containing T_S, by (T₁=0 K,and T₂=T_*.

Replacing T=T_Sand ρ=ρ_Sin 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_Scan be calculated respectively by using the relations (15), (22) and (19) by replacing T=T_S, ρ=ρ_S, P=P_S, a=a_Sand A=A_S.

The integration results of the ratios ρ_*/ ρ₀, ρ_S/ρ₀and A_S/A_*primarily depend on the values of N, b₁and b₂.

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)

$$C_F=\frac{F}{P_0{A}_{*}}$$E37

Where

$$F=m{V}_E=m{M}_E{a}_E$$E38

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

$$C_F=\gamma \left(T_0\right)M_E\left(\frac{a_E}{a_0}\right)\left(\frac{{\rho }_{*}}{{\rho }_{*}}\right)\left(\frac{a_{*}}{a_0}\right)$$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_Pand ratio γas 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 PGmodel compared to our HTmodel can be calculated for each parameter. Then, for each value (T₀, M), the εerror can be evaluated by the following relationship:

$${\epsilon }_y\text{(}{T}_0,M\text{)}=\left|1-\frac{y_{PG}\text{(}{T}_0,M\text{)}}{y_{HT}\text{(}{T}_0,M\text{)}}\right|\times 100$$E40

The letter yin 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 PGmodel given a supersonic uniform Mach number M_Sat the exit section of rockets, degrades the desired performances (exit Mach number, pressure force), especially if the temperature T₀of 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 T₀. 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.

$$M_S\text{(}HT\text{)}=M_S\text{(}PG\text{)}$$E41

$$M_S\text{(}PG\text{)}=\frac{\sqrt{2H\text{[}{T}_{S\text{(}HT\text{)}}\text{]}}}{a\text{[}{T}_{S(HT)}\text{]}}$$E42

$$\frac{A_S}{A_{*}}\text{(}HT\text{)}=e^{{\displaystyle {\int }_{T_{S\text{(}HT\text{)}}}^{T_{*}}{F}_A\left(T\right)dT}}>\frac{A_S}{A_{*}}\text{(}PG\text{)}$$E43

The relation (36) indicates that the Mach number of the PGmodel 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 PGmodel as present equation (38). Then the shape of the nozzle obtained by PGmodel is included in the nozzle obtained by our model. The temperature T₀presented in equation (38) is that correspond to the temperature T₀for 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 HTmodel.

$$\frac{A_S}{A_{*}}\text{(}HT\text{)}=\frac{A_S}{A_{*}}\text{(}PG\text{)}$$E44

$$M_S\text{(}HT\text{)}<M_S\text{(}PG\text{)}$$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 HTmodel is initially made by the determination of the temperature T_Sas 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.

5. Results and comments

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 Kfor HTand PGmodels. 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 T₀increases, 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 temperatureT₀.

Figure 5.

Variation of the specific heats ratio versusT₀.

5.1. Results for the critical parameters

Figures 6, 7 and 8 represent the variation of the critical thermodynamic ratios versus T₀. It can be seen that with enhancement T₀,the critical parameters vary, and this variation becomes considerable for high values of T₀unlike to the PGmodel, where they do not depend on T_0.. For example, the value of the temperature ratio given by the HTmodel is always higher than the value given by the PGmodel. The ratios are determined by the choice of N=300000, b₁=0.1 and b₂=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 ofT_*/T₀versusT₀.

Figure 7.

Variation ofρ_*/ρ₀versusT₀.

Figure 8.

Variation ofP_*/P₀versusT₀.

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 HTmodel, especially for values of T₀.

Figure 9.

Variation of the non-dimensional critical mass flow rate withT₀.

Figure 10 presents the variation of the critical sound velocity ratio versus T₀. The influence of the T₀on this parameter can be found.

Figure 10.

Effect ofT₀on the velocity sound ratio.

	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 T₀=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 T₀influences on the value of the thermodynamic and geometrical parameters of flow unlike the PGmodel.

The curve 4 of figure 11 is under the curves of the HTmodel, 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

Figure 11.

Variation ofT/T₀versus Mach number.

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

Figure 12.

Variation ofρ/ρ₀versus Mach number.

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.

Figure 13.

Variation ofP/P₀versus Mach number.

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 T₀, 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 T₀. This result shows that the PGmodel can be used for T₀<1000 K.

Figure 15 presents the variation of the sound velocity ratio versus Mach number at high temperature. T₀value influences on this parameter.

Figure 16 shows the variation of the thrust coefficient versus exit Mach number for various values of T₀. It can be seen the effect of T₀on this parameter. We can found that all the four curves are almost confounded when M_E<2.00approximately. 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

Figure 16.

Variation ofC_Fversus exit Mach number.

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 PGand the HTmodels for several T₀values.

It can be seen that the error depends on the values of T₀and M. For example, if T₀=2000 K and M=3.00, the use of the PGmodel 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 Mand T₀, the error εis 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 PGmodel when T₀<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 T₀.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 ofPGversus Mach number.

5.4. Results for the supersonic nozzle application

Figure 18 presents the variation of the Mach number through the nozzle for T₀=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 PGmodel. If T₀is taken into account, we will see a fall in Mach number of the dimensioned nozzle in comparison with the PGmodel. The more is the temperature T₀, 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 PGmodel. It should be noticed that the difference becomes considerable if the value T₀exceeds 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 PGmodel for various values of T₀.

One can see that the curves confound until Mach number M_S=2.0for the whole range of T₀. 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 T₀is lower than 1000 K. For example, if the nozzle delivers a Mach number M_S=3.00at the exit section, on the assumption of the PGmodel, the HTmodel gives Mach number equal to M_S=2.93, 2.84 and 2.81 for T₀=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)	1.5000	2.0000	3.0000	4.0000	5.0000	6.0000
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	5.9040
MS (T0=1500 K)	1.4830	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	4.5360	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 HTmodel use for T₀=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 PGmodel use. The nozzle that is calculated according to the PGmodel provides less cross-section area in comparison with the HTmodel.

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 givenM_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 T₀is lower than 1000 K for any value of Mach number, or when the Mach number is lower than 2.0for any value of T₀up to approximately 3000 K.

The PGmodel 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 PGmodel is the Mach number because of a nonlinear implicit equation connecting the parameters Tand 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 C_P 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.